Patent Publication Number: US-2006010689-A1

Title: Automated drill process for two-diameter holes in multi-layer variable thickness composite materials

Description:
STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH OR DEVELOPMENT  
      This invention was developed in the course of work under U.S. government contract No. F/A-18 E/F N00019-99-C-1226; FY00-04. The U.S. government may possess certain rights in the invention. 
    
    
     FIELD OF THE INVENTION  
      This invention relates generally to the manufacture of mobile platforms and, more particularly, the assembly of skin panels to the airframes of aircraft.  
     BACKGROUND OF THE INVENTION  
      Aircraft bodies are assembled by fabricating a frame and fastening panels to the frame. Typically, the frame is an aluminum or titanium structure with ribs, stringers, and the like to distribute the loads imposed by the aircraft&#39;s weight and aerodynamic forces that act on the aircraft. The assembly process continues with panels being fastened to the structure to form the skin of the aircraft. Because these skin panel fasteners carry much of the load on the aircraft as shear stress, the fastening of the skin panels to the aircraft frame is a factor in how efficiently the aircraft carries the loads.  
      Increasingly, composite skin panels are being used to lighten the aircraft and improve its load carrying capability. Assembling the composite panels to the aircraft requires that a liquid shim be applied to the cured panel to fill any gaps that might otherwise exist between the panel and the frame. These gaps arise because the composite panels generally will not match the shape of the frame exactly. Rather, some variation will exist between the frame and the panel that may be several thousands of an inch in magnitude. Thus, the liquid shim compensates for the variation. Once applied to the composite panel, the liquid shim begins curing and eventually forms a portion of the panel. Next, the panel is mounted to the airframe and fastener holes are drilled through the panel and frame for subsequent installation of a fastener.  
      If the material of the structure and the panel are the same, the hole may be sized with one constant diameter through the two sub-assemblies. However, metals and composites behave differently when subjected to loads. For instance, interference fits are frequently selected between metallic structures and fasteners to improve fatigue life in the tensilely loaded panels and frames. On the other hand composites generally require a clearance between the composite and the fastener to prevent the composite from delaminating while installing interference fit fasteners through the composite panel.  
      To provide the interference fit and the clearance, the panel must therefore be removed from the airframe after the fastener hole is drilled to the interference diameter. The hole in the composite panel is then reamed to a slightly larger size to create the clearance. Thereafter, the panel is placed on the structure once again. Then, the fastener is placed in the hole having the two diameters and tightened into place.  
      Unfortunately, the process of providing the clearance causes several disadvantageous results. First, removing and reinstalling the panel consumes time and resources that could be employed for other useful activities. Second, because the panel has been moved after the initial hole was formed, the reamer used to enlarge the hole in the composite panel may be positioned off of the longitudinal axis of the hole. Accordingly, the reamed enlargement may be off-center, or eccentric, with respect to the axis of the hole (through the panel). Moreover, perfect re-alignment between the panel and the structure may not be re-acquired either. Further with these previous assembly methods, any interlaminar metallic burrs generated during the enlargement process must be removed manually from the airframe structure before fastening the panel to the structure.  
      In the alternative, the reamer could be brought to the panel while it is still mounted on the airframe and the hole enlarged. In practice, this alternative has produced poor assemblies because the thickness of the composite panel varies from the theoretical thickness across the panel (both by design and due to variations inherent to fabrication of composite parts). Moreover, the liquid shim applied to the panel may vary in thickness because of the gaps between the airframe and the panel and because of variations in how the liquid shim is applied. Thus, the operator (or numerically controlled machine programmer) does not know at what depth to stop the reamer before it engages the airframe. If the reamer is advanced too far, it creates a clearance within the metal and weakens the joint. If the reamer does not advance far enough, it leaves an interference fit in the composite and weakens the joint.  
      Thus a need exists to improve the assembly of composite members to metallic structures.  
     SUMMARY OF THE INVENTION  
      It is in view of the above problems that the present invention was developed. The invention provides improved mobile platform skin panel assemblies and methods of assembling the same.  
