Variable-pitch blade comprising a centring and retention zone
Patent Information
- Authority / Receiving Office
- EP · EP
- Patent Type
- Applications
- Current Assignee / Owner
- SAFRAN SA
- Filing Date
- 2024-08-27
- Publication Date
- 2026-07-08
Smart Images
Figure FR2024051122_06032025_PF_FP_ABST
Abstract
Description
[0001] DESCRIPTION
[0002] Variable pitch vane with centering and retention zone
[0003] TECHNICAL FIELD OF THE INVENTION
[0004]
[0001] The present invention relates to the general field of turbomachines, and more specifically to the field of variable-pitch propeller blades for these turbomachines.
[0002] The invention applies to any type of land-based or aeronautical turbomachines, and in particular to aircraft turbomachines such as turbojets and turboprops. More specifically, the invention finds a preferred application in the field of aircraft turbomachines comprising at least one unducted propeller, and also a pair of unducted co-rotating or contra-rotating propellers, this type of turbomachine also being called “unducted fan(s)”, or also bearing the English names “open rotor” or “propfan”.
[0005]
[0003] The general architecture of a turbomachine of the “open rotor” type is distinguished from that of a conventional turbomachine by the particular arrangement of the fan arranged outside the shroud of the turbomachine. More precisely, two types of turbomachines with unshrouded fan(s) are distinguished, namely the pusher type (“open rotor pusher” in English) and the tractor type (“open rotor puller” in English).
[0006]
[0004] In the case of a turbomachine with unducted fan(s) of the pusher type, the rotary propeller or the co-rotating or contra-rotating propellers are arranged downstream of the turbomachine, that is to say at the rear of the turbomachine following the direction of movement of the aircraft.
[0007]
[0005] In the case of a tractor-type unducted fan turbomachine, the rotary propeller or the co-rotating or counter-rotating propellers are located upstream of the turbomachine, i.e. at the front.
[0008]
[0006] More particularly, but not exclusively, the invention relates to a blade intended to be used in an unducted fan rotor of an aircraft turbomachine, such as an "open rotor" type turbomachine having two rotating propellers, a USF (for "Unducted Single Fan" in English) type engine having a moving blade and a fixed blade, or a turboprop having an architecture with a single propeller.
[0009]
[0007] The invention thus proposes a variable-pitch blade comprising a centering and retention zone, as well as a gas turbine engine comprising a plurality of such blades. STATE OF THE ART
[0010]
[0008] Climate change is a major concern for many legislative and regulatory bodies around the world. Indeed, various restrictions on carbon emissions have been, are being, or will be adopted by various states. In particular, an ambitious standard applies both to new types of aircraft and those in circulation requiring the implementation of technological solutions in order to make them compliant with current regulations. Civil aviation has been mobilizing for several years now to make a contribution to the fight against climate change.
[0011]
[0009] Technological research efforts have already made it possible to significantly improve the environmental performance of aircraft. The Applicant takes into consideration the impact factors in all phases of design and development to obtain less energy-intensive, more environmentally friendly aeronautical components and products whose integration and use in civil aviation have moderate environmental consequences with the aim of improving the energy efficiency of aircraft.
[0012]
[0010] Consequently, the Applicant is constantly working to reduce its negative climate impact by using methods and operating virtuous development and manufacturing processes and minimizing greenhouse gas emissions to the minimum possible in order to reduce the environmental footprint of its activity.
[0013]
[0011] This sustained research and development work covers new generations of aircraft engines, the lightening of aircraft, particularly through the materials used and lighter on-board equipment, the development of the use of electrical technologies to ensure propulsion, and, as essential complements to technological progress, aeronautical biofuels.
[0014]
[0012] By way of example, the Applicant's international application WO 2022 / 018357 A1 discloses a prior art embodiment of an unducted fan engine comprising variable pitch blades.
[0015]
[0013] The advantage of unducted fan engines is that the diameter of the fan is not limited by the presence of a shroud, so that it is possible to design an engine with a high bypass ratio (BPR), and consequently reduced fuel consumption. Thus, in this type of engine, the fan blades can have a large span.
[0016]
[0014] In addition, these engines generally include a mechanism for modifying the pitch angle of the blades in order to adapt the thrust generated by the fan according to the different phases of flight.
[0015] However, the design of such blades involves several disciplines and requires taking into account opposing constraints.
[0017]
[0016] On the one hand, the dimensioning of these blades must allow optimal aerodynamic performances (maximize efficiency and provide thrust while minimizing losses). The improvement of the aerodynamic performances of the fan tends towards an increase in the bypass ratio (BPR), which results in an increase in the external diameter, and therefore in the span of these blades.
[0018]
[0017] On the other hand, it is also necessary to guarantee mechanical strength of the blades, i.e. resistance to mechanical stresses resulting from static and dynamic loads, while limiting their acoustic signature.
[0019]
[0018] Furthermore, on architectures with unducted fans, engine start-up is generally carried out with a very open pitch, known as a feather pitch. Indeed, a very open pitch allows power to be consumed by the torque, which ensures machine safety by guaranteeing low fan speeds. More precisely, according to simple considerations, the power is proportional to the product of the speed and the torque. However, the torque increases with the incidence which can be increased via the pitch. Indeed, those skilled in the art of aerodynamics understand that the resulting force on a blade profile is, as a first approximation, perpendicular to the chord and can be broken down into two components: the thrust along the engine axis and the drag of the blade in the plane of the propeller.Thus, with increasing blade pitch, the resulting force shifts towards the propeller plane which has the effect of increasing the drag of the aerodynamic profile and decreasing the thrust.
