Blade comprising a structure made of composite material and related manufacturing method

Abstract

The invention relates to a blade (7) of a turbomachine, comprising: - an aerodynamic profile structure (20); - a main beam (21) comprising a blade root portion (24), an airfoil portion (25) arranged inside the aerodynamic profile structure (20) and comprising a main body (27) connected to the blade root portion (24) and two branches extending radially from the main body (27), and a strut portion (26); and - a structural reinforcement (31) extending from the blade root portion (24) to the main body (27) of the airfoil portion (25), the structural reinforcement (31) being rigidly connected to the blade root portion (24) and being configured to form a different force path than the strut portion (26) in case of breakage of the main beam (21) within the strut portion (26).

Description

Technical Field

[0001] The present invention relates to a blade comprising a structure made of composite materials.

[0002] The present invention relates more particularly, but not exclusively, to blades intended for use in ductless fans or blades for wind turbine rotors in aircraft engines (e.g., open rotor type engines with two rotating propellers (i.e., engines whose fans have no ducts), or unducted single fan (USF) type engines with blade assemblies and wheel blade assemblies, or turboprop engines with an architecture containing a single propeller). Background Technology

[0003] The design of fan blades needs to take into account opposing stresses.

[0004] On the one hand, the dimensions of these blades must be such that they achieve optimal aerodynamic performance (maximizing efficiency and providing thrust while minimizing losses). Improvements in fan aerodynamic performance tend to increase the bypass ratio (BPR), which leads to an increase in the outer diameter, and consequently, an increase in the span of these blades.

[0005] On the other hand, it is also necessary to ensure resistance to mechanical stresses that may be applied to these blades, while limiting the acoustic characteristics of these blades.

[0006] The advantage of ductless fan engines is that the diameter of the fan is not limited by the presence of a cowl, which allows for the design of engines with high bypass ratios and reduced fuel consumption.

[0007] Therefore, in this type of engine, the fan blades can have a large span.

[0008] The manufacture of these blades from metallic materials was proposed. Although blades made of metallic materials have good mechanical resistance, they also have the disadvantage of relatively large mass.

[0009] To reduce this mass, it is hoped that these blades can be manufactured using composite materials.

[0010] In addition, these engines typically include mechanisms that allow for changes in the blade pitch angle to adjust the thrust generated by the fan according to different phases of flight.

[0011] Furthermore, in ductless fan architectures, engine startup is typically feathered at a very wide pitch angle. In fact, this very wide pitch angle allows power to be consumed through torque, which ensures machine safety by maintaining low fan speeds.

[0012] However, at very wide pitch angles, the blades experience strongly separated turbulent aerodynamic flows, which generates broadband vibration excitation. Particularly on blades with wide chords and long spans, the bending forces are strong even at low engine speeds.

[0013] However, the strong aerodynamic forces acting on these blades can damage the blades at the level of the blade strut (which connects the root to the airfoil) and / or damage the hub in the interface region between these blades and the hub of the fan rotor.

[0014] In ductless architectures, strong vibrational excitations can occur at very high rotational speeds due to the influence of the engine mounted on the aircraft and the direction of the infinite upstream flow. In fact, ductless engines are affected by the ground and fuselage, causing the power supply to the propeller (in terms of flow velocity) to be distorted according to the engine's orientation. This results in a vibrational response of the blades, particularly at the first engine stage 1N, 2N, and 3N. On the other hand, without an air inlet sleeve, the direction of the air flowing through the blades is not parallel to the drive axis, forming a sideslip angle with it. This sideslip angle results in a force known as the "1P" force, which causes the vibrational response of the blades at engine stage 1N. Similarly, these 1P forces can also occur during the aircraft's climb or approach phases, as air flows over the airfoil at an angle of incidence. Summary of the Invention

[0015] The object of this invention is to provide a blade comprising a composite material, which is suitable for use with a variable pitch mechanism and in an "open rotor" type environment, while being able to withstand strong aerodynamic forces under limited space requirements and minimal mass pressure.

[0016] Another object of the present invention is to provide a blade comprising a composite material, which is suitable for use with a variable pitch mechanism and in an "open rotor" type environment, and which is capable of maintaining the airfoil, particularly in high-stress strut regions, in the event of a partial breakage of the blade.

[0017] Another object of the present invention is to provide a blade comprising a composite material, which is suitable for use with a variable pitch mechanism and in an "open rotor" type environment, and which can be manufactured simply and quickly without requiring a large amount of work.

[0018] Therefore, according to a first aspect of the present invention, a turbine blade is provided, the blade comprising:

[0019] - An aerodynamic profile structure comprising two facing skins, each skin comprising fiber reinforcements densified by a matrix;

[0020] - The main sparsity includes a blade root portion, an airfoil portion, and a strut portion. The blade root portion is configured to be mounted on the hub of the turbine rotor. The airfoil portion is arranged inside the aerodynamic profile structure and located between two skins. The strut portion extends outside the aerodynamic profile structure and is located between the blade root portion and the airfoil portion.

