Metal matrix composite structural panel construction

Claim 1

A method of fabricating lightweight structural panels suitable for use in aerospace vehicular construction comprising: providing a metal exterior skin for the structural panel, brazing doublers along the interior periphery of the exterior skin with the doublers being of the same composition and of essentially the same thickness as the exterior skin, brazing a plurality of equally spaced parallel honeycomb stringers along the interior surface of the exterior skin with one end surface of the honeycomb stringers being confined within the area defined by the doublers and the other end of each honeycomb stringer being capped so as to leave the honeycomb cell sidewalls exposed; and wherein the metal exterior skin is a metal matrix composite laminate sandwich construction having a first layer of titanium, three adjacent layers of silicon carbide coated boron fibers embedded within an aluminum alloy matrix, an intermediate layer of Ti-6Al-4V titanium alloy, three additional layers of silicon coated boron fibers embedded within an aluminum alloy matrix, and a second layer of titanium.

Claim 2

A method of fabricating lightweight structural panels suitable for use in aerospace vehicular construction comprising: providing a metal exterior skin for the structural panel, brazing doublers along the interior periphery of the exterior skin with the doublers being of the same composition and of essentially the same thickness as the exterior skin, brazing a plurality of equally spaced parallel honeycomb stringers along the interior surface of the exterior skin with one end surface of the honeycomb stringers being confined within the area defined by the doublers and the other end of each honeycomb stringer being capped so as to leave the honeycomb cell sidewalls exposed; and wherein the metal exterior skin is a metal matrix composite laminate sandwich construction having a first layer of titanium, three adjacent layers of silicon carbide coated boron fibers embedded within an aluminum alloy matrix, an intermediate layer of Ti-6Al-4V titanium alloy, three additional layers of silicon coated boron fibers embedded within an aluminum alloy matrix, and a second layer of titanium.

Claim 4

A method of making elongated stringer stiffeners for use in fabricating lightweight structural panels and the like comprising the steps of: providing a honeycomb core material constructed of a titanium alloy, chemically cleaning the honeycomb core by immersion in a nitric-hydrofluoric acid solution, positioning the cleaned honeycomb core with one end of the open cells thereof in contact with a sheet of 3004 aluminum braze alloy, heating the honeycomb core and braze alloy to a temperature of 1250° F. and a pressure of 1.33 mPa (1×10 -5 torr) for ten minutes to cause the braze alloy to melt and prewet the honeycomb core nodes therewith due to capillary action, providing a titanium-clad, silicon-carbide coated boron fiber/aluminum alloy matrix laminate sheet, positioning a 718 aluminum braze alloy sheet between and in contact with the prewet honeycomb core and the titanium-clad laminate sheet, vacuum brazing this assembly into a unitary mass at a temperature of 1095° F. and a pressure of 1.33 mPa (1×10 -5 torr) for three minutes to provide a cap for the honeycomb core, and cooling the brazed assembly and cutting individual elongated stringer stiffeners of the desired width from the capped honeycomb along a length of the unitary brazed mass.

Claim 5

A method of making elongated stringer stiffeners for use in fabricating lightweight structural panels and the like comprising the steps of: providing a honeycomb core material constructed of a titanium alloy, chemically cleaning the honeycomb core by immersion in a nitric-hydrofluoric acid solution, positioning the cleaned honeycomb core with one end of the open cells thereof in contact with a sheet of 3004 aluminum braze alloy, heating the honeycomb core and braze alloy to a temperature of 1250° F. and a pressure of 1.33 mPa (1×10 -5 torr) for ten minutes to cause the braze alloy to melt and prewet the honeycomb core nodes therewith due to capillary action, providing a titanium-clad, silicon-carbide coated boron fiber/aluminum alloy matrix laminate sheet, positioning a 718 aluminum braze alloy sheet between and in contact with the prewet honeycomb core and the titanium-clad laminate sheet, vacuum brazing this assembly into a unitary mass at a temperature of 1095° F. and a pressure of 1.33 mPa (1×10 -5 torr) for three minutes to provide a cap for the honeycomb core, and cooling the brazed assembly and cutting individual elongated stringer stiffeners of the desired width from the capped honeycomb along a length of the unitary brazed mass.