CA2769217A1 - Outer shell sector for a bladed ring for an aircraft turbomachine stator, including vibration damping shims - Google Patents
Outer shell sector for a bladed ring for an aircraft turbomachine stator, including vibration damping shims Download PDFInfo
- Publication number
- CA2769217A1 CA2769217A1 CA2769217A CA2769217A CA2769217A1 CA 2769217 A1 CA2769217 A1 CA 2769217A1 CA 2769217 A CA2769217 A CA 2769217A CA 2769217 A CA2769217 A CA 2769217A CA 2769217 A1 CA2769217 A1 CA 2769217A1
- Authority
- CA
- Canada
- Prior art keywords
- sector
- outer shell
- shim
- vibration damping
- elementary sectors
- 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.)
- Abandoned
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Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
- F04D—NON-POSITIVE-DISPLACEMENT PUMPS
- F04D29/00—Details, component parts, or accessories
- F04D29/40—Casings; Connections of working fluid
- F04D29/52—Casings; Connections of working fluid for axial pumps
- F04D29/54—Fluid-guiding means, e.g. diffusers
- F04D29/541—Specially adapted for elastic fluid pumps
- F04D29/542—Bladed diffusers
-
- 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
- F01D11/001—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between stator blade and rotor
-
- 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
- F01D25/04—Antivibration arrangements
- F01D25/06—Antivibration arrangements for preventing blade vibration
-
- 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
- F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
- F04D—NON-POSITIVE-DISPLACEMENT PUMPS
- F04D29/00—Details, component parts, or accessories
- F04D29/60—Mounting; Assembling; Disassembling
- F04D29/64—Mounting; Assembling; Disassembling of axial pumps
- F04D29/644—Mounting; Assembling; Disassembling of axial pumps especially adapted for elastic fluid pumps
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
- F04D—NON-POSITIVE-DISPLACEMENT PUMPS
- F04D29/00—Details, component parts, or accessories
- F04D29/66—Combating cavitation, whirls, noise, vibration or the like; Balancing
- F04D29/661—Combating cavitation, whirls, noise, vibration or the like; Balancing especially adapted for elastic fluid pumps
- F04D29/668—Combating cavitation, whirls, noise, vibration or the like; Balancing especially adapted for elastic fluid pumps damping or preventing mechanical vibrations
-
- 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
- F01D9/042—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector fixing blades to stators
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Structures Of Non-Positive Displacement Pumps (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
The invention relates to an assembly forming an outer shell sector (28) for a bladed ring sector (20) intended for a turbine or compressor stator of an aircraft turbine engine, comprising a plurality of elementary sectors (30) and vibration-damping blocks (34) disposed between the two elementary sectors associated therewith. According to the invention, the profile of each vibration-damping block (34) is substantially identical to that of the elementary sectors (30).
Description
OUTER SHELL SECTOR FOR A BLADED RING FOR AN AIRCRAFT
TURBOMACHINE STATOR, INCLUDING VIBRATION DAMPING SHIMS
DESCRIPTION
This invention generally relates to an aircraft turbomachine, preferably of the turbojet or turboprop type.
More particularly, the invention relates to the compressor or turbine stator of such a turbomachine, and more precisely to a bladed ring sector comprising a plurality of stator blades and two concentric shells supporting the blades and designed to radially delimit a primary flow passing through the turbomachine, inwards and outwards respectively. Such a bladed ring is usually made using several sectors arranged end to end, is usually used in the compressor or the turbine as a guide vane or a nozzle.
Turbomachines usually comprise a low pressure compressor, a high pressure compressor, a combustion chamber, a high pressure turbine and a low pressure turbine, in series. Compressors and turbines comprise several rows of mobile blades at a circumferential spacing, these rows being separated by rows of fixed blades also at a circumferential spacing.
In modern turbomachines, high dynamic stresses are applied to the guide vanes and nozzles. Technological progress leads to a reduction in the number of stages for equal or better performances, resulting in a higher load for each stage. Furthermore, changes to production technologies have led to a reduction in the number of parts, which reduces the damping effect of connections between parts. This is the case particularly when an abradable cartridge brazing technology is used which eliminates a large source of dissipation of vibration energy.
Document FR-A-2 902 843 discloses a means of solving this vibration problem by breaking the outer shell sector down into elementary sectors at a fixed spacing from each other along the tangential direction by the use of slits or radial cuts, oblique or in another direction, each elementary sector supporting a single blade of the bladed ring sector. Furthermore, damping inserts in the form of strips are inserted between the elementary sectors. The operating principle is based on the introduction of a stiffness non-linearity in the dynamic behaviour of the structure.
This non-linearity is triggered by a threshold vibration level of the system. This vibration activity causes a relative movement between the elementary sectors of the blades and the damping inserts. This relative movement causes successive loss and recovery of adhesion of the damping inserts and consequently a continuous variation of the local stiffness of the system. Consequently, the mode(s) causing the vibration activity are disorganised by the permanent variation of the associated natural frequencies. Resonance of the system cannot be set up because of the continuous variation in the state of the dynamic system. This reduces vibration amplitudes in the system.