[0021] The airfoil section includes a main body connected to the root portion of the blade and two branches extending radially from the main body; and

[0022] - A structural reinforcement extending from the blade root portion to the airfoil body, the structural reinforcement being fixed to the blade root portion and configured to form a different force path from the strut portion in the event of a fracture of the main beam within the strut portion.

[0023] According to the first aspect, some preferred but non-limiting features of the blade are the following features, used alone or in combination:

[0024] - The structural reinforcement includes a pair of plates that are fixed to the airfoil portion to make contact with the body relative to the blade root portion;

[0025] - The structural reinforcement includes at least one rod that is tightly mounted between the blade root portion and the plate, preferably including two parallel rods;

[0026] - At least one rod is inserted into a through slot on the radial inner surface of the blade root portion;

[0027] - At least one rod is straight;

[0028] - The structural reinforcement includes a shank that is integral with the blade root portion and extends through a channel formed in the strut portion and, where appropriate, in the body of the airfoil portion;

[0029] - The plate is attached to one end of the handle;

[0030] - The passage opens in the support section of the main beam to arrange an annular space around the handle in the support section;

[0031] - The main beam is metal; and / or

[0032] - The blade also includes a filling component comprising an inner cavity housed within the aerodynamic profile structure between two branches of the airfoil portion.

[0033] According to a second aspect, the present invention proposes a fan comprising a hub and blades according to a first aspect, the blades extending radially from the hub, each blade being rotatably mounted relative to the hub about a corresponding pitch axis.

[0034] According to a third aspect, the present invention provides an engine comprising an actuation mechanism and a fan according to a second aspect, the actuation mechanism being adapted to be controlled to rotate the blades about a pitch axis, thereby changing the blade pitch angle.

[0035] According to a fourth aspect, the present invention proposes an aircraft comprising a gas turbine engine according to a third aspect.

[0036] According to a fifth aspect, the present invention provides a method for manufacturing a blade according to a first aspect, the method comprising the following steps:

[0037] S1: Production of main beams and structural reinforcements;

[0038] S3: For example, fiber reinforcements with aerodynamic profiles are produced through three-dimensional weaving;

[0039] S4: Insert the main beam into the fiber reinforcement, such that the blade root portion is located outside the fiber reinforcement, and the airfoil portion is located inside the fiber reinforcement; and

[0040] S5: Arrange the assembly formed by the fiber reinforcement and the main beam in the mold, and inject the matrix into the assembly to obtain the blade.

[0041] Optionally, prior to step S4, the method further includes a step of positioning a filling component made of material, the filling component including an inner cavity located between branches of the airfoil portion of the main beam. Attached Figure Description

[0042] Other features, objects, and advantages of the invention will become apparent from the following description, which is entirely illustrative and non-limiting and should be read in conjunction with the accompanying drawings:

[0043] Figure 1 An example of an engine including a ductless fan is shown schematically.

[0044] Figure 2 The fan blades and the actuation mechanism that enables the change of the blade pitch angle are schematically shown, with the fan blades shown in partial cross-section to show the main beam, filling components and structural reinforcements.

[0045] Figure 3 This is an isometric view of the main beam of the blade according to a first exemplary embodiment of the present invention.

[0046] Figure 4 yes Figure 3 Side view of the main beam.

[0047] Figure 5yes Figure 3 A cross-sectional view of the main beam along a plane passing through the pitch axis Y and the main beam branch.

[0048] Figure 6 This is a cross-sectional view of the main beam of the blade according to a second exemplary embodiment of the present invention.

[0049] Figure 7 This is a flowchart illustrating the steps of a method for manufacturing a blade according to one embodiment.

[0050] Figure 8 An example of an aircraft including an engine according to an embodiment of the present invention is shown.

[0051] In all the figures, similar elements have the same reference numerals. Detailed Implementation

[0052] exist Figure 1 The engine 1 shown is an "open rotor" type engine, a configuration commonly referred to as a "thruster" configuration (i.e., a ductless fan is arranged at the rear of the power generator, with the air inlet located at...). Figure 1 (The right side of the middle).

[0053] Engine 1 includes a nacelle 2 and a ductless fan 3 (or propeller), the nacelle being designed to be fixed to the fuselage of aircraft 100. Fan 3 includes two counter-rotating fan rotors 4 and 5. In other words, when engine 1 is running, fan rotors 4 and 5 are driven to rotate in opposite directions relative to nacelle 2 about the same axis of rotation X (which coincides with the main axis of the engine).

[0054] exist Figure 1 In the example shown, engine 1 is an "open rotor" type engine with a "pusher" configuration, which has fan rotors rotating in opposite directions. However, the invention is not limited to this configuration. The invention is also applicable to "open rotor" type engines with a "puller" configuration (i.e., the fan is arranged upstream of the power generator, with the air inlet located in front of, between, or behind the two fan rotors).

[0055] Furthermore, the present invention is also applicable to engines with different architectures, such as those including a fan rotor (including blades) and a fan stator (including impeller blades), or an architecture including a single fan rotor.

