Abstract: A pressure bulkhead for an aircraft comprising a pressurized zone, the pressure bulkhead made of a composite material and comprising a front panel intended to be faced towards the pressurized zone; and a rear panel, substantially parallel to the front panel and intended to be located farther from the pressurized zone than the front panel. The pressure bulkhead further comprises a plurality of longerons located between the front panel and the rear panel, each one of them being attached to the front panel and to the rear panel; and a plurality of stiffeners, some of them being attached to the front panel and a remainder of them being attached to the rear panel.

Abstract: A fuselage rear end of an aircraft, comprising a structural part comprising a skin and longitudinal and transversal reinforcing members and a fairing. The structural part longitudinally spans over the whole rear end and comprises a first portion in which the transversal reinforcing members occupy the whole perimeter of the corresponding fuselage section and at least a second portion in which the transversal reinforcing members occupy only a portion of the perimeter of the corresponding fuselage section. The fairing is located below the second portion of the structural part.

Abstract: A frame assembly for a rear section of an aircraft, comprising at least one frame having a plane of symmetry, wherein the frame assembly also comprises at least one supporting element having two ends, wherein each one of the ends is attached to said at least one frame at different sides of the plane of symmetry. This arrangement permits a more effective transfer of loads coming from the vertical tail plane of the aircraft to the fuselage.

Abstract: A structure of a control surface for aircrafts, such as elevators, rudders, landing flaps, ailerons, and other similar lifting surfaces. The control surface includes a front spar, a rear spar, a plurality of ribs, at least one hinge fitting, and at least one actuator fitting, the hinge and actuator fittings joined with the front spar. A first rib is joined with a hinge fitting to form a first oblique angle with the front spar; a second rib is joined with an actuator fitting to form a second oblique angle with the front spar. The rear spar, located between the front spar and trailing edge of the control surface, extends only between the second rib and an outboard end of the control surface. The ribs are located in correspondence with the main torsion and bending load paths created in the control surface during flight, landing or takeoff.

Abstract: A fuselage rear end of an aircraft, comprising a structural part comprising a skin and longitudinal and transversal reinforcing members and a fairing. The structural part longitudinally spans over the whole rear end and comprises a first portion in which the transversal reinforcing members occupy the whole perimeter of the corresponding fuselage section and at least a second portion in which the transversal reinforcing members occupy only a portion of the perimeter of the corresponding fuselage section. The fairing is located below the second portion of the structural part.

Abstract: A method for manufacturing a monolithic torsion box inner structure of an aircraft lifting surface including: a) providing a set of laminated preforms of a composite material for forming said torsion box inner structure, each laminated preform being configured for constituting a part of one component of the torsion box inner structure; b) arranging the laminated preforms in a curing tooling and subjecting the assembly to an autoclave cycle to co-cure the said laminated preforms; c) demoulding the curing tooling in a vertical direction.

Abstract: A pressure bulkhead for an aircraft comprising a pressurized zone, the pressure bulkhead made of a composite material and comprising a front panel intended to be faced towards the pressurized zone; and a rear panel, substantially parallel to the front panel and intended to be located farther from the pressurized zone than the front panel. The pressure bulkhead further comprises a plurality of longerons located between the front panel and the rear panel, each one of them being attached to the front panel and to the rear panel; and a plurality of stiffeners, some of them being attached to the front panel and a remainder of them being attached to the rear panel.

Abstract: The invention refers to a composite structure for an aircraft comprising a panel and at least one stringer joined to a surface of the skin panel, the stringer having a foot and a web protruding from the foot, and a run-out section at one of its ends. The panel includes a plurality of stacked and co-cured plies of composite material, and at least part of the foot of the stringer at the run-out section, is inserted between two plies of the panel, and it is co-bonded, co-cured or secondary bonded with said two plies. The structure of the invention provides a reinforced join between the foot of the stringer and the skin panel, avoiding peeling or de-bonding problems, without substantially modifying the manufacturing process.

Abstract: A method for manufacturing a leading edge profile section of an aircraft lifting surface is provided. The method comprises the following steps: a) providing a set of laminated preforms of a composite material configured with a suitable shape for constituting the leading edge profile section; b) arranging said laminated preforms in a curing tooling and subjecting the assembly to an autoclave cycle to co-cure said laminated preforms; c) demolding the curing tooling in a spanwise direction towards the aircraft symmetry plane. The invention also comprises a leading edge profile section manufactured by said method comprising in addition to the skin of the leading edge profile section, one or more of the following structural elements: an auxiliary spar, a longitudinal stiffener reinforcing an auxiliary spar, a longitudinal stringer reinforcing the skin of the leading edge profile.

Abstract: A frame of a fuselage section of an aircraft that receives external loads, such as the rear section that receives the loads of the tail planes, comprising a first sector configured specifically to withstand these external loads and a second sector along the remaining perimeter of the fuselage, said first sector comprising the following structural elements a) a foot piece to be joined to the fuselage skin; b) a head piece; c) a web comprising one or more fitting-pieces for receiving said external loads and a plurality of X-shaped pieces joined by its ends to said foot piece and said head piece; and d) two connecting pieces of the ends of said foot piece and said head piece with the second sector of the frame.

Abstract: The invention refers to a composite structure for an aircraft comprising a panel and at least one stringer joined to a surface of the skin panel, the stringer having a foot and a web protruding from the foot, and a run-out section at one of its ends. The panel includes a plurality of stacked and co-cured plies of composite material, and at least part of the foot of the stringer at the run-out section, is inserted between two plies of the panel, and it is co-bonded, co-cured or secondary bonded with said two plies. The structure of the invention provides a reinforced join between the foot of the stringer and the skin panel, avoiding peeling or de-bonding problems, without substantially modifying the manufacturing process.

