WO1997025198A1

WO1997025198A1 - Composite honeycomb sandwich structure

Info

Publication number: WO1997025198A1
Application number: PCT/US1997/000075
Authority: WO
Inventors: Dale E. Hartz; William B. Hopkins; Christopher L. Pederson; Dave G. Erickson; Darrell H. Corbett; Stuart A. Smith
Original assignee: The Boeing Company
Priority date: 1996-01-11
Filing date: 1997-01-06
Publication date: 1997-07-17
Also published as: CN1101751C; JP3913275B2; JP2007015385A; EP0883484A1; JP4407964B2; CN1211947A; JP2000502968A; AU1822997A

Abstract

We eliminate resin flow into the cells of honeycomb in sandwich structure by using an unsupported film adhesive (108), a barrier layer (110), and a scrim supported adhesive layer (112) between the composite laminate (102) and the core (106). We produce superior panels with lighter weights, improved mechanical properties, and more predictable structural performance by keeping resin in the laminate rather than losing it to the core cells. We reduce core crush and ply wrinkling in composite honeycomb sandwich structure by preventing slipping of tiedown plies relative to the mandrel and to one another during autoclave curing. We produce superior panels with lighter weights, improved mechanical properties, and more predictable structural performance. The method involves applying a film adhesive to the tiedown plies in the margin of the part outside the net trim line. During heating of the autoclave and prior to the application of high pressure to the composite structure, the film adhesive cures to form a strong bond between the plies and to the mandrel. When pressure is applied, the tiedown plies are locked together and to the mandrel to prevent slippage between any layers in the panel.

Description

COMPOSITE HONEYCOMB SAND WICH STR UCTURE

Technical Field

The present invention relates to composite honeycomb sandwich structure, and particularly to resin impregnated fabric sheets forming outer skins adhered on opposed surfaces of a honeycomb core with an intermediate barrier to eliminate resin flow from the skins to the core.

Background Art

Aerospace honeycomb core sandwich panels (having composite laminate skins cocured with adhesives to the core through autoclave processing) find widespread use today because of the high stiffness-to-weight (i.e., "specific stiffness) and strength-to-weight (i.e., specific strength) ratios the panels afford. Typical honeycomb core sandwich panels are described in U.S. Patents 5,284,702; 4,622,091; and 4,353,947. Alteneder et al., Processing and Characterization Studies of Honeycomb Composite Structures, 38th Int'l SAMPE Symposium, May 10-13, 1993 (PCL Internal No. 200-01/93- A WA) discusses common problems with these panels, including core collapse (i.e., core crush), skin laminate porosity, and poor tool surface finish.

U.S. Patent 5,445,861 describes composite sandwich structure for sound absorption (acoustic insulation) and other applications. The sandwich structures have seven layers as follows:

(1) an outer skin;

(2) a small celled honeycomb or foam core;

(3) a frontside inner septum;

(4) a large celled middle honeycomb core;

(5) a backside, inner septum;

(6) a backside, small celled honeycomb or foam core; and (7) an inner skin. Tuned cavity absorbers in the middle honeycomb core absorb sound. Performance of this structure suffers from resin flow to the cells of the honeycomb cores during fabrication for the reasons already discussed and because such flow alters the resonance ofthe structure.

Summary of the Invention

With a high flow resin system, large amounts of resin can flow into the core during the autoclave processing cycle. Such flow robs resin from the laminate, introduces a weight penalty in the panel to achieve the desired performance, and forces over design of the laminate plies to account for the flow losses. The resin loss from the laminate plies also reduces the thickness of the cured plies which compromises the mechanical performance. To achieve the desired performance and the corresponding laminate thickness, additional plies are necessary with resulting cost and weight penalties. Because the weight penalty is severe in terms of the impact on vehicle performance and cost in modern aircraft and because the flow is a relatively unpredictable and uncontrolled process, aerospace design and manufacture dictates that flow into the core be eliminated or significantly reduced. In addition to the weight penalty from resin flow to the core, we discovered that microcracking that originated in the migrated resin could propagate to the bond line and degrade mechanical performance. Such microcracking potential poses a catastrophic threat to the integrity of the panel and dictates that flow be eliminated or, at least, controlled.