[0020]
[0019] Therefore, in the case of a feathered start, the thrust generated by the propeller is zero, the torque is maximum and the speed is minimum. However, with such a very open pitch, the blades undergo a turbulent aerodynamic flow, completely detached, which generates a broadband vibration excitation. In particular on blades with a wide chord and a large span, the bending force is intense, although the engine speed is not maximum.
[0021]
[0020] In normal operation, during the ground and flight phases, the pitch is modified (the pitch angle is more closed). The aerodynamic flow is therefore perfectly healthy (adhered to the aerodynamic profile). The broadband stresses disappear, the rotation speed being higher, and the bending force is controlled.
[0021] Currently, these blades are generally made of metallic material. Although blades made of metallic material have good mechanical strength, they nevertheless have the disadvantage of having a relatively large mass.
[0022] In order to reduce this mass, it is possible to manufacture these blades from composite material. In particular, organic matrix composites (OMCs) and ceramic matrix composites (CMCs) replace metallic parts in certain parts of turbomachines. Furthermore, their use contributes to optimizing aircraft performance, in particular by improving the efficiency of the turbomachine and reducing the overall mass of the turbomachine, significantly reducing harmful emissions to the environment (CO, CO2, NO X , etc.).
[0022]
[0023] However, the intense aerodynamic forces to which these blades are subjected can damage the blade and / or the hub in the interface area between these blades and the fan rotor hub. This problem also arises due to the vibration levels on engine orders 1 N, 2 N and 3 N (possibly higher). Furthermore, in the absence of a direction of the air flowing through the blades that is parallel to the engine axis, so-called "1 P" bending-type forces or loads are generated which cause a vibration response of the blades on engine order 1 N. Similarly, these 1 P forces can also appear during the climb or approach phases of the aircraft because the air flows through the blades with an angle of incidence. These vibration excitations at high rotational speed can cause significant friction damage.
[0024] There is therefore a need to design a blade architecture that can withstand these loads and vibration levels. In particular, there is a need to ensure, within the blade, a connection and support of the composite blade by the associated spar, as well as support at all points of the skins of the composite blade, independently of the stress modes in operation, such as centrifugal stresses, and incidents of use, such as impacts from bird ingestion.
[0023] SUMMARY OF THE INVENTION
[0024]
[0025] The invention aims to at least partially remedy the needs mentioned above and the drawbacks relating to the achievements of the prior art.
[0025]
[0026] To this end, the invention is the result of technological research aimed at significantly improving aircraft performance and, in this sense, contributes to reducing the environmental impact of aircraft.
[0026]
[0027] In particular, the invention aims to propose a blade including a composite material, suitable for use with a variable pitch mechanism and in an “open rotor” type environment, while being capable of withstanding intense aerodynamic forces, under the constraint of limited bulk and minimal mass. Specifically, the invention aims to optimize the retention of the blade preform of such a composite blade to ensure cohesion with the associated spar.
[0028] The invention thus relates, according to one of its aspects, to a variable-pitch blade for a gas turbine engine, comprising a root and a blade connected to the root, the root comprising a body and a blade root attachment part, the body being housed in the blade root attachment part extending around a blade pitch axis, the body comprising: a bulb having a convex rounded shape in cross section, housed in the blade root attachment part, a stilt connecting the bulb and the blade, the blade comprising a composite material structure comprising: an aerodynamic profile structure comprising facing skins obtained by three-dimensional weaving of a fiber reinforcement by a matrix forming a three-dimensional woven preform comprising warp threads linked by weft threads, a spar, the spar comprising the body, extending outside the aerodynamic profile structure,and a blade portion disposed inside the aerodynamic profile structure between the skins, characterized in that the blade comprises a first interface defining a centering and retention zone between the aerodynamic profile structure and the blade portion, extending radially at a first part of the blade portion juxtaposed with the stilt and of axial dimension substantially equal to the axial dimension of the stilt in contact with the blade portion, in that the aerodynamic profile structure comprises, at the centering and retention zone, a transverse thickness, measured between the spar and the intrados or the extrados of the blade, variable axially and radially respectively as a function of the axial and radial evolution of the transverse thickness of the spar, and in that the three-dimensional woven preform forming the aerodynamic profile structure comprises, at the centering and retention zone,weaving arrangements including transverse outlets of weft threads.,
[0027]
[0029] Thanks to the invention, it is possible to better distribute the stresses, and also relieve the skins of the composite structure, by changing the stiffness distribution in the blade, which goes against the classic solutions stiffening the skins with the aim of having maximum rigidity.
[0028]
[0030] Advantageously, varying the transverse thickness of the aerodynamic profile structure can optimize the retention of the skins on the spar.
[0031] The blade according to the invention may further comprise one or more of the following characteristics taken in isolation or in any possible technical combination.
[0029]
[0032] The centering and retention zone may be located radially along the blade pitch axis at a height, measured along the blade pitch axis from the inner end of the blade root, of between 1% and 5% of the radial height of the blade, measured along the blade pitch axis between the inner end of the blade root and the tip of the blade.
[0030]
[0033] Advantageously, such positioning of the centering and retention area can provide improved retention of the three-dimensional woven preform on the spar.
[0031]
[0034] In addition, the axial thickness variation gradient of the transverse thickness of the aerodynamic profile structure, at the centering and retention zone, between the abscissa of the start of the spar and the master couple of the spar can be between 2.3 and 2.5 mm / mm.