      More particularly, the present invention provides an assembly that includes a skin panel made from a composite material and a structure made from a metal. The assembly defines a hole through the panel and the structure with a first diameter through the metal and a second diameter through the panel, which has a thickness profile known in advance. The transition between the two diameters occurs within a pre-selected tolerance from the surface of the metal that is adjacent to the panel and another pre-selected tolerance from the surface of the panel that abuts the structure. In other words, the clearance reliably extends into the metal only to the extent of the first tolerance, whereas the interference fit reliably extends into the composite only to the extent of the second tolerance. Thus, the present invention provides superior joints between composite skin panels and metallic structures. In addition, the present invention provides assemblies with a structure having a first material and a member having a second material and a fastener hole therethrough, the materials requiring the hole to have two different diameters.  
      In another preferred embodiment, the present invention provides a method of assembling composite panels to metallic structures. The method includes leaving the panel on the structure while a hole, which extends through the panel and structure, is enlarged down to between a pre-selected tolerance from the surface of the panel and another pre-selected tolerance in the structure. Thus, the present invention provides a superior method of assembling composite panels to metallic structures.  
      Further features and advantages of the present invention, as well as the structure and operation of various embodiments of the present invention, are described in detail below with reference to the accompanying drawings. 
    
    
     BRIEF DESCRIPTION OF THE DRAWINGS  
      The accompanying drawings, which are incorporated in and form a part of the specification, illustrate the embodiments of the present invention and together with the description, serve to explain the principles of the invention. In the drawings:  
       FIG. 1  illustrates a wing assembly in accordance with the principals of the present invention;  
       FIG. 2  illustrates a mechanical joint of the wing of  FIG. 1  in cross section taken along the line  2 - 2 ;  
       FIG. 3  illustrates test results of mechanical joints constructed in accordance with the principles of the present invention;  
       FIG. 4  illustrates an instrument for verifying the mechanical joint of  FIG. 2 ;  
       FIG. 5  illustrates another instrument for verifying the mechanical joint of  FIG. 2 ; and  
       FIG. 6  illustrates a method in accordance with a preferred embodiment of the present invention. 
    
    
     DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS  
      Referring to the accompanying drawings in which like reference numbers indicate like elements,  FIG. 1  illustrates an assembly including a composite panel fastened to a metallic structure in accordance with the principals of the present invention. More particularly,  FIG. 1  shows a wing  10  for use on a mobile platform (e.g. an aircraft) that has a structure  12  (including, for example ribs and spars) and a plurality of skin panels  14 . The panels  14  define the outer mold line of the wing  10  and are fastened to the structure  12  with fasteners  16  (e.g. rivets) as shown. Also, a liquid shim  18  is applied to the inner surface of the panels  14 . When the panel is pressed against the structure  12 , the liquid shim  18  flows to fill gaps between the structure  12  and the panel  14 . In a few isolated areas, the panel  14  will fit directly against the structure  12  and the liquid shim  18  will essentially flow completely out of these areas. In other areas, gaps, on the order of several thousandths of an inch, will exist where the liquid shim  18  will gather. A sealant  20  is also applied to the outer surface of the structure  12  to prevent moisture and other environmental contaminants from reaching the joint (i.e. the region where the panels  14  are adjacent the structure  12  with some quantity of liquid shim  18  and sealant  20  there between) following formation of the joint. Generally, the term “outer” refers to the panel  14  side of the joint, whereas “inner” refers to the structure  12  side of the joint. Further, “depth” refers to the distance from the outer surface of the panel  14  measured generally perpendicular thereto.  
       FIG. 2  illustrates the joint between the structure  12  and the panel  14  of  FIG. 1  in cross section. In  FIG. 2 , the fastener  16  is not shown, though it fits in a hole  22 . Generally within the structure  12 , the hole  22  has a first diameter  24 , whereas generally within the panel  14 , the hole  22  has a second diameter  26 . A transition region  28  connects the portion of the hole  22  having the first diameter  24  and the portion having the second diameter  26 . While a transition  28  is shown that makes an acute angle with the axis  30  of the hole  22 , step and curvilinear transitions  28  are also within the scope of the present invention.  