[0056] This invention is applicable to turboprop engine type architectures (including single fan rotors) and wind turbine rotors.

[0057] In this application, the axis of rotation of the fan rotors 4 and 5 of the fan 3 (or propeller) is referred to as axis X. The axial direction corresponds to the direction of axis X, and the radial direction is the direction perpendicular to and passing through axis X. Each blade 7 is rotatably mounted relative to the hub 6 about a corresponding pitch axis Y: the pitch axis Y extends radially in general relative to axis X. Finally, "inner" (or internal) and "outer" (or external) are used with reference to the radial direction, such that the inner portion or inner surface of the element is closer to axis X than the outer portion or outer surface of the same element.

[0058] Therefore, blade 7 will be defined relative to the axis X of the fan rotor (on which the blade is intended to be mounted) and the pitch axis Y of the blade. The “chord” here will be understood as the line segment, in approximately axial direction, connecting the leading edge of blade 7 to the trailing edge of the blade for a given section of blade 7 (and therefore for a given point on the pitch axis Y).

[0059] exist Figure 1 In this configuration, each fan rotor 4, 5 includes a hub 6 (or blade assembly hub) rotatably mounted relative to the nacelle 2 and a plurality of blades 7 fixed to the hub 6. The blades 7 extend generally radially relative to the axis of rotation X of the fan rotor.

[0060] like Figure 2 As shown, the fan 3 also includes an actuation mechanism 8 that enables a coordinated change in the pitch angle of the fan rotor blades 7 to adjust the engine performance to suit different flight phases. For this purpose, each blade 7 includes an attachment member 9 (or blade hub) disposed at the blade root. The attachment member 9 is rotatably mounted relative to the hub 6 about the pitch axis Y. More precisely, the attachment member 9 is rotatably mounted within a receiving portion 10 disposed in the hub 6 via balls 11 or other rolling elements.

[0061] The attachment member 9 includes a wall having an outer surface with a rotary shape. The outer surface has two circular slots adapted to form a rolling path for balls or other rolling elements.

[0062] As a variation, each blade 7 may include a cylindrical blade root configured to be directly connected to the hub 6 via a bearing.

[0063] The actuation mechanism 8 includes, for example, an actuator 12 comprising a body 13 fixed to the hub 6 and a rod 14 adapted to be driven to translate relative to the body 13. The actuation mechanism 8 also includes an annular slider 15 mounted to the rod 14 and a pin 16 mounted to the attachment member 9. The pin 16 is adapted to slide in and rotate relative to the annular slider 15 to switch between translational movement of the rod 14 and rotational movement of the attachment member 9, thereby corresponding to rotational movement of the blade 7 relative to the hub 6 about the blade's pitch axis Y.

[0064] The blade 7 includes an aerodynamic profile structure 20 and a main beam 21, which is adapted to be arranged in the airflow to generate lift when the engine 1 is running.

[0065] The aerodynamic profile structure 20 includes two skins 22 connected to each other and extending generally facing each other. The skins 22 are shaped to together define the inner arc surface, outer arc surface, leading edge, and trailing edge of the blade 7. In a manner known per se, the leading edge is configured to extend facing the airflow entering the engine 1. The leading edge corresponds to the front portion of the aerodynamic profile, which faces the airflow and divides it into an inner arc surface flow and an outer arc surface flow. The trailing edge corresponds to the rear portion of the aerodynamic profile, where the inner and outer arc surface flows converge.

[0066] The skin 22 of the aerodynamic profile structure 20 is made of a composite material, which includes fiber reinforcements 23 densified through a matrix (see in particular). Figure 2 Therefore, the skin is monolithic and, according to a non-limiting embodiment, is manufactured as a single piece. As variations not shown, one type of fiber reinforcement for the inner curved surface and another type of fiber reinforcement for the outer curved surface can be considered.

[0067] The fiber reinforcement 23 may be formed from a single fiber preform having varying thickness. The fibers of the fiber reinforcement 23 include at least one of the following materials: carbon, glass, aramid, polypropylene, and / or ceramic. The fiber reinforcement 23 may comprise a woven (two-dimensional or three-dimensional woven), braided, knitted, or laminated fiber arrangement. The matrix typically comprises an organic (thermosetting, thermoplastic, or elastomer) material or a carbon matrix. For example, the matrix may comprise a plastic material, typically a polymer (e.g., epoxide, bismaleimide, or polyimide).

[0068] The main sparsity 21 includes a blade root portion 24 extending outside the aerodynamic profile structure 20, an airfoil portion 25 disposed inside the aerodynamic profile structure 20 between two skins 22, and a strut portion 26 extending between the blade root portion 24 and the airfoil portion 25. The blade root portion 24 is configured to be inserted into the hub 6 via an attachment portion 9 where appropriate. The strut portion 26 corresponds to the area of ​​the main sparsity 21 extending between the outlet of the hub 6 (at the level of the support) and the aerodynamic structure 20. The airfoil portion 25, together with the aerodynamic profile structure 20, forms the airfoil of the blade 7.

