Abstract:
Shims used to join part assemblies are automatically designed and fabricated without the need for fitting part assemblies together in order to determine the exact dimensions of voids filled by the shims. The locations of key features on part assemblies are surveyed using a merged photogrammetry and laser tracking technique that generate the dimensions of a virtual shim. The dimensions of the virtual shim are contained in a digital file that can be used to automatically fabricate the shim using automated fabrication equipment such as a CNC machining center. The automated virtual shim design may be modified to reflect the effect of part assembly fit on performance characteristics of the aircraft.

Description:
TECHNICAL FIELD 
       [0001]    This disclosure generally relates to manufacturing processes used to join parts, and deals more particularly with a method for fitting, aligning and joining large, complex part assemblies. 
       BACKGROUND 
       [0002]    Shims are commonly used in fitting and assembling parts and subassemblies in order to compensate for dimensional variations. In the aircraft industry, shims are used extensively in fitting and joining fuselage sections, and in attaching wings and tail assemblies (vertical fin and horizontal stabilizer assemblies) to the fuselage. The shims, sometimes referred to as fillers, are used to fill voids between the joined assemblies which may be caused by tolerance build up in parts. The use of shims to fill voids between mating surfaces on part assemblies results in a more structurally sound aircraft. Shims are also used to bring parts into proper alignment. 
         [0003]    Design and fabrication of unique shims for each aircraft can be a time consuming and labor intensive process. A skilled technician must manually measure and record each void in order to determine the dimensions and shape of a particular shim that will fill the void. The recorded dimensions are then sent to a machine shop where the shim is fabricated. 
         [0004]    The shim design and installation process described above may materially slow down aircraft assembly, especially where the assemblies are manufactured in different geographic locations and are shipped to a final assembly location. This is due, in part, to the fact that the shims cannot be designed and manufactured until the assemblies are fitted together at the final assembly destination so that the size and shape of the voids can be determined. 
         [0005]    Efforts have been made to reduce the time required for determining shim dimensions as exemplified in U.S. Pat. No. 6,618,505 issued Sep. 9, 2003 and assigned to the Boeing Company. This prior patent discloses a method and apparatus for determining the dimension of a shim, using digital photogrammetry to measure the profile of the voids requiring shims. The shim dimensions are calculated based on the void profile measurements referenced against an engineering standard defining an ideal fit between the assemblies. While this prior process reduces the time required for shim design, further efficiency improvements are possible. 
         [0006]    Accordingly, a need exists for a method for fitting and joining part assemblies in which the shims are automatically designed and fabricated without the need for physically fitting the part assemblies to determine the location and profile of potential voids. Embodiments of the disclosure are directed towards satisfying this need. 
       SUMMARY 
       [0007]    Illustrated embodiments of the disclosure provide a method of automatically designing and fabricating shims without the need for joining part assemblies in order to determine the exact dimensions of voids filled by the shims. The locations of key features on part assemblies are surveyed using a merged photogrammetry and laser tracking technique that generate the dimensions of a virtual shim. The dimensions of the virtual shim are contained in a digital file that can be used to automatically fabricate the shim using automated fabrication equipment such as a CNC machining center. The automated virtual shim design may be modified to reflect the effect of part assembly fit on performance characteristics of the aircraft. For example, the virtual shim dimensions can be adjusted to alter the incidence, sweep, or dihedral of wings relative to a fuselage. 
         [0008]    According to one embodiment of the disclosure, a method is provided for fitting two parts together, comprising the steps of: measuring the location of a first set of features on a first part; measuring the location of a second set of features on a second part; generating a virtual fit between the first and second parts based on the location measurements; and, generating dimensions of shims to be positioned between the first and second parts based on the generated virtual fit. Feature location measurement may be performed using both laser tracker and photogrammetry processes. Generating the virtual fit may include performing a virtual nominal fit and then optimizing the virtual nominal fit. The virtual fit may also include generating computer models of the first and second parts and then comparing the computer models to determine the shape of voids requiring shims. 