      The hole  22  includes an overall depth  32  that includes a first depth  34  defined by the first diameter  24  portion, a second depth  36  defined by the second diameter  26  portion, and a third depth  38  defined by the transition  28 . Since the transition  28  may be a step, the depth  38  will henceforth be treated as being generally negligible. Similarly, the assembly  10  (e.g. the wing of  FIG. 1 ) includes an overall thickness  40  associated with the locale of the hole  22 . The overall thickness  40  includes a thickness  42  of the panel (which is measured in advance of the hole drilling and reaming method described herein), a thickness  44  of the liquid shim, a thickness  46  of the sealant, and a thickness  48  of the structure. The overall thickness  40  of the panel, the liquid shim  44 , and (to an extent) the sealant  46  vary with each hole  22  location and, in general, between panels  12 . Thus, the overall thickness  40  in the locale of each of the holes  22  is not known until measured at each hole location. For the same reason, the overall depth  32  of the holes are not known until the overall thickness  40  is measured. Various methods exist for measuring the overall thickness  40  (and therefore the overall depth  32 ) including sonar and laser-based measurements.  
      Also,  FIG. 2  shows two tolerances associated with the hole  22  including a first tolerance  50  and a second tolerance  52 . The tolerance  50  begins at the inner surface of the panel  14 , extends into the panel  14  therefrom, and defines the allowable depth to which the first diameter  24  portion of the hole  22  may extend into the panel  14 . Similarly, the tolerance  52  begins at the outer surface of the structure  12 , extends into the structure  12  therefrom, and defines the depth to which the second diameter  26  portion of the hole  22  may extend into the structure  12 . Between the tolerances  50  and  52  a variable thickness  54  portion of the joint exists defined by the local variations in the liquid shim  18  and the sealant  20 .  
      As mentioned previously, the thickness  42  of the panel  14  also varies. The variations in the thickness  42  of the panel(s) arise from the complexity of the composite panels  14  and the design and manufacturing requirements, or preferences, for a given panel. For instance, the number of plies (e.g. carbon epoxy plies) in the panel typically varies with location on the panel  14 . Also, manual lay-up methods and autoclave cure cycles are likely to cause further deviations from the theoretical thickness of the panel  14 . To account for these variations, the following data may be loaded into a database for a numerically controlled machine prior to performing the drilling and reaming operations described for the present embodiment: a hole identification number, a hole location on the panel, and the actual panel  14  thickness  42 . The methods discussed herein may also be performed manually, although automated machining is preferred. Whether the machining is performed manually or automatically, the machining parameters (e.g. drill feeds and speeds) can be changed as the tool(s) progress through the various layers of the assembly based on the stack-up information and the theoretical and actual thicknesses associated with a given hole identification number. As a result, it is possible to machine at the optimum rate for each material in the assembly and, thereby, reduce the cycle time for each hole. In particular, the machining speed may increase as progress is made through the softer materials (e.g. composites and aluminum) as compared to the speed preferred for the harder materials (e.g. titanium).  
      In accordance with the principals of the present invention, the variable thickness  54  region is defined by materials (the liquid shim  18  and the sealant  20 ) for which neither an interference fit, nor a clearance, with the fastener  16  (see  FIG. 1 ) is preferred. That is to say that the liquid shim  18  and the sealant  20  will allow for either an interference fit or a clearance with the fastener  16 . Accordingly, the transition  28  may occur anywhere within the portion of the joint defined by the combination of the tolerances  50  and  52  and the variable thickness  54 . It has been found, in experimental operation, that the second diameter  26  portion of the hole  22  can be enlarged to the second diameter  26  without exceeding either tolerance  50  and  52  by using the variable thickness  54  portion to advantage, as illustrated in  FIG. 3 .  
      More particularly,  FIG. 3  shows the results of experimental holes produced in accordance with the present invention. In particular,  FIG. 3A  shows a realistic worst-case example of a joint between a structure  112  and a composite panel  114 . In this realistic worst-case scenario, only a small gap (of up to about 0.005 inches containing liquid shim  118 ) exists between the structure  112  and the panel  114  because almost all of the liquid shim was forced from between the structure  112  and the panel  114 . Thus, the variable thicknesses employed to advantage by the present invention are at a minimum. Because experience has shown that the application of the sealant  120  is generally uniform, the present discussion will assume that the sealant is typically about 0.0056 inches thick and generally uniform. It will be understood, though, that the sealant thickness can vary. Further, the structure  112  could vary in thickness, although the structure  112  usually has a uniform thickness. However, in operation, whether the structure  112  is of uniform thickness, or not, has no bearing on the quality of joints constructed in accordance with the principles of the present invention.  
