Variable pitch blade for aircraft turbine engine propeller

Abstract

The invention relates to a variable pitch blade (10) for an aircraft turbine engine propeller, comprising a blade (12) connected by means of a strut (30) to a root (14) centered on a pitch axis (A) of the blade, in which the blade (10) is made of a composite material based on at least one metal body (50) and a fibrous preform (52), the fibrous preform (52) is obtained by three-dimensional fibre weaving and is embedded in a polymer matrix which ensures that the preform (52) is rigidly attached to the metal body (50), and wherein the metal body (50) forms at least the root (14) and the strut (30) and the fibrous preform (52) forms at least the blade (12), the blade comprises a pressure side (12a) and a suction side (12b) connected to each other by a leading edge (12c) and a trailing edge (12d) of the blade (12).

Description

Technical Field

[0001] This invention relates to the field of aircraft turbine engines, and more particularly to propulsion propellers of such turbine engines including variable pitch blades. Background Technology

[0002] Prior art specifically includes documents FR-A1-3 017 163, FR-A1-3 080 322, FR-A1-3 112 819, FR-A1-3 121 474, US-A-4, 524, 499, FR-A1-3 120 249, US-A1-2023 / 0801843, and WO-A1-2023 / 031522.

[0003] To minimize pollutant emissions from air transport, it is necessary to improve the efficiency of all aspects of the turbine engine propulsion system, more specifically, propulsion efficiency, which is the efficiency of converting the energy imparted to the air passing through the engine into useful thrust.

[0004] The first-order components affecting propulsion efficiency are those associated with the low-pressure section of the propulsion system, which directly contribute to thrust generation: the low-pressure turbine, the low-pressure transmission system, the fan, and the secondary flow that guides the fan airflow. A known guiding principle for improving propulsion efficiency is to reduce the fan's compression ratio, thereby reducing the flow velocity at the engine exit and the associated kinetic energy losses.

[0005] One of the main consequences of the reduced flow velocity at the engine outlet is that a larger mass of air must be handled in the low-pressure section (secondary flow) to ensure a given thrust level set according to the aircraft's characteristics; therefore, this leads to an increase in the engine's bypass ratio. The bypass ratio, or BPR, is defined as the ratio between the mass flow rate through the secondary flow (cold flow) and the mass flow rate through the main flow (hot flow) that specifically supplies the combustion chamber.

[0006] The direct effect of this increase in secondary airflow is the need to increase the fan diameter, and therefore the external dimensions of the surrounding retaining casing and the nacelle that constitute the aerodynamic envelope of the casing in question. To achieve a high bypass ratio, the casing would become too large and heavy (and generate significant drag), thus necessitating its removal and replacement with a configuration featuring a ductless propeller. Several concepts for ductless turbine engines can be envisioned, such as the ductless single-fan (USF) architecture: a ductless turbine engine comprising (at least) an upstream propeller wheel (or “open fan”) with variable pitch and a downstream stator impeller (stator) with fixed or variable pitch.

[0007] The technical field of this invention relates to ducted or unducted variable pitch rotating (rotor) propeller or fan blades with potential applications in the aerospace propulsion industry. Other examples of architectures of particular interest include: counter-rotating open rotors (CROR) and turboprop engines.

[0008] In ductless engines (USF, CROR, and turboprop engines), strong vibration excitation occurs at high speeds due to the engine mounting on the aircraft and the direction of the upstream flow at infinity. In fact, ductless engines are affected by the ground and fuselage, causing a distortion in the propeller supply velocity with the engine azimuth angle. This results in vibrational responses in the propeller blades at the first (1N), second (2N), and third (3N) orders (potentially higher) of the engine.

[0009] On the other hand, without an intake sleeve, the direction of air flowing through the blades is not parallel to the engine shaft. This sideslip angle generates a force known as "1P" (1 / revolution), which causes the propeller blades to produce a vibrational response at the "1N" engine order. In a similar manner, these "1P" forces may also occur during the aircraft's climb or approach phase, as air flows over the blades at an angle of incidence. These high-speed vibrations excite very high stress cycles throughout the blades. In particular, the section at the blade root between the hub and the duct (also known as the "strut") is a load-bearing and critical area due to its function of holding the blades in place.

[0010] The purpose of this invention is to provide a blade that can maintain a "1P" load at its root while keeping the blade's mass to a minimum. Summary of the Invention

[0011] This invention proposes a variable-pitch blade for an aircraft turbine engine propeller, the blade comprising a blade connected by a strut to a root centered on the blade's pitch axis, the blade being made of a composite material consisting of at least one metal body and a fiber preform, the fiber preform being obtained by three-dimensional fiber weaving and embedded in a polymer matrix, the polymer matrix ensuring the preform is fixed to the metal body, the metal body forming at least the root and the strut, and the fiber preform forming at least the blade, the blade comprising a pressure side and a suction side connected together by the blade's leading and trailing edges, characterized in that the metal body further forms a first portion of a sparsity extending inside the blade along the pitch axis, the sparsity comprising a second portion formed by the preform and extending inside the blade along the pitch axis, the second portion of the sparsity having a longitudinal end located on the root side, the longitudinal end engaging the first portion of the sparsity.

