Abstract:
A device and methods for using the device, that permit the rapid and accurate inspection of aircraft wing attachment fittings, including those wing to fuselage attachments modified according to a Structural Life Extension Program (SLEP). Such aircraft life extension programs often result in the placement of fitting stack-up components that tend to challenge the ability of standard inspection sensors and techniques to achieve accurate readings. A specially designed, contact compliant, Electric Current Perturbation (ECP) probe is used. The ECP probe positions a receive coil in conjunction with a drive coil (and its ferrite core) in a manner that minimizes steel interferences in the inspection area. The ECP probe works with conventional eddy current instrumentation with an index scanner to allow for flaw location within a particular stack-up layer and/or within the area around the attachment aperture. Data acquired through the use of the system and method of the present invention allows for the rapid discernment of flaws and defects in the area adjacent the probe placement.

Description:
CROSS REFERENCE TO RELATED APPLICATIONS 
   This application claims the benefit under Title 35 United States Code § 119(e) of U.S. Provisional Application No. 60/842,839 filed Sep. 7, 2006, the full disclosure of which is incorporated herein by reference. 

   GOVERNMENT RIGHTS 
   The U.S. government has a paid-up license in this invention and the right in limited circumstances to require the patent owner to license others on reasonable terms as provided for by the terms of contract number F42620-00-D-0037 awarded by the Department of Defense (DOD-USAF). 

   BACKGROUND OF THE INVENTION 
   1. Field of the Invention 
   The present invention relates generally to devices and methods for the non-destructive inspection of structural systems. The present invention relates more specifically to devices and methods for the inspection of aircraft structural systems, particularly those associated with the attachment of a wing structure to an aircraft fuselage. 
   2. Description of the Related Art 
   Many of today&#39;s aircraft, including military aircraft, are being utilized past their original designed lifetime. This utilization creates inherent problems related to the structural limitations brought about because of the age of the aircraft. Repairs and structural modifications can extend the life of the aircraft but in the process can position structural components that affect the ability of existing test systems to perform nondestructive inspection (NDI) on areas where these modifications have been made. 
   An example of the above described concern is in the area of wing attachment fittings that serve to connect the wing structure to the aircraft fuselage. If a defect such as a crack is identified in an attachment fitting, the repair protocol calls for drilling or reaming out the fastener hole in the attachment fitting “stack-up area” (described in more detail below) and bushing the hole to allow for the original size fastener to be reinstalled. This procedure, however, presents an inspection problem, generally requiring the removal of the wing to gain access to the areas of interest. The process of wing removal and reinstallation is costly and not always practical given the location of the aircraft and the availability of equipment and trained personnel. 
   A method and a system to inspect the fastener hole of an attachment fitting stack-up without wing removal are therefore needed. Conventional NDI methods and systems are not capable of inspecting through a bushed hole with the kind of sensitivity-to-defect that is required to identify and locate flaws and defects of the size that must be identified (or confirmed absent) in order to keep the aircraft in service. 
   The present invention addresses the above described issues through the use of not only an inspection procedure and an inspection probe structure, but also a specific application of each. This specific application may generally be referred to as a 2 nd  layer bushing inspection of an aircraft wing attachment fitting using ECP (electric current perturbation) sensing. 
   SUMMARY OF THE INVENTION 
   It is therefore an object of the present invention to provide systems and methods for the inspection of bushed fastener holes in an attachment fitting stack-up for an aircraft wing structure, which do not require wing removal. In fulfillment of this and further objectives, the present invention provides a device, and methods for using the device, that permit the rapid and accurate inspection of aircraft wing attachment fittings, including those wing to fuselage attachments modified according to a Structural Life Extension Program (SLEP). Such aircraft life extension programs often result in the placement of fitting stack-up components that tend to challenge the ability of standard inspection sensors and techniques to achieve accurate readings. 
