HEAT RESISTANT COMPOSITE STRUCTURE

Abstract: A heat resistant composite structure has a structural or load bearing portion formed of a composite material such as graphite reinforced organic matrix composite and a pre-ceramic polymer composite heat shield formed upon the load bearing portion. The composite structure defining the load bearing portion and the pre-ceramic polymer composite formed thereon are co-cured. The pre-ceramic polymer composite layer protects the primary structure from heat damage by absorbing heat as the pre-ceramic polymer transforms into a ceramic. The heat resistant composite structure finds particular application in forming structures of the weapons' bay of a combat aircraft so as to mitigate damage to the aircraft due to inadvertent ordnance deflagration therein.

Description: HEAT RESISTANT COMPOSITE STRUCTURE Field of the Invention The present invention relates generally to composite structures and more particularly to a heat resistant composite structure for an aircraft weapons ' bay or the like, for mitigating the danger associated with undesirable ordnance deflagration. Background of the Invention The potential for ordnance deflagration within the weapons' bay of a combat aircraft represents a serious hazard to both the aircraft and the crew thereof. Such undesirable ordnance deflagration may occur as a result of the ordnance being struck by an artillery shell or shrapnel, for example. As those skilled in the art will appreciate, such ordnance deflagration is likely to result in catastrophic damage to the aircraft structure. For example, it is estimated that the inadvertent ignition of a rocket motor within the weapons' bay of an aircraft will result in a thermal flux of approximately 400 BTϋ/ft 2 /second. Such heat is sufficient to result in substantial erosion of weapons' bay structures, e.g., the sides and bulkheads thereof. In view of the foregoing it is desirable to provide a means for mitigating damage to aircraft structures during such deflagration, at least for sufficient time to facilitate jettisoning of the deflagrating ordnance and/or ejection of the aircraft's crew. One known method for providing heat resistance to aircraft structures involves forming a heat shield upon the aircraft structures via the addition of silicone or rubber ablatives thereto. However, as those skilled in the art will appreciate, such silicone and rubber ablatives are parasitic, i.e., they merely contribute weight and have no load bearing capacity. Thus, such contemporary silicone and rubber ablatives inherently possess a high weight and volume penalty and are therefore undesirable for use in aircraft structures where weight and volume are severely constrained. Contemporary weapons' bay structures may be formed of graphite-reinforced organic matrix composite or the like. One known method for forming a heat shield upon such a contemporary composite structure comprises bonding fiber reinforced pre-ceramic polymer to the graphite reinforced organic matrix composite. This contemporary process for forming a heat resistant composite structure thus comprises the steps of forming and curing the graphite reinforced organic matrix composite structure and separately forming and curing the fiber reinforced pre-ceramic polymer. After both the graphite reinforced organic matrix composite and the fiber reinforced pre- ceramic polymer have been separately cured, then they are bonded to one another to form a heat resistant structure. As those skilled in the art will appreciate, such a heat shield resists heat by absorbing heat as the pre- ceramic transforms into a ceramic. Thus, substantial heat is absorbed by the pre-ceramic in the transformation process, thereby substantially mitigating heat damage to the underlying graphite reinforced organic matrix composite. The pre-ceramic also resists heat by ablating, i.e., blistering, de-laminating, and/or peeling away. Although such bonded heat resistant composite structures have proven generally suitable for their intended purpose, such bonded heat resistant composite structures suffer from inherent deficiencies. For example, the construction of a bonded heat resistant composite structure is comparatively complex in that it involves the separate construction of the composite structure and the pre-ceramic polymer covering therefore, as well as the bonding of the pre-ceramic polymer to the structural composite. As those skilled in the art will appreciate, such increased complexity increases the cost and the likelihood of defects occurring in the heat resistant composite structure. As such, it is