An all-composite assembly such as a composite laminate aircraft empennage has vertical and horizontal stabilizers with differing sets of interlaminar fracture toughnesses and differing stiffnesses to improve flight characteristics. Composite laminate skins are bonded to unitized and stiffened unders
An all-composite assembly such as a composite laminate aircraft empennage has vertical and horizontal stabilizers with differing sets of interlaminar fracture toughnesses and differing stiffnesses to improve flight characteristics. Composite laminate skins are bonded to unitized and stiffened understructure to reduce weight and improve damage containment.
대표청구항▼
1. A composite assembly, comprising: a first composite structure, the first composite structure including a composite first understructure and a composite first laminate skin bonded to the first understructure, the composite first laminate skin having a first set of pre-selected interlaminar fractur
1. A composite assembly, comprising: a first composite structure, the first composite structure including a composite first understructure and a composite first laminate skin bonded to the first understructure, the composite first laminate skin having a first set of pre-selected interlaminar fracture toughnesses, wherein the first understructure comprises: a plurality of longitudinally extending composite spars, anda plurality of Z-shaped composite stiffeners extending between and bonded to the spars, and a plurality of longitudinally extending, composite stringers passing through the Z-shaped composite stiffeners and bonded to at least one of the composite first laminate skin and composite second laminate skin; andat least one second composite structure, the second composite structure including a composite second understructure and a composite laminate second skin bonded to the second understructure, the composite laminate second skin having a second set of pre-selected interlaminar fracture toughnesses. 2. The composite assembly of claim 1, further comprising: each of the composite first laminate skin and the composite laminate second skin being subject to Mode I, II and III loading, andthe first set of pre-selected interlaminar fracture toughness and the second set of preselected interlaminar fracture toughness differ from each other in the Modes I, II, and III loading. 3. The composite assembly of claim 1, wherein: the first composite structure comprises a first torsional stiffness, andthe second composite structure comprises a second torsional stiffness, the first torsional stiffness being greater than the second torsional stiffness. 4. The composite assembly of claim 3, wherein: the first torsional stiffness comprises a range of approximately 45.0 to 52.0 million pounds per square inch, andthe second torsional stiffness comprises a range of approximately 40.0 to 50.2 million pounds per square inch. 5. The composite assembly of claim 1, wherein the first set of pre-selected interlaminar fracture toughnesses of the composite laminate first skin of the first composite structure comprises: a Mode I interlaminar fracture toughness within a range of approximately 4.0 to 6.5 inch-pounds per square inch,a Mode II interlaminar fracture toughness within a range of approximately 12.0 to 15.5 inch-pounds per square inch, anda Mode III interlaminar fracture toughness within a range of approximately 16.0 to 18.5 inch-pounds per square inch. 6. The composite assembly of claim 5, wherein the second set of pre-selected interlaminar fracture toughnesses of the composite laminate second skin of the second composite structure comprises: a Mode I interlaminar fracture toughness within a range of approximately 2.5 to 3.5 inch-pounds per square inch,a Mode II interlaminar fracture toughness within a range of approximately 7.5 to 9.5 inch-pounds per square inch, anda Mode III interlaminar fracture toughness within a range of approximately 18.0 to 20.5 inch-pounds per square inch. 7. The composite assembly of claim 1, wherein: each of the composite spars and the Z-shaped composite stiffeners is generally I-shaped in cross-section and includes a pair of caps, andthe caps of the composite spars and the caps of the Z-shaped composite stiffeners are bonded together. 8. The composite assembly of claim 1, further comprising: a plurality of substantially straight, composite cross-beams respectively passing through the Z-shaped composite stiffeners and extending substantially normal to the composite spars. 