CPC B64C 1/40 (2013.01) [B29C 66/727 (2013.01); B29C 66/72525 (2013.01); B32B 3/12 (2013.01); B32B 5/12 (2013.01); B32B 5/18 (2013.01); B32B 7/12 (2013.01); B32B 15/046 (2013.01); B32B 15/14 (2013.01); B32B 15/20 (2013.01); B32B 37/12 (2013.01); B32B 37/146 (2013.01); B32B 38/0004 (2013.01); G10K 11/168 (2013.01); B29C 44/186 (2013.01); B29C 66/72523 (2013.01); B32B 2260/023 (2013.01); B32B 2260/046 (2013.01); B32B 2262/106 (2013.01); B32B 2266/0214 (2013.01); B32B 2305/022 (2013.01); B32B 2305/024 (2013.01); B32B 2307/102 (2013.01); B32B 2307/50 (2013.01); B32B 2605/18 (2013.01); B64C 2001/0072 (2013.01)] | 20 Claims |
1. An aircraft acoustic structural panel (10) made by the method comprising:
compressing together under pressure and heat a stack including a preformed uniform foam layer (3), a layer of adhesive (2), and a core honeycomb (1) having a plurality of honeycomb cells (14) and thereby pressing said preformed uniform foam layer (3) substantially flush atop said core honeycomb (1) to form corresponding foam plugs from said preformed uniform foam layer (3) being thermally bonded inside said honeycomb cells (14) by said adhesive (2), with said preformed uniform foam layer (3) being initially thicker than said core honeycomb (1) and fully compressed into said honeycomb cells (14) to preform a core honeycomb laminate (12);
stacking said preformed core honeycomb laminate (12) between opposite top and bottom structural outer laminates (16, 18) having corresponding outer skins (20, 22); and
compressing together under heat and pressure said stacked preformed core honeycomb laminate (12) and said structural outer laminates (16, 18) into a unitary structural panel (10) defining said aircraft acoustic structural panel having said preformed core honeycomb laminate (12) integrally bonded between said outer skins (20, 22).
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