Abstract

The C-141 System Program Directorate (SPD) at Warner Robins Air Logistics Center (WR-ALC) has identified second-layer cracking of C-141 lower inner-wing spanwise splice-joints as the life limiting feature of the C-141 aircraft. Presently, splice-joint nondestructive inspection (NDI) is accomplished with an eddy current surface scan, repeated at field-level every 120 days. The surface inspection detects only surface breaking cracks, which dictates the short inspection interval. An existing bolt hole eddy current (BHEC) inspection technique is capable of detecting smaller second-layer splice-joint cracks, although the expense and aircraft downtime associated with the BHEC technique are unacceptable to the C-141 users. Including fastener removal and reinstallation operations, BHEC inspection of the C-141 splice-joints is estimated to require 9,000 man-hours per aircraft. The C-141 SPD requested the WR-ALC Materials Analysis Team (TIEDM) to develop an alternate, improved inspection method for the splice-joint. The C-141 SPD goal is to conduct a one-time, field-level fleet inspection and subsequently establish a five- year, depot-level inspection interval. TIEDM determined the best alternative inspection method available, with potential to meet the specific C-141 SPD requirements, was an automated ultrasonic scanning technique.

A prototype inspection system was designed and manufactured by SAIC and is currently undergoing acceptance testing. Laboratory and on-aircraft testing to date have proven the prototype system has a detection threshold potential of (0.050 inch crack lengths in the C-141 splice-joint second layer. A Probability-of-Detection (POD) Study has been designed to formally quantify the inspection reliability of the prototype process. POD data will allow the C-141 SPD to confidently determine future inspection intervals and requirements. The new automated ultrasonic scanning process has the potential to significantly reduce splice-joint inspection labor costs as compared to the costs of BHEC inspection. This paper describes the inspection development process, including feasibility studies, design and fabrication of prototype, functional testing, POD testing, and Air Force implementation. Typical ultrasonic C-scan images will be presented, and the POD test plan will be discussed.

TABLE OF CONTENTS

• Background
• Inspection Area Technical Description
• Ultrasonic Method - Feasibility Study
• Process Development and Prototype Delivery
• Prototype System Description
  Scanner, Scanner Utilities, Portable Cart,
  Computer Control System, Tandem Transducer Head
• Functional Testing
• Probability-of-Detection Experiment
  Specimen Design, Specimen, Characterization, Calibration, Standards,
  POD Testing & Variables, Inspectors & Training, POD Protocols, Data Analysis
• Specimen Cracking and Characterization Results
• Air Force Implementation Plan
• REFERENCES

Background

Risk analysis assessment of the C-141 spanwise splices indicates that the inner wing risk could become unacceptable for some C-141 aircraft with over 43,000 flight equivalent damage hours (Ref. 1). This assessment is supported by data gathered during a C-141 wing teardown, performed by Lockheed in 1993, to determine the condition of the inner wing, lower surface, spanwise splices (Ref. 2). Second-layer cracking in the C-141 lower inner wing potentially affects over 7,000 fastener sites per aircraft.

The spanwise splice cracking problem dictated the need for inspection of the second-layer or inner tab of the splice-joint. Bolt hole eddy current (BHEC) inspection is the only immediately fieldable method available, capable of providing a reliable inspection of the splice-joint inner tab. BHEC has the proven capability of detecting 0.070" cracks in the inner tab fastener holes, but its use would require the removal of every splice-joint fastener. Cost estimates for BHEC testing of the C-141 spanwise splice-joints have ranged up to 9,000 man-hours per aircraft. Potential risks associated with BHEC include inducing damage to fastener holes during large scale removal and reinstallation operations, and the degradation of fuel tank sealant integrity. Based on the associated costs and risks, BHEC is an unacceptable alternative for large-scale inspection of C-141 spanwise splice-joints. As an interim safety measure, the C-141 SPD issued a Time Compliance Technical Order (TCTO) in July '96 which directs the performance of an eddy current surface scan on the exterior surface of the lower inner wing spanwise joints. This inspection will remain in effect until a new method; capable of detecting smaller inner tab cracks can be developed and validated.

The C-141 SPD requested that WR-ALC/TIEDM develop an alternative inspection method with the following capabilities: a) proven detection of 0.125 inch second-layer cracks, b) inspection from aircraft exterior without fastener removal, c) archived inspection records, d) automation to remove operator subjectivity, e) portability, f) significant return-on-investment. Based on inner wing damage tolerance analysis data, 0.125 inch crack detection capability would allow a five-year inspection interval to be established for the spanwise joint.

