The present invention relates to a process for joining honeycomb sandwich panels with thermoplastic welds, particularly double interleaf staggered joints connecting the face sheets of the respective panels.
The use of composites in primary structure in aerospace applications is limited today because of the relatively high cost. A significant contribution to the total cost is the assembly cost where the precured composite elements are assembled, drilled, and fastened. The necessary design for mechanical fastening complicates the structure, especially in thin sections, because of the need for access to both sides of the bond line.
While composites might be adhesively bonded, cocured, or welded, these connecting processes generally produce bonds that rely upon the resin matrix for strength. The bond line lacks any reinforcing material to help with load transfer. These bonds generally have modest strength, and are susceptible to disbanding with shock impact or other xe2x80x9cout of planexe2x80x9d forces affecting the assembly. Such forces often arise in environments prone to vibration.
1. Composite Manufacturing
Fiber-reinforced organic resin matrix composites have a high strength-to-weight ratio (specific strength) or a high stiffness-to-weight ratio (specific stiffness) and desirable fatigue characteristics that make them increasingly popular as a replacement for metal in aerospace applications where weight, strength, or fatigue is critical. Thermoplastic or thermoset organic resin composites would be more economical with improved manufacturing processes that reduced touch labor and forming time.
Prepregs combine continuous, woven, or chopped reinforcing fibers with an uncured matrix resin, and usually comprise fiber sheets with a thin film of the matrix. Sheets of prepreg generally are placed (laid-up) by hand or with fiber placement machines directly upon a tool or die having a forming surface contoured to the desired shape of the completed part or are laid-up in a flat sheet which is then draped and formed over the tool or die to the contour of the tool. Then the resin in the prepreg lay up is consolidated (i.e. pressed to remove any air, gas, or vapor) and cured (i.e., chemically converted to its final form usually through chain-extension or fused into a single piece) in a vacuum bag process in an autoclave (i.e., a pressure oven) to complete the part.
The tools or dies for composite processing typically are formed to close dimensional tolerances. They are massive, must be heated along with the workpiece, and must be cooled prior to removing the completed part. The delay caused to heat and to cool the mass of the tools adds substantially to the overall time necessary to fabricate each part. These delays are especially significant when the manufacturing run is low rate where the dies need to be changed frequently, often after producing only a few parts of each kind. An autoclave has similar limitations; it is a batch operation.
In hot press forming, the prepreg is laid-up to create a preform, which is bagged (if necessary), and placed between matched metal tools that include forming surfaces to define the internal, external, or both mold lines of the completed part. The tools and composite preform are placed within a press and then the tools, press, and preform are heated.
The tooling in autoclave or hot press fabrication is a significant heat sink that consumes substantial energy. Furthermore, the tooling takes significant time to heat the composite material to its consolidation temperature and, after curing the composite, to cool to a temperature at which it is safe to remove the finished composite part.
As described in U.S. Pat. No. 4,657,717, a flat composite prepreg panel was sandwiched between two metal sheets made from a superplastically formable alloy and was formed against a die having a surface precisely contoured to the final shape of the part.
Attempts have been made to reduce composite fabrication times by actively cooling the tools after forming the composite part. These attempts have shortened the time necessary to produce a composite part, but the cycle time for heating and cooling remains long. Designing and making tools to permit their active cooling also increases their cost.
Boeing described a process for organic matrix forming and consolidation using induction heating in U.S. Pat. No. 5,530,227. There, Boeing laid up prepregs in a flat sheet sandwiched between aluminum susceptor sheets. The susceptor sheets were susceptible to heating by induction and formed a retort to enclose the prepreg preform. To ensure an inert atmosphere around the preform during curing and to permit withdrawing volatiles and outgassing during the consolidation, the face sheets are welded around their periphery. Such welding unduly increased the preparation time and the cost for part fabrication. It also ruined the susceptor sheets (i.e., prohibited their reuse) which added a significant cost penalty to each part fabricated with this approach. Boeing described in U.S. Pat. No. 5,599,472 a technique that readily and reliably sealed the susceptor sheets of the retort without the need for welding and permitted reuse of the susceptor sheets in certain circumstances. This xe2x80x9cbag-and-sealxe2x80x9d technique applies to both resin composite and metal processing.
