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CN-122029102-A - Variable pitch vane for an aircraft turbine engine propeller

CN122029102ACN 122029102 ACN122029102 ACN 122029102ACN-122029102-A

Abstract

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

Inventors

  • Vincent juden
  • Matteo Minavino
  • Francois Charlie
  • Laurent Jablonsky
  • Arnold Puggis

Assignees

  • 赛峰飞机发动机公司

Dates

Publication Date
20260512
Application Date
20241016
Priority Date
20231019

Claims (20)

  1. 1. A variable pitch blade (10) for an aircraft turbine engine propeller, the blade (10) comprising a blade (12) connected by a strut (30) to a root (14) centred on a pitch axis (a) of the blade, the blade (10) being made of a composite material consisting of at least one metal body (50) and a fibrous preform (52), the fibrous preform being obtained by three-dimensional fibre braiding and being embedded in a polymer matrix ensuring that the preform (52) is fixed to a metal body (50), the metal body (50) forming at least the root (14) and the strut (30) and the fibrous preform (52) forming at least the blade (12), the blade comprising a pressure side (12 a) and a suction side (12 b) connected together by a leading edge (12 c) and a trailing edge (12 d) of the blade (12), characterized in that the metal body (50) also forms a first portion (54 a) of a spar (54), the first portion (50) forming at least the root (14) and the strut (30) forming at least the blade (12) and the blade (12) comprising a pressure side (12 b) connected together by a leading edge (12 d) and a suction side (12 b) along the pitch axis (54) extending inside the spar (54), the second portion (54 b) of the spar (54) has a longitudinal end (54 b 1) on the root (14) side, which is engaged in a cavity (56) of the first portion (54 a) of the spar (54).
  2. 2. The blade (10) of claim 1, wherein the preform (52) includes at least one fiber untwisted portion of the preform thereof, the at least one fiber untwisted portion enabling definition of: An upper section (52 a) of the blade (12) in which the fibres are fixed to each other throughout the thickness of the blade, -A lower section (52 b) of the blade (12) connected to the upper section (52 a) and comprising two overlapping skins (52 b1,52b 2) on one side respectively of the pressure side (12 a) and suction side (12 b) of the blade (12), the fibres of each of the skins (52 b1,52b 2) being fixed to each other over the whole thickness of the skins (52 b1,52b 2), and A second portion (54 b) of the spar (54) arranged between two skins (52 b1,52b 2) of the lower section (52 b) and comprising a longitudinal end (54 b 1) opposite the root (14) connected to the upper section (52 a), -A first portion (54 a) of the spar (54), which is arranged between the two skins (52 b1,52b 2).
  3. 3. The blade (10) according to any of the preceding claims, wherein the blade is additionally made of at least one foam block (58, 60) located inside the blade (12).
  4. 4. A blade (10) according to claim 3 when appended to claim 2, wherein a first foam block (58) is arranged between the skins (52 b1,52b 2), between the second portion (54 b) of the spar (54) and the leading edge (12 c) of the blade (12), and a second foam block (60) is arranged between the skins (52 b1,52b 2), between the second portion (54 b) of the spar (54) and the trailing edge (12 d) of the blade (12).
  5. 5. The blade (10) according to any of the preceding claims, wherein the first portion (54 a) of the spar (54) has a dimension (L1) measured along the chord of the blade (12), the dimension representing 80 to 120% of the dimension (L2) of the second portion (54 b) of the spar (54) measured in the same way.
  6. 6. The blade (10) according to any of the preceding claims, wherein the first portion (54 a) of the spar (54) comprises two wings (62 a,62 b) located on one of a pressure side (12 a) and a suction side (12 b) of the blade (12), respectively, and defining the cavity (56) therebetween, which cavity is open on the second portion (54 b) side of the spar (54) in the direction of the pitch axis (a) to form a receiving opening for receiving the second portion (54 b).
