WINDOW SKIN PANEL AND METHOD OF MAKING SAME

Abstract

A lightweight, structurally strong skin panel having one or more optically transparent areas forming see-through windows, and a method of making same. A plurality of layers of pre-preg fiber tape comprised of a plurality of optically transparent fibers pre-impregnated with an optically transparent resin is positioned over a plurality of metal sheets, with each metal sheet having a plurality of openings where windows are to be formed. The pre-preg tape layers and the metal sheets are layered onto one another such that one or more of the metal sheets is sandwiched between a pair of the pre-preg tape layers. The assembly is placed in a molding tool, and the tool placed within a vacuum bag. A vacuum assisted resin transfer forming process is used, together with heating of the molding tool, to produce a high strength, lightweight, integrated skin panel having optically transparent window portions. The skin panel eliminates the bulky and heavy frame structure traditionally employed on aircraft windows and has sufficient structural strength to be used as a portion of the skin of a fuselage of an aircraft without the need for reinforcing frame-like elements around the window areas.

Description

CROSS-REFERENCE TO RELATED APPLICATIONS

The present application is a continuation-in-part application of U.S. application Ser. No. 10,654,765 filed Sep. 4, 2003, and presently pending, which is incorporated by reference into the present application. The present application is also related in general subject matter to U.S. application Ser. No. 10/655,257, filed Sep. 4, 2003, and U.S. application Ser. No. 11/316,173, filed Dec. 22, 2005, the disclosures of which are also both incorporated by reference into the present application.

FIELD

The present disclosure relates to transparent window skin panels, and more particularly to a laminated transparent window skin panel and method of making same particularly well adapted for use in forming structurally strong yet lightweight, optically transparent skin panels for use with mobile platforms such as aircraft and spacecraft.

BACKGROUND

Passenger windows in most commercial aircraft are relatively small in size. This is due, in part, to the limited capabilities of current transparent window materials and also due to the heavy and complex support structure needed to support the windows within the frame of the aircraft. Such present day window assemblies used on various forms of airborne mobile platforms such as aircraft, spacecraft, rotorcraft, etc., often are of double pane construction. The use of double pane construction (involving two distinct, optically transparent window panels) has typically been required to meet structural strength goals. The use of two distinct window panels, however, adds weight that limits the payload capacity of the mobile platform.

Typically, the transparent window materials used in the above-described double panel window assemblies consist of a transparent polymer. While very successful and exhibiting such useful qualities as high durability and easy formation of complex shapes, these polymer windows do have a limited strength capability.

Windows made from transparent materials also typically require a supplemental support structure, often termed in the industry as a “doubler”, that extends about the periphery of the transparent window portion. The doubler interfaces the transparent window panel to the skin panel of the mobile platform and provides the needed structural strength at the peripheral region of the transparent window panels to enable them to be secured to the surrounding skin panel structure. This doubler support structure generally is made up of one or more window forgings, window panes, and/or stringers. Each component is designed to strengthen the skin panel which surrounds and supports the window. However, the doubler structure increases the cost and weight of the completed window assembly, thereby providing an incentive to keep passenger windows relatively small.

Accordingly, it would be highly desirable to provide a new, single pane window panel and method of making same that is both lightweight and sufficiently structurally strong to act as a structural skin panel on a mobile platform. In particular, it would be highly desirable to provide a new, lightweight, integrated, single pane window panel assembly especially well suited for an aircraft or spacecraft, or other form of mobile platform, that has sufficient structural strength to function as a skin panel for the mobile platform. Furthermore, it would be desirable to provide an integrated window panel assembly that is sufficiently structurally strong to act as a skin panel without the need to incorporate the traditional doubler support structure at the peripheral area of the window panel portion of the assembly. Such a lightweight, integrated window panel assembly would enable even larger windows to be used on present day aircraft and spacecraft without incurring significant additional weight.

