Fixed bucket assembly including variable pitch bucket for turbine engine

Abstract

The invention relates to a vane (7) for a stationary vane assembly (6), comprising:-a vane root (11) comprising a first vane root portion (11a) intended to be connected to a hub (8) of a turbine engine (1), and a second vane root portion (11b) intended to be connected to the hub (8) of the turbine engine (1); the first and second vane root portions (11a, 11b) are connected by a vane root engagement portion (llc) having a first outer surface (17a) and an opposite second outer surface (17b); -a first reinforcement (19a) and a second reinforcement (19b), the first reinforcement (19a) having a first inner surface (23a) configured to engage with the first outer surface (17a) and the second reinforcement (19b) having a second inner surface (23b) configured to engage with the second outer surface (17b); and-a first fastening member (24a) configured to mechanically fasten the first reinforcement body (19a) to the first bucket root portion (11a), and-a second fastening member (24b) configured to mechanically fasten the second reinforcement body (19b) to the first bucket root portion (11a).

Description

Technical Field

[0001] This invention relates to the field of gas turbine engines. In particular, it relates to a stationary blade assembly for a gas turbine engine, such as a stationary blade assembly comprising stator blades with variable pitch, each stator blade being pivotally mounted. This invention is particularly applicable to stationary blade assemblies for ductless or ducted turbine engines. Background Technology

[0002] A gas turbine engine that includes at least one ductless propeller is called an "open rotor" turbine engine or an "unducted fan" turbine engine. Such a gas turbine engine may include two ductless counter-rotating propellers (Contra-Rotating Open Rotor, CROR) or a single ductless propeller (Unducted Single Fan, USF) and a stationary blade assembly comprising multiple stator blades. The propeller can be positioned at the rear of the gas generator (or engine) as a thruster type, or at the front of the gas generator as a puller type. These gas turbine engines are turboprop engines, which differ from turbojet engines in that they use an external (ductless) propeller outside the nacelle instead of an internal fan. This allows for a significantly higher bypass ratio without being adversely affected by the weight of the casing or nacelle designed to surround the propeller or fan blades.

[0003] The stator blades of the stationary blade assembly are typically mounted on a hub, which carries nozzles for separating the main and secondary flows, which flow in the main flow path and around the inlet casing, respectively. Unlike the upstream propeller of the USF gas turbine engine, the stator blades of the stationary blade assembly are fixed and do not rotate relative to the axis of rotation of the upstream propeller, and therefore are not subjected to any centrifugal force.

[0004] The stator blades extend from the inlet housing and can have variable pitch if desired. In this case, each stator blade root is pivotally mounted along the pitch axis and connected to a pitch-changing system installed in the gas turbine engine. However, the area where the stator blade roots and the pivot are joined is severely constrained by the numerous devices surrounding them.

[0005] However, existing technologies for fastening propeller blades cannot meet the requirements for fastening stator blades in gas turbine engines, because propeller blades use centrifugal force generated by the rotation of the propeller to press the root of the blade into its fastener. However, in stationary blade assemblies, the absence of rotation and therefore the absence of centrifugal force prevents the use of such fastening devices.

[0006] Therefore, the blades of stationary blade assemblies are typically fastened to metal fasteners, including bolts extending axially or radially, to secure the blade root to the fastener. However, this type of fastener has a large tangential footprint, especially when the bolts are installed radially, making it difficult to integrate this type of fastener into gas turbine engines.

[0007] In addition, the fasteners are subjected to high bending stress, and the area where the bolts are inserted at the root of the blade forms a stress concentration area. Due to the pressure difference between the belly and back of the blade, this stress concentration area is subjected to asymmetrical stress. Summary of the Invention

[0008] One objective of this application is to overcome the aforementioned disadvantages by providing a system for securing the blades of a stationary blade assembly of a gas turbine engine, which has a reduced footprint while ensuring that the blades of the stationary blade assembly are properly held.

[0009] Therefore, according to the first aspect, a blade for a stationary blade assembly of a gas turbine engine is proposed, the blade comprising: - Blade root, the blade root includes a first blade root portion and a second blade root portion, the first blade root portion is intended to be connected to the hub of the gas turbine engine, the second blade root portion extends radially from the first blade root portion relative to the central axis of the gas turbine engine when the first blade root portion is connected to the hub of the gas turbine engine, the first blade root portion and the second blade root portion are connected by a blade root engagement portion, the blade root engagement portion has a first outer surface and a second outer surface opposite to the first outer surface; - A first reinforcing body and a second reinforcing body, the first reinforcing body and the second reinforcing body being arranged on both sides of the blade root joint portion, the first reinforcing body having a first inner surface configured to closely fit the first outer surface of the blade root joint portion, the second reinforcing body having a second inner surface configured to closely fit the second outer surface of the blade root joint portion; and - A first fastening member and a second fastening member, the first fastening member being configured to mechanically fasten the first reinforcement to the root portion of the first blade, and the second fastening member being configured to mechanically fasten the second reinforcement to the root portion of the first blade.

[0010] Advantageously, the first outer surface is concave, the second outer surface is concave, the first inner surface of the first reinforcement is convex, and the second inner surface of the second reinforcement is convex.

