US20090016890A1 - Turbomachine rotor assembly - Google Patents
Turbomachine rotor assembly Download PDFInfo
- Publication number
- US20090016890A1 US20090016890A1 US12/171,775 US17177508A US2009016890A1 US 20090016890 A1 US20090016890 A1 US 20090016890A1 US 17177508 A US17177508 A US 17177508A US 2009016890 A1 US2009016890 A1 US 2009016890A1
- Authority
- US
- United States
- Prior art keywords
- disk
- shim
- rotor assembly
- operating temperature
- bearing surfaces
- 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
Links
- 239000000463 material Substances 0.000 claims abstract description 30
- 239000002131 composite material Substances 0.000 claims description 4
- 239000011159 matrix material Substances 0.000 claims description 4
- 229910001092 metal group alloy Inorganic materials 0.000 claims description 2
- 229910000601 superalloy Inorganic materials 0.000 description 7
- PXHVJJICTQNCMI-UHFFFAOYSA-N Nickel Chemical compound [Ni] PXHVJJICTQNCMI-UHFFFAOYSA-N 0.000 description 5
- 230000000694 effects Effects 0.000 description 4
- 239000003822 epoxy resin Substances 0.000 description 3
- 229910000816 inconels 718 Inorganic materials 0.000 description 3
- 229910052751 metal Inorganic materials 0.000 description 3
- 239000002184 metal Substances 0.000 description 3
- 229920000647 polyepoxide Polymers 0.000 description 3
- 239000004642 Polyimide Substances 0.000 description 2
- 238000013016 damping Methods 0.000 description 2
- 239000011521 glass Substances 0.000 description 2
- 229920001721 polyimide Polymers 0.000 description 2
- 229920001296 polysiloxane Polymers 0.000 description 2
- OKTJSMMVPCPJKN-UHFFFAOYSA-N Carbon Chemical compound [C] OKTJSMMVPCPJKN-UHFFFAOYSA-N 0.000 description 1
- 241001274660 Modulus Species 0.000 description 1
- 229910001069 Ti alloy Inorganic materials 0.000 description 1
- 239000000853 adhesive Substances 0.000 description 1
- 230000001070 adhesive effect Effects 0.000 description 1
- 230000000712 assembly Effects 0.000 description 1
- 238000000429 assembly Methods 0.000 description 1
- 229910052799 carbon Inorganic materials 0.000 description 1
- 230000005489 elastic deformation Effects 0.000 description 1
- 239000000835 fiber Substances 0.000 description 1
- 229910001026 inconel Inorganic materials 0.000 description 1
- 238000000034 method Methods 0.000 description 1
- 229910052759 nickel Inorganic materials 0.000 description 1
- 238000006116 polymerization reaction Methods 0.000 description 1
- 239000012783 reinforcing fiber Substances 0.000 description 1
- 230000000717 retained effect Effects 0.000 description 1
- 238000000926 separation method Methods 0.000 description 1
- 125000006850 spacer group Chemical group 0.000 description 1
- 230000003068 static effect Effects 0.000 description 1
- 230000001052 transient effect Effects 0.000 description 1
- 238000004073 vulcanization Methods 0.000 description 1
Images
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
- F04D29/26—Rotors specially for elastic fluids
- F04D29/32—Rotors specially for elastic fluids for axial flow pumps
- F04D29/321—Rotors specially for elastic fluids for axial flow pumps for axial flow compressors
- F04D29/322—Blade mountings
-
- 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
- F01D5/30—Fixing blades to rotors; Blade roots ; Blade spacers
- F01D5/3007—Fixing blades to rotors; Blade roots ; Blade spacers of axial insertion type
-
- 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
- F01D5/30—Fixing blades to rotors; Blade roots ; Blade spacers
- F01D5/3092—Protective layers between blade root and rotor disc surfaces, e.g. anti-friction layers
-
- 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
- F05D2230/00—Manufacture
- F05D2230/50—Building or constructing in particular ways
- F05D2230/54—Building or constructing in particular ways by sheet metal manufacturing
-
- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T50/00—Aeronautics or air transport
- Y02T50/60—Efficient propulsion technologies, e.g. for aircraft
Definitions
- the invention relates to a turbomachine rotor assembly, of the type comprising: a rotor disk presenting slots in its outer periphery; blades fastened via their roots in said slot; and shims mounted between the blade roots and the disk, each shim comprising two branches and a base part interconnecting the branches, each branch being disposed between the bearing surface of a blade root and the corresponding bearing surface of the disk.
- the invention is applicable to any type of turbomachine whether for terrestrial or aviation purposes (turbojet, turboprop, terrestrial gas turbine, etc.).
- turbomachine whether for terrestrial or aviation purposes (turbojet, turboprop, terrestrial gas turbine, etc.).
- the invention can apply to the fan, to the low pressure compressor (or “booster”), to the high pressure compressor, to the high pressure turbine, or to the low pressure turbine of the turbojet.
- boost low pressure compressor
- the axial direction corresponds to the direction of the axis of rotation A of the turbomachine rotor
- a radial direction is a direction perpendicular to the axis A.
- adjectives such as “inner” and “outer” are used relative to a radial direction such that the (radially) inner portion of an element is closer to the axis A than is the (radially) outer portion of the same element.
- the (moving) blades are fastened to the disk of the rotor via fastener systems, that may be constituted by shank fasteners that are rectilinear or curved, hammerhead-shaped, or Christmas-tree-shaped.
- fastener systems can be defined as devices in which the blade roots form the male portions of the system and are retained radially in the female portions of the system, which female portions are formed in the outer periphery of the disk and are commonly referred to as “slots”.
- the blades When the rotor is set into rotation, the blades are subjected mainly to centrifugal forces and also to axial aerodynamic forces, and the blade roots are pressed in abutment against portions of the disk lying on either side of the outer opening of each slot, under the effect of centrifugal forces.