   In the present invention, a specially designed, contact compliant, Electric Current Perturbation (ECP) probe is used. The ECP probe positions a receive coil in conjunction with a drive coil (and its ferrite core) in a manner that minimizes steel interferences in the inspection area. The ECP probe works with conventional eddy current instrumentation with an index scanner to allow for flaw location within a particular stack-up layer. Data acquired through the use of the system and method of the present invention allows for the rapid discernment of flaws and defects in the area adjacent the probe placement. 
   The methods and probe structures of the present invention therefore provide an inspection technique that allows for an accurate assessment of the integrity of an aircraft wing attachment fitting, even one that has been modified to extend its life, without the need to remove the entire wing structure from the aircraft fuselage. 

   
     BRIEF DESCRIPTION OF THE DRAWINGS 
       FIG. 1  is a perspective view of the assembled compliant ECP probe of the present invention. 
       FIG. 2  is a partially exploded perspective view of the compliant ECP probe of the present invention. 
       FIG. 3  is a detailed perspective view of the sensor components of the ECP probe of the present invention. 
       FIG. 4  is a detailed plan view of the differential receive coil component of the ECP probe of the present invention. 
       FIG. 5  is a detailed partial cross sectional view showing the probe of the present invention utilized in conjunction with a fastener hole on a typical wing structure. 
       FIG. 6  is a detailed partial cross sectional view of a wing structure bearing five (5) fastener holes appropriate for use in conjunction with the probe and system of the present invention. 
       FIG. 7  is a perspective view of a simulated wing attachment fitting utilized as a test specimen for use in conjunction with the present invention. 
       FIG. 8  is a detailed perspective view of a simulated inspection hole representing a typical fastener hole for a wing structure applicable to the present invention. 
       FIG. 9  is a graphic plot of data collected with the system of the present invention showing evidence of a detected flaw furthest from the interfering steel. 
       FIG. 10  is a graphic plot of data collected with the system of the present invention showing evidence of a detected flaw closest to the interfering steel. 
   

   DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS 
   The basis of the system and method of the present invention revolves around the design of an Electric Current Perturbation (ECP) compliant probe (see in particular  FIGS. 1 and 2 ), capable of adjusting to the surface inside dimension of the fastener hole. This probe uses a drive coil and a differential receive coil arrangement (see in particular  FIGS. 3 and 4 ), that significantly improve the sensor discrimination. The receive coil is advantageously positioned below the drive coil. The positioning of the receive coil to the drive coil helps minimize magnetic steel interferences of the inspection area at the receive coil by using the ferrite core of the drive coil as a shield. This arrangement increases the sensitivity to defects of the probe with both a standard attachment fitting configuration and with a Structural Life Extension Program (SLEP)-modified wing attachment fitting stack-up (see for example  FIGS. 5 and 6 ). 
   In standard practice, a structural crack adjacent a fitting may be initially detected during implementation of a conventional NDI method after wing removal. The repair of such an attachment fitting hole stack-up configuration, calls for the fastener hole to be reamed or enlarged to remove the suspected crack, and a bushing to be installed, thus allowing for the original sized fastener to be reinstalled. Once bushing installation is completed, the wing is reinstalled and mated to the fuselage. Force Structural Maintenance Programs (FSMP) based on damage tolerance analysis dictate the intervals for recurring inspection. Typically, the next inspection interval would require a costly wing removal; however, through the use of the ECP probe of the present invention, a standard or SLEP-modified attachment fitting configuration could be re-inspected with only a fastener removal. 
   Tests conducted on a test specimen representing the stack-up area of a wing attachment fitting area (see  FIGS. 7 and 8 ), clearly show the layers of material needing inspection after the fastener has been removed. The layers may preferably include a SLEP modification with a 17-7 ph steel addition for structural enhancement. As in the real world situation, this additional material creates problems with a conventional eddy current probe because of the magnetic interference of the steel on the receive coil. The addition and positioning of the ferrite core in association with the drive coil, with respect to the receive coil in the ECP probe (see  FIG. 3 ) helps isolate the induced noise of the steel, thus increasing sensitivity to defects. 