TURBOMACHINE STATOR, INCLUDING VIBRATION DAMPING SHIMS
DESCRIPTION
This invention generally relates to an aircraft turbomachine, preferably of the turbojet or turboprop type.
More particularly, the invention relates to the compressor or turbine stator of such a turbomachine, and more precisely to a bladed ring sector comprising a plurality of stator blades and two concentric shells supporting the blades and designed to radially delimit a primary flow passing through the turbomachine, inwards and outwards respectively. Such a bladed ring is usually made using several sectors arranged end to end, is usually used in the compressor or the turbine as a guide vane or a nozzle.
Turbomachines usually comprise a low pressure compressor, a high pressure compressor, a combustion chamber, a high pressure turbine and a low pressure turbine, in series. Compressors and turbines comprise several rows of mobile blades at a circumferential spacing, these rows being separated by rows of fixed blades also at a circumferential spacing.
In modern turbomachines, high dynamic stresses are applied to the guide vanes and nozzles. Technological progress leads to a reduction in the number of stages for equal or better performances, resulting in a higher load for each stage. Furthermore, changes to production technologies have led to a reduction in the number of parts, which reduces the damping effect of connections between parts. This is the case particularly when an abradable cartridge brazing technology is used which eliminates a large source of dissipation of vibration energy.
Document FR-A-2 902 843 discloses a means of solving this vibration problem by breaking the outer shell sector down into elementary sectors at a fixed spacing from each other along the tangential direction by the use of slits or radial cuts, oblique or in another direction, each elementary sector supporting a single blade of the bladed ring sector. Furthermore, damping inserts in the form of strips are inserted between the elementary sectors. The operating principle is based on the introduction of a stiffness non-linearity in the dynamic behaviour of the structure.
This non-linearity is triggered by a threshold vibration level of the system. This vibration activity causes a relative movement between the elementary sectors of the blades and the damping inserts. This relative movement causes successive loss and recovery of adhesion of the damping inserts and consequently a continuous variation of the local stiffness of the system. Consequently, the mode(s) causing the vibration activity are disorganised by the permanent variation of the associated natural frequencies. Resonance of the system cannot be set up because of the continuous variation in the state of the dynamic system. This reduces vibration amplitudes in the system.
Nevertheless, even if this solution is satisfactory in terms of reducing vibrations, it can be improved. Furthermore, in this solution disclosed in document FR-A-2 902 843, the damping inserts are held in contact against the friction surfaces of the elementary sectors due to the effect of the pressure gradient between the aerodynamic flowpath and the outside of the compressor, applying a radially inwards force on these inserts. The disadvantage is that this pressure gradient cannot be sufficient to satisfactorily force the inserts into contact with the friction surfaces. In this case, the result is firstly a reduction in the vibration damping performances, but also a possible loss of leak tightness of the air flowpath.
Another disadvantage with this solution is the fact that one of the blades in the bladed ring sector will be overloaded. Aerodynamic forces applied on the blades include a tangential component that cannot be resisted in the outer shell sector, due to its segmentation into tangentially spaced elementary sectors. Thus, these tangential components are combined and pass through the inner shell sector of the bladed ring sector before passing through the blade located adjacent to the anti-rotation stop fitted on the ring sector. Therefore, this blade is very highly loaded due to the incapability of the outer shell sector to transmit static forces along the tangential direction.
Therefore, the purpose of the invention is to at least partially overcome the problems mentioned Y
above that arise with embodiments according to prior art.
The first purpose of the invention to achieve this is an assembly forming an outer shell sector for a bladed ring sector that will be used on a compressor or turbine stator in an aircraft turbomachine, said outer shell sector comprising firstly a plurality of elementary sectors at a spacing from each other along a tangential direction of said assembly, and secondly vibration damping shims, each of them being inserted between two elementary sectors associated with it, placed directly consecutively along said tangential direction.
According to the invention, the profile of each vibration damping shim is approximately the same as the profile of the elementary sectors.
Due to the particular profile of the shims, the friction interface between the shims and the elementary sectors is large which results in an improved damping effect.
Furthermore, the fact that the shims are forced into contact with the friction surfaces of the elementary sectors can result in a perfect seal between these elements, independent of the pressure difference between the aerodynamic flowpath and the outside of the compressor or the turbine. This seal is obtained by construction, with shims applying forces on the friction surfaces of the elementary sectors approximately along the tangential direction. Note that this seal is further reinforced during operation, because the forces bringing the friction surfaces and the shims into contact with each other are accentuated by application of the tangential component of aerodynamic forces applied on the stator blades, on the elementary sectors.
Another disadvantage with this solution is the fact that one of the blades in the bladed ring sector will be overloaded. Aerodynamic forces applied on the blades include a tangential component that cannot be resisted in the outer shell sector, due to its segmentation into tangentially spaced elementary sectors. Thus, these tangential components are combined and pass through the inner shell sector of the bladed ring sector before passing through the blade located adjacent to the anti-rotation stop fitted on the ring sector. Therefore, this blade is very highly loaded due to the incapability of the outer shell sector to transmit static forces along the tangential direction.
