A composite laminate, method of forming same, and use for same are disclosed. One example of a composite laminate has multiple layers or plies (305A-305E) composed of generally parallel reinforcing fibers (315A-315E) embedded in a matrix (305M). The reinforcing fibers have orientations in the ranges
A composite laminate, method of forming same, and use for same are disclosed. One example of a composite laminate has multiple layers or plies (305A-305E) composed of generally parallel reinforcing fibers (315A-315E) embedded in a matrix (305M). The reinforcing fibers have orientations in the ranges of 3 to 8 degrees, −3 to −8 degrees, 10 to 40 degrees, −10 to −40 degrees, and approximately 90 degrees, the orientations being with respect to a predetermined axis (320), such as an axis of tension (T). A method of manufacturing a composite laminate includes laying a resin and fibers having these orientations and then curing the resulting laminate. One example of a use is for the skin on the fuselage or wing of an aircraft.
대표청구항▼
1. A wing skin, comprising: a composite laminate comprising a matrix; andreinforcing fibers embedded in the matrix, a first plurality of the reinforcing fibers generally being parallel with an orientation in a range of 3 to 8 degrees, a second plurality of the reinforcing fibers generally being para
1. A wing skin, comprising: a composite laminate comprising a matrix; andreinforcing fibers embedded in the matrix, a first plurality of the reinforcing fibers generally being parallel with an orientation in a range of 3 to 8 degrees, a second plurality of the reinforcing fibers generally being parallel with an orientation in a range of −3 to −8 degrees, a third plurality of the reinforcing fibers generally being parallel with an orientation in a range of 21 to 40 degrees, a fourth plurality of the reinforcing fibers generally being parallel with an orientation in a range of −21 to −40 degrees, and a fifth plurality of the reinforcing fibers generally being parallel with an orientation of approximately 90 degrees, the orientations being with respect to a predetermined axis, the reinforcing fibers of the composite laminate collectively having a total volume, the first and the second pluralities of the reinforcing fibers making up about 50 percent of the total volume, and the third and the fourth pluralities of the reinforcing fibers making up about 40 percent of the total volume. 2. The wing skin of claim 1 wherein the reinforcing fibers of the first plurality have an orientation of approximately 5 degrees and the reinforcing fibers of the second plurality have an orientation of approximately −5 degrees. 3. The wing skin of claim 1 wherein the reinforcing fibers of the third plurality have an orientation of approximately 30 degrees and the reinforcing fibers of the fourth plurality have an orientation of approximately −30 degrees. 4. The wing skin of claim 1 wherein the reinforcing fibers of the third plurality have an orientation of approximately 40 degrees and the reinforcing fibers of the fourth plurality have an orientation of approximately −40 degrees. 5. The wing skin of claim 1 wherein the third plurality of the reinforcing fibers have an orientation in a range of 21 to 29 degrees, and wherein the fourth plurality of the reinforcing fibers have an orientation in a range of −21 to −29 degrees. 6. The wing skin of claim 1 wherein the third plurality of the reinforcing fibers have an orientation in a range of 31 to 40 degrees, and wherein the fourth plurality of the reinforcing fibers have an orientation in a range of −31 to −40 degrees. 7. The wing skin of claim 1 wherein the composite laminate is a portion of a skin of an aircraft. 8. The wing skin of claim 1 wherein the predetermined axis is an axis of tension on the composite laminate. 9. The wing skin of claim 1 wherein two of the pluralities, both having reinforcing fibers having an identical orientation, are separated by at least one plurality which has reinforcing fibers having a different orientation. 10. The wing skin of claim 1 wherein the composite laminate has a midplane, and the pluralities are arranged symmetrically about the midplane. 11. A method of manufacturing a composite laminate of a wing skin, comprising: embedding reinforcing fibers in a matrix, a first plurality of the reinforcing fibers generally being parallel with an orientation in a range of 3 to 8 degrees, a second plurality of the reinforcing fibers generally being parallel with an orientation in a range of −3 to −8 degrees, a third plurality of the reinforcing fibers generally being parallel with an orientation in a range of 31 to 40 degrees, a fourth plurality of the reinforcing fibers generally being parallel with an orientation in a range of −31 to −40 degrees, and a fifth plurality of the reinforcing fibers generally being parallel with an orientation of approximately 90 degrees, the orientations being with respect to a predetermined axis, wherein the reinforcing fibers collectively have a total volume, the first and the second pluralities of the reinforcing fibers making up about 50 percent of the total volume, and the third and the fourth pluralities of the reinforcing fibers making up about 40 percent of the total volume. 12. The method of claim 11 wherein embedding the reinforcing fibers comprises positioning the reinforcing fibers of the first plurality to have an orientation of approximately 5 degrees, and positioning the reinforcing fibers of the second plurality to have an orientation of approximately −5 degrees. 13. The method of claim 11 wherein embedding the reinforcing fibers comprises positioning the reinforcing fibers of the third plurality to have an orientation of approximately 31 degrees, and positioning the reinforcing fibers of the fourth plurality to have an orientation of approximately −31 degrees. 