         [0009]    According to another embodiment, a method is provided for producing shims used in fitting aircraft part assemblies together. The method includes the steps of: generating first and second sets of data respectively representing the location of features on first and second part assemblies; performing a virtual fit between the first and second part assemblies using the first and second sets of data; analyzing characteristics of the aircraft based on the virtual fit; modifying the virtual fit based on the results of the analysis; generating the dimensions of at least one shim based on the modified virtual fit; and, fabricating the shim using the generated dimensions. One of the part assemblies may comprise a wing and the analyzed characteristics may include one or more of the angle of incidence of the wing, the sweep angle of the wing or the dihedral of the wing. The generated dimensions of the shim may include generating a set of digital data representing the dimensions, and the fabricating step may include using the digital data set to control a machine used to fabricate the shim. Performing the virtual fit may include providing a set of data representing a nominal fit between the first and second part assemblies, including key geometric features, and aligning the key geometric features of the first and second part assemblies. The virtual fit may also include aligning certain features in a first set of features on the first and second part assemblies, and then performing a best fit between features in a second set of features on the first and second part assemblies. 
         [0010]    In accordance with still another embodiment, a method is provided for manufacturing an aircraft comprising the steps of: manufacturing a first part assembly; generating a first set of data representing the position of features on the first part assembly; manufacturing a second part assembly; generating a second set of data representing the position of features on the second part assembly; performing a virtual fit between the first and second part assemblies using the first and second sets of data; generating the dimensions of shims used to fit the first and second part assemblies together based on the virtual fit; fabricating shims based on the generated dimensions; and, assembling the first and second part assemblies using the fabricated shims. The first and second part assemblies may be manufactured respectively in first and second geographic locations, and the final assembly step may be performed in a third geographic location. The method may further include the steps of analyzing characteristics of the aircraft based on the virtual fit and then modifying the virtual fit based on the results of the analysis. The step of performing the virtual fit may include aligning the features in a first set of features on the first and second part assemblies, and performing a best fit between features in a second set of features on the first and second part assemblies. 
         [0011]    In accordance with still another embodiment of the disclosure, a method is provided for manufacturing an aircraft, comprising the steps of: fabricating a first part assembly in a first manufacturing process; generating a first set of data representing the position of features on the first part assembly; fabricating a second part assembly in a second manufacturing process; generating a second set of data representing the position of features on the second part assembly; performing a virtual fit between the first and second part assemblies using the first and second sets of data; analyzing characteristics of the aircraft based on the virtual fit; modifying the virtual fit based on the results of the analysis; and, altering at least one of the first and second manufacturing processes based on the results of the modified virtual fit. The first and second part assemblies may be manufactured in differing geographic regions. 
         [0012]    Other features, benefits and advantages of the disclosed embodiments will become apparent from the following description of embodiments, when viewed in accordance with the attached drawings and appended claims. 
     
    
     
       BRIEF DESCRIPTION OF THE ILLUSTRATIONS 
         [0013]      FIG. 1  is a side view illustration of two aircraft fuselage sections being fitted together. 
           [0014]      FIG. 2  is a plan view illustration of a wing assembly being fitted to a section of the fuselage shown in  FIG. 1 . 
           [0015]      FIG. 3  is a perspective illustration showing a wing assembly being fitted to a fuselage. 
           [0016]      FIG. 4  is a diagrammatic illustration showing in cross section, the major components joined between the wing assembly and fuselage shown in  FIG. 3 . 
           [0017]      FIG. 5  is a side view illustration of an aircraft showing potential adjustments in the angle of wing incidence using shims. 
           [0018]      FIG. 6  is a side view illustration of a section of the fuselage showing key reference points used to adjust the angle of incidence of the wing shown in  FIG. 5 . 
           [0019]      FIG. 7  is a sectional view taken along the line  7 - 7  in  FIG. 6 . 
           [0020]      FIG. 8  is a perspective illustration of a typical shim. 
           [0021]      FIG. 9  is a perspective illustration of an alternate form of a shim. 
           [0022]      FIG. 10  is a plan view illustration showing a potential range of adjustment in the sweep angle of a wing using shims. 