- the surfaces of the blade roots and of the disks that come into abutment against each other are commonly referred to as “bearing surfaces”. These bearing surfaces are subjected to pressure (as a result of said forces applied to said bearing surfaces). To a first approximation, it can be estimated that this pressure depends on the square of the speed of rotation of the rotor.
- anti-wear solutions can be adopted, i.e. solutions that slow down the appearance of wear at the contact interfaces, and these solutions include those based on inserting a third body, referred to as “shim”, between the blade roots and the disk.
- the shim serves in particular to double the number of contact interfaces (going from a single blade/disk interface to a pair of interfaces, blade/shim and shim/disk), and to reduce the relative movements between the parts that are in contact, thus enabling wear to be reduced in operation.
- shim of the above-mentioned type is described in document FR 2 890 684. That shim is made entirely out of metal, and it is constituted by a sheet of metal that is folded appropriately.
- An object of the invention is to provide a shim that is more effective than the above-mentioned known shim in terms of performing the “anti-wear” function, so as to provide better protection to the bearing surfaces of the blades and of the disk.
- each branch is made of a first material presenting a first Young's modulus of value E, at any operating temperature in the operating temperature range of the shim, the rotor assembly including pads adhering to the bearing surfaces of the disk and/or to the bearing surfaces of the blade root, the pads being made of a second material presenting a second Young's modulus of value lying in the range E/20 to E/5 at said operating temperature.
- operating temperature is used to mean the temperature to which the shim is subjected while the turbomachine is in operation under normal conditions of use.
- the relationship between said first and second Young's moduluses, as defined above, needs to be satisfied for all of the temperatures in the range of operating temperatures of the shim.
- the shim belongs to the fan or to the low pressure compressor of a bypass two-spool airplane turbojet
- its operating temperature lies in the range 20° C. to 150° C.
- its operating temperature lies in the range 1500C to 500° C.
- it belongs to the high pressure turbine of a bypass two-spool airplane turbojet its operating temperature lies in the range 400° C. to 700° C.
- the present invention thus relates to adopting said multilayer structure in which the (isotropic or anisotropic) elasticity characteristics of the second material are better than the (isotropic or anisotropic) elasticity characteristics of the first material in the desired operating temperature range.
- said first material is a metal alloy or an organic matrix composite material
- said second material is non-metallic.
- the second material may be made of rubber, of silicone, of polyimide, of glass, or of epoxy resin.
- the invention also provides a turbomachine including such a rotor assembly.
- FIG. 1 is a fragmentary exploded and diagrammatic view showing a turbomachine rotor assembly comprising a rotor disk, an example of shim in accordance with the invention, and a blade root;
- FIG. 2 is a (fragmentary) radial section view on plane II-II of the FIG. 1 rotor assembly, once it has been assembled;
- FIG. 3 is a (fragmentary) section view analogous to that of FIG. 2 showing another example of a rotor assembly of the invention.
- FIGS. 1 and 2 show: a rotor disk 2 having numerous grooves or “slots” 4 in its periphery that define housings, each suitable for receiving the root 16 of a blade 14 , the root 16 being surrounded by a shim 20 .
- the blade root 16 and the fan disk 2 are made out of titanium alloy, for example.
- assemblies also exist (not shown) that have a spacer placed between the blade root 16 and the bottom of the slot 4 .
- the blades 14 When the disk 2 is set into rotation, the blades 14 are subjected to centrifugal forces, and the bearing surfaces 16 A on the blade root 16 become pressed against bearing surfaces 22 A of the disk 2 .
- the surfaces 16 A constitute the flanks of the blade root 16
- the surfaces 22 A constitute the bottom faces of the lip-shaped portions 22 of the disk that extend on either side of the outer opening of each slot 4 .
- the shim 20 has two side branches 20 A for coming against the bearing surfaces 16 A of the blade root 16 and a base part 20 B, here a base plate, interconnecting the branches and extending under the blade root 16 .
- Each branch 20 A is made of a first material that at any temperature in the operating temperature range of the shim, presents a corresponding first Young's modulus of value E.
- the shim 20 is a wear piece having the main function of limiting wear of the blade root 16 and of the fan disk 2 .
- the shim 20 needs to present a certain amount of stiffness in order to have adequate mechanical strength and perform its anti-wear function.
- the value of E is preferably greater than or equal to 110,000 megapascals (MPa) for a shim made of metal (e.g.: 210,000 MPa for shim made of a nickel-based superalloy, of the type sold under the name “Inconel”), and greater than or equal to 70,000 MPa for a shim made of an organic matrix composite material.
- MPa megapascals
- the rotor assembly includes pads 40 that adhere to the bearing surfaces 22 A of the disk, these pads 40 being made of a second material presenting a second Young's modulus of value lying in the range E/20 to E/5 at said temperature.
- pads are fastened on the bearing surfaces 16 A of the blade root, or on both the bearing surfaces 22 A of the disk and the bearing surfaces 16 A of the blade root.
- the rotor assembly belongs to the fan or the low pressure compressor of a bypass two-spool airplane turbojet, it is subjected to operating temperatures lying in the range 20° C. to 150° C.
- the first material can be selected as a Ni-based superalloy with more than 15% by weight Fe and Cr, such as the superalloy sold under the name “Inconel 718”; while the second material can be rubber (natural or synthetic).
- the first material it is also possible for the first material to be a composite material using an epoxy resin matrix with reinforcing fibers, e.g. made of carbon; the second material could then be an epoxy resin on its own (with the difference in Young's modulus between the first and second materials being associated with the absence of fibers).
- the assembly belongs to the high pressure compressor of a bypass two-spool airplane turbojet, it is subjected to operating temperatures lying in the range 150° C. to 500° C. Under such circumstances, and by way of example, it is possible to select for the first material an Ni-based superalloy having more than 15% by weight of Fe and Cr, such as the superalloy sold under the name “Inconel 718”; the second material could be a silicone or polyimide.