[0012] The proposed solution consists of 3D-woven composite blades and a metal root. The unique aspect of this concept is that the blade root and struts are made of metal below the aerodynamic duct. Another feature is the construction of the spars located inside the blade. Instead of a wedge effect on the outside of the root, the composite blade is held by the metal body manufactured inside the blade by joining a portion of the preform to a portion of the metal body. This embodiment provides an alternative mode of holding the composite blade by the metal root. This holding mode reveals that a portion of the composite blade, acting as a spars, transmits forces to the metal root. Therefore, there is less composite skin structure on the aerodynamic surfaces (pressure and suction sides) of the preform near the blade, as the composite skin only needs to reconstruct the aerodynamic profile in areas of lower pressure load.

[0013] Therefore, compared with the prior art, the blade spars of the present invention comprise two parts, specifically two materials. A first metal part is formed by the body and is integrally formed with or fixed to the root and strut. A second composite part is formed by a preform and is integrally formed with or fixed to the remainder of the preform. The components (body and preform) are also fixed together using a polymer matrix, for example by RTM or an equivalent method.

[0014] The blades according to the invention may include one or more of the following features, individually or in combination with each other: - The preform includes at least one fiber unraveling portion therein, which enables the definition of: +The upper section of the blade, in which the fibers are fixed to each other across the entire thickness of the blade. + The lower section of the blade, connected to the upper section, and comprising two overlapping skins located on one side of the blade's pressure and suction sides, respectively, wherein the fibers of each skin are fixed to each other over the entire thickness of the skin, and The second part of the spar, which is arranged between the two skins of the lower section and includes a longitudinal end opposite the root, the longitudinal end connecting to the upper section, +The first part of the wing spars, which is arranged between the two skins; - The impeller is also made of at least one foam block located inside the blade; - The first foam block is arranged between the skins, between the second part of the spar and the leading edge of the blade, and the second foam block is arranged between the skins, between the second part of the spar and the trailing edge of the blade; - The first part of the spar has a dimension measured along the chord of the blade, which represents 80% to 120% of the dimension of the second part of the spar, measured in the same manner; - The first part of the spar includes two wings located on one side of the pressure side and the suction side of the blade, respectively, and defines the cavity between the two wings. The cavity opens on the second part of the spar along the direction of the blade pitch axis to form a receiving opening for receiving the second part. - The cavity opens at its two opposite ends, which are located on the leading and trailing edges of the blade, respectively; - The first part of the spar includes two baffles located on the leading and trailing edge sides of the blade, respectively. These baffles extend between the wing portions and connect the wing portions together to close the two opposite ends of the cavity located on the leading and trailing edge sides of the blade, respectively. - The reception opening is defined by two longitudinal edges of the wing, which are tapered; -The wing has: +A constant thickness along the pitch axis at at least 50% or even 80% of its height, or +thickness, which varies along the pitch axis and is maximum, for example, at the root side or half-height of the wing; -The wing has: +A constant thickness along the chord of the blade at at least 50% or even 80% of its length, or + The thickness varies along the chord of the blade and is greatest, for example, in the middle of the airfoil; -Wings: + indicates straight lines and parallel lines, or The plus sign is curved and has concave sections that point toward each other. +o has a first boss oriented toward each other and a second boss oriented toward one of the pressure side and suction side of the blade, respectively; - The first part of the spar also includes at least one partition wall of the cavity, which extends between and is spaced apart from the wing portions, or extends from one wing portion to another and connects to these wing portions; - The wall has a height along the pitch axis, which represents between 30% and 60% of the height of the wing along that axis; -The wall is connected to the partition; -Metal body: + forms a part, + It is formed by at least two parts, the first part including a root, a pillar and a wing, and the second part including the other wing; - The second part of the spar includes at least one fiber unraveling portion over at least a portion of its length, and the first part of the spar includes at least one protrusion that engages in the second part of the spar at the level of the unraveling portion; - The second part of the wing spars is connected to the skin via a fiber mesh.

[0015] The present invention also relates to a turbine engine, particularly a turbine engine for an aircraft, the turbine engine comprising a propeller comprising at least one blade as described above.