           [0023]      FIG. 11  is a sectional illustration showing the attachment of a wing to the fuselage, and depicting the use of a shim. 
           [0024]      FIG. 12  is a frontal illustration of an aircraft, depicting the range of adjustment of the dihedral of the wings using shims. 
           [0025]      FIG. 13  is an enlarged illustration of a wing attached to a fuselage, showing alternate dihedral positions of the wing, and key reference points used to adjust the dihedral angle. 
           [0026]      FIG. 14  is a perspective view of the inboard end of a wing and combined laser tracking/photogrammetry equipment used to survey the location of features on the wing. 
           [0027]      FIG. 15  is an enlarged illustration of a section of the wing shown in  FIG. 14 , better depicting two reflective targets used in the feature location survey process. 
           [0028]      FIG. 16  is a block diagram of a process used to assemble an aircraft, including an automated shim dimension method. 
           [0029]      FIG. 17  is a simplified flow chart of a method and software used in the automated shim dimension process shown in  FIG. 16 . 
           [0030]      FIG. 18  is a flow diagram of a method for assembling an aircraft in which part assemblies are fabricated in differing geographic locations. 
       
    
    
     DETAILED DESCRIPTION 
       [0031]    Referring first to  FIGS. 1-3 , embodiments of the disclosure relate to a method and manufacturing process for fitting and attaching parts, or assemblies of parts. As used herein the term “parts” or “part assemblies” is intended to include a wide range of structures and components that are to be fitted and or joined together, and may comprise individual parts, assemblies of parts or subassemblies. The method is particularly useful in fitting relatively large complex parts or part assemblies in which gaps or voids may be present between the assembled parts that require the use of shims to fill these voids. In the illustrated embodiments, the parts comprise large assemblies used in constructing aircraft, however it is to be understood that the method and process may be employed and fitting various other types of part assemblies for a wide range of applications. 
         [0032]    Commercial aircraft  20  are typically manufactured by assembling large, modular sections. In  FIG. 1 , two fuselage sections  22   a ,  22   b  carried on wheel lift systems  26  are moved into end-to-end contact and are joined together using various types of fasteners and connections. This joining and attachment process includes the need to fit certain mating parts of the two fuselage sections  22   a ,  22   b  together. Because of accumulated or “stacked” tolerances in the parts forming each of the fuselage sections  22   a ,  22   b , mating portions of the sections  22   a ,  22   b  may not be perfectly fitted, resulting in gaps or voids between the two mating surfaces. These voids must be filled with later discussed shims in order to assure that the two sections  22   a ,  22   b  have sufficient structural integrity at the joints between them. 
         [0033]    As shown in  FIG. 2 , a starboard wing assembly  24  supported on wheel lifts  26  is moved into position for attachment to one of the fuselage sections  22   a .  FIG. 3  also shows the wing assembly  24  having been moved into position relative to the fuselage  22 , ready for attachment. The fuselage  22  is supported on body cradles  28  that are moveable along a production line track  30 . The wing assembly  24  is supported on positioners  34  which are capable of adjusting the position of the wing assembly  24  along X (fore and aft), Y (inboard-outboard) and Z (up and down) directions so that the wing assembly  24  is properly positioned when the attachment process is completed. A laser tracker  22  or similar non-contact measuring device is used to assess the position of key reference points on the wing assembly  24  and the fuselage  22  during the final fitting process. A computer based controller  38  may receive measurement data collected by the laser tracker  22  and is operative to control the positioners  34  during the final fitting process. 
         [0034]    Although not specifically shown in the drawings, the vertical fin and horizontal stabilizers (not shown) are fitted and attached to the fuselage  22  in a manner similar to that of the wing assembly  24 . 
         [0035]    The wing assembly  24  is attached to the fuselage  22  by laterally extending mating components of the wing assembly  24  and the fuselage  22 . These mating components, which must be fitted together in a desired alignment, are diagrammatically shown in  FIG. 4 . Laterally extending components of the fuselage  22  referred to as “stub” components are shown in cross hatch. The stub components of the wing assembly include an upper flange  48 , lower flange  50 , forward spar terminal fitting  52  and rear spar terminal fitting  54 . These stub components respectively mate with wing components comprising an upper wing panel  40 , lower wing panel  42 , wing forward spar  44  and wing rear spar  46 . 