- the assembly belongs to the high pressure turbine of a bypass two-spool airplane turbojet, it is subjected to operating temperatures lying in the range 400° C. to 700° C.
- the first material to be selected as an Ni-based superalloy with more than 15% by weight Fe and Cr, such as the superalloy sold under the name “Inconel 718”; the second material may be glass (which in this operating temperature range presents viscoelastic behavior).
- the pads 40 can be fastened to the bearing surfaces 22 A of the disk in various ways, and in particular:
- the fastening obtained must be sufficient to prevent the pads 40 from detaching from the bearing surfaces 22 A in operation.
- FIG. 3 is a section view analogous to that of FIG. 2 showing (in part) another example of a rotor assembly 120 of the invention. Elements or element portions that are analogous between FIGS. 2 and 3 are identified by the same numerical references plus 100.
- FIG. 3 differs from that of FIG. 2 in that the base part 120 B of the shim 120 extends over the outer periphery of the rotor disk 202 , between two adjacent slots 204 , with each branch 120 A of the shim entering into a respective slot 204 and being received between the bearing surface 216 A of the blade root 216 and the corresponding bearing surface 222 A of the disk 202 .
- the positions of the pads 140 remain the same.
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)
- Gasket Seals (AREA)
Abstract
A turbomachine rotor assembly comprising: a rotor disk (2) presenting slots (4) in its outer periphery; blades (14) fastened via their roots (16) in said slot (4); and shims (20) mounted between the blade roots (16) and the disk (2), each shim (20) comprising two branches (20A) and a base part (20B) interconnecting the branches, each branch (20A) being disposed between the bearing surface (16A) of a blade root (16) and the corresponding bearing surface (22A) of the disk (2); each branch (20A) being made of a first material presenting a first Young's modulus of value E, at any operating temperature in the operating temperature range of the shim (20). The rotor assembly includes pads (40) adhering to the bearing surfaces (22A) of the disk and/or to the bearing surfaces (16A) of the blade root, the pads (40) being made of a second material presenting a second Young's modulus of value lying in the range E/20 to E/5 at said operating temperature.
Description
- The invention relates to a turbomachine rotor assembly, of the type comprising: a rotor disk presenting slots in its outer periphery; blades fastened via their roots in said slot; and shims mounted between the blade roots and the disk, each shim comprising two branches and a base part interconnecting the branches, each branch being disposed between the bearing surface of a blade root and the corresponding bearing surface of the disk.
- The invention is applicable to any type of turbomachine whether for terrestrial or aviation purposes (turbojet, turboprop, terrestrial gas turbine, etc.). In the particular example of a bypass two-spool airplane turbojet, the invention can apply to the fan, to the low pressure compressor (or “booster”), to the high pressure compressor, to the high pressure turbine, or to the low pressure turbine of the turbojet.
- In the present application, the axial direction corresponds to the direction of the axis of rotation A of the turbomachine rotor, and a radial direction is a direction perpendicular to the axis A. Furthermore, unless specified to the contrary, adjectives such as “inner” and “outer” are used relative to a radial direction such that the (radially) inner portion of an element is closer to the axis A than is the (radially) outer portion of the same element.
- In a rotor assembly (i.e. an assembly forming part of the rotor) of the above-specified type, the (moving) blades are fastened to the disk of the rotor via fastener systems, that may be constituted by shank fasteners that are rectilinear or curved, hammerhead-shaped, or Christmas-tree-shaped. These fastener systems can be defined as devices in which the blade roots form the male portions of the system and are retained radially in the female portions of the system, which female portions are formed in the outer periphery of the disk and are commonly referred to as “slots”.
- When the rotor is set into rotation, the blades are subjected mainly to centrifugal forces and also to axial aerodynamic forces, and the blade roots are pressed in abutment against portions of the disk lying on either side of the outer opening of each slot, under the effect of centrifugal forces. The surfaces of the blade roots and of the disks that come into abutment against each other are commonly referred to as “bearing surfaces”. These bearing surfaces are subjected to pressure (as a result of said forces applied to said bearing surfaces). To a first approximation, it can be estimated that this pressure depends on the square of the speed of rotation of the rotor.
- It can thus be understood that the variations in the speed of rotation of the rotor during an operating cycle of the turbomachine: from stationary to full throttle, passing through various particular intermediate speeds (idling, taxiing, cruising, descending, for an aviation turbomachine), give rise to variations in the pressure acting on the above-defined bearing surfaces. These pressure variations associated with elastic deformations of the contacting parts give rise to relative movements between the blade roots and the disk. When they are repeated, these relative movements, known as slip or as separation depending on their nature, give rise to wear phenomena in the bearing surfaces of the blades or of the disks. It is also possible for the dynamic movements of the blades at a given speed of rotation (response of the blades to alternating stresses of harmonic or transient nature) to contribute to the phenomenon of said bearing surfaces becoming worn. These wear phenomena are naturally penalizing on the lifetime of a turbomachine.
- Various so-called “anti-wear” solutions can be adopted, i.e. solutions that slow down the appearance of wear at the contact interfaces, and these solutions include those based on inserting a third body, referred to as “shim”, between the blade roots and the disk. The shim serves in particular to double the number of contact interfaces (going from a single blade/disk interface to a pair of interfaces, blade/shim and shim/disk), and to reduce the relative movements between the parts that are in contact, thus enabling wear to be reduced in operation.
- A known example of shim of the above-mentioned type is described in
document FR 2 890 684. That shim is made entirely out of metal, and it is constituted by a sheet of metal that is folded appropriately. - An object of the invention is to provide a shim that is more effective than the above-mentioned known shim in terms of performing the “anti-wear” function, so as to provide better protection to the bearing surfaces of the blades and of the disk.