2. Processing in an Induction Press
The dies or tooling for induction processing in Boeing""s induction heating workcell are ceramic because a ceramic is not susceptible to induction heating and, preferably, is a thermal insulator (i.e., a relatively poor conductor of heat). Ceramic tooling usually is strengthened and reinforced internally with fiberglass rods or other appropriate reinforcements and externally with metal or other durable strongbacks to permit it to withstand the temperatures and pressures necessary to form, to consolidate, or otherwise to process the composite materials or metals. Ceramic tools cost less to fabricate than metal tools of comparable size and have less thermal mass than metal tooling, because they are unaffected by the induction field. Because the ceramic tooling is not susceptible to induction heating, it is possible to embed induction heating elements in the ceramic tooling and to heat the composite or metal retort without significantly heating the tools. Thus, induction heating can reduce the time required and energy consumed to fabricate a part.
While graphite or boron fibers can be heated directly by induction, most organic matrix composites require a susceptor in or adjacent to the composite material preform to achieve the necessary heating for consolidation or forming. The susceptor is heated inductively and transfers its heat principally through conduction to the preform or workpiece that, in Boeing""s prior work, is sealed within the susceptor retort. Enclosed in the metal retort, the workpiece does not experience the oscillating magnetic field which instead is absorbed in the retort sheets. Heating is by conduction from the retort to the workpiece.
Induction focuses heating on the retort (and workpiece) and eliminates wasteful, inefficient heat sinks. Because the ceramic tools in the induction heating workcell do not heat to as high a temperature as the metal tooling of conventional, prior art presses, problems caused by different coefficients of thermal expansion between the tools and the workpiece are reduced. Furthermore, Boeing""s induction heating press is energy efficient because significantly higher percentages of input energy go to heating the workpiece than occurs with conventional presses. The reduced thermal mass and ability to focus the heating energy permits change of the operating temperature rapidly which improves the products produced. Finally, the shop environment is not heated as significantly from the radiation of the large thermal mass of a conventional press. The shop is a safer and more pleasant environment for the press operators.
In induction heating for consolidating and forming organic matrix composite materials, Boeing generally places a thermoplastic organic matrix composite preform of PEEK or ULTEM, for example, within the metal susceptor envelope (i.e., retort). These thermoplastics have a low concentration of residual volatile solvents and are easy to use. The susceptor face sheets of the retort are inductively heated to heat the preform. Consolidation and forming pressure consolidate and, if applicable, form the preform at its curing temperature. The sealed susceptor sheets form a pressure zone in the retort in a manner analogous to conventional vacuum bag processes for resin consolidation. The retort is placed in an induction heating press on the forming surfaces of dies having the desired shape of the molded composite part. After the retort and preform are inductively heated to the desired elevated temperature, differential pressure (while maintaining the vacuum in the pressure zone around the preform) across the retort which functions as a diaphragm in the press forms the preform against the die into the desired shape of the completed composite panel.
The retort often includes three, stacked susceptor sheets sealed around their periphery to define two pressure zones. The first pressure zone surrounds the composite panel/preform or metal workpiece and is evacuated and maintained under vacuum. The second pressure zone is pressurized (i.e., flooded with gas) at the appropriate time to help form the composite panel or workpiece. The shared wall of the three layer sandwich that defines the two pressure zones acts as the diaphragm.
Boeing can perform a wide range of manufacturing operations in its induction heating press. These operations have optimum operating temperatures ranging from about 350xc2x0 F. (175xc2x0 C.) to about 1850xc2x0 F. (1010xc2x0 C.) or even above. For each operation, Boeing usually holds the temperature relatively constant for several minutes to several hours to complete the operations. While temperature can be controlled by controlling the input power fed to the induction coil, a better and simpler way capitalizes on the Curie temperature. Judicious selection of the metal or alloy in the retort""s susceptor face sheets avoids excessive heating irrespective of the input power. With improved control and improved temperature uniformity in the workpiece, Boeing produces better products. The method capitalizes on the Curie temperature phenomenon to control the absolute temperature of the workpiece and to obtain substantial thermal uniformity in the workpiece by substantially matching the Curie temperature of the susceptor to the desired temperature of the induction heating operation being performed. The Curie temperature is generally slightly above the processing temperature. With thermoplastic welding, for example, it might be the melt temperature of the matrix resin plus about 1-75xc2x0 F. (preferably 5-25xc2x0 F.) so that the bond line remains in the processing window without excessive heating. This temperature control method is explained in greater detail in Boeing""s U.S. Pat. Nos. 5,723,849 or 5,645,744.
