Method for fitting part assemblies

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.

TECHNICAL FIELD

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

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.

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.

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.

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.

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

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.

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.

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.

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.

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.

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.

DETAILED DESCRIPTION

Referring first toFIGS. 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.

Commercial aircraft20are typically manufactured by assembling large, modular sections. InFIG. 1, two fuselage sections22a,22bcarried on wheel lift systems26are 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 sections22a,22btogether. Because of accumulated or “stacked” tolerances in the parts forming each of the fuselage sections22a,22b, mating portions of the sections22a,22bmay 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 sections22a,22bhave sufficient structural integrity at the joints between them.

As shown inFIG. 2, a starboard wing assembly24supported on wheel lifts26is moved into position for attachment to one of the fuselage sections22a.FIG. 3also shows the wing assembly24having been moved into position relative to the fuselage22, ready for attachment. The fuselage22is supported on body cradles28that are moveable along a production line track30. The wing assembly24is supported on positioners34which are capable of adjusting the position of the wing assembly24along X (fore and aft), Y (inboard-outboard) and Z (up and down) directions so that the wing assembly24is properly positioned when the attachment process is completed. A laser tracker22or similar non-contact measuring device is used to assess the position of key reference points on the wing assembly24and the fuselage22during the final fitting process. A computer based controller38may receive measurement data collected by the laser tracker22and is operative to control the positioners34during the final fitting process.

Although not specifically shown in the drawings, the vertical fin and horizontal stabilizers (not shown) are fitted and attached to the fuselage22in a manner similar to that of the wing assembly24.

The wing assembly24is attached to the fuselage22by laterally extending mating components of the wing assembly24and the fuselage22. These mating components, which must be fitted together in a desired alignment, are diagrammatically shown inFIG. 4. Laterally extending components of the fuselage22referred to as “stub” components are shown in cross hatch. The stub components of the wing assembly include an upper flange48, lower flange50, forward spar terminal fitting52and rear spar terminal fitting54. These stub components respectively mate with wing components comprising an upper wing panel40, lower wing panel42, wing forward spar44and wing rear spar46.

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 assembly24relative to the fuselage22along any of three axes: X (fore and aft), Y (inboard-outboard) and Z (up and down). In the embodiment shown inFIG. 4, a gap60is present between the wing forward spar44and the forward spar terminal fitting52. Similarly, a gap62is present between the lower flange50and the lower wing panel42. These two gaps60,62require the introduction of shims in order to fill the gaps and fix the final position of the wing assembly24relative to the fuselage22.

Referring now toFIGS. 5-8, the angle of incidence64of the wing assembly24depends on the fit between the components of the fuselage22and wing assembly24discussed earlier with reference toFIG. 4. The angle of incidence64may be adjusted during the final fitting and attachment process using shims72to fill the gaps. The exact dimensions and shape of the shims72are determined according to a method that will be discussed later below, however for purposes of this description, a flat, rectangularly shaped shim72(FIG. 8) is shown.

Adjusting the angle of incidence64of the wing assembly24is carried out using measurements of the positions of reference points, such as the two reference points66shown inFIG. 6. A line connecting the reference points66forms an angle relative to horizontal equal to that of the angle of incidence64. 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 fitting54are received within upper rear and lower rear cords48,50respectively. Upper and lower stub panels70,76are respectively connected to cords48,50. A splice plate78covers a splice in the stub lower panel76. The backside of the rear spar terminal fitting54is secured to a stub rear spar web80.

The shims72fill gaps between the cords48,50and the rear spar terminal54depending upon the size of the gaps, and the dimensions of the shims72, the angle of incidence64of the wing assembly24may be adjusted.

Although flat, rectangularly shaped shims72are often used in fitting and joining aircraft assemblies, the shims72may be of any various profiles, shapes and dimensions. For example, as shown inFIG. 9, a shim72ais rectangularly shaped in footprint, but is wedge shaped in cross section.

FIG. 10shows a port wing assembly in various angles of sweep67. The sweep angle67, which is determined by measurement of reference points66, can be adjusted using the shims72, with the thickness of the shim72affecting the sweep angle67.

FIG. 11shows the use of a shim72for connecting the components of the wing assembly24with the fuselage22. The wing assembly24is connected to the fuselage22using an upper, double plus chord48and a lower chord50. Chords48,50are connected together through a web94and stiffener92. The wing assembly24includes upper and lower panel stringers82,84respectively. The upper panel stringer82, which is covered by panel40, is secured to tabs on the upper chord48by means of fasteners51. The lower panel stringer84, which is covered by lower panel42, is connected through paddle fittings90and fasteners51to a tab on the lower chord50. The fuselage22includes upper and lower panel stringers86,88respectively. The upper panel stringer86is secured by fasteners51to the upper chord48. The lower panel stringer88is attached via a paddle fitting90and fasteners51to a tab on the lower chord50. A tab on the upper double chord48is secured to a stringer98on the fuselage22by means of fasteners53. A body skin100is also secured to a tab on the upper chord48, and is reinforced by a strap102.