[0016] Finally, the present invention relates to a method for manufacturing the impeller as described above, wherein the method comprises the following steps: a) The fibers are woven in three dimensions to form a preform, and a portion of the preform is unraveled to form the second part of the spar. b) Compact the second part of the wing spars. c) Join the second part of the spar into the cavity of the first part of the spar. d) Compact the components formed by the preform and the main body. e) RTM-type consolidation of components in a mold, including, for example, impregnating preforms with resin in a mold. Attached Figure Description

[0017] Further features and advantages will become apparent from the following description of non-limiting embodiments of the invention with reference to the accompanying drawings, in which: [ Figure 1 ] Figure 1 This is a schematic perspective view of a variable-pitch blade used in a turbine engine propeller for an aircraft. [ Figure 2 ] Figure 2 yes Figure 1 A larger scale view of a portion, showing the root of the blades. [ Figure 3 ] Figure 3 This is a schematic axial cross-sectional view of an angular pitch setting system for variable pitch blades. [ Figure 4 ] Figure 4 This is a schematic cross-sectional view of the blade according to the first embodiment of the present invention. [ Figure 5 ] Figure 5 It is along Figure 4 A schematic partial cross-section of line VV in the diagram. [ Figure 6 ] Figure 6 It is along Figure 4 or Figure 5 A schematic cross-sectional view of line VI-VI. [ Figure 7 ] Figure 7 It is similar to Figure 6 The schematic diagram illustrates an alternative embodiment. [ Figure 8 ] Figure 8 It is similar to Figure 6 The schematic diagram illustrates another variation of the embodiment. [ Figure 9 ] Figure 9 It is similar to Figure 6 The schematic diagram illustrates another variation of the embodiment. [ Figure 10 ] Figure 10 It is similar to Figure 5 The schematic diagram illustrates another variation of the embodiment. [ Figure 11 ] Figure 11 It is similar to Figure 5 The schematic diagram illustrates another variation of the embodiment. [ Figure 12 ] Figure 12 This is a schematic cross-sectional view of a blade according to another embodiment of the present invention. [ Figure 13 ] Figure 13 It is along Figure 12 A schematic cross-section taken by line XIII-XIII in the diagram. [ Figure 14 ] Figure 14 It is along Figure 12 A schematic partial cross-section of line XIV-XIV in the diagram. [ Figure 15 ] Figure 15 It is similar to Figure 13 The schematic diagram illustrates another variation of the embodiment. [ Figure 16 ] Figure 16 This is a schematic cross-sectional view of a blade according to another embodiment of the present invention. [ Figure 17 ] Figure 17 It is along Figure 16 A schematic cross-section taken by line XVII-XVII in the diagram. [ Figure 18 ] Figure 18 It is similar to Figure 17 The schematic diagram illustrates another variation of the embodiment. [ Figure 19 ] Figure 19 It is similar to Figure 17 The schematic diagram illustrates another variation of the embodiment. [ Figure 20 ] Figure 20 This is a schematic cross-sectional view of a blade according to another embodiment of the present invention, and [ Figure 21 ] Figure 21 It is similar to Figure 5The schematic diagram illustrates another variation of the embodiment. Detailed Implementation

[0018] Figure 1 A variable-pitch blade 10 for an aircraft turbine engine propeller is shown; the propeller may be ducted or unducted.

[0019] The whorl 10 includes blades 12 connected to the root 14.

[0020] The blade 12 has an aerodynamic profile and includes a pressure side 12a and a suction side 12b, which are connected by an upstream leading edge 12c and a downstream trailing edge 12d. The terms "upstream" and "downstream" refer to the airflow direction around the blade during operation.

[0021] The blade 12 has a free upper end, referred to as the top, and a lower end connected to the root 14.

[0022] In the example shown, the blade 10 is made of composite material by an injection molding process known as RTM (Resin Transfer Molding). This method involves preparing a fiber preform 18 through three-dimensional weaving, then placing the preform in a mold and injecting a polymerizable resin (e.g., epoxy resin) that impregnates the preform. After the blade 12 has cured and hardened, the leading edge 12c of the blade 12 is typically reinforced by a metal shield 20, which is assembled and attached, for example, by gluing.

[0023] The blade 10 here includes a sparsity 22, which includes a core forming the blade 12 and a portion configured to be inserted into the preform 18 prior to resin injection, and a portion extending from the side opposite the top of the blade 12 to form a portion of the root 14 (referred to as body 24).

[0024] The spar 22 is typically made of a composite material having an epoxy organic matrix reinforced with 3D woven carbon fibers, wherein the warp direction is predominantly radially oriented and the latitudinal direction is predominantly oriented along the blade chord at the height of the aerodynamic duct. However, the spar can also be a mechanically more advantageous component made of different organic matrix composite materials (thermosetting, thermoplastic, or elastomers) reinforced with long fibers (carbon, glass, aramid, polypropylene) in different fiber arrangements (woven, braided, knitted, unidirectional).

[0025] Although not shown, the blade 12 may be hollow or solid and includes an inner cavity filled with a foam or honeycomb filler material. This filler material is mounted around the spars 22 and covered with an organic matrix composite skin to increase the blade's impact resistance.

[0026] The shield 20 can be made of titanium or titanium alloy, stainless steel, steel, aluminum, nickel, etc. The pressure side 12a or even the suction side 12b of the blade 12 can be covered with a polyurethane membrane to prevent corrosion.

[0027] A is the elongation axis of the blade 10 and the blade 12, and in particular the pitch axis of the blade 10, i.e., the axis around which the angular position of the blade is adjusted. This axis is also typically the radial axis, and thus extends radially from the axis of rotation of the propeller equipped with the blade.