9. The composite assembly of claim 1, where the first composite structure and the second composite structure are arranged to form an aircraft empennage. 10. The composite assembly of claim 1, wherein: the first composite structure is an aircraft vertical stabilizer andthe second composite structure is an aircraft horizontal stabilizer. 11. An aircraft empennage, comprising: a vertical stabilizer having a composite laminate first skin and a composite first understructure bonded to the first skin, such that the composite first understructure comprises a first integrated grid that comprises first composite spars, first composite cross-beams, and first composite stiffeners bonded together, wherein the composite laminate first skin comprises: a Mode I interlaminar fracture toughness within a range of approximately 4.0 to 6.5 inch-pounds per square inch;a Mode II interlaminar fracture toughness within a range of approximately 12.0 to 15.5 inch-pounds per square inch; anda Mode III interlaminar fracture toughness within a range of approximately 16.0 to 18.5 inch-pounds per square inch; anda pair of horizontal stabilizers, each of the horizontal stabilizers including a composite laminate second skin and a composite second understructure bonded to the first skin, the composite second understructure comprising a second integrated grid that comprises second composite spars, second composite cross-beams, and second composite stiffeners bonded together. 12. The aircraft empennage of claim 11, wherein the composite laminate second skin comprises: a Mode I interlaminar fracture toughness within a range of approximately 2.5 to 3.5 inch-pounds per square inch,a Mode II interlaminar fracture toughness within a range of approximately 7.5 to 9.5 inch-pounds per square inch, anda Mode III interlaminar fracture toughness within a range of approximately 18.0 to 20.5 inch-pounds per square inch. 13. The aircraft empennage of claim 12, wherein the vertical stabilizer comprises a torsional stiffness in a range of approximately 45.0 to 52.0 million pounds per square inch. 14. The aircraft empennage of claim 13 wherein each of the horizontal stabilizers comprises a bending stiffness in a range of approximately 30.0 to 36.5 million pounds per square inch. 15. The aircraft empennage of claim 11, wherein each of the spars comprises a bending stiffness of approximately 45 million pounds per square inch. 16. The aircraft empennage of claim 11, wherein: each of the composite stiffeners comprises a Z-shape, andeach of the composite cross-beams passes through one of a Z-shaped composite stiffener. 17. An aircraft empennage, comprising: a vertical stabilizer, the vertical stabilizer including a composite first understructure and a composite laminate first skin bonded to the first understructure, the composite laminate first skin being subject to Mode I, II and III loading and having a first set of interlaminar fracture toughnesses in Modes I, II, and III; andat least one horizontal stabilizer, the horizontal stabilizer including a composite second understructure and a composite laminate second skin bonded to the second understructure, the composite laminate second skin being subject to Mode I, II and III loading and having a second set of interlaminar fracture toughnesses in Modes I, II, and III that are lesser in value than the first set of interlaminar fracture toughnesses. 18. The aircraft empennage of claim 17, wherein the composite laminate first skin comprises: a Mode I interlaminar fracture toughness within a range of approximately 4.0 to 6.5 inch-pounds per square inch;a Mode II interlaminar fracture toughness within a range of approximately 12.0 to 15.5 inch-pounds per square inch; anda Mode III interlaminar fracture toughness within a range of approximately 16.0 to 18.5 inch-pounds per square inch. 19. The aircraft empennage of claim 17, wherein the vertical stabilizer comprises a torsional stiffness in a range of approximately 45.0 to 52.0 million pounds per square inch. 20. The aircraft empennage of claim 17, wherein the horizontal stabilizer comprises a torsional stiffness in a range of approximately 40.0 to 50.2 million pounds per square inch. 21. The aircraft empennage of claim 17, wherein at least one understructure comprises a Z-shaped composite stiffener.
Westre Willard N. ; Allen-Lilly Heather C. ; Ayers Donald J. ; Cregger Samuel E. ; Evans David W. ; Grande Donald L. ; Hoffman Daniel J. ; Rogalski Mark E. ; Rothschilds Robert J., Titanium-polymer hybrid laminates.
※ AI-Helper는 부적절한 답변을 할 수 있습니다.