As shown inFIGS. 12 and 13, shims72can be used to adjust the dihedral angle69. Three positions of the wing assembly are shown inFIG. 13, respectively designated by the numerals24,24aand24b. The dihedral angle69is adjusted using a pair of reference points66which define the dihedral angle69.

Attention is now directed toFIGS. 14 and 15which depict the use of non-contact measuring equipment to measure the three dimensional position of parts or features of the wing assembly24as well as the fuselage22.FIG. 14depicts the use of both a laser tracker104and photogrammetry apparatus106for measuring features such as reference points66on the wing assembly24. In the illustrated embodiment, a merged photogrammetry and laser tracking technique is used to determine the special location of laser targets such as targets66aand66bshown inFIG. 15. The two sets of measurement data generated by the laser tracker104and photogrammetry equipment106are loaded into a computer (not shown) and are combined using commercially available spatial analyzer software.

The merged laser tracker and photogrammetry technique mentioned above is described in more detail in U.S. patent application Ser. No. 11/518,471, 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 target66bshown inFIG. 15may be coded, by uniquely arranging reflective squares and dots69which may be “read” by a computer to uniquely identify the position of the targets66b. For example, the uniquely positioned targets66bcan be used to establish the position of the reference points66shown inFIGS. 6,10and13. 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 assembly24and the fuselage22, 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 assembly24and the fuselage22, in a common coordinate system.

Reference is now made toFIGS. 16 and 17which depict the steps and related software flow charts for joining and fitting large complex part assemblies such as the previously described attachment of fuselage sections22a,22band wing assemblies24. As shown at108, the laser/photogrammetry process108is used to measure the spatial position of the fuselage sections at114and the wing assembly24shown at116. A set of data is generated that defines airplane configuration model based definition at118. The configuration definition at118essentially 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 data118is combined with the spatial position data generated at118, and is used in an automated shim dimension process110.

The shim dimension process110begins by performing a virtual nominal join at120. The virtual join120essentially comprises an initial virtual fit between the assemblies to be joined, using the configuration definition data118. Then, at step122, the initial virtual join or fit performed at120is optimized, again using the configuration definition data118. The optimization performed at122may include analyzing the structural and aerodynamic relationships between various assemblies on the aircraft so that flight performance is optimized within the airplane configuration definition118. For example, the inclination, sweep and dihedral of the wing assembly24may be adjusted within certain ranges determined by the configuration definition118in order to optimize aircraft performance. Then, at step124, 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 step122.

The details of the automated shim dimension process110are shown inFIG. 17. The virtual nominal join or fit process120requires 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 process108(FIG. 16) is imported as three dimensional data into a CAD program such as CATIA at148. The fitting process includes the alignment of key geographic features which are typically fixed at150. The parameters used in the virtual join process are optimized at152in 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 at122.

The preliminary virtual fit is initially optimized using the configuration definition data118(FIG. 16), resulting in a set of interim data156that is then analyzed at158. At160, 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 at170. However, if the analyzed results are not valid, a determination is made at step162of 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 board168. However, if the results appear to be correctable, the optimization parameters are revised at step164and a determination is made at166whether to approve the revised optimization parameters. If approval is obtained at step166, the optimization of the fit is repeated at step154using revised optimization parameters.

When the fit is accepted at step170, a set of data is developed and stored at step174comprising empirical shim data and alignment data. The data developed at step174may 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 at172which may be stored as CAD shim models at176. The shim models176may be automatically delivered as digital data files to equipment (not shown) such as a CNC machining center which automatically machines the shims72to the dimensions which fill gaps based on the accepted fit at170.

Attention is now directed toFIG. 18which depicts the steps of fitting and assembling wing and fuselage assemblies that have been fabricated in different geographic locations. The wing is assembled at step184, following which a survey is performed to measure the location of features on the assembled wing at186, using, for example, the merged laser tracker photogrammetry technique described earlier. At step188, the feature location data is transmitted to a second geographical location180where this, along with other data relating to the location of features on the fuselage, nominal engineering data, etc are loaded at194. An initial virtual fit is performed at196, following which the virtual fit is optimized at198, as described earlier. The optimized virtual fit data is transmitted back to the first geographic location178where modifications to the wing assembly are carried out, if required. At step92, the wing assembly is shipped to the final assembly location180.

At a second geographic location182, the fuselage is assembled at step204, following which a survey is made to; measure the location of fuselage features at206using the previously described laser tracker/photogrammetry techniques. At step208, the surveyed location data is transmitted to the final assembly location180and is used as part of the data loaded at194employed to carry out the initial virtual fit at196. The optimized virtual fit information is transmitted back to the fuselage assembly location182where it is used to carry out any modification of the fuselage, if required. At step212, the fuselage assembly is shipped to the final assembly location180based on the optimized virtual fit at198, shims are fabricated at step200which are then used to assemble the wing and fuselage at step202.

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 site180. 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 location180. It should be noted here that although performing the steps of loading the data194, performing the virtual fit196and optimizing the fit at198have been indicated as being carried out at the final assembly site180, these steps may be performed at any location, in which case the final shim dimensions are delivered to the final assembly site180where the shims are fabricated at step200.