beneficial to provide a method for forming a heat resistant composite structure which is comparatively less complex, and thus less expensive and less subject to defect. Summary of the Invention The present invention specifically addresses and alleviates the above-mentioned deficiencies associated with the prior art. More particularly, the present invention comprises a heat resistant composite structure, comprising a structure formed of an organic matrix composite material and a pre-ceramic polymer formed upon the structure so as to provide heat resistance therefor. Thus, the pre-ceramic polymer functions as a heat shield for the composite structure. The composite material of the structure and the pre-ceramic polymer formed thereon are co-cured so as to reduce the complexity of the fabrication process, thereby reducing the cost and increasing the reliability thereof. The pre-ceramic polymer protects the composite structure from heat damage by absorbing heat as the pre- ceramic transforms into a ceramic. The pre-ceramic polymer also protects the composite structure from heat damage by ablating. As those skilled in the art will appreciate, the transformation of a pre-ceramic polymer into a ceramic material is an endother ic reaction requiring substantial heat absorption and the ablation process carries heat away from the ablative covered structure as the heated ablative both undergoes chemical decomposition and wears away. As used herein the term structure is defined to include panels, bulkheads, supports, etc., such as those utilized as the structural components of an aircraft. However, those skilled in the art will appreciate that the heat resistant composite structure of the present invention is suitable for use in various different applications. As such, description of the heat resistant composite structure herein as a structure of an aircraft weapons' bay is of by way illustration only and is not by way of limitation. As those skilled in the art will appreciate, a pre- ceramic polymer material is capable of absorbing a substantial quantity of heat during the transformation thereof into a ceramic material, as well as during the ablation thereof. Thus, the pre-ceramic polymer material is therefore capable of substantially mitigating heat damage to an underlying structure, such as when inadvertent deflagration of ordnance occurs in an aircraft weapons' bay. Such mitigation of heat damage is intended to provide sufficient time for the deflagrating ordnance to be jettisoned and/or for the aircraft crew to eject. The present invention may provide sufficient heat protection to allow smaller items of ordnance to deflagrate completely within the weapons' bay without substantial damage to the aircraft. The structure preferably comprises a graphite reinforced organic matrix composite and at least one layer of fiber reinforced pre-ceramic polymer is formed upon the graphite reinforced organic matrix composite to provide a heat resistant composite structure therefor. Thus, the layer(s) of fiber reinforced pre-ceramic polymer are co-cured along with the graphite reinforced organic matrix composite. In this manner a heat resistant composite structure is formed wherein the layer (s) of fiber reinforced pre-ceramic polymer protect the graphite reinforced organic matrix composite structure from heat damage by absorbing heat as the pre- ceramic transforms into a ceramic . The layers of fiber reinforced pre-ceramic polymer preferably comprise ceramic matrix composite precursor. They preferably comprise continuous fiber reinforced pre- ceramic polymer. Those skilled in the art will appreciate that various other pre-ceramic polymers are likewise suitable. The composite structure preferably comprises carbon fiber reinforced bismaleimide (BMI) composite. Those skilled in the art will appreciate that various other composite materials are likewise suitable. Optionally, the modulus of elasticity of the pre- ceramic is selected so as to provide desired damping for acoustic loading. In this manner, structures may be formed which are less susceptible to vibration. That is, the damping characteristics of the structure may be tailored so as to mitigate vibration thereof by selecting the modulus of elasticity of the pre-ceramic polymer so as to provide the desired damping. According to the preferred embodiment of the present invention a plurality of layers of fiber reinforced pre- ceramic polymer are formed upon the structure, so as to enhance the heat resistance thereof. The pre-ceramic polymer preferably comprises a pre-ceramic polymer which forms a ceramic when subjected to