[0069] The main beam 21 can be made of metal and is a single piece: the blade root portion 24, the airfoil portion 25, and the strut portion 26 are thus integral. The metal material of the main beam 21 can include at least one of the following materials: steel, titanium, titanium alloys (especially TA6V, which includes titanium, aluminum, vanadium, and trace amounts of carbon, iron, oxygen, and nitrogen), nickel-based superalloys (such as Inconel), and aluminum alloys. The manufacture of the metal main beam 21 can involve a variety of specific methods, such as machining, forging, forming, casting, or additive manufacturing (3D printing).

[0070] As a variation, the main beam 21 may include a composite material comprising fiber reinforcements densified by a matrix. Similar to the aerodynamic profile structure 20, the matrix of the main beam 21 typically comprises an organic (thermosetting, thermoplastic, or elastomer) material or a carbon matrix. For example, the matrix may comprise a plastic material, typically a polymer (e.g., epoxide, bismaleimide, or polyimide). The fibers of the fiber reinforcements of the main beam comprise at least one of the following materials: carbon, glass, aramid, polypropylene, and / or ceramic. The fiber reinforcement 23 may comprise a woven (two-dimensional or three-dimensional woven), braided, knitted, or laminated fiber arrangement. The matrix of the main beam 21 and the matrix of the aerodynamic profile structure may be the same where appropriate. The fibers of the fiber reinforcements of the main beam 21 may be made of the same or different materials as the fibers of the fiber reinforcement 23.

[0071] To resist strong aerodynamic forces, the airfoil portion 25 of the main sparsity 21 includes a main body 27 connected to the blade root portion 24 and two branches 28 extending radially from the main body 27. The main body 27 and branches 28 of the airfoil portion 25 together form the core of the blade 7.

[0072] Branch 28 extends radially from the top of the main body 27, gradually moving away from the pitch axis Y. Therefore, branch 28 is slightly offset in the direction of its free ends 30 from the radially inner ends 29 fixed to the main body 27. Consequently, the distance between the free ends 30 of branch 28 is greater than the distance between the radially inner ends 29 of branch 28. For example, the maximum distance between the free ends 30 of branch 28 can be between 50% and 80% of the maximum chord of blade 7, while the maximum distance between the radially inner ends 29 (measured at the level of the blade as described below) can be between 20% and 50% of the maximum chord.

[0073] From a mechanical perspective, the two branches 28 of the main beam 21 allow for increased tensile stiffness of the blade 7 structure while maintaining reduced mass. Another advantage of this geometry of the main beam 21 (where the branches 28 extend radially and in the chord direction) is that it retains the skin 22 of the aerodynamic profile structure 20 through cross-sectional constraints. In fact, under the action of radially oriented centrifugal forces, the skin 22 is pressed against the inclined surfaces of the branches 28. Therefore, the retention of the skin 22 is ensured not only by the strength of a “weak” interface, such as an adhesive or resin interface that ensures cohesion between the two substrates.

[0074] In one embodiment, the width of each branch 28 (the dimension along the chord of the blade 7 for a given height) is substantially constant between the radially inner end 29 of the branch and the free end 30 of the branch, with a difference of less than 10%. The thickness of the branch (the dimension along the axis perpendicular to the chord for a given height) decreases from the radially inner end 29 of the branch in the direction of the free end 30 of the branch.

[0075] The geometry of branch 28 is chosen to follow the geometry of cortex 22, and thus the geometry of blade 7. In other words, when blade 7 is twisted, branch 28 is also twisted to follow the geometry of blade 7 (see, for example, [link to relevant documentation]). Figure 4 Therefore, the two branches 28 do not necessarily extend in the same plane, and preferably follow the curvature of the blade 7.

[0076] The height of the airfoil portion 25 of the main sparsity 21 (the dimension along the pitch axis Y) can be between 20% and 75% of the height of the aerodynamic structure, for example, about 35%.

[0077] Alternatively, the main beam 21 may include a greater number of branches 28. The additional branches 28 then extend radially from the body 27.

[0078] When the main beam 21 incorporates composite materials, its construction is adapted to the type of material used. Specifically, compared to a metal main beam 21, the branches 28 can extend further in the chord direction of the blades 7. Furthermore, various fiber reinforcements can be used to produce the main beam 21, thus avoiding machining, and the working cross-section of the main beam 21 will be larger compared to a metal main beam 21.

[0079] The blade root portion 24 may have a bulbous shape, that is, a generally raised or curved shape, which extends around the pitch axis Y. If a mid-plane perpendicular to the pitch axis Y and passing through the maximum cross-section of the bulbous portion is defined, the bulbous portion of the blade root portion 24 may, for example, have a generally circular cross-section.

[0080] The main beam 21 may be hollow. As a variation, only a portion of the main beam 21 (e.g., branch 28) may be hollow, while the rest of the main beam 21 may be solid. According to another variation, the entire main beam 21 may be solid.