   Data acquired using the test specimen shown and described above can be seen in  FIGS. 9 and 10 ; both of which show the ability of the probe to detect cracks in the aluminum skin layer even when positioned next to the 17-7 ph steel layer. As can be seen in the aforementioned figures, the probe of the present invention can detect flaws in the skin when positioned closest to the steel and furthest from the steel. Although the magnetic interferences from the steel can be seen in the upper region of both graph figures, with the shielding effect of the drive coil&#39;s ferrite core, the interference is reduced enough to allow for the detection of the cracks. 
   The ECP probe design of the present invention, with the drive coil and receive coil arrangement described, allows for sensitivity to small defects in aluminum skin with a repaired, bushed, fastener hole when positioned next to a steel strap in an SLEP-modified wing attachment fitting configuration. 
   In addition, the ECP probe design of the present invention, with the drive coil and receive coil arrangement described, also allows for sensitivity to small defects in aluminum skin with a repaired, bushed, fastener hole in a standard wing attachment fitting configuration. 
   The ECP probe of the present invention works with conventional eddy current instrumentation with an indexing scanner. This allows for an inspector to verify flaw location within the particular stack-up layer. 
   Finally, the unique compliant probe design, with its integrated spring mechanism, maintains the surface of the probe face in contact with the inside dimension wall of the attachment fitting fastener hole. As a result of this compliant contact, the data acquired allows for the discrimination of flaw information outside of the interference caused by added structural components. 
   Reference is made to  FIG. 1  for a brief description of the overall structure of the compliant Electric Current Perturbation (ECP) probe  10  of the present invention. The size and shape of probe  10  is generally that of the fastener used to attach the wing to the fuselage of the aircraft. The generally cylindrical probe  10  has a handle end  12  and a sensor end  14 . Positioned on the sensor end  14  is a proximity compliance mechanism  16  that serves to assure sufficient signal transmission proximity to the wall of the fastener hole being investigated. 
     FIG. 2  discloses the same ECP probe  10  in a partially exploded view, wherein the compliant mechanism  16  is segregated into its functional components as part of the sensor end  14  of the probe. In this view, compliant spring mechanisms  18   a  and  18   b  are shown positioned in a manner such that they preference sensor contact face  20  against the interior wall of the fastener hole to be investigated. The manner in which the sensor elements of the probe investigate the structures within the fastener aperture is described in more detail below. It will be understood by those skilled in the art that the necessary electrical signal connections for the probe (omitted for clarity in these views) would extend from the probe in the vicinity of the handle end, to connect to the appropriate instrumentation for the inspection system. 
     FIG. 3  is a detailed view of the sensor end  14  of ECP probe  10 . At this end of the probe, sensor contact face  20  is shown to incorporate the various electrical/electronic components (coils) that make up the sensor structure of the probe. At the sensor end  14  of sensor contact face  20  are positioned drive coil  22  (which surrounds ferrite core  24 ) and differential receive coil  26 . Drive coil connection conductors  28  extends from drive coil  22  through interior channels within probe  10  to a point where they extend lengthwise, interior to the probe  10 , towards the handle end  12  of the probe. In similar fashion, receive coil connection conductors  30  extend from differential receive coil  26  in a direction apart from drive coil  22  again to a point where they extend down along the length of the probe interior toward the handle end  12  of the probe  10 . 
   As mentioned, the above described probe structure allows for an appropriate level of shielding to occur over differential receive coil  26 . The ferrite core  24  and drive coil  22  positioned adjacent to differential receiver coil  26  provide the necessary shielding to reduce the effects of the steel plate components of the inspected structures, which steel components are typically placed as part of a SLEP-modified wing attachment fitting configuration. 
