JP4407964B2 - Composite honeycomb sandwich structure - Google Patents

Description

  The present invention relates to a composite honeycomb sandwich structure, and more particularly, a resin that forms an external skin that is bonded on both opposing surfaces of a honeycomb core with an intermediate barrier to eliminate resin flow from the skin to the core. The present invention relates to an impregnated fabric sheet.

  An aerospace honeycomb core sandwich panel (with a composite laminated skin that is cured with an adhesive to the core through autoclaving) provides a high stiffness to weight ratio (ie, “specific stiffness”) that this panel provides and Due to the strength to weight ratio (ie specific strength), it is widely used today. Typical honeycomb core sandwich panels are described in US Pat. Nos. 5,284,702, 4,622,091 and 4,353,947. Alteneder et al., “Processing and Characterization Studies of Honeycomb Composite Structures,” 38th International SAMPE Symposium, May 10-13, 1993 (PCL Internal No. 200). -01 / 93-AWA) discusses common problems with these panels, including core collapse (ie, core crush), porosity of skin laminates and poor surface finish of the tool.

  US Pat. No. 5,445,861 describes a composite sandwich structure for sound absorption (acoustic insulation) and other applications. This sandwich structure has the following seven layers.

(1) External skin (2) Honeycomb or foam core in small cells (3) Front internal septum (4) Intermediate honeycomb core in large cells (5) Rear internal septum (6) Small cells on the back The resulting honeycomb or foam core and (7) the tuned cavity absorber in the inner skin intermediate honeycomb core absorbs sound. The performance of this structure is adversely affected by the resin flow to the cells of the honeycomb core that occurs during fabrication for reasons already described, and because such flow changes the resonance of the structure.

SUMMARY OF THE INVENTION In a high flow resin system, a large amount of resin can flow into the core during an autoclave treatment cycle. Resin is lost from the laminate due to such a flow, and the weight of the panel needs to be added to achieve the desired performance, and an excessive design of the laminate ply is needed to deal with the loss due to the flow. Since the resin is lost from the laminate ply, the thickness of the cured ply is reduced and the mechanical performance is degraded. In order to achieve the desired performance and the corresponding laminate thickness, additional plies are required, resulting in additional cost and weight. In aerospace design and manufacturing, the addition of weight is serious in terms of impact on vehicle performance and the cost of modern aircraft, and flow is a relatively unpredictable and uncontrollable process. It is required to eliminate or greatly reduce the flow to the core. In addition to the weight addition due to the resin flow to the core, we have discovered that microcracks generated in the transferred resin can propagate to the bond line, degrading mechanical performance. The possibility of such microcracks poses a serious threat to the integrity of the panel and requires the flow to be eliminated or at least controlled.

  The flow from the laminate to the core occurs because the viscosity of the resin is reduced (ie thinned) at high processing temperatures. Thus, prior art attempts to solve the flow problem generally focused on maintaining the ambient temperature viscosity of the resin at the cure temperature. For example, the processing cycle could be altered to initiate resin curing during the slow heat up and low pressure phase to cause resin chain growth prior to completion of the high temperature and high pressure. In this staged cure cycle, one would attempt to maintain resin viscosity by accumulating molecular weight at low temperatures. Higher molecular weight resins have inherently higher viscosities and therefore remain thicker and resistant to harmful core flow. Unfortunately, the staged cure cycle still produces too much flow and there are still many potential problems with microcracks. Also, the porosity of the facesheet may increase beyond acceptable limits. Furthermore, the autoclave treatment time is increased by changing the curing cycle. As processing time increases, production costs increase considerably, with the risk that expensive parts will be rejected due to improperly understood and uncontrolled factors.

  We use a high flow resin system to eliminate the flow of resin (matrix) to the honeycomb core for sandwich structures, reproducibility and predictability in sandwich panel fabrication, and confidence in the structural performance of the resulting panel Get sex. We use a barrier film supported by a scrim between a resin composite laminate reinforced with fibers and a honeycomb core. This sandwich structure is the same because the resin stays in the laminate (skin) that gives structural strength, rather than the resin flowing into the core where it makes sense, resulting in excess weight and potential panel defects. Lighter than prior art panels for performance characteristics. We also generally use a support-free film adhesive between the barrier film and the laminate to bond the laminate to the barrier film. These layers (which may be combined into one product) allow designers to achieve improved performance, hold the resin in the laminate, and otherwise address the resin flow to the core Reduce the excess resin that needs to be incorporated into the design and make the panel with structural reliability reliable.