a temperature of between approximately 600 °F and approximately 2000 °F. Co-curing the pre-ceramic polymer along with the structural composite substantially simplifies the fabrication of a heat resistant composite structure, thereby substantially reducing the costs associated therewith, as well as substantially reducing the likelihood of defects and thus enhancing the reliability thereof. These, as well as other advantages of the present invention will be more apparent from the following description and drawings. It is understood that changes in the specific structure shown and described may be made within the scope of the claims without departing from the spirit of the invention. Brief Description of the Drawings Figure 1 is a fragmentary cross-sectional view showing a bismaleimide bagging profile wherein the composite structure is fabricated and cured separately from the pre-ceramic polymer (not shown) according to contemporary methodology; Figure 2 is a fragmentary cross-sectional view showing the BLACKGLAS/bismaleimide co-bonding bagging profile according to a first embodiment of the present invention; and Figure 3 is a fragmentary cross-sectional view showing the Applied Poleramic CO-2/bismaleimide co- bonding bagging profile according to a second embodiment of the present invention. Detailed Description of the Preferred Embodiments The detailed description set forth below in connection with the appended drawings is intended as a description of the presently preferred embodiments of the invention, and is not intended to represent the only forms in which the present invention may be constructed or utilized. The description sets forth the functions and the sequence in steps for constructing and operating the invention in connection with the illustrated embodiments. It is to be understood, however, that the same or equivalent functions may be accomplished by different embodiments that are also intended to be encompassed within the spirit and scope of the invention. The fire shield of the present invention is illustrated in Figures 2 and 3 which depict two presently preferred embodiments thereof. Figure 1 depicts the prior art method for forming a composite structure to which a pre-formed and already cured pre-ceramic polymer may be bonded to form a heat resistant composite structure. As described in detail below, various different materials and methodologies are suitable for use in the practice of the present invention. Two specific examples of materials and methodology suitable for use in the practice of the present invention are described in detail in relation to Figures 2 and 3 which depict the exemplary processes and materials. Referring now to Figure 1 , a composite structure is formed according to contemporary methodology by placing a vacuum bag 10 over N-10 AIRWEAVE breather (1 ply) 12, which is disposed over 1 ply 181 glass 14, which is disposed over 1 ply fluorinated ethylene propylene (FEP) form 16, which is disposed over 1 ply porous ARMALON 18, which is disposed over bismaleimide prepreg 20, which is disposed over 1 ply porous ARMALON 22, forms a layer over 1 ply FEP film 24. An example of the bismaleimide prepreg fiber is type 5250-4 woven AS4-6K-5HS manufactured by Cytec of Aneheim, California. ARMALON is manufactured by Fairprene Industrial Products, Inc. of Charleston, South Carolina. The entire layered profile is disposed upon flat tool 26. Double sided adhesive tape 28 in cooperation with silicone dam 30 and flash breaker tape 34 applied to the edges of the silicone rubber dam 30 define a form of the end of the structural composite. Vacuum bag sealer 32 provides a vacuum seal between the vacuum bag 10 and the flat tool 26. According to contemporary methodology, the bismaleimide lay-up profile of Figure 1 is subjected to a cure cycle comprising: 1. Apply and maintain full vacuum throughout cure. 2. Apply 45 ±5 psi autoclave pressure. 3. Heat at 5°F/Minute to 350 °F. 4. Hold at 350 °F for six hours. 5. Cool under pressure and vacuum to below 150 °F. The profile is then subjected to a post cure cycle comprising the steps of: 1. Heat at 5°F to 440° F. 2. Hold for six hours. 