      In the example shown by  FIG. 3A , the joint includes a ⅜″ composite panel  114  and a ⅜″ aluminum structure  112 . The hole  122  was drilled through the joint with a bit to produce the interference fit diameter  124 . Tolerances  150  and  152 , of 0.010 inches each, were selected for the joint based on good engineering practice to prevent delamination of the panel  114  and to create a satisfactory interference fit with the structure  112 . Next, the hole was reamed (from the panel side of the joint with the panel  114  remaining on the structure  112 ) to enlarge the portion of the hole  122  in the panel  120  to the clearance diameter  126 . The reamer was allowed to advance into the panel  120  a distance equal to the measured thickness of the panel  114  and liquid shim  118  and then withdrawn. The hole  122  was thereafter inspected to determine where the transition  128  actually occurred with respect to the inner surface of the panel  114  and the outer surface of the structure  112 . The process was repeated for the remainder of the holes  122  desired for securely fastening the panel  114  to the structure  112 .  
      The results showed that the average transition  128  occurred within the liquid shim  118  in accordance with the principals of the present invention. That is to say, the transition  128  lies within the variable thickness of the liquid shim  118  where the joint requires neither a clearance nor an interference fit. The process was also shown to possess a Cpk (process capability index) of 1.45 that is satisfactory for most applications, even here in the realistic worst-case example.  
       FIG. 3B  shows a more typical joint. Here, the liquid shim  118  (gap) is about 0.10 inches thick. The resulting average transition  128  again fell within the variable thickness of the liquid shim  118 , thereby creating a satisfactory joint. The Cpk of the process operating on a typical joint was found to be about 2.05 and more than adequate to ensure the quality of the joint. In all cases, the depth of the reaming operation possessed a standard deviation of about 0.0028 inches using otherwise conventional drilling and reaming techniques.  
      For purposes of demonstration, an abnormal situation, wherein little (or no) liquid shim (gap) exists in the joint at the locale of the hole  122  was also tested, as shown by  FIG. 3C . The resulting transition  128  occurred at a depth 0.0012 inches into the composite panel  114 , well within the 0.010 inch tolerance  150 . With a standard deviation of 0.0028 and a Cpk of about 0.86 the vast majority of the transitions  128  were found to be within the tolerance  150  in the panel  112 , or in the sealant  120 . A few transitions  128  were found to lie just outside of the tolerance  150 . Given the number of holes  122  employed per joint, the factors of safety utilized, and the availability of inspection tools for the holes  122 , the worst-case scenario of  FIG. 3C  also produces a satisfactory joint. Of course, controlling the drilling and reaming subprocesses to a smaller standard deviation will eliminate the existence of even those few transitions  128  that lie outside of the tolerance  150 . In summary of the current embodiment, the present invention provides suitable joints even in the worst-case scenario involving no liquid shim  118  at a particular hole  122  location in the joint.  
      Moreover, because the panel  114  and the structure  112  remain generally adjacent each after the application of the liquid shim  118  and the sealant  120 , no manufacturing debris, or other contaminants, will be found in the joint. Thus, the overall joint is stronger than provided by previous methods of assembly. Moreover, the drilling and reaming may be performed by using the same conventional 6-degree of freedom robot that can remain stationary relative to the axis  130  of the hole  122  (except as it traverse the axis  130 ). Thus, eccentricity of the clearance diameter  126  portion of the hole  122  (with respect to the interference diameter  124  portion) is significantly reduced over that of previous processes. For the joints illustrated by  FIG. 3 , the vast majority of holes  122  had an eccentricity of less than about 0.0013 inches, thereby yielding improved fit between the fastener and the panel  114  and structure  112  assembly. The “one-up” method of the present embodiment is, thus, particularly well suited to applications wherein the panels  114  are sealed, or bonded, to the structure  112  and thereafter left in place for the remainder of the aircraft assembly. Of course, the term “one up” refers to one of the advantages of the present invention in that the panel  114  may be lifted “up” into place on the structure  112  (or otherwise moved into place) only once with no subsequent disassembly required.  
      In the alternative to the one-up method, the present invention may also be employed where no sealant, or bonding agent, is applied between the panel  14  and the structure. For instance, the panel  114  may be clamped to the structure  112 , an initial one-diameter hole drilled, and the hole may then be enlarged through the panel  114 . Thereafter, the panel  114  may be unclamped from the structure  112  and, if desired, removed for de-burring and other operations prior to subsequent assembly operations.  