  When attempting to cure a panel having a barrier film supported by a scrim, particularly when attempting to use a lightweight core material, core crushing often occurred in the chamfer region of the honeycomb core. By including a fixing ply in contact with the core under the barrier film (and adhesive), the fixing ply reduces the displacement of the barrier film relative to the core during curing, thus reducing core crashes in these panels Can do.

  By controlling core misalignment in this invention, a lower density honeycomb core can be used to produce a structure without expensive scrap due to core crash. We both reduce time, materials and recycling / scrap, and improve the reliability of the manufacturing process to produce high quality aerospace panels with the highest specific strength and stiffness Reduce manufacturing costs.

An additional fixing ply means that three or more fixing plies will be included in the final preform of the panel. In conventional practice, there will also be a fixing ply on the outer surface of the panel and possibly between the laminate and the adhesive barrier film. Each fixing ply extends outwardly from the part beyond the net trim line of the finished product. Conventionally, the fixing plies are individually and sequentially fixed to the layup mandrel with tape. Especially when using low density cores, it is important to establish relationships between plies and to mandrels. Tape defects cause face sheet ply wrinkles or core crashes. A core crash may still sometimes occur when the fixing ply in contact with the core is pulled backwards from the tape that secures it to the mandrel and deviates from the other fixing plies. To overcome the force acting on the core in the panel when applying autoclave pressure, the adhesive force of the tape alone is not sufficient. We have discovered an easy and inexpensive way to bond the fixing plies together in a reliable manner. When the plies are bonded together, the force acting on any individual ply is distributed to all the fixing plies, reducing the maximum force on the tape that bonds the fixing plies to the mandrel. Although described in connection with a composite honeycomb sandwich structure, this adhesion method is generally applicable to all applications involving fastening plies in a composite structure.

  Accordingly, in one aspect, the present invention relates to improvements in the manufacture of composite structures, and more particularly to composite honeycomb sandwich structures in which a fixing ply is used to fix parts during autoclave curing under high temperature and pressure. In order to lock the fixing plies together so that there is no movement of one ply relative to the other ply, to cure several plies and connect them together during the initial stage of autoclave curing before applying pressure, Use a curing adhesive. We can apply adhesive to the outside of the net trim line of the part so that it can be removed during the finishing of the part.

  In another aspect, the present invention relates to the bonding of fixing plies together during the construction of a composite structure, in particular during autoclaving under high temperature and pressure of a composite honeycomb sandwich structure. Since the tape attachment must be sufficient to prevent any ply or one ply from shifting relative to each other, the conventional practice of attaching the fixing ply only to the mandrel is inadequate. We can effectively bond the plies together so that the maximum force on the tape is reduced by attaching a cold-curing film adhesive between the fixing plies just outside the part's net trim line I found In the autoclave, this film adhesive melts and cures at a lower temperature than the resin in the laminate, and therefore the laminate resin flows and cures before the autoclave pressure increases at higher temperatures. The fixing plies are bonded together. The film adhesive eliminates movement of the fixing plies relative to each other. In our preferred embodiment for a bismaleimide (BMI) sandwich panel, we cured at about 375 degrees Fahrenheit (191 degrees Celsius) and about 250 degrees Fahrenheit (121 degrees Celsius) for BMI cured at about 440 degrees Fahrenheit. Prefer to use an adhesive that cures at a degree).

Detailed Description of the Preferred Embodiments As the basis for this description, we first describe a typical composite honeycomb sandwich structure. Next, the method of our invention for reliably bonding the fixing plies together is described.

  The composite honeycomb sandwich panel minimizes, eliminates or significantly reduces resin flow from the laminate to the core, thereby enabling a stronger and easier processing cycle for the manufacture of aerospace structures. Such a sandwich panel 100 (FIG. 1) typically has an outer face sheet or skin 102 that is bonded to a central honeycomb core 106. The finished skin 102 comprises a laminate of layers of organic matrix resin reinforced with fibers in a composite form that is cured and consolidated. The core 106 can be paper, synthetic paper, metal, composite, etc., suitable for this application. In the panel of the present invention, we incorporate at least one securing ply between the core 106 and the skin 102 to reduce the detrimental shift between the core and skin that often occurs otherwise. Since we reduce core crush during autoclave curing, we get a higher specific strength and a higher specific stiffness.

