WO2002028709A1

WO2002028709A1 - Composite skin panel opening edge and method for manufacture

Info

Publication number: WO2002028709A1
Application number: PCT/US2001/030299
Authority: WO
Inventors: Timothy Myron Hazen; John Verdi Howard; Josef Antonin Fila; Steven Joel Addison
Original assignee: Bell Helicopter Textron Inc.
Priority date: 2000-10-02
Filing date: 2001-09-27
Publication date: 2002-04-11
Also published as: CA2423665A1

Abstract

The present invention disclosed herein provides a component frame (18) for supporting a component (28) in a vehicle. The component (28) may be a door, a window or another piece of equipment on an aircraft that requires a flush-fitting interface with the aircraft. The component frame (18) comprises an outer skin (20) having a recessed edge (16) disposed generally around the perimeter of the component frame (18). An inner skin (24) is disposed within the aircraft and serves to sandwich a honeycomb core (22) between itself and the outer skin (20).

Description

COMPOSITE SKIN PANEL OPENING EDGE AND METHOD FOR MANUFACTURE

TECHNICAL FIELD OF THE INVENTION This invention relates in general to the field of aircraft manufacturing and, in particular to, a method for manufacturing an edge closure for door and window openings in composite skin panels.

BACKGROUND OF THE INVENTION Without limiting the scope of the invention, its background will be described with reference to door openings in composite skin helicopter panels as an example.

Helicopter door openings have remained essentially unchanged since the 1960 's. The openings have not changed because helicopter bodies are manufactured using many of the same techniques that were developed during that time. Typical helicopter manufacturing techniques require many shaped and fitted panels to be riveted to an airframe. One disadvantage to this method is that riveting individual panels is a costly, labor intensive process. Another disadvantage is that the aesthetics of the aircraft are degraded because the helicopter has the appearance of a patchwork piece of riveted panels. Exposed rivet heads and seams at panel joints also reduce aerodynamic efficiency of the helicopter. Door openings are fabricated by a process similar to the typical helicopter body manufacturing process. A hole in the aircraft skin is fitted with several flange pieces that form a sealing surface for doors and windows. These pieces are usually riveted to the helicopter skin or frame. Riveting door frame pieces creates the same problems of costs and aesthetics as riveting skin panels. The multi-piece door frame, however, creates an additional fitting problem.

Because several door frame pieces are riveted to the helicopter, slight variances in the frame assembly process create door fitting problems. As a result, doors and openings must be hand fitted together. Hand fitting further increases labor costs because doors must be individually crafted to fit a particular opening. The doors, consequently, often are not interchangeable among helicopters of the same model, which increases future maintenance costs if a door must be replaced. The fitting process is also not consistent among assembly line workers. Doors that have been hand-fitted to their frames occasionally have environmental leaks and operational problems. Helicopter owners and passengers perceive poor quality and possibly unsafe conditions if they encounter a door that is difficult to operate or has air and water leaks.

Other factors also create problems if old door frame technology is incorporated into modern helicopter manufacturing. New manufacturing materials and techniques have been combined to produce helicopters that are stronger and lighter than helicopters manufactured using older technology. Helicopters made with composite materials, for example, have a smooth appearance because a solid piece of material may be formed into required complex shapes without using rivets or other mechanical fasteners. Riveting a multi-piece door frame to a sleek airframe negatively affects the aesthetics of the helicopter.

Therefore, a need has arisen for an edge closure for door and window openings in structural composite skin panels. A need has also arisen for such an edge closure that is not susceptible to environmental leaks because of poorly fitting doors and windows. A need has also arisen for an edge closure for door and window openings that reduces installation labor costs. Further, a need has arisen for such an edge closure for door and window openings that is compatible with current airframe manufacturing techniques.