[0028] The main body 24 of the root 14 has in Figure 2 The specific shape can be seen more clearly in the middle.

[0029] The main body 24 basically consists of three parts: -Free end 28, located on the opposite side of blade 12. - Support 30, support 30 is located on the side of blade 12, and - Spherical body 32, located between the free end 28 and the support 30.

[0030] In the example shown, the free end 28 has a generally parallelepiped shape. This end 28 is preferably misaligned or offset relative to axis A to provide keying or indexing.

[0031] The support column 30 can have a relatively complex shape and can be considered to include: - Two lateral surfaces 30a and 30b are located on the pressure side 12a and suction side 12b of the blade 12, respectively. These two lateral surfaces 30a and 30b converge toward each other along axis A and in the direction of the tip of the blade 12. - Two edges, namely upstream edge 30c and downstream edge 30d, which are along axis A and diverge from each other toward the top of blade 12.

[0032] The spherical body 32 has a generally convex or dome shape that extends about axis A.

[0033] The sphere 32 has two peripheral support surfaces extending around axis A, namely a lower support surface 32a and an upper support surface 32b. In the example shown, due to the shape of the sphere, the lower support surface 32a faces downward (i.e., on the opposite side of the blade 12) and radially outward relative to axis A, while the upper support surface 32b faces upward (i.e., toward the blade 12) and radially outward relative to axis A.

[0034] Figure 3An example of a system 34 for setting the pitch angle of the blade 10 is shown. The system 34, which generally comprises a cup-shaped member 36, a ring 38, a stop member 40, and a nut 42, is described in detail in an earlier application WO-A1-2023 / 031522.

[0035] This invention proposes a novel design for a variable-pitch blade 10 for an aircraft turbine engine propeller, the first embodiment of which is described in... Figures 4 to 6 As shown in the image.

[0036] The blade 10 includes blades 12 connected to a root 14 via a strut 30, the root 14 being centered on the blade's pitch axis A. The blade 10 is made of a composite material consisting of at least one metal body 50 and a fiber preform 52, the fiber preform 52 being obtained through three-dimensional fiber weaving and embedded in a polymer matrix, the polymer matrix ensuring the preform 52 is fixed to the metal body 50. The metal body 50 at least forms the root 14 and the strut 30 of the blade 10. The body 50 can be a single component. The fiber preform 52 at least forms the blade 12.

[0037] The particularity of this invention stems from the fact that the metal body 50 also forms a first portion 54a of a spar 54 extending inside the blade 12 along the pitch axis A. The spar 54 includes a second portion 54b formed by a preform 52 and extending inside the blade 12 along the pitch axis A. Furthermore, the second portion 54b of the spar 54 has a longitudinal end 54b1 located on one side of the root 14, engaging in the cavity 56 of the first portion 54a of the spar 54.

[0038] The following Figure 4 Several variations of the embodiment of the blade 10 are shown. In these figures, the orthogonal reference frame XYZ includes an axis X parallel to the axis of rotation of the blade and the propeller carrying the blade (and extending, for example, along the chord of the blade 10), an axis Z extending parallel to the pitch axis A, and an axis Y transverse to and perpendicular to axes X and Z.

[0039] These common elements of the blades 10 bear the same reference numerals and use the reference numerals already used above, provided that these elements are for... Figures 1 to 3 The impeller 10 is public and has already been described above.

[0040] Advantageously, the preform 52 includes at least one fiber unraveling portion, which enables the definition of: - The upper section 52a of the blade 12, in which the fibers are fixed to each other over the entire thickness of the blade 12. - The lower section 52b of the blade is connected to the upper section 52a and includes two overlapping skins 52b1 and 52b2 located on one side of the pressure side 12a and the suction side 12b of the blade 12, respectively, with the fibers of each skin 52b1 and 52b2 fixed to each other over the entire thickness E0 of the skin. Figure 5 ),as well as - The second part 54b of the wing spade 54 is arranged between the two skins 52b1 and 52b2 of the lower section 52b and includes a longitudinal end 54b2 opposite to the root 14 connected to the upper section 52a. The first part 54a of the wing spars 54 is arranged between the two skins 52b1 and 52b2.

[0041] Advantageously, the second part 54b of the spar 54 is stiffer along axis A than the skins 52b1 and 52b2 to avoid transmitting force to the skins 52b1 and 52b2 when the blade 10 bends or is subjected to centrifugal force. On the other hand, the stiffness of the spar 54 along the chord of the blade 12 can be less than the stiffness of the skins 52b1 and 52b2. This can be achieved, for example, by adjusting the warp / weft ratio or the properties of the fibers in the preform 52. As a reminder, in the preform 52, the weft yarns (or weft threads) of the preform 52 are "horizontal" (in the direction of the blade chord), and the warp yarns (or warp threads) of the preform are "vertical" (in the direction of axis A).

