IPC분류정보
국가/구분 |
United States(US) Patent
등록
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국제특허분류(IPC7판) |
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출원번호 |
US-0624322
(2009-11-23)
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등록번호 |
US-8292226
(2012-10-23)
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발명자
/ 주소 |
- Sankrithi, Mithra M. K. V.
- Retz, Kevin
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출원인 / 주소 |
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대리인 / 주소 |
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인용정보 |
피인용 횟수 :
9 인용 특허 :
5 |
초록
▼
A method for minimizing the weight of an internally pressurized aircraft fuselage of a type that includes an elongated tubular shell having a near-elliptical cross-section with a radius R(φ)- and a curvature Curv(φ), where φ is a roll elevation angle of the shell, includes tailoring at least one str
A method for minimizing the weight of an internally pressurized aircraft fuselage of a type that includes an elongated tubular shell having a near-elliptical cross-section with a radius R(φ)- and a curvature Curv(φ), where φ is a roll elevation angle of the shell, includes tailoring at least one structural attribute of the shell as a function of at least one of the elevation angle φ, R(φ) and Curv(φ) so as to reduce the weight of the fuselage relative to an identical fuselage shell in which the same at least one structural attribute has not been so tailored.
대표청구항
▼
1. A method for minimizing the weight of an internally pressurized aircraft fuselage of a type that includes an elongated tubular shell having a central axis x, opposite nose and tail ends, and a near-elliptical cross-section having a radius R(φ) at substantially every point along the x axis between
1. A method for minimizing the weight of an internally pressurized aircraft fuselage of a type that includes an elongated tubular shell having a central axis x, opposite nose and tail ends, and a near-elliptical cross-section having a radius R(φ) at substantially every point along the x axis between the two ends, wherein: φ is a roll elevation angle of the shell varying from 0 degrees to +360 degrees about the x axis;R(φ) varies radially by no more than ±7% from a radius r(φ) of a true elliptical cross-section having a radius r(φ), a major axis of dimension 2·rmax and a minor axis of 2·rmin,a curvature Curv(φ) of the shell is defined as the inverse of the local radius of curvature of a surface of the shell and is associated with R(φ), anda curvature κ(φ) of the true ellipse is given by: κ(φ)=[r2+2·(∂r∂φ)2-r·∂2r∂φ2][r2+(∂r∂φ)2]1.5,the method comprising:tailoring at least one structural attribute of the shell as a function of at least one of the elevation angle φ, R(φ) and Curv(φ) so as to reduce the weight of the fuselage relative to an identical fuselage shell in which the same at least one structural attribute has not been so tailored, wherein the tailored function is periodic for φ=0 to 360 degrees and has a period of 360/n degrees,n is an integer, andthe tailored function includes at least two local extrema located within 15 degrees of the values of φ corresponding to a maximum R(φ) located near the major axis of the true elliptical cross-section. 2. The method of claim 1, wherein: the shell comprises a circumferential skin having a thickness; and,the tailoring of the at least one structural attribute comprises tailoring the thickness of the skin as a function of at least one of φ, R(φ) and Curv(φ). 3. The method of claim 2, wherein: the circumferential skin comprises a multi-ply composite structure incorporating at least one of non-metallic and metallic materials;each ply is oriented at a selected angle relative to the other plies; and,the tailoring of the at least one structural attribute further comprises tailoring the plies with respect to at least one of the number of plies, the angular orientation of at least one of the plies, and the material of the plies. 4. The method of claim 1, wherein: the shell comprises a plurality of generally parallel, longitudinally spaced circumferential frames; and,the tailoring of the at least one structural attribute comprises tailoring a radial depth of the frames as a function of at least one of φ, R(φ) and Curv(φ). 5. The method of claim 4, wherein: each circumferential flange comprises at least one of an inner and an outer circumferential flange; and,the tailoring of the at least one structural attribute further comprises tailoring a radial depth of the flange substantially as a function of at least one of φ, R(φ) and Curv(φ). 6. The method of claim 4, wherein: each circumferential frame comprises a radial web; and,the tailoring of the at least one structural attribute further comprises tailoring a thickness of the web as a function of at least one of φ, R(φ) and Curv(φ). 7. The method of claim 6, wherein: the thicknesses of the webs are variable in a radial direction; and,the tailoring of the at least one structural attribute further comprises tailoring the radial distribution of the web thicknesses as a function of at least one of φ, R(φ) and Curv(φ). 8. The method of claim 4, wherein: each circumferential frame comprises a multi-ply composite structure made of at least one of non-metallic and metallic materials;each ply is oriented at a selected angular orientation relative to the other plies; and,the tailoring the at least one structural attribute further comprises tailoring the plies with respect to at least one of the number of plies, the relative angular orientation of the plies, and the material of the plies. 