[0032]
[0035] Furthermore, the blade may comprise a second interface defining a bearing area between the aerodynamic profile structure and the blade portion, extending radially at a second portion of the blade portion outside the centering and retention area, the axial dimension of the bearing area being greater than the axial dimension of the centering and retention area and increasing radially away from the blade root, the aerodynamic profile structure comprising, at the bearing area, a transverse thickness, measured between the spar and the intrados or the extrados of the blade, variable axially and radially respectively as a function of the axial and radial evolution of the transverse thickness of the spar, being less than the transverse thickness of the aerodynamic profile structure at the centering and retention area, and the three-dimensional woven preform forming the aerodynamic profile structure comprising,at the level of the span area, weaving arrangements including transverse outlets of weft threads.,
[0033]
[0036] Additionally, the bearing area may be located radially along the blade pitch axis at a height, measured along the blade pitch axis from the inboard end of the blade root, of between 5% and 10% of the radial height of the blade, measured along the blade pitch axis between the inboard end of the blade root and the blade tip.
[0034]
[0037] The blade may further comprise a third interface defining a median zone between the aerodynamic profile structure and the blade portion, extending radially outwardly to the bearing zone, the aerodynamic profile structure comprising, at the median zone, a transverse thickness, measured between the spar and the intrados or the extrados of the blade, variable axially and radially respectively as a function of the axial and radial evolution of the transverse thickness of the spar, being greater than the transverse thickness of the aerodynamic profile structure at the bearing zone, and the three-dimensional woven preform forming the aerodynamic profile structure comprising, at the median zone, weaving arrangements comprising transverse outlets of weft threads.
[0035]
[0038] Furthermore, the transverse weft thread outlets may be formed axially at a position between 0 and 20% of the blade chord, the value of 0 corresponding to the leading edge of the blade, and / or between 40% and 100% of the blade chord.
[0039] Transverse weft thread outlets and / or transverse weft thread inlets may further be formed axially at a position between 40% and 60% of the blade chord, the value of 0 corresponding to the leading edge of the blade.
[0036]
[0040] Furthermore, the three-dimensional woven preform forming the aerodynamic profile structure may comprise weaving arrangements comprising transverse outlets of warp threads, in particular located radially along the pitch axis of the blade at a height, measured along the pitch axis of the blade from the inner end of the blade root, of between 10% and 20% of the radial height of the blade, measured along the pitch axis of the blade between the inner end of the blade root and the tip of the blade, and / or transverse inlets of warp threads, in particular located radially along the pitch axis of the blade at a height, measured along the pitch axis of the blade from the inner end of the blade root, of between 20% and 40% of the radial height of the blade, measured along the pitch axis of the blade between the inner end of the blade root and the tip of the blade.
[0037]
[0041] The spar can be made of metal, composite, then including a fibrous reinforcement obtained by three-dimensional weaving and densified by a matrix, or hybrid, both metal and composite.
[0038]
[0042] In addition, the composite material structure may comprise a lightweighting foam arranged inside the aerodynamic profile structure between the skins, in particular at least partially around the blade portion of the spar.
[0039]
[0043] Advantageously, the presence of the lightening foam can help to ensure a satisfactory ratio between strength and mass of the blade.
[0040]
[0044] In addition, the radial height of the blade, measured along the blade pitch axis between the inner end of the blade root and the tip of the blade, may be between 1500 mm and 2100 mm, preferably between 1650 mm and 1950 mm.
[0045] Furthermore, the chord of the blade at the inner end of the blade, corresponding to the distance between the leading edge and the trailing edge, may be between 300 mm and 600 mm, preferably between 350 mm and 500 mm.
[0041]
[0046] Furthermore, the invention also relates, according to another of its aspects, to a gas turbine engine comprising a fan, the fan comprising a hub and blades extending radially from the hub, the blades being such as that defined previously, each blade being rotatably mounted relative to the hub around a respective pitch axis, the fan further comprising an actuating mechanism capable of being controlled to rotate the blades around their pitch axes so as to modify the pitch angle of the blades.
[0042]
[0047] Preferably, the diameter of the blower may be less than or equal to 6 m, in particular between 3 m and 5 m, more preferably between 3.5 m and 4.5 m.
[0048] Additionally, the number of blades per propeller can be between 12 and 18.
[0043]
[0049] Advantageously, the flow velocity relative to the blower may be greater than a Mach number of 0.7 and less than a Mach number of 0.9.
[0044]
[0050] Furthermore, the fan propeller hub ratio, defined as the ratio of the inner radius to the outer radius of the blade, the inner radius being measured from the point on the leading edge of the blade closest to the axis of rotation and the outer radius being measured from the point on the leading edge furthest from the axis of rotation, may be between 0.25 and 0.35.
[0045]
[0051] In particular, for a hub ratio greater than or equal to 0.25 and less than 0.30, the number of blades per propeller can be between 12 and 16.
[0046]
[0052] For a hub ratio greater than or equal to 0.30 and less than or equal to 0.35, the number of blades per propeller can be between 1 and 18.
[0047]
[0053] Furthermore, the invention also relates, according to another of its aspects, to an aircraft comprising a gas turbine engine as defined previously.
