Stock material or miscellaneous articles – Structurally defined web or sheet – Honeycomb-like
Reexamination Certificate
2000-11-30
2002-11-05
Cole, Elizabeth M. (Department: 1771)
Stock material or miscellaneous articles
Structurally defined web or sheet
Honeycomb-like
C428S119000, C428S219000, C428S208000, C428S220000, C442S179000, C442S195000
Reexamination Certificate
active
06475596
ABSTRACT:
FIELD OF THE INVENTION
The invention relates to composite sandwich structures, preferably honeycomb core, composite sandwich structures, and to fabric and prepreg components for such composite structures. More particularly, the invention relates to core crush resistant, honeycomb core composite sandwich structures, particularly those composite structures used in the aerospace industry, and to fabrics and prepregs for making such composite sandwich structures.
BACKGROUND OF THE INVENTION
Honeycomb core composite sandwich structures find widespread use in the aerospace industry as panel components in various aerospace structures. The honeycomb core composites are formed from a lay-up of prepreg skin plies encompassing a honeycomb core, the latter typically having beveled edges. The prepreg plies may be fabrics, tapes, or non-wovens that have been pre-impregnated with a thermosetting, thermoplastic or other polymeric resin. The fabrics used to form the prepregs are woven fabrics, formed primarily or entirely of high modulus, reinforcing fibers in the form of continuous filament tows. Curing of the lay-up is carried out in a high temperature, high pressure environment, typically in an autoclave.
The technical requirements of aerospace end uses generally dictate that the prepregs and prepreg components meet a rigid set of chemical, physical, and mechanical specifications including overall prepreg basis weight, fiber modulus, and resin flow rate. The basis weight of the prepreg and the high strength properties of the fibers and the resin, in combination with the strength properties of the honeycomb core component, impart high strength-to-weight, and high stiffness-to-weight ratios to the final composite structure. In addition, the flow rate characteristics of the resin and the high pressures used to cure the composite, minimize porosity, i.e., the inclusion of voids and through holes, that might impair strength, the desired impervious nature, and/or surface smoothness of the final honeycomb panel sandwich structure.
Even though honeycomb core composite panels have long been used in the aerospace industry, manufacture of these structures is still plagued by high reject scrap levels, generating substantial quantities of unusable scrap and impacting negatively on manufacturing economics. Partial collapse of the honeycomb core during curing of the composite, known in the industry as “core crush”, is a particularly common reason for rejection of cured panels. Core crush is typically observed in the beveled edge or chamfer region of the honeycomb structural part.
Substantial effort and research extending over many years have been directed to the core crush problem. For example, U.S. Pat. No. 5,685,940 to Hopkins discloses an improved tiedown method to produce or prevent core crush and ply wrinkling in honeycomb sandwich structures. A scrim-supported barrier film is placed between the fiber-reinforced resin composite laminates and honeycomb core to prevent resin flow from the prepreg into the honeycomb core. A tiedown ply between the core and the barrier film is used to reduce slippage of the barrier film relative to the core during curing. In addition, a film adhesive having a curing temperature lower than that of the laminate resin is placed between the tiedown plies just outside the net trim line. During the curing process, cured film adhesive bonds the tiedown plies to one another before the curing of the prepreg laminates, thus strengthening the tiedown and reducing core crush. The Hopkins patent also discusses other methods and structural modifications which have been proposed for minimizing or eliminating the core crush. Nevertheless, core crush remains a significant problem in the industry.
SUMMARY OF THE INVENTION
The present invention provides a core crush resistant prepreg for use in making a fiber reinforced composite sandwich structure. Use of the prepreg of this invention, can significantly reduce the degree of core crush as compared to conventional structures.
In accordance with a first aspect of the invention, it has been found that the core crush problem associated with honeycomb core composite sandwich structures can be significantly reduced by controlling construction of the fabric used to prepare the prepreg. In particular, it has been found that core crush can be substantially reduced by controlling the cross sectional aspect ratio of the carbon fiber tow in the prepreg, the average thickness of the prepreg and the openness of the prepreg, as measured by visual inspection. In particular, prepregs according to the invention are made from fabrics having an areal weight range of from about 150 to 400 grams per square meter. The prepreg has an average tow aspect ratio of less than about 15.5, an average prepreg thickness of at least about 0.245 mm, and/or an openness of at least about 1.2% but less than about 10.0%.
While not wishing to be bound by theory, it is believed that average tow aspect ratio, prepreg openness and prepreg thickness determine the frictional force between prepreg plies during the curing step in the manufacture of honeycomb core composite sandwich structures. When prepreg properties of tow aspect ratio, prepreg thickness and prepreg openness are maintained within the ranges set forth above, sufficient frictional force is provided between prepreg plies such that the innermost prepreg plies, adjacent the honeycomb core, are restrained from slipping during the curing process to thereby eliminate or minimize core crush.
It has also been found, according to the present invention, that when the tow aspect ratio, prepreg thickness and prepreg openness are optimized to minimize core crush, the porosity of the final honeycomb core composite sandwich structure can be unacceptable. The porosity problems can be especially prevalent in thicker composite sandwich structures and especially when a low flow resin system is used to impregnate the prepreg. The present invention employs a hardenable polymeric resin composition having a flow rate higher than the flow rate of resins traditionally used in commercial practice in the aerospace industry in prepregs for honeycomb core composite sandwich structures in order to maintain acceptable porosity in the final composite structure. Therefore, prepregs according to the present invention are impregnated with a hardenable polymeric resin composition having rheology which is predominately viscous in nature, such that the ratio of viscous to elastic components of the viscosity, i.e., tan &dgr;, is within the following defined ranges.
Prior to significant resin cross-linking or curing, the resin composition used in this invention preferably has a tan &dgr; of between about 1.2 and about 2.0, preferably between about 1.5 and about 1.8, more preferably about 1.35, at 70° C.; or, a tan &dgr; of between about 0.7 and about 2.0, preferably between about 0.9 and about 1.8, more preferably about 1.35, at 100° C.; or, a tan &dgr; of between about 0.5 and about 1.7, preferably between about 0.7 and about 1.5, more preferably about 1.35, at 140° C.
Preferably, the tan &dgr; of the resin composition is from about 0.5 to about 2.0, more preferably between about 1.0 and about 1.8, most preferably about 1.35, throughout the elevated temperature range of from about 70° C. to about 140° C., or if the minimum viscosity temperature is below 140° C., the range of from about 70° C. to the minimum viscosity temperature.
More preferably, prior to significant resin cross-linking or curing, the resin composition has a tan &dgr; of between about 1.0 and about 2.0, more preferably between about 1.2 and about 1.8 at about 70° C.; between about 0.7 and about 2.0, more preferably between about 1.0 and about 1.7 at about 100° C.; and, between about 0.5 and about 2.0, more preferably between about 0.6 and about 1.7 at about 140° C., or at the minimum viscosity temperature, if the minimum viscosity temperature is below 140° C.
Preferably, the resin composition comprises an epoxy resin and has an average epoxy functionality of greater than 2.0.
According to a
Buyny Robert Albert
Hsiao Hao-Ming
Lee Shaw Ming
Martin Cary Joseph
Alston & Bird LLP
Cole Elizabeth M.
Hexcel Corporation
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