Therefore, the purpose of the invention is to at least partially overcome the problems mentioned Y
above that arise with embodiments according to prior art.
The first purpose of the invention to achieve this is an assembly forming an outer shell sector for a bladed ring sector that will be used on a compressor or turbine stator in an aircraft turbomachine, said outer shell sector comprising firstly a plurality of elementary sectors at a spacing from each other along a tangential direction of said assembly, and secondly vibration damping shims, each of them being inserted between two elementary sectors associated with it, placed directly consecutively along said tangential direction.
According to the invention, the profile of each vibration damping shim is approximately the same as the profile of the elementary sectors.
Due to the particular profile of the shims, the friction interface between the shims and the elementary sectors is large which results in an improved damping effect.
Furthermore, the fact that the shims are forced into contact with the friction surfaces of the elementary sectors can result in a perfect seal between these elements, independent of the pressure difference between the aerodynamic flowpath and the outside of the compressor or the turbine. This seal is obtained by construction, with shims applying forces on the friction surfaces of the elementary sectors approximately along the tangential direction. Note that this seal is further reinforced during operation, because the forces bringing the friction surfaces and the shims into contact with each other are accentuated by application of the tangential component of aerodynamic forces applied on the stator blades, on the elementary sectors.
5 Concerning the tangential component of the aerodynamic forces applied on the blades, note that one of the essential advantages of this invention lies in the fact that this component can transit through the assembly forming an outer shell sector because the outer shell sector is very much stiffened along the tangential direction due to the particular positioning of vibration damping shims, even though it is separated into sectors along this direction. The result is that there is no overload of the blades that are therefore loaded approximately uniformly.
Finally, note that by adopting approximately the same profile as the profile of the elementary sectors, the outer radial delimitation of the primary annular flow, also called the air flowpath, is perfectly recreated between the elementary sectors at a spacing from each other.
Preferably, said shim bears in contact with two parallel plane friction surfaces facing each other along said tangential direction and provided on said two elementary sectors associated with said shim, and said shim has two complementary plane friction surfaces parallel to each other and cooperating with the two corresponding friction surfaces of the elementary sectors. The plane contacts between the friction surfaces and the complementary friction surfaces give satisfactory damping of vibrations by friction. It is also possible to make the two friction surfaces simultaneously during a single machining operation, for example by a single cutting operation, in order to obtain straight slits, in other words slits in a determined plane, inside which the corresponding shims will subsequently be housed. This makes it very much simpler to fabricate the assembly according to the invention, which results in a significant cost and time saving.
Preferably, said shim is provided with hooks to hold it in place on the compressor or turbine stator, therefore these hooks have the same profile as the hooks fixed on the elementary sectors.
Preferably, the elementary sectors are separated from each other by radial slits completely filled in by said vibration damping shims.
Preferably, said vibration damping shims extend approximately along an axial or oblique direction of said assembly.
Another purpose of this invention applies to a bladed ring sector designed to be installed on a compressor or turbine stator of an aircraft turbomachine comprising an assembly forming an outer shell sector like that described above, an inner shell sector and a plurality of blades at a tangential spacing from each other and inserted between the assembly forming the outer shell sector and the inner shell sector. In this case, each elementary sector will carry a single stator blade, or possibly several blades, without going outside the scope of the invention.
Finally, note that by adopting approximately the same profile as the profile of the elementary sectors, the outer radial delimitation of the primary annular flow, also called the air flowpath, is perfectly recreated between the elementary sectors at a spacing from each other.
Preferably, said shim bears in contact with two parallel plane friction surfaces facing each other along said tangential direction and provided on said two elementary sectors associated with said shim, and said shim has two complementary plane friction surfaces parallel to each other and cooperating with the two corresponding friction surfaces of the elementary sectors. The plane contacts between the friction surfaces and the complementary friction surfaces give satisfactory damping of vibrations by friction. It is also possible to make the two friction surfaces simultaneously during a single machining operation, for example by a single cutting operation, in order to obtain straight slits, in other words slits in a determined plane, inside which the corresponding shims will subsequently be housed. This makes it very much simpler to fabricate the assembly according to the invention, which results in a significant cost and time saving.
Preferably, said shim is provided with hooks to hold it in place on the compressor or turbine stator, therefore these hooks have the same profile as the hooks fixed on the elementary sectors.
Preferably, the elementary sectors are separated from each other by radial slits completely filled in by said vibration damping shims.
Preferably, said vibration damping shims extend approximately along an axial or oblique direction of said assembly.
Another purpose of this invention applies to a bladed ring sector designed to be installed on a compressor or turbine stator of an aircraft turbomachine comprising an assembly forming an outer shell sector like that described above, an inner shell sector and a plurality of blades at a tangential spacing from each other and inserted between the assembly forming the outer shell sector and the inner shell sector. In this case, each elementary sector will carry a single stator blade, or possibly several blades, without going outside the scope of the invention.
r The bladed ring may form a guide vane of a compressor or a nozzle of a turbine.
