WO1998056574A1 - Structure composite resistant a la chaleur - Google Patents

Structure composite resistant a la chaleur Download PDF

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Publication number
WO1998056574A1
WO1998056574A1 PCT/US1998/010211 US9810211W WO9856574A1 WO 1998056574 A1 WO1998056574 A1 WO 1998056574A1 US 9810211 W US9810211 W US 9810211W WO 9856574 A1 WO9856574 A1 WO 9856574A1
Authority
WO
WIPO (PCT)
Prior art keywords
ceramic
ceramic polymer
composite structure
composite
layer
Prior art date
Application number
PCT/US1998/010211
Other languages
English (en)
Inventor
Eric T. Sorenson
Michael T. Hahn
David E. Daws
Original Assignee
Northrop Grumman Corporation
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Northrop Grumman Corporation filed Critical Northrop Grumman Corporation
Priority to EP98937933A priority Critical patent/EP0988144A1/fr
Priority to IL13343598A priority patent/IL133435A0/xx
Priority to AU86563/98A priority patent/AU8656398A/en
Publication of WO1998056574A1 publication Critical patent/WO1998056574A1/fr

Links

Classifications

    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B29—WORKING OF PLASTICS; WORKING OF SUBSTANCES IN A PLASTIC STATE IN GENERAL
    • B29C—SHAPING OR JOINING OF PLASTICS; SHAPING OF MATERIAL IN A PLASTIC STATE, NOT OTHERWISE PROVIDED FOR; AFTER-TREATMENT OF THE SHAPED PRODUCTS, e.g. REPAIRING
    • B29C70/00—Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts
    • B29C70/04—Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts comprising reinforcements only, e.g. self-reinforcing plastics
    • B29C70/28—Shaping operations therefor
    • B29C70/30—Shaping by lay-up, i.e. applying fibres, tape or broadsheet on a mould, former or core; Shaping by spray-up, i.e. spraying of fibres on a mould, former or core
    • B29C70/34—Shaping by lay-up, i.e. applying fibres, tape or broadsheet on a mould, former or core; Shaping by spray-up, i.e. spraying of fibres on a mould, former or core and shaping or impregnating by compression, i.e. combined with compressing after the lay-up operation
    • B29C70/342—Shaping by lay-up, i.e. applying fibres, tape or broadsheet on a mould, former or core; Shaping by spray-up, i.e. spraying of fibres on a mould, former or core and shaping or impregnating by compression, i.e. combined with compressing after the lay-up operation using isostatic pressure
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B32—LAYERED PRODUCTS
    • B32B—LAYERED PRODUCTS, i.e. PRODUCTS BUILT-UP OF STRATA OF FLAT OR NON-FLAT, e.g. CELLULAR OR HONEYCOMB, FORM
    • B32B37/00—Methods or apparatus for laminating, e.g. by curing or by ultrasonic bonding
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
    • B64D7/00—Arrangements of military equipment, e.g. armaments, armament accessories, or military shielding, in aircraft; Adaptations of armament mountings for aircraft
    • C—CHEMISTRY; METALLURGY
    • C09—DYES; PAINTS; POLISHES; NATURAL RESINS; ADHESIVES; COMPOSITIONS NOT OTHERWISE PROVIDED FOR; APPLICATIONS OF MATERIALS NOT OTHERWISE PROVIDED FOR
    • C09K—MATERIALS FOR MISCELLANEOUS APPLICATIONS, NOT PROVIDED FOR ELSEWHERE
    • C09K21/00—Fireproofing materials
    • C09K21/14—Macromolecular materials
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B29—WORKING OF PLASTICS; WORKING OF SUBSTANCES IN A PLASTIC STATE IN GENERAL
    • B29K—INDEXING SCHEME ASSOCIATED WITH SUBCLASSES B29B, B29C OR B29D, RELATING TO MOULDING MATERIALS OR TO MATERIALS FOR MOULDS, REINFORCEMENTS, FILLERS OR PREFORMED PARTS, e.g. INSERTS
    • B29K2995/00—Properties of moulding materials, reinforcements, fillers, preformed parts or moulds
    • B29K2995/0012—Properties of moulding materials, reinforcements, fillers, preformed parts or moulds having particular thermal properties
    • B29K2995/0016—Non-flammable or resistant to heat
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C1/00—Fuselages; Constructional features common to fuselages, wings, stabilising surfaces or the like
    • B64C2001/0054—Fuselage structures substantially made from particular materials
    • B64C2001/0072—Fuselage structures substantially made from particular materials from composite materials
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/40—Weight reduction