Flow from the laminates to the core occurs because of viscosity reduction of the resin (i.e., thinning) at the elevated processing temperatures. Therefore, prior art attempts to solve the flow problem have generally focused on retaining the ambient temperature viscosity of the resin at the curing temperatures. For example, one might alter the processing cycle to initiate curing ofthe resin during a slow heat-up, low pressure step to induce resin chain growth before high temperature, high pressure completion. In this staged cure cycle, one would try to retain the resin's viscosity by building molecular weight at low temperatures. Higher molecular weight resins have inherently higher viscosity so they remain thicker and are resistant to damaging flow to the core. Unfortunately, with a staged cure cycle, too much flow still occurs, and the potential problems of microcracking still abound. Also, facesheet porosity might increase beyond acceptable limits. Furthermore, a modified cure cycle increases autoclave processing time. Increased processing time translates to a significant fabrication cost increase with risk of rejection of high value parts at the mercy of uncontrolled and inadequately understood factors.

We eliminate resin (matrix) flow into d e honeycomb core for sandwich structure using high flow resin systems and results in reproducibility and predictability in sandwich panel fabrication and confidence in the structural performance ofthe resulting panel. We use a scrim-supported barrier film between the fiber-reinforced resin composite laminates and the honeycomb core. This sandwich structure is lighter for die same performance characteristics man prior art panels because the resin remains in the larninate (skin) where it provides structural strength ratf er than flowing to the core where it is worthless, introducing excess weight and potential panel failure. We also generally use an unsupported film adhesive between the barrier film and the laminates to bond the laminates to the barrier film. With these layers (which might be combined into one product), they achieved improved performance, retained d e resin in d e laminates and thereby reduced excess resin that designers otherwise needed to design into the panels to account for resin flow into d e core, and reliably fabricated panels in which tiiey had structural confidence.

Core crush frequently occurred in the chamfer region of honeycomb core when we cured a panel having a scrim-supported barrier film, particularly when d ey tried to use lighter weight core materials. We can reduce core crush in diese panels by including a tiedown ply in contact with die core beneath the barrier film (and adhesive) because die tiedown ply reduced slippage of die barrier film relative to die core during curing.

Controlling core slippage in d e present invention allows us to use lighter density honeycomb core to produce structures widiout costly scrap due to core crush. We reduce manufacturing costs botii by saving time, materials, and rework/scrap and by improving die reliability of die manufacturing process to produce aerospace-quality panels having d e highest specific strength and specific stiffness.

The added tiedown ply means mat three or more tiedown plys will be included in d e final preform of the panel. In conventional practice, tiiere will also be tiedown plys on die outer surfaces of die panel and possibly between d e laminate and die adhesive barrier film. Each tiedown ply extends outwardly from the part beyond die net trim line of die finished product. Conventionally, die tiedown plies are secured individually and sequentially to die layup mandrel witii tape. Especially when using low density core it is important to fix the relation of the plies to one another and to the mandrel. Failure of die tape results in facesheet ply wrinkles or core crush. Core crush could still occasionally occur when die tiedown ply in contact witii d e core pulled away from d e tape securing it to the mandrel, slipping relative to die other tiedown plies. The adhering strengtii of die tape alone was insufficient to overcome die forces acting on the core in a panel when we applied autoclave pressure. We discovered how to adhere the tiedown plies to each otiier reliably, easily, and inexpensively. Adhering the plies to each other distributes die forces acting on any individual ply among all the tiedown plies, reducing die maximum force seen by d e tape adhering the tiedown plies to die mandrel. While described with respect to a composite honeycomb sandwich structure, d e adhering method is generally applicable to all applications involving tiedown plies in composite construction.