[0048] BRIEF DESCRIPTION OF THE FIGURES
[0049]
[0054] Other advantages, aims and particular characteristics of the invention will emerge from the following non-limiting description of at least one embodiment of the present invention, with reference to the appended figures, in which: Figure 1 schematically represents an example of an engine according to the invention including an unducted fan; Figure 2 schematically represents a fan blade according to the invention and an actuating mechanism making it possible to modify the pitch angle of the fan blades; Figure 3 schematically represents in perspective an unducted fan blade according to the invention; Figure 4 is a larger-scale view of a part of Figure 3 and shows the root of the blade; Figure 5 schematically represents in perspective the body of the root of the blade of Figure 3;Figure 6 schematically represents another example of an unducted fan blade according to the invention; Figure 6A is a sectional view along ÀÀ of Figure 6; Figure 6B is a sectional view along BB of Figure 6; Figure 6C is a sectional view along CC of Figure 6; Figure 6D is an enlarged detail view of zone D of Figure 6; Figure 6E is a sectional view along EE of Figure 6; Figure 7A is a sectional view showing the configuration of the preform at section AÀ of Figure 6; Figure 7B is a sectional view showing the configuration of the preform at section BB of Figure 6; Figure 7C is a sectional view showing the configuration of the preform at section CC of Figure 6; and Figure 7D is a sectional view showing the configuration of the preform at section EE of Figure 6.;
[0050]
[0055] Throughout these figures, like references may designate identical or similar elements.
[0051]
[0056] Furthermore, the different parts represented in the figures are not necessarily shown on a uniform scale, in order to make the figures more readable.
[0052] DETAILED DESCRIPTION OF THE INVENTION
[0053]
[0057] Throughout the description, given as a non-limiting example of embodiment, it is noted that the terms upstream and downstream are to be considered with respect to a main direction F of normal gas flow (from upstream to downstream) for a turbomachine 1. Furthermore, the axis X of the turbomachine 1 is called the axis of radial symmetry of the turbomachine 1. The axial direction of the turbomachine 1 corresponds to the axis of rotation X of the turbomachine 1. A radial direction of the turbomachine 1 is a direction perpendicular to the axis X of the turbomachine 1.
[0054]
[0058] Furthermore, unless otherwise specified, the adjectives and adverbs axial, radial, axially and radially are used with reference to the aforementioned axial and radial directions. Furthermore, unless otherwise specified, the terms inner and outer are used with reference to a radial direction such that the inner part of an element is closer to the X axis of the turbomachine 1 than the outer part of the same element.
[0055]
[0059] Furthermore, a cross-section is understood to mean a section perpendicular to the Y-axis of the blade setting. A bulb is also understood to mean a swollen or curved part, i.e. one comprising a bulge or a bulge which extends around the Y-axis of setting in this case.
[0056]
[0060] In Figure 1, the engine 1 shown is an “open rotor” type engine, in a configuration commonly referred to as a “pusher” (i.e. the unducted fan is placed at the rear of the power generator with an air inlet located on the side, on the right in Figure 1) with counter-rotating fan rotors. However, the invention is not limited to this configuration. It also applies to “open rotor” type engines in a “puller” configuration (i.e. the fan is placed upstream of the power generator with an air inlet located before, between or just behind the two fan rotors).
[0057]
[0061] Furthermore, the invention also applies to engines having different architectures, such as an architecture comprising a fan rotor comprising moving blades and a fan stator comprising fixed blades, or a single fan rotor. The invention is applicable to turboprop-type architectures (comprising a single fan rotor).
[0058]
[0062] The engine 1 comprises a nacelle 2 intended to be fixed to a fuselage of an aircraft, and an unducted fan 3. The fan 3 comprises two counter-rotating fan rotors 4 and 5. In other words, when the engine 1 is in operation, the rotors 4 and 5 are rotated relative to the nacelle 2 around the same axis of rotation X (which coincides with a main axis of the engine), in opposite directions.
[0059]
[0063] Each fan rotor 4, 5 comprises a hub 6 rotatably mounted relative to the nacelle 2 and a plurality of blades 7 fixed to the hub 6. The blades 7 extend substantially radially relative to the axis of rotation X of the hub 6.
[0060]
[0064] As illustrated in Figure 2, the fan 3 further comprises an actuating mechanism 8 for collectively modifying the pitch angle of the blades 7 of the rotors 4, 5, in order to adapt the performance of the engine 1 to the different flight phases. For this purpose, each blade 7 comprises an attachment part 9 (or hub) arranged at the blade root. The attachment part 9 is rotatably mounted relative to the hub 6 around a pitch axis Y. More precisely, the attachment part 9 is rotatably mounted inside a housing 10 formed in the hub 6, by means of balls 11 or other rolling elements.
[0065] The actuating mechanism 8 comprises, for example, here an actuator 19 comprising a body 13 fixed to the hub 6 and a rod 14 capable of being driven in translation relative to the body 13. The actuating mechanism 8 further comprises an annular slide 15 mounted integral with the rod 14 and a pin 16 mounted integral with the attachment part 9. The pin 16 is capable of sliding in the slide 15 and of rotating relative to the slide 15, so as to convert a translational movement of the rod 14 into a rotational movement of the attachment part 9, and consequently a rotational movement of the blade 7 relative to the hub 6 around its setting axis Y.
[0061]
[0066] With reference to Figures 3 to 5, an example of a blade 7 for a propeller of an aircraft turbomachine is shown, for example a blade 7 as described previously with reference to Figures 1 and 2.
[0062]
[0067] The blade 7 comprises a blade 12 connected to a blade root 17. The blade 12 has an aerodynamic profile and comprises a lower surface 12a and an upper surface 12b which are connected by an upstream leading edge 12c and by a downstream trailing edge 12d. The blade 12 has an upper end which is free, called the tip, and a lower end which is connected to the root 17.
[0063]
[0068] In the example shown, the blade 7 is made of composite material, for example by an injection process called RTM (for “Resin Transfer Molding” in English). This process consists of preparing a fiber preform 18 by three-dimensional weaving then placing this preform in a mold and injecting a polymerizable resin such as an epoxy resin, which will impregnate the fiber preform 18. After polymerization and hardening of the blade 12, its leading edge 12c is generally reinforced by a metal shield 20 added and fixed, for example by gluing.