Furthermore, the ring sector preferably extends around an angular range of between 5 and 60 , but can be as much as 360 so as to form the entire bladed ring.
Another purpose of the invention is an aircraft turbomachine comprising a compressor or turbine stator equipped with at least one bladed ring sector like that described above.
Other advantages and characteristics of the invention will appear in the detailed non-limitative description given below.
This description will be made with reference to the appended drawings among which;
- figure 1 shows a diagrammatic sectional view of a turbomachine that will be equipped with one or several bladed ring sectors according to this invention;
- figure 2 shows a sectional view representing part of the high pressure compressor of the turbomachine shown in figure 1, and including a bladed ring sector according to this invention;
figure 3 shows a perspective view of the bladed ring sector shown in the previous figure, the sector being in the form of a preferred embodiment of this invention;
- figure 4 shows an axial view of part of the bladed ring sector shown in the previous figure;
- figure 5 shows a profile view of the shims and the elementary sectors of the bladed ring sector shown in the previous figures, along line V-V in figure 4; and - figures 6a to 6c show views diagrammatically showing the different steps in a fabrication process of the bladed ring sector shown in the previous figures.
With reference firstly to figure 1, the figure shows an aircraft turbojet 100 to which the invention is applicable. It comprises, in order along the upstream to downstream direction, a low pressure compressor 2, a high pressure compressor 4, an annular combustion chamber 6, a high pressure turbine 8 and a low pressure turbine 10.
Figure 2 shows part of the high pressure compressor 4. In a known manner, the compressor comprises rows 14 of stator blades and rows 16 of rotor blades alternating on an axial direction parallel to the axis 12 of the compressor. The stator blades 18 distributed circumferentially/tangentially around the axis 12, are included in a part of the stator called the bladed ring 20, preferably constructed in sectors along the circumferential direction 22. Thus, in the following we will refer to a bladed ring sector 20, it being understood that this sector 20 preferably extends over an angular range of between 5 and 60 , but possibly as much as 360 so as to form the entire bladed ring.
The sector 20, therefore forming all or part of a turbine nozzle or a compressor guide vane, comprises an inner shell sector 24 forming the inner surface radially delimiting a primary annular flow 26 passing through the turbomachine, this shell sector 24 supporting the fixed roots of the stator blades 18. In addition to these blades 18, the sector 20 also comprises an assembly forming an outer shell sector 28 forming the outer surface radially delimiting the primary annular flow, and supporting the fixed heads of the blades 18.
In this respect, note that the sector 20 also comprises known additional elements fitted on the shell sector 24, such as a radially internal abradable coating 29 forming the annular sealing track contacted by a sealing device 31 supported by the rotor stage 16 supporting the rotating blades and arranged on the downstream side of the sector 20 concerned. The rotating sealing device 31 is a known labyrinth or lip seal type sealing device.
Figure 3 shows the bladed ring sector 20.
In the preferred embodiment described, the entire turbine nozzle or compressor guide vane is obtained by end to end placement of a plurality of these sectors 20, therefore each forming an angular or circumferential portion of this bladed ring. The angular sectors 20 (only one of which can be seen in figure 3) are preferably deprived of any rigid direct mechanical links connecting them to each other, their adjacent ends being simply placed facing each other with or without clearance.
More specifically with reference to figures 3 and 4, the figures show that the inner ring sector 24 is made in a single part and is not segmented. On the other hand, the assembly 28 forming the outer shell sector 28 is segmented into elementary sectors 30 at a spacing from each other along the tangential direction 22, by straight radial or slightly oblique slits 32, therefore creating clearances between the directly 5 consecutive sectors 30. Each slit 32 is made along a median straight line between two directly consecutive blades 18, such that each elementary sector 30 supports a single fixed stator blade 18. One of the two elementary sectors 30 located at the ends of the sector 10 20 supports a rotation stop 33 projecting radially outwards and that will cooperate with another part of the compressor stator in a known manner.
The assembly 28 also comprises vibration damping shims 34 housed between directly consecutive elementary sectors 30.
More precisely, each vibration damping shim 34 is housed between two plane parallel friction surfaces 38 facing each other along the tangential direction 22, and provided on the corresponding tangential ends facing each other on the two elementary sectors associated with the shim. Similarly, each shim 34 has two complementary plane friction surfaces 40 parallel to each other and also parallel and in contact with the two corresponding plane friction surfaces 38 with which they cooperate.
Therefore, each shim 34 is squeezed between two directly consecutive elementary sectors 30, having a shape complementary with the shape of the friction surfaces 38.
The contact between the two friction surfaces 38, 40 of each pair is preferably obtained as ti soon as the shim 34 is put into position between its two associated elementary sectors 30. The shims 34 thus apply forces oriented approximately along the tangential direction in contact with the friction surfaces 38 of the elementary sectors, with their complementary plane friction surfaces 40. These forces are advantageously increased during operation by the additional application of the tangential component of aerodynamic forces applied on the stator blades, on the elementary sectors.