[0081] The blade 7 also includes a structural reinforcement 31 extending from the blade root portion 24 to the body 27 of the airfoil portion 25. This structural reinforcement 31 is specifically fixed to the blade root portion 24 to create a different force path than the strut portion 26 of the main beam 21 should the main beam 21 break within it. Due to this different force path, the structural reinforcement 31 allows the airfoil of the blade 7 to be maintained even if the main beam 21 breaks within its strut portion 26 (which is subjected to significant stress). The structural reinforcement 31 particularly reduces the risk of one or more cracks / fissures arising from fatigue within the strut portion 26 or during bird strikes.

[0082] The structural reinforcement 31 includes a contreplaque 32 fixed to the airfoil portion 25 to contact the body 27 opposite to the blade root portion 24. Preferably, the contreplaque 32 is generally symmetrical with respect to the pitch axis Y and centered on the pitch axis Y. As an example, the contreplaque 32 is pressed against the radially outer surface 271 of the body 27, between the radially inner ends 29 of the branches 28 of the airfoil portion 25 of the main sparsity 21.

[0083] In a first embodiment, the structural reinforcement 31 includes at least one rod 34 tightly mounted between the blade root portion 24 and the counterplate 32. Preferably, the structural reinforcement 31 includes two rods 34 mounted in parallel within the main beam 21.

[0084] In this embodiment, a slot 33 is formed in the main beam 21, configured to accommodate a corresponding rod 34 of the structural reinforcement 31. Each slot 33 generally extends along the pitch axis Y. In the case where the structural reinforcement 31 includes two slots 33, the slots 33 extend symmetrically on both sides of the pitch axis Y. Furthermore, each slot 33 extends from the blade root portion 24 to the core of the airfoil portion 25 and leads to the radially inner surface 241 of the blade root portion 24 (corresponding to the surface abutting against the bottom of the attachment member 9 or the bottom of the hub 6) and the radially outer surface 271 of the body 27 facing the opposing plate 32. Therefore, the slot 33 is a through slot.

[0085] Preferably, the rod 34 and the slot 33 are straight to facilitate the manufacture and assembly of the blade 7. The slot 33 can be made, for example, by drilling holes in the main beam 21, since the main beam is metal.

[0086] Each rod 34 may include a screw, which can be secured to the plate 32 using a nut 35.

[0087] To create an alternative force path in the event of a fracture of the support portion 26 of the main beam 21, the plate 32 includes two through holes, each accommodating a rod 34. The plate is applied against the radially outer surface 271 of the body 27, with each through hole facing the outlet of a corresponding slot 33. Each screw is then inserted into one of the through holes and into the corresponding slot 33, such that the head of each screw rests against the plate 32. The free end of each screw then protrudes from the corresponding slot 33. Nuts 35 are then screwed onto the free end of each screw and tightened to apply prestress to the plate 32 by applying pressure in the direction of the body 27 (without crushing the plate). Therefore, in the event of a fracture of the support portion 26, the force through the rod 34 (which extends radially relative to the fan rotor) increases significantly under the action of centrifugal force, which has the effect of pressing against the body 27 against the plate 32 and thus causing the airfoil to move slightly radially inward. Therefore, the airfoil is maintained while generating an imbalance (mechanical and / or aerodynamic imbalance), which can be detected by a system used to monitor engine behavior.

[0088] Of course, the screw and nut 35 can be installed in the other direction, and then the head of the screw contacts the radial inner surface 241 of the blade root portion 24, while the free end of the screw is fixed to the plate 32 by the nut 35.

[0089] In this embodiment, the plate 32 may include a compressible material. The plate 32 may be made of an elastomer. As a variation, the plate 32 may include an organic honeycomb structure (e.g., Nomex). ®The honeycomb structure can be of the type comprising aramid fibers calendered into sheets and coated with phenolic resin, or a honeycomb structure comprising one of the following materials: (poly(terephthalamide) (Kevlar type), glass fiber, aluminum). According to another variation, plate 32 is made of metal, for example, of the same material as main beam 21.

[0090] exist Figure 6 In the second embodiment illustrated by way of example, the structural reinforcement 31 is integral with the main beam 21. For this purpose, the structural reinforcement 31 includes a shank 36 integral with the blade root portion 24, and the shank extends along the pitch axis Y through a channel 37 formed in the strut portion 26 and optionally in the body 27 of the airfoil portion 25. Preferably, the shank 36 also extends through the body 27.

[0091] Therefore, the shank 36 forms a section of the support portion 26 and the main body 27. The shank is integral with the blade root portion 24. On the other hand, the shank is separate from and distinct from the rest of the support portion 26 and the airfoil portion 25 of the main beam 21. The shank 36 also protrudes from the radially outer surface 271 of the main body 27, and the counterplate 32 is connected to the free end of the shank 36.

[0092] The handle 36 and the counter plate 32 can be integral. As a variation, the counter plate 32 can be added and fixed to the handle 36, for example, by welding or using fastening members such as screws.

[0093] As long as this embodiment does not require the application of prestress to the plate 32, the plate can be made of the same material as the main beam 21 (and the handle 36).