         [0036]    The accumulated tolerances in the mating components discussed immediately above are such that gaps between these two sets of components may be present. These gaps allow slight movement or adjustment of the wing assembly  24  relative to the fuselage  22  along any of three axes: X (fore and aft), Y (inboard-outboard) and Z (up and down). In the embodiment shown in  FIG. 4 , a gap  60  is present between the wing forward spar  44  and the forward spar terminal fitting  52 . Similarly, a gap  62  is present between the lower flange  50  and the lower wing panel  42 . These two gaps  60 ,  62  require the introduction of shims in order to fill the gaps and fix the final position of the wing assembly  24  relative to the fuselage  22 . 
         [0037]    Referring now to  FIGS. 5-8 , the angle of incidence  64  of the wing assembly  24  depends on the fit between the components of the fuselage  22  and wing assembly  24  discussed earlier with reference to  FIG. 4 . The angle of incidence  64  may be adjusted during the final fitting and attachment process using shims  72  to fill the gaps. The exact dimensions and shape of the shims  72  are determined according to a method that will be discussed later below, however for purposes of this description, a flat, rectangularly shaped shim  72  ( FIG. 8 ) is shown. 
         [0038]    Adjusting the angle of incidence  64  of the wing assembly  24  is carried out using measurements of the positions of reference points, such as the two reference points  66  shown in  FIG. 6 . A line connecting the reference points  66  forms an angle relative to horizontal equal to that of the angle of incidence  64 . The relative position of the reference points may be measured using a variety of techniques, however will be discussed later, laser tracking and/or photogrammetry techniques are particularly useful in performing these measurements. The upper and lower ends of the rear spar terminal fitting  54  are received within upper rear and lower rear cords  48 ,  50  respectively. Upper and lower stub panels  70 ,  76  are respectively connected to cords  48 ,  50 . A splice plate  78  covers a splice in the stub lower panel  76 . The backside of the rear spar terminal fitting  54  is secured to a stub rear spar web  80 . 
         [0039]    The shims  72  fill gaps between the cords  48 ,  50  and the rear spar terminal  54  depending upon the size of the gaps, and the dimensions of the shims  72 , the angle of incidence  64  of the wing assembly  24  may be adjusted. 
         [0040]    Although flat, rectangularly shaped shims  72  are often used in fitting and joining aircraft assemblies, the shims  72  may be of any various profiles, shapes and dimensions. For example, as shown in  FIG. 9 , a shim  72   a  is rectangularly shaped in footprint, but is wedge shaped in cross section. 
         [0041]      FIG. 10  shows a port wing assembly in various angles of sweep  67 . The sweep angle  67 , which is determined by measurement of reference points  66 , can be adjusted using the shims  72 , with the thickness of the shim  72  affecting the sweep angle  67 . 
         [0042]      FIG. 11  shows the use of a shim  72  for connecting the components of the wing assembly  24  with the fuselage  22 . The wing assembly  24  is connected to the fuselage  22  using an upper, double plus chord  48  and a lower chord  50 . Chords  48 ,  50  are connected together through a web  94  and stiffener  92 . The wing assembly  24  includes upper and lower panel stringers  82 ,  84  respectively. The upper panel stringer  82 , which is covered by panel  40 , is secured to tabs on the upper chord  48  by means of fasteners  51 . The lower panel stringer  84 , which is covered by lower panel  42 , is connected through paddle fittings  90  and fasteners  51  to a tab on the lower chord  50 . The fuselage  22  includes upper and lower panel stringers  86 ,  88  respectively. The upper panel stringer  86  is secured by fasteners  51  to the upper chord  48 . The lower panel stringer  88  is attached via a paddle fitting  90  and fasteners  51  to a tab on the lower chord  50 . A tab on the upper double chord  48  is secured to a stringer  98  on the fuselage  22  by means of fasteners  53 . A body skin  100  is also secured to a tab on the upper chord  48 , and is reinforced by a strap  102 . 