As shown diagrammatically in figure 5, one of the special features of this invention lies in the fact that the profile of the shims 34 is approximately the same as the profile of the elementary sectors, this same profile corresponding to the profile of the outer shell sector. In this disclosure, profile refers to the global shape of the element seen along the tangential direction 22, although a sectional view is shown in figure 5.
Thus, the lower surface 46 of each shim 34, like the elementary sectors 30, acts as the outer radial delimitation of the air flowpath. Consequently, the global annular delimitation surface of the air flowpath composed of the sequence of these surfaces 46 formed on the shims 34 and the sectors 30, is approximately continuous from an aerodynamic point of view because there is no step between the successive surfaces 46.
Each shim 34 and each sector 30 also comprises hooks to hold it in place on another part of the compressor stator, and more precisely a fixing hook 48 projecting forwards, and a fixing hook 50 projecting backwards. As shown in figure 2, the hooks 48, 50 are housed in the corresponding annular slits 52, 54 provided in another part of the compressor stator, to fix the sector 20 onto this other part of the stator.
The shims 34, entirely filling in the slits 32, perform a vibration damping function by friction in contact with the friction surfaces 38, based on the physical principle described above for the shims disclosed in document FR-A-2 902 843. They also perform a seal function, and a function to allow the tangential component of aerodynamic forces applied on the stator blades to pass through. More generally in this respect, each shim 34 is capable of transmitting tangential forces between the two elementary sectors 30 between which it is inserted.
The natures of the materials used for the elementary sectors 30 and for the shims 34 are approximately the same, preferably metallic, and are chosen such that the shims wear preferentially rather than the elementary sectors 30.
Note also that the ratio between the extent of each shim and the extent of each elementary sector along the tangential direction that also correspond to the thicknesses, is between 0.5 and 1.
Figures 6a to 6c diagrammatically show a process for fabrication of the bladed ring sector 20.
Firstly as can be seen in figure 6a, a single-piece assembly 100 is made by pouring or machining forming the inner shell sector 24, the outer shell sector 28 and the stator blades 18. The next step is to make straight radial slits 32 on the outer shell sector 28 so as to obtain the elementary sectors 30 as shown diagrammatically in figure 6b, by simple and inexpensive machining. For example, these slits 32 can be made simply by cutting the sector 28.
Finally, figure 6c shows the final step that consists of putting the vibration damping shims 34 into position in the slits 32 forming the friction surfaces, simply by sliding the shims into their corresponding holes.
Note that a precise sliding adjustment clearance is preferred to make it relatively easy to insert of each shim in its associated slit while holding this shim in its slit solely by the squeezing force between the two friction surfaces 38.
Obviously, those skilled in the art could make various modifications to the invention as described above, solely using non-limitative examples.
Furthermore, the ring sector preferably extends around an angular range of between 5 and 60 , but can be as much as 360 so as to form the entire bladed ring.
Another purpose of the invention is an aircraft turbomachine comprising a compressor or turbine stator equipped with at least one bladed ring sector like that described above.
Other advantages and characteristics of the invention will appear in the detailed non-limitative description given below.
This description will be made with reference to the appended drawings among which;
- figure 1 shows a diagrammatic sectional view of a turbomachine that will be equipped with one or several bladed ring sectors according to this invention;
- figure 2 shows a sectional view representing part of the high pressure compressor of the turbomachine shown in figure 1, and including a bladed ring sector according to this invention;
figure 3 shows a perspective view of the bladed ring sector shown in the previous figure, the sector being in the form of a preferred embodiment of this invention;
- figure 4 shows an axial view of part of the bladed ring sector shown in the previous figure;
- figure 5 shows a profile view of the shims and the elementary sectors of the bladed ring sector shown in the previous figures, along line V-V in figure 4; and - figures 6a to 6c show views diagrammatically showing the different steps in a fabrication process of the bladed ring sector shown in the previous figures.
With reference firstly to figure 1, the figure shows an aircraft turbojet 100 to which the invention is applicable. It comprises, in order along the upstream to downstream direction, a low pressure compressor 2, a high pressure compressor 4, an annular combustion chamber 6, a high pressure turbine 8 and a low pressure turbine 10.
Figure 2 shows part of the high pressure compressor 4. In a known manner, the compressor comprises rows 14 of stator blades and rows 16 of rotor blades alternating on an axial direction parallel to the axis 12 of the compressor. The stator blades 18 distributed circumferentially/tangentially around the axis 12, are included in a part of the stator called the bladed ring 20, preferably constructed in sectors along the circumferential direction 22. Thus, in the following we will refer to a bladed ring sector 20, it being understood that this sector 20 preferably extends over an angular range of between 5 and 60 , but possibly as much as 360 so as to form the entire bladed ring.
The sector 20, therefore forming all or part of a turbine nozzle or a compressor guide vane, comprises an inner shell sector 24 forming the inner surface radially delimiting a primary annular flow 26 passing through the turbomachine, this shell sector 24 supporting the fixed roots of the stator blades 18. In addition to these blades 18, the sector 20 also comprises an assembly forming an outer shell sector 28 forming the outer surface radially delimiting the primary annular flow, and supporting the fixed heads of the blades 18.
