Patent Document

FIELD OF THE INVENTION  
       [0001]     The present invention relates generally to a method for detecting and repairing scratches and cracks located proximate aircraft lap joints formed by overlapped fuselage skin panels, and more particularly to a method for detecting and repairing such scratches and cracks in inner panels adjacent lap joints.  
       BACKGROUND OF THE INVENTION  
       [0002]     Lap joints are created when two or more aircraft skin panels are joined, and a portion of one panel (i.e. an inner panel) is overlapped by portion of another panel (i.e. an outer skin panel). The term lap joint, as used herein, refers both to longitudinal joints, as formed when outer (e.g. an upper longitudinal panel) and inner (e.g. a lower longitudinal panel) fuselage skin panels are joined, and to circumferential or butt joints, as formed when two curved skin panel assemblies are joined by a structural panel (e.g. a splice plate). Similarly, the term inner panel, as used herein, may refer any structural panel (e.g. a splice plate or an inner fuselage skin panel) that is at least partially overlapped at a lap joint. Lap joint panels are typically joined together utilizing an anti-corrosive sealant and a plurality (e.g. two or three) rows of rivets disposed proximate the outer skin panel&#39;s overlapping edge.  
         [0003]     It has been discovered that the surface of inner panels may be scratched proximate the lap joints during routine maintenance (e.g. during removal of excess lap joint sealant). This is problematic because such scratches may lead to the formation of cracks in panels over time that may structurally compromise the aircraft&#39;s fuselage. If cracks have not yet formed, the scratches may be blended out by abrasively removing a shallow volume of material along the panel&#39;s surface providing that scratches are visible and accessible and that the scratched skin panel is sufficiently thick. If cracks have formed, however, the cracked panel may require the excision of the cracked portion thereof and the installment of a replacement panel such as a repair doubler. Unfortunately, this is a relatively costly and cumbersome process.  
         [0004]     Repair of scratches and cracks on or in the lapped area of an inner panel is further complicated because access thereto is prevented by the overlapping outer skin panel. For this reason, detection of such inner skin cracks and scratches typically requires the use of expensive ultrasonic and subsurface eddy current detection methods as opposed to other, less expensive detection methods (e.g. visual detection for scratches and high frequency eddy current detection for cracks).  
         [0005]     It should thus be appreciated that it would be desirable to provide an improved method for detecting and, if necessary, repairing cracks and scratches present in inner aircraft fuselage panels proximate the aircraft&#39;s lap joints.  
       BRIEF SUMMARY OF THE INVENTION  
       [0006]     According to a broad aspect of the invention there is provided a method for detecting scratches proximate an aircraft lap joint formed where an outer skin panel overlaps an inner panel wherein a portion of the outer skin panel overlapping the inner panel is trimmed to expose a previously overlapped region thereof, and the previously overlapped region is inspected to detect scratches.  
         [0007]     According to a further aspect of the invention there is provide a method for detecting scratches and cracks proximate an aircraft fuselage lap joint formed where an outer skin panel overlaps an inner panel wherein a portion of the outer skin panel overlapping the inner panel is trimmed to expose a previously overlapped region thereof, and the previously overlapped region is inspected to detect scratches and tested to detect cracks.  
         [0008]     According to a still further aspect of the invention there is provided a method for detecting and repairing scratches and cracks proximate an aircraft fuselage lap joint formed where an outer skin panel having an overlapping edge overlaps an inner panel wherein at least a first scratch is found in the inner panel proximate the lap joint, and a portion of the outer skin panel overlapping the inner panel including at least a portion of the overlapping edge is trimmed to expose a previously overlapped region of the inner panel. The previously overlapped region is inspected to detect scratches and tested to detect cracks and at least one detected crack is repaired.  