      With respect to  FIG. 4 , a pair of instruments  200  and  300  are shown as used for inspecting holes  222  and  322  in accordance with the principles of the present invention. Generally, the instruments  200  and  300  are used to verify the end point of the clearance diameter portion of the hole. The instrument  200  includes a boroscope eyepiece at the proximal end (not shown) and a mirror  202  at the distal end. The distal end is shown inserted in the hole  222  and further includes a pair of diametrically opposed apertures  204 , an index mark  206 , and a flat distal end  208 . The flat distal end  308  is adapted to engage the transition  228  as shown. Thus, when the flat distal end  208  engages the transition  228  the instrument  200  is held in a fixed position relative to the transition  228 . Preferably the outside diameter (of at least a portion) of, the distal end is the same as the clearance diameter  226 , thereby further facilitating holding the instrument  200  in fixed relationship to the transition  228 . The mirror  202  is positioned so that it allows a user looking through the eyepiece to see out through the aperture  204  and inspect the side of the hole  222  in the vicinity of the transition  228 . In particular, the index mark  206  is also viewable by the user and set (inside the distal end of the instrument  200 ) at a distance from the distal end equal to the tolerance  252  associated with the structure  212 . Thus, as the user views the side of the hole  222 , via mirror  202 , the user can judge whether the index mark  206  is deeper than the outside surface of the structure  212 . If so, the clearance diameter  126  portion penetrates too deeply into the structure  112  indicating an unacceptable hole  222 . Such a situation is illustrated in  FIG. 4B . In contrast,  FIG. 4A  shows a hole  222  wherein the index mark  206  is at the outer surface of the structure  222 . Thus,  FIG. 4A  illustrates an acceptable hole  222 . Accordingly, the instrument  200  is referred to as a “too deep” gage  200 .  
       FIGS. 4C and 4D  illustrate the corresponding “too shallow” gage  300  in a hole  322  that is acceptable ( FIG. 4C ) and in an unacceptable hole  322  ( FIG. 3D ). The differences between the “too deep” gage  200  and the “too shallow” gage  300  include the presence of a step  308  on the sides of the distal end and the location of the index mark  306 . The step  308  is adapted to engage the transition  328  and generally corresponds in shape to the transition  328 . Thus, the index mark  306  may be positioned at a distance equal to the tolerance  350  associated with the panel  314  from the step  308  as shown. A user may view the index mark  306  and the side of the hole  322  via the mirror  302  to judge whether any of the liquid shim  318 , sealant  320 , or structure  312  is visible between the index mark  306  and the distal end. If so, the hole  322  is deep enough (i.e. is within or deeper than the tolerance  350 ). If not, then the transition  328  is too shallow (i.e. the clearance diameter  326  portion of the hole  322  does not extend far enough through the panel  320  to produce a satisfactory joint).  
      While the exemplary fastener holes previously discussed were generally orientated perpendicularly to the mating surfaces of the structure  12  and panel  14  (see  FIG. 2 ), the present invention is not thereby limited. Rather, in a preferred embodiment, a joint with a “ramped” pair of mating surfaces is provided. That is to say that the axis  30  (see  FIG. 2 ) of the hole is orientated at an acute angle with respect to the mating surfaces. In the present embodiment, the depth  36  of the clearance diameter  26  portion of the hole  22  is measured along the axis  30  of the hole  22 . Thus, the intersection of the axis  30  with the plane defining the end of the clearance diameter  26  portion lies within the acceptable depth range  56 . Ramped holes  22  are inspected by, for example, rotating the “too deep/too shallow” gage  400  in the hole  22  until the views through the pair of diametrically opposed apertures  404  are the same. If both views are neither too deep, nor too shallow, the hole  22  is considered acceptable.  
      Lighting may also be provided internal to the instruments  200  and  300  to enable the user to view the visible differences between the structure  212 , the panel  214 , the liquid shim  218 , and the sealant  220 . The instruments  200  and  300  also enable the user to inspect the sides of the hole  222  to determine whether any chips or cuttings were caught between the outer pitch of the drill bit (and reamer) and the sides of the hole  322  by turning the instrument while traversing the axis of the hole  322 . Thus, the composite panels  220  and  320  may be inspected for internal machining damage that would otherwise be hidden.  