TIEDM immediately explored and evaluated low frequency eddy current (LFEC) and manual shear wave ultrasonic inspection techniques to satisfy the C-141 spanwise joint inspection need. A LFEC technique developed by TIEDM and a manual shear wave ultrasonic technique developed by Lockheed were both proven to be acceptable for detection of a 0.375 inch crack that extends forward from an inner tab fastener hole to the panel edge. Due to a combination of joint geometry, panel thickness changes, and/or fastener variables, neither technique has yet proven to be an effective means for detection of shorter length cracks in the forward direction or for any crack length in the aft direction. Subsequently, neither technique will satisfy the specific C-141 SPD need for 0.125 inch crack detection in the spanwise splice-joint inner tab.

At the time of the C-141 SPD request, the FAA, under its Aging Aircraft Program, was funding the development of a multi-layer corrosion detection technique for the DC-9 wing box (Ref. 3). Successful on-aircraft demonstration of the DC-9 wing box inspection was performed in the first quarter '95. Based on similarities between the DC-9 wing box and the C-141 splice joint configuration, i.e., multi-layers bonded with aircraft sealant, TIEDM decided to pursue a proof-of-concept study to explore the feasibility of using a scanning, ultrasonic imaging technique to solve the C-141 spanwise splice-joint problem.

Inspection Area Technical Description

FIGURE 1. C-141 Lower Inner Wing Detail. FIGURE 2. Splice-Joint Configuration.

The C-141 wing is constructed of five major assemblies, the center wing box and the left- and right-hand inner and outer wing sections (see Figure 1). The inner wings are joined to the center wing at Wing Station (WS) 77.7 and to the outer wing at WS 405. The inner wing is designed to carry fuel and is divided into four separate fuel tanks. The inner wing lower surface consists of 11 wing panels numbered 1 through 11, AFT to FWD. Panels 1 through 6 span the entire inner wing length (approximately 27 feet). The length of Panels 7 through 11, taper with the leading edge of the wing

The lower inner wing panels are attached together at spanwise splices with 0.250 inch to 0.375 inch taper-lok fasteners. The fasteners are installed wet with sealant, and a faying surface sealant is applied in the panel lap joint. The sealant applied at manufacture was MIL-S-81733. The splice-joint configuration is shown in Figure 2. Total joint thickness (two layers) ranges from approximately 0.275 inch to a maximum of approximately 0.825 inch. The thickest locations are in the transition zones near the outer wing and center wing attach points. Inner tab cracking initiates at the forward and/or aft side of the splice-joint fastener holes, typically on the interface side mated with the outer tab. Cracks initiating on the forward side of a splice joint fastener hole propagate until the edge distance is traversed and the ligament is severed. Cracks initiating on the aft side of a splice-joint fastener hole propagate until they are surface breaking, as shown in Figure 2.

The current TCTO inspection performed on the C-141 spanwise splices is an eddy current surface scan which detects these surface breaking cracks. The procedure consists of the following steps: 1) Removal of the lap-joint sealant bead as necessary in the area of inspection, 2) Calibration on lap-joint standards, and 3) Surface inspection. The surface inspection entails scanning a surface probe along the aft lap-joint panel the length of each spanwise joint, #'s 2 through 9, as noted in Figure 2. The procedure is designed to detect a surface crack < 0.100 inch in length measured from the panel edge. Total crack length at this point of detection (as measured from the inner tab fastener hole) is approximately 1.0 inch. This inspection is repeated every 120 days on every C-141 aircraft and has become a major manpower burden on field-level NDI personnel. Some high time aircraft are inspected every 30 days. To date, one splice-joint crack has been found with this TCTO.

Ultrasonic Method - Feasibility Study

WR-ALC/TIEDM contracted with SAIC/Ultra Image International in October '95 to perform a feasibility study on the use of automated ultrasonic imaging technology for inspection of C-141 spanwise splice second layer cracking. The impetus for this request was based on SAIC's participation in the successful development of a multi-layer corrosion inspection technique for the DC-9 wing box. The wing box inspection utilizes a pitch-catch angle-beam methodology developed at Northwestern University (Ref. 4) to inspect through multiple metallic layers bonded with sealant. SAIC proposed to use a pulse-echo, angle beam technique (see Figure 2) combined with automated scanning and imaging technologies for the C-141 spanwise splice-joint configuration. SAIC predicted this technique would provide repeatable, accurate results with a high probability for detection of small second-layer fatigue cracks.