     
    
     BRIEF DESCRIPTION OF THE DRAWINGS  
       [0009]     The present invention will hereinafter be described in conjunction with the following figures, wherein like numerals denote like elements, and:  
         [0010]      FIG. 1  is a plan view of an aircraft fuselage;  
         [0011]      FIGS. 2 and 3  are cross-sectional views of a circumferential and a longitudinal lap joint, respectively;  
         [0012]      FIGS. 4 and 5  are isometric views of an untrimmed and trimmed longitudinal lap joint of the type depicted in  FIG. 3 , respectively;  
         [0013]      FIG. 6  is a magnified photographic cross-sectional view of the untrimmed lap joint shown in  FIG. 4 ;  
         [0014]      FIG. 7  is an isometric view of a lap joint of the type depicted in  FIGS. 3-6  and a trimming tool for use thereon;  
         [0015]      FIG. 8  is a side view a lap joint of the type depicted in  FIGS. 3-7  and an isometric view a high frequency eddy current (HEFC) device for detecting cracks therein;  
         [0016]      FIG. 9  is a schematic view illustrating the generation of eddy currents in a test article (e.g. an inner structural panel such as an inner fuselage skin panel) by way of an HFEC device of the type shown in  FIG. 8 ; and  
         [0017]      FIG. 10  is a flow chart illustrating an exemplary inventive method for detecting and, if necessary, repairing inner panel cracks and scratches.  
     
    
     DETAILED DESCRIPTION OF THE INVENTION  
       [0018]     The following detailed description of the invention is merely exemplary in nature and is not intended to limit the scope, applicability, or configuration of the invention in any way. Rather, the following description provides a convenient illustration for implementing an exemplary embodiment of the invention. Various changes to the described embodiment may be made in the function and arrangement of the elements described herein without departing from the scope of the invention.  
         [0019]      FIG. 1  is a plan view of an aircraft fuselage  140 , comprising multiple structural panels (e.g. fuselage skin panels). When coupled together, the panels comprising fuselage  140  are joined utilizing at least two types of joints: circumferential lap joints  142  and longitudinal lap joints  100 . Referring to  FIG. 2 , a circumferential lap joint  142  is formed when first and second outer skin panels  102  and  150 , respectively, are coupled by first and second pluralities of rivets  110  and  146 , respectively, to an inner structural panel  144  (e.g. a splice plate).  
         [0020]     In contrast to circumferential lap joints longitudinal lap joints typically join an outer fuselage skin panel to an inner overlapped fuselage skin panel. A longitudinal lap join  100  is shown in cross-section in  FIG. 3  and in an isometric view in  FIGS. 4 and 5 . Lap joint  100  comprises an outer skin panel  102 , a doubler  104 , and an inner skin panel  106 . Outer skin panel  102  may be bonded by way of an anti-corrosion sealant (not shown in  FIGS. 3-5 ) to doubler  104 , which may be, in turn, bonded by way of anti-corrosion sealant (not shown) to inner skin panel  106 . Each of these three layers is further coupled together by a plurality of rivets  110  (e.g. three rows of counter-sunk rivets). Outer skin panel  102 , bonded doubler  104 , and inner panel  106  may be manufactured from a lightweight material (e.g. aluminum) and may have a base metal component comprising an alloy (e.g. aluminum-copper).  
         [0021]     Referring to  FIGS. 4 and 5 , an area  300  (shown exaggerated for clarity) of the outer surface of inner panel  106  is prone to scratching during aircraft maintenance (e.g. during removal of excess sealant). Area  300  is disposed proximate the overlapping edge of lap joint  100  and roughly corresponds to the location of excess sealant that may have been removed during maintenance. When lap joint  100  is untrimmed as shown in  FIG. 4 , only a portion of surface area  300  may be seen. After lap joint  100  is trimmed as shown in  FIG. 5 , however, surface area  300  may be seen in its entirety.  