U.S. Pat. No. 5,587,098 describes joining large structures in a Boeing induction heating press with localized heating along the bond line. The die is modified to include a xe2x80x9csmartxe2x80x9d susceptor corresponding to the region of the bond line. The susceptor heats when the embedded induction coil is energized. The temperature of the susceptor is controlled by its Curie point. The susceptor heats the bond line selectively, and the surrounding ceramic dies trap the heat. Residual stress at the joint can be relieved by using a susceptor that has segments of different Curie temperature materials extending in a successive pattern outwardly from the bond line to produce a thermal gradient from ambient to the bond line temperature. We prefer two or three temperature zones with these segments, typically a central welding or bonding zone with outer zones each 50-100xc2x0 F. lower in temperature than the welding zone.
3. Thermoplastic Welding
Three major joining technologies exist for aerospace composite structure: mechanical fastening; adhesive bonding; and welding. Both mechanical fastening and adhesive bonding are costly, time consuming assembly steps that introduce excess cost even if the parts that are assembled are fabricated from components produced by an emerging, cost efficient process. Mechanical fastening requires expensive hole locating, drilling, shimming, and fastener installation, while adhesive bonding often requires complicated surface pretreatments.
Thermoplastic welding eliminates fasteners. It joins thermoplastic composite components at high speeds with minimum touch labor and little, if any, pretreatments. A conventional welding interlayer tape (compromising the susceptor and surrounding thermoplastic resin either coating the susceptor or sandwiching it) also can simultaneously take the place of shims required in mechanical fastening. As such, composite welding promises to be an affordable joining process. For xe2x80x9cweldingxe2x80x9d a combination of thermoplastic and thermoset composite parts together, the resin that the susceptor melts functions as a hot melt adhesive. If fully realized, thermoplastic-thermoset bonding in addition to true thermoplastic welding will further reduce the cost of composite assembly.
There is a significant stake in developing a successful induction thermoplastic welding process. Its advantages versus traditional composite joining methods are:
reduced parts count versus fasteners
minimal surface preparation, in most cases a simple solvent wipe to remove surface contaminants
indefinite shelf life at room temperature
short process cycle time, typically measured in minutes
enhanced joint performance, especially hot/wet and fatigue
the possibility of rapid field repair of composites or other structures.
There is little or no loss of bond strength after prolonged exposure to environmental influences.
U.S. Pat. No. 4,673,450 describes a method to spot weld graphite fiber reinforced PEEK composites using a pair of electrodes. After roughening the surfaces of the prefabricated PEEK composites in the region of the bond, Burke placed a PEEK adhesive ply along the bond line, applied a pressure of about 50-100 psi through the electrodes, and heated the embedded graphite fibers by applying a voltage in the range of 20-40 volts at 30-40 amps for approximately 5-10 seconds with the electrodes. Access to both sides of the assembly was required in this process which limited its application.
U.S. Pat. Nos. 3,966,402 and 4,120,712 describe thermoplastic welding with induction heating. In these patents, conventional metallic susceptors are used and have a regular pattern of openings of traditional manufacture. Achieving a uniform, controllable temperature in the bond line, which is crucial to preparing a thermoplastic weld of adequate integrity to permit use of welding in aerospace primary structure, is difficult with those conventional susceptors.
Thermoplastic welding is a process for forming a fusion bond between two faying thermoplastic faces of two parts. A fusion bond is created when the thermoplastic on the surface of the two thermoplastic composite parts is heated to the melting or softening point and the two surfaces are brought into contact, so that the molten thermoplastic mixes. The surfaces are held in contact while the thermoplastic cools below the softening temperature.
The same process parameters apply essentially to hot melt thermoplastic adhesive bonds between prefabricated thermoset composites.