[0064]
[0069] The blade 7 here comprises a spar 22 which comprises a part forming a core of the blade 12 and which is intended to be inserted into the preform 18 before the injection of resin, and a part which extends on the side opposite the top of the blade 12 to form a part of the root 17, called body 24.
[0065]
[0070] The spar 22 is for example metallic here. However, the spar 22 can also be made of composite material, in particular of epoxy organic matrix composite (OMC) material reinforced by 3D woven carbon fibers with the warp direction predominantly oriented radially and the weft predominantly oriented along the chord of the blade 12 at the aerodynamic vein height. The spar 22 can also be a more mechanically advantageous assembly of different organic matrix composite materials (thermosetting, thermoplastic or elastomer) reinforced by long fibers (carbon, glass, aramid, polypropylene) in several fiber arrangements (woven, braided, knitted, unidirectional).
[0071] Although not shown, the blade 12 may be hollow or solid and has an internal cavity filled with a foam or honeycomb type filler material, for example such as the lightweight foam 42 visible in FIG. 6 described later. This filler material is installed around the spar 22 and is covered with a skin made of organic matrix composite material to increase the impact resistance of the blade 12.
[0066]
[0072] The shield 20 may be made of titanium or titanium alloy, stainless steel, steel, aluminum, nickel, etc. The intrados 12a, or even the extrados 12b, of the blade 12 may be covered with a polyurethane film for protection against erosion.
[0067]
[0073] The foot 17 essentially comprises two parts, namely this body 24, which is part of the spar 22, and the attachment piece 9, or annular shaft, which extends around the body 24 and the Y axis of the blade 7. The Y axis is the setting axis of the blade 7, that is to say the axis around which the angular position of the blade 7 is adjusted. It is generally also a radial axis which therefore extends along a radius relative to the axis of rotation X of the propeller equipped with this blade 7.
[0068]
[0074] The body 24 of the foot 17 has a particular shape better visible in figure 5. Here it essentially comprises three parts, namely: a free end 28 located on the side opposite the blade 12; a stilt 30 located on the side of the blade 12; and a bulb 32 located between the free end 28 and the stilt 30.
[0069]
[0075] The free end 28 has a generally parallelepiped shape in the example shown.
[0070]
[0076] The stilt 30 has a relatively complex shape and can be considered as comprising: two lateral flanks 30a, 30b, located respectively on the side of the intrados 12a and the extrados 12b of the blade 12, which converge towards each other along the Y axis and in the direction of the top of the blade 12; and two edges, respectively upstream 30c and downstream 30d, which on the contrary diverge from each other along the Y axis and in the direction of the top of the blade 12.
[0071]
[0077] The bulb 32 has a generally swollen or domed shape, this bulge or doming extending all around the Y axis.
[0072]
[0078] With reference to Figure 6 and the following Figures 6A to 7D, an example of a variable-pitch blade 7 according to the invention for a gas turbine engine 1 is shown in more detail. The blade 7 comprises a root 17 and a blade 12 connected to the root 17. The root 17 comprises a body 24 and a blade root attachment part 9. The body 24 is housed in the blade root attachment part 9 and extends around the pitch axis Y of the blade 7.
[0079] As described previously, the body 24 comprises a bulb 32 having a convex rounded shape in cross section, housed in the blade root attachment part 9, and a stilt 30 connecting the bulb 32 and the blade 12.
[0073]
[0080] In addition, the blade 7 comprises a composite material structure 40 which comprises an aerodynamic profile structure 41 comprising facing skins obtained by three-dimensional weaving of a fiber reinforcement by a matrix forming a three-dimensional woven preform 45, visible in FIGS. 7A, 7B, 7C and 7D, comprising warp threads linked by weft threads; a spar 22 comprising the body 24, extending outside the aerodynamic profile structure 41, and a blade portion 25 arranged inside the aerodynamic profile structure 41 between the skins; and in this example but in no way limiting, a lightening foam 42 arranged inside the aerodynamic profile structure 41 between the skins, partially around the blade portion 25 of the spar 22.
[0074]
[0081] The spar 22 is preferably made of metal, but can also be made of composite, then comprising a fibrous reinforcement obtained by three-dimensional weaving and densified by a matrix, or be hybrid, both made of metal and composite.
[0075]
[0082] With reference to figures 6, 6A, 6D and 7A, the blade 7 comprises a first interface defining a centering and retention zone Z cr between the aerodynamic profile structure 41 and the blade portion 25, extending radially at the level of a first part P1 of the blade portion 25, visible in FIG. 6D, juxtaposed with the stilt 30 and of axial dimension di substantially equal to the axial dimension of the stilt 30 in contact with the blade portion 25.
[0076]
[0083] The centering and retention zone Z cr is advantageously located radially along the Y axis of the blade 7 at a height h cr, visible in Figure 6D, measured along the Y axis of the blade 7 from the inner end of the blade root 17, which is between 1% and 5% of the radial height H of the blade 7, measured along the Y axis of the blade 7 between the inner end of the blade root 17 and the tip of the blade 12.
[0077]
[0084] In addition, at the level of this centering and retention zone Z cr , the aerodynamic profile structure 41 has a transverse thickness ei, measured between the intrados 12a and the extrados 12b of the blade 12, variable axially and radially respectively as a function of the axial and radial evolution of the transverse thickness ei of the spar 22.