[0042] Therefore, the separation between the skins 52b1, 52b2 and the spar 54 outside the cavity 56 is preferably achieved only through a change in material stiffness. There is no interlayer interface between the skins 52b1, 52b2 and the spar 54.

[0043] The blade 10 may also include at least one foam block 58, 60 located inside the blade 12.

[0044] The first foam block 58 can be arranged between the skins 52b1 and 52b2, between the second part 54b of the spar 54 and the leading edge 12c of the blade 12, and the second foam block 60 can be arranged between the skins 52b1 and 52b2, between the second part 54b of the spar 54 and the trailing edge 12d of the blade 12.

[0045] The first portion 54a of the spar 54 may have a dimension L1 measured along the chord of the blade 14, which is 80% to 120% of the dimension L2 of the second portion 54b of the spar 54, measured in the same manner. Figure 4 ).

[0046] Figure 4 It is the cross-section that passes through the blade's skeleton line and is located in the middle between the pressure side surface and the suction side surface of the blade.

[0047] The first part 54a of the spar 54 may include two wings 62a and 62b, which are located on one side of the pressure side 12a and the suction side 12b of the blade 12, respectively, and define a cavity 56 between the two wings. The cavity 56 opens on the second part 54b side of the spar 54 in the direction of the pitch axis A to form a receiving opening for receiving the second part 54b.

[0048] The receiving opening is defined by two longitudinal edges 62a1 and 62b1 of the wings 62a and 62b, which may be tapered, as shown in the attached figure.

[0049] exist Figures 4 to 6 In the embodiment shown, cavity 56 is open at its two opposite ends located on the leading edge 12c side and the trailing edge 12d side of blade 12, respectively.

[0050] This configuration allows the second portion 54b of the spar 54 to be inserted into the first portion 54a of the spar 54 via lateral insertion. In this case, the root 14 provides the following retention function for the blade 12: - Absorbing forces Fy and moments Mx and Mz through complementary shapes. - The forces Fx, Fz and torque My are absorbed through the adhesive interfaces between the various parts 54a and 54b of the spar 54.

[0051] X is oriented along the engine axis, Z is oriented along the blade pitch variation axis, and Y is the vector product of X and Z.

[0052] The thickness E1 of the wing sections 62a and 62b varies along the pitch axis A, and is greatest, for example, on the root 14 side. Figure 5 ).

[0053] The thickness E2 of the wings 62a and 62b varies along the chord of the blade 12, and is greatest, for example, in the middle of the wing. Figure 6 ).

[0054] The wings 62a and 62b are basically straight and parallel. The wings 62a and 62b have protrusions oriented toward the pressure side 12a and suction side 12b of the blade 12, respectively.

[0055] Figure 7 The embodiment shown differs from the previous embodiment in that the first part 54a of the spar 54 includes two baffles 64a and 64b, which are located on the leading edge 12c side and the trailing edge 12d side of the blade 12, respectively.

[0056] These baffles 64a and 64b extend between the wings 62a and 62b and connect the wings 62a and 62b together to close the two opposite ends of the cavity 56 located on the leading edge 12c side and the trailing edge 12d side of the blade 12, respectively.

[0057] In this variant, the root 14 retains the leaf 12 as follows: - Absorbing forces Fx, Fy and moments Mx, My, Mz through complementary shapes. - Absorbs force Fz through the adhesive interfaces between the various parts of the spar.

[0058] Figure 8 The variant of the illustrated embodiment differs from the first embodiment in that the wings 62a and 62b are curved and have recesses facing each other. The wings 62a and 62b have a constant thickness E2 for at least 50% or even 80% of their length along the chord of the blade 12.

[0059] In this variant, cavity 56 faces the leading edge 12c and trailing edge 12d. However, the second portion 54b of the spar 54 is inserted into cavity 56 from above. In this case, the root 14 retains the blade 12 as follows: - Absorbing forces Fx, Fy and moments Mx, My and Mz through complementary shapes - Absorbs force Fz through the adhesive interfaces between the various parts of the spar.

[0060] Figure 9 The variant of the illustrated embodiment differs from the first embodiment in that the wings 62a and 62b have first bosses oriented toward each other and second bosses oriented toward one of the pressure side and suction side of the blade, respectively. The thickness of the wings 62a and 62b varies along the chord of the blade and is, for example, greatest at the middle of the wing.

[0061] Figure 10 The variation of the embodiment shown differs from the first embodiment in that the thickness E1 of the wings 62a and 62b varies along the pitch axis A and is maximum at half the height of the wing.

[0062] In this variant, cavity 56 is open towards the leading edge 12c and trailing edge 12d. This means that the second portion 54b of the spar 54 can be inserted into cavity 56 laterally (from the side) without compressing that portion 54b. In this case, the root 14 retains the blade 12 as follows: - Absorb forces Fy, Fz and moments Mx, My, Mz through complementary shapes.

[0063] - The force Fx is absorbed through the adhesive interfaces between the various parts of the spar.

[0064] Figure 11 The variant shown differs from the first embodiment in that the wings 62a, 62b have a constant thickness at least 50% or even 80% of their height along the pitch axis.