- This object is achieved by a turbomachine rotor assembly of the type defined in the introduction, in which each branch is made of a first material presenting a first Young's modulus of value E, at any operating temperature in the operating temperature range of the shim, the rotor assembly including pads adhering to the bearing surfaces of the disk and/or to the bearing surfaces of the blade root, the pads being made of a second material presenting a second Young's modulus of value lying in the range E/20 to E/5 at said operating temperature.
- It should be observed that the Young's modulus of a material varies as a function of the temperature of the material, and consequently that the values E and E′ depend on temperature.
- The term “operating temperature” is used to mean the temperature to which the shim is subjected while the turbomachine is in operation under normal conditions of use. In the present invention, the relationship between said first and second Young's moduluses, as defined above, needs to be satisfied for all of the temperatures in the range of operating temperatures of the shim.
- For example, when the shim belongs to the fan or to the low pressure compressor of a bypass two-spool airplane turbojet, its operating temperature lies in the
range 20° C. to 150° C. When it belongs to the high pressure compressor of a bypass two-spool airplane turbojet, its operating temperature lies in the range 1500C to 500° C. When it belongs to the high pressure turbine of a bypass two-spool airplane turbojet, its operating temperature lies in the range 400° C. to 700° C. - The present invention thus relates to adopting said multilayer structure in which the (isotropic or anisotropic) elasticity characteristics of the second material are better than the (isotropic or anisotropic) elasticity characteristics of the first material in the desired operating temperature range.
- In an embodiment, said first material is a metal alloy or an organic matrix composite material, while said second material is non-metallic. For example, and in non-exhaustive manner, the second material may be made of rubber, of silicone, of polyimide, of glass, or of epoxy resin.
- The invention has the following effects:
-
- uniformly distributing contact pressures by the pads accommodating due to the elasticity of the second material;
- during a change in speed of rotation, limiting relative movements of the parts due to centrifugal forces by “static” shear in said pads; and
- damping any dynamic movements of the blade by “dynamic” shear in said pads.
- A particular consequence of these effects is to prevent or limit wear phenomena in the bearing surfaces, thereby increasing the lifetimes of blade roots and of disks.
- These effects are reinforced when the second material presents viscoelastic behavior in the operating temperature range of the shim, more particularly for the purpose of damping any dynamic movements of the blade.
- The invention also provides a turbomachine including such a rotor assembly.
- The invention and its advantages can be better understood on reading the following detailed description. The description refers to the accompanying figures, in which:
-
FIG. 1 is a fragmentary exploded and diagrammatic view showing a turbomachine rotor assembly comprising a rotor disk, an example of shim in accordance with the invention, and a blade root; -
FIG. 2 is a (fragmentary) radial section view on plane II-II of theFIG. 1 rotor assembly, once it has been assembled; and -
FIG. 3 is a (fragmentary) section view analogous to that ofFIG. 2 showing another example of a rotor assembly of the invention. -
FIGS. 1 and 2 show: arotor disk 2 having numerous grooves or “slots” 4 in its periphery that define housings, each suitable for receiving theroot 16 of ablade 14, theroot 16 being surrounded by ashim 20. Theblade root 16 and thefan disk 2 are made out of titanium alloy, for example. - It should be observed that assemblies also exist (not shown) that have a spacer placed between the
blade root 16 and the bottom of the slot 4. - When the
disk 2 is set into rotation, theblades 14 are subjected to centrifugal forces, and thebearing surfaces 16A on theblade root 16 become pressed against bearingsurfaces 22A of thedisk 2. In the example shown, thesurfaces 16A constitute the flanks of theblade root 16, while thesurfaces 22A constitute the bottom faces of the lip-shaped portions 22 of the disk that extend on either side of the outer opening of each slot 4. - The
shim 20 has twoside branches 20A for coming against thebearing surfaces 16A of theblade root 16 and abase part 20B, here a base plate, interconnecting the branches and extending under theblade root 16. Eachbranch 20A is made of a first material that at any temperature in the operating temperature range of the shim, presents a corresponding first Young's modulus of value E. Theshim 20 is a wear piece having the main function of limiting wear of theblade root 16 and of thefan disk 2. Theshim 20 needs to present a certain amount of stiffness in order to have adequate mechanical strength and perform its anti-wear function. Thus, the value of E is preferably greater than or equal to 110,000 megapascals (MPa) for a shim made of metal (e.g.: 210,000 MPa for shim made of a nickel-based superalloy, of the type sold under the name “Inconel”), and greater than or equal to 70,000 MPa for a shim made of an organic matrix composite material. - In accordance with the invention, the rotor assembly includes
pads 40 that adhere to thebearing surfaces 22A of the disk, thesepads 40 being made of a second material presenting a second Young's modulus of value lying in the range E/20 to E/5 at said temperature. - In other embodiments (not shown), pads are fastened on the
bearing surfaces 16A of the blade root, or on both thebearing surfaces 22A of the disk and thebearing surfaces 16A of the blade root. - So far as the choice of materials is concerned, it naturally depends on the operating temperature of the shim.
- When the rotor assembly belongs to the fan or the low pressure compressor of a bypass two-spool airplane turbojet, it is subjected to operating temperatures lying in the
range 20° C. to 150° C. Under such circumstances, and by way of example, it is possible for the first material to be selected as a Ni-based superalloy with more than 15% by weight Fe and Cr, such as the superalloy sold under the name “Inconel 718”; while the second material can be rubber (natural or synthetic). In these circumstances, it is also possible for the first material to be a composite material using an epoxy resin matrix with reinforcing fibers, e.g. made of carbon; the second material could then be an epoxy resin on its own (with the difference in Young's modulus between the first and second materials being associated with the absence of fibers). - When the assembly belongs to the high pressure compressor of a bypass two-spool airplane turbojet, it is subjected to operating temperatures lying in the range 150° C. to 500° C. Under such circumstances, and by way of example, it is possible to select for the first material an Ni-based superalloy having more than 15% by weight of Fe and Cr, such as the superalloy sold under the name “Inconel 718”; the second material could be a silicone or polyimide.