SUMMARY OF THE INVENTION The present invention disclosed herein provides a panel in a vehicle such as an aircraft, a boat, an automobile, or a railcar, that has an outer skin. The outer skin has a component opening and a recessed edge that is disposed around the component opening. The panel also has an inner skin that corresponds to the outer skin. A core disposed between the outer skin and the inner skin provides structural rigidity to the panel.

In one embodiment of the present invention, the outer skin is a composite material that may be pre-cured before assembling the panel. The outer skin may be made from resin impregnated glass, carbon, aramid woven fiber cloth or other suitable materials. The outer skin may also be comprised of several pieces of fiber cloth that are oriented and layered to best support flight and landing loads. The outer skin forms a solid, single structure after it is cured. In another embodiment of the invention, the core may be a structural honeycomb material. The honeycomb material may be made from a metallic or non-metallic material and have cells that are a hexagonal or rectangular shape. The honeycomb material may, alternatively, have multiple other shapes to facilitate bending to achieve complex contours while continuing to provide structural integrity. Materials and shapes of

, the honeycomb material may be chosen to create a vehicle that has certain desirable structural properties. In yet another embodiment, the core is bonded between the outer skin and the inner skin. The inner skin has a component opening that corresponds to the outer skin. Although the outer skin may be pre-cured, the core and the inner skin may be assembled in an uncured state. The entire assembly may then be cured and bonded together in an autoclave or by other curing methods . In another embodiment, space between an edge of the core and the outer skin may be filled with an expanding adhesive. The expanding adhesive may fill any gaps between cut cells of the core and the outer skin. After the expanding adhesive has cured, the adhesive forms a hardened structure that bonds the core to the outer skin and inner skin. The cured adhesive also may support torsional and compressive loads on the edge of the component opening.

In yet another embodiment, an airframe section has a panel and a frame integral with the panel. A component is supported by the panel and fits within the frame and substantially flush with the outer surface of the airframe section.

BRIEF DESCRIPTION OF THE DRAWINGS For a more complete understanding of the present invention, including its features and advantages, reference is now made to the detailed description of the invention taken in conjunction with the accompanying drawings in which like numerals identify like parts and in which: Figure 1 is a perspective view of a section of a helicopter airframe incorporating certain embodiments of the invention; Figure 2 is a cross-section of a composite skin panel incorporating certain embodiments of the invention;

Figure 3 is a cross-section of a door opening and a door assembly incorporating certain embodiments of the invention; and Figure 4 is a cross-section of a door opening and a door assembly incorporating certain embodiments of the invention.

DETAILED DESCRIPTION OF THE INVENTION

While the making and using of various embodiments of the present invention are discussed in detail below, it should be appreciated that the present invention provides many applicable inventive concepts which can be embodied in a wide variety of specific contexts. The specific embodiments discussed herein are merely illustrative of specific ways to make and use the invention and do not delimit the scope of the invention.

Referring now to Figure 1, therein is depicted a perspective view of a section of a helicopter airframe section 10. Airframe section 10 has several openings 12 in a panel 14 for windows, doors, baggage compartments, steps, fuel funnels and other necessary components. Component openings 12 have recessed edges 16 that allow the components to fit substantially flush with panel 14. Flush-fitting components maintain aerodynamic surfaces of airframe section 10 and aid fuel efficiency. Flush- fitting components also contribute to design aesthetics that create an image of quality to owners, users and passengers .