  In order to prevent resin flow from the composite laminate skin to the core, the unsupported film-like adhesive 108 (FIG. 2) is placed between the skin 102 and the core 106 to keep the resin outside the cells 114 of the core 106. ), We use a barrier film 110 and a film adhesive 112 supported by a scrim.

  FIG. 3 illustrates a core filling problem that can occur when the film adhesive 112 is used alone without the barrier film 110 and the film adhesive 108. The honeycomb cells 114 are filled with a resin 118 that moves from the laminate and thereby reduces the resin in the skin 102. This adversely affects structural performance since the thickness of the ply is reduced by the reduction in resin. Since the resin 118 in the cell is merely wasted, the total weight increases due to the decrease in resin. In any case, uncontrolled resin flow and reduction may begin in the resin 118 in the cell, especially during thermal cycling, and the fiber reinforced skin 102, particularly the bond between skin 102 and core 106. The panel becomes suspicious for microcracks that can move into the line.

  FIG. 4 illustrates the undesired bulges that can occur if the barrier film 110 is used without the film adhesive 112 supported by the scrim in an attempt to eliminate the cell resin 118. Here, the waste resin bulge 120 protrudes downward into the cells 114 of the honeycomb core 106. Although resin is contained within the bulge 120, the resin of the skin 102 still decreases. Resin flow to the bulge 120 imposes structural performance penalties and weight additions comparable to the uncontrolled condition shown in FIG.

  As shown in FIG. 2 with film adhesive 108, barrier film 110 and film adhesive 12 supported on a scrim, the resin flow without cell resin 118 or resin bulge 120 is examined. However, we have found that the barrier film creates a sliding surface between the laminate skin and the core, which often causes a core crash during the autoclaving cycle. In fact, in our initial test, 22 of 31 test panels experienced core crashes. This defect rate is unacceptable from a cost and schedule perspective. Our fastening plies in the chamfer area reduce or eliminate harmful core misalignments and the frequency of core crashes due to such misalignments.

  RIGIDITE (registered trademark) 5250-4-W-IM7-GP-CSW, RIGIDITE (registered trademark) 5250-4- from Cytec Engineered Materials, Inc (Cytec) For bismaleimide laminated skins made with W-IM7-GP-CSX and RIGIDITE® 5250-4-W-IM7-GP-PW prepreg, film adhesive 108 is also available from Cytec. It is preferably a .015 psf METTLBOND® 2550U adhesive. The film adhesive provides additional resin to promote a high quality bond between the laminate and the barrier film 110. The barrier film 110 is 0.001 inch thick, that is, can be bonded to withstand a cure cycle to provide a resin impermeable membrane between the skin 102 and the core 106. A KAPTON (registered trademark) polyimide barrier film is preferred. The scrim is preferably a glass fiber, “Style 104” fiber cloth, and the film adhesive 112 is 0.06 psf METLBOND® 2550G adhesive available from Cytec. The film-like adhesive supported by the scrim prevents the barrier film from expanding into the core cell, thereby holding the resin in the laminate (ie skin layer) so that the thickness of the cured ply is maximized, thereby We achieve maximum performance with minimum weight for the panel.

Film adhesive 108, barrier film 110 and film adhesive 112 are M
ETLBOND (registered trademark) 2550B-. It can be purchased as a single item from Cytec as 082 36 ″.

  The ply of the skin 102 is typically a carbon fiber prepreg impregnated with a bismaleimide thermosetting resin, although the invention applies to other resin systems. Tow may be used instead of the prepreg. Film adhesive 108 must be adjusted to achieve a proper bond between skin 102 and barrier film 110. The honeycomb core is typically an HRP Fiberglass Reinforced Phenolic honeycomb available from Hexcel.

  The support-like film adhesive and barrier film layer in the sandwich structure is formed between the skin 102 and the core 106 when the core is a metal such as aluminum and the skin includes a galvanically electrically dissimilar material such as carbon fiber. It also functions as a barrier against corrosion.

  Additional information on preferred panels can be found in the technical paper we have incorporated by reference, Harz et al., “Development of a Bismaleimade / Carbon Honeycomb Sandwich Structure”, SAMPE, 1996. Shown in March of the year. This article describes both the Harz et al. Barrier film improvement, the fixing ply method and the bonding method of the present invention.