Simple as the thermoplastic welding process sounds, it is difficult to perform reliably and repeatably in a real factory on full-scale parts to build a large structure such as an airplane wing box. One difficulty is heating the bond line properly without overheating the entire structure. Another difficulty is achieving intimate contact of the faying surfaces of the two parts at the bond line during heating and cooling because of (1) the normal imperfections in the flatness of composite parts, (2) thermal expansion of the thermoplastic during heating to the softening or melting temperature, (3) flow of the thermoplastic out of the bond line under pressure (i.e., squeeze out), and (4) contraction of the thermoplastic in the bond line during cooling.
The exponential decay of the strength of magnetic fields with distance from their source dictates that, in induction welding processes, the susceptible structure closest to the induction coil will be the hottest, since it experiences the strongest field. Therefore, it is difficult to obtain adequate heating at the bond line between two graphite or carbon fiber reinforced resin matrix composites relying on the susceptibility of the fibers alone as the source of heating in the assembly. For the inner plies to be hot enough to melt the resin, the outer plies closer to the induction coil and in the stronger magnetic field are too hot. The matrix resin in the entire piece of composite melts. The overheating results in porosity in the product, delamination, and, in some cases, destruction or denaturing of the resin. To avoid overheating of the outer plies and to insure adequate heating of the inner plies, we use a susceptor of significantly higher conductivity than the fibers to peak the heating selectively at the bond line instead of in the composites themselves. To create a weld, an electromagnetic induction coil heats a susceptor to melt and cure a thermoplastic resin (also sometimes referred to as an adhesive) to bond the elements of the assembly together.
The current density in the susceptor may be higher at the edges of the susceptor than in the center because of the nonlinearity of the coil, such as occurs when using a cup core induction coil like that described in U.S. Pat. No. 5,313,037. Overheating the edges of the assembly can result in underheating the center, either condition leading to inferior welds because of non-uniform curing. An open or mesh pattern in the susceptor embedded at the bond line allows the resin to create the fusion bond between the composite elements of the assembly when the resin heats and melts.
a. Moving coil welding processes
In U.S. Pat. No. 5,500,511, Boeing described a tailored susceptor for approaching the desired temperature uniformity. Designed for use with the cup coil of U.S. Pat. No. 5,313,037, this susceptor relied upon carefully controlling the geometry of openings in the susceptor (both their orientation and their spacing) to distribute the heat evenly. The susceptor had a regular array of anisotropic, diamond shaped openings with a ratio of the length (L) to the width (W) greater than 1. This susceptor produced a superior weld by producing a more uniform temperature than obtainable using a susceptor having a similar array, but one where the L/W ratio was one. Changing the length to width ratio (the aspect ratio) of the diamond-shaped openings in the susceptor produced a large difference in the longitudinal and transverse conductivity in the susceptor, and, thereby, tailored the current density within the susceptor. A tailored susceptor having openings with a length (L) to width (W) ratio of 2:1 has a longitudinal conductivity about four times the transverse conductivity. In addition to tailoring the shape of the openings to tailor the susceptor, Boeing altered the current density in regions near the edges by increasing the foil density (i.e., the absolute amount of metal). Increasing the foil density along the edge of the susceptor increased the conductivity along the edge and reduced the current density and the edge heating. Boeing increased foil density by folding the susceptor to form edge strips of double thickness or by compressing openings near the edge of an otherwise uniform susceptor. These susceptors were difficult to reproduce reliably. Also, they required careful placement and alignment to achieve the desired effect.
The tailored susceptor was designed to use with the cup coil of U.S. Pat. No. 5,313,037, where the magnetic field is strongest near the edges because the central pole creates a null at the center. Therefore, the tailored susceptor was designed to counter the higher field at the edges by accommodating the induced current near the edges. The high longitudinal conductivity encouraged induced currents to flow longitudinally.
The selvaged susceptor for thermoplastic welding which is described in U.S. Pat. No. 5,508,496 controls the current density pattern during eddy current heating by an induction coil to provide substantially uniform heating to a composite assembly and to insure the strength and integrity of the weld in the completed part. This susceptor is particularly desirable for welding ribs between prior welded spars using an asymmetric induction coil (described in U.S. Pat. No. 5,444,220), because it provides (1) a controllable area of intense, uniform heating under the poles of the coil; (2) a trailing region with essentially no heating; and (3) a leading region with minor preheating.