14. The method of claim 11 wherein embedding the reinforcing fibers comprises positioning the reinforcing fibers of the third plurality to have an orientation of approximately 40 degrees, and positioning the reinforcing fibers of the fourth plurality to have orientations of approximately −40 degrees. 15. The method of claim 11 wherein the predetermined axis is an axis of tension. 16. The method of claim 11 wherein two of the pluralities of reinforcing fibers, both having reinforcing fibers having an identical orientation, are separated by at least one plurality of reinforcing fibers which has reinforcing fibers having a different orientation. 17. The method of claim 11 wherein the composite laminate has a midplane, and the pluralities of reinforcing fibers are arranged symmetrically about the midplane. 18. An aircraft, comprising: a fuselage;a wing assembly operatively connected to the fuselage; anda composite laminate incorporated into at least a selected portion of a skin of the fuselage or the wing assembly, the composite laminate comprising reinforcing fibers embedded in a matrix, a first plurality of the reinforcing fibers generally being parallel with an orientation in a range of 3 to 8 degrees, a second plurality of the reinforcing fibers generally being parallel with an orientation in a range of −3 to −8 degrees, a third plurality of the reinforcing fibers generally being parallel with an orientation in a range of 21 to 40 degrees, a fourth plurality of the reinforcing fibers generally being parallel with an orientation in a range of −21 to −40 degrees, and a fifth plurality of the reinforcing fibers generally being parallel with an orientation of approximately 90 degrees, the orientations being with respect to a predetermined axis such that none of the orientations are parallel with the predetermined axis, the reinforcing fibers collectively having a total volume, the first and the second pluralities of the reinforcing fibers making up about 50 percent of the total volume, and the third and the fourth pluralities of the reinforcing fibers making up about 40 percent of the total volume,wherein the first, second, third, fourth, and fifth pluralities of the reinforcing fibers abut one another to comprise a first set of layers, and wherein the composite laminate comprises a plurality of sets of layers abutting one another, each set of layers equivalent to the first set of layers. 19. The aircraft of claim 18 wherein the reinforcing fibers of the first plurality have an orientation of approximately 5 degrees and the reinforcing fibers of the second plurality have an orientation of approximately −5 degrees. 20. The aircraft of claim 18 wherein the reinforcing fibers of the third plurality have an orientation of approximately 30 degrees and the reinforcing fibers of the fourth plurality have an orientation of approximately −30 degrees. 21. The aircraft of claim 18 wherein the reinforcing fibers of the third plurality have an orientation of approximately 40 degrees and the reinforcing fibers of the fourth plurality have an orientation of approximately −40 degrees. 22. The aircraft of claim 18 wherein the third plurality of the reinforcing fibers have an orientation in a range of 21 to 29 degrees, and wherein the fourth plurality of the reinforcing fibers have an orientation in a range of −21 to −29 degrees. 23. The aircraft of claim 18 wherein the third plurality of the reinforcing fibers have an orientation in a range of 31 to 40 degrees, and wherein the fourth plurality of the reinforcing fibers have an orientation in a range of −31 to −40 degrees.
연구과제 타임라인
LOADING...
LOADING...
LOADING...
LOADING...
LOADING...
이 특허에 인용된 특허 (76)
Pridham Barry J,GBX ; Duffy Roger P,GBX ; Jones Christopher C. R.,GBX, Adhesively bonded joints in carbon fibre composite structures.
Frosch Robert A. Administrator of the National Aeronautics and Space Administration ; with respect to an invention of ( Mercer Island WA) Robinson Robert K. (Mercer Island WA) Tomlinson Harry M. (Bel, Fuselage structure using advanced technology fiber reinforced composites.
Kavesh Sheldon (Whippany NJ) Prevorsek Dusan C. (Morristown NJ), High tenacity, high modulus polyethylene and polypropylene fibers and intermediates therefore.
Roseburg Lawrence E. (Bellevue WA), Optimum aircraft body frame to body skin shear tie installation pattern for body skin/stringer circumferential splices.
Chang Fu-Kuo (Palo Alto CA) Reifsnider Kenneth (Blacksburg VA) Davidson James A. (Germantown TN) Georgette Frederick S. (Memphis TN), Orthopedic device of biocompatible polymer with oriented fiber reinforcement.
van Weperen, Karst Jan; Elemans, Norbertus Franciscus Jacobus; Jonkers, Thomas Maria, Printing form for rotary screen printing made from fiber-reinforced plastics material.
Guckert, Werner; Rauch, Siegfried, Process for producing structural parts, structural part produced by the process, thermal insulation cylinder, protective tube, heating element, stay pipe, hot-press die and thermal insulation element.
Fischer Josef (Ried im Innkreis ATX) Stephan Walter A. (Braunau ATX), Removable or hinged component for covering openings in the fuselage of an aircraft.
Garesch Carl E. (New Kensington PA) Roebroeks Gerandus H. J. J. (Den Bommel NLX) Greidanus Buwe V. W. (Delft NLX) Oost Rob C. V. (Heerjansdam NLX) Gunnink Jan W. (Nieuwerkerk a/d Ijssel NLX), Spliced laminate for aircraft fuselage.
Grantham,Kent; Harrison,Thomas D.; Kay,Robert M.; Kuss,Michael R.; Turnmire,William W.; Venskus,Mark K.; Whitcomb,Tory R.; Christie, legal representative,Rose; Christie, deceased,Peter S., Structural panels for use in aircraft fuselages and other structures.
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.
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.