         [0022]     A series of scratches  130  (e.g. scribe marks made, perhaps, by a cutting tool used to remove excess sealant) is present on inner surface  106  within surface area  300 . Prior to trimming ( FIG. 4 ), scratches  130  are only partially visible. After trimming ( FIG. 5 ), however, scratches  130  are entirely visible. The presence of scratches  130  suggests that inner panel  106  may have additional scratches within or proximate area  300  that are hidden by overlapping edge  114  of outer skin panel  102  and doubler  104 . Edge  114  may be trimmed (i.e. removed) without weakening lap joint  100  to reveal area  300  in its entirety and thus permit further inspection thereof. As can be seen in  FIG. 5 , trimming of edge  114  reveals a second series of scratches  132  that was hidden by overlapping edge  114  prior to trimming.  
         [0023]      FIG. 6  is a magnified photographic cross-sectional view of lap joint  100 . Scratches may initiate the formation of cracks that penetrate into inner panel  106  and weaken lap joint  100 . As can be seen, a plurality of scratches  404  including a scratch  402  is present on the outer surface of inner panel  106 . A crack  400  has initiated from scratch  402  and extends downward therefrom into inner panel  106 . When lap joint  100  is untrimmed ( FIG. 4 ), scratch  402  is hidden from view by overlapping edge  114 . Trimming of edge  114 , however, may reveal scratch  402 .  
         [0024]     Trimming of edge  114  may be accomplished by the means of a trim tool  300 , for example, of the type shown in  FIG. 7 . Trim tool  300  comprises a handle  302  and a rotary cutting shaft  306  having a distal cutting head (not shown). Trim tool  300  has a guide edge  301  that may be moved along lap joint  100  in the direction of arrow  308  so as to trim edge  114  away from outer skin panel  102  and doubler  104 . Cutting shaft  306  is offset from edge  301  and inner panel  106  so as to trim lap joint  100  to a predetermined depth. The trimming depth may be controlled by adjusting the depth of the distal cutting head relative to guiding edge  301 . It is desirable that trim tool  300  remove most or all of edge  114  while leaving inner panel  106  unscathed. Generally, trim tool  300  should be of the type capable of trimming 0.070 inch plus or minus 0.010 inch. If preferred, however, trim tool  300  may be configured to leave behind a thin layer of doubler  104  and sealant  410 , which can later be removed with a manual scrapping tool (e.g. a nylon scrapper). If preferred, however, trim tool  300  may be configured to trim more then 0.070 inches of the lap joint or less then 0.070 inches of the lap joint edge and leave behind a thin layer of doubler  104  and sealant  410 , which can later be removed with a manual scrapping tool (e.g. a nylon scrapper). Trimming of edge  114  may also be accomplished by means of an automated computer controlled machine trim tool rather then a hand operated tooling. The computer controlled trim tool machine may use a laser for trim feed back control of the lap joint edge location and thickness.  
         [0025]     Due to the configuration of trim tool  300  (e.g. the tapering of the cutting head), residual material  122  may be left after trimming. The residual material  122  ( FIG. 5 ) may be disposed along the base of edge  120  ( FIG. 5 ) proximate inner panel  106  and may comprise a remnant foil of doubler  104  and, perhaps, outer skin panel  102 . To prevent interference with crack detection, residual material  122  ( FIG. 5 ) should be of a relatively small width and thickness (e.g. 0.020 and 0.003 inch, respectively).  
         [0026]     After edge  114  has been trimmed away from lap joint  100 , the newly exposed section of inner panel  106  including area  300  may be examined for scratches and cracks. Scratches may be detected by, for example, visual observation. Cracks, which may extend further below the surface of skin  106 , may be detected using a non-destructive inspection (NDI) method; for example, high frequency eddy current (HFEC) inspection.  
         [0027]     An exemplary HFEC device  600  is illustrated in  FIG. 8 . HFEC device  600  comprises a probe  602  coupled by way of a connection  604  to an inspection instrument  606  having a display  608 . As illustrated in  FIG. 9 , HFEC device  600  creates eddy currents within an article  704  (e.g. a structural panel such as inner skin panel  106 ) by delivering thereto an alternating current  700  via a conductive coil/s  702  contained within probe/s  602  ( FIG. 8 ). Alternating current  700  induces an alternating magnetic field  706  in article  704 , which, in turn, induces eddy currents  710  to flow therethrough. The strength of eddy currents  710  are measured by probe  602  and the results displayed on display  608 . If the observed conductivity is significantly below a predicted value, current-impeding cracks are likely present in the tested article and crack repair may be undertaken.  