FIGURE 3. Feasibility Study Test-Article Configuration.

TIEDM fabricated a test article from a 17-inch segment of an actual C-141 splice-joint containing 14 fasteners. The taper-lok fasteners were removed, and the panels were separated and cleaned free of sealant. Bolt hole eddy current inspection was performed on the outer and inner tab fastener holes to determine if any actual fatigue cracks were present; none were detected. To simulate fatigue cracks, 24 saw-cuts were placed into the fastener holes of the inner tab. Sixteen through thickness saw-cuts ranged from 0.020 inch to 0.135 inch. Eight 45( corner notches ranging from 0.020 inch to 0.080 inch were also cut into the inner tab. Figure 3 shows the test article configuration and saw-cut lengths and locations. The test article was resealed with MIL-S-8802 sealant, and taper-lok fasteners were reinstalled. The test article was subsequently shipped to SAIC.

The feasibility study commenced with an ultrasonic immersion evaluation of the splice-joint test article. Immersion scanning was used to take advantage of the resolving capabilities of a focused probe and to eliminate mechanical contact scanning affects. Initial evaluations focused on the sealant bondline. The influence of the sealant bondline on the propagation of ultrasonic shear waves ultimately controls the inspection sensitivity capable by ultrasound on the splice joint configuration. Immersion testing yielded vital bondline information. It was determined that the sealant bondline provided ample sound coupling through to the inner tab of the splice-joint. Further immersion evaluations performed on the test article determined acceptable probe angle, incident angle, transducer frequency and depth gating requirements. A 20( incident angle provided the best signals from the fastener holes and saw-cuts. A typical immersion C-scan image of the test article is shown in Figure 4.

FIGURE 4. Typical Immersion Scan Results. FIGURE 5. Zoomed Image, Hole #'s 6 through 10. FIGURE 6. Immersion Scan Data From Test Article. FIGURE 7. Contact Scan Amplitude and Depth Plots.

The scan was performed using a 5 MHz, 0.50 inch diameter, spherically focused (1.5 inch focus) immersion transducer, and a 0.010 inch scan resolution. Separate scans were performed with the transducer aimed inboard, then outboard. The separate inboard and outboard files are pasted together in Figure 4. Time-of-flight and amplitude data are collected with each scan. Amplitude data only is shown in Figure 4. All artificial defects in the test article were successfully detected, as noted in Figure 4, and in the zoomed image shown in Figure 5.

To illustrate the quality of the ultrasonic signals received from the inner tab of the test article, captured A-scan results from Hole #9 are shown in Figure 6.

The C-scan is comprised of peak amplitude information collected from a precisely gated signal in the second layer of the test article splice-joint. Hole #9 has two saw-cuts in the inner tab fastener hole; an 0.045 inch, 45( corner notch at the interface surface, and a 0.040 inch through-the-thickness cut. The image was generated with the transducer aimed in the outboard direction. In the C-scan, higher amplitudes are shown as lighter shades. The cursor is located over the corner notch image with the corresponding A-scan shown above. The test yields excellent signal-to-noise ratio for this small indication. The through-the-thickness 0.045 inch cut image is located in the lower portion of the C-scan, just to the left of the vertical cursor. The large oval image centered in the C-scan corresponds to a signal from the inner tab fastener hole. The small oval image is skip information, inherent to angle-beam inspections, from the first-layer (outer tab) fastener hole. Note that the saw-cut images are separated and appear at an angle to the fastener image. This distinct separation and location will allow crack signals to be readily identified. This phase of the study proved that ultrasonic inspection coupled with imaging technology provided adequate inspection sensitivity for detection of small fatigue cracks in the splice joint inner tab, provided the sealant bondline is intact.

The next phase of the project involved determining the feasibility of converting the successful laboratory immersion techniques to a field-level process. This conversion involved the use of automated scanning equipment and contact transducers. Contact inspection introduces a variety of disadvantages versus immersion inspection, including mechanical scanning affects and loss of the focused probe resolution. These disadvantages must be overcome with appropriate selection and design of high precision scanning equipment and contact probes.