SUMMARY

A transparent window skin panel and method of making same for use in a mobile platform is provided. In one embodiment the transparent window skin panel includes a plurality of metal sheets. A fiber reinforced resin at least partially surrounds the plurality of metal sheets and sandwiches at least a portion of one metal sheet therebetween. The fiber reinforced resin is optically transparent. A cutout is formed within each of the plurality of metal sheets. The cutout corresponds to a window area in the fully formed transparent window skin panel. The window skin panel is sufficiently structurally strong to act as an integral portion of the fuselage of a mobile platform, for example as a portion of a fuselage of an aircraft or spacecraft. The window skin panel has a single pane construction that is an important factor in achieving its light weight, as compared to double pane windows.

A method of manufacturing the transparent window skin panel is also provided. The method includes using a pre-preg tape layer comprised of a plurality of fibers pre-impregnated with a resin and a metal sheet. The pre-preg tape layer and the metal sheet are layered onto a tool such that the metal sheet and the pre-preg tape layer are aligned one atop the other, with the pre-preg tape layer extending over an aligned cutout or window opening area in the metal sheet. The tool, metal sheet, and pre-preg tape layer are heated such that the resin flows to partially cover the metal sheet and the fibers. When the assembly of the metal sheet and pre-preg tape layer is cured, the fibers of the pre-preg tape layer and the resin are substantially optically transparent and thus form a see-through window portion in the skin panel.

In one specific method of manufacture, a plurality of metal sheets are incorporated and a plurality of pre-preg tape layers are arranged to sandwich at least one of the metal sheets. The assembly is placed in a heated mold and formed in a single step into a lightweight, integrated skin panel having a generally optically transparent, single pane window area.

In its various embodiments, the skin panel forms a lightweight yet structurally strong panel that provides the important benefit of an integrally formed window portion. Since the window portion of the assembly effectively forms a single pane window portion that does not require any separate doubler structure around the perimeter of the transparent window portion, the assembly provides a significant weight savings over conventional double pane window assemblies presently used in many aircraft and spacecraft. The weight savings increases the payload of the aircraft or spacecraft. Alternatively, the weight savings allows significantly larger windows to be employed in a mobile platform without incurring any additional weight penalty over conventional double pane windows.

BRIEF DESCRIPTION OF THE DRAWINGS

The present disclosure will become more fully understood from the detailed description and the accompanying drawings, wherein:

FIG. 1 is a partial view of a front of an aircraft having an embodiment of a transparent window skin panel constructed in accordance with the present disclosure;

FIG. 2 is a side cross sectional view of the transparent window skin panel taken in the direction of arrow 2-2 in FIG. 1; and

FIG. 3 is an exploded perspective view of the materials used to construct the transparent window skin panel of FIG. 2.

DETAILED DESCRIPTION

The following description of various embodiment(s) is merely exemplary in nature and is in no way intended to limit the present disclosure, its application or its uses.

Referring to FIG. 1, there is illustrated one embodiment of a transparent window skin panel 10 constructed according to principles of the present disclosure. The skin panel 10 is shown mounted to and forming a fuselage portion of an aircraft 12. The aircraft 12 generally includes a skin 13. The window skin panel 10 includes a frame area 14 and a plurality of single pane windows 16. While in the particular example provided, the transparent window skin panel 10 is illustrated as including three side windows of the aircraft 12, it is to be understood that the transparent window skin panel 10 may be used in any portion of the aircraft 12 and have a single window or any plurality of windows. The single pane windows 16 may be significantly larger in area than conventional, double pane windows. However, because of the significant weight savings provided by the transparent window skin panel 10 over conventional double pane window assemblies, no additional weight penalty is incurred when using larger windows.

With reference to FIG. 2, a cross section of the transparent window skin panel 10 is shown. In one embodiment the frame 14 includes a plurality of structural sheets, for example, metal sheets, that form rigid structural panels. One or more pre-preg fiber tape layers 22 form the transparent single pane windows. The pre-preg fiber tape layers 22 are saturated with an optically transparent resin, to be described in greater detail in the following paragraphs. In the embodiment of FIG. 2, at least one of the metal sheets 20 is sandwiched between the fiber pre-preg tape layers 22. In this example three metal sheets 20 are illustrated, however, a greater or lesser number of metal sheets 20 may be used as needed to provide the desired degree of structural strength and rigidity. While the metal sheets 20 in this example are shown having a uniform thickness, it will be appreciated that different thicknesses could just as easily be used. The single pane window skin panel 10 has an allowable tension strength of preferably about 40,000-60,000 pounds per square inch per ply of fiber reinforced resin material, and more preferably about 50,000 pounds per square inch per ply of fiber reinforced resin material. This makes the transparent window skin panel 10 especially well suited for the demanding needs of commercial aircraft, military aircraft and aerospace applications.