  The Harz-type panel provides mechanical and physical edgeband features equivalent to a solid BMI / carbon laminate (cured at 859 psig). Our tests confirm that the edge band cured ply thickness is equal to the solid laminate in our panel and that the edge band 160 (FIGS. 5 and 6) meets the requirements of the solid laminate nondestructive inspection specification. The mechanical performance of the edgeband and facesheet is improved over the results we have achieved with a sandwich structure that lacks a combination of adhesive, barrier film, and adhesive supported by the scrim. Penetration tensile mechanical performance also meets the design requirements.

  Especially for phenolic cores, by preconditioning the core so that no volatilization proceeds while the core is cured by heating to about 235 degrees Celsius (455 degrees Fahrenheit) before laying up the sandwich panel, This eliminates the separation between the core and the laminated plate caused by the outgassing from the core.

  When the autoclave pressure is applied and the resin melts, the core crush 200 (FIG. 5) occurs in the chamfer portion region 155 when the barrier film 110 and the core 106 are displaced from the face sheet 102. As illustrated in FIG. 5, the barrier film 100 and the core 106 move to the right to compress the core in the chamfer region 155, creating a core crash 200. The skin 102 bends in the edge band region 160 where the core has moved.

  Referring to FIG. 6, the improved honeycomb sandwich panel includes at least one fixing ply 150 that contacts the core 106 along the chamfer portion 155. Such chamfers (ie, angled transitions in the core, often in edgeband 160) typically occur near the periphery of the panel, but the assembled structure requires fasteners or penetrations Will occur at the joint line or hardened point in the middle of the panel.

Typically, we describe a single ply of carbon fiber or glass fiber fabric with conventional 0/90 fiber orientation in making bismaleimide panels with 5 or 8 lb / ft 3 HRP cores, as described by Harz et al. 150 is used. The anchoring ply 150 functions to inhibit or limit skin displacement relative to the core so as to reduce core crashes that would otherwise be caused by the displacement. When the preform is heated during the autoclaving cycle, the matrix resin softens and melts, and the high flow resin essentially melts, the fixing ply 150 clamps the core with the inherent roughness of the fibers. . These panels allow us to save 2.5 to 4 lb / ft 3 of the core because we can use a lower density honeycomb core without suffering from core crashes. For fighters, this change can save 25 lbs per vehicle.

  As shown in FIG. 6, the locking ply 150 is in contact with the core 106 along at least a portion of the chamfer portion 155 that overlaps the core 106 about 1 inch and extends beyond the trim line 165 of this part to the edge band 160. A narrow peripheral strip that extends outward. The fixing ply 150 will be either on the flat side of the chamfer section or on an angled surface (as shown in FIG. 6). An important factor is that the fixing ply 150 contacts the core under the adhesive and barrier film 110 used to bond the laminate skin to the core. The fixing ply 150 is cut everywhere in the body of the part except for the narrow peripheral area in the chamfer area to form a peripheral frame around the edge of the panel. In this way, the fixing ply 150 allows an adhesive interface between the core 106 and the skin 102 in the panel area.

  Traditionally, in the manufacture of Harz-type panels, we used four complete cover sheet fastening plies 175 to secure the layers and cores, all of which are illustrated in FIG. These conventional plies 175 are commonly used in making sandwich panels prior to the introduction of the Harz-type barrier film, and we usually use them all, but now except for the fixing ply 150 that contacts the outer ply and the surrounding core. I believe you can remove all plies. That is, a total of three plies would be used instead of five as illustrated in FIG.

  The fixing plies 150 and 175 extend through the edge band 160 and beyond the net trim line 165 to the fastening point we tape to the layup mandrel. In addition, to prevent misalignment of the fixing ply, we incorporated a low temperature cure (ie 121 degrees Celsius for BMI panels) film adhesive 180 between the fixing plies just outside the net trim line of the part. Film adhesive 180 eliminates movement of one ply relative to the other ply when pressure is applied during the autoclave cure cycle. By curing at a temperature about 100 to 150 degrees Fahrenheit below the curing temperature of the laminate resin, the fixing adhesive hardens before we have to increase the autoclave pressure, and the cured adhesive becomes the fixing ply. Are combined with each other. By using this bonding method, there is no relative ply movement, face sheet wrinkling and core crash that could otherwise occur.

