Structural composite panel for an aircraft including a protection against high energy impacts
A structural panel consisting of a stratified composite material includes one face exposed to impacts and further includes a layer consisting of a hyper-elastic material bonded adhesively to its other face. According to this embodiment, debris striking the exposed face of this composite panel will have some of its energy dissipated by the local rupture of the composite skin, while the rest of the energy is absorbed by the deformation of the layer of hyper-elastic material that captures the debris and expels it again.
1. A structural panel, comprising:
a 2-4 mm thick composite skin with fiber reinforcement formed from continuous fibers, the composite skin is a carbon-epoxy resin composite with a continuous fiber volume ratio of greater or equal to 50%, one surface of said skin being exposed to impacts;
a layer formed from a hyper-elastic material formed from a chloroprene elastomer with a thickness equal to or less than the thickness of said composite skin and bonded adhesively with epoxy glue to a first surface of the composite skin, the hyper-elastic material having elongation capabilities of about 500%; and
stiffeners where the layer of hyper-elastic material is bonded to the skin between said stiffeners,
wherein the structural panel is resistant to operational static and fatigue stresses imposed on a second surface of the composite skin opposite the first surface, and debris striking the second surface with an incident energy of 1000 to 4000 joules will have energy dissipated by local rupture of the composite skin and no peeling off of the layer of hyper-elastic material by glue rupturing along the composite skin, and
the panel is configured such that some energy of the debris is dissipated by local rupture of the composite skin, and energy absorbed by deformation of the layer of hyper-elastic material expels the debris.
2. A method for manufacturing a structural element exposed to impacts of an aircraft fuselage according to claim 1 , comprising:
manufacturing skin panels made of composite material;
assembling said panels to form the fuselage element; and
bonding a layer of hyper-elastic material on inner surfaces of panels exposed to impacts after assembly.
3. The method according to claim 2 , further comprising:
bonding stiffeners by co-firing or by gluing to at least one of the skin panels before assembly and that the layer of hyper-elastic material is bonded between the stiffeners after the panels are assembled.
4. An aircraft fuselage comprising a panel, said panel comprising:
a 2-4 mm thick composite skin with fiber reinforcement formed from continuous fibers, the composite skin is a carbon-epoxy resin composite with a continuous fiber volume ratio of greater or equal to 50%, one surface of said skin being exposed to impacts;
a layer formed from a hyper-elastic material formed from a chloroprene elastomer with a thickness equal to or less than the thickness of said composite skin and bonded adhesively with epoxy glue to a first surface of the composite skin, the hyper-elastic material having elongation capabilities of about 500%; and
stiffeners where the layer of hyper-elastic material is bonded to the skin between said stiffeners,
wherein the structural panel is resistant to operational static and fatigue stresses imposed on a second surface of the composite skin opposite the first surface, and debris striking the second surface with an incident energy of 1000 to 4000 joules will have energy dissipated by local rupture of the composite skin and no peeling off of the layer of hyper-elastic material by glue rupturing along the composite skin, and
the panel is configured such that some energy of the debris is dissipated by local rupture of the composite skin, and energy absorbed by deformation of the layer of hyper-elastic material expels the debris.