Abstract: Structures and methods for joining composite fuselage sections and other panel assemblies together are disclosed herein. In one embodiment, a shell structure configured in accordance with the present invention includes a first panel portion positioned adjacent to a second panel portion. The first panel portion can include a first stiffener attached to a first composite skin, and the second panel portion can include a second stiffener attached to a second composite skin. The shell structure can further include a fitting extending across a first edge region of the first panel portion and a second edge region of the second panel portion. A first end portion of the fitting can be attached to the first stiffener and the first composite skin, and a second end portion of the fitting can be attached to a second stiffener and a second composite skin, to join the first panel portion to the second panel portion.

Abstract: Structures and methods for joining composite fuselage sections and other panel assemblies together are disclosed herein. In one embodiment, a shell structure configured in accordance with the present invention includes a first panel portion positioned adjacent to a second panel portion. The first panel portion can include a first stiffener attached to a first composite skin, and the second panel portion can include a second stiffener attached to a second composite skin. The shell structure can further include a fitting extending across a first edge region of the first panel portion and a second edge region of the second panel portion. A first end portion of the fitting can be attached to the first stiffener and the first composite skin, and a second end portion of the fitting can be attached to a second stiffener and a second composite skin, to join the first panel portion to the second panel portion.

Abstract: Structures and methods for joining composite fuselage sections and other panel assemblies together are disclosed herein. In one embodiment, a shell structure configured in accordance with the present invention includes a first panel portion positioned adjacent to a second panel portion. The first panel portion can include a first stiffener attached to a first composite skin, and the second panel portion can include a second stiffener attached to a second composite skin. The shell structure can further include a fitting extending across a first edge region of the first panel portion and a second edge region of the second panel portion. A first end portion of the fitting can be attached to the first stiffener and the first composite skin, and a second end portion of the fitting can be attached to a second stiffener and a second composite skin, to join the first panel portion to the second panel portion.

Abstract: Structures and methods for joining composite fuselage sections and other panel assemblies together are disclosed herein. In one embodiment, a shell structure configured in accordance with the present invention includes a first panel portion positioned adjacent to a second panel portion. The first panel portion can include a first stiffener attached to a first composite skin, and the second panel portion can include a second stiffener attached to a second composite skin. The shell structure can further include a fitting extending across a first edge region of the first panel portion and a second edge region of the second panel portion. A first end portion of the fitting can be attached to the first stiffener and the first composite skin, and a second end portion of the fitting can be attached to a second stiffener and a second composite skin, to join the first panel portion to the second panel portion.

Abstract: Structures and methods for joining composite fuselage sections and other panel assemblies together. In one embodiment, a shell structure configured in accordance with the present invention includes a first panel portion positioned adjacent to a second panel portion. The first panel portion can include a first stiffener attached to a first composite skin, and the second panel portion can include a second stiffener attached to a second composite skin. The shell structure can further include a fitting extending across a first edge region of the first panel portion and a second edge region of the second panel portion. A first end portion of the fitting can be attached to the first stiffener and the first composite skin, and a second end portion of the fitting can be attached to a second stiffener and a second composite skin, to join the first panel portion to the second panel portion.

Abstract: Structures and methods for joining composite fuselage sections and other panel assemblies together are disclosed herein. In one embodiment, a shell structure configured in accordance with the present disclosure includes a first panel portion positioned adjacent to a second panel portion. The first panel portion can include a first stiffener attached to a first composite skin, and the second panel portion can include a second stiffener attached to a second composite skin. The shell structure can include a fitting extending across a first edge region of the first panel portion and a second edge region of the second panel portion. A first end portion of the fitting can be attached to the first stiffener and the first composite skin, and a second end portion of the fitting can be attached to a second stiffener and a second composite skin, to join the first panel portion to the second panel portion.

Abstract: Structures and methods for joining composite fuselage sections and other panel assemblies together are disclosed herein. In one embodiment, a shell structure configured in accordance with the present disclosure includes a first panel portion positioned adjacent to a second panel portion. The first panel portion can include a first stiffener attached to a first composite skin, and the second panel portion can include a second stiffener attached to a second composite skin. The shell structure can include a fitting extending across a first edge region of the first panel portion and a second edge region of the second panel portion. A first end portion of the fitting can be attached to the first stiffener and the first composite skin, and a second end portion of the fitting can be attached to a second stiffener and a second composite skin, to join the first panel portion to the second panel portion.

Abstract: A hybrid ice protection system (HIPS) including both an electro-expulsive de-icing system (EEDS) and electrothermal heaters to first break ice away from an outer surface of an engine nacelle inlet lip and then melt remaining residual ice from the outer surface of the inlet lip. The EEDS may have a plurality of EEDS actuators positioned to provide striking force to inner surfaces of both outer and inner walls of the inlet lip. The electrothermal heaters may be positioned to heat the inlet lip at areas between locations where the EEDS actuators provide striking force. The HIPS may also include a control system for actuating the EEDS actuators to strike the inner surface of the inlet lip when the inlet lip is at or below a predetermined temperature and then activating the electrothermal heaters to remove residual ice left on the inlet lip after actuation of the EEDS actuators.

Abstract: Structures and methods for joining composite fuselage sections and other panel assemblies together are disclosed herein. In one embodiment, a shell structure configured in accordance with the present invention includes a first panel portion positioned adjacent to a second panel portion. The first panel portion can include a first stiffener attached to a first composite skin, and the second panel portion can include a second stiffener attached to a second composite skin. The shell structure can further include a fitting extending across a first edge region of the first panel portion and a second edge region of the second panel portion. A first end portion of the fitting can be attached to the first stiffener and the first composite skin, and a second end portion of the fitting can be attached to a second stiffener and a second composite skin, to join the first panel portion to the second panel portion.

Abstract: Structures and methods for joining composite fuselage sections and other panel assemblies together are disclosed herein. In one embodiment, a shell structure configured in accordance with the present invention includes a first panel portion positioned adjacent to a second panel portion. The first panel portion can include a first stiffener attached to a first composite skin, and the second panel portion can include a second stiffener attached to a second composite skin. The shell structure can further include a fitting extending across a first edge region of the first panel portion and a second edge region of the second panel portion. A first end portion of the fitting can be attached to the first stiffener and the first composite skin, and a second end portion of the fitting can be attached to a second stiffener and a second composite skin, to join the first panel portion to the second panel portion.

Abstract: Structures and methods for joining composite fuselage sections and other panel assemblies together are disclosed herein. In one embodiment, a shell structure configured in accordance with the present invention includes a first panel portion positioned adjacent to a second panel portion. The first panel portion can include a first stiffener attached to a first composite skin, and the second panel portion can include a second stiffener attached to a second composite skin. The shell structure can further include a fitting extending across a first edge region of the first panel portion and a second edge region of the second panel portion. A first end portion of the fitting can be attached to the first stiffener and the first composite skin, and a second end portion of the fitting can be attached to a second stiffener and a second composite skin, to join the first panel portion to the second panel portion.