[0094] An integral main beam 21, having a handle 36 and, where appropriate, a counterplate 32, can be obtained specifically by 3D printing. Where appropriate, layers made of compressible material can be fixed to the radially inner surface of the counterplate 32 and / or the radially outer surface 271 of the body 27 to cushion the impact of the counterplate 32 on the body 27 of the main beam 21 when a secondary force path is initiated (i.e., in the event of breakage of the support portion 26). Where appropriate, layers made of compressible material can also be fabricated by 3D printing and integrally formed with the counterplate 32 and / or the body 27. For example, layers made of compressible material can include organically derived foams (polyacrylamide, polyethylene terephthalate (PET), polyvinyl chloride (PVC), polyetherimide (PEI), polyethylene, carbon, polyisocyanurate, polyurethane, etc.) or metal-derived foams (particularly made of aluminum alloy), or Nomex made of Kevlar, glass fiber, or aluminum. ® A honeycomb structure of this type.

[0095] Optionally, the channel 37 may open within the strut portion 26 of the main beam 21 to arrange a space 38 around the shank 36. The space 38 may, in particular, have an annular shape with a generally circular cross-section near the blade root portion 24, and the cross-section of the space 38 (in a plane perpendicular to the pitch axis Y) decreases in the direction of the body 27. The space 38 may be empty, or, as a variation, may be wholly or partially filled with a filling component that may include an inner cavity.

[0096] The cross section of the stem 36 (in a plane perpendicular to the pitch axis Y) is generally constant in the support section 26 and the body 27, and then gradually flares outward near the counterplate 32.

[0097] It should be noted that regardless of the embodiment of the structural reinforcement 31, the shape of the blade root portion 24 of the main beam 21 remains unchanged. In other words, the channel 37 can be formed only in the support portion 26 and the main body 27 of the main beam 21. Furthermore, regardless of the embodiment of the structural reinforcement 31, the external shapes of the support portion 26 and the airfoil portion 25 of the main beam 21 can be the same. On the other hand, when the structural reinforcement 31 is integral with the blade root portion 24, the main body 27 of the airfoil portion 25 is divided into three parts (see in particular). Figure 6 The first part of the main body 27 is formed by the shank 36 of the structural reinforcement 31. The second and third parts correspond to the remainder of the main body 27 and extend around the shank 36. These second parts then each extend in the extension portion of the associated branch 28, so that these second parts can all be separated along the shank 36 and joined together only at the level of the blade root portion of the main beam 21 (see...). Figure 6 ).

[0098] Optionally, the blade 7 also includes a filler component 39 disposed between the two skins 22 of the aerodynamic profile structure 20, between the two branches 28 and the radial outer surface 271 of the body 27 of the airfoil portion 25 of the main beam 21. The filler component 39 serves as a support for the skins 22 of the aerodynamic profile structure 20.

[0099] The filling component 39 can be made of a material including the inner cavity, such as organically derived foam (polyacrylamide, polyethylene terephthalate (PET), polyvinyl chloride (PVC), polyetherimide (PEI), polyethylene, carbon, polyisocyanurate, polyurethane, etc.) or metal-derived foam (especially made of aluminum alloy), or Nomex made of Kevlar, glass fiber, or aluminum. ® A honeycomb structure of this type.

[0100] Examples of manufacturing methods

[0101] The blade 7 according to the present invention can be obtained according to the following steps.

[0102] During step S1, the main beam 21 and structural reinforcement 31 are manufactured.

[0103] The main beam 21 can be produced using any conventional method, including machining, forging, forming, casting, or additive manufacturing (3D printing).

[0104] When the structural reinforcement 31 includes rods 34, one or more slots 33 are formed in the main beam 21, for example, by machining. A previously machined mating plate 32 is applied to the radially outer surface 271 of the body 27, and then rods 34 (typically screws) are inserted into each hole and each associated slot 33. Nuts 35 are then screwed onto the free end of each rod 34 to block one or more rods 34 and mating plates 32 relative to the main beam 21 and to apply prestress to the mating plates 32.

[0105] As a variation, when the main beam 21 is obtained by 3D printing, the structural reinforcement 31 can be produced simultaneously with the manufacturing of the main beam 21. Then, the structural reinforcement 31 and the main beam 21 are obtained by 3D printing, and the structural reinforcement and the main beam are integral.

[0106] When the main beam 21 comprises composite materials, the fiber reinforcement can be produced by three-dimensional weaving on a jacquard loom. During weaving, bundles of warp strands (or warp threads) are arranged in multiple layers. Weft strands T interweave with warp strands C to connect different layers of warp strands C together. The three-dimensional weaving can be a weaving utilizing "interlocking" weaving. "Interlocking" refers to a weaving method in which each layer of weft strands interconnects multiple layers of warp strands with all strands of the same weft column that have the same movement in the weaving plane. Branches 28 can be obtained, for example, by non-interlocking at the level of the radial outer surface 271 of the body 27.

[0107] Other types of known three-dimensional weavings can be used, such as those described in particular in document WO 2006 / 136755.