Boeing achieved better performance (i.e., more uniform heating) in rib welding by using a selvaged susceptor having edge strips without openings. The resulting susceptor, then, has a center portion with a regular pattern of openings and solid foil edges, referred to as selvage edge strips. The susceptor is embedded in a thermoplastic resin to make a susceptor/resin tape that is easy to handle and to use in preforming the composite pieces prior to welding. Also, with a selvaged susceptor, the impedance of the central portion should be anisotropic with a lower transverse impedance than the longitudinal impedance. Here, the L/W ratio of diamond shaped openings should be less than or equal to one. With this selvaged susceptor in the region immediately under the asymmetric induction work coil, current flows across the susceptor to the edges where the current density is lowest and the conductivity, highest.
Generally, the selvaged susceptor is somewhat wider than normal so that the selvage edge strips are not in the bond line. Boeing sometimes removes the selvage edge strips after forming the weld, leaving only a perforated susceptor foil in the weld. This foil has a relatively high open area fraction.
Another difficulty remaining in perfecting the thermoplastic welding process for producing large scale aerospace structures in a production environment involved control of the surface contact of the faying surfaces of the two parts to be welded together. The timing, intensity, and schedule of heat application must be controlled so the material at the faying surfaces are brought to and maintained within the proper temperature range for the requisite amount of time for an adequate bond to form. Intimate contact is maintained while the melted or softened material hardens in its bonded condition.
Large scale parts, such as wing spars and ribs, and the wing skins that are bonded to the spars and ribs, are typically on the order of 20-30 feet long at present, and potentially as much as 100 feet in length when the process is perfected for commercial transport aircraft. Parts of this magnitude are difficult to produce with perfect flatness. Instead, the typical part will have various combinations of surface deviations from perfect flatness, including large scale waviness in the direction of the major length dimension, twist about the longitudinal axis, dishing or sagging of xe2x80x9cIxe2x80x9d beam flanges, and small scale surface defects such as asperities and depressions. These irregularities interfere with full surface area contact between the faying surfaces of the two parts and actually result in surface contact only at a few xe2x80x9chigh pointsxe2x80x9d across the intended bond line. Applying pressure to the parts to force the faying surfaces into contact achieves additional surface contact, but full intimate contact is difficult or impossible to achieve in this way. Applying heat to the interface by electrically heating the susceptor in connection with pressure on the parts tends to flatten the irregularities further, but the time needed to achieve full intimate contact with the use of heat and pressure is excessive, can result in deformation of the top part, and tends to raise the overall temperature of the xe2x80x9cIxe2x80x9d beam flanges to the softening point, so they begin to yield or sag under the application of the pressure needed to achieve a good bond.
Boeing""s multipass thermoplastic welding process described in U.S. Pat. No. 5,486,684 (which we incorporate by reference) enables a moving coil welding process to produce continuous or nearly continuous fusion bonds over the full area of the bond line. The result is high strength welds produced reliably, repeatably, and with consistent quality. This process produces improved low cost, high strength composite assemblies of large scale parts fusion bonded together with consistent quality. It uses a schedule of heat application that maintains the overall temperature of the structure within the limit in which it retains its high strength. Therefore, it does not require internal tooling to support the structure against sagging which otherwise could occur when the bond line is heated above the high strength temperature limit. The process also produces nearly complete bond line area fusion on standard production composite parts having the usual surface imperfections and deviations from perfect flatness. The multipass welding process (1) eliminates fasteners and the expense of drilling holes, inspecting the holes and the fasteners, inspecting the fasteners after installation, sealing between the parts and around the fastener and the holes; (2) reduces mismatch of materials; and (3) eliminates arcing from the fasteners.
In the multipass process, an induction heating work coil is passed multiple times over a bond line while applying pressure in the region of the coil to the components to be welded, and maintaining the pressure until the resin hardens. The resin at the bond line is heated to the softening or melting temperature with each pass of the induction work coil and pressure is exerted to flow the softened/melted resin in the bond line and to reduce the thickness of the bond line. The pressure improves the intimacy of the faying surface contact with each pass to improve continuity of the bond. The total time at the softened or melted condition of the thermoplastic in the faying surfaces is sufficient to attain deep interdiffusion of the polymer chains in the materials of the two faying surfaces throughout the entire length and area of the bond line. The process produces a bond line of improved strength and integrity in the completed part. Dividing the time that the faying surfaces are at the melting temperature allows time for the heat in the interface to dissipate without raising the temperature of the entire structure to the degree at which it loses its strength and begins to sag. The desired shape and size of the final assembly is maintained.