The transparent window skin panel 10 is preferably lap spliced to the skin 13 of the aircraft 12. This lap splice (not shown) results in a high strength coupling wherein the transparent window skin panel 10 is mechanically fastened to an adjacent skin panel (not shown) of the aircraft skin 14.

Turning now to FIG. 3, one preferred method of constructing the transparent window skin panel 10 will now be described. A molding tool 24 is provided, illustrated schematically in FIG. 3, capable of receiving the components of the transparent window skin panel 10. The tool 24 has a smooth polished surface 26 shaped to form the outer surface of the transparent window skin panel 10. Alternatively, a glass mold may be used to form the smooth outer surface of the tool 24. The shape of the single pane windows 16, while illustrated as essentially rectangular and flat in FIGS. 1 and 2, may comprise any shape. For example, the single pane windows 16 could comprise round, square, oval or hexagon shapes if desired. Virtually any shape of single pane window 16 could be formed. For an aircraft application, the transparent window skin panel 10 will ideally be made with single pane windows 16 that are substantially rectangular or oval in shape, and which have a slight cross sectional curvature to match the overall curvature of the fuselage into which the transparent window skin panel 10 will be integrated.

With further reference to FIG. 3, a plurality of metal sheets 28 and a plurality of fiber pre-preg tape layers 30 are then provided. Each metal sheet 28 includes a plurality of spaced apart openings 34 formed therethrough. The metal sheets 28 are further aligned so that one of the single pane windows 16 is able to be formed in within each of the openings 34. Again, while the openings 34 (and therefore the windows 16) are illustrated as rectangular, it is to be understood that any shape may be employed. The shape of the openings 34 will dictate the shape of the single pane windows 16.

The metal sheets 28 are preferably made of aluminum due to its light weight and high strength. However, various other metals may just as easily be employed including, for example, titanium, stainless steel, magnesium or carbon steel. Preferably, the metal sheets 28 are constructed from metal foil tape laid out to form and meet the preferred shape and dimensions of the metal sheet 28. Alternatively, a single sheet of metal may be substituted for the use of a plurality of the metal sheets 28.

The pre-preg tape layers 30 each include a plurality of fiber plies 36 that are woven together to form a fiber mesh. The orientations of the fiber plies 36 are based on the desired directional strength of the transparent window skin panel 10. The fiber plies 36 may be arranged to provide unidirectional or bi-directional strength (e.g., the fiber plies 36 may run either in one direction or a plurality of directions). In one form the fiber plies 36 may be comprised of a weave of glass fibers each having a rectangular cross section. Fibers having other cross sectional shapes besides a rectangular cross sectional shape may also be used.

For commercial aircraft applications, in order to carry the loads in the fuselage, the fiber plies 36 are preferably arranged in a plurality of different orientations. Typical layup orientations are designated in degrees with zero degrees being along the longitudinal axis of the fuselage and 90 degrees being around the circumference of the fuselage. In one embodiment, the fiber plies 36 are arranged with about 25% of the plies oriented in the zero degree direction, about 25% in the 90 degree direction, about 25% in the +45 degree direction and about 25% in the −45 degree direction. The resin 38 may comprise an aliphatic epoxy resin, although various other resins that are generally transparent when fully cured may be employed. The resin 38 is also preferable selected to be highly resistant to ultraviolet degradation, and aliphatic epoxy resin meets this criterion well. The index of refraction of the resin 38 is also preferably matched to the index of refraction of the fiber plies 36.

In one embodiment, the pre-preg tape layers 30 may each be about 0.125 inch (3.175 mm) to about 12.0 inches wide (304.8 mm). However, tape layers of other suitable dimensions could just as easily be employed.