         [0028]     HFEC inspection devices, such as that just described, are well known and further discussion is not deemed necessary at this time; however, the interested reader is referred generally to U.S. Pat. No. 4,706,020 entitled “High Frequency Eddy Current Probe with Planar, Spiral-like Coil on Flexible Substrate for Detecting Flaws in Semi-Conductive Material” issued to Viertl, et al. on Nov. 10, 1987, and U.S. Pat. No. 3,963,980 entitled “Ultrasonic Instrument for Non-Destructive Testing of Articles with Current-Conducting Surfaces” issued to Shkarlet on Jun. 15, 1976.  
         [0029]      FIG. 10  is a flow chart illustrating an exemplary embodiment of an inventive process for detecting and, if necessary, repairing inner panel (e.g. inner fuselage skin panel) scratches and cracks of the type described above in connection with lap joint  100 . To begin, the exposed portion of the inner panel proximate the lap joint may be examined for scratches ( 802 ). This may be done, for example, through visual inspection. If no scratches are found on the exposed portion of the inner panel, the process may be halted ( 816 ) and no further action taken until the next scheduled inspection. If, however, scratches are found on the exposed inner panel, the outer skin panel may be partially trimmed away in the above described manner ( 804 ). An inspection for scratches on the newly exposed area of the inner panel may then be conducted ( 806 ). If no scratches are detected on the newly exposed area, the scratches previously detected in the inner panel may be repaired (e.g. blended out) ( 814 ) and the process may be halted ( 816 ) or re-inspect the scratch locations for crack development at an interval dependant on the airplane landing cycles. Alternatively, if scratches are detected on the newly exposed area, a high frequency eddy current (HFEC) inspection may be performed (as described above) on the inner panel to determine if cracks have formed therein ( 808 ). If no cracks are detected, the existing scratches may be repaired ( 814 ) and the process may be halted ( 816 ). If cracks are detected, the cracked area of the panel may be replaced ( 812 ) by way of, for example, a repair doubler and the process may be halted ( 816 ).  
         [0030]     It should be understood that the exemplary process described in conjunction with  FIG. 10  is only suggestive in nature. Other processes comprising steps similar to the exemplary process may be implemented. For example, it may be desirable to replace inner panels that are scratched but not cracked, especially if the scratches are relatively deep. Conversely, if the inner panel is sufficiently thick, it may be possible and desirable to blend out relatively shallow cracks. It should further be understood that repair may not take place immediately after detection as it may be desirable to monitored detected scratches for repair at a later time (e.g. after a predefined number of flight cycles).  
         [0031]     It should be appreciated that, although the inventive method has been primarily described above in conjunction with longitudinal lap joint  100 , the method may be used to repair cracks and scratches proximate circumferential lap joints as well (e.g. circumferential lap joint  142  of  FIG. 2 ). It will be clear to one skilled in the art that, as circumferential joints generally comprise two overlapping edges (e.g. overlapping edges  114  and  148  shown in  FIG. 2 ), many or all of the steps of the inventive method must be performed twice (once for each overlapping edge) to inspect and, if desired, repair the entire circumferential joint.  
         [0032]     While an exemplary embodiment has been presented in the foregoing detailed description, it should be appreciated that a vast number of variations exist. It should also be appreciated that the exemplary embodiment is only an example, and is not intended to limit the scope, applicability, or configuration of the invention in any way. Rather, the foregoing detailed description will provide those skilled in the art with a convenient road map for implementing the exemplary embodiment. It should be understood that various changes can be made in the function and arrangement of elements without departing from the scope of the invention as set forth in the appended claims and the legal equivalents thereof.

Technology Category: 7