3. Cool to below 150°F (postcure freestanding). In this manner, a structural composite is formed which is suitable for use in aircraft applications, such as for a structural panel of an aircraft weapons' bay. However, as those skilled in the art will appreciate, such a structural panel suffers from the inherent deficiency of being susceptible to damage caused by ordnance deflagration within the aircraft weapons' bay. As discussed above, such a structural composite may be shielded from heat via the use of silicone or rubber ablatives. However, such silicone and rubber ablatives are parasitic, i.e., contribute weight but not structural strength, and also carry a high weight and volume penalty. Therefore, the use of silicone or rubber adhesives is not desirable in aircraft structures where weight and volume constraints are strict. It is also known to apply one or more layers of preformed pre-ceramic polymer to such a preformed structural composite (like that formed according to Figure 1) via bonding. Although substantially enhanced heat resistance is provided by such a structure, the process of separately forming the composite structure and the pre-ceramic polymer and then bonding the two together is comparatively complex. Such complexity undesirably increases costs and inhibits the production of a defect free heat resistant composite structure. Referring now to Figure 2, according to first exemplary embodiment of the present invention a vacuum bag 110 is applied over an N-10 AIRWEAVE breather (1 ply) 112, which forms a layer over 2 plies of 181 glass 114, which forms a layer over of 5 mil TEFLON film 116, which forms a layer over BLACKGLAS laminate 118, which forms a layer over bismaleimide substrate 120, which forms a layer over non-porous ARMALON 122. The entire layered assembly is disposed upon flat tool 124 and vacuum bag sealant 126 facilitates the application of a vacuum to the layered assembly via vacuum bag 110. The BLACKGLAS laminate matrix is composed of type 489C resin manufactured by Allied Signal Advanced Materials of Santa Clara, California. The fiber is NEXTEL 312 AF10 fabric manufactured by 3M of St. Paul, Minnesota. Prior to prepregging, the Nextel 312 is nitrided by Allied Signal to produce a boron nitride surface structure on the fibers. TEFLON is a federally registered trademark of Du Pont de Nemours and Co. of Wilmington, Delaware. AIRWEAVE is a federally registered trademark of Airtech International, Inc. of Carson, California. As those skilled in the art will appreciate, the BLACKGLAS laminate 118 comprises a pre-ceramic polymer which will be co-cured along with the bismaleimide substrate 120 when subjected to the following curing schedule: 1. Apply and maintain full vacuum. 2. Heat from room temperature to 350° F at 5° F/minute maximum. 3. At 150° F, apply 100 psi autoclave pressure. 4. Vent vacuum to atmosphere at 10 psi autoclave pressure. 5. Hold at 350 °F for 2 hours minimum. 6. Cool to below 150 °F, then vent pressure. Referring now to Figure 3, according to a second exemplary embodiment of the present invention a vacuum bag 210 is applied over 2 plies of 181 glass 212, which forms a layer over 5 mil-thick TEFLON film 214, forms a layer over Applied Poleramic CO-2 laminate 216, which forms a layer over bismaleimide substrate 218, which forms a layer over non-porous ARMALON 220. The layered assembly is disposed upon flat tool 224 and the vacuum bag sealant 226 facilitates the application of the vacuum to the layered assembly via vacuum bag 210. The layered assembly is subjected to the following cure cycle: 1. Apply 60 psi immediately, allowing autoclave air temperature to rise. 2. Vent vacuum at 10 psi autoclave pressure. 3. Heat from room temperature 200 °F/minute. 4. Hold at 220 °F for 15 minutes. 5. Heat from 220 T to 300 °F at 5 °F/minute. 6. Hold at 300 "F for 30 minutes. 7. Cool to below 150 °F at 5-10 °F/minute. 8. Vent pressure and cool to room temperature. It will be appreciated by those skilled in the art that co-curing the composite structure along with the pre-ceramic polymer substantially simplifies the process of forming a heat resistant composite structure so as to substantially reduce the costs associated therewith and also so as to increase the reliability thereof by mitigating the occurrence of defects in the finished product. It is understood that the exemplary fire shield described herein and shown in the drawings represents only a presently preferred embodiments of the invention. Indeed, various modifications and additions may be made to such embodiments without departing from the spirit and scope of the invention. For example, the pre-ceramic polymer may be preformed upon a variety of different types of composite structures and co-cured along therewith. Indeed, the fire shield of the present invention may find application in a variety of different fields. Thus, these and other modifications and additions may be obvious to those skilled in the art and may be implemented to adapt the present invention for use in a variety of different applications.