  This securing method uses a “picture frame” perimeter securing ply 150 (excluding the conventional inner sheet), thus saving material, reducing costs and saving weight. The normal fastening procedure requires a ply on the outer surface of the skin and between the skin and the base adhesive (FIG. 5).

  Conventional fastening systems will not work well without the “picture frame” ply because the barrier film 110 allows for core misalignment. The Corbett and Smith methods will sometimes not work well without the bonding method of this invention.

  For a lightweight core (ie, 5-8 lb / ft 3) with the previously described bismaleimide prepreg and adhesive system, we maintain a chamfer angle of 20 ° ± 2 °.

By the term “chamfer part” we mean an angled cutting area (curved part) of the honeycomb core that tapers at a constant slope from full thickness to no thickness. The chamfer part is used in the edge band of a composite honeycomb sandwich panel to provide a smooth transition between the structural body of the panel with the embedded honeycomb and the connecting edge band entirely covered by the honeycomb core. The method of the present invention makes it possible to use a chamfer part angle that is often much steeper than that required in conventional implementations if it is desired to avoid core crashes without one locking ply. To do. We prefer the 20 ° chamfer section, but we believe that this angle can be increased to any angle suitable for the design requirements of the panel.

  By the term “autoclaving” we mean a cycle where high temperature and high pressure are applied to the panel to consolidate and cure the resin in the laminate and simultaneously bond or otherwise bond the cured laminate to the honeycomb core. . Our preferred cycle is illustrated in FIG. Our adhesive for fixing plies cures at about 250 degrees Fahrenheit (121 degrees Celsius), so it cures before an increase in autoclave pressure that can introduce relative movement between the layers in the panel.

  If a core crash occurs, the damage to the panel is usually so extensive that it cannot be repaired and the part is discarded. Today's advanced composite resins and reinforcing fiber costs require processing that is substantially free of core crush. Otherwise, processing costs are prohibitive. Core crashes are a big problem when panels are designed as close to design limits as possible. The method of the present invention reduces core crash and ply movement or wrinkles.

  While the preferred embodiment has been described, variations, changes and modifications that may be made without departing from the spirit of the invention will be readily apparent to those skilled in the art. Accordingly, the claims should be construed freely based on this specification, with the full scope of equivalents known to those skilled in the art. The examples here are given to illustrate the invention and are not intended to limit the invention. Accordingly, the invention is defined by the following claims and should be limited only as necessary in light of the pertinent prior art.

1 is a diagram showing a typical composite honeycomb sandwich structure. FIG. FIG. 4 is a schematic partial cross-sectional view of a skin-core interface in a sandwich structure having a barrier film supported on a scrim to prevent resin flow from the skin to the core. 1 is a schematic partial cross-sectional view of a prior art honeycomb sandwich structure that is adversely affected by resin flow to the core using a film adhesive with support without a barrier film. FIG. FIG. 6 is another schematic partial cross-sectional view showing a sandwich structure in which the resin in the skin is reduced, but the curved unsupported barrier film prevents the resin from reaching the core. FIG. 4 is a schematic cross-sectional elevational view showing a core crush of a honeycomb sandwich panel caused by a shift between a core and a barrier film. FIG. 6 is another schematic cross-sectional elevation view illustrating the use of a fixing ply to reduce core crash. 2 is a graph of a typical autoclave cure cycle to make a composite honeycomb sandwich panel showing that our anchoring adhesive cures before high pressure is applied in the cycle.

Claims (17)