         [0043]    As shown in  FIGS. 12 and 13 , shims  72  can be used to adjust the dihedral angle  69 . Three positions of the wing assembly are shown in  FIG. 13 , respectively designated by the numerals  24 ,  24   a  and  24   b . The dihedral angle  69  is adjusted using a pair of reference points  66  which define the dihedral angle  69 . 
         [0044]    Attention is now directed to  FIGS. 14 and 15  which depict the use of non-contact measuring equipment to measure the three dimensional position of parts or features of the wing assembly  24  as well as the fuselage  22 .  FIG. 14  depicts the use of both a laser tracker  104  and photogrammetry apparatus  106  for measuring features such as reference points  66  on the wing assembly  24 . In the illustrated embodiment, a merged photogrammetry and laser tracking technique is used to determine the special location of laser targets such as targets  66   a  and  66   b  shown in  FIG. 15 . The two sets of measurement data generated by the laser tracker  104  and photogrammetry equipment  106  are loaded into a computer (not shown) and are combined using commercially available spatial analyzer software. 
         [0045]    The merged laser tracker and photogrammetry technique mentioned above is described in more detail in U.S. patent application Ser. No. 11/518,417, filed Sep. 8, 2006, assigned to the Boeing Company, the entire contents of which are incorporated by reference herein. Some of the reflective targets such as target  66   b  shown in  FIG. 15  may be coded, by uniquely arranging reflective squares and dots  69  which may be “read” by a computer to uniquely identify the position of the targets  66   b . For example, the uniquely positioned targets  66   b  can be used to establish the position of the reference points  66  shown in  FIGS. 6 ,  10  and  13 . It should be noted here that although a merged laser tracker/photogrammetry technique has been illustrated to locate key features which determine the fit between the wing assembly  24  and the fuselage  22 , a variety of other contact and non-contact technologies can be used to develop digital data sets representing the location of parts or features on the wing assembly  24  and the fuselage  22 , in a common coordinate system. 
         [0046]    Reference is now made to  FIGS. 16 and 17  which depict the steps and related software flow charts for joining and fitting large complex part assemblies such as the previously described attachment of fuselage sections  22   a ,  22   b  and wing assemblies  24 . As shown at  108 , the laser/photogrammetry process  108  is used to measure the spatial position of the fuselage sections at  114  and the wing assembly  24  shown at  116 . A set of data is generated that defines airplane configuration model based definition at  118 . The configuration definition at  118  essentially comprises nominal design information for the aircraft including data which may include tolerances and ranges for key parameters, such as wing inclination, sweep and dihedral. The configuration definition data  118  is combined with the spatial position data generated at  118 , and is used in an automated shim dimension process  110 . 
         [0047]    The shim dimension process  110  begins by performing a virtual nominal join at  120 . The virtual join  120  essentially comprises an initial virtual fit between the assemblies to be joined, using the configuration definition data  118 . Then, at step  122 , the initial virtual join or fit performed at  120  is optimized, again using the configuration definition data  118 . The optimization performed at  122  may include analyzing the structural and aerodynamic relationships between various assemblies on the aircraft so that flight performance is optimized within the airplane configuration definition  118 . For example, the inclination, sweep and dihedral of the wing assembly  24  may be adjusted within certain ranges determined by the configuration definition  118  in order to optimize aircraft performance. Then, at step  124 , virtual shim measurements are calculated to determine the size (dimensions) and shape of the shims required to fill voids or gaps between the assemblies, based on the optimized fit completed at step  122 . 
         [0048]    The details of the automated shim dimension process  110  are shown in  FIG. 17 . The virtual nominal join or fit process  120  requires the generation and loading of engineering models for the assembly fit, which comprises nominal fit data. The measurement data generated by the merged laser/photogrammetry process  108  ( FIG. 16 ) is imported as three dimensional data into a CAD program such as CATIA at  148 . The fitting process includes the alignment of key geographic features which are typically fixed at  150 . The parameters used in the virtual join process are optimized at  152  in order to obtain a best fit. The preliminary, virtual nominal join or fit data is then used in a process for optimizing the structural and aero relationships at  122 . 