Claims: WHAT IS CLAIMED IS: 1. A heat resistant composite structure comprising: a) a structure formed of a composite material; and b) a pre-ceramic polymer formed to the structure; c) wherein the composite material of the structure and the pre-ceramic polymer formed thereto are co-cured; and d) wherein the pre-ceramic polymer protects the structure from heat damage by absorbing heat as the pre-ceramic polymer transforms into a ceramic. 2. The heat resistant composite structure as recited in Claim 1 wherein: a) the structure comprises graphite reinforced organic matrix composite; and b) at least one layer of fiber reinforced pre-ceramic polymer is formed upon the graphite reinforced organic matrix composite. 3. The heat resistant composite structure as recited in Claim 1 wherein the pre-ceramic polymer comprises ceramic matrix composite precursor. 4. The heat resistant composite structure as recited in Claim 1 wherein the pre-ceramic polymer comprises continuous fiber reinforced pre-ceramic polymer. 5. The heat resistant composite structure as recited in Claim 1 wherein the pre-ceramic polymer is selected so as to have a modulus of elasticity which provides desired damping for acoustic loading. 6. The heat resistant composite structure as recited in Claim 1 wherein the pre-ceramic polymer comprises a pre-ceramic polymer which forms a ceramic when subjected to a temperature of between approximately 600 °F and approximately 2000 . 7. The heat resistant composite structure as recited in Claim 1 wherein the composite material of the structure comprises carbon fiber reinforced bismaleimide composite. 8. The heat resistant composite structure as recited in Claim 2 wherein the layer (s) of fiber reinforced pre-ceramic polymer comprise a plurality of layers of fiber reinforced pre-ceramic polymer. 9. A heat resistant composite structure for an aircraft weapons' bay, the heat resistant composite structure comprising: a) a graphite reinforced organic matrix composite structure b) at least one layer of fiber reinforced pre-ceramic polymer formed upon the graphite reinforced organic matrix composite structure, the graphite reinforced organic matrix composite structure defining at least a portion of the aircraft's weapons' bay; c) wherein the layer(s) of fiber reinforced pre-ceramic polymer mitigate heat damage to at least a portion of the aircraft's weapon's bay during ordnance deflagration by absorbing heat as the pre- ceramic polymer transforms into a ceramic. 10. A method for forming a heat resistant composite structure, the method comprising the steps of: a) forming a structure comprised of a composite; b) forming at least one layer of pre-ceramic polymer upon the composite material of the structure; and c) co-curing the layer(s) of pre-ceramic polymer along with the composite structure. 11. The method as recited in Claim 10 wherein the step of forming at least one layer of pre-ceramic polymer upon the composite structure comprises forming a ceramic matrix composite precursor laminate upon a graphite reinforced organic matrix composite. 12. The method as recited in Claim 10 wherein the step of forming at least one layer of pre-ceramic polymer composite upon the composite structure comprises forming at least one layer of continuous fiber reinforced pre- ceramic polymer upon a graphite reinforced organic matrix composite. 13. The method as recited in Claim 10 wherein the step of forming at least one layer of pre-ceramic polymer upon the composite structure comprises forming at least one layer of fiber reinforced pre-ceramic polymer having a modulus of elasticity which is selected so as to provide desired damping for acoustic loading. 14. The method as recited in Claim 10 wherein the step of forming at least one layer of pre-ceramic upon the composite structure comprises forming a plurality of layers of fiber reinforced pre-ceramic upon a graphite reinforced organic matrix composite structure. 15. The method as recited in Claim 10 wherein the step of forming at least one layer of pre-ceramic polymer upon the composite structure comprises forming at least one layer of fiber reinforced pre-ceramic polymer which forms a ceramic composite when subjected to a temperature of between approximately 600 °F and approximately 2000 °F upon the graphite reinforced organic matrix composite structure. 16. The method as recited in Claim 10 wherein the step of forming at least one layer of pre-ceramic polymer composite upon the primary composite structure comprises forming ceramic matrix composite/graphite fiber reinforced bismaleimide composite upon a graphite reinforced organic matrix composite structure.

Inventors: DAWS DAVID E, HAHN MICHAEL T, SORENSON ERIC T, DAWS, DAVID, E., HAHN, MICHAEL, T., SORENSON, ERIC, T.
Applicants: NORTHROP GRUMMAN CORP, NORTHROP GRUMMAN CORPORATION
Application year: 1998
Main classification: B32B5
IPC codes: B29C70, B64C1, B64D7, C04B35, C09K21