With further reference to FIG. 3, the metal sheets 28 and the pre-preg tape layers 30 are then laid atop the tool 24 in an order corresponding to the desired order of lamina in the transparent window skin panel 10. In the particular example provided, the metal sheets 28 alternate with double layers of the pre-preg tape layers 30 such that at least one of the metal sheets 28 is sandwiched between a pair of the pre-preg tape layers 30.

A flexible caul plate 40 having a polished surface, to form a high quality optical surface for the finished windows 16 (illustrated schematically in FIG. 3) is then closed onto the components. A vacuum bag 42 is used to seal the tool 24, the pre-preg tape layer 30 and the metal sheets 28. The air trapped within the vacuum bag 42 is then removed under suction. Finally, the components are placed in an autoclave 44 (illustrated schematically in FIG. 3).

The components may be heated to preferably approximately 250 degrees Fahrenheit under a pressure of preferably approximately 100-200 psi. Within the autoclave, the resin 38 melts and flows through the fiber plies 36 to fully wet (e.g. fully covering and saturating) the fiber plies 36 and metal sheets 28. The transparent window skin panel 10 is then cured at a suitable temperature, for example about 250° F., over a period of time, for example about 3-5 hours, until the resin 38 hardens. The components are then removed from the autoclave 44, vacuum bag 42, and the tool 24 and caul plate 40, and the transparent window skin panel 10 is removed. The metal sheets 28 correspond to the metal sheets 20 within the frame 14 (FIG. 2) and the resin 38 and fiber plies 36 make up the pre-preg fiber tape layers layers 22 (FIG. 2). The fiber plies 36 and resin 38 form the single pane windows 16 within each of the openings 34.

As noted above, the single pane windows 16 (FIGS. 1 and 2) are generally optically transparent. To impart transparency, the resin 38 is transparent and the fibers of the fiber plies 36 have an index of refraction such that they are substantially transparent. The index of refraction of the fiber used in the fiber plies 36 is matched to the index of refraction of the resin 38. In this way, the transparent window skin panel 10 is generally optically transparent in the areas of the openings 34 in the metal sheets 28.

By integrally forming the optically transparent resin 22 and fiber plies 36 of the single pane window 16 with the metal sheets 20 of the frame 14 area, the solid and high strength transparent window skin panel 10 is provided. Simultaneously, the heavy doubler or like support structure typically used as a reinforcing frame structure for aircraft windows is substantially eliminated, thus reducing the weight of the aircraft. This allows for larger windows to be employed, if desired, without increasing the weight of the aircraft.

In present day commercial aircraft construction, the weight savings provided by the single window pane construction of the transparent window skin panel 10 is substantial. In a large, commercial passenger jet aircraft having about 200 windows, the construction of the single pane windows 16 can produce a weight savings of about 2000 pounds, or roughly the equivalent of about 10 passengers, over a fuselage constructed with the same number of, and comparably sized, double pane windows. For a commercial passenger jet aircraft amount having about 75 windows, the weight savings is estimated to be about 500 pounds, or approximately about 2.5 passengers. This weight savings amounts to a significant fuel savings for a commercial aircraft, or alternatively can allow the payload to be increased over what could be achieved with an aircraft having conventional double pane windows,

While the present disclosure has been described in connection with aircraft windows, it will be appreciated that the various embodiments described herein can be incorporated on other forms of mobile platforms such as buses, trains, ships, rotorcraft, spacecraft, etc., where composite panels may be employed. The weight savings and structural strength provided by the window skin panel 10 is especially advantageous for use with the fuselage or body portions of mobile platforms, where the overall weight of the mobile platform is an important consideration for performance or fuel economy reasons. The present invention can also be implemented on fixed structures where lightweight panels having window portions are needed.

The description of the various embodiments herein is merely exemplary in nature. Thus, variations that do not depart from the gist of the present disclosure are intended to be within the scope of the appended claims.