A structural susceptor includes fiber reinforcement within the weld resin to alleviate residual tensile strain otherwise present in an unreinforced weld. This susceptor includes alternating layers of thin film thermoplastic resin sheets and fiber reinforcement (usually woven fiberglass fiber) sandwiching the conventional metal susceptor that is embedded in the resin. While the number of total plies in this structural susceptor is usually not critical, Boeing prefers to use at least two plies of fiber reinforcement on each side of the susceptor. This structural susceptor is described in greater detail in U.S. Pat. No. 5,717,191.
The structural susceptor permits gap filling between the welded composite laminates which tailors the thickness (number of plies) in the structural susceptor to fill the gaps, thereby eliminating costly profilometry of the faying surfaces and the inherent associated problem of resin depletion at the faying surfaces caused by machining the surfaces to have complementary contours. Standard manufacturing tolerances produce gaps as large as 0.120 inch, which are too wide to create a quality weld using the conventional susceptors.
It is easy to tailor the thickness of the structural susceptor to match the measured gap by scoring through the appropriate number of plies of resin and fiber reinforcement and peeling them off. In doing so, a resin rich layer will be on both faying surfaces and this layer should insure better performance from the weld.
b. Fixed coil induction welding
Thermoplastic welding using Boeing""s induction heating workcell differs from the moving coil processes because of the coil design and resulting magnetic field. The fixed coil workcell presents promise for welding at faster cycle times than the moving coil processes because it can heat multiple susceptors simultaneously. The fixed coil can reduce operations to minutes where the moving coil takes hours. The keys to the process, however, are achieving controllable temperatures at the bond line in a reliable and reproducible process that assures quality welds of high bond strength. The fixed coil induces currents to flow in the susceptor differently from the moving coils and covers a larger area. Nevertheless, proper processing parameters permit welding with the induction heating workcell using a susceptor at the bond line, as described in U.S. Pat. No. 5,641,422.
Another advantage with the fixed coil process is that welding can occur using the same tooling and processing equipment used to consolidate the skin, thereby greatly reducing tooling costs. Finally, the fixed coil heats the entire bond line at one time to eliminate the need for shims that are currently used with the moving coil. To control the temperature and to protect against overheating, xe2x80x9csmartxe2x80x9d susceptors are used as a retort, as the bond line susceptor material, or both.
The need for a susceptor in the bond line poses many obstacles to the preparation of quality parts. The metal which is used because of its high susceptibility differs markedly in physical properties from the resin or fiber reinforcement so dealing with it becomes a significant issue. A reinforced susceptor overcomes problems with conventional susceptors by including the delicate metal foils (0.10-0.20 inch widexc3x970.005-0.010 inch thick; preferably 0.10xc3x970.007 inch) in tandem with the warp fibers of the woven reinforcement fabric. The weave fibers hold the foils in place longitudinally in the fabric in electrical isolation from each other yet substantially covering the entire width of the weld surface. This arrangement still has adequate space between the foils for the flow and fusion of the thermoplastic resin. Furthermore, in the bond line, the resin can contact, wet, and bond with the reinforcing fiber rather than confronting the resinphilic metal of the conventional systems. There will be a resin-fiber interface with only short runs of a resin-metal interface. The short runs are the length of the diameter of two weave fibers plus the spatial gap between the weave fibers, which is quite small. Thus, the metal is shielded within the fabric and a better bond results. In this woven arrangement to foil can assume readily the contour of the reinforcement. Finally, the arrangement permits efficient heat transfer from the foil to the resin in the spatial region where the bond will form.
The reinforced susceptor might be an analog of the structural, selvaged, or tailored susceptors of Boeing""s other applications (i.e. a tape encased in resin and placed along the bond line) or may be fabricated as part of the facing plies of the prefabricated composites that abut along the bond line.
