A wing panel structure for an aerospace vehicle or the like may include an outer layer of material having a predetermined thickness. A core structure may be placed on at least a portion of the outer layer of material. An inner layer of material may be placed at least on the core structure. The inner
A wing panel structure for an aerospace vehicle or the like may include an outer layer of material having a predetermined thickness. A core structure may be placed on at least a portion of the outer layer of material. An inner layer of material may be placed at least on the core structure. The inner layer of material may have a selected thickness less than the predetermined thickness of the outer layer of material.
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What is claimed is: 1. A panel structure for an aerospace vehicle, comprising: a first portion of the panel structure; a second portion of the panel structure; an outer layer of material having a predetermined thickness and extending across both the first portion and the second portion of the panel
What is claimed is: 1. A panel structure for an aerospace vehicle, comprising: a first portion of the panel structure; a second portion of the panel structure; an outer layer of material having a predetermined thickness and extending across both the first portion and the second portion of the panel structure; a core structure placed on the outer layer of material and included in the first portion of the panel structure; an inner layer of material formed at least on the core structure, wherein the inner layer of material has a selected thickness less than the predetermined thickness of the outer layer of material; a stiffener formed on the outer layer of material and included in the second portion of the panel structure, wherein the stiffener has a predetermined structural shape comprising a flange formed on the outer layer and a member extending from the flange on an opposite side from the outer layer; and a support rib formed on the inner layer of material between the stiffener and an assembly including the core structure and the inner layer of material, wherein the inner layer of material extends under the support rib and overlaps a portion of the flange of the stiffener. 2. The panel structure of claim 1, wherein the outer layer of material comprises a structure to predominantly support a load. 3. The panel structure of claim 1, wherein the outer layer of material comprises a multiplicity of plies of material. 4. The panel structure of claim 3, wherein the multiplicity of plies of material are cured and processed to a higher strength specification than the core structure and the inner layer of material, wherein the higher strength specification comprises a curing temperature above about 300 degrees F. and a pressure above about 80 psi. 5. The panel structure of claim 3, wherein the multiplicity of plies of material comprise a multiplicity of epoxy unidirectional tape plies. 6. The panel structure of claim 3, wherein the plies of material are continuous for an extent of the panel. 7. The panel structure of claim 1, further comprising a layer of a non-destructive inspection (NDI) reflective material formed between the outer layer of material and the core structure. 8. The panel structure of claim 1, wherein the core structure comprises a honeycomb type structure. 9. The panel structure of claim 1, wherein the outer layer, the core structure and the inner layer are cured at a curing temperature between about 300 and about 400 degrees F. and a pressure between about 80 and about 100 psi. 10. The panel structure of claim 1, wherein the inner layer of material comprises a plurality of plies of a fabric. 11. The panel structure of claim 1, wherein the stiffener includes a group comprising an I section stiffener and a T section stiffener. 12. The panel structure of claim 1, wherein the stiffener is inboard of the assembly including the core structure and the inner layer of material. 13. A panel structure for an aerospace vehicle, comprising: a first portion of the panel structure; a second portion of the panel structure; an outer layer of material having a predetermined thickness and extending across both the first portion and the second portion of the panel structure; a core structure placed on the outer layer of material and included in the first portion of the panel structure; an inner layer of material formed at least on the core structure; a stiffener placed on the outer layer and included in the second portion of the panel structure, wherein the stiffener has a predetermined structural shape comprising a flange formed on the outer layer and a member extending from the flange on an opposite side from the outer layer; and a support rib placed on the inner layer of material between the stiffener and an assembly including the core structure and the inner layer of material, wherein the inner layer of material extends under the support rib and overlaps a portion of the flange of the stiffener. 14. The panel structure of claim 13, wherein the outer layer of material comprises a structure to predominantly support a load. 15. The panel structure of claim 13, wherein the outer layer of material comprises a multiplicity of plies of material. 16. The panel structure of claim 15, wherein the multiplicity of plies of material are cured and processed to a higher strength specification than the core structure and inner layer of material, wherein the higher strength specification comprises a curing temperature above about 300 degrees F. and a pressure above about 80 psi. 17. The panel structure of claim 13, further comprising a layer of a non-destructive inspection (NDI) reflective material disposed between the outer layer of material and the core structure. 