[0011] Advantageously, the impeller includes a platform comprising a radially outer surface and a radially inner surface, the radially outer surface being configured to define an airflow path through the impeller assembly, the platform being fixedly mounted on a first reinforcement and / or a second reinforcement.

[0012] The impeller preferably includes a first support arm and / or a second support arm, the first support arm including a first end fastened to a first reinforcement and a opposite second end fastened to a platform, and the second support arm including a third end fastened to a second reinforcement and a opposite fourth end fastened to a platform.

[0013] Advantageously, the first reinforcement includes a first reinforcement portion that contacts the root portion of the first blade, a second reinforcement portion that contacts the root portion of the second blade, and a first reinforcement engagement portion that connects the second reinforcement portion to the first reinforcement portion, the convex first inner surface being the surface of the first reinforcement engagement portion, and the first end of the first support arm being fastened to the first reinforcement portion.

[0014] Advantageously, the second reinforcement includes a third reinforcement portion that contacts the root portion of the first blade, a fourth reinforcement portion that contacts the root portion of the second blade, and a second reinforcement joining portion that connects the fourth reinforcement portion to the third reinforcement portion, the convex second inner surface being the surface of the second reinforcement joining portion, and the third end of the second support arm being fastened to the third reinforcement portion.

[0015] Advantageously, each of the reinforcements includes a plurality of first support arms distributed along the dorsal surface of the blade and a plurality of second support arms distributed along the ventral surface of the blade.

[0016] In one embodiment, each of the reinforcements is made of metal.

[0017] Advantageously, the platform includes a first platform portion fastened to a first reinforcement and a second platform portion fastened to a second reinforcement.

[0018] Advantageously, at least one second end of the first support arm and / or at least one fourth end of the second support arm are fastened to the platform by a third fastening member.

[0019] The blade preferably includes a first skin, a second skin, and a forming element. The first and second skins are made of a composite material composed of fiber preforms and are connected to each other at the leading and trailing edges of the blade. The forming element extends between the first and second skins within the root joint portion of the blade.

[0020] The present invention also relates to a gas turbine engine blade assembly, the gas turbine engine blade assembly including at least one blade as defined above and a hub, the blade being fixedly mounted on the hub by a first fastening member and a second fastening member.

[0021] Advantageously, the blade assembly includes a fastener adapted to be pivotally mounted relative to the hub about the pitch axis, wherein the blade is fixedly mounted on the fastener by a first fastening member and a second fastening member. Attached Figure Description

[0022] Other features, objects, and advantages of the invention will become apparent from the following description, which is given only in an illustrative and non-limiting manner and should be read with reference to the accompanying drawings, in which: Figure 1 A schematic diagram of a partial axial cross-section of an example gas turbine engine is shown, which may include the blades of a stationary blade assembly according to one embodiment; Figure 2 This is a schematic cross-sectional view of an example of a stationary impeller assembly according to one embodiment; Figure 3 This is a top view of an embodiment of a platform for wheel blades according to an embodiment of the present invention; and Figure 4 It shows Figure 3 A partial schematic diagram of an exemplary embodiment of the fastening of the motherboard and secondary board of the platform shown.

[0023] In all the accompanying drawings, similar elements have the same reference numerals. Detailed Implementation

[0024] Figure 1 The gas turbine engine 1 shown is an aircraft gas turbine engine. The gas turbine engine 1 includes a fan or propeller 2, a compressor section 3, a combustion chamber 4, a turbine section 5 downstream of the combustion chamber 4, and an exhaust casing 50. The gas turbine engine 1 also includes a stationary blade assembly 6.

[0025] exist Figure 1In the example shown, the stationary blade assembly 6 is a stationary blade assembly arranged to rectify the airflow downstream of the propeller 2. However, the invention applies to any stationary (i.e., non-rotating) blade assembly of the gas turbine engine 1, whether it is a rectifying blade assembly for a fan or propeller, a rectifying blade assembly for the compression section, or a nozzle blade assembly for the turbine section. The stationary blade assembly 6 includes a hub 8 and at least one blade 7, the hub being fixedly mounted relative to the housing 9 of the gas turbine engine 1, and at least one blade extending from the hub 8. As an example, the gas turbine engine 1 may particularly include a USF-type turboprop engine comprising a ductless propeller 2, in which case the stationary blade assembly 6 is ductless and extends downstream of the propeller 2 (see [link to documentation]). Figure 1 In another example, the gas turbine engine 1 may be a turbojet engine including a ductless fan, in which case the stationary blade assembly 6 may correspond to a ductless stator blade (referred to as an "outlet guide blade") extending downstream of the fan.

[0026] In this specification, the terms "upstream" and "downstream" are defined relative to the direction of the gas flow passing through the stationary blade assembly 6. The axis of rotation of the propeller 2 (or correspondingly, the fan) rotor is referred to as the X-axis. In other words, the X-axis is the central axis of the gas turbine engine 1. The axial direction corresponds to the direction of the X-axis, and the radial direction is the direction perpendicular to and passing through the X-axis. Furthermore, the circumferential (or tangential) direction corresponds to the direction perpendicular to the longitudinal axis X and does not pass through the longitudinal axis X.