The reference sample was scanned using an SAIC manufactured, 45( shearwave, 0.250 inch x 0.250 inch, 5.0 MHz contact transducer, mounted in a standard Ultra Image semi-automatic scanner assembly. Motion encoders automatically provide the computer with precise transducer location while scanner movement is performed manually. A transducer holder was assembled that would hold transducer skew to ( 10(. The holder was designed such that the transducer is slightly recessed from the inspection surface, allowing the transducer to slide on a thin film of couplant applied by an automated couplant delivery system. A portable ULTRA IMAGE IV system was programmed for scan parameters, and ultrasonic settings were optimized for contact transducer scanning. The scan resolution was set for 0.020 inch. Again, two scans were performed, one with the sound beam aimed in the inboard direction, the other with the sound beam aimed outboard. Redundant inspection from both sides of the fastener reduces the probability of missing data due to a variety of variables, including poor surface conditions, bondline voids, or crack skew. Figure 7 shows typical C-scan data collected during a semi-automatic contact scan.

The image was generated with the transducer aimed in the inboard direction. Due to the physical limitations of contact scanning at the end of the test article, only 13 fasteners could be inspected per scan. The image displays time-of-flight (TOF) data at the top and peak amplitude data at the bottom. In the TOF scan, fastener holes and saw-cuts are clearly shown with the gradient scale corresponding to decreasing TOF, light to dark. In the amplitude scan, the saw-cut images are not as distinct, but can be distinguished by their lateral position at the top and bottom side of the fastener hole image. Note that saw-cut image positioning is the same as detected during immersion scanning (Ref. Figure 4), although signal separation between the fastener hole and saw-cut images is not as distinct. All through-the-thickness saw-cuts ( 0.040 inch were detected with the semi-automatic scanning set-up. Corner notch detection proved more difficult, as only one corner notch (0.075") at the interface was consistently detected.

The feasibility study was completed by SAIC in December '95, and their results were reported to TIEDM. The study conclusively demonstrated the feasibility of fielding an automated ultrasonic inspection process capable of second-layer fatigue crack detection in the C-141 lower inner wing spanwise splice-joints. It was further demonstrated that, with further process development, fatigue cracks as small as 0.050 inch may be targeted for on-aircraft inspection. These results far exceeded the C-141 requirement for 0.125 inch crack detection. TIEDM reported the results to the C-141 SPD and concluded that automated scanning, coupled with ultrasonic imaging technology, offered the best near-term solution for C-141 spanwise splice-joint inspection needs. TIEDM recommended pursuing further development and optimization of this inspection technology for the C-141 application.

Process Development and Prototype Delivery

Based on the successful results of the feasibility study, WR-ALC/TIEDM pursued funding for the development of a spanwise splice-joint inspection process. Funding was provided to TIEDM by HQAMC with the subsequent cognizance of the Aging Aircraft & Systems Office. A Contract Engineering Task was awarded to SAIC under the WR-ALC Design Engineering Program. The primary purpose of this contract effort is to develop a more reliable nondestructive inspection process. Initial contract objectives were:
• To design, develop, test, and prototype a portable automated ultrasonic scanning inspection process. The process shall have a detection threshold of ( 0.050 inch for chordwise fatigue cracks in the inner tab (second-layer) fastener holes of the C-141 lower inner wing panel, spanwise splice-joints. The process shall be accomplished from the wing exterior and shall require no pre-inspection aircraft preparation, such as coating removal, fastener removal, or couplant application.
• The process shall be automated to the maximum extent and shall be operable by one Level II Ultrasonic NDI technician. The process shall be self-contained and portable to the maximum extent to allow travel to and around the aircraft in a maintenance hangar environment. Any components attaching to the aircraft shall be as lightweight as possible.
• The process shall provide for high speed ultrasonic scanning inspection, capable of inspecting a minimum of 100 splice joint fastener holes per hour. Fewer than 450 man-hours shall be required for complete aircraft inspection.

Prototype System Description

Once the ultrasonic parameters were established, a field-level inspection system had to be designed and constructed to meet the established operational requirements. A prototype system and preliminary inspection procedure were delivered to TIEDM in February '97. The prototype system consists of a scanner, transducer head, and portable cart which houses all computer controls and accessories. Descriptions of all components are detailed below:

Scanner

Scanner design efforts focused on reducing weight while maintaining speed, resolution, and reliability. The scan envelope (36" X 6") was designed to accommodate the long, narrow area of interest. The width of the splice-joint area in need of inspection is only 2 inches, so the Y-axis extension was intentionally kept to a minimum to reduce the weight of the scanner. Because of the procedural numbering sequence chosen, all scans will be less than 36 inches in length. The scanner is vacuum attached to the aircraft skin with three, six-inch cups spaced along the scanner length. The scanner axes are powered by two AC-induction motors. The motors and accompanying controller provide high resolution and positional accuracy to ( 0.001" at speeds up to twelve inches per second. Integral scanner safety measures include visual gauge indicators and audible alarms. Supplemental safety measures include safety straps and a portable catch basin. Total scanner weight is approximately 40 pounds.