[0065] Figures 12 to 17 The difference between this embodiment and the first embodiment is that the second portion 54b of the spar 54 includes at least one fiber unraveling portion over at least a portion of its length.

[0066] Figures 12 to 14 Another variation of the embodiment is shown, which differs from the first embodiment in that the first portion 54a of the spar 54 further includes at least one partition wall 66 for separating the cavity 56.

[0067] The wall 66 extends between and at a distance from the wings 62a and 62b, and is substantially parallel to the wings.

[0068] Wall 66 is inserted into the disassembly portion of the second part 54b of wing beam 54.

[0069] The wall 66 has a height H1 along the pitch axis A, which represents 30% to 60% of the height H2 of the wings 62a and 62b along the axis A.

[0070] The wall 66 has a dimension L3 measured along the chord of the blade 14, where dimension L3 represents 80% to 120% of dimension L1. In other words, the wings 62a, 62b and the wall 66 can have the same dimensions measured in this way.

[0071] The wall 66 may have a thickness E3, which varies along the pitch axis A and is maximum, for example, on the root 14 side.

[0072] Wall 66 can have a constant thickness E4 along the chord of the blade.

[0073] In this variant, cavity 56 is open towards the leading edge 12c and trailing edge 12d. This allows the second portion 54b of the spar 54 to be inserted laterally into cavity 56 of the first portion 54a. However, since the second portion 54b has a constant thickness along the pitch axis A before being inserted into cavity 56, it can also be inserted from above. The root 14 retains the blade 12 as follows: - Absorb forces Fy, Fz and moments Mx, My, Mz through complementary shapes.

[0074] - The force Fx is absorbed through the adhesive interfaces between the various parts of the spar, since the thickness is constant in the chord direction.

[0075] Figure 15The embodiment shown differs from the previous embodiment in that the first part 54a of the spar 54 includes two baffles 64a and 64b, which are located on the leading edge 12c side and the trailing edge 12d side of the blade 12, respectively.

[0076] The partitions 64a and 64b that extend between the wings 62a and 62b and connect the wings 62a and 62b together are also connected to the wall 66 and close the two opposite ends of the cavity 56 located on the leading edge 12c side and the trailing edge 12d side of the blade 12, respectively.

[0077] In this variant, the second portion 54b of the preform 54 can be inserted through the top of the cavity 56, thus allowing for non-open mouths on both the leading edge 12c and trailing edge 12d sides. The root 12's retaining effect on the blade 12 is as follows: - Absorb forces Fx, Fy, Fz and torques Mx, My, Mz through complementary shapes.

[0078] Figure 16 and Figure 17 Another variation of the embodiment is shown, which differs from the first embodiment in that the first portion 54a of the spar 54 further includes at least one partition wall 68 for separating the cavity 56.

[0079] The wall 68 extends from one wing 62a to another wing 62b and is connected to these wings 62a, 62b.

[0080] The wall 68 is inserted into the disassembly portion of the second part 54b of the wing beam 54.

[0081] The wall 68 has a height H1 along the pitch axis A, which represents 30% to 60% of the height H2 of the wings 62a and 62b along the axis A.

[0082] The wall 68 may have a thickness E3, which varies along the pitch axis A and is maximum, for example, on the root 14 side.

[0083] The wall 66 can have a constant thickness E4 between the pressure side and the suction side of the blade 12.

[0084] The first part 54a of the wing spars 54 also includes two septa 64a and 64b, which are located on the leading edge 12c side and the trailing edge 12d side of the blade 12, respectively.

[0085] These septa 64a, 64b extend between and connect the wings 62a, 62b to close the two opposite ends of the cavity 56, located on the leading edge 12c side and the trailing edge 12d side of the blade 12, respectively. The septa 64a, 64b and the wall 68 are substantially parallel.

[0086] Figure 18 and Figure 19 The alternative embodiment shown differs from the first embodiment in that the second portion 54b of the spar 54 includes at least one fiber unraveling portion 70 over at least a portion of its length.

[0087] The first portion 54a of the spar 54 includes at least one protrusion 72, which engages at the disengagement portion 70 in the second portion 54b of the spar 54. In the example shown, the second portion 54b of the spar 54 includes a first disengagement portion 70 and a second disengagement portion 70, the first disengagement portion 70 being located on the leading edge side of the blade 12 and engaging the first protrusion 72 of the first portion 54a of the spar 54 in the first disengagement portion 70, and the second disengagement portion 70 being located on the trailing edge side of the blade 12 and engaging the second protrusion 72 of the first portion 54a of the spar 54 in the second disengagement portion 70.

[0088] Protrusion 72 is roughly pointed.

[0089] exist Figure 18 and Figure 19 As can be seen, the cross-section of the first part 54b of the preform 54 is approximately H-shaped or X-shaped. The cavity 56 also has a similar overall cross-sectional shape.