[0027] exist Figure 1 In the example shown, the stationary blade assembly 6 includes a plurality of blades 7 extending radially from the hub 8. Optionally, the blades 7 have variable pitch, i.e., each blade 7 is mounted to pivot about a corresponding pitch axis Y on the hub 8 of the stationary blade assembly 6 (the pitch axis Y is fixed and is generally radial relative to axis X). However, this is not limiting, as the blades 7 can be fixed relative to the hub 8 (particularly when the fan is ducted), in which case the hub 8 corresponds to a shroud for the intermediate housing (located between the low-pressure compressor housing and the high-pressure compressor housing).

[0028] The stationary blade assembly 6 therefore includes a hub 8 fixedly mounted relative to the housing 9 of the gas turbine engine 1. Thus, the stationary blade assembly does not rotate. The blades 7 of the blade assembly 6 extend, for example, approximately radially along the pitch axis Y relative to the X-axis.

[0029] In one embodiment where the blade 7 has a variable pitch, the blade assembly 6 includes an actuation mechanism 10 for changing the pitch angle of the blades 7 of the blade assembly 6 to adapt the performance of the gas turbine engine 1 to different flight phases. Furthermore, the blade assembly 6 includes a fastener 8a (or pivot) connecting the blades 7 to the hub 8. More specifically, each blade 7 includes a blade root 11, which is mounted in the hub 8 by the fastener 8a. The fastener 8a is mounted such that it is rotatable relative to the hub 8 about the pitch axis Y of the blade 7. More specifically, the fastener 8a is rotatably mounted, for example, by a rolling element (e.g., a ball bearing) within a receiving portion disposed in the hub 8.

[0030] In the following description, for simplicity, the invention will be described in the case of a variable pitch blade 7, in which the blade root 11 is mounted in the hub 8 by fasteners 8a. However, this disclosure is not limited to this configuration, and the disclosure is adapted, with necessary modifications, to a fixed blade 7 whose root is directly mounted on the hub 8.

[0031] In addition to the blade root 11 which is intended to be connected to the hub 8, the blade 7 also includes an aerodynamic profile airfoil 12, which is designed to be arranged in the airflow to generate lift when the gas turbine engine 1 is running.

[0032] The aerodynamic profile airfoil 12 has a domed back 7a, a flared front 7b, a leading edge 7c, and a trailing edge 7d. The leading edge 7c is configured to extend facing the airflow entering the gas turbine engine 1. The leading edge 7c corresponds to the front portion of the aerodynamic profile, which faces the airflow and divides the airflow into a flared front flow and a domed back flow. The trailing edge 7d corresponds to the rear portion of the aerodynamic profile 24, where the flared front flow and the domed back flow meet.

[0033] The blade root 11 includes a lower first blade root portion 11a and an upper second blade root portion 11b. The lower first blade root portion is intended to be connected to the hub 8 via a fastener 8a. When the first blade root portion 11a is connected to the hub 8 of the gas turbine engine 1, the upper second blade root portion extends radially from the first portion 11a relative to the axis X. Therefore, the second portion 11b may be integrally connected to the aerodynamic profile airfoil portion 12. When the blade 7 is rigidly attached to the hub 8, the first portion 11a extends facing the hub 8, while the second portion 11b extends radially relative to the hub 8.

[0034] The blade root 11 also includes a joining portion 11c that connects the first portion 11a and the second portion 11b of the blade root 11.

[0035] In a first embodiment, the blade 7 is made of a composite material. The blade 7 includes a fiber reinforcement and a matrix (e.g., a polymer matrix), with the fiber reinforcement embedded in the matrix. The fiber reinforcement may include a three-dimensional braided structure or a woven fiber structure. Furthermore, the fiber reinforcement is manufactured such that it includes warp yarns that extend continuously within the aerodynamic profile airfoil portion 12 and within the blade root portion 11. Alternatively, the fiber reinforcement may include a stacked two-dimensional fiber structure. The fibers of the fiber reinforcement include at least one of the following materials: carbon (typically silicon carbide), glass, aramid, polypropylene, and / or ceramic (typically oxide ceramic). The matrix typically includes an organic (thermosetting, thermoplastic, or elastomer) matrix or a carbon matrix. For example, the matrix includes a plastic material, typically a polymer such as epoxy, bismaleimide, or polyimide.

[0036] Specifically, the blade 7 may include a first skin 13a and a second skin 13b, which are connected to each other and extend generally opposite to each other. Specifically, the skins 13a and 13b are connected at the blade tip along the entire chord of the blade 7, at the leading edge 7c, and at the trailing edge 7d. The skins 13a and 13b may be integral and integrally formed from fiber preforms of variable thickness. Alternatively, the first skin 13a may be formed from a first portion of fiber reinforcement to form an arched back 7a, and the second skin 13b may be formed from a second portion of fiber reinforcement to form an arched belly 7b, and the first and second portions of fiber reinforcement may then be connected, for example, near the blade tip.

[0037] The first skin 13a and the second skin 13b can be separated by a cavity 14, which opens on the lower surface of the blade root 11 to reduce the mass of the blade 7. The lower surface of the blade root 11 corresponds to the surface of the blade root 11 intended to face the fastener 8a. The cavity 14 does not open at the blade tip, nor at the leading edge 7c or trailing edge 7d of the blade 7.