  7. 7. The blade (10) according to claim 6, wherein the cavity (56) is open at two opposite ends thereof, the two opposite ends being located on a leading edge (12 c) side and the trailing edge (12 d) side of the blade (12), respectively.
  8. 8. The blade (10) according to claim 6, wherein the first portion (54 a) of the spar (54) comprises two bulkheads (64 a,64 b) located on the leading edge (12 c) side and the trailing edge (12 d) side of the blade (12), respectively, these bulkheads (64 a,64 b) extending between and connecting the wing portions (62 a,62 b) to close two opposite ends of the cavity (56) located on the leading edge (12 c) side and the trailing edge (12 d) side, respectively.
  9. 9. The blade (10) according to any one of claims 6 to 8, wherein the receiving opening is defined by two longitudinal edges (62 a1,62b 1) of the airfoil (62 a,62 b), which are tapered.
  10. 10. The blade (10) according to any one of claims 6 to 9, wherein the airfoil (62 a,62 b) has: -a constant thickness (E1) along the pitch axis (a) over at least 50% or even 80% of its height, or -A thickness (E1) varying along the pitch axis (a) and being maximum, for example, on the root (14) side or half height side of the wing.
  11. 11. The blade (10) according to any one of claims 6 to 10, wherein the airfoil (62 a,62 b) has: -a constant thickness (E2) along the chord of the blade (12) over at least 50% or even 80% of its length, or -A thickness (E2) varying along the chord of the blade (12) and for example being greatest in the middle of the wing portions (62 a,62 b).
  12. 12. The wheel blade (10) according to any of claims 6 to 11, wherein the wing (62 a,62 b): linear and parallel, or Is curved, has recesses oriented towards each other, -Having a first boss oriented towards each other and a second boss oriented towards one of a pressure side (12 a) and a suction side (12 b) of the blade (12), respectively.
  13. 13. The blade (10) according to any one of claims 6 to 12, wherein the first portion (54 a) of the spar (54) further comprises at least one dividing wall (66, 68) of the cavity (56), the wall (66, 68) extending between and spaced apart from the wings (62 a,62 b) or extending from one wing to the other and being connected thereto.
  14. 14. The blade (10) according to claim 13, wherein the wall (66, 68) has a height (H1) along the pitch axis (a) representing between 30% and 60% of the height (H2) of the airfoil (62 a,62 b) along the axis (a).
  15. 15. The blade (10) according to claim 13 or 14 when dependent on claim 8, wherein the wall (66, 68) is connected to the partition (64 a, 64 b).
  16. 16. The blade (10) according to any one of the preceding claims, wherein the metal body (50): -the formation of the one part is carried out, -Formed by at least two portions (50 a,50 b), a first portion (50 a) comprising the root (14), the strut (30) and one of the wings (62 a), and a second portion (50 b) comprising the other of the wings (62 b).
  17. 17. The blade (10) according to any one of the preceding claims, wherein the second portion (54 b) of the spar (54) comprises at least one fiber untwisted portion (70) thereof over at least a portion of its length, the first portion (54 a) of the spar (54) comprising at least one protrusion (72) engaged in the second portion (54 b) of the spar at the level of the untwisted portion (70).
  18. 18. The blade (10) according to any one of the preceding claims, wherein the second portion (54 b) of the spar (54) is connected to the skin (52 b1,52b 2) by a web (74).
  19. 19. Turbine engine, in particular for an aircraft, comprising a propeller comprising at least one vane (10) according to any one of the preceding claims.
  20. 20. A method for manufacturing a vane (10) according to any one of claims 1 to 18, wherein the method comprises the steps of: a) Braiding fibers in three dimensions to form the preform (52) and untwisting a portion of the preform (52) to form the second portion (54 b) of the spar (54), B) Compacting a second portion (54 b) of the spar (54), C) -joining a second portion (54 b) of the spar (54) in a cavity (56) of a first portion (54 a) of the spar (54), D) Compacting the assembly formed by the preform (52) and the body (50), E) The assembly is RTM-type consolidated in a mold.

