The Boeing Company

Willard N Westre

David W Evans

Max Urban Kismarton

Marc Joel Piehl

Mark E Rogalski

Samuel Eric Cregger

Donald J Ayers

Edward Li

Donald L Grande

Stuart M Buchan

Heather C Allen-Lilly

Eric Sager

Daniel J Hoffman

Robert J Rothschilds

Jeffery Lee Marcoe

Kevin Matthew Retz

Stephen T Brown

Paul Stanley Nordman

Charles W Newquist

Antonio Carlos Rufin

William G Roeseler

Harry W Dursch

Sven E Axter

William B Grace

David J Carbery

Roy Ikegami

Thomas D Martin

Robert A Burns

Leanna M Micona

Terry K Isomura

Earl E Van Beek

Jefferson F Newton

William G Reeseler

A Lew Wingert

Gary D Tuss

Moe K Soleiman

Judy S Chen

There is provided a box structure for carrying load having upper and lower composite integrated sandwich panels. The panels have facesheets sandwiching one or more core portions and adjacent dense packs oriented in an axial direction. The box structure further has a plurality of spars. Each spar has a web and web attachments and has a spar length in the axial direction. The plurality of spars are connected to the panels with the web attachments located at the dense packs. The facesheets are configured to carry primarily torsion and pressure loads in shear and no significant axial loads. The dense packs are configured to carry all significant box bending in axial tension and compression loads.

Box structures for carrying loads and methods of making the same

The invention relates to aircraft floor assemblies and to methods for their assembly. In one embodiment, an aircraft comprises at least one fuselage section and at least one floor section. At least one system component is installed to the floor section prior to installation of the floor section in the fuselage of the aircraft. In another embodiment, a floor section adapted to be installed in an aircraft includes at least one system component installed to the floor section while it is outside of the aircraft. In yet another embodiment, a method is disclosed for assembling an aircraft. The method comprises installing at least one system component to a floor section, and installing the floor section into a fuselage section of the aircraft.

Aircraft floor and method of assembly

An integrated floor for an aircraft fuselage includes a composite panel forming a floor surface and composite beams bonded to the floor panel. The beams extend longitudinally within the fuselage and support the floor.

Integrated aircraft floor with longitudinal beams

The invention relates to airplane floor assemblies and to methods for their assembly. In one embodiment, an airplane comprises at least one fuselage section and at least one floor section. At least one system component is installed to the floor section prior to installation of the floor section in the fuselage of the airplane. In another embodiment, a floor section adapted to be installed in an airplane includes at least one system component installed to the floor section while it is outside of the airplane. In yet another embodiment, a method is disclosed for assembling an airplane. The method comprises installing at least one system component to a floor section, and installing the floor section into a fuselage section of the airplane.

Airplane floor assembly

Embodiments of integral composite panels and joints for composite structures are described In one implementation, an integrated panel spanning substantially the entire wingspan of an aircraft, includes at least a center portion and a pair of outwardly projecting wing portions. The portions may include a skin formed from successive layers or plies of composite material which overlap and offset at the joint between respective sections creating a pad-up area to carry loads between the portions. In a particular implementation, the skin is laid over one or more structural stringers which are transitioned into the joints between sections such as by tapering of the thickness and/or stiffness of the stringer.

Composite wing-body joint

Laminate structures and methods for forming same are disclosed. In one embodiment, a laminate structure includes a metal-polymer composite lamina. The metal-polymer composite lamina has a first face and a second face spaced apart, and extends to a terminal edge. The lamina includes a ply of fiber-reinforced polymer that extends between the first face and the second face and has an interior edge. The interior edge defines at least one cutout. A ply of metal foil extends between the first face and the second face substantially from the interior edge filling the at least one cutout.

Methods of forming metal foil ply replacement in composite structures

A method is provided for altering residual stresses in a fiber-metal-laminate structure. At least a first fiber laminate substrate layer having first and second surfaces is provided. At least a first outer metallic laminate layer having third and fourth surfaces is provided, and the first and fourth surface is provided, and the first and fourth surfaces are bonded together. The third surface is peened, such that the at least first fiber laminate substrate layer is subjected to residual tensile stresses and the at least first outer metallic laminate layer is subjected to residual compression stresses. Peening defines a peened region that extends from the third surface into the at least first outer metallic laminate layer a finite distance that is closer to the third surface than the fourth surface.

Method for improving crack resistance in fiber-metal-laminate structures

Titanium foil ply replacement in layup of composite skin

The invention provides a hybrid laminate and skin panels of hybrid laminate structure that are suitable for a supersonic civilian aircraft. The hybrid laminates include layups of layers of titanium alloy foil and composite plies, that are optimally oriented to counteract forces encountered in use, that are bonded to a central core structure, such as titanium alloy honeycomb. The reinforcing fibers of the composite plies are selected from carbon and boron, and the fibers are continuous and parallel oriented within each ply. However, some plies may be oriented at angles to other plies. Nevertheless, in a preferred embodiment of the invention, a substantial majority of, or all of, the fibers of the hybrid laminates are oriented in a common direction. The outer surfaces of the laminates include a layer of titanium foil to protect the underlying composite-containing structure from the environment, and attack by solvents, and the like.

Titanium-polymer hybrid laminates

Disclosed is a lightweight honeycomb structure that comprises a top panel (34) and bottom panel (32) and a large-celled honeycomb layer (36) disposed between the top and bottom panels. The honeycomb panel structure according to the present invention can be used in such applications as a sound-absorbing payload shroud for launch vehicles, flooring and walls in aerospace applications, aircraft wings and in structures that will be exposed to high heat. Additionally, a method of making a honeycomb structure of thermoplastic material is disclosed.

Lightweight honeycomb panel structure

An apparatus for attaching a sandwich panel to another structure. A hollow pin passes through the panel near the edge to be joined to the structure. A bolt, passing through the edge of the panel, and barrel nut, inside the hollow of the pin, are used to draw the edge of the panel against a flange attached to the structure. The panel may be reinforced with doublers or an end cap.

Apparatus to attach a sandwich panel

A high-temperature, low-thermal expansion fastener for joining low thermal expansion refractory materials includes a pin with an inset portion on the shank of the pin that forms a head-facing shoulder, at least one shell insert which is seatable about the inset portion, and an internally threaded member which can be screwed onto the threaded outer surface of the shell inserts to maintain the inserts in abutting relationship with the inset portion. Forces within the fastener are transferred from the internally threaded member to the shell inserts which in turn press against the head-facing shoulder on the pin shank to place the shank in tension. The fastener parts, including a spacer, may be selected to compensate for thermal expansion or contraction of the fastened elements.

High temperature low thermal expansion fastener

A seal for preventing extremely hot gases from passing between a movable engine ramp and an engine splitter wall in a hypersonic aircraft. The seal moves with the engine ramp, in a direction parallel to the surface of the splitter wall. The seal includes a first flange which makes sliding contact with the outer engine ramp surface and is attached to a flexible sealing portion that fills the gap between the ramp and wall. The sealing portion maintains sliding contact with the wall as the gap changes dimension due to variations in the ramp's temperature. The attachment region can be a concave surface to radiate the seal's heat. In a second embodiment, the seal includes a second flange through which the seal is attached to the cooler side of the engine ramp.

Slidable high-temperature seal

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Bellevue, WA, United States of America

Willard N Westre