In this respect, note that the sector 20 also comprises known additional elements fitted on the shell sector 24, such as a radially internal abradable coating 29 forming the annular sealing track contacted by a sealing device 31 supported by the rotor stage 16 supporting the rotating blades and arranged on the downstream side of the sector 20 concerned. The rotating sealing device 31 is a known labyrinth or lip seal type sealing device.
Figure 3 shows the bladed ring sector 20.
In the preferred embodiment described, the entire turbine nozzle or compressor guide vane is obtained by end to end placement of a plurality of these sectors 20, therefore each forming an angular or circumferential portion of this bladed ring. The angular sectors 20 (only one of which can be seen in figure 3) are preferably deprived of any rigid direct mechanical links connecting them to each other, their adjacent ends being simply placed facing each other with or without clearance.
More specifically with reference to figures 3 and 4, the figures show that the inner ring sector 24 is made in a single part and is not segmented. On the other hand, the assembly 28 forming the outer shell sector 28 is segmented into elementary sectors 30 at a spacing from each other along the tangential direction 22, by straight radial or slightly oblique slits 32, therefore creating clearances between the directly 5 consecutive sectors 30. Each slit 32 is made along a median straight line between two directly consecutive blades 18, such that each elementary sector 30 supports a single fixed stator blade 18. One of the two elementary sectors 30 located at the ends of the sector 10 20 supports a rotation stop 33 projecting radially outwards and that will cooperate with another part of the compressor stator in a known manner.
The assembly 28 also comprises vibration damping shims 34 housed between directly consecutive elementary sectors 30.
More precisely, each vibration damping shim 34 is housed between two plane parallel friction surfaces 38 facing each other along the tangential direction 22, and provided on the corresponding tangential ends facing each other on the two elementary sectors associated with the shim. Similarly, each shim 34 has two complementary plane friction surfaces 40 parallel to each other and also parallel and in contact with the two corresponding plane friction surfaces 38 with which they cooperate.
Therefore, each shim 34 is squeezed between two directly consecutive elementary sectors 30, having a shape complementary with the shape of the friction surfaces 38.
The contact between the two friction surfaces 38, 40 of each pair is preferably obtained as ti soon as the shim 34 is put into position between its two associated elementary sectors 30. The shims 34 thus apply forces oriented approximately along the tangential direction in contact with the friction surfaces 38 of the elementary sectors, with their complementary plane friction surfaces 40. These forces are advantageously increased during operation by the additional application of the tangential component of aerodynamic forces applied on the stator blades, on the elementary sectors.
As shown diagrammatically in figure 5, one of the special features of this invention lies in the fact that the profile of the shims 34 is approximately the same as the profile of the elementary sectors, this same profile corresponding to the profile of the outer shell sector. In this disclosure, profile refers to the global shape of the element seen along the tangential direction 22, although a sectional view is shown in figure 5.
Thus, the lower surface 46 of each shim 34, like the elementary sectors 30, acts as the outer radial delimitation of the air flowpath. Consequently, the global annular delimitation surface of the air flowpath composed of the sequence of these surfaces 46 formed on the shims 34 and the sectors 30, is approximately continuous from an aerodynamic point of view because there is no step between the successive surfaces 46.
Each shim 34 and each sector 30 also comprises hooks to hold it in place on another part of the compressor stator, and more precisely a fixing hook 48 projecting forwards, and a fixing hook 50 projecting backwards. As shown in figure 2, the hooks 48, 50 are housed in the corresponding annular slits 52, 54 provided in another part of the compressor stator, to fix the sector 20 onto this other part of the stator.
The shims 34, entirely filling in the slits 32, perform a vibration damping function by friction in contact with the friction surfaces 38, based on the physical principle described above for the shims disclosed in document FR-A-2 902 843. They also perform a seal function, and a function to allow the tangential component of aerodynamic forces applied on the stator blades to pass through. More generally in this respect, each shim 34 is capable of transmitting tangential forces between the two elementary sectors 30 between which it is inserted.
The natures of the materials used for the elementary sectors 30 and for the shims 34 are approximately the same, preferably metallic, and are chosen such that the shims wear preferentially rather than the elementary sectors 30.
Note also that the ratio between the extent of each shim and the extent of each elementary sector along the tangential direction that also correspond to the thicknesses, is between 0.5 and 1.
Figures 6a to 6c diagrammatically show a process for fabrication of the bladed ring sector 20.
Firstly as can be seen in figure 6a, a single-piece assembly 100 is made by pouring or machining forming the inner shell sector 24, the outer shell sector 28 and the stator blades 18. The next step is to make straight radial slits 32 on the outer shell sector 28 so as to obtain the elementary sectors 30 as shown diagrammatically in figure 6b, by simple and inexpensive machining. For example, these slits 32 can be made simply by cutting the sector 28.
Finally, figure 6c shows the final step that consists of putting the vibration damping shims 34 into position in the slits 32 forming the friction surfaces, simply by sliding the shims into their corresponding holes.
