Hinged composite sandwich panels

A method comprising:

• bonding a seal to a first over-crush edge region of a composite sandwich panel, wherein a thickness of the composite sandwich panel decreases within the first over-crush edge region in a direction towards a first edge of the composite sandwich panel within the first over-crush edge region;
• connecting a hinge to the first over-crush edge region of the composite sandwich panel;
• applying heat and a compressive force to the composite sandwich panel to form the first over-crush edge region;
• allowing volatiles to escape through a fourth edge of the composite sandwich panel while forming a second over-crush edge region and a third over-crush edge region of the composite sandwich panel;
• and
• applying a decorative laminate over a first face and the first edge of the composite sandwich panel.

A method for rotating a composite panel within an aircraft, the method comprising:

• overlaying a first composite skin with a hollow cell core and a second composite skin;
• forming the composite panel by applying an over-crush pressure into the second composite skin and the hollow cell core from a first edge of the second composite skin into a border of a first over-crush edge region of the composite panel along an entire length of the first edge of the second composite skin while sandwiching the hollow cell core between the first composite skin and the second composite skin and reducing a thickness of a first edge of the first over-crush edge region to a lesser dimension than a thickness at the border of the first over-crush edge region of the composite panel comprising the hollow cell core sandwiched between the first composite skin and the second composite skin;
• bonding a seal to the second composite skin along the entire length of the first over-crush edge region and extending off the first over-crush edge region beyond the first edge of the first over-crush edge region;
• connecting a hinge to a structure in the aircraft and along a full length of the second composite skin such that an axis of rotation of the hinge is over the second composite skin in the first over-crush edge region;
• and
• applying the over-crush pressure into the second composite skin and the hollow cell core from a second edge of the second composite skin into a border of a second over-crush edge region along an entire length of the second edge of the second composite skin while sandwiching the hollow cell core between the first composite skin and the second composite skin and reducing a thickness of the second edge to a lesser dimension than a thickness at the border of the second over-crush edge region of the composite panel comprising the hollow cell core sandwiched between the first composite skin and the second composite skin.

A method for rotating a composite panel within an aircraft, the method comprising:

• overlaying a first composite skin with a hollow cell core and a second composite skin;
• forming the composite panel by applying an over-crush pressure into the second composite skin and the hollow cell core from a first edge of the second composite skin into a border of a first over-crush edge region of the composite panel along an entire length of the first edge of the second composite skin while sandwiching the hollow cell core between the first composite skin and the second composite skin and reducing a thickness of a first edge of the first over-crush edge region to a lesser dimension than a thickness at the border of the first over-crush edge region of the composite panel comprising the hollow cell core sandwiched between the first composite skin and the second composite skin;
• bonding a seal to the second composite skin along the entire length of the first over-crush edge region and extending off the first over-crush edge region beyond the first edge of the first over-crush edge region;
• connecting a hinge to a structure in the aircraft and along a full length of the second composite skin such that an axis of rotation of the hinge is over the second composite skin in the first over-crush edge region;
• and
• applying the over-crush pressure into the second composite skin and the hollow cell core from a third edge of the second composite skin into a border of a third over-crush edge region along an entire length of the third edge of the second composite skin while sandwiching the hollow cell core between the first composite skin and the second composite skin and reducing a thickness of the third edge to a lesser dimension than a thickness at the border of the third over-crush edge region of the composite panel comprising the hollow cell core sandwiched between the first composite skin and the second composite skin.