Multilayer ceramic composite and method of production

Abstract

An aircraft part has a multilayer ceramic composite that includes an inner skin, a ceramic insulation core, and an outer skin. The inner skin includes a plurality of layers of fiberglass impregnated with a thermoset resin. The ceramic insulation core is sufficiently flame resistant to prevent penetration of a flame with a temperature of about 2000 degrees F. for at least 15 minutes when the aircraft part is directly exposed to the flame. The outer skin includes a plurality of layers of fiberglass impregnated with a thermoset resin. The multilayer ceramic composite is manufactured using a novel method that provides superior performance at a reduced cost.

Claims

1. An aircraft part with integral thermal and flame protection, the aircraft part comprising: an inner skin comprising a plurality of layers of fiberglass impregnated with a thermoset resin; a ceramic insulation core that is sufficiently flame resistant to prevent penetration of a flame with a temperature of about 2000 degrees F. for at least 15 minutes when the aircraft part is directly exposed to the flame; and an outer skin comprising a plurality of layers of fiberglass impregnated with a thermoset resin.

2. The aircraft part of claim 1, wherein the inner skin comprises a continuously woven 8-harness satin e-glass, and is saturated with a phenolic resin.

3. The aircraft part of claim 1, wherein the outer skin comprises a continuously woven 8-harness satin e-glass, and is saturated with a phenolic resin.

4. The aircraft part of claim 1, wherein the ceramic insulation core comprises a non-woven alumina-silicate fiber.

5. The aircraft part of claim 1, wherein the inner skin comprises at least three layers of a continuously woven 8-harness satin e-glass, each saturated with a phenolic resin; wherein the outer skin comprises at least three layers of a continuously woven 8-harness satin e-glass, each saturated with a phenolic resin; and wherein the ceramic insulation core comprises at least two layers of a non-woven alumina-silicate fiber.

6. The aircraft part of claim 5, wherein the aircraft part is a generally annular component of a plenum.

7. The aircraft part of claim 6, further comprising plenum fittings operably mounted through the annular component.

8. A method for manufacturing an aircraft part, the method comprising the steps of: providing a mold of the aircraft part; forming an inner skin over the mold by applying a plurality of layers of fiberglass impregnated with a thermoset resin; covering the inner skin with a ceramic insulation core; forming an outer skin over the ceramic insulation core by applying a plurality of layers of fiberglass impregnated with a thermoset resin; positioning the mold and its contents in a vacuum bag; pumping air out of the vacuum bag so that approximately one atmosphere of pressure is applied to the contents of the vacuum bag; placing the vacuum bag into a non-pressurized curing oven; and baking the contents of the vacuum bag at about 250-300 degrees F. in accordance with the cure rate schedule to cure the multilayer ceramic composite.

9. The aircraft part of claim 8, wherein the inner skin comprises a continuously woven 8-harness satin e-glass, and is saturated with a phenolic resin.

10. The aircraft part of claim 8, wherein the outer skin comprises a continuously woven 8-harness satin e-glass, and is saturated with a phenolic resin.

11. The aircraft part of claim 8, wherein the ceramic insulation core comprises a non-woven alumina-silicate fiber.

12. The aircraft part of claim 8, wherein the ceramic insulation core is sufficiently flame resistant to prevent penetration of a flame with a temperature of about 2000 degrees F. for at least 15 minutes when the aircraft part is directly exposed to the flame.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

(1) The accompanying drawings illustrate the present invention. In such drawings:

(2) FIG. 1 is a perspective view of a non-structural aircraft part, in this case a plenum component, that is manufactured using a multilayer ceramic composite according to one embodiment of the present invention;

(3) FIG. 2 is a sectional view of the multilayer ceramic composite taken along line 2-2 in FIG. 1;

(4) FIG. 3 is a perspective view of a mold over which the multilayer ceramic composite has been spread, the mold having been placed in a vacuum bag prior to curing; and

(5) FIG. 4 is a perspective view of another non-structural aircraft part, in this case another embodiment of a plenum.

DETAILED DESCRIPTION OF THE INVENTION

(6) The above-described drawing figures illustrate the invention, a multilayer ceramic composite and method of manufacturing a non-structural aircraft part using the composite.

(7) FIG. 1 is a perspective view of one embodiment of the non-structural aircraft part 10 that is manufactured using the multilayer ceramic composite 20. As shown in FIG. 1, the multilayer ceramic composite is particularly useful in the construction of non-structural aircraft parts, such as a plenum component. The plenum component may include a generally annular construction (or, in this case, a portion of such an annular construction) that includes an outer rim 12 and an inwardly extending flange 14. Fittings 16 may be installed on the plenum component for operative interconnection with an auxiliary power unit (APU) of an aircraft (not shown), and mounting elements 17 may be used for mounting the plenum component, or, in this case, for connecting a second plenum component (not shown). The plenum component may also include apertures 18 and similar cutouts for directing air flow. The plenum component may be installed around an engine (not shown) for directing air flow, as is known in the art. While a plenum component is illustrated herein, alternative non-structural aircraft parts (e.g., ducting, etc.) may also be constructed using the teachings of the present invention.