Thus, in one aspect, die present invention relates to an improvement in the manufacture of composite structure, especially composite honeycomb sandwich structure, where tiedown plys are used to secure the part during autoclave curing at elevated temperature and pressure. To lock the tiedown plies togedier so tiiat tiiere is no movement of one ply relative to another, we use a lower temperature curing adhesive to cure and to connect the several plys together during the early stages of autoclave curing prior to applying pressure. We apply the adhesive outside die net trim line for d e part, so that it is removed during finishing of the part.

In another aspect, the present invention relates to adhering tiedown plys to one anod er during the construction of composite structure, especially during the autoclave processing at elevated temperature and pressure of composite honeycomb sandwich structure. The conventional practice of taping d e tiedown plys to the mandrel alone is unsatisfactory, because the taping must be sufficient to prevent slippage of any ply or of one ply relative to another. We discovered tiiat we could adhere the plies effectively to one another to reduce maximum forces on d e tape by applying a low temperature curing film adhesive between die tiedown plys just outside the net trim line for the part. In die autoclave, this film adhesive melts and cures at a lower temperature tiian die resin in the laminates so tiiat it bonds die tiedown plies together prior to increasing the autoclave pressure at d e higher temperature where d e laminate resin flows and cures. The film adhesive eliminates movement of die tiedown plys relative to one another. In our preferred embodiment for a bismaleimide (BMI) sandwich panel, we prefer to use an adhesive that cures at about 250°F (121°C) for a BMI that cures around 375°F (191°C), and post-cures around 440 °F. Brief Description of the Drawings.

Fig. 1 illustrates a typical composite honeycomb sandwich structure.

Fig. 2 is a schematic, partial sectional view of die skin-core interface in sandwich structure having a scrim-supported barrier film to prevent resin flow from the skin to the core.

Fig. 3 is a schematic, partial sectional view of prior art honeycomb sandwich structure, suffering resin flow to d e core, using a supported film adhesive witi out a barrier film.

Fig. 4 is another schematic, partial sectional view showing sandwich structure with resin depletion in die skin, but where d e resin is prevented from reaching die core witii a bulging, unsupported barrier film.

Fig. 5 is a schematic, sectional elevation showing core crush of a honeycomb sandwich panel caused by core and barrier film slippage.

Fig. 6 is anotiier schematic, sectional elevation showing the use of a tiedown ply to reduce core crush.

Fig. 7 is a graph of a typical autoclave curing cycle for making composite honeycomb sandwich panels, showing that our tiedown adhesive cures prior to the application of high pressure in the cycle.

Detailed Description of a Preferred Embodiment

As a frame of reference for d is description, we will initially describe typical composite honeycomb sandwich structure. Then, we will turn to our invention of a method of reliably adhering the tiedown plies together.

A composite honeycomb sandwich panel minimizes, eliminates, or significantly reduces resin flow from the laminates to die core, thereby permitting a simpler processing cycle tiiat is more robust for die manufacture of aerospace structure. Such a sandwich panel 100 (Fig. 1) generally has outer facesheets or skins 102 adhered to a central honeycomb core 106. The finished skins 102 comprise laminates of layers of fiber-reinforced organic matrix resin in a cured and consolidated composite form. The core 106 can be paper, synthetic paper, metal, composite, or the like, as appropriate for the application. In panels of die present invention, we obtain higher specific strengths and higher specific stiffnesses because we reduce core crush during autoclave curing by incorporating at least one tiedown ply between d e core 106 and skin 102 to reduce damaging slippage between die core and skin tiiat otherwise often occurs.

To prevent flow of resin from the composite laminate skin to the core, we use an unsupported film adhesive 108 (Fig. 2), a barrier film 1 10, and a scrim-supported film adhesive 112 between d e skin 102 and the core 106 to keep resin out of me cells 114 of the core 106.