[0038] The first skin 13a has a first lower end portion 15a, which faces the tip of the blade 7 and thus the aerodynamic profile airfoil portion 12. Similarly, the second skin 13b has a second lower end portion 15b, which faces the tip of the blade 7 and thus the aerodynamic profile airfoil portion 12. At the first end portion 15a and the second end portion 15b of the skin facing the blade tip, non-interconnected portions are formed in the fiber reinforcement, such that the first end portion 15a and the second end portion 15b are free. Thus, the first end portion 15a and the second end portion 15b form a first portion 11a and are spaced apart to extend tangentially against the fastener 8a (or, if necessary, by means of the composite opposing plate 18 described below).

[0039] Alternatively, the blade 7 is made of metal: in this case, the blade root 11 is shaped such that the end of the blade root opposite the blade tip includes an arched back portion and an arched belly portion, which flare outward from the second portion 11b to extend substantially tangentially against a fastener to form two wings constituting the first portion 11a of the blade root 11.

[0040] The blade 7 may also include a formed element 16 extending between the first skin 13a and the second skin 13b (or between the two airfoils in the case of a metal blade) within the blade root joint 11c in the cavity 14. Therefore, the formed element 16 strengthens the rigidity of the blade 7 at the bottom of the airfoil. The formed element 16 also ensures a gradual transition of rigidity in the radial direction to avoid stress concentration at the exit of the clamping area.

[0041] Specifically, the molded part 16 can be made of a composite material. The molded part 16 includes a fiber reinforcement and a matrix, preferably a matrix compatible with (typically the same matrix) the fiber reinforcement of the skins 13a, 13b, with the fiber reinforcement embedded within the matrix. The fiber reinforcement of the molded part 16 may include fibers with a Young's modulus greater than 15 GPa, such as carbon fibers, aramid fibers, glass fibers, or basalt fibers.

[0042] The blade 7 may also include one or more fillers 50, which are typically made of foam or any other suitable material with a lower density than the composite material of the first skin 13a and the second skin 13b. The fillers are arranged within the blade 7 between the first skin 13a and the second skin 13b, between the forming member 16 and the blade tip (the upper end of the blade 7), i.e., in the portion of the cavity 14 extending within the airfoil 12, to reinforce the blade 7 and / or impart the final shape to the airfoil 12 of the first skin 13a and the second skin 13b.

[0043] In one embodiment, particularly when the impeller 7 is made of a composite material, the joining portion 11c has a first concave outer surface 17a disposed on the arched back side 7a and a second concave outer surface 17b disposed opposite the first concave outer surface 17a on the arched belly side 7b. The first and second outer surfaces 17a and 17b may be integral with each other and may be curved to limit the risk of breakage between the first and second portions 11a and 11b when forces are transmitted between them. The first outer surface 17a may comprise two portions joined together by a common joining line, such that the first outer surface 17a forms a non-right angle with the joining line as its vertex. In this case, the second outer surface 17b may also comprise two portions joined together by a common joining line, such that the second outer surface 17b forms a non-right angle with the joining line as its vertex. Therefore, the two portions of the joining portion 11c may, for example, form a right angle.

[0044] The fastener 8a has an upper radial surface 8b and is designed to receive a first portion 11a, more specifically, a lower radial end 11e of the blade root 11. Therefore, the first portion 11a is arranged abutting against the upper radial surface 8b of the fastener 8. Alternatively, and as... Figure 2 As shown in the example, the composite counterplate 18 can be arranged between the first portion 11a and the upper radial surface 8b to close the blade root 11, thereby absorbing forces and increasing the rigidity of the blade 7.

[0045] The blade 7 also includes a first reinforcing body 19a and a second reinforcing body 19b, which are arranged on both sides of the blade root 11, more specifically, on both sides of the joint portion 11c. In particular, the first reinforcing body 19a is arranged on the arched back side 7a of the blade 7, and the second reinforcing body 19b is arranged on the arched belly side 7b.

[0046] The first reinforcement 19a includes a first reinforcement portion 20a, a second reinforcement portion 22a, and a first reinforcement joining portion 21a, which connects the first reinforcement portion 20a and the second reinforcement portion 22a. The first reinforcement portion 20a contacts the first blade root portion 11a, the second reinforcement portion 22a contacts the second blade root portion 11b, and the first reinforcement joining portion 21a contacts the blade root joining portion 11c.

[0047] More specifically, the first reinforcement 19a includes a first convex inner surface 23a, which is configured to closely fit the first concave outer surface 17a of the engagement portion 11c of the blade root 11. More specifically, the first convex inner surface 23a is the surface of the first reinforcement engagement portion 21a, such that the first reinforcement 19a closely fits the arched back 7a of the blade root 11, thereby strengthening the blade 7 and absorbing as much force as possible along the length of the first reinforcement 19a. The first inner surface 23a may include two portions joined together by a common engagement line, such that the first inner surface 23a forms a non-right angle with the engagement line as its apex.

[0048] Similarly, the second reinforcement 19b includes a third reinforcement portion 20b, a fourth reinforcement portion 22b, and a second reinforcement joining portion 21b, which connects the third reinforcement portion 20b and the fourth reinforcement portion 22b. The third reinforcement portion 20b contacts the first blade root portion 11a, the fourth reinforcement portion 22b contacts the second blade root portion 11b, and the second reinforcement joining portion 21b contacts the blade root joining portion 11c.