[0078]
[0085] As seen in Figure 6A, the transverse thickness ei is modified, on each side of the spar 22, with the evolution of the thickness ei of the spar 22. For example, on the leading edge side (on the left in Figure 6A), the reasoning being similar on the trailing edge side, the total transverse thickness goes from a value of 2 x ei to a value e c which is greater than 2 x ei. Preferably, the axial thickness variation gradient of the transverse thickness ei of the aerodynamic profile structure 41, at the centering and retention zone Z cr , between the abscissa of the start of the spar 22 and the master torque of the spar 22 is between 2.3 and 2.5 mm / mm.
[0079]
[0086] Furthermore, at the level of this centering and retention zone Z cr, and as visible in Figure 7A, the three-dimensional woven preform 45 forming the aerodynamic profile structure 41 comprises weaving arrangements comprising transverse outlets of weft threads St, and possibly inlets of weft threads E t .
[0080]
[0087] Thus, in the example of Figure 7A, transverse outlets of weft threads S t are formed axially at a position between 0 and 20% of the chord of the blade 7 and between 40% and 100% of the chord of the blade 7.
[0081]
[0088] In addition, transverse outlets of weft threads S t and / or transverse entries of weft threads E t are formed axially at a position between 40% and 60% of the chord of the blade 7.
[0082]
[0089] The centering and retention zone Z cr, corresponds to a zone called the bottom of the blade. In this zone, a shape complementarity is put in place both in the direction of the chord and out of plane so as to allow the skins to be retained even in the presence of a degraded interface between the spar 22 and the skins. A delinking Zd is made in the preform 45 over a given width close to that of the spar 22 in terms of perimeter, as visible in FIG. 7A, this delinking Zd being made between two layers of weft threads by not passing warp threads through the delinking zone so as not to link threads of weft layers located on either side of the delinking.
[0083]
[0090] Further, with reference to Figures 6, 6B, 6C and 7B, the blade 7 also comprises a second interface defining a bearing area Z pbetween the aerodynamic profile structure 41 and the blade portion 25, extending radially at a second portion P2 of the blade portion 25 outside the centering and retention zone Z cr , the axial dimension d2 of the bearing area Z p being greater than the axial dimension di of the centering and retention zone Z cr and increasing radially away from the foot of the blade 17.
[0084]
[0091] As seen in Figure 6B, at the Z range area p , the aerodynamic profile structure 41 has a transverse thickness e2, measured between the spar 22 and the intrados 12a or the extrados 12b of the blade 12, variable axially and radially respectively as a function of the axial and radial evolution of the transverse thickness ei of the spar 22, being less than the transverse thickness ei of the aerodynamic profile structure 41 at the centering and retention zone Z cr .
[0092] In addition, the three-dimensional woven preform 45 forming the aerodynamic profile structure 41 comprises, at the level of the bearing area Z p , weaving arrangements comprising transverse outlets of weft threads S t and possibly weft thread entries E t .
[0085]
[0093] Thus, in the example of Figure 7B, transverse outlets of weft threads S t are formed axially at a position between 0 and 20% of the chord of the blade 7 and between 40% and 100% of the chord of the blade 7.
[0086]
[0094] In addition, transverse outlets of weft threads S t and / or transverse entries of weft threads E t are formed axially at a position between 40% and 60% of the chord of the blade 7.
[0087]
[0095] The Z range zone pcan allow the centering of the spar 22 along the chord to be adjusted as well as possible, as well as adding to the retention by reinforcing this area. Furthermore, this can allow the aerodynamic profile structure 41 not to stop on the spar 22 at an angle, which would be a source of stress concentrations.
[0088]
[0096] Advantageously, the range zone Z p is located radially along the Y axis of the blade 7 at a height h p , measured along the pitch axis Y of the blade 7 from the inner end of the blade root 17, which is between 5% and 10% of the radial height H of the blade 7, measured along the pitch axis Y of the blade 7 between the inner end of the blade root 17 and the tip of the blade 12.
[0089]
[0097] Furthermore, with reference to figures 6, 6C, 6D and 7C, the blade 7 here comprises a third interface defining a median zone Z mbetween the aerodynamic profile structure 41 and the blade portion 25, extending radially outwardly to the bearing area Z p .
[0090]
[0098] At the level of the middle zone Z m , with reference to Figure 6C, the aerodynamic profile structure 41 has a transverse thickness e3, measured between the spar 22 and the intrados 12a or the extrados 12b of the blade 12, which is variable axially and radially respectively as a function of the axial and radial evolution of the transverse thickness ei of the spar 22, being greater than the transverse thickness e2 of the aerodynamic profile structure 41 at the level of the bearing area Z p .
[0091]
[0099] In addition, as seen in Figure 7C, the three-dimensional woven preform 45 forming the aerodynamic profile structure 41 comprises, at the level of the median zone Z m, weaving arrangements comprising transverse outlets of weft threads St, and possibly inlets of weft threads E t .
[0092]
[0100] Thus, in the example of Figure 7C, transverse outlets of weft threads S t are formed axially at a position between 0 and 20% of the chord of the blade 7 and between 40% and 100% of the chord of the blade 7.
[0101] In addition, transverse outlets of weft threads S t and / or transverse entries of weft threads E t are formed axially at a position between 40% and 60% of the chord of the blade 7.
[0093]
[0102] Furthermore, as visible in Figure 6E, it can be seen that the transverse thickness ei, e2 or e3 of the aerodynamic profile structure 41 advantageously varies along the spar 22 in the radial direction.
[0094]
[0103] Furthermore, with reference to FIG. 7D, it can be seen that the three-dimensional woven preform 45 forming the aerodynamic profile structure 41 comprises weaving arrangements comprising transverse outlets of warp threads S c , located radially along the Y axis of the blade 7 at a height hs, measured along the Y axis of the blade 7 from the inner end of the blade root 17, which is between 10% and 20% of the radial height H of the blade 7.