- When the assembly belongs to the high pressure turbine of a bypass two-spool airplane turbojet, it is subjected to operating temperatures lying in the range 400° C. to 700° C. Under such circumstances, and by way of example, it is possible for the first material to be selected as an Ni-based superalloy with more than 15% by weight Fe and Cr, such as the superalloy sold under the name “Inconel 718”; the second material may be glass (which in this operating temperature range presents viscoelastic behavior).
- In general, it should be observed that the
pads 40 can be fastened to thebearing surfaces 22A of the disk in various ways, and in particular: -
- by natural adhesion;
- during polymerization of the pad (during its vulcanization it if is made of rubber); or
- by adhesive;
- or by combining the above-mentioned techniques.
- Naturally, the fastening obtained must be sufficient to prevent the
pads 40 from detaching from the bearing surfaces 22A in operation. -
FIG. 3 is a section view analogous to that ofFIG. 2 showing (in part) another example of arotor assembly 120 of the invention. Elements or element portions that are analogous betweenFIGS. 2 and 3 are identified by the same numerical references plus 100. - The example of
FIG. 3 differs from that ofFIG. 2 in that thebase part 120B of theshim 120 extends over the outer periphery of the rotor disk 202, between two adjacent slots 204, with eachbranch 120A of the shim entering into a respective slot 204 and being received between the bearing surface 216A of the blade root 216 and the corresponding bearing surface 222A of the disk 202. The positions of thepads 140 remain the same.
Claims (6)
1. A turbomachine rotor assembly comprising: a rotor disk presenting slots in its outer periphery; blades fastened via their roots in said slots; and shims mounted between the blade roots and the disk, each shim comprising two branches and a base part interconnecting the branches, each branch being disposed between the bearing surface of a blade root and the corresponding bearing surface of the disk; each branch being made of a first material presenting a first Young's modulus of value E, at any operating temperature in the operating temperature range of the shim, said rotor assembly including pads adhering to the bearing surfaces of the disk and/or to the bearing surfaces of the blade root, the pads being made of a second material presenting a second Young's modulus of value lying in the range E/20 to E/5 at said operating temperature.
2. A rotor assembly according to claim 1 , in which said first material is a metal alloy or an organic matrix composite material, while said second material is non-metallic.
3. A rotor assembly according to claim 1 , in which said second material has viscoelastic behavior in the operating temperature range of the shim.
4. A rotor assembly according to claim 1 , in which the base part of each shim extends under a respective blade root.
5. A rotor assembly according to claim 1 , in which the base part of each shim extends over the outer periphery of the disk between two adjacent slots.
6. A turbomachine including a rotor assembly according to claim 1 .
Applications Claiming Priority (2)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
FR0756467A FR2918703B1 (en) | 2007-07-13 | 2007-07-13 | ROTOR ASSEMBLY OF TURBOMACHINE |
FR0756467 | 2007-07-13 |
Publications (1)
Publication Number | Publication Date |
---|---|
US20090016890A1 true US20090016890A1 (en) | 2009-01-15 |
Family
ID=39106159
Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
US12/171,775 Abandoned US20090016890A1 (en) | 2007-07-13 | 2008-07-11 | Turbomachine rotor assembly |
Country Status (8)
Country | Link |
---|---|
US (1) | US20090016890A1 (en) |
EP (1) | EP2014874B1 (en) |
JP (1) | JP2009019630A (en) |
CN (1) | CN101344013B (en) |
CA (1) | CA2636924A1 (en) |
DE (1) | DE602008005524D1 (en) |
FR (1) | FR2918703B1 (en) |
RU (1) | RU2465464C2 (en) |
Cited By (24)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US20120107125A1 (en) * | 2009-04-29 | 2012-05-03 | Snecma | Reinforced fan blade shim |
US20120257981A1 (en) * | 2011-04-11 | 2012-10-11 | Rolls-Royce Plc | Retention device for a composite blade of a gas turbine engine |
US20120263596A1 (en) * | 2011-04-14 | 2012-10-18 | Rolls-Royce Plc | Annulus filler system |
WO2013148445A1 (en) * | 2012-03-26 | 2013-10-03 | United Technologies Corporation | Blade wedge attachment |
US20140294597A1 (en) * | 2011-10-10 | 2014-10-02 | Snecma | Cooling for the retaining dovetail of a turbomachine blade |
WO2014143364A3 (en) * | 2013-03-14 | 2014-11-27 | United Technologies Corporation | Co-formed element with low conductivity layer |
US20150110636A1 (en) * | 2013-02-12 | 2015-04-23 | United Technologies Corporation | Wear pad to prevent cracking of fan blade |
US20150192144A1 (en) * | 2014-01-08 | 2015-07-09 | United Technologies Corporation | Fan Assembly With Fan Blade Under-Root Spacer |
US9506356B2 (en) | 2013-03-15 | 2016-11-29 | Rolls-Royce North American Technologies, Inc. | Composite retention feature |