[0049] More specifically, the second reinforcement 19b includes a second convex inner surface 23b, which is configured to closely fit the second concave outer surface 17b of the engagement portion 11c of the blade root 11. More specifically, the second convex inner surface 23b is the surface of the third reinforcement engagement portion 21b, such that the second reinforcement 19b substantially closely fits the arched surface 7b of the blade root 11, thereby strengthening the blade 7 and absorbing as much force as possible along the length of the second reinforcement 19b. The second inner surface 23b may include two portions joined together by a common engagement line, such that the second inner surface 23b forms a non-right angle with the engagement line as its apex.

[0050] Therefore, the first reinforcement 19a and the second reinforcement 19b can strengthen the joint portion, which is an area with a high risk of fracture when the gas turbine engine 1 is in use.

[0051] The impeller 7 also includes a first fastening member 24a, which is configured to mechanically fasten the first reinforcement 19a to the root portion 11a of the first impeller. More specifically, the first fastening member 24a can fasten the first reinforcement 19a to the first portion 11a and, where applicable, to the composite opposed plate 18.

[0052] In addition, in order to limit the number of components, the first fastening member 24a may also be configured to fasten the blade 7 (and more specifically, the portion of the first skin 13a that forms part of the root portion 11a of the first blade) to the fastener 8a and thus to the hub 8.

[0053] The blade 7 is also connected to the hub 8, for example, via a first fastening member 24a.

[0054] For example, the first fastening member 24a includes one or more bolts with bolt shanks passing through the first part 11a (more specifically, the first skin 13a), the composite opposing plate 18, and the fastener 8a.

[0055] Therefore, the first fastening member 24a enables the first reinforcement 19a to be fastened to the blade 7 (on the first portion 11a) firstly, and then the blade 7 (the first portion 11a and the composite opposed plate 18, if applicable) to be connected to the hub 8.

[0056] Furthermore, and more specifically, the first fastening member 24a also fastens the first reinforcement portion 20a to the first blade root portion 11a. Additionally, as described above, the first fastening member 24a fastens the first blade root portion 11a to the hub portion 8 via fastener 8a. Therefore, the first fastening member 24a is located at a distance from the blade root engagement portion 11c and the first reinforcement engagement portion 21a, which are portions with a high risk of fracture.

[0057] Similarly, the impeller 7 also includes a second fastening member 24b, which is configured to mechanically fasten the second reinforcement 19b to the root portion 11a of the first impeller. More specifically, the second fastening member 24b can fasten the second reinforcement 19b to the first portion 11a and, where applicable, to the composite opposed plate 18.

[0058] In addition, to limit the number of components, the second fastening member 24b can also be configured to fasten the wheel blade 7 (and more specifically, the portion of the second skin 13b that forms part of the root portion 11a of the first wheel blade) to the fastener 8a and thus to the hub 8.

[0059] The blade 7 is also connected to the hub 8, for example, via a second fastening member 24b.

[0060] For example, the second fastening member 24b includes one or more bolts with bolt shanks passing through the first part 11a (more specifically, the second skin 13b), the composite opposing plate 18, and the fastener 8a.

[0061] Therefore, the second fastening member 24b enables the second reinforcement 19b to be fastened to the blade 7 (on the first part 11a) and to connect the blade 7 (and, if applicable, the first part 11a and the composite opposed plate 18) to the hub 8.

[0062] Furthermore, and more specifically, the second fastening member 24b also fastens the third reinforcement portion 20b to the first blade root portion 11a. Additionally, as described above, the second fastening member 24b fastens the first blade root portion 11a to the hub portion 8 via fastener 8a. Therefore, the second fastening member 24b is located at a certain distance from the blade root joint portion 11c and the second reinforcement joint portion 21b, which are portions with a high risk of fracture.

[0063] Therefore, on the one hand, the blade 7 is fixedly mounted on the hub 8 by the first fastening member 24a and the second fastening member 24b.

[0064] On the other hand, the first reinforcement 19a and the second reinforcement 19b thus enable the reinforcement of the area where the blade 7 is fastened to the hub 8, which is an area with a high risk of fracture during the use of the gas turbine engine 1, because this is where forces are concentrated to transmit power between the blade 7 and the hub 8.

[0065] The first reinforcement 19a and the second reinforcement 19b can be formed as a single integral part around the blade root 11, or alternatively as two separate parts.

[0066] In any case, the contact and fastening of the first reinforcement 19a and the second reinforcement 19b with the blade root 11 (particularly with the blade root engagement portion 11c (with the first concave outer surface 17a and the second concave outer surface 17b, respectively) allows the first skin 13a and the second skin 13b to be pressed against the forming member 16, thereby limiting out-of-plane stress in the concave region of the blade root 11, which is highly detrimental to the integrity of the interface between the skins 13a, 13b and the forming member 16. Therefore, the first reinforcement 19a and the second reinforcement 19b can reduce localized stress by better distributing the transmitted forces. The reduction in stress is even greater when the first outer surface 17a and the second outer surface 17b bend.

[0067] Advantageously, the first reinforcement 19a and the second reinforcement 19b are metallic, particularly made of titanium.

[0068] Additionally, the blade 7 includes an aerodynamic platform 25, an example of which is shown in... Figure 2 (cross section), Figure 3 (Top view) and Figure 4 As shown in the image.