[0095]
[0104] These weaving arrangements also include transverse entries of warp threads E c , located radially along the Y axis of the blade 7 at a height he, measured along the Y axis of the blade 7 from the inner end of the blade root 17, which is between 20% and 40% of the radial height H of the blade 7.
[0096]
[0105] Advantageously, the radial height H of the blade 7 is between 1500 mm and 2100 mm, preferably between 1650 mm and 1950 mm. In addition, the chord Ci of the blade 7, visible in FIG. 6, at the inner end of the blade 12, corresponding to the distance between the leading edge 12c and the trailing edge 12d, is between 300 mm and 600 mm, preferably between 350 mm and 500 mm.
[0097]
[0106] Furthermore, the diameter DH of the blower 3, visible in figure 1, is less than or equal to 6 m, in particular between 3 m and 5 m, more preferably between 3.5 m and 4.5 m.
[0098]
[0107] Advantageously, the number of blades 7 per propeller is between 12 and 18, and the hub ratio of the fan propeller 3, defined as the ratio between the inner radius and the outer radius of the blade 7, the inner radius being measured from the point on the leading edge 12c of the blade 12 closest to the axis of rotation X and the outer radius being measured from the point on the leading edge 12c furthest from the axis of rotation X, is between 0.25 and 0.35.
[0099]
[0108] In particular, a number of blades 7 per propeller between 12 and 16 will be chosen for a hub ratio greater than or equal to 0.25 and less than 0.30, and a number of blades 7 per propeller between 14 and 18 will be chosen for a hub ratio greater than or equal to 0.30 and less than or equal to 0.35.
[0109] Of course, the invention is not limited to the embodiments which have just been described. Various modifications can be made to it by those skilled in the art.
Claims
CLAIMS 1. Variable pitch vane (7) for a gas turbine engine (1), having a pitch axis (Y) and comprising a root (17) and a blade (12) connected to the root (17), the root (17) comprising a body (24) and a blade root attachment part (9), the body (24) being housed in the blade root attachment part (9) extending around the pitch axis (Y) of the vane (7), the body (24) comprising: a bulb (32) having in cross section, perpendicular to the pitch axis (Y) of the vane (7), a convex rounded shape, housed in the blade root attachment part (9), a stilt (30) connecting the bulb (32) and the blade (12), the vane (7) comprising a composite material structure (40) comprising: an aerodynamic profile structure (41) comprising facing skins obtained by three-dimensional weaving of a fibrous reinforcement by a matrix forming a three-dimensional woven preform (45) comprising warp threads linked by weft threads, a spar (22),the spar (22) comprising the body (24), extending outside the aerodynamic profile structure (41), and a blade portion (25) arranged inside the aerodynamic profile structure (41) between the skins, characterized in that the blade (7) comprises a first interface defining a centering and retention zone (Z, cr ) between the aerodynamic profile structure (41) and the blade portion (25), extending radially, parallel to the pitch axis (Y) of the blade (7), at a first part (P1) of the blade portion (25) juxtaposed with Péchasse (30) and of axial dimension (di), perpendicular to the pitch axis (Y) of the blade (7), substantially equal to the axial dimension, perpendicular to the pitch axis (Y) of the blade (7), of Péchasse (30) in contact with the blade portion (25), in that the aerodynamic profile structure (41) comprises, at the centering and retention zone (Z cr), a transverse thickness (ei), perpendicular to the pitch axis (Y) of the blade (7) and measured between the spar (22) and the intrados (12a) or the extrados (12b) of the blade (12), variable axially, perpendicular to the pitch axis (Y) of the blade (7), and radially, parallel to the pitch axis (Y) of the blade (7), respectively as a function of the axial evolution, perpendicular to the pitch axis (Y) of the blade (7), and radial, parallel to the pitch axis (Y) of the blade (7), of the transverse thickness (ei) of the spar (22), and in that the three-dimensional woven preform (45) forming the aerodynamic profile structure (41) comprises, at the centering and retention zone (Z cr ), weaving arrangements including transverse outlets of weft threads (S t ).
2. Variable-pitch blade (7) according to claim 1, in which the centering and retention zone (Z cr) is located radially along the blade (7) setting axis (Y) at a height (h cr ), measured along the pitch axis (Y) of the blade (7) from the inner end of the blade root (17), between 1% and 5% of the radial height (H) of the blade (7), measured along the pitch axis (Y) of the blade (7) between the inner end of the blade root (17) and the tip of the blade (12).
3. Variable-pitch blade (7) according to claim 1 or 2, in which the axial thickness variation gradient of the transverse thickness (ei) of the aerodynamic profile structure (41), at the centering and retention zone (Z cr ), between the abscissa of the start of the spar (22) and the master couple of the spar (22) is between 2.3 and 2.5 mm / mm.
4. Variable-pitch blade (7) according to one of the preceding claims, in which the blade (7) comprises a second interface defining a bearing area (Z p) between the aerodynamic profile structure (41) and the blade portion (25), extending radially at a second portion (P2) of the blade portion (25) outside the centering and retention zone (Z cr ), the axial dimension (dz) of the bearing area (Z p ) being greater than the axial dimension (di) of the centering and retention zone (Z cr ) and increasing radially away from the blade root (17), the aerodynamic profile structure (41) comprising, at the level of the bearing area (Z p ), a transverse thickness (ez), measured between the spar (22) and the intrados (12a) or the extrados (12b) of the blade (12), variable axially and radially respectively as a function of the axial and radial evolution of the transverse thickness (ei) of the spar (22), being less than the transverse thickness (ei) of the aerodynamic profile structure (41) at the centering and retention zone (Z cr), and the three-dimensional woven preform (45) forming the aerodynamic profile structure (41) comprising, at the level of the bearing area (Z p ), weaving arrangements including transverse outlets of weft threads (S t ).