US20170167439A1 (en) * | 2015-12-14 | 2017-06-15 | Rohr, Inc. | Cascade assembly for a thrust reverser of an aircraft nacelle |
US9745856B2 (en) | 2013-03-13 | 2017-08-29 | Rolls-Royce Corporation | Platform for ceramic matrix composite turbine blades |
US9975815B2 (en) | 2015-02-26 | 2018-05-22 | General Electric Company | Methods for forming ceramic matrix composite articles |
US20180245475A1 (en) * | 2015-08-19 | 2018-08-30 | Siemens Aktiengesellschaft | Gas turbine blade or compressor blade having anti-fretting coating in the blade root region and rotor |
US10329912B2 (en) | 2012-09-03 | 2019-06-25 | Safran Aircraft Engines | Turbine rotor for a turbomachine |
US10519788B2 (en) | 2013-05-29 | 2019-12-31 | General Electric Company | Composite airfoil metal patch |
US10577961B2 (en) | 2018-04-23 | 2020-03-03 | Rolls-Royce High Temperature Composites Inc. | Turbine disk with blade supported platforms |
CN111022126A (en) * | 2019-11-19 | 2020-04-17 | 中国航发沈阳黎明航空发动机有限责任公司 | Rotor sealing vibration reduction structure |
EP2960009B1 (en) * | 2014-06-24 | 2020-05-13 | Rolls-Royce plc | Rotor blade manufacture |
US10767498B2 (en) | 2018-04-03 | 2020-09-08 | Rolls-Royce High Temperature Composites Inc. | Turbine disk with pinned platforms |
US10890081B2 (en) | 2018-04-23 | 2021-01-12 | Rolls-Royce Corporation | Turbine disk with platforms coupled to disk |
US11426963B2 (en) * | 2019-04-17 | 2022-08-30 | Mitsubishi Heavy Industries, Ltd. | Composite blade and method of forming composite blade |
WO2023175256A1 (en) * | 2022-03-18 | 2023-09-21 | Safran Aircraft Engines | Method for maintaining a bladed wheel of a high-pressure turbine of a turbomachine |
US11988602B2 (en) | 2018-05-11 | 2024-05-21 | Carrier Corporation | Surface plasmon resonance detection system |
US20240200460A1 (en) * | 2021-04-01 | 2024-06-20 | Safran Aircraft Engines | Foil for a turbomachine rotor blade, assembly for a turbomachine rotor, and turbomachine |
Families Citing this family (7)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US8251667B2 (en) * | 2009-05-20 | 2012-08-28 | General Electric Company | Low stress circumferential dovetail attachment for rotor blades |
US8740573B2 (en) * | 2011-04-26 | 2014-06-03 | General Electric Company | Adaptor assembly for coupling turbine blades to rotor disks |
US11156090B2 (en) | 2013-03-14 | 2021-10-26 | Raytheon Technologies Corporation | Low speed fan for gas turbine engines |
CN104500446A (en) * | 2014-12-14 | 2015-04-08 | 惠阳航空螺旋桨有限责任公司 | Connection structure of composite material blade roots of wind tunnel axial flow compressor fan rotor |
CN104500447A (en) * | 2014-12-14 | 2015-04-08 | 惠阳航空螺旋桨有限责任公司 | Wind tunnel axial flow compressor fan |
US10605100B2 (en) * | 2017-05-24 | 2020-03-31 | General Electric Company | Ceramic matrix composite (CMC) turbine blade assembly, dovetail sleeve, and method of mounting CMC turbine blade |
FR3099213B1 (en) * | 2019-07-23 | 2021-07-16 | Safran Aircraft Engines | BLOWER ROTOR FOR AN AIRCRAFT TURBOMACHINE |
Citations (11)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US2950083A (en) * | 1954-07-23 | 1960-08-23 | Thompson Ramo Wooldridge Inc | Blade assembly |
US3784320A (en) * | 1971-02-20 | 1974-01-08 | Motoren Turbinen Union | Method and means for retaining ceramic turbine blades |
US4169694A (en) * | 1977-07-20 | 1979-10-02 | Electric Power Research Institute, Inc. | Ceramic rotor blade having root with double curvature |
US4207029A (en) * | 1978-06-12 | 1980-06-10 | Avco Corporation | Turbine rotor assembly of ceramic blades to metallic disc |
US5087174A (en) * | 1990-01-22 | 1992-02-11 | Westinghouse Electric Corp. | Temperature activated expanding mineral shim |
US5160243A (en) * | 1991-01-15 | 1992-11-03 | General Electric Company | Turbine blade wear protection system with multilayer shim |
US5462408A (en) * | 1992-12-23 | 1995-10-31 | Europcopter France | Blade made of thermoplastic composite, in particular for ducted tail rotor of a helicopter, and its method of manufacture |
US6290466B1 (en) * | 1999-09-17 | 2001-09-18 | General Electric Company | Composite blade root attachment |
US6398499B1 (en) * | 2000-10-17 | 2002-06-04 | Honeywell International, Inc. | Fan blade compliant layer and seal |
US6751863B2 (en) * | 2002-05-07 | 2004-06-22 | General Electric Company | Method for providing a rotating structure having a wire-arc-sprayed aluminum bronze protective coating thereon |
US6860722B2 (en) * | 2003-01-31 | 2005-03-01 | General Electric Company | Snap on blade shim |
Family Cites Families (13)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
GB709636A (en) * | 1951-05-09 | 1954-06-02 | Rolls Royce | Improvements in or relating to compressor and turbine bladed rotors |
US2874932A (en) * | 1952-02-25 | 1959-02-24 | Maschf Augsburg Nuernberg Ag | Steel turbine rotors with ceramic blades |
SU595521A1 (en) * | 1976-10-08 | 1978-02-28 | Научно-производственное объединение по технологии машиностроения | Turbine machine runner |
JPS62113804A (en) * | 1985-11-13 | 1987-05-25 | Toshiba Corp | Steam turbine |
DE3815977A1 (en) * | 1988-05-10 | 1989-11-30 | Mtu Muenchen Gmbh | INTERMEDIATE FILM FOR JOINING MACHINE COMPONENTS HAZARDOUS TO FRICTION |