A method for bonding and bonding fixing plies together in the manufacture of a composite structure,
Comprising the steps of assembling the composite preform in the form of a composite structure on top of (a) lay-up mandrel, the composite preform has at least one resin impregnated laminate and at least two fixing plies, said method comprising further,
(B) comprises the step of bonding together the fixing plies by a film-like adhesive provided on the fixed plies outside the net trim line of the composite structure,
The film adhesive is cured at a temperature lower than that of the resin in the resin-impregnated laminate, and is a method for bonding together the fixing ply in the production of a composite structure. The method for bonding together a fixing ply in the production of a composite structure according to claim 1, wherein the resin-impregnated laminate comprises a bismaleimide matrix resin. The composite preform comprises a barrier film made of polyimide surface-treated so as to be able to adhere to other layers adjacent to the resin-impregnated laminate. A method for bonding together. Composite preform, honeycomb core and comprises an adhesive between the barrier film and the honeycomb core, a method for bonding together the locking ply in the manufacture of a composite structure according to claim 2. Composite preform comprises a film-like adhesive layer between the barrier film and the resin-impregnated laminate, a method for bonding together the locking ply in the manufacture of a composite structure of claim 4. Composite preform, to prevent bending of the core cell barrier film, comprising a scrim to support the adhesive between the Bariyafiru arm and the honeycomb core, the fixing plies in the manufacture of a composite structure according to claim 4 A method for bonding together. Composite preform comprises a fixing plies in contact with the honeycomb co <br/> A between the adhesive and the honeycomb core is supported by the scrim, the fixing plies in the manufacture of a composite structure according to claim 6 A method for bonding together. A method for reducing core crush in a composite honeycomb sandwich panel with a chamfer part having a resin-impregnated laminate bonded to a honeycomb core, wherein the honeycomb core has a chamfer part, the method comprising:
(A) contacting the honeycomb core of the panel with the fixing ply in the chamfer part region in order to prevent displacement between the honeycomb core and the resin-impregnated laminate;
(B) assembling a fixing ply on the outer surface of the resin-impregnated laminate;
(C) adhering the fixing plies to each other and to the layup mandrel with a lower temperature curing adhesive applied to the fixing ply outside the net trim line of the composite honeycomb sandwich panel;
A method for reducing core crushing in a composite honeycomb sandwich panel with a chamfer part, wherein the adhesive melts and hardens before the autoclave pressure is applied and before the resin in the resin-impregnated laminate melts and flows. The resin-impregnated laminate includes a barrier film for preventing resin flow from the face sheet of the resin-impregnated laminate to the core cell, and one fixing ply is located between the barrier film and the honeycomb core. A method for reducing core crush in a composite honeycomb sandwich panel with a chamfer part as described. A composite honeycomb sandwich structure with improved resistance to core crash,
(A) a honeycomb core having a core cell and a peripheral chamfer part;
(B) at least one composite laminate having a matrix resin ply reinforced with fibers bonded to the honeycomb core;
(C) by combining the composite laminate and the honeycomb core, and, to eliminate the flow of resin to the core cell from composite laminate, and Bariyafiru arm between the composite laminate and the honeycomb core,
(D) A composite honeycomb sandwich structure including a peripheral fixing ply that contacts the chamfer portion of the honeycomb core under the adhesive to eliminate barrier film displacement with respect to the honeycomb core and thereby reduce core crash. The composite honeycomb sandwich structure according to claim 10, wherein the composite laminate includes a bismaleimide matrix resin. The composite honeycomb sandwich structure according to claim 10, wherein the barrier film is a polyimide surface-treated so as to be bonded to other layers .   The composite honeycomb sandwich structure of claim 11, wherein the adhesive comprises bismaleimide. 14. The composite honeycomb sandwich structure according to claim 13, further comprising a film adhesive layer between the barrier film and the composite laminate. The composite honeycomb sandwich structure according to claim 10, further comprising a scrim for supporting an adhesive in order to prevent the barrier film from bending into the core cell. A composite honeycomb sandwich structure that is resistant to core crush caused by displacement of the composite laminate along the chamfer portion of the honeycomb core,
(A) a honeycomb core having a chamfer part;
(B) a fixing ply that contacts the chamfer part;
(C) including at least one composite laminate bonded to the honeycomb core through a fixing ply at the chamfer part;
The composite laminate has a barrier film for preventing resin flow from the composite laminate to the core cell,
Fixing plies will produce core crush during the autoclave curing of the structure for adhering the honeycomb core composite laminate, to prevent adverse deviation of the composite laminate against the honeycomb core, composite Honeycomb sandwich structure. A method for reducing core crash in a composite honeycomb sandwich panel with a chamfer part having a composite laminate bonded to a honeycomb core, wherein the honeycomb core has a chamfer part and the composite laminate is from the composite laminate to the core cell. Having a barrier film to prevent resin flow , said method comprising:
Core crushing in a composite honeycomb sandwich panel with a chamfer including a step of contacting a fixing ply to the honeycomb core of the composite honeycomb sandwich panel in the region of the chamfer to prevent deviation between the honeycomb core and the composite laminate. A way to reduce.

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