The present disclosure will become more fully understood from the detailed description and the accompanying drawings, wherein:
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
With reference to
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
With further reference to
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
A flexible caul plate 40 having a polished surface, to form a high quality optical surface for the finished windows 16 (illustrated schematically in
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 (
As noted above, the single pane windows 16 (
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.
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.
Number | Date | Country | |
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Parent | 10654765 | Sep 2003 | US |
Child | 11612512 | US |