Scanner Utilities

Support functions required by the scanner are provided by the scanner utility system. The utility system supplies regulated air pressure, automatic coupling, and vacuum control. These systems are integrated inside the portable cart as a slide-out module. All utilities are monitored and controlled by the operator from a control panel on the cart exterior. Utility specifications are provided as follows: Air (0 to 45 psi), Water (0 to 35 psi), Vacuum (18 to 22 inches Hg).

Portable Cart

The cart is totally self-contained and requires only access to 120V external power. The cart houses the system computer, scanner motion controller, couplant reservoir and utility box. The cart also provides for monitor, cables, and miscellaneous parts storage during transport. The cart also contains an integral utility control panel and pullout keyboard. Auxiliary power outlets and a printer port are also provided. While in operation, cooling is provided by two fans which circulate air through the cart. Thirty-five foot scanner attachment cables provide numerous options for cart placement during inspection operations. The cart has large, steerable pneumatic wheels and a tow bar to aid in portability.

Computer Control System

The system controlling software is ULTRA IMAGE IV/Scanmaster Version 9. The software provides the operator with a Windows-like operating environment for all motion control, ultrasonic, data analysis, and reporting functions. The computer is 120 MHz Pentium-based with a 540 MB hard disk. The ultrasonic system supports three channels with fully programmable pulser, receiver, and gate parameters. The system is capable of real-time display of A-scan, B-scan, and C-scan information of all three channels. Each channel can be programmed for A-scan waveform capture. All inspection results are completely archived on hard disk or on a 1 GB JAZ drive provided for external data storage.

Tandem Transducer Head

A tandem transducer head (See Figure =>) is used to hold two opposing shearwave, pulse-echo transducers. Each transducer is separately gimbaled to allow travel over minor surface irregularities without loss of coupling. Transducer spacing is adjustable, which allows ultrasonic sound paths to be optimized for shear wave inspections. The transducer housing is ported to provide couplant flow to the inspection surface. The housing also allows the transducer to be recessed so that it slides on a thin film of couplant. The holder also provides two inches of vertical displacement by means of an integral pneumatic actuator which provides slight contour- following capabilities. Air pressure provided by the actuator maintains constant transducer pressure against the inspection surface. The transducer head is configurable to a single or triple transducer arrangement. FIGURE 8. Transducers in scanning position on-aircraft.

Functional Testing

Functional testing was used to verify system operations under conditions of an actual field inspection environment. On-aircraft functional tests were conducted at WR-ALC , February '97, on C-141 aircraft S/N 0273. The aircraft was at WR-ALC undergoing PDM inspection and had previously been defueled and depainted. Engine pylons had also been previously removed. A preliminary inspection procedure, based primarily on laboratory testing was used to conduct the test. First, the lower wing skin surface was surveyed to determine the plan of action. From the survey it was determined that the first twelve to eighteen inches of each splice joint outboard of the center wing/inner wing joint were inaccessible to the automatic scanner. These areas will be scanned in a semi-automatic mode with the aid of a hand-operated manual scanner. Data is acquired in the same manner as in the automatic mode. A manual scanner was placed into position to verify full coverage of the joint area. Approximately one hundred fasteners per wing will require inspection in this manner. It was also determined that with engine pylons installed, approximately 50 fasteners underneath the pylon could not be inspected. Also, the pylon would restrict automatic scanner placement on the inboard and outboard sides, requiring approximately 100 additional fasteners to be inspected with the manual scanner.

FIGURE 9. System On-Aircraft.

The equipment was delivered to the hangar and forklifted onto the wing stand. Figure 9 shows the scanner and cart in place during scanning. Three spanwise splice-joints, #'s 3, 5 and 7, were arbitrarily selected for inspection. Splice-joints #3 and #5 span the length of the inner wing and contain approximately 300 fasteners each. Splice-joint #7 is slightly shorter and contains approximately 260 fasteners. Each splice-joint was laid out and marked, in accordance with the preliminary inspection procedure, to define the scan areas and to facilitate scanner placement. The equipment was energized and scanning started on splice-joint #3. A previously stored calibration scan file was used for ultrasonic parameters. Joints #5 and #7 were subsequently scanned in the same manner. On-aircraft scanning was completed over a period of three days. A total of 51 scans was performed, with a total of 850 fastener sites inspected. An average of 160 holes were inspected per hour. Twenty to twenty-four fasteners were inspected per set-up. Scanner positioning, scanning and data analysis took approximately 7 minutes per scan. The original goal of the program was 100 fasteners per hour.