Fig. 3 illustrates the core-filling problems that can result when a film adhesive 112 is used alone without the barrier film 110 and film adhesive 108. Cells 114 of die honeycomb fill with resin 118 which migrates from the laminates and which thereby depletes d e resin in the skin 102. Resin depletion impacts structural performance because it reduces ply thickness. Resin depletion increases total weight since the cell resin 118 is simply waste. In all cases, uncontrolled resin flow and depletion makes d e panel suspect, especially to microcracking that can begin in the cell resin 1 18 during thermal cycling and migrate to the fiber-reinforced skin 102, especially at the bond line between the skin 102 and core 106.

Fig. 4 illustrates undesirable bulging tiiat can occur if a barrier film 110 is used witiiout a scrim-supported film adhesive 112 to try to eliminate cell resin 118. Here, a waste resin bulge 120 protrudes downwardly into the cells 114 of the honeycomb core 106. While d e resin is contained in the bulge 120, the skin 102 is still depleted in resin. The flow of resin to bulge 120 imposes structural performance and weight penalties comparable to the uncontrolled condition illustrated in Fig. 3. As shown in Fig. 2 witii the film adhesive 108, barrier fihn 110, and scrim-supported film adhesive 12, resin flow is checked witiiout cell resin 118 or resin bulges 120. We discovered, however, tiiat the barrier film produced a slip plane between the laminate skins and d e core which often resulted in core crush during the autoclave processing cycle. In 22 of 31 test panels, in fact, we experienced core crush in our initial trials. This rate of failure was unacceptable from a cost and schedule perspective. Our tiedown plys in the chamfer region reduce the frequency of or eliminate damaging core slippage and d e core crush attributable to such slippage.

For bismaleimide laminated skins made with RIGIDITE® 5250-4-W- IM7-GP-CSW, RIGIDITE® 5250-4- W-IM7-GP-CSX, and RIGIDITE® 5250- 4-W-IM7-GP-PW prepreg from Cytec Engineered Materials, Inc. (Cytec), d e film adhesive 108 preferably is 0.015 psf METLBOND® 2550U adhesive, also available from Cytec. The film adhesive provides additional resin to promote a quality bond between the laminate and barrier film 1 10. The barrier film 110 preferably is a 0.001 inch thick, bondable grade, surface treated KAPTON® polyimide barrier film capable of withstanding d e cure cycle to provide a resin impermeable membrane between d e skin 102 and core 106. The scrim preferably is fiberglass, "Style 104" fiber cloth and the film adhesive 1 12 is 0.06 psf METLBOND® 2550G adhesive, available from Cytec. The scrim- supported film adhesive prevents the barrier film from bulging into the core cells, thereby retaining the resin in die laminate (i.e., skin layers) so that the cured ply thickness is maximized and thereby, we achieve maximum performance at minimum weight for the panels.

The film adhesive 108, barrier film 110, and film adhesive 112 can be purchased as a single item from Cytec as METLBOND® 2550B- 082 36".

The plys of the skin 102 typically are prepregs of carbon fiber impregnated with bismaleimide tiiermoset resin, although the present invention applies to other resin systems. Tows might be used in place of the prepreg. The film adhesive 108 should be tailored to achieve an adequate bond between the skin 102 and barrier film 1 10. The honeycomb core generally is HRP Fiberglass Reinforced Phenolic honeycomb available from Hexcel.

The supported film adhesive and barrier film layers in the sandwich structure also function as corrosion barriers between the skin 102 and core 106 in the case where the core is metal, such as aluminum, and d e skin includes a galvanically dissimilar material, such as carbon fiber.

Additional information concerning preferred panels is presented in die technical paper: Hartz et al., "Development of a Bismaleimade/Carbon Honeycomb Sandwich Structure," SAMPE, March, 1996, which we incorporate by reference. This paper describes botii the Hartz et al. barrier film improvement, the tiedown ply method, and the adhering method ofthe present invention.