The susceptor may be a multistrip susceptor having two or more parallel foil strips that extend the full length of the strip. The foil is usually about 0.007 inch thick and each strip is about 0.10-0.20 inch wide. The strips are separated by gaps of comparable width or slightly wider dimension which we etch or ablate from a solid foil. Along the length of the susceptor, periodically, transverse spacer strips span the gap and keep the carrier strips apart. The foil can be virtually any width. It can be about two-four inches wide to match the spar cap width or might even be the full width of sheets of the composite prepreg used to form the skins. Dimensions given are typical and could be varied.
The strength and durability of adhesive bonds or thermoplastic welds connecting composite structure is improved by adding Z-pin mechanical reinforcement to the bond line. Weld strength can also be improved with a post-weld anneal to control cooling of the bond line.
4. Joining Honeycomb Sandwich Panels
Thermoplastic honeycomb sandwich panels have strength-to-weight ratios and use temperature capabilities unmatched by metals or thermoset composites. These properties make the panels ideally suited for high performance aircraft and spacecraft, especially for cryogenic tanks. But, in making large structures, often two or more subparts must be joined. The joint usually is designed to withstand twice the load of the bulk material, and is impenetrable to liquid and vapor. Traditional joints used a butt or single lap splice combined with composite doublers fastened or adhered on each side of the face sheets. The following problems characteristically plagued these joints:
1. Drilling fastener holes in the graphite reinforced thermoplastic composite was costly and greatly compromised its load carrying capabilities.
2. Fastener holes required extensive sealing to ensure liquid and vapor integrity.
3. Fasteners added significant weight.
4. Film adhesives are restricted to a limited range of temperatures.
5. In highly loaded applications, film adhesives usually age, peel, and crack.
6. Costly modifications to core ends (ramping up or down) or the addition of a backup structure, or both, were often required to support fastener installation.
7. Butt and single lap joints are inherently weak because they do not share fibers between face sheet elements. This discontinuity creates a stress concentration area with limited load carrying capabilities.
Developing a structural joint that is easily and reliable formed without autoclave processing would greatly enhance the application of thermoplastic composite sandwich panels to large structures. Such joints need to be able to withstand tensile stresses in the range of 12,000 lb/inch.
The present invention is a thermoplastic composite sandwich panel having a thermoplastic weld on at least one face sheet made without autoclave processing in a double interleaf stagger joint or equivalent configuration. The invention eliminates fasteners and produces a fluid impervious, sealed joint. The panels are particularly suited for use as cryogenic tanks in spacecraft or launch vehicles.
A double interleaf staggered joint for joining thermoplastic fiber-reinforced composites with a thermoplastic seam weld, has at least two fully compacted, laminated face sheets substantially free of volatiles having a plurality of plies of a thermoplastic resin reinforced with fibers. Each face sheet has an edge with a plurality of fingers adapted for interleaving into the joint. The face sheets also have a right hand and a left hand configuration suitable for forming the joint. Resin film between fingers on each joint mating interface insures that fusion of the fingers produces a resin-to-resin bond. Optionally, a metal foil on one surface of the face sheets in the region of the joint reinforces the joint. Each finger is a plurality of plies of fiber-reinforced resin composite with the plies staggered to reduce areas of stress concentration in the joint.
Our preferred process for making the joint includes the steps of:
(a) laying up upper and lower, left and right laminated face sheets having a plurality of plies of fiber-reinforced thermoplastic matrix resin;
(b) positioning separator plies in the edges of the face sheets that will form the joint to create an interleaf split of fiber-reinforced laminated composite fingers;
(c) consolidating the face sheets to produce composites substantially free of volatiles or porosity;
(d) bonding the upper and lower face sheets, respectively, to a honeycomb core with a resin rich layer on the surface of the face sheet that contacts the honeycomb core to form left and right honeycomb core sandwich panels;
(e) removing the separator plies;
(f) interleaving fingers of the left and right face sheets of the left and right panels above and below the core to define double staggered interleaf joints;
(g) optionally, positioning a metal foil between each face sheet and the core in the region of the joint to reinforce the joint; and
(h) melting the resin in the joints to form thermoplastic fusion welds.
Our preferred apparatus for forming a fusion weld in a joint between interleaved fingers of at least two, preconsolidated, fiber-reinforced thermoplastic resin matrix composite sandwich panels, includes clamping elements adapted for applying a pressure to the joint, and insulation associated with edges of the platen to contain heat at the joint. The elements including heated, conformal platens matching the configuration of the panels.