JP2718131B2 (en) * | 1989-01-23 | 1998-02-25 | 石川島播磨重工業株式会社 | Gas turbine disk |
US5240375A (en) * | 1992-01-10 | 1993-08-31 | General Electric Company | Wear protection system for turbine engine rotor and blade |
JPH07247804A (en) * | 1993-01-07 | 1995-09-26 | General Electric Co <Ge> | Rotor and moving vane assembly for gas-turbine engine and multilayer covering shim |
JP3216956B2 (en) * | 1994-06-08 | 2001-10-09 | 株式会社日立製作所 | Gas turbine blade fixing device |
WO1996041068A1 (en) * | 1995-06-07 | 1996-12-19 | National Research Council Of Canada | Anti-fretting barrier |
US6132175A (en) * | 1997-05-29 | 2000-10-17 | Alliedsignal, Inc. | Compliant sleeve for ceramic turbine blades |
FR2890126B1 (en) * | 2005-08-26 | 2010-10-29 | Snecma | ASSEMBLY AND METHOD FOR THE FOOT ASSEMBLY OF A TURBOMACHINE, BLOWER, COMPRESSOR AND TURBOMACHINE BLADE COMPRISING SUCH AN ASSEMBLY |
FR2890684B1 (en) * | 2005-09-15 | 2007-12-07 | Snecma | CLINKING FOR TURBOREACTOR BLADE |
-
2007
- 2007-07-13 FR FR0756467A patent/FR2918703B1/en not_active Expired - Fee Related
-
2008
- 2008-07-10 CA CA002636924A patent/CA2636924A1/en not_active Abandoned
- 2008-07-11 JP JP2008181059A patent/JP2009019630A/en active Pending
- 2008-07-11 EP EP08160249A patent/EP2014874B1/en active Active
- 2008-07-11 DE DE602008005524T patent/DE602008005524D1/en active Active
- 2008-07-11 RU RU2008128359/06A patent/RU2465464C2/en active
- 2008-07-11 US US12/171,775 patent/US20090016890A1/en not_active Abandoned
- 2008-07-14 CN CN2008101379626A patent/CN101344013B/en active Active
Patent Citations (12)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US2950083A (en) * | 1954-07-23 | 1960-08-23 | Thompson Ramo Wooldridge Inc | Blade assembly |
US3784320A (en) * | 1971-02-20 | 1974-01-08 | Motoren Turbinen Union | Method and means for retaining ceramic turbine blades |
US4169694A (en) * | 1977-07-20 | 1979-10-02 | Electric Power Research Institute, Inc. | Ceramic rotor blade having root with double curvature |
US4207029A (en) * | 1978-06-12 | 1980-06-10 | Avco Corporation | Turbine rotor assembly of ceramic blades to metallic disc |
US5087174A (en) * | 1990-01-22 | 1992-02-11 | Westinghouse Electric Corp. | Temperature activated expanding mineral shim |
US5160243A (en) * | 1991-01-15 | 1992-11-03 | General Electric Company | Turbine blade wear protection system with multilayer shim |
US5462408A (en) * | 1992-12-23 | 1995-10-31 | Europcopter France | Blade made of thermoplastic composite, in particular for ducted tail rotor of a helicopter, and its method of manufacture |
US6290466B1 (en) * | 1999-09-17 | 2001-09-18 | General Electric Company | Composite blade root attachment |
US6398499B1 (en) * | 2000-10-17 | 2002-06-04 | Honeywell International, Inc. | Fan blade compliant layer and seal |
US6431835B1 (en) * | 2000-10-17 | 2002-08-13 | Honeywell International, Inc. | Fan blade compliant shim |
US6751863B2 (en) * | 2002-05-07 | 2004-06-22 | General Electric Company | Method for providing a rotating structure having a wire-arc-sprayed aluminum bronze protective coating thereon |
US6860722B2 (en) * | 2003-01-31 | 2005-03-01 | General Electric Company | Snap on blade shim |
Cited By (36)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US8870545B2 (en) * | 2009-04-29 | 2014-10-28 | Snecma | Reinforced fan blade shim |
US20120107125A1 (en) * | 2009-04-29 | 2012-05-03 | Snecma | Reinforced fan blade shim |
US20120257981A1 (en) * | 2011-04-11 | 2012-10-11 | Rolls-Royce Plc | Retention device for a composite blade of a gas turbine engine |
US9039379B2 (en) * | 2011-04-11 | 2015-05-26 | Rolls-Royce Plc | Retention device for a composite blade of a gas turbine engine |
US20120263596A1 (en) * | 2011-04-14 | 2012-10-18 | Rolls-Royce Plc | Annulus filler system |
US20140294597A1 (en) * | 2011-10-10 | 2014-10-02 | Snecma | Cooling for the retaining dovetail of a turbomachine blade |
US9631495B2 (en) * | 2011-10-10 | 2017-04-25 | Snecma | Cooling for the retaining dovetail of a turbomachine blade |
WO2013148445A1 (en) * | 2012-03-26 | 2013-10-03 | United Technologies Corporation | Blade wedge attachment |
US9611746B2 (en) | 2012-03-26 | 2017-04-04 | United Technologies Corporation | Blade wedge attachment |
US10329912B2 (en) | 2012-09-03 | 2019-06-25 | Safran Aircraft Engines | Turbine rotor for a turbomachine |
US20150110636A1 (en) * | 2013-02-12 | 2015-04-23 | United Technologies Corporation | Wear pad to prevent cracking of fan blade |
US10273816B2 (en) * | 2013-02-12 | 2019-04-30 | United Technologies Corporation | Wear pad to prevent cracking of fan blade |
US9745856B2 (en) | 2013-03-13 | 2017-08-29 | Rolls-Royce Corporation | Platform for ceramic matrix composite turbine blades |
WO2014143364A3 (en) * | 2013-03-14 | 2014-11-27 | United Technologies Corporation | Co-formed element with low conductivity layer |
US10309230B2 (en) | 2013-03-14 | 2019-06-04 | United Technologies Corporation | Co-formed element with low conductivity layer |
US9506356B2 (en) | 2013-03-15 | 2016-11-29 | Rolls-Royce North American Technologies, Inc. | Composite retention feature |