[0069] In one embodiment, platform 25 further includes a first platform portion 25a and a second platform portion 25b. The first platform portion 25a includes a main plate 28a configured to support and abut against a first surface of the blade 7, for example, against the arched back surface 7a of the blade 7, particularly against a second portion 11b of the blade root on the arched back surface 7a. The second platform portion 25b includes a secondary plate 28b extending as an extension of the main plate 28a. Furthermore, the secondary plate 28b is fastened to the main plate 28a and configured to support and abut against a second surface of the blade 7, for example, against the arched surface 7b of the blade 7, particularly against the second blade root portion 11b on the arched surface 7b.

[0070] Platform 25 includes a radially outer surface 26 (relative to the X-axis) and a radially inner surface 27, the radially outer surface being configured to define an airflow path through the impeller assembly 6. Therefore, platform 25 can define a radially lower portion of the airflow path through the impeller assembly 6.

[0071] Platform 25 can be made of plastic or metal (such as polyetheretherketone (PEEK), polyetherimide (PEI), aluminum-based alloys) or composite materials.

[0072] The first platform portion 25a is preferably cantilevered on the second platform portion 25b, for example... Figure 4 As shown. The first platform portion 25a is preferably mechanically fastened to the second platform portion 25b, which allows the platform 25 to be removed from the wing and the blades 7 to be interchanged, while minimizing mass and geometric imperfections in the airflow path (i.e., upward or downward steps, depending on the direction of the aerodynamic flow). More specifically, it is sufficient to remove either the first platform portion 25a or the second platform portion 25b to access the fastener 8a, the blade root 11, and the hub 8.

[0073] Specifically, the first platform portion 25a can be riveted to the second platform portion 25b by means of rivets 47. For this purpose, the first platform portion 25a includes a first connecting edge 29a, and the second platform portion 25b includes a second connecting edge 29b. The first connecting edge 29a and the second connecting edge 29b are assembled by riveting and can be tapered, such that the thickness of the platform 25 is approximately constant (e.g., ...). Figure 4 (As shown). In other words, the sum of the thicknesses of the first connecting edge 29a and the second connecting edge 29b at the joint is approximately equal to the thicknesses of the main board 28a and the secondary board 28b away from the joint.

[0074] When platform 25 is made of one of the materials listed above, the thickness of platform 25 can be between 5 mm and 15 mm, preferably between 10 mm and 15 mm. Furthermore, when the main board 28a and the secondary board 28b are riveted, they overlap on a strip whose width is at least 1.5 times the diameter of the rivet 47. For example, a rivet 47 with a diameter of 6 mm can be used.

[0075] In addition, to prevent air from re-flowing into the root of the impeller 11, the impeller 7 also includes a seal 48 installed between the main plate 28a and the secondary plate 28b, between the main plate 28a and the arch back 7a of the impeller 7, and between the secondary plate 28b and the arch belly 7b of the impeller 7. The seal 48 can be co-injected with the platform 25, or it can be fastened and glued after the platform 25 has been manufactured. The seal 48 can be made of, for example, rubber (typically ethylene propylene diene monomer (EPDM)), silicone, or any suitable material.

[0076] The main board 28a and the secondary board 28b extend in a generally circumferential direction relative to the X-axis, that is, in a plane orthogonal to the pivot axis Y, to define the flow path radially inward.

[0077] Alternatively, platform 25 can be a single unit that encloses the blade root 11.

[0078] Furthermore, platform 25 is fixedly mounted on a first reinforcement 19a and / or a second reinforcement 19b that supports platform 25. More specifically, blade 7 (more particularly, the first reinforcement 19a) includes a first support arm 30a, and / or blade 7 (more particularly, the second reinforcement 19b) includes a second support arm 30b. The first support arm 30a and the second support arm 30b are arranged on both sides of blade root 11. For example, the first support arm 30a is arranged on the arched back side 7a of blade root 11, and the second support arm 30b is arranged on the arched belly side 7b of blade root 11. The first support arm 30a includes a first lower end portion 31a fastened to the first reinforcement 19a and a corresponding second upper end portion 32a fastened to platform 25. Similarly, the second support arm 30b includes a third lower end portion 31b fastened to the second reinforcement 19b and a corresponding fourth upper end portion 32b fastened to platform 25.

[0079] More specifically, the first lower end 31a of the first support arm 30a is fastened to the first reinforcement portion 20a. Therefore, the first support arm 30a is fastened to the first reinforcement 19a at a distance from the first reinforcement engagement portion 21a to protect that area. Furthermore, the second upper end 32a of the first support arm 30a is fastened to the platform 25 on its inner surface 27 at a point smaller than the distance from the first lower end 31a from the Y-axis, and more specifically, to the first platform portion 25a. Therefore, the support of the platform 25 is strengthened and it can transmit large forces, such as forces caused by the action of air in contact with and flowing through the platform 25.

[0080] The second upper end 32a of the first support arm is fastened to the platform, for example, by a third fastening member 33a.

