Abstract: A turboprop includes a rotary propeller upstream from an engine and an air intake that is not coaxial to the propeller, said air intake defining a conduit for supplying air to the engine and further defining a bypass to said conduit, the bypass having an outlet oriented substantially axially towards the downstream of the engine. The turboprop further includes a nacelle surrounding the engine and the air intake, wherein the air intake is secured to a housing of the engine and is not rigidly connected to the nacelle, so as to allow, during operation, relative movements between the air intake and the nacelle. The outlet is connected by a flexible link to an intake of an air circuit carried by the nacelle.

Abstract: The present disclosure relates to a propulsion assembly having a nacelle surrounding a turbojet engine, the nacelle having an upstream structure with an air inlet including an air inlet lip, the air inlet lip defining, with the air inlet lip, an air inlet chamber, the upstream structure being capable of translational movement along an axis parallel to a longitudinal axis of the nacelle, between a closed position and an open position that allows access to the turbojet engine. The propulsion assembly further includes at least one locking/unlocking device allowing and preventing movement of the upstream structure that in one form is fixed to the air inlet lip and to an upstream flange of the turbojet engine.

Abstract: The invention relates to a molding device (1) for receiving fibers onto which a resin is injected, including a base (3), side plates (5) confining the base (3), and a cover (7) capable of engaging with a free end (9) of the side plates (5) in order to define an inner space (11) for receiving the fibers and the resin, wherein said device (1) comprises, in the inner space (11), a skin (21) that is configured to be complementary to the inner surface of the cover (7) and to at least a portion of the inner surface of the side plates (5). The invention also relates to a method for manufacturing a resin part using such a device (1).

Abstract: The present disclosure relates to a propulsion assembly having a nacelle surrounding a turbojet engine, the nacelle having an upstream structure with an air inlet including an air inlet lip, the air inlet lip defining, with the air inlet lip, an air inlet chamber, the upstream structure being capable of translational movement along an axis parallel to a longitudinal axis of the nacelle, between a closed position and an open position that allows access to the turbojet engine. The propulsion assembly further includes at least one locking/unlocking device allowing and preventing movement of the upstream structure that in one form is fixed to the air inlet lip and to an upstream flange of the turbojet engine.

Abstract: The invention related to a device for making a part of a composite material by resin-transfer molding, that includes a manufacturing mold in which is placed a blank of the part to be impregnated with resin, characterized in that the manufacturing mold is divided into a matrix and one or more mobile structural member to be indexed on the matrix, the whole defining, after assembly, an imprint corresponding to the shape of the part to be manufactured.

Abstract: A turboprop includes a rotary propeller upstream from an engine and an air intake that is not coaxial to the propeller, said air intake defining a conduit for supplying air to the engine and further defining a bypass to said conduit, the bypass having an outlet oriented substantially axially towards the downstream of the engine. The turboprop further includes a nacelle surrounding the engine and the air intake, wherein the air intake is secured to a housing of the engine and is not rigidly connected to the nacelle, so as to allow, during operation, relative movements between the air intake and the nacelle. The outlet is connected by a flexible link to an intake of an air circuit carried by the nacelle.

Abstract: The invention relates to a composite structuring panel (1) for the trailing edge of an aircraft element, having: an upper surface (3); a lower surface (5); an edge (7) connecting said upper (3) and lower (5) surfaces; the upper surface (3) and the lower surface (5) being connected by transverse stiffeners (9) and the structuring panel being made of a unitary part forming the upper surface (3), the lower surface (5), the edge (7), and the transverse stiffeners (9). The invention also relates to a method for manufacturing such a panel (1), and to an aircraft element comprising such a panel (1).

Abstract: The invention relates to a composite sandwich panel (1) comprising at least a web (2) provided between an inner skin (3) and an outer skin (4), each skin being made of a composite material from a plurality of fibrous plies (5, 6, 7, 8), characterized in that a portion of the plies of the inner skin and/or of the outer skin each extends at least partially through the web at an uninterrupted area thereof in order to form together a through-reinforcement (9) extending from the inner skin to the outer skin of the composite sandwich panel. The invention further relates to an aircraft element including such a composite sandwich panel.

Abstract: A method for manufacturing a part in composite material with a hollow core is provided that includes applying at least one adhesive layer on an open surface of the hollow core. The adhesive layer is a blocking polymerizable adhesive layer having, after polymerization, sealing properties relatively to a resin and capable of preventing its diffusion towards an inside of the hollow core. Subsequently, the blocking adhesive layer is polymerized so as to achieve sealing of the hollow core. Additionally, various parts manufactured according to the methods of the present disclosure are also provided.

Abstract: A device for manufacturing a part in composite material by resin transfer moulding and method is provided according to the present disclosure. The device includes a manufacturing mould in which at least one preform is intended to be arranged of the part intended to be impregnated with resin during a resin transfer process. The mould is subdivided into a die and at least one movable structural element intended to be indexed on the die to form an assembly, the assembly forming a cavity corresponding to a shape of the part to be manufactured. The mould is associated with at least one positioning frame having a low coefficient of thermal expansion in at least one dimension, the positioning frame being designed to allow the indexing and holding in position of the at least one movable structural element, corresponding to at least one singularity, on the die.

Abstract: The invention relates to a method for producing a structural composite trailing edge panel (1) for an aircraft element having: an upper surface (3), a lower surface (5), and a trailing edge (7) connecting said upper (3) and lower (5) surfaces, the upper surface (3) and the lower surface (5) being connected by transverse stiffeners (9), at least one longitudinal spar (10) being positioned in such a way that the directrix (?10) of each longitudinal spar (10) and the directrix (?9) of the transverse stiffeners (9) are not collinear and the structuring panel (1) being made up of a one-piece component forming the upper surface (3), the lower surface (5), the trailing edge (7), the transverse stiffeners (9) and the spar or spars (10).

Abstract: The invention relates to a molding device (1) for receiving fibers onto which a resin is injected, including a base (3), side plates (5) confining the base (3), and a cover (7) capable of engaging with a free end (9) of the side plates (5) in order to define an inner space (11) for receiving the fibers and the resin, wherein said device (1) comprises, in the inner space (11), a skin (21) that is configured to be complementary to the inner surface of the cover (7) and to at least a portion of the inner surface of the side plates (5). The invention also relates to a method for manufacturing a resin part using such a device (1).

Abstract: The invention relates to a composite structuring panel (1) for the trailing edge of an aircraft element, having: an upper surface (3); a lower surface (5); an edge (7) connecting said upper (3) and lower (5) surfaces; the upper surface (3) and the lower surface (5) being connected by transverse stiffeners (9) and the structuring panel being made of a unitary part forming the upper surface (3), the lower surface (5), the edge (7), and the transverse stiffeners (9). The invention also relates to a method for manufacturing such a panel (1), and to an aircraft element comprising such a panel (1).

Abstract: The invention relates to a composite sandwich panel (1) comprising at least a web (2) provided between an inner skin (3) and an outer skin (4), each skin being made of a composite material from a plurality of fibrous plies (5, 6, 7, 8), characterised in that a portion of the plies of the inner skin and/or of the outer skin each extends at least partially through the web at an uninterrupted area thereof in order to form together a through-reinforcement (9) extending from the inner skin to the outer skin of the composite sandwich panel. The invention further relates to an aircraft element including such a composite sandwich panel.