9. An aircraft, comprising: a fuselage, including an elongated internally pressurized tubular shell having a centerline axis, opposite closed nose and tail ends, and a near-elliptical cross-section having a radius R(φ), where φ is an elevation angle defined by an angular coordinate of a cylindrical coordinate system concentric with the centerline axis, a curvature Curv(φ), where Curv(φ) is the inverse of a local radius of curvature of a surface of the shell, and a circumference that varies radially by no more than ±7% from the circumference of a true elliptical cross-section at substantially every position along the centerline axis between the nose and tail ends thereof, wherein:the shell of the fuselage includes at least one structural attribute that has been tailored as a function of at least one of the elevation angle φ, R(φ) and Curv(φ) so as to reduce the weight of the fuselage relative to an identical fuselage shell in which the same at least one structural attribute has not been so tailored,the tailoring function is periodic for φ=0 to 360 degrees, with a period of 360/n degrees,n is an integer, andthe tailored function includes at least two local extrema located within 15 degrees of the values of φ corresponding to a maximum R(φ) located near the major axis of the true elliptical cross-section. 10. The aircraft of claim 9, wherein: the shell comprises a circumferential outer skin and circumferentially spaced longitudinal stringers disposed adjacent to an inner surface of the skin; and,the at least one tailored structural attribute comprises at least one of a cross-sectional shape and size, number, and material of the stringers. 11. The aircraft of claim 10, wherein: each of at least one of the circumferential skin and the stringers comprises a composite of a plurality of plies, each having a selected angular orientation relative to the others; and,the at least one tailored structural attribute comprises at least one of the number, relative angular orientation, and material of the plies. 12. The aircraft of claim 9, wherein: the shell comprises a sandwich structure including a circumferential outer skin attached to a rigid core of at least one of a foam material and a plurality of rigid, interconnected cells; and,the at least one tailored structural attribute comprises at least one of a thickness of the outer skin, a thickness of the core, a core cell density and a core material. 13. The aircraft of claim 9, wherein: the shell comprises an isogrid structure having at least one external face sheet attached to a grid comprising internal stiffening members; and,the at least one tailored structural attribute comprises at least one of grid spacing, grid thickness, grid geometry, grid material, face sheet thickness and face sheet material. 14. The aircraft of claim 9, wherein the shell comprises a filament-wound structure. 15. The aircraft of claim 9, wherein the shell comprises a tape-laid composite structure. 16. The aircraft of claim 9, wherein the shell comprises at least one of an autoclave-cured composite structure, a microwave-cured composite structure and an E-beam cured composite structure. 17. The aircraft of claim 9, wherein the shell includes at least one of a carbon-fiber-in-resin composite structure and a combination of composite and metallic materials. 18. The aircraft of claim 9, wherein the shell includes at least one of stitched multiply composite structure, a stitched resin-film-infused (RFI) composite structure and a stapled multiply composite structure. 19. The aircraft of claim 9, wherein the shell comprises a composite structure including electrically conductive elements for mitigating at least one of electromagnetic effects (EME) and lightning effects acting upon the aircraft. 20. The aircraft of claim 9, wherein the shell comprises a composite structure having an outer surface with a colored, electrically conductive riblet film disposed thereon for providing a decorative color, reduced aerodynamic drag, and mitigation of lightning and electromagnetic effects (EME) acting the aircraft. 21. The aircraft of claim 9, wherein the shell comprises a composite skin having some longitudinally oriented fiber plies having an orientation of zero degrees, plus or minus 20 degrees, relative to a local fuselage surface axis system, and other plies wound circumferentially around the shell and having orientations varying within a range of 90 degrees, plus or minus 20 degrees, relative to the local fuselage surface axis system. 22. The aircraft of claim 21, wherein the shell further comprises first angled plies having orientations varying within a range of +45 degrees, plus or minus 20 degrees, relative to the local fuselage surface axis system, and second angled plies with orientations varying within in a range of −45 degrees, plus or minus 20 degrees, relative the local fuselage surface axis system. 23. The aircraft of claim 22, wherein the angular orientations of the first and second angled plies vary periodically between selected angular values in correspondence with periodically varying pressure-induced circumferential loads in the shell. 24. The aircraft of claim 22, wherein the first and second angled plies are laid down around the shell during its construction along steered paths such that the magnitude of their respective orientations vary relative to 45 degrees for regions of φ wherein longitudinal loads incident on the shell exceed circumferential loads incident on the shell by a selected amount. 25. The aircraft of claim 21, wherein additional longitudinal plies having orientations in a range of zero degrees, plus or minus 20 degrees relative to the local fuselage surface axis system are placed in at least one of a crown and a keel region of the fuselage during its construction for efficiently reacting fuselage bending moments induced by at least one of horizontal tail loads, elevator loads and nose gear slapdown loads incident thereon. 26. The aircraft of claim 9, further comprising at least one additional composite ply layer in a crown region of the shell for reducing a risk of hail damage in the fuselage crown area. 27. The aircraft of claim 9, further comprising at least one additional composite ply layer in a window belt area of upper sides of the shell.
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