FIGURE 10. Scan Image From Splice 7 During Functional Test. FIGURE 11. Zoomed Image of Hole #7420 and Indication.

Sample images of the inspection results are given in Figures 10 and 11, which depict the analysis mode of the inspection process. Figure 10 is an image from the scan of Splice 7, Segment 4, holes 18 through 34. Both channels' data have been pasted together; therefore, two rows of fastener site images appear. All fastener sites are displayed within the scan area, indicating complete coverage. Figure 11 is a zoomed view of Figure 10. Hole #'s 7420 and 7421 are shown; Hole 7420 has a crack indication as noted, which was subsequently confirmed with BHEC. After the image is zoomed, the operator can pan and scroll the entire scan length. Transducer location coordinates are maintained, with the image so that any point along the scan can be correlated back to a specific location on the aircraft.

A total of seven indications were detected. An area of what appeared to be corrosion was also identified. Five of the seven fasteners where indications were noted were removed so that a confirmation bolt hole eddy current inspection could be accomplished. Two of the indications were confirmed as cracks with bolt hole eddy current. Further investigation is pending on the source of the remaining three indications. Several areas of procedural improvement were identified and remedied.

Probability-of-Detection Experiment

A C-141 Ad Hoc Safety Review Team formed in March '96 to advise the C-141 SPD and the Air Mobility Command on issues relative to the continued airworthiness of the wing of the C-141 aircraft. The team goal was to ensure that the maintenance actions on the wing of the C-141 aircraft are adequate to control the risk so that loss of an aircraft would not be expected. The Ad Hoc Team completed their review in July '96 and reported the following: The current C-141 maintenance and inspection program is adequate to protect the safety of the C-141 wing, although it was suggested that immediate action was needed to upgrade the current inspection program. A major part of this upgrade was to develop, validate and implement an automated NDI system for all inspections which are deemed critical, to prevent structural failure of the C-141 wing. The lower inner wing spanwise splice-joint inspection has been deemed a critical inspection.

TIEDM briefed the Ad Hoc Team in August '96 on the status of the ongoing spanwise splice joint inspection development. Based on the maturity of the technologies to be used for the splice-joint process and favorable inspection results already achieved, the Ad Hoc Team determined the technical risk of failure of the process was low. The Ultra Image process was approved by the team as the long-term NDI solution for C-141 splice-joint inspections. The team suggested a Probability-of-Detection (POD) study be performed on the inspection process and approved the expenditure of funds on a POD study effort.

A modification was made to the existing SAIC contract, requiring the development and performance of a POD study. An initial POD Technical Coordination meeting was held in October '96, with attendees from WR-ALC/LJ/TI, SAIC, Lockheed/Martin, and Sandia National Laboratories. SAIC presented an initial outline of the POD study plan. All in attendance approved the basic elements of the proposed POD plan.

SAIC will use FAA Report No. DOT/FAA/CT-92/12,I "Reliability Assessment at Airline Inspection Facilities, Volume 1: Generic Protocol for Inspection Reliability Experiments" (Ref. 5) as a guideline in the design and execution of the Spanwise Splice POD Study. The generic protocol will be modified specifically for the C-141 spanwise splice-joint inspection process. SAIC subcontracted with Sandia National Laboratory and its FAA Airworthiness Assurance NDI Validation Center as consultants on the POD study. Specific elements of the POD Plan are detailed below:

Specimen Design: The specimens used for inspection during the POD experiment will be actual spanwise joint segments cut from C-141 A/C 66-0186 as part of the Lockheed Teardown Inspection (Ref. 2). The segments cover the entire lower inner wing spanwise joint thickness range. The segments are approximately 40inches in length, and contain approximately 35 spanwise fastener sites each (see Figure 12).

FIGURE 12. Typical POD Test Specimen. Second Layer Only Shown.

Twelve to fifteen segments will undergo cyclic fatigue testing to obtain a 4:1 uncracked-to-cracked fastener site ratio. Approximately 60 total crack sites will be propagated to cover a crack length range of 0.030" to 0.250". Some larger cracks will also be included, even one or two cracks to the tab edge. Distribution of the number of cracks and crack lengths in each segment has been statistically determined. Uncracked segments will also be included in the experiment.