(8) FIG. 2 is a sectional view of the multilayer ceramic composite 20 taken along line 2-2 in FIG. 1. As shown in FIG. 2, the multilayer ceramic composite includes an inner skin 22 and an outer skin 26 that are both constructed of fiber reinforcement saturated (wet out) with a thermoset resin. The inner and outer skins surround and sandwich a ceramic insulation core. 24.

(9) In one embodiment, the inner skin and the outer skin are constructed of a continuously woven 8-harness satin e-glass, and is saturated with a phenolic resin (e.g., a fiberglass phenolic prepreg). While these embodiments are provided to illustrate the present invention, those skilled in the art may select similar or equivalent materials consistent with the teachings of the present invention, and such similar or equivalent materials should be considered within the scope of the present invention.

(10) In this embodiment, the ceramic insulation core includes a non-woven alumina-silicate fiber that is flexible and capable of being formed in or around a suitable mold. In alternative embodiments, alternative insulating materials may be used, so long as they provide the necessary insulation at an acceptable cost. Equivalent insulation materials should be considered within the scope of the present invention.

(11) In the embodiment of FIG. 2, during production, multiple layers of fiberglass phenolic prepreg are stacked together to form the inner skin. In this case, six layers are used, although the specific number of layers may be varied depending upon the particular requirements of each product. Then, the ceramic insulation core is disposed on the inner skin. In this embodiment, one or more layers of alumina-silicate fiber are stacked together, and may be held in place with an adhesive. This ceramic insulation core comprises a flexible non-woven material, so that the material may be formed in mold into the non-structural aircraft part (such as shown in FIG. 1).

(12) Finally, multiple layers of fiberglass phenolic prepreg are stacked together to form the outer skin, to form the multilayer ceramic composite. In this case, three layers are used; although, as noted before, the specific number of layers may be varied depending upon the particular requirements of each product. This process may obviously be performed in different orders, or using equivalent techniques known in the art, and such alternatives should be considered within the scope of the present invention.

(13) In this embodiment, the overall cured phenolic and insulation material is about 3/16 inch thick. The ceramic insulation core should be thick enough to prevent a 2000 degree F. flame applied from the exterior side of the multilayer ceramic composite from raising the temperature on the opposite side to beyond 250 degrees F. The described ceramic insulation core has been found to keep the temperature on the non-burning side of the multilayer ceramic composite around or below 220 degrees F. By keeping the temperature on the non-burning side to 250 degrees F. or lower, this prevents any volatiles in the prepreg materials on the non-burning side from flashing or igniting for a short period (and thus failing to meet government safety requirements).

(14) FIG. 3 is a perspective view a mold 30 over which the multilayer ceramic composite 20 has been spread during production of the non-structural aircraft part 10 (of FIG. 1). The mold and multilayer ceramic composite are also shown having been placed in a vacuum bag 32 prior to curing. As shown in FIG. 3, the multilayer ceramic composite is flexible enough to be spread or otherwise positioned over or within the mold so as to form the correct shape. The mold is placed in the vacuum bag, which is operatively connected to a vacuum pump 34. Once air has been evacuated from the vacuum bag, the multilayer ceramic composite is subjected to a low pressure (14 to 14.5 psi) by the atmosphere, without any additional pressure being required from another source, such as an autoclave. The mold may then be placed in a curing oven (autoclave not required) at about 250-300 degrees F. and cured in accordance with the cure rate schedule to cure the multilayer ceramic composite.

(15) This process enables the inexpensive production of the plenum component with a complex non-geometric shape, without the need for complicated and expensive specialty molds or tooling, or an autoclave. This combination of materials and process has been demonstrated to produce a component which will meet the requirements of the FAA's AC20-135 and EASA's version of the Fire Test Requirements for commercial aircraft. The multilayer ceramic composite further does not require expensive carbon fiber material, or other expensive reinforcement materials such as Nextel, Kevlar, etc., to be fireproof.

(16) Once cured, the multilayer ceramic composite may be removed from the mold, fabricated in accordance with the customer's requirement, and the fittings and other components added (as shown in FIG. 1). However, a ceramic blanket is not included, and there are no added protective metal components to allow the part to pass flammability tests, as the multilayer ceramic composite is able by itself to pass the required flammability tests.

(17) For purposes of this application, the term about is defined to mean +/10%. This does not preclude alternatives outside of the described ranges, unless specifically required in the claims, and no limitations beyond the claims should be inferred from the embodiments described in the specification.

(18) FIG. 4 is a perspective view of another non-structural aircraft part 40, in this case another embodiment of a plenum. The particular shape of the plenum will vary depending upon the particular engine used, and alternative configurations should be considered within the scope of the present invention. Furthermore, while the plenum illustrates one embodiment of the non-structural aircraft part, other aircraft components known in the art may also utilize the multilayer ceramic composite in their construction, consistent with the teachings on the present invention.

(19) As used in this application, the words a, an, and one are defined to include one or more of the referenced item unless specifically stated otherwise. Also, the terms have, include, contain, and similar terms are defined to mean comprising unless specifically stated otherwise. Furthermore, the terminology used in the specification provided above is hereby defined to include similar and/or equivalent terms, and/or alternative embodiments that would be considered obvious to one skilled in the art given the teachings of the present patent application. While the invention has been described with reference to at least one particular embodiment, it is to be clearly understood that the invention is not limited to these embodiments, but rather the scope of the invention is defined by the following claims.