[0090] exist Figure 18 In a variant of the illustrated embodiment, the second portion 54b of the spar 54 is connected to the skins 52b1 and 52b2 via fiber meshes 74. These fiber meshes 74 may extend along all or part of the length of the second portion 54b along the pitch axis A.

[0091] For example, Figure 20 The variant embodiments shown are similar to Figure 16 and Figure 17 The difference in the variant shown is that the partition wall 68 is replaced by a rod 76. The rod 76 extends from one wing 62a to the other wing 62b and connects to these wings 62a, 62b.

[0092] The rod 76 is a certain distance from the bottom of the cavity 56, and its cross-section can be circular or elliptical, as shown in the figure.

[0093] The rod 76 passes through an arrangement whose shape is complementary to the end 54b1 of the second part 54b of the spar 54, such as an orifice.

[0094] Figure 21 The variant embodiments shown are similar to Figure 16 and Figure 17 The difference in the variant shown is that the main body 50 is formed by at least two parts 50a and 50b.

[0095] The first part 50a of the main body 50 includes the root 14, the pillar 30, one of the wings 62a or even the wall 66.

[0096] The second part 50b of the main body 50 includes another wing 62b and is attached to the first part by any suitable means or technique (such as welding).

[0097] In other variations not shown: - If the adhesive properties provided by the polymer matrix are insufficient, an adhesive layer can be provided at the interface between portions 54a and 54b of the spar to enhance the adhesive properties. - The impeller can be foam-free and include a relatively thick skin 52b1, 52b2 that occupies a foam volume. - A space or gap may be provided between the end 54b1 of the spar 54 and the bottom of the cavity 56.

[0098] The present invention also relates to a method for manufacturing the impeller 10 according to the invention, the method comprising the following steps: a) The fibers are three-dimensionally woven to form a preform 52, and a portion of the preform 52 is unraveled to form the second part 54b of the spar 54. b) Compact the second part 54b of the wing spars 54. c) The second part 54b of the spar 54 is joined to the cavity 56 of the first part 54a of the spar 54. d) Compact the assembly formed by the preform 52 and the main body 50. e) The components are RTM-type consolidated in the mold.

[0099] RTM (resin transfer molding) bonding is well known to those skilled in the art and will not be elaborated upon here.

[0100] The method may also include one or more of the following steps: waterjet cutting the profile of the preform, inserting one or more foams 58, 60 between the skins of the preform 52, attaching the shield 20 to the leading edge 12c, machining the pressure side 12a and the suction side 12b, coating the pressure side 12a and the suction side 12b with a PU (polyurethane) layer, attaching a heating pad and / or an anti-wear strip to the root 14, etc.

[0101] This invention has many advantages, including: - The roots and supports of the blades are all made of metal. - The composite blade is held within the cavity of the root by the metal root, rather than by a wedge effect outside the root. - Part of the blade acts as a sparsity, transferring forces toward the metal root. - Preforms have fewer skin structures because these skins only need to reconstruct the aerodynamic profile in areas of lower pressure load. -etc.

Claims

1. A variable-pitch blade (10) for an aircraft turbine engine propeller, the blade (10) comprising a blade (12) connected by a strut (30) to a root (14) centered on the pitch axis (A) of the blade, the blade (10) being made of a composite material consisting of at least one metal body (50) and a fiber preform (52), the fiber preform being obtained by three-dimensional fiber weaving and embedded in a polymer matrix, the polymer matrix ensuring that the preform (52) is fixed to the metal body (50), the metal body (50) forming at least the root (14) and the strut (30), and the fiber preform (52) forming at least the blade (12), the blade comprising a pressure side (12a) and a suction side (12b) connected together by a leading edge (12c) and a trailing edge (12d) of the blade (12), characterized in that, The metal body (50) also forms a first portion (54a) of a spar (54) extending inside the blade (12) along the pitch axis (A). The spar (54) includes a second portion (54b) formed by the preform (52) and extending inside the blade (12) along the pitch axis (A). The second portion (54b) of the spar (54) has a longitudinal end (54b1) located on the root (14) side, which engages in a cavity (56) of the first portion (54a) of the spar (54).

2. The blade (10) according to claim 1, wherein, The preform (52) includes at least one fiber unraveling portion thereof, the at least one fiber unraveling portion enabling the definition of: - The upper section (52a) of the blade (12), in which the fibers are fixed to each other over the entire thickness of the blade, - The lower section (52b) of the blade (12), the lower section being connected to the upper section (52a) and comprising two overlapping skins (52b1, 52b2) located on one side of the pressure side (12a) and suction side (12b) of the blade (12), respectively, wherein the fibers of each skin (52b1, 52b2) are fixed to each other over the entire thickness of the skin (52b1, 52b2), and - The second portion (54b) of the wing spar (54), the second portion being disposed between the two skins (52b1, 52b2) of the lower section (52b) and including a longitudinal end (54b1) opposite to the root (14), the longitudinal end being connected to the upper section (52a), The first part (54a) of the wing beam (54) is arranged between the two skins (52b1, 52b2).