The Hartz-t pe panels provide mechanical and physical edgeband properties equivalent to solid BMI/carbon laminate (cured at 0.59 MPa (85 psig)). Our tests confirm that in our panels the edgeband cured-ply-thickness is equivalent to a solid laminate and that the edgeband 160 (Figs. 5 & 6) met the requirements of d e solid laminate nondestructive inspection specification. The edgeband and facesheet mechanical performance improved over results we achieved with sandwich structure lacking the scrim-supported adhesive, barrier film, adhesive combination. The flatwise tensile mechanical performance also met design requirements.

Preconditioning the core to eliminate volatile evolution during curing by heating the core to about 235° C (455° F), prior to laying up the sandwich panel, especially for phenolic core, eliminates core-laminate disbonding otiierwise caused by outgassing from the core.

Core crush 200 (Fig. 5) occurs in the chamfer region 155 when the barrier film 1 10 and core 106 slip relative to the facesheets 102 when autoclave pressure is applied and when the resin is melted. As shown in Fig. 5, the barrier films 100 and core 106 have moved toward the right to compress the core in the chamfer region 155 to produce die core crush 200. The skin 102 has sagged in the edgeband region 160 where d e core moved away.

Referring now to Fig. 6, die improved honeycomb sandwich panel includes at least one tiedown ply 150 in contact with the core 106 along a chamfer 155. Such a chamfer (i.e. an angled transition in the core, often at the edgeband 160) typically occurs around the periphery ofthe panel, but it might also occur intermediate of the panel at join lines or hard points where fasteners or pass-throughs might be necessary in the assembled structure.

Typically we use a single ply 150 of carbon fiber or fiberglass fabric witii a conventional 0/90 fiber orientation in the fabrication of bismaleimide panels having 5 or 8 Ib/ft^ HRP core, like Hartz et al. describe. The tiedown ply 150 functions to prohibit or to limit slippage ofthe skin relative to the core so as to reduce core crush otherwise attributable to the slippage. The tiedown ply 150 anchors the core with the inherent roughness of the fabric when the preform is heated during the autoclave processing cycle and me matrix resin softens, melts, and, for high flow resins, essentially liquefies. Witii these panels, we can save between 2.5-4 lb/ft^ of core because we can use lighter density honeycomb core without suffering core crush. For a fighter, this change can save as much as 25 lbs per vehicle.

As shown in Fig. 6, the tiedown ply 150 is a narrow, peripheral strip tiiat contacts the core 106 along at least a portion of the chamfer 155 for about 1 inch overlap with the core 106 and extends outward into the edgeband 160 beyond d e trimline 165 of the part. The tiedown ply 150 might be on either the flat side of the chamfer or the angled surface (which is how we show it in Fig. 6). The key factor is that the tiedown ply 150 contact die core beneath the adhesive and barrier film 110 which is used to bond d e laminate skin to the core. The tiedown ply 150 is cutaway everywhere in the body of die part otiier than a narrow peripheral area in the chamfer region, and forms a peripheral frame around d e edge of the panel. In this way, the tiedown ply 150 allows an adhesive interface between the core 106 and die skins 102 in the panel region.

Traditionally, in making a Hartz-type panel, we use four complete cover sheet tiedown plies 175 in an effort to anchor the layers and die core, and we show all these plies in Fig. 6. These traditional plies 175 were commonly used in sandwich panel fabrication prior to introducing the Hartz-type barrier film, and we commonly use them all, although we believe we can now eliminate all but the outer plies and the peripheral, core contacting tiedown ply 150. That is, we would use tiiree total plies rather than five, as Fig. 6 shows.