US10519788B2 (en) | 2013-05-29 | 2019-12-31 | General Electric Company | Composite airfoil metal patch |
US20150192144A1 (en) * | 2014-01-08 | 2015-07-09 | United Technologies Corporation | Fan Assembly With Fan Blade Under-Root Spacer |
EP2960009B1 (en) * | 2014-06-24 | 2020-05-13 | Rolls-Royce plc | Rotor blade manufacture |
US9975815B2 (en) | 2015-02-26 | 2018-05-22 | General Electric Company | Methods for forming ceramic matrix composite articles |
US11414354B2 (en) | 2015-02-26 | 2022-08-16 | General Electric Company | Ceramic matrix composite articles and methods for forming same |
US10093586B2 (en) | 2015-02-26 | 2018-10-09 | General Electric Company | Ceramic matrix composite articles and methods for forming same |
US10954168B2 (en) | 2015-02-26 | 2021-03-23 | General Electric Company | Ceramic matrix composite articles and methods for forming same |
US20180245475A1 (en) * | 2015-08-19 | 2018-08-30 | Siemens Aktiengesellschaft | Gas turbine blade or compressor blade having anti-fretting coating in the blade root region and rotor |
US11352893B2 (en) * | 2015-08-19 | 2022-06-07 | Siemens Energy Globall Gmbh & Co. Kg | Gas turbine blade or compressor blade having anti-fretting coating in the blade root region and rotor |
US20170167439A1 (en) * | 2015-12-14 | 2017-06-15 | Rohr, Inc. | Cascade assembly for a thrust reverser of an aircraft nacelle |
US10514004B2 (en) * | 2015-12-14 | 2019-12-24 | Rohr, Inc. | Cascade assembly for a thrust reverser of an aircraft nacelle |
US10767498B2 (en) | 2018-04-03 | 2020-09-08 | Rolls-Royce High Temperature Composites Inc. | Turbine disk with pinned platforms |
US10890081B2 (en) | 2018-04-23 | 2021-01-12 | Rolls-Royce Corporation | Turbine disk with platforms coupled to disk |
US10577961B2 (en) | 2018-04-23 | 2020-03-03 | Rolls-Royce High Temperature Composites Inc. | Turbine disk with blade supported platforms |
US11988602B2 (en) | 2018-05-11 | 2024-05-21 | Carrier Corporation | Surface plasmon resonance detection system |
US11426963B2 (en) * | 2019-04-17 | 2022-08-30 | Mitsubishi Heavy Industries, Ltd. | Composite blade and method of forming composite blade |
CN111022126A (en) * | 2019-11-19 | 2020-04-17 | 中国航发沈阳黎明航空发动机有限责任公司 | Rotor sealing vibration reduction structure |
US20240200460A1 (en) * | 2021-04-01 | 2024-06-20 | Safran Aircraft Engines | Foil for a turbomachine rotor blade, assembly for a turbomachine rotor, and turbomachine |
WO2023175256A1 (en) * | 2022-03-18 | 2023-09-21 | Safran Aircraft Engines | Method for maintaining a bladed wheel of a high-pressure turbine of a turbomachine |
FR3133640A1 (en) * | 2022-03-18 | 2023-09-22 | Safran Aircraft Engines | Method of maintaining a high-pressure turbine bladed wheel of a turbomachine |
Also Published As
Publication number | Publication date |
---|---|
RU2008128359A (en) | 2010-01-20 |
JP2009019630A (en) | 2009-01-29 |
FR2918703A1 (en) | 2009-01-16 |
DE602008005524D1 (en) | 2011-04-28 |
CA2636924A1 (en) | 2009-01-13 |
RU2465464C2 (en) | 2012-10-27 |
CN101344013B (en) | 2013-11-06 |
CN101344013A (en) | 2009-01-14 |
EP2014874B1 (en) | 2011-03-16 |
FR2918703B1 (en) | 2009-10-16 |
EP2014874A1 (en) | 2009-01-14 |
Similar Documents
Publication | Publication Date | Title |
---|---|---|
US20090016890A1 (en) | Turbomachine rotor assembly | |
US20090060745A1 (en) | Shim for a turbomachine blade | |
US8500410B2 (en) | Blade made of composite material comprising a damping device | |
US8061997B2 (en) | Damping device for composite blade | |
US9657577B2 (en) | Rotor blade with bonded cover | |
EP2837772B1 (en) | Annulus filler and corresponding stage and gas turbine engine | |
US8834105B2 (en) | Structural low-ductility turbine shroud apparatus | |
US20110206530A1 (en) | Vibration damper device for turbomachine blade attachments, associated turbomachine and associated engines | |
EP3093435B1 (en) | Rotor damper | |
US20150050150A1 (en) | Annulus filler | |
US20150322807A1 (en) | Multi-ply finger seal | |
US20140086751A1 (en) | Annulus filler for axial flow machine | |
JP2008002460A (en) | Sector of compressor guide vanes assembly or sector of turbomachine nozzle assembly | |
GB2504035A (en) | Turbine engine comprising a metal protector for a composite part | |
JP2016501339A (en) | Apparatus and method for reducing CMC and metal attachment and interface wear and friction | |
US20140064938A1 (en) | Rub tolerant fan case | |
US11346233B2 (en) | Damping device | |
CN108026785B (en) | Turbine of a turbine engine, turbojet engine and aircraft | |
US9803648B2 (en) | Retainer plate | |
US11536157B2 (en) | Damping device | |
CN113818933A (en) | Turbine engine with floating interstage seal |
Legal Events
Date | Code | Title | Description |
---|---|---|---|
AS | Assignment |
Owner name: SNECMA, FRANCE Free format text: ASSIGNMENT OF ASSIGNORS INTEREST;ASSIGNORS:DOUGUET, CHARLES JEAN-PIERRE;JACO, CHRISTOPHE;LOMBARD, JEAN-PIERRE FRANCOIS;REEL/FRAME:021528/0883 Effective date: 20080626 |
|
STCB | Information on status: application discontinuation |
Free format text: ABANDONED -- FAILURE TO RESPOND TO AN OFFICE ACTION |