[0108] During step S2, a filler component 39 is produced and arranged between the two branches 28 of the main beam 21 and positioned to support against the airfoil portion 25 of the main beam 21. Where appropriate, an adhesive film may be applied at the interface between the filler component 39 and the main beam 21.

[0109] The filling component 39 can be produced by injecting foam between the branches 28 of the main beam 21 or by adding and fixing pre-machined filling components 39.

[0110] As a variation, when the main beam 21 is obtained by 3D printing, the infill component 39 can be obtained simultaneously with step S1 (especially when the infill component is metal). In this variation, the infill component 39, the structural reinforcement 31, and the main beam 21 can then be integrated.

[0111] During step S3, a fiber reinforcement 23 of an aerodynamic profile structure 20 is produced. The fiber reinforcement 23 can be produced by three-dimensional weaving on a jacquard loom. During weaving, bundles of warp strands (or warp yarns) are arranged in multiple layers. Weft strands T interweave with warp strands C to interconnect different layers of warp strands C. Three-dimensional weaving can be a weaving method utilizing "interlocking." "Interlocking" refers to a weaving method in which each layer of weft strands connects multiple layers of warp strands to all strands of the same weft post that have the same movement in the weaving plane.

[0112] Other types of known three-dimensional weavings can be used, such as those described in particular in document WO 2006 / 136755.

[0113] To allow the main beam 21 to be inserted (step S4) into the aerodynamic profile structure 20, non-interconnection is performed in the fiber reinforcement 23. Non-interconnection can be performed at the level of the leading or trailing edge of the blade 7, or at the head (opposite to the blade root). Non-interconnection areas are obtained by not connecting the warp strands of two consecutive layers at two different locations, and the non-interconnection areas are separated by the weft strands.

[0114] Then, the fiber reinforcement 23 is shaped (three-dimensionally deformed to give the fiber reinforcement a twisted shape corresponding to its final shape).

[0115] As a variation, the fiber reinforcement 23 of the aerodynamic profile structure 20 may include a two-dimensional woven, braided, knitted, or laminated fiber arrangement. Specifically, the fiber reinforcement 23 of the skin 22 of the aerodynamic profile structure 20 may include a prepreg-laminated composite material. In this variation of the embodiment, non-interconnection is not required.

[0116] During step S4, the fiber reinforcement 23 of the aerodynamic profile structure 20 is arranged around the main beam 21 and the filling member 39 such that the blade root portion 24 and the strut portion 26 are located outside the fiber reinforcement 23, and the airfoil portion 25 is located inside the fiber reinforcement 23.

[0117] When the fiber reinforcement 23 is produced by three-dimensional weaving, it is advantageous to insert the main beam 21 and the filler 39 through non-interconnection of the fiber reinforcement, typically inserting the main beam and the filler from above when non-interconnection is performed at the blade head.

[0118] When the fiber reinforcement 23 is laminated, the prepreg composite sheet can be laminated onto the main beam 21 and the filler 39.

[0119] During step S5, the assembly thus obtained, formed by the main beam 21, the fiber reinforcement 23 of the aerodynamic profile structure 20, and the filling component 39, is arranged in a mold having a cavity with the shape of the final molded component (i.e., blade 7), and plastic material (the “matrix” of the aerodynamic profile structure) is injected into the mold to impregnate the fiber reinforcement 23. The injection of the plastic material can be performed using injection techniques of the RTM or VARRTM type. The injected plastic material is, for example, a thermosetting liquid composition containing an organic precursor of the matrix material. The organic precursor is typically in the form of a polymer (e.g., a resin), optionally diluted in a solvent.

[0120] The plastic material is heated in a manner known per se to induce polymerization, for example, through crosslinking. For this purpose, a mold is arranged in an oven. The resulting part is then demolded and optionally machined to remove excess length and obtain a part with the desired shape, although the fibers of the fiber reinforcement 23 may shrink during the polymerization of the plastic material.

[0121] During step S6, the attachment member 9 can be added and secured to the root of the blade 7, where appropriate, around the blade root portion 24 of the main beam 21. Specifically, the attachment member can be machined to form a cavity whose shape and size correspond to the shape and size of the blade root portion 24. Optionally, the attachment member 9 can be manufactured in two parts so that it can be added and secured around the blade root portion 24 by means of two dedicated rings (e.g., by shrink-fit, threaded connection, welding, or by using clamping collars). Therefore, step S6, which secures the attachment member 9, can be performed before or after injection (step S5). Advantageously, the material constituting the attachment member 9 can be different from the material of the main beam 21.

[0122] Then routine operations can be performed on the blade tip within the 7-range, such as rework by machining, joining anti-friction strips, inserting a defrosting system, or adding metal leading and / or trailing edge guards.