The tiedown plies 150 and 175 extend through die edgeband 160 beyond the net trim line 165 to anchoring points that we tape to the layup mandrel. To further prevent slippage ofthe tiedown plies, we have incorporated a low curing (i.e. 121°C for BMI panels) film adhesive 180 between die tiedown plies just outside d e net trim line of the part. The film adhesive 180 eliminates movement of one ply relative to the others when we apply pressure during the autoclave curing cycle. Curing at a temperature of about 100 - 150°F below the curing temperature ofthe laminate resin, the tiedown adhesive cures before we need to increase d e autoclave pressure and die cured adhesive bonds the tiedown plys to one another. Using the adhering method eliminates relative movement of die plys and eliminates facesheet wrinkles and core crush that otherwise can occur.

The tiedown method saves material, reduces cost, and saves weight, because it use die "picture frame" peripheral tiedown ply 150 (with the traditional, internal sheets omitted). The normal tiedown procedure entails plys on the outer surfaces ofthe skins and internally between the skin and underlying adhesive (Fig. 5). A traditional tiedown system will fail without the "picture frame" ply because the barrier film 110 permits the core to slip. The Corbett and Smith method will fail occasionally witiiout the adhering method of die present invention. For lightweight core (i.e. 5-8 lb/ft^) with the bismaleimide prepreg and adhesive system previously described, we hold die chamfer angle to 20° ± 2°.

By "chamfer" we mean an angled, cut region (a ramp) of the honeycomb core tapering from full thickness to no thickness with a steady slope. A chamfer is used at the edge band of a composite honeycomb sandwich panel to provide a smooth transition between the structural body ofthe panel that has the embedded honeycomb and a connecting edge band lacking any honeycomb core. The method of d e present invention allows us to use much steeper chamfer angles than traditional practices often require if one is to avoid core crush without one tiedown ply. While we prefer a 20° chamfer, we believe that we could increase die angle to whatever angle suited the panel design requirements.

By "autoclave processing" we mean the cycle of elevated temperature and pressure applied to the panel to consolidate and cure resin in die laminate while bonding or otherwise adhering die cured laminate to the honeycomb core. Our preferred cycle is illustrated in Fig. 7. Our adhesive for the tiedown plies cures at about 250°F (121°C) so it cures prior to the increase in autoclave pressure that can introduce relative motion between layers in the panel.

If core crush occurs, the damage to the panel is generally so extensive that repair is impossible so the part is scrapped. The cost of today's advanced composite resins and reinforcing fibers requires a process that virtually eliminates core crush. Otherwise, the processing costs are prohibitive. With panels being designed as close to d e design edge as possible, core crush is a significant issue. The mediod of die present invention reduces cores crush and ply movement or wrinkling.

While we have described preferred embodiments, those skilled in the art will readily recognize alterations, variations, and modifications, which might be made without departing from the inventive concept. Therefore, interpret the claims liberally with the support of die full range of equivalents known to those of ordinary skill based upon this description. The examples are given to illustrate the invention and are not intended to limit it. Accordingly, defme die invention by the claims and limit die claims only as necessary in view ofthe pertinent prior art.

Claims

We claim:

1. Composite honeycomb sandwich structure, comprising:

(a) a honeycomb core, having core cells;

(b) at least one composite laminate having plies of fiber-reinforced matrix resin adhered to the core;

(c) a film barrier layer between the laminate and the core to bond the laminate and core and to eliminate resin flow from the laminate into the core cell; and

(d) a film adhesive with supporting scrim between the barrier layer and the core to eliminate resin flow to or sagging of the barrier film into the core cells.

2. The structure of claim 1 wherein the laminate includes bismaleimide matrix resin.

3. The structure of claim 1 wherein the barrier layer is a bondable grade, polyimide.

4. The structure of claim 2 wherein the film adhesive includes bismaleimide.

5. The structure of claim 4 further comprising an unsupported film adhesive layer between d e barrier layer and the laminate.

6. A method for eliminating the flow of resin from laminate skins of a composite honeycomb sandwich panel to cells of the honeycomb comprising the step of: containing me resin in the skin with a scrim supported barrier film tiiat is impermeable to the resin and tiiat is adhered between the skin and honeycomb.