Note that a precise sliding adjustment clearance is preferred to make it relatively easy to insert of each shim in its associated slit while holding this shim in its slit solely by the squeezing force between the two friction surfaces 38.
Obviously, those skilled in the art could make various modifications to the invention as described above, solely using non-limitative examples.
Claims (7)
1. Assembly forming an outer shell sector (28), for a bladed ring sector (20) designed to be used on a compressor stator in an aircraft turbomachine, said outer shell sector comprising firstly a plurality of elementary sectors (30) at a spacing from each other along a tangential direction (22) of said assembly, and secondly vibration damping shims (34) each of them being inserted between two elementary sectors associated with it, placed directly consecutively along said tangential direction, characterised in that the profile of each vibration damping shim (34) is approximately the same as the profile of the elementary sectors (30).
2. Assembly according to claim 1, characterised in that said shim is forced in contact with two parallel plane friction surfaces (38) facing each other along said tangential direction (22) and provided on said two elementary sectors (30) associated with said shim, and in that said shim (34) has two complementary plane friction surfaces (40), parallel to each other and cooperating with the two corresponding friction surfaces of the elementary sectors.
3. Assembly according to claim 1 or claim 2, characterised in that said shim (34) is provided with hooks (48, 50) to hold it in place on the compressor or turbine stator.
4. Assembly according to any one of the previous claims, characterised in that the elementary sectors (30) are separated from each other by radial slits (32) completely filled in by said vibration damping shims (34).
5. Assembly according to any one of the previous claims, characterised in that said vibration damping shims (34) extend approximately along an axial or oblique direction of said assembly.
6. Bladed ring sector (20) designed to be installed on a compressor stator of an aircraft turbomachine, comprising an assembly forming an outer shell sector (28) according to any one of the previous claims, an inner shell sector (24), and a plurality of blades (18) at a tangential spacing from each other and inserted between the assembly forming the outer shell sector and the inner shell sector.
7. Aircraft turbomachine comprising a compressor stator equipped with at least one bladed ring sector according to claim 6.
Applications Claiming Priority (3)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| FR0955439A FR2948736B1 (en) | 2009-07-31 | 2009-07-31 | EXTERNAL VIROLE SECTOR FOR AIRBORNE TURBOMACHINE AIRBORNE STATOR CROWN, COMPRISING SHOCK ABSORBING MOUNTS |
| FR0955439 | 2009-07-31 | ||
| PCT/EP2010/061037 WO2011012679A2 (en) | 2009-07-31 | 2010-07-29 | Outer shell sector for a bladed stator ring of an aircraft turbine engine, comprising vibration-damping blocks |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CA2769217A1 true CA2769217A1 (en) | 2011-02-03 |
Family
ID=41800367
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CA2769217A Abandoned CA2769217A1 (en) | 2009-07-31 | 2010-07-29 | Outer shell sector for a bladed ring for an aircraft turbomachine stator, including vibration damping shims |
Country Status (9)
| Country | Link |
|---|---|
| US (1) | US20120128482A1 (en) |
| EP (1) | EP2459884B1 (en) |
| JP (1) | JP5697667B2 (en) |
| CN (1) | CN102472297A (en) |
| BR (1) | BR112012002304A2 (en) |
| CA (1) | CA2769217A1 (en) |
| FR (1) | FR2948736B1 (en) |
| RU (1) | RU2537997C2 (en) |
| WO (1) | WO2011012679A2 (en) |
Families Citing this family (16)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| FR2971022B1 (en) | 2011-02-02 | 2013-01-04 | Snecma | COMPRESSOR RECTIFIER STAGE FOR A TURBOMACHINE |
| US9610644B2 (en) * | 2011-02-08 | 2017-04-04 | United Technologies Corporation | Mate face brazing for turbine components |
| US9546557B2 (en) * | 2012-06-29 | 2017-01-17 | General Electric Company | Nozzle, a nozzle hanger, and a ceramic to metal attachment system |
| US9334756B2 (en) | 2012-09-28 | 2016-05-10 | United Technologies Corporation | Liner and method of assembly |
| USRE48980E1 (en) | 2013-03-15 | 2022-03-22 | Raytheon Technologies Corporation | Acoustic liner with varied properties |