Specimen Characterization: Each crack will be completely characterized and documented for length and orientation with a combination of high magnification optics (during and after cracking) and contact and immersion ultrasonic imaging techniques. Following initial characterization, the splice segments will be sealed, assembled, and painted at WR-ALC with established C-141 maintenance procedures and materials. A second characterization will be carried out with immersion ultrasonic inspection following assembly.

Calibration Standards: A calibration standard has been manufactured from a typical spanwise joint segment (as shown in Figure 12) using EDM notches to simulate second-layer fatigue cracks. EDM notching allows closer control of crack lengths, orientation, and surface condition. Notch lengths in the standard range from 0.030 inch to 0.100 inch. This standard will be used to calibrate the equipment prior to conducting the POD inspections and will insure the consistency needed from calibration to calibration. The calibration standard will be assembled and painted in the same manner as the cracked specimens.

POD Testing & Variables: Since the splice-joint inspection process is automated to a great extent, human and environmental factors will not be included in the POD study. Initially, a Laboratory Baseline of the process will be conducted to characterize the detection reliability under extreme favorable conditions with an extremely skilled operator. This exercise will focus on procedural and process variables that will require inspector input or adjustment during the course of the POD study. Variables identified to be included in the laboratory baseline are: transducer, receiver gain, calibration standard, couplant pressure, gate parameters, probe pressure, and scanner misalignment.

Actual POD inspections will take place on crack specimens mounted overhead to simulate the C-141 lower inner wing. An actual full length C-141 wing panel will be mounted on a framework to simulate work on a typical C-141 wing stand platform. The crack specimens will be attached to the wing panel to allow specimen placement to be varied along the wing panel length. The inspector will mount the scanner to the wing panel and inspect the crack specimens in the same manner as would be accomplished on an aircraft.

Inspectors & Training: Approximately 15 inspectors will participate in the POD experiment. The inspectors will represent a cross-section of experience and training to determine if and how these factors affect inspection results. The inspector pool will be a mix of the following: Six experienced equipment operators will receive only procedural specific training (3 days); Six Air Force NDI technicians with no ultrasonic scanning/imaging equipment experience will receive extensive equipment operation training plus procedural training (2 weeks); Three Air Force NDI technicians with limited ultrasonic scanning/imaging equipment experience will receive equipment operation refresher training plus procedural training (1 week). Specific training details and curriculum will be detailed in the final plan.

POD Protocols: The protocol package will detail all aspects of the experiment to ensure consistency throughout all phases. The protocols are comprised of step-by-step procedures and descriptions which control the following: Monitors Duties, Inspectors Duties, In-briefs & Out-briefs, Specimen arrangements & control, Data Recording, Data Analysis.

Data Analysis: Data will be presented using the methodologies as covered in MIL-STD-1823 and will include POD curves and ROC curves. Data will be presented in such a manner that any number of variables affecting the inspection results can be extracted.

Specimen Cracking and Characterization Results

Sixteen splice-joint segments underwent initial characterization, including hole sizing and spacing measurements, tab thickness measurements, and bolt hole eddy current testing. Thirteen of the sixteen segments underwent cyclic fatigue testing to grow cracks in the second-layer tab. The approach to obtaining controlled crack growth in the segments was to sequentially initiate cracks, largest to smallest, with 0.020 inch starter notches. The notches were cut into the designated fastener sites and cracks grown to the desired length. Doublers were installed in some instances to retard crack growth, once the desired crack length was achieved. The goal was to generate 60 cracked fastener sites, with target lengths ranging from 0.030 inch to 0.250 inch.

Cracks were initially characterized by visual, optical means and ultrasonic immersion testing. Following crack measurement, cracked holes are oversized to remove starter notches. The panels will then be reassembled with sealant and the appropriate size taper-lok fasteners and painted per C-141 standards. A final immersion ultrasound characterization is scheduled for each assembled panel. A complete characterization package will accompany each test specimen.

Typical crack length characterization results are shown in Table 1 for segment #7200. As noted in Table 1, cracks did not initiate in all notched holes. There were also instances in other segments when cracks initiated in unnotched holes. For these reasons, the initial crack distribution plans which specified crack locations and sizes within the segments was adjusted as necessary to achieve the original crack distribution goals. A total of 60 cracks have been generated. Following fastener hole reaming operations to remove the crack starter notches, the final crack length range will be 0.025 inch to 0.375 inch. Cracks are evenly distributed in the forward and aft directions.