18. The panel structure of claim 13, wherein the core structure comprises a honeycomb type structure. 19. The panel structure of claim 13, wherein the stiffener comprises a stringer of composite material. 20. The panel structure of claim 13, wherein the stiffener is inboard of the assembly including the core structure and the inner layer of material. 21. An aerospace vehicle, comprising: a fuselage; and an airfoil extending from the fuselage, wherein the airfoil includes at least one panel structure, the at least one panel structure including: first portion of the panel structure; second portion of the panel structure; an outer layer of material having a predetermined thickness and extending across both the first portion and the second portion of the panel structure; a core structure placed on the outer layer of material and included in the first portion of the wing panel structure; an inner layer of material formed at least on the core structure, wherein the inner layer of material has a selected thickness less than the predetermined thickness of the outer layer of material; and a stiffener formed on the outer layer of material and included in the second portion of the wing panel structure, wherein the stiffener has a predetermined structural shape comprising a flange formed on the outer layer and a member extending from the flange on an opposite side from the outer layer; and a support rib formed on the inner layer of material between the stiffener and an assembly including the core structure and the inner layer of material, wherein the inner layer of material extends under the support rib and overlaps a portion of the flange of the stiffener. 22. The aerospace vehicle of claim 21, wherein the outer layer of material of the at least one panel structure comprises a structure to predominantly support a load. 23. The aerospace vehicle of claim 21, wherein the outer layer of material of the at least one panel structure comprises a multiplicity of plies of material and wherein the multiplicity of plies of material are cured and processed to a higher strength specification than the core structure, wherein the higher strength specification comprises a curing temperature above about 300 degrees F. and a pressure above about 80 psi. 24. The aerospace vehicle of claim 21, further comprising a layer of a non-destructive inspection (NDI) material disposed between the outer layer of material and the core structure of the at least one panel structure. 25. The aerospace vehicle of claim 21, wherein the core structure of the at least one panel structure comprises a honeycomb type structure. 26. A method of making a panel structure for an aerospace vehicle, comprising: forming an first portion of the panel structure; forming a second portion of the panel structure; forming an outer layer of material having a predetermined thickness and extending across both the first portion and the second portion of the panel structure; placing a core structure on at least a portion of the outer layer of material to form the first portion of the panel structure; forming an inner layer of material disposed at least on the core structure, wherein the inner layer of material has a selected thickness less than the predetermined thickness of the outer layer of material; disposing a stiffener on the outer layer of material to form the second portion of the panel structure, wherein the stiffener has a predetermined structural shape comprising a flange formed on the outer layer and a member extending from the flange on an opposite side from the outer layer; and forming a support rib on the inner layer of material between the stiffener and an assembly including the core structure and the inner layer of material, wherein the inner layer of material extends under the support rib and overlaps a portion of the flange of the stiffener. 27. The method of claim 26, wherein forming the outer layer of material comprises forming a structure to predominantly support a load. 28. The method of claim 26, wherein forming the outer layer of material comprises: depositing a multiplicity of plies of material; curing and processing the multiplicity of plies of material to a higher strength specification than the core structure and inner layer of material, wherein the higher strength specification comprises a curing temperature above about 300 degrees F. and a pressure above about 80 psi. 29. The method of claim 28, wherein the multiplicity of plies of material of the outer layer of material are cured and processed before the core structure and inner layer of material are disposed on the panel structure. 30. The method of claim 26, further comprising forming a layer of NDI reflective material between the outer layer of material and the core structure. 31. The method of claim 26, wherein placing the core structure comprises placing a honeycomb type structure. 32. The method of claim 26, wherein forming the inner layer of material comprises laying a plurality of plies of a fabric. 33. The method of claim 26, further comprising curing the panel structure after forming the inner layer of material. 34. The method of claim 33, wherein curing the panel structure comprises applying a temperature between about 300 and about 400 degrees F. and a pressure between about 80 and about 100 psi.
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.
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