[0123] Therefore, this invention enables the blade 7 to resist mechanical loads encountered during flight that could excite vibration modes in the blade 7. The production of the metal main sparsity 21 also allows for an aerodynamic lift-to-drag ratio at the bottom of the airfoil, enabling proper supply to the low-pressure compressor downstream of the fan using a conventional annular air inlet sleeve and avoiding the use of an offset fan sleeve. Furthermore, the skin 22 of the aerodynamic profile structure is maintained not only by the matrix injected into the fiber reinforcement after insertion of the main sparsity 21, but also by limiting the cross-section surrounding the metal main sparsity 21. From a mechanical perspective, the use of this metal main sparsity 21 is also advantageous in terms of torsional stiffness relative to its finite mass. From an industrial perspective, manufacturing, monitoring, and assembly operations are significantly reduced compared to using a main sparsity 21 made of composite materials. Finally, the structural reinforcement 31 enables prevention of the consequences of strut breakage of the blade 7 and allows for pilot warning of malfunctions.

Claims

1. A blade (7) of a turbomachine, comprising: - an aerodynamic profile (20) comprising two facing skins (22) comprising fibrous reinforcements (23) densified by a matrix; - a main beam (21) comprising a blade root portion (24) configured to be mounted on a hub of a rotor of the turbomachine, an airfoil portion (25) arranged inside the aerodynamic profile (20) and between the two skins (22), and a strut portion (26) extending outside the aerodynamic profile (20) and between the blade root portion (24) and the airfoil portion (25), - the airfoil portion (25) comprising a main body (27) connected to the blade root portion (24) and two branches (28) extending radially from the main body (27); - the blade (7) being characterized in that it further comprises a structural reinforcement (31) extending from the blade root portion (24) to the main body (27) of the airfoil portion (25), the structural reinforcement (31) being fixed to the blade root portion (24) and configured to form a force path different from the strut portion (26) in the event of breakage of the main beam (21) within the strut portion (26).

2. The vane (7) according to claim 1, wherein - the structural reinforcement (31) comprising a counterplate (32) fixed to the airfoil portion (25) to come into contact with the main body (27) opposite the blade root portion (24).

3. The vane (7) according to claim 2, wherein - the structural reinforcement (31) comprising at least one rod (34) tightly mounted between the blade root portion (24) and the counterplate (32).

4. The vane (7) according to claim 3, wherein - the at least one rod (34) being inserted into a through slot (33) opening onto a radially inner surface (241) of the blade root portion (24).

5. The vane (7) according to claim 3 or 4, wherein - the at least one rod (34) being rectilinear.

6. The vane (7) according to claim 1 or 2, wherein - the structural reinforcement (31) comprising a shank (36) integral with the blade root portion (24) and extending through a passage (37) formed in the strut portion (26).

7. The vane (7) according to claim 2, wherein - the structural reinforcement (31) comprising a shank (36) integral with the blade root portion (24) and extending through a passage (37) formed in the strut portion (26), the counterplate (32) being connected to one end of the shank (36).

8. The vane (7) according to claim 6, wherein - the passage (37) opening in the strut portion (26) of the main beam (21) to arrange an annular space (38) around the shank (36) in the strut portion (26).

9. The vane (7) according to any one of claims 1 to 4, wherein - the main beam (21) being metallic.

10. The vane (7) according to any one of claims 1 to 4, wherein - the blade further comprising a filler member (39) comprising an internal cavity housed in the aerodynamic profile (20) between the two branches (28) of the airfoil portion (25).

11. The vane (7) according to claim 3, wherein The structural reinforcement (31) comprises two parallel bars (34).

12. The vane (7) according to claim 6, wherein The shank extends through a channel (37) formed in the strut portion (26) and in the main body (27) of the airfoil portion (25).

13. The vane (7) according to claim 7, wherein The shank extends through a channel (37) formed in the strut portion (26) and in the main body (27) of the airfoil portion (25).

14. A fan (3) comprising a hub (6) and a blade (7) according to any one of claims 1 to 13, the blade extending radially from the hub (6), each blade (7) being rotatably mounted with respect to the hub about a respective pitch axis (Y).

15. A gas turbine engine (1) comprising an actuation mechanism (8) and a fan (3) according to claim 14, the actuation mechanism being adapted to be controlled to rotate the blades (7) about the pitch axis (Y) so as to vary the pitch angle of the blades (7).

16. An aircraft (100) comprising at least one gas turbine engine (1) according to claim 15.

17. A method for manufacturing a blade (7) according to any one of claims 1 to 13, the method comprising the following steps: S1 : producing the main spar (21) and the structural reinforcement (31); S3: producing a fiber reinforcement (23) of the aerodynamic profile structure (20); S4: inserting the main spar (21) into the fiber reinforcement (23) so that the blade root portion (24) is located outside the fiber reinforcement (23) and so that the airfoil portion (25) is located inside the fiber reinforcement (23); and S5: arranging the assembly formed by the fiber reinforcement (23) and the main spar (21) in a mold and injecting a matrix into the assembly, thereby obtaining the blade (7).

18. The method of claim 17, wherein, The method further comprises, before step S4, a step of positioning a filler part made of material, the filler part comprising an internal cavity located between the branches (28) of the airfoil portion (25) of the main spar (21).

19. The method of claim 17, wherein, In step S3, the fiber reinforcement (23) of the aerodynamic profile structure (20) is produced by three-dimensional weaving.

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