Claims

1. A method of forming a high strength, structural window skin panel, comprising:

using a plurality of metal sheets to form a frame structure, wherein each said metal sheet includes an opening defining a window area;

aligning said metal sheets with one another so that said openings are aligned to form a uniform window opening;

laying a plurality of generally optically transparent, fiber pre-preg tape layers pre-impregnated with an optically transparent resin against said metal sheets such that at least one of the metal sheets is sandwiched between a pair of said fiber pre-preg tape layers;

further arranging said fiber pre-preg tape layers to fully cover said uniform window opening while overlaying said metal sheets;

heating the metal sheet and the fiber pre-preg tape layers such that said optically transparent resin flows and wets the fiber pre-preg tape layers and the metal sheets; and

curing, the fiber pre-preg tape layers and the metal sheets such that the fiber pre-impregnated resin tape layers and metal sheets form an integrated, lightweight window skin panel having an optically transparent, single pane window portion, and having an allowable tension strength of at least about 40,000 pounds per square inch per ply of and in the frame panel.

2. The method of claim 1, wherein the fiber pre-preg tape layers are each are comprised of glass fibers.

3. The method of claim 1, wherein the resin comprises an optically transparent aliphatic epoxy resin.

4. The method of claim 1, wherein the fiber pre-preg tape layers comprise fibers having an index of refraction matching an index of refraction of the resin.

5. The method of claim 1, wherein at least one of the metal sheets comprises a metallic foil strip.

6. The method of claim 1, wherein at least one of the metal sheets is comprised of aluminum.

7. The method of claim 1, wherein at least one of the metal sheets is comprised of titanium.

12. The method of claim 1, wherein the fiber pre-impregnated resin tape has a width of approximately 0.125 inch (3.175 mm) to about 12.0 inches (304.8 mm).

13. A method of manufacturing a lightweight, structurally strong, integrated transparent window skin panel, comprising:

providing a plurality of pre-preg tape layers each having fibers pre-impregnated with an optically transparent resin, the resin and the fibers being selected to have substantially the same index of refraction;

providing a plurality of metal sheets each having a plurality of spaced apart openings formed therein;

arranging said metal sheets such that said spaced apart openings in each of said sheets are aligned to form a corresponding plurality of spaced apart window opening areas;

layering the pre-preg tape layers and the metal sheets onto a tool such that the metal sheets and the pre-preg tape layers are aligned one atop the other so that the pre-preg tape layers completely cover the openings and overlay a periphery of the metal sheets, with at least a pair of the pre-preg tape layers sandwiching at least one of the metal sheets;

heating the tool, the metal sheets, and the pre-preg tape layers so that the resin flows to cover portions of the metal sheets and wets the fibers in each of the pre-preg tape layers, the resin and fibers being substantially transparent to form a plurality of see-through window portions in the skin panel, with the skin panel having a structural strength sufficiently strong to be used as a portion of a fuselage skin panel of an aircraft.

14. The method of claim 13, wherein providing the pre-preg tape layers pre-impregnated with a resin comprises providing a plurality of pre-preg tape layers each pre-impregnated with a transparent, aliphatic epoxy resin.

15. The method of claim 13, wherein providing a plurality of metal sheets comprises providing a plurality of aluminum sheets.

16. The method of claim 13, wherein providing a plurality of metal sheets comprises providing a plurality of titanium sheets.

17. The method of claim 13, wherein the fibers are comprised of glass fibers.

18. The method of claim 13, wherein providing a plurality of pre-preg tape layers comprises providing a plurality of pre-preg tape layers each having a width of approximately ⅛″ (3.175 mm) to about 12″ (304.8 mm).

19. The method of claim 13, further comprising placing a caul plate atop the metal sheets, the pre-preg tape layers and tool.

20. The method of claim 19, further comprising placing the caul plate, the metal sheets, the pre-preg tape layers, and the tool into a vacuum bag and removing the air therein.

21. The method of claim 13, using an autoclave to heat the tool, the metal sheets and the pre-preg tape layers.

22. The method of claim 21, wherein the autoclave heats the tool, metal sheets, and the pre-preg tape layers to approximately 250 degrees Fahrenheit under approximately 100 to 200 psi of pressure; and

wherein the tool, metal sheets and the pre-preg tape layers are allowed to cure at a temperature of about 250° F. for approximately 3-5 hours.