7. A method for adhering tiedown plies together in the manufacture of composite structure, comprising the steps of:

(a) assembling on the layup mandrel a composite preform in the shape of die composite structure, the preform having at least one resin-impregnated laminate and at least two tiedown plys; and

(b) adhering the tiedown plys to one another witii a film adhesive applied to die plys outside a net trim line of d e composite structure.

wherein d e film adhesive cures at a temperature lower than the resin in the laminate.

8. The mediod of claim 7 wherein the laminate includes bismaleimide matrix resin.

9. The mediod of claim 7 wherein the preform includes a barrier film made from a bondable grade, polyimide adjacent the laminate.

10. The method of claim 8 wherein the preform includes a honeycomb core and an adhesive between the barrier film and core.

11. The mediod of claim 10 wherein the preform includes a film adhesive layer between the barrier film and the laminate.

12. The method of claim 10 wherein the preform includes a supporting scrim between d e barrier film adhesive and die core to prevent sagging of the barrier film into the core cells.

13. The method of claim 12 wherein d e preform includes a tiedown ply in contact witii the core between the adhesive and core.

14. A method for reducing core crush in a chamfered composite honeycomb sandwich panel having a resin-impregnated laminate adhered to a honeycomb core, die core having a chamfer, comprising the steps of:

(a) contacting a tiedown ply witii the honeycomb core of the panel in the region of d e chamfer to prevent slippage between the core and the laminate; and

(b) assembling tiedown plies over outer surfaces of the laminate;

(c) adhering d e tiedown plies together and to d e layup mandrel with a lower temperature curing adhesive applied to the tiedown plies outside a net trim line of the panel,

wherein the adhesive melts and cures prior to the application of autoclave pressure and prior to die melt and flow of die resin in d e laminate.

15. The mediod of claim 14 wherein the laminate includes a barrier fihn to prevent resin flow from facesheets of the laminate to the core cells, and wherein one tiedown ply is between the barrier film and core.

16. Composite honeycomb sandwich structure having improved resistance to core crush, comprising:

(a) a honeycomb core, having core cells and a peripheral chamfer;

(b) at least one composite laminate having plies of fiber-reinforced matrix resin adhered to die core;

(c) a barrier film adhesive between d e laminate and d e core to bond die laminate and core and to eliminate resin flow from the laminate into d e core cells; and

(d) a peripheral tiedown ply in contact with the chamfer of the core beneath d e adhesive to eliminate slippage of the barrier film relative to the core and, in so doing, to reduce core crush.

17. The structure of claim 16 wherein the laminate includes bismaleimide matrix resin.

18. The structure of claim 16 wherein the barrier fihn is a bondable grade, polyimide.

19. The structure of claim 17 wherein the adhesive includes bismaleimide.

20. The structure of claim 19 further comprising a film adhesive layer between the barrier film and die laminate.

21. The structure of claim 16 further comprising a supporting scrim between the barrier film adhesive and the core to prevent sagging of the barrier film into the core cells.

22. Composite honeycomb sandwich structure resistant to core crush caused by slippage of a composite laminate along a chamfer of a honeycomb core, comprising:

(a) a honeycomb core having a chamfer;

(b) a tiedown ply contacting the chamfer; and

(c) at lease one laminate adhered to die core through the tiedown ply at d e chamfer wherein die tiedown ply prevents damaging slippage of the laminate relative to the core that would produce core crush during autoclave curing of the structure to adhere the core to the laminate.

23. A method for reducing core crush in a chamfered composite honeycomb sandwich panel having a laminate adhered to a honeycomb core, d e core having a chamfer, comprising the step of: contacting a tiedown ply witii die honeycomb core of the panel in the region of the chamfer to prevent slippage between the core and the laminate.

24. The method of claim 23 wherein die laminate includes a barrier film to prevent resin flow from the laminate to the core cells.