5. Variable pitch vane (7) according to claim 4, in which the bearing area (Z p ) is located radially along the blade (7) setting axis (Y) at a height (h p ), measured along the blade setting axis (Y) (7) from the inner end of the root of blade (17), between 5% and 10% of the radial height (H) of the blade (7), measured along the axis (Y) of the blade (7) between the inner end of the blade root (17) and the tip of the blade (12).
6. Variable-pitch blade (7) according to claim 4 or 5, wherein the blade (7) comprises a third interface defining a middle zone (Z m) between the aerodynamic profile structure (41) and the blade portion (25), extending radially outwardly to the bearing area (Z p ), the aerodynamic profile structure (41) comprising, at the level of the middle zone (Z m ), a transverse thickness (es), measured between the spar (22) and the intrados (12a) or the extrados (12b) of the blade (12), variable axially and radially respectively as a function of the axial and radial evolution of the transverse thickness (ei) of the spar (22), being greater than the transverse thickness (ez) of the aerodynamic profile structure (41) at the level of the bearing area (Z p ), and the three-dimensional woven preform (45) forming the aerodynamic profile structure (41) comprising, at the level of the middle zone (Z m ), weaving arrangements including transverse outlets of weft threads (S t ).
7. A blade according to any one of the preceding claims, wherein the transverse outlets of weft threads (S t ) are formed axially at a position between 0 and 20% of the chord of the blade (7), the value of 0 corresponding to the leading edge (12c) of the blade (12), and / or between 40% and 100% of the chord of the blade (7).
8. Blade according to any one of claims 1 to 5, in which transverse outlets of weft threads (S t ) and / or transverse entries of weft threads (E t ) are formed axially at a position between 40% and 60% of the blade chord (7), the value of 0 corresponding to the leading edge (12c) of the blade (12).
9. Blade according to any one of the preceding claims, in which the three-dimensional woven preform (45) forming the aerodynamic profile structure (41) comprises weaving arrangements comprising transverse outlets of warp threads (Sc ), in particular located radially along the axis (Y) of the blade (7) at a height (hs), measured along the axis (Y) of the blade (7) from the inner end of the blade root (17), between 10% and 20% of the radial height (H) of the blade (7), measured along the axis (Y) of the blade (7) between the inner end of the blade root (17) and the tip of the blade (12), and / or transverse entries of warp threads (E c ), in particular located radially along the axis (Y) of the blade setting (7) at a height (he), measured along the pitch axis (Y) of the blade (7) from the inner end of the blade root (17), between 20% and 40% of the radial height (H) of the blade (7), measured along the pitch axis (Y) of the blade (7) between the inner end of the blade root (17) and the tip of the blade (12).
10. Variable pitch blade (7) according to any one of the preceding claims, in which the spar (22) is made of metal, composite, then comprising a fibrous reinforcement obtained by three-dimensional weaving and densified by a matrix, or hybrid, both metal and composite.
11. Variable-pitch blade (7) according to any one of the preceding claims, in which the composite material structure (40) further comprises a lightening foam (42) arranged inside the aerodynamic profile structure (41) between the skins, in particular at least partially around the blade portion (25) of the spar (22).
12. Variable pitch blade (7) according to any one of the preceding claims, in which the radial height (H) of the blade (7), measured along the pitch axis (Y) of the blade (7) between the inner end of the blade root (17) and the tip of the blade (12), is between 1500 mm and 2100 mm, in particular between 1650 mm and 1950 mm.
13. Variable-pitch blade (7) according to any one of the preceding claims, in which the chord (Ci) of the blade (7) at the inner end of the blade (12), corresponding to the distance between the leading edge (12c) and the trailing edge (12d), is between 300 mm and 600 mm, in particular between 350 mm and 500 mm.
14. A gas turbine engine (1) comprising a fan (3), the fan (3) comprising a hub (6) and blades (7) extending radially from the hub (6), the blades (7) being according to any one of the preceding claims, each blade (7) being rotatably mounted relative to the hub (6) about a respective pitch axis (Y), the fan (3) further comprising an actuating mechanism (8) capable of being controlled to rotate the blades (7) about their pitch axes (Y) so as to modify the pitch angle of the blades (7).
15. Gas turbine engine (1) according to claim 14, wherein the diameter (DH) of the fan (3) is less than or equal to 6 m, in particular between 3 m and 5 m, in particular still between 3.5 m and 4.5 m.
16. Gas turbine engine (1) according to claim 14 or 15, wherein the number of blades (7) per propeller is between 12 and 18.
17. A gas turbine engine (1) according to one of claims 14 to 16, wherein the hub ratio of the fan propeller (3), defined as the ratio of the inner radius to the outer radius of the blade (7), the inner radius being measured from the point on the leading edge (12c) of the blade (12) closest to the axis of rotation (X) and the outer radius being measured from the point on the leading edge (12c) furthest from the axis of rotation (X), is between 0.25 and 0.
35.
18. Gas turbine engine (1) according to claim 17, wherein for a hub ratio greater than or equal to 0.25 and less than 0.30, the number of blades (7) per propeller is between 12 and 16.
19. Gas turbine engine (1) according to claim 17, wherein for a hub ratio greater than or equal to 0.30 and less than or equal to 0.35, the number of blades (7) per propeller is between 14 and 18.