A manufacturing method of a control surface of an aircraft, the control surface including an upper skin, a lower skin, ribs joining the upper skin and the lower skin and located along a chordwise direction of the control surface. The manufacturing method includes the steps of providing a single composite preform comprising the upper skin, the lower skin and the ribs, and curing the single composite preform such that an integrated box comprising the upper skin, the lower skin and the ribs is formed.

A vehicle, including a high voltage energy consumer; an electrical energy source; an electrical energy distribution network having an electrical energy distribution element configured to carry electrical power from the electrical energy source to the high voltage energy consumer; and a cooling system for cooling the high voltage energy consumer, the cooling system having: coolant, and a coolant distribution network having a coolant distribution element for delivering the coolant to the high voltage energy consumer to remove heat generated by the high voltage energy consumer; wherein the coolant distribution element provides mechanical support to the electrical energy distribution element, and the coolant within in the coolant distribution element removes heat generated in the electrical energy distribution element as it carries electrical power from the electrical energy source to the high voltage energy consumer

In order to integrate an aerodynamic box without a solid shim, a spar joint is provided. The spar joint comprises a first spar having a first spar contact surface and a second spar having a second spar contact surface. When the first and second spars are fixed to each other, the first and second contact surfaces engage each other in a manner forming a contact plane. The first and second contact surfaces are inclined such that a first imaginary plane that is an extension of the contact plane in a chordwise direction forms at least one intersection with a second imaginary plane that is originating from a chordwise extremal end of an aerodynamic skin. The spar joint is used in aerodynamic boxes, such as flap boxes and TE boxes. The spar joint is also used in airfoil profiles.

In order to integrate an aerodynamic box without a solid shim, a spar joint is provided. The spar joint comprises a first spar having a first spar contact surface and a second spar having a second spar contact surface. When the first and second spars are fixed to each other, the first and second contact surfaces engage each other in a manner forming a contact plane. The first and second contact surfaces are inclined such that a first imaginary plane that is an extension of the contact plane in a chordwise direction forms at least one intersection with a second imaginary plane that is originating from a chordwise extremal end of an aerodynamic skin. The spar joint is used in aerodynamic boxes, such as flap boxes and TE boxes. The spar joint is also used in airfoil profiles.

An apparatus is provided for analysis of a leading edge rib of a fixed leading edge section of an aircraft wing. The apparatus may identify geometric or inertial properties of a plurality of stiffeners of the rib in which respective stiffeners are represented by a collection of geometry within a solid model of the rib, and perform an analysis to predict a failure rate of the leading edge rib under an external load. From the failure rate, the apparatus may determine a structural integrity of the leading edge rib under the external load. Identifying the properties may include, extracting a section cut of the geometry that corresponds to and has one or more properties of the respective stiffener, and identifying the properties of the cross-section and thereby the respective stiffener based on a correlation of the cross-section to a generic profile of a plurality of different cross-sections.

An apparatus is provided for analysis of a leading edge rib of a fixed leading edge section of an aircraft wing. The apparatus may identify geometric or inertial properties of a plurality of stiffeners of the rib, and based thereon perform an analysis to predict a failure rate of the leading edge rib under an external load. From the failure rate, the apparatus may determine a structural integrity of the leading edge rib under the external load. Performing the analysis may include importing a plurality of section cuts into a finite element model of the rib and thereby identifying nodes proximate the section cuts. Under an external load, internal load distributions may be extracted from elements proximate the nodes and elements, and the failure rate of the leading edge rib under the external load may be predicted based on the internal load distributions of the elements.

A leading edge structure for a flow control system of an aircraft is disclosed having a leading edge panel that surrounds a plenum, wherein the leading edge panel has a first side portion, a second side portion opposite the first side portion, an inner surface facing the plenum and an outer surface in contact with an ambient flow, and wherein the leading edge panel comprises a plurality of micro pores forming a fluid connection between the plenum and the ambient flow, wherein the plenum is connected to an air outlet arrangement configured for causing an underpressure in the plenum, so that air from the ambient flow is drawn through the micro pores into the plenum and from there discharged through the air outlet arrangement into the ambient flow.

A trailing edge for a composite lifting surface is disclosed having a spar, an upper panel and a lower panel each having a free edge, a seal, and an elongated profile located on the panels following a spanwise direction of the trailing edge. An elongated profile including a web extending along the spanwise direction of the trailing edge, and having first and second flange portions, and a transition zone in between the first and second flange portions extending at different heights with respect to each other. The first flange portion is configured to hold at least a part of the seal underneath, and the second flange portion is configured to contact the panel so that the first flange portion secures the seal to the panel, and the second flange portion secures the elongated profile to the panel.

A trailing edge for a composite lifting surface is disclosed having a spar, an upper panel and a lower panel each having a free edge, a seal, and an elongated profile located on the panels following a spanwise direction of the trailing edge. An elongated profile including a web extending along the spanwise direction of the trailing edge, and having first and second flange portions, and a transition zone in between the first and second flange portions extending at different heights with respect to each other. The first flange portion is configured to hold at least a part of the seal underneath, and the second flange portion is configured to contact the panel so that the first flange portion secures the seal to the panel, and the second flange portion secures the elongated profile to the panel.

The present disclosure relates to a structural component of an aircraft wing body and an aircraft including the structural component. According to an aspect of the present disclosure, a structural component of an aircraft wing body is provided. The structural component includes a body part and a profile. The body part includes an edge portion formed by end portions of a first skin and a second skin of the body part superposed together. The profile is attached to the edge portion. The profile has an outer profile conforming to an outer profile of the body part such that the structural component, as a whole, exhibits an aerodynamic outer profile after the profile is attached to the edge portion. The profile is attached to the edge portion via a plurality of separate intermediate members or by contacting the edge portion.