Description

Variable pitch vane for an aircraft turbine engine propeller Technical Field The present invention relates to the field of aircraft turbine engines, and in particular to propulsive propellers of these turbine engines comprising variable pitch blades. Background The prior art includes in particular document FR-A1-3 017 163、FR-A1-3 080 322、FR-A1-3 112 819、FR-A1-3 121 474、US-A-4, 524, 499、FR-A1-3 120 249、US-A1-2023/0801843 and WO-A1-2023/031522. In order to minimize pollutant emissions from air transportation, it is desirable to increase the efficiency of various aspects of the turbine engine propulsion system, and more particularly to increase the propulsive efficiency, which is the efficiency of converting the energy imparted to the air passing through the engine into useful thrust. The first-order elements that affect this propulsive efficiency are elements associated with the low-pressure portion of the propulsion system that directly contribute to the generation of thrust, the low-pressure turbine, the low-pressure drive train, the fan, and the secondary flow that directs the fan airflow. A known guideline for improving propulsive efficiency is to reduce the compression ratio of the fan, thereby reducing the flow rate and associated kinetic energy losses at the engine outlet. One of the main consequences of the reduced flow rate at the engine outlet is that a greater mass of air flow must be treated at the low pressure section (secondary flow) to ensure a given thrust level set according to the characteristics of the aircraft, which therefore results in an increased bypass ratio of the engine. Bypass ratio or BPR is defined as the ratio between the mass flow through the secondary stream (cold stream) and the mass flow through the primary stream (hot stream) and specifically the mass flow that supplies the combustion chamber. A direct effect of this increase in the secondary air flow is that the diameter of the fan needs to be increased, and thus the external dimensions of the surrounding holding casing and of the nacelle constituting the aerodynamic envelope of the casing in question. To achieve a high bypass ratio, the housing may become too large and heavy (and create significant drag), so the housing is removed and replaced with a configuration with an unducted propeller. A number of concepts for an unducted turbine engine, such as an Unducted Single Fan (USF) architecture, are contemplated, including (at least) an upstream propeller wheel (or "open fan") with variable pitch and a downstream stator wheel impeller (stator) with fixed or variable pitch. The technical field of the present invention relates to ducted or unducted pitch-variable rotary (rotor) propellers or fan blades of potential application in the aeronautical propulsion industry. Other examples of architectures of particular interest are counter-rotating open rotor (CROR) and turboprop engines. On unducted architectures (USF, CROR and turboprop engines), strong vibration excitations can occur at high rotational speeds due to the influence of the direction of incoming flow on board the aircraft and upstream of infinity. In fact, unducted engines are affected by the ground and fuselage, which results in flow rate distortions in propeller supply with engine azimuth. This results in the blades of the propeller producing a vibratory response in the first (1N), second (2N), third (3N) (and possibly higher) stages of the engine. On the other hand, without the air intake sleeve, the direction of air flowing through the vanes is not parallel to the engine shaft. This slip angle produces a force called "1P" (once per revolution, i.e., 1/revolution), which causes the propeller blades to produce a vibratory response at the "1N" engine order. In a similar manner, these "1P" forces may also occur during the climb or approach phase of the aircraft, as air flows over the blades at an incident angle. These high-speed vibratory excitations create very high stress cycles throughout the vane. In particular, the portion of the blade root between the hub and the duct (also referred to as the "strut") is the load and critical area due to its retaining function on the blade. It is an object of the present invention to provide a blade whose root is capable of maintaining a "1P" load while maintaining the mass of the blade at a minimum. Disclosure of Invention The invention proposes a variable pitch blade for an aircraft turbine engine propeller, the blade comprising a blade connected by a strut to a root centred on the pitch axis of the blade, the blade being made of a composite material consisting of at least one metal body and a fibrous preform obtained by three-dimensional fibre braiding and embedded in a polymer matrix ensuring that the preform is fixed to the metal body, the metal body forming at least the root and the strut and the fibrous preform forming at least the blade, the blade comprising a pressure side and a suction side connected together by a lea