| DE102013212252A1 (en) * | 2013-06-26 | 2014-12-31 | Siemens Aktiengesellschaft | Turbine and method of squeal detection |
| FR3008455B1 (en) * | 2013-07-09 | 2015-08-21 | Snecma | COMPRESSOR RECTIFIER HAVING GAME RETRIEVAL MEANS |
| CN104440153B (en) * | 2014-11-04 | 2017-06-06 | 中国南方航空工业(集团)有限公司 | Casing intra vane processes damping unit |
| FR3029242B1 (en) | 2014-11-28 | 2016-12-30 | Snecma | TURBOMACHINE TURBINE, COMPRISING INTERCROSSED PARTITIONS FOR AIR CIRCULATION IN DIRECTION OF THE LEAK EDGE |
| US10655482B2 (en) * | 2015-02-05 | 2020-05-19 | Rolls-Royce Corporation | Vane assemblies for gas turbine engines |
| JP6689117B2 (en) * | 2016-03-31 | 2020-04-28 | 三菱日立パワーシステムズ株式会社 | Stator blade ring and axial flow rotary machine equipped in the axial flow rotary machine |
| CN106988794B (en) * | 2017-06-02 | 2018-12-14 | 中国航发南方工业有限公司 | Stator sub-assembly clamping means and stator sub-assembly |
| CN107747563B (en) * | 2017-09-30 | 2020-04-10 | 中国航发沈阳发动机研究所 | Fan casing with damping |
| US11242762B2 (en) * | 2019-11-21 | 2022-02-08 | Raytheon Technologies Corporation | Vane with collar |
| FR3115819B1 (en) * | 2020-11-02 | 2023-04-14 | Safran Aircraft Engines | Aircraft turbomachine stator assembly, comprising an external structure formed of two annular sections surrounding a bladed stator crown |
| FR3119196B1 (en) * | 2021-01-27 | 2022-12-23 | Safran Aircraft Engines | Sectorized annular row of fixed vanes |
Family Cites Families (13)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US2661147A (en) * | 1949-01-19 | 1953-12-01 | Ingersoll Rand Co | Blower blade fastening device |
| SU453486A1 (en) * | 1973-04-11 | 1974-12-15 | DEVICE FOR DAMPING THE OSCILLATIONS OF WORK BLADDES OF AXIAL TURBO DUMPERS | |
| JPS5239807A (en) * | 1975-09-25 | 1977-03-28 | Mitsubishi Heavy Ind Ltd | Moving vane vibration controlling apparatus |
| US5201850A (en) * | 1991-02-15 | 1993-04-13 | General Electric Company | Rotor tip shroud damper including damper wires |
| DE4436731A1 (en) * | 1994-10-14 | 1996-04-18 | Abb Management Ag | compressor |
| FR2831615B1 (en) * | 2001-10-31 | 2004-01-02 | Snecma Moteurs | SECTORIZED FIXED RECTIFIER FOR A TURBOMACHINE COMPRESSOR |
| US6984108B2 (en) * | 2002-02-22 | 2006-01-10 | Drs Power Technology Inc. | Compressor stator vane |
| US6733237B2 (en) * | 2002-04-02 | 2004-05-11 | Watson Cogeneration Company | Method and apparatus for mounting stator blades in axial flow compressors |
| EP1510654A1 (en) * | 2003-08-25 | 2005-03-02 | Siemens Aktiengesellschaft | Unitary turbine blade array and method to produce the unitary turbine blade array. |
| US7104752B2 (en) * | 2004-10-28 | 2006-09-12 | Florida Turbine Technologies, Inc. | Braided wire damper for segmented stator/rotor and method |
| FR2902843A1 (en) * | 2006-06-23 | 2007-12-28 | Snecma Sa | COMPRESSOR RECTIFIER AREA OR TURBOMACHINE DISTRIBUTOR SECTOR |
| US7591634B2 (en) * | 2006-11-21 | 2009-09-22 | General Electric Company | Stator shim welding |
| US7806655B2 (en) * | 2007-02-27 | 2010-10-05 | General Electric Company | Method and apparatus for assembling blade shims |
-
2009
- 2009-07-31 FR FR0955439A patent/FR2948736B1/en active Active
-
2010
- 2010-07-29 WO PCT/EP2010/061037 patent/WO2011012679A2/en active Application Filing
- 2010-07-29 BR BR112012002304A patent/BR112012002304A2/en not_active IP Right Cessation
- 2010-07-29 CA CA2769217A patent/CA2769217A1/en not_active Abandoned
- 2010-07-29 EP EP10739591.5A patent/EP2459884B1/en active Active
- 2010-07-29 CN CN2010800340319A patent/CN102472297A/en active Pending
- 2010-07-29 US US13/386,496 patent/US20120128482A1/en not_active Abandoned
- 2010-07-29 RU RU2012107522/06A patent/RU2537997C2/en not_active IP Right Cessation
- 2010-07-29 JP JP2012522172A patent/JP5697667B2/en not_active Expired - Fee Related
Also Published As
| Publication number | Publication date |
|---|---|
| RU2012107522A (en) | 2013-09-10 |
| FR2948736B1 (en) | 2011-09-23 |
| BR112012002304A2 (en) | 2016-05-31 |
| WO2011012679A2 (en) | 2011-02-03 |
| EP2459884A2 (en) | 2012-06-06 |
| JP2013501181A (en) | 2013-01-10 |
| JP5697667B2 (en) | 2015-04-08 |
| WO2011012679A3 (en) | 2011-04-21 |
| EP2459884B1 (en) | 2018-06-27 |
| US20120128482A1 (en) | 2012-05-24 |
| FR2948736A1 (en) | 2011-02-04 |
| CN102472297A (en) | 2012-05-23 |
| RU2537997C2 (en) | 2015-01-10 |
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Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| FZDE | Discontinued |
Effective date: 20160729 |