Panel 14 may have a frame 18 that imparts additional structural rigidity to airframe section 10. Frame 18 may be formed from a composite material or frame 18 may be a metal, such as aluminum or titanium. Frame 18 may be bonded to the interior of airframe section 10 so that no mechanical fasteners penetrate panel 14. This method of attachment retains the smooth, solid appearance of the exterior of airframe section 10. A cross section of panel 14 is depicted in Figure 2. Panel 14 may be made from a composite material such as fiberglass or carbon fiber, for example, which may be comprised of the three separate components bonded together to form airframe section 10. Components of panel 14 may be oriented and designed to support all of the loads generated by or imposed on the helicopter. The construction of panel 14 may comprise an outer skin 20, which may be molded from several pieces of composite material to form airframe section 10 as a single, continuous piece. Outer skin 20 may be resin- impregnated, woven graphite cloth, which is pre-cured before assembly of panel 14. Panel 14 may also be made from an aramid such as the aramid fiber cloth sold under the trade name KEVLAR®. Other composite materials suitable for making panel 14 will be apparent to those with ordinary skill in the art. In some areas, airframe section 10 may have multiple layers of composite material that exhibit specific structural properties if oriented in a particular direction. Areas that are subjected to greater flight or landing loads may be buttressed or reinforced using extra layers of material to sustain those loads. Some of those loads are directional. Material layers, consequently, may be oriented to best support directional loads. For example, component openings 12 in panel 14 may create weak points in airframe section 10 because flight and landing loads must be transferred around component openings 12. Transferring loads around component openings 12 may create high stresses in the structures adjacent to component openings 12. Recessed edges 16 of component openings 12, consequently, may experience the same high loads as other areas of airframe section 10. In prior door component openings, flight and landing loads deformed rivets and flange pieces and caused door fit problems. The unique construction of panel 14 and recessed edge 16 creates component opening 12 that retains its shape and provides a consistent sealing surface. Recessed edge 16 may be reinforced by composite tape, which is designed and oriented to support loads in a particular direction.

A honeycomb core 22 may be bonded between outer skin 20 and an inner skin 24. Honeycomb core 22 may be carbon-based, metallic, or a more exotic material, all of which may be suitable to impart structural rigidity and strength to panel 14. Honeycomb core 22 is typically comprised of a material having multiple cells that reduce total weight and add strength in particular directions. Sometimes, these cells have a hexagonal cross-section. Honeycomb core 22 may, alternatively, have multiple other shapes to facilitate bending to achieve complex contours while continuing to provide structural integrity to panel 14. Materials and shapes of honeycomb core 22 may be chosen to enhance desirable structural properties in airframe section 10.

Cutting honeycomb core 22 to fit within outer skin 20 leaves jagged edges on honeycomb core 22 because cells will usually not be perfectly severed. These jagged edges on honeycomb core 22 result in air pockets between honeycomb core 22 and outer skin 20. These air pockets reduce the strength of recessed edge 16. Inner skin 24 may be uncured before outer skin 20, honeycomb core 22 and inner skin 24 are bonded together.

The assembly may be placed in an autoclave for final curing and bonding. Airframe section 10 emerges from the autoclave process as one continuous piece. Of course, other methods of curing the composite material such as infrared curing, for example, may be used to produce airframe section 10. Recessed edge 16 may be reinforced by using an expanding adhesive 26 to fill the space between the cut end of honeycomb core 22 and recessed edge 16. Air pockets are present between outer skin 20 and honeycomb core 22 because cutting honeycomb core 22 to fit component openings 12 does not result in a perfectly matched edge joint. As expanding adhesive 26 cures, it bonds the cut edge of honeycomb core 22 to outer skin 20. Expanding adhesive 26 also cures to form a rigid support for recessed edge 16.

Because recessed edge 16 is supported by a cured expanding adhesive 26, which is a rigid structure, it is able to support compressive and torsional forces. As a result, component opening 12 is structurally tied to panel 14 and recessed edge 16 provides a uniform and stable platform to support a door 28, which is illustrated in Figures 3 and 4.

Recessed edge 16 provides a vast improvement over prior component openings because component opening 12 and recessed edge 16 may be formed to very close tolerances.

As a result of these tolerances, doors 28 fit much better and may not require labor intensive hand fitting. Doors 28, consequently, provide better tactile feedback and withstand wind and water better than hand-fitted and filled doors. Manufacturing processes are also simplified because any door 28 may be manufactured to perfectly fit component opening 12 on any helicopter. As a result, manufacturing quality is improved while saving labor costs associated with trial and error hand-fitting.