Definitions

  • 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.
  • the inadvertent ignition of a rocket motor within the weapons' bay of an aircraft will result in a thermal flux of approximately 400 BT ⁇ /ft 2 /second.
  • Such heat is sufficient to result in substantial erosion of weapons' bay structures, e.g., the sides and bulkheads thereof.
  • 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.
  • 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.
  • such a heat shield resists heat by absorbing heat as the pre- ceramic transforms into a ceramic.
  • 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.
  • bonded heat resistant composite structures have proven generally suitable for their intended purpose, such bonded heat resistant composite structures suffer from inherent deficiencies.
  • 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.
  • increased complexity increases the cost and the likelihood of defects occurring in the heat resistant composite structure.
  • 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.
  • 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.
  • 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.
  • the term structure is defined to include panels, bulkheads, supports, etc., such as those utilized as the structural components of an aircraft.
  • the heat resistant composite structure of the present invention is suitable for use in various different applications.
  • 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.
  • 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.
  • 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.
  • the layer(s) of fiber reinforced pre-ceramic polymer are co-cured along with the graphite reinforced organic matrix composite.
  • 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.
  • BMI bismaleimide
  • the modulus of elasticity of the pre- ceramic is selected so as to provide desired damping for acoustic loading.
  • 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.
  • 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.
  • 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.
  • 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.
  • FIG. 2 and 3 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.
  • 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.
  • FEP fluorinated ethylene propylene
  • An example of the bismaleimide prepreg fiber is type
  • ARMALON is manufactured by Fairprene
  • 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.
  • the bismaleimide lay-up profile of Figure 1 is subjected to a cure cycle comprising:
  • a structural composite is formed which is suitable for use in aircraft applications, such as for a structural panel of an aircraft weapons' bay.
  • a structural panel suffers from the inherent deficiency of being susceptible to damage caused by ordnance deflagration within the aircraft weapons' bay.
  • such a structural composite may be shielded from heat via the use of silicone or rubber ablatives.
  • 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.
  • 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.
  • 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:
  • 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.
  • 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.
  • the pre-ceramic polymer may be preformed upon a variety of different types of composite structures and co-cured along therewith.
  • the fire shield of the present invention may find application in a variety of different fields.
  • 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.

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  • Chemical & Material Sciences (AREA)
  • Engineering & Computer Science (AREA)
  • Materials Engineering (AREA)
  • Organic Chemistry (AREA)
  • Aviation & Aerospace Engineering (AREA)
  • Composite Materials (AREA)
  • Mechanical Engineering (AREA)
  • Laminated Bodies (AREA)

Abstract

Une structure composite résistant à la chaleur comporte une partie structurale ou porteuse de charge formée d'un matériau composite, tel qu'un composite à matrice organique renforcée par du graphite, et une protection thermique composite en polymère pré-céramique formée sur la partie porteuse de charge. La structure composite définissant la partie porteuse de charge ainsi que le composite polymère pré-céramique formé sur cette dernière sont co-durcies. La couche composite en polymère pré-céramique protège la structure primaire des détériorations provoquées par la chaleur étant donné qu'elle absorbe la chaleur lorsque le polymère pré-céramique se transforme en céramique. La structure composite résistant à la chaleur s'applique tout particulièrement pour former des structures de la soute d'armes d'un avion de combat de manière à limiter les détériorations sur l'avion qui sont dues à la déflagration accidentelle des pièces d'armement dans cette même soute.
PCT/US1998/010211 1997-06-12 1998-05-19 Structure composite resistant a la chaleur WO1998056574A1 (fr)