3. The blade (10) according to any one of the preceding claims, wherein, The blade is further made of at least one foam block (58, 60) located inside the blade (12).

4. The blade (10) according to claim 3, which is dependent on claim 2, wherein, A first foam block (58) is arranged between the skins (52b1, 52b2) and between the second part (54b) of the spar (54) and the leading edge (12c) of the blade (12), and a second foam block (60) is arranged between the skins (52b1, 52b2) and between the second part (54b) of the spar (54) and the trailing edge (12d) of the blade (12).

5. The blade (10) according to any one of the preceding claims, wherein, The first portion (54a) of the spar (54) has a dimension (L1) measured along the chord of the blade (12), which represents 80% to 120% of the dimension (L2) of the second portion (54b) of the spar (54) measured in the same manner.

6. The blade (10) according to any one of the preceding claims, wherein, The first portion (54a) of the spar (54) includes two wings (62a, 62b) located on one side of the pressure side (12a) and suction side (12b) of the blade (12), respectively, and defines the cavity (56) between the two wings. The cavity opens on the second portion (54b) side of the spar (54) along the direction of the pitch axis (A) to form a receiving opening for receiving the second portion (54b).

7. The blade (10) according to claim 6, wherein, The cavity (56) is open at its two opposite ends, which are located on the leading edge (12c) side and the trailing edge (12d) side of the blade (12), respectively.

8. The blade (10) according to claim 6, wherein, The first part (54a) of the spar (54) includes two baffles (64a, 64b) located on the leading edge (12c) side and the trailing edge (12d) side of the blade (12), respectively. These baffles (64a, 64b) extend between the wings (62a, 62b) and connect the wings together to close the two opposite ends of the cavity (56) located on the leading edge (12c) side and the trailing edge (12d) side of the blade (12), respectively.

9. The blade (10) according to any one of claims 6 to 8, wherein, The receiving opening is defined by two longitudinal edges (62a1, 62b1) of the wings (62a, 62b), which are tapered.

10. The blade (10) according to any one of claims 6 to 9, wherein, The wings (62a, 62b) have: - A constant thickness (E1) along the pitch axis (A) at at least 50% or even 80% of its height, or -Thickness (E1), which varies along the pitch axis (A) and is maximum, for example, at the root (14) side or half-height side of the wing.

11. The blade (10) according to any one of claims 6 to 10, wherein, The wings (62a, 62b) have: -A constant thickness (E2) along the chord of the blade (12) at at least 50% or even 80% of its length, or - The thickness varies along the chord of the blade (12) and is greatest (E2) in the middle of the airfoil (62a, 62b), for example.

12. The wheel blade (10) according to any one of claims 6 to 11, wherein, The wings (62a, 62b): - is straight and parallel, or - It is curved, with concave sections oriented toward each other. - It has a first boss oriented toward each other and a second boss oriented toward one of the pressure side (12a) and suction side (12b) of the blade (12), respectively.

13. The blade (10) according to any one of claims 6 to 12, wherein, The first portion (54a) of the wing beam (54) also includes at least one partition wall (66, 68) of the cavity (56), the wall (66, 68) extending between and spaced apart from the wings (62a, 62b) or extending from one wing to another and connecting to the wings.

14. The blade (10) according to claim 13, wherein, The walls (66, 68) have a height (H1) along the pitch axis (A), which represents between 30% and 60% of the height (H2) of the wings (62a, 62b) along the axis (A).

15. The blade (10) according to claim 13 or 14, which is dependent on claim 8, wherein, The walls (66, 68) are connected to the partitions (64a, 64b).

16. The blade (10) according to any one of the preceding claims, wherein, The metal body (50): - Forming a part, - It is formed by at least two parts (50a, 50b), the first part (50a) including the root (14), the pillar (30) and one of the wings (62a), and the second part (50b) including the other wing (62b).

17. The blade (10) according to any one of the preceding claims, wherein, The second portion (54b) of the spar (54) includes at least one fiber unraveling portion (70) over at least a portion of its length, and the first portion (54a) of the spar (54) includes at least one protrusion (72) which engages in the second portion (54b) of the spar at the level of the unraveling portion (70).

18. The blade (10) according to any one of the preceding claims, wherein, The second part (54b) of the wing spars (54) is connected to the skin (52b1, 52b2) via a fiber mesh (74).

19. A turbine engine, particularly a turbine engine for an aircraft, the turbine engine comprising a propeller, the propeller comprising at least one blade (10) according to any one of the preceding claims.

20. A method for manufacturing a blade (10) according to any one of claims 1 to 18, wherein, The method includes the following steps: a) The fibers are woven in a three-dimensional manner to form the preform (52), and a portion of the preform (52) is unraveled to form the second portion (54b) of the spar (54). b) Compact the second part (54b) of the wing beam (54), c) The second part (54b) of the spar (54) is joined in the cavity (56) of the first part (54a) of the spar (54). d) Compact the assembly formed by the preform (52) and the body (50), e) The component is RTM-type consolidated in the mold.

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