       FIG. 5  shows a combined “too deep/too shallow” instrument  400 . The instrument  400  differs from the instruments  200  and  300  in that, the instrument  400  includes both of the index marks  406  and the step  408 . In particular, the index mark  406 ′ is positioned a distance from the step  408  equal to the tolerance  452  associated with the structure  412 . Thus, it indicates whether the hole  422  is too deep. The index mark  406 ′, on the other hand, is positioned a distance equal to the clearance  450  associated with the panel  414  from the distal end, thereby indicating whether the transition is too shallow. Accordingly, the one tool  400  may be used to simultaneously determine whether the hole  422  is deep enough and whether the hole  422  is shallow enough (i.e. within an acceptable range of depth).  
      In another preferred embodiment, the present invention provides a dial indicator for inspecting the depth of the transitions. The dial indicator includes a plunger operatively connected to a depth dial gage. Further, the end of the plunger is adapted to engage the transition of a hole, thereby enabling the inspection. To inspect a hole, the indicator is zeroed by fully depressing the plunger against a hard surface. Then the plunger is inserted into the hole and allowed (by a biasing member such as a spring) to extend to the depth at which it stops. Generally, the depth at which the plunger stops indicates the location of the transition. However, debris in the hole may cause the dial indicator to indicate a transition depth shallower than the true transition depth. Also, erosion (particularly of the liquid shim and sealant) caused by chips being caught between the drill bit, or reamer, during the machining of the hole, may allow the plunger to extend beyond the true transition, thereby indicating a transition depth larger than the true transition depth. Accordingly the boroscope based instruments  200 ,  300 , and  400  are preferred over the dial indicator of the current embodiment.  
      With reference now to  FIG. 6 , a method in accordance with yet another preferred embodiment is illustrated. As shown, the method  500  includes laying up a plurality of composite panels and subsequently curing them in operation  502 . Usually in parallel with operation  502 , an airframe is fabricated (or assembled) as in operation  504 . Liquid shim is then applied to the cured panels. Before the liquid shim cures, the panels are pressed against the airframe to cause the liquid shim to fill the gaps there between. See operation  506 . The liquid shim is allowed to cure. Before final assembly of the panels to the airframe, operation  508  applies sealant to the airframe. Thereafter, the panels are mounted (with jigs or other support equipment) to the airframe as in operation  510 . Holes having a diameter that will cause an interference fit with the fasteners are drilled through the panels and through the structure at pre-selected locations. See operation  512 . While leaving the panels on the airframe (as in operation  514 ) the holes are then reamed to the larger clearance diameter to a depth that is pre-selected to cause an acceptable joint in operation  516 . While many types of tools may be employed to enlarge the hole, either step or flat-bottom reamers are employed in preferred embodiments of the present invention. Depending on the tool selected, the set-up values (e.g. step reamer lengths—to the step and the total length) are measured in advance and included in the machining program to enable the machine to ream the hole to the specified depth (e.g. to the bottom of the flat reamer or to the step of the step reamer). Of course, the overall thickness of the panel(s) is known or measured before pre-selecting the depth of the enlargement. At reference  518 , the holes are then inspected to determine whether the diameter transitions are within the acceptable ranges (neither too deep, nor, too shallow) with, for example, the instruments provided by the present invention. Fasteners are inserted into the holes and the panels are fastened to the structure as shown at operation  520 . Thus, the superior joints discussed herein may be assembled by use of the present embodiment.  
      In view of the foregoing, it will be seen that the several advantages of the invention are achieved and attained. In particular, satisfactory joints are provided with the panels remaining on the airframe once placed thereon. Thus, the present invention reduces the cost of assembling aircraft. For the same reason, the present invention provides joints having superior mechanical properties (e.g. strength, fit, noise or rattling because of poor “fit up”). Additionally, the present invention provides improved inspection tools over those previously available.  
      The embodiments were chosen and described in order to best explain the principles of the invention and its practical application to thereby enable others skilled in the art to best utilize the invention in various embodiments and with various modifications as are suited to the particular use contemplated.  
      As various modifications could be made in the constructions and methods herein described and illustrated without departing from the scope of the invention, it is intended that all matter contained in the foregoing description or shown in the accompanying drawings shall be interpreted as illustrative rather than limiting. For example, whereas the foregoing discussion involved composite members being assembled to metallic structures, the present invention is not limited thereby. Rather, any assembly with materials requiring having fastener holes with two different diameters is within the scope of the present invention (e.g. titanium and aluminum). Thus, the breadth and scope of the present invention should not be limited by any of the above-described exemplary embodiments, but should be defined only in accordance with the following claims appended hereto and their equivalents.