Priority Applications (3)

Application Number Priority Date Filing Date Title
EP98937933A EP0988144A1 (fr) 1997-06-12 1998-05-19 Structure composite resistant a la chaleur
IL13343598A IL133435A0 (en) 1997-06-12 1998-05-19 Heat resistant composite structure
AU86563/98A AU8656398A (en) 1997-06-12 1998-05-19 Heat resistant composite structure

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
US87357097A 1997-06-12 1997-06-12
US08/873,570 1997-06-12

Publications (1)

Publication Number Publication Date
WO1998056574A1 true WO1998056574A1 (fr) 1998-12-17

Family

ID=25361900

Family Applications (1)

Application Number Title Priority Date Filing Date
PCT/US1998/010211 WO1998056574A1 (fr) 1997-06-12 1998-05-19 Structure composite resistant a la chaleur

Country Status (4)

Country Link
EP (1) EP0988144A1 (fr)
AU (1) AU8656398A (fr)
IL (1) IL133435A0 (fr)
WO (1) WO1998056574A1 (fr)

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB2447848A (en) * 2007-03-22 2008-10-01 Richard Kenneth Mcainsh Laminate material
CN112265347A (zh) * 2020-09-18 2021-01-26 航天特种材料及工艺技术研究所 一种结构承载-烧蚀防热一体化复合材料及其制备方法

Families Citing this family (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5084575A (en) * 1987-07-31 1992-01-28 American Home Products Corporation Quinoline substituted naphthalenepropionic acid derivatives as anti-inflammatory/antiallergic agents
US4960892A (en) * 1988-06-10 1990-10-02 American Home Products Corporation Naphthalenepropionic acid derivatives as anti-inflammatory/anti-allergic agents

Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5122226A (en) * 1987-08-12 1992-06-16 United Technologies Corporation Method of making hybrid composite structures of fiber reinforced glass and resin matrices
US5558932A (en) * 1994-09-22 1996-09-24 Auto-Air Composites, Inc. Integrated structural composite and ceramic flame barrier
US5582784A (en) * 1995-08-16 1996-12-10 Northrop Grumman Corporation Method of making ceramic matrix composite/ceramic foam panels
US5804306A (en) * 1996-11-27 1998-09-08 Northrop Grumman Corporation Ceramic matrix composite/organic matrix composite hybrid fire shield

Patent Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5122226A (en) * 1987-08-12 1992-06-16 United Technologies Corporation Method of making hybrid composite structures of fiber reinforced glass and resin matrices
US5558932A (en) * 1994-09-22 1996-09-24 Auto-Air Composites, Inc. Integrated structural composite and ceramic flame barrier
US5582784A (en) * 1995-08-16 1996-12-10 Northrop Grumman Corporation Method of making ceramic matrix composite/ceramic foam panels
US5804306A (en) * 1996-11-27 1998-09-08 Northrop Grumman Corporation Ceramic matrix composite/organic matrix composite hybrid fire shield

Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB2447848A (en) * 2007-03-22 2008-10-01 Richard Kenneth Mcainsh Laminate material
GB2447848B (en) * 2007-03-22 2011-06-15 Richard Kenneth Mcainsh Laminate material and method of making the same
CN112265347A (zh) * 2020-09-18 2021-01-26 航天特种材料及工艺技术研究所 一种结构承载-烧蚀防热一体化复合材料及其制备方法

Also Published As

Publication number Publication date
EP0988144A1 (fr) 2000-03-29
IL133435A0 (en) 2001-04-30
AU8656398A (en) 1998-12-30

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