TABLE 1. Segment 7200 Fatigue Crack Data
Section & Hole #	Direction	Starter Notch	Hole Dia.	Length Actual	Taper-Lok Dia.	Length Final
7203	FWD	0.018	0.250	0.118	0.375	0.0555
7205	FWD	0.022	0.250
7208	FWD	0.020	0.250	0.124	0.375	0.0615
7210	AFT	0.020	0.250	0.106	0.375	0.0435
7219	AFT	0.022	0.250	0.151	0.375	0.0885
7220	FWD	0.022	0.250
7225	FWD	0.020	0.250
7227	AFT	0.020	0.250
7229	AFT	0.020	0.250	0.300	0.375	0.2375
7231	FWD	0.022	0.250	0.289	0.375	0.2265
7232	AFT	0.020	0.250

Air Force Implementation Plan

The POD study is expected to be completed in September '97. With POD data, the C-141 SPD can confidently determine the appropriate inspection interval for implementation of the new splice-joint inspection process. It is expected that the POD study will prove the new process to have a detection threshold equivalent to bolt hole eddy current (~ 0.075 inch) and potentially as low as 0.050 inch. The SPD goal is to perform a one-time, fleet-wide inspection to assess fleet splice-joint condition. Once commenced, the fleet inspection is expected to be completed within a 6 to 12-month time period, at an estimated cost of $75K per aircraft. It is anticipated, that a savings of over $100 Million will be realized by accomplishing fleet-wide inspection with the new process versus BHEC. The SPD plans to begin fleet inspection as soon as funding is identified and approved at Command Level. Numerous options exist to accomplish fleet inspection operations including: Air Force field teams, contractor field teams, or establishment of a depot inspection speed-line. The SPD must determine the most cost effective option of inspection with minimal aircraft downtime and affect on aircraft availability.

When fleet-wide inspection has been completed, the SPD will subsequently establish a 5-year depot-level inspection interval concurrent with the C-141 Programmed Depot Maintenance (PDM). Once a PDM requirement is implemented, field-level inspection of splice-joints will no longer be required. The new splice-joint inspection process can be accomplished during PDM without any addition of flow-days. Additional splice-joint inspection costs at PDM will be more than offset from savings generated by elimination of the ongoing 120 day interval field-level inspection requirement.

References

1. Lockheed Aeronautical Systems Company, "C-141 Aircraft Structural Integrity Program (ASIP) 1994," Lockheed Report No. LG95ER0091, prepared for Warner Robins Air Logistics Center under contract FO9603-93-G-0011, May 1995
2. Lockheed Aeronautical Systems Company, "Wing Teardown Inspection of C-141 Aircraft 66-0186," Lockheed Report No. LG93ER0087, prepared for Warner Robins Air Logistics Center under contract FO9603-93-G-0011, August 1993
3. Komsky, I., Achenbach, J., Andrew, G., Grills, R., Register, J., Linkert, G., Hueto, G., Steinberg, A., Ashbaugh, M., Moore, D., and Weber, H., "An Ultrasonic Technique to Detect Corrosion in DC-9 Wing Box from Concept to Field Application," Mat. Eval., Vol. 53, No. 7, pp. 848-852, 1995.
4. Achenbach, Jan D., Komsky, Igor N., Lee, Y.C., and Angel, Yves C., "Self Calibrating Ultrasonic Technique For Crack-Depth Measurement," Journal of Nondestructive Evaluation, Vol. 11, 1992, pp. 103-108.
5. Spencer, F., Borgonovi, G., Roach, D., Schurman, D., and Smith, R., "Reliability Assessment at Airline Inspection Facilities, Volume I: A Generic Protocol for Inspection Reliability Experiments", DOT/FAA/CT-92/12,I, Department of Transportation, FAA Technical Center, May 1993.
6. Andrew, G., MacInnis, T., Mullis, T., "Second Layer Ultrasonic Inspection of C-141 Splice Joints", presented at the SPIE Conference on Nondestructive Evaluation Techniques for Aging Infrastructure and Manufacturing, Scottsdale, AZ, December 3 - 5, 1996.

Authors:

Mr. Roy T. Mullis tmullis@ti.robins.af.mil Warner Robins Air Logistics Center Technology & Industrial Support Directorate Materials Analysis Team (TIEDM) Robins Air Force Base, GA 31098

Mr. Timothy MacInnis TIMOTHY.J.MACINNIS@cpmx.saic.com Ultra Image International A Division of Science Applications International Corporation New London, CT 06320 For product information e-mail Glenn A. Andrew: andrewg@cpva.saic.com SAIC Home Page: http://www.saic.com

For more information see: NDT in Aerospace - UTonline 11/97