Abstract: An optimized structure of a control surface for aircrafts, such as elevators, rudders, landing flaps, ailerons, and other similar lifting surfaces. The control surface includes a front spar, a rear spar, a plurality of ribs and at least one hinge fitting, as well as and an actuator fitting joined with the front spar. At least one rib is connected with the actuator fitting and it is arranged so as to define an oblique angle with the front spar. One or more ribs are located in correspondence with the main torsion and bending load paths created in the control surface during flight, landing or takeoff, so the same structural behavior of the control surface is achieved, but, with a reduced number of ribs, so that the weight of the control surface is reduced.

Abstract: Horizontal tail plane of an aircraft which comprises a single torsion box which also comprises a front spar (1), a rear spar (2), an upper cover and a lower cover, ribs and two closing ribs (3), the front spar (1), the rear spar (2) and the upper and lower covers extending between said two closing ribs (3) and along the whole span of the horizontal tail plane in a symmetric layout with respect to the symmetry axis of the aircraft.

Abstract: A reinforcement structure for an opening in the primary structure of an aircraft, the structure comprising a skin, frame members which are transverse with respect to the flying direction of the aircraft, and stringers which are longitudinal with respect to the flying direction of the aircraft. The reinforcing structure includes a perimetral reinforcing element situated along the edge of the opening and reproducing the geometrical form thereof; at least one pair of transverse reinforcing elements arranged on both transverse sides of the opening; and at least one pair of longitudinal reinforcing elements arranged on both longitudinal sides of the opening.

Abstract: A method for manufacturing a leading edge profile section of an aircraft lifting surface is provided. The method comprises the following steps: a) providing a set of laminated preforms of a composite material configured with a suitable shape for constituting the leading edge profile section; b) arranging said laminated preforms in a curing tooling and subjecting the assembly to an autoclave cycle to co-cure said laminated preforms; c) demoulding the curing tooling in a spanwise direction towards the aircraft symmetry plane. The invention also comprises a leading edge profile section manufactured by said method comprising in addition to the skin of the leading edge profile section, one or more of the following structural elements: an auxiliary spar, a longitudinal stiffener reinforcing an auxiliary spar, a longitudinal stringer reinforcing the skin of the leading edge profile.

Abstract: The invention provides a method for manufacturing a monolithic torsion box inner structure (40) of an aircraft lifting surface comprising the following steps: a) providing a set of laminated preforms of a composite material for forming said torsion box inner structure (40), each laminated preform being configured for constituting a part of one component of the torsion box inner structure (40); b) arranging said laminated preforms in a curing tooling and subjecting the assembly to an autoclave cycle to co-cure said laminated preforms; c) demoulding the curing tooling in a vertical direction. The invention also provides an aircraft lifting surface wherein: the inner structure (40) of each torsion box (13) is manufactured by said method and the upper and lower skins (31, 33) are attached to the torsion box inner structure (40) with mechanical attachment means.

Abstract: Aircraft fuselage frame in composite material with stabilising ribs which, in at least a first sector (31), comprises a frame (9) with an omega-shaped cross-section (11) formed by a top element (23), two webs (25, 25?), two feet (27, 27?) which includes at least one internal rib (41) formed by a top element (43) joined to the top element (23) of the frame (9), a web (45), two flanges (49, 49?) joined to the webs (25, 25?) of the frame (9) and a foot (47) aligned with the feet (27, 27?) of the frame (9). The invention also relates to a manufacturing procedure for the first sector (31) of the frame with stabilised web (10) comprising steps for: a) providing a frame (9) and at least one internal rib (41) with the configurations indicated; b) joining at least said internal rib (41) to the frame (9).

Abstract: Load introduction zone in the rear end of an aircraft comprising receiving elements (3, 4) of the loads of the horizontal and vertical tail stabilizers connected to structural elements (2, 8, 9) wherein: a) said structural elements are the cladding (8), two contiguous load-bearing frames (2) with a ? or omega cross-section and, at least, a third frame (9); b) the receiving element of the horizontal tail stabilizer load is a first fitting (3) structured as a torque box, transversely arranged between said load-bearing frames (2) and connected to them, and the receiving elements of the vertical tail stabilizer loads are some second fittings (4) coupled to these frames (2, 9); c) said frames (2, 9) and said first and second fittings (3, 4) are entirely made of composite material.

Abstract: Torsion box of an aircraft stabilizing surface comprising a front spar (15), a rear spar (17), ribs (21) and upper and lower stiffened skins, in which at least one of the skins (31) is stiffened with a plurality of stringers (33, 43), preferably omega-shaped stringers, extending all of them along the full skin span, preferably in a convergent distribution, and having a decreasing cross section towards the skin outer edge.

Abstract: The present invention relates to a curved rear pressure bulkhead (1) made of a composite material, comprising three layers: an outer layer (2) and an inner layer (3) symmetrical with one another, which are made of a fiber laminate, and the intermediate layer or core (4) made of a lightweight material. With such a sandwich-type shape, the resulting bulkhead (1) is able to withstand the bending loads coming from the strain and pressurization of the fuselage without having to increase its stiffness by means of using some type of stiffener. The bulkhead (1) comprises a ring (6) that is attached to the ends of the sandwich shape. This ring (6) is split into several parts (8) to facilitate the assembly of the bulkhead (1) made of composite material to the fuselage skin (5).