   Reference is now made briefly to  FIG. 4  for a detailed view of the differential receive coil structure of the present invention. The differential receive coils are, as known in the art, counter wound coils positioned coaxial with each other. In this view of  FIG. 4 , receive coil core  32   a  is wound in a clockwise direction (when viewed from the left looking towards the right in  FIG. 4 ) by way of receive coil windings  34   a . In similar but opposite fashion, receive coil core  32   b  is counter wound (counter clockwise in the view mentioned above) with receive coil windings  34   b . With this structure, differential receive coil  26  operates to detect a return signal from the material initiated by the drive coil as is known in the art. 
   Reference is now made to  FIGS. 5 and 6  for a brief description of the manner in which ECP probe  10  is utilized in conjunction with the inspection of fastener hole apertures associated with the attachment of a wing to an aircraft fuselage. In  FIG. 5  a detailed view of the area and structure immediately surrounding fastener hole  38  is shown. Probe  10  is shown positioned for insertion into fastener hole  38 , which is itself positioned on wing structure  40 . The wing is structured to include stringer  42  (as an example) and rib  44  (as an example) in the manner shown in  FIG. 5  which is typical for the environment surrounding a wing attachment fastener hole  38 . Immediately adjacent and overlaying rib  44  is lower skin  46 . Outside of lower skin  46  is SLEP strap  48  placed and positioned as described above. Overlaying SLEP strap  48  is Longeron splice fitting  50 . Overlaying Longeron splice fitting  50  are a number of shims  52 . Bushing  54  is shown positioned within fastener hole  38  as a retrofit for the purpose of extending the life of the aircraft structure. 
     FIG. 6  shows (in partial cross sectional detail) a number of fastener holes  38  (labeled A through E in  FIG. 6 ). The partial cross sectional view shown in  FIG. 5  may be recognized as a portion of the view shown in  FIG. 6  associated with fastener hole D as an example. In addition to the above mentioned structures associated with lower skin  46 , SLEP strap  48 , Longeron splice fitting  50 , and the various shims  52 , the view shown in  FIG. 6  additionally discloses the placement of front spar web  58  as well as front spar cap  56 . Associated with a number of fastener holes  38  on the interior of the wing structure  40  are radius blocks  62 . A primary wing fuselage fitting  60  is also disclosed in this view. In this detail shown in  FIG. 6  it can be seen how the individual and sequential use of probe  10  progressively through each of the fastener holes  38  can readily accomplish an inspection of these fittings associated with the attachment system. 
   A bench test set up appropriate for confirmation of the functionality of the probe and the associated methodology of the present invention is shown in the views presented in  FIGS. 7 and 8 .  FIG. 7  represents a test set up, descried briefly above, designed to accurately represent the stack up area of a wing attachment fitting system as the same might appear after a fastener has been removed. The layers of material in the assembly include a SLEP modification with a 17-7 small ph steel addition for structural enhancement.  FIG. 8  is a close up detailed view of one the fastener apertures on the test specimen set up. 
     FIG. 9  as described briefly above, comprises test data results showing evidence of a crack  102  positioned furthest from the steel addition in the assembly. In this case the crack is a 0.095 inch×0.050 inch flaw positioned furthest from the steel component in the area adjacent the aperture.  FIG. 10  displays a similar set of data showing a 0.095 inch×0.050 inch flaw positioned closest to the steel component. In each case, the graphic data shown in  FIGS. 9 and 10  includes evidence of magnetic interferences from the steel component, which can be seen in the top portion of both figures in range  100  on the vertical graphic scale. As indicated above, due to the shielding effect provided by the drive coil ferrite core, this interference is reduced enough to readily permit the detection of signal data associated with the flaws in the area under investigation. 
   Although the present invention has been described in terms of the foregoing preferred embodiments, this description has been provided by way of explanation only, and is not intended to be construed as a limitation of the invention. Those skilled in the art will recognize modifications of the present invention that might accommodate specific environments and specific aircraft fitting geometries. Such modifications as to size, and even configuration, where such modifications are merely coincidental to the specific application of the devices and methods, do not necessarily depart from the spirit and scope of the invention.