Figures 3 and 4 depict examples of interfaces between panel 14 and door 28. The interface may include a seal 30, which seals rain and wind from the interior of the helicopter. Seal 30 may be bonded to recessed edge 16 or to door 28. If seal 30 is bonded to recessed edge 16, seal 30 may have a contour that also acts as a rain gutter when door 28 is opened. Seal 30 may be made from rubber, plastic or other resilient material. A lip on seal 30 may catch water and direct it around component opening 12 to help keep passengers and the helicopter interior dry.

Frame 18 may be bonded to the interior of panel 14 to add structural rigidity and provide mounting points for other airframe section 10 components. A bulkhead 32 may be fastened to frame 18 to add rigidity to airframe section 10. Bulkhead 32 also provides stiffness to recessed edge 16 and consequently may support various door 28 loads. Combining panel 14, frame 18 and bulkhead 32 creates a structure that may support large flight and landing loads. This structure also provides solid mounting points for doors 28, windows and other components. Solid mounting points give doors 28 a tactile property that users associate with safety and quality. While this invention has been described with reference to illustrative embodiments, this description is not intended to be construed in a limiting sense. Various modifications and combinations of the illustrative embodiments, as well as other embodiments of the invention, will be apparent to persons skilled in the art upon reference to the description. It is, therefore, intended that the appended claims encompass any such modifications or embodiments.

Claims

What is claimed is : 1. A panel in a vehicle comprising: an outer skin; a recessed edge disposed within the outer skin and around the perimeter of a component opening; an inner skin; and an core disposed between the outer skin and the inner skin.

2. The panel as recited in claim 1 wherein the outer skin is a composite material.

3. The component frame as recited in claim 1 wherein the core is a structural honeycomb material.

4. The panel as recited in claim 3 wherein the outer skin is bonded to a first side of the honeycomb material and the inner skin is bonded to a second side of the honeycomb material.

5. The panel as recited in claim 1 wherein the recessed edge is comprised of the outer skin bonded to the inner skin.

6. The component frame as recited in claim 1 further comprising an expanding adhesive disposed between an edge of the core and the outer skin.

7. A method for forming a panel in an aircraft, the method comprising the steps of: forming an outer skin having a component opening, the component opening having a recessed edge; forming an inner skin having a component opening that corresponds to the outer skin; and providing a core between the outer skin and the inner skin.

8. The method recited in claim 7 wherein the outer skin is a pre-cured composite material.

9. The method recited in claim 7 wherein the core has a honeycomb structure.

10. The method recited in claim 7 wherein the core is bonded between the outer skin and the inner skin.

11. The method recited in claim 7 further comprising the step of providing an expanding adhesive between an edge of the core and the outer skin.

12. The method recited in claim 7 wherein the recessed edge is formed by bonding a portion of the outer skin to a portion of the inner skin.

13. An airframe section comprising: a panel; a frame integral with the panel; and a component supported by the panel, the component fitting within the frame and substantially flush with the outer surface of the airframe section.

14. The airframe section of claim 13 wherein the panel is a composite structure comprising: an outer shell; an inner shell; and a core disposed between the outer shell and the inner shell; and a recessed edge integral with the frame.

15. The airframe section of claim- 14 further comprising an expanding adhesive disposed between an edge of the core and the outer shell.

16. The airframe section of claim 14 wherein the outer shell is bonded to a first side of the core and the inner shell is bonded to a second side of the core.

17. The airframe section of claim 14 wherein the outer shell is cured before the inner shell.

18. The airframe section of claim 14 wherein the recessed edge is formed by bonding the inner shell to the outer shell.

19. The airframe section of claim 14 wherein the component is a door.

20. The airframe section of claim 14 wherein the component is a window.

21. The airframe section of claim 14 wherein the frame is a composite structure that structurally supports a component opening.

22. The airframe section of claim 14 wherein the frame is a composite structure that structurally supports a bulkhead.