         [0049]    The preliminary virtual fit is initially optimized using the configuration definition data  118  ( FIG. 16 ), resulting in a set of interim data  156  that is then analyzed at  158 . At  160 , a determination is made of whether the analyzed results are valid. If the results are valid, the purposed fit is accepted and data representing this fit is stored at  170 . However, if the analyzed results are not valid, a determination is made at step  162  of whether the fit may be corrected. If the fit is not correctable, the fit results may be referred to an authority for determining corrective action, such as the manufacturing review board  168 . However, if the results appear to be correctable, the optimization parameters are revised at step  164  and a determination is made at  166  whether to approve the revised optimization parameters. If approval is obtained at step  166 , the optimization of the fit is repeated at step  154  using revised optimization parameters. 
         [0050]    When the fit is accepted at step  170 , a set of data is developed and stored at step  174  comprising empirical shim data and alignment data. The data developed at step  174  may be used in improving the process for generating shim dimensions for future assemblies, and to alter manufacturing processes used to produce subsequent part assemblies so as to reduce the size or number of gaps and potentially eliminate the gaps, thus eliminating the need for shims. The accepted fit data is used to create shim models at  172  which may be stored as CAD shim models at  176 . The shim models  176  may be automatically delivered as digital data files to equipment (not shown) such as a CNC machining center which automatically machines the shims  72  to the dimensions which fill gaps based on the accepted fit at  170 . 
         [0051]    Attention is now directed to  FIG. 18  which depicts the steps of fitting and assembling wing and fuselage assemblies that have been fabricated in different geographic locations. The wing is assembled at step  184 , following which a survey is performed to measure the location of features on the assembled wing at  186 , using, for example, the merged laser tracker photogrammetry technique described earlier. At step  188 , the feature location data is transmitted to a second geographical location  180  where this, along with other data relating to the location of features on the fuselage, nominal engineering data, etc are loaded at  194 . An initial virtual fit is performed at  196 , following which the virtual fit is optimized at  198 , as described earlier. The optimized virtual fit data is transmitted back to the first geographic location  178  where modifications to the wing assembly are carried out, if required. At step  92 , the wing assembly is shipped to the final assembly location  180 . 
         [0052]    At a second geographic location  182 , the fuselage is assembled at step  204 , following which a survey is made to; measure the location of fuselage features at  206  using the previously described laser tracker/photogrammetry techniques. At step  208 , the surveyed location data is transmitted to the final assembly location  180  and is used as part of the data loaded at  194  employed to carry out the initial virtual fit at  196 . The optimized virtual fit information is transmitted back to the fuselage assembly location  182  where it is used to carry out any modification of the fuselage, if required. At step  212 , the fuselage assembly is shipped to the final assembly location  180  based on the optimized virtual fit at  198 , shims are fabricated at step  200  which are then used to assemble the wing and fuselage at step  202 . 
         [0053]    From the above, it may be appreciated that large, complex assemblies such as the wings and fuselage of an aircraft may be fabricated at different manufacturing sites, and that the shims required to fit and join these assemblies can be fabricated in advance of the arrival of the subassemblies at the final assembly site  180 . Thus, measurements and the generation of shim data need not be delayed until the assemblies can be physically fitted to determine the size and location of gaps and voids which need to be shimmed. Instead, the generation of an optimized, virtual fit between the assemblies allows the shims to be dimensioned and fabricated so as to carry out just-in-time assembly at the final assembly location  180 . It should be noted here that although performing the steps of loading the data  194 , performing the virtual fit  196  and optimizing the fit at  198  have been indicated as being carried out at the final assembly site  180 , these steps may be performed at any location, in which case the final shim dimensions are delivered to the final assembly site  180  where the shims are fabricated at step  200 . 
         [0054]    Although the embodiments of this disclosure have been described with respect to certain exemplary embodiments, it is to be understood that the specific embodiments are for purposes of illustration and not limitation, as other variations will occur to those of skill in the art.