[0081] In one embodiment, the third fastening member 33a includes a screw that engages with a countersunk hole manufactured in the platform 25 (more precisely, in the first platform portion 25a) and a threaded hole manufactured in the second upper end portion 32a of the first support arm 30a, the threaded hole being open at the second upper end portion 32a of the first support arm 30a. The countersunk hole is accessible from the upper surface 26 of the platform 25 such that the screw can be inserted or removed once the various portions of the platform 25 have been positioned. To improve aerodynamics, a plug can be inserted into the countersunk hole after the third fastening member 33a has been inserted.

[0082] In another embodiment, not shown, the third fastening member 33a includes a ring with at least partially threaded interior. Preferably, the interior of the ring includes a first internal thread at one end and a second internal thread in opposite or the same direction at opposite ends. The ring is fastened to the platform 25 via the first internal thread (e.g., to a protrusion formed on the inner surface 27) and to the second upper end portion 32a via the second internal thread. In other words, the ring surrounds the protrusion and the second upper end portion 32a to rigidly attach the first support arm 30a to the platform.

[0083] In any case, the third fastening member 33a is configured to allow the fastener to be easily installed and removed.

[0084] Similarly, the third lower end 31b of the second support arm 30b is fastened to the third reinforcing body portion 20b. Thus, the second support arm 30b is fastened to the second reinforcing body 19b at a distance from the second reinforcing body engagement portion 21b to protect that area. Furthermore, the fourth upper end 32b of the second support arm 30b is fastened to the platform 25 on the inner surface 27 at a point smaller than the distance from the third lower end 31b of the second support arm 30b to the Y-axis, and more specifically, to the second platform portion 25b. Therefore, the support of the platform 25 is strengthened and it can better transmit greater forces, such as those caused by the action of air in contact with and flowing through the platform 25.

[0085] The fourth upper end 32b of the second support arm 30b is fastened to the platform 25, for example, by a fourth fastening member 33b.

[0086] In one embodiment, the fourth fastening member includes a screw that engages with a countersunk hole manufactured in platform 25 (more precisely, in the second platform portion 25b) and a threaded hole manufactured in the fourth upper end portion 32b of the first support arm 30b, the threaded hole being open at the fourth upper end portion 32b of the second support arm 30b. To improve aerodynamics, a plug can be inserted into the countersunk hole after the third fastening member 33a has been inserted.

[0087] In another embodiment, not shown, the fourth fastening member 33 includes a ring with at least partially threaded interior. Preferably, the interior of the ring includes a first internal thread at one end and a second internal thread in opposite or the same direction at opposite ends. The ring is fastened to the platform 25 via the first internal thread (e.g., to a protrusion formed on the inner surface 27) and to the fourth upper end portion 32b via the second internal thread. In other words, the ring surrounds the protrusion and the fourth upper end portion 32b to rigidly attach the second support arm 30b to the platform 25.

[0088] In any case, the fourth fastening member is configured to allow the fastener to be easily installed and removed.

[0089] To increase rigidity, the first support arm 30a may include a first rod 34a extending between the first support arm and the second reinforcement portion 21a to always protect the reinforcement joint portion 21a while improving the force absorption of the first reinforcement 19a. Similarly, the second support arm 30b may include a second rod 34b extending between the second support arm 30b and the fourth reinforcement portion 22b to always protect the reinforcement joint portion 21a while improving the force absorption of the second reinforcement 19b.

[0090] In one embodiment, the blade 7 may include a plurality of first support arms 30a and / or a plurality of second support arms 30b. For example, the blade 7 may include at least two first support arms 30a and two second support arms 30b. Considering the conventional dimensions (particularly the chord dimensions) of the stator blade 7 and therefore the first reinforcement 19a and the second reinforcement 19b, the blade 7 may include up to four first support arms 30a and four second support arms 30b. For example, the blade includes three first reinforcement arms 30a and three second reinforcement arms 30b. Thus, the first support arms 30a and the second support arms 30b are substantially identical to each other, i.e., the first support arms 30a are identical to each other and the second support arms 30b are identical to each other. In all embodiments having a plurality of first support arms 30a and / or second support arms 30b, the first support arms 30a and the second support arms 30b are arranged to extend along the first reinforcement 19a and the second reinforcement 19b, respectively, and thus substantially along the blade 7 between the leading edge 7c and the trailing edge 7d, i.e., along the chord of the blade 7. For example, the first support arm 30a can be arranged on the arched back side 7a of the blade root 11, and the second support arm 30b can be arranged on the arched belly side 7b of the blade root 11. In other words, the first support arm 30a and the corresponding second support arm 30b are aligned approximately along the blade root 11a between the leading edge 7c and the trailing edge 7d. Preferably, therefore, two consecutive first support arms 30a or two consecutive second support arms 30b along the blade root 11 are spaced apart at a constant distance to uniformly distribute force absorption along the first reinforcement 19a and the second reinforcement 19b.

[0091] Therefore, platform 25 can be connected and secured to the first support arm 30a and / or the second support arm 30b. Alternatively, platform 25 can be integral with at least one of the first support arm 30a or the second support arm 30b, and connected and secured to the other of the first support arm 30a and the second support arm 30b. Finally, platform 25 can be integral with all of the first support arms 30a and the second support arms 30b.

[0092] Finally, each of the first support arms 30a may include a first rod 34a as defined above. Similarly, each of the second support arms 30b may include a second rod 34b as defined above.

[0093] Therefore, the proposed blade assembly 6 enables the reduction of the tangential and radial space occupied by the blade root 11 while maintaining the integrity of the blade root 11, ensuring the absorption of forces between the blade 7 and the hub 8, and ensuring the support of the aerodynamic platform 25.

Claims

1. A blade (7) for a stationary blade assembly (6) of a gas turbine engine (1), the blade comprising: - Blade root (11), the blade root includes a first blade root portion (11a) and a second blade root portion (11b), the first blade root portion is configured to be connected to the hub (8) of the gas turbine engine (1), the second blade root portion extends radially from the first blade root portion (11a) relative to the central axis (X) of the gas turbine engine (1) when the first blade root portion (11a) is connected to the hub (8) of the gas turbine engine (1), the first blade root portion and the second blade root portion (11a, 11b) are connected by a blade root joining portion (11c), the blade root joining portion (11c) has a first outer surface (17a) and a second outer surface (17b) opposite to the first outer surface (17a). - A first reinforcement (19a) and a second reinforcement (19b) are arranged on both sides of the blade root joint portion (11c). The first reinforcement (19a) has a first inner surface (23a) configured to fit tightly against the first outer surface (17a) of the blade root joint portion (11c). The second reinforcement (19b) has a second inner surface (23b) configured to fit tightly against the second outer surface (17b) of the blade root joint portion (11c). as well as - A first fastening member (24a) and a second fastening member (24b), the first fastening member being configured to mechanically fasten the first reinforcement (19a) to the root portion of the first blade (11a), and the second fastening member being configured to mechanically fasten the second reinforcement (19b) to the root portion of the first blade (11a).

2. The blade according to claim 1, wherein, The first outer surface (17a) is concave, the second outer surface (17b) is concave, the first inner surface (23a) of the first reinforcement is convex, and the second inner surface (23b) of the second reinforcement is convex.

3. The impeller (7) according to claim 2, comprising a platform (25) including a radially outer surface (26) and a radially inner surface (27), the radially outer surface being configured to define an airflow path through the impeller assembly, the platform (25) being fixedly mounted on the first reinforcement (19a) and / or the second reinforcement (19b).

4. The impeller (7) according to claim 3, comprising a first support arm (30a) and / or a second support arm (30b), the first support arm comprising a first end (31a) fixed to the first reinforcement (19a) and an opposing second end (32a) fixed to the platform (25), the second support arm comprising a third end (31b) fixed to the second reinforcement (19b) and an opposing fourth end (32b) fixed to the platform (25).

5. The blade (7) according to claim 4, wherein, The first reinforcement (19a) includes a first reinforcement portion (20a) that contacts the first blade root portion (11a), a second reinforcement portion (22a) that contacts the second blade root portion (11b), and a first reinforcement engagement portion (21a) that connects the second reinforcement portion (22a) to the first reinforcement portion (20a). The convex first inner surface (23a) is the surface of the first reinforcement engagement portion (21a), and the first end (31a) of the first support arm is fixed to the first reinforcement portion (20a).

6. The blade (7) according to claim 4 or 5, wherein, The second reinforcement (19b) includes a third reinforcement portion (20b) that contacts the first blade root portion (11a), a fourth reinforcement portion (22b) that contacts the second blade root portion (11b), and a second reinforcement joining portion (21b) that connects the fourth reinforcement portion (22b) to the third reinforcement portion (20b). The convex second inner surface (23b) is the surface of the second reinforcement joining portion (21b), and the third end (31b) of the second support arm is fixed to the third reinforcement portion (20b).

7. The blade according to any one of claims 4 to 6, wherein, Each of the reinforcements (19a; 19b) includes a plurality of first support arms (30a) distributed along the arched back surface (7a) of the blade (7) and a plurality of second support arms (30b) distributed along the arched belly surface (7b) of the blade (7).

8. The blade (7) according to any one of claims 3 to 7, wherein, The platform (25) includes a first platform portion (25a) fixed to the first reinforcement (19a) and a second platform portion (25b) fixed to the second reinforcement (19b).

9. The blade (7) according to any one of claims 4 to 8, wherein, The second end (32a) of at least one of the first support arms (30a) and / or the fourth end (32b) of at least one of the second support arms (30b) are secured to the platform (25) by a third fastening member (33a; 33b).

10. The blade (7) according to any one of claims 1 to 9, wherein, Each of the reinforcements (19a; 19b) is metallic.

11. The blade (7) according to any one of claims 1 to 10, comprising a first skin (13a), a second skin (13b), and a forming member (16), the first skin and the second skin being made of a composite material composed of fiber preforms and being joined to each other at the leading edge (7c) and the trailing edge (7d) of the blade (7), the forming member extending between the first skin (13a) and the second skin (13b) at the root joint portion (11c) of the blade.

12. A blade assembly (6) of a gas turbine engine (1) comprising at least one blade (7) according to any one of claims 1 to 11 and a hub (8), the blade (7) being fixedly mounted on the hub (8) by a first fastening member (24a) and a second fastening member (24b).

13. The blade assembly (6) according to claim 12, comprising a fastener (8a) adapted to be pivotally mounted relative to the hub (8) about a pitch axis (Y), and wherein, The blade (7) is fixedly mounted on the fastener (8a) by the first fastening member (24a) and the second fastening member (24b).

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