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88/SB

- 707C WING

- WIDESPREAD FATIGUE DAMAGE INSPECTION AND MODIFICATION OF WING LOWER SURFACE PANELS AND JOINT AT BL 65

1. PLANNING INFORMATION

A. EFFECTIVITY

This service bulletin is applicable to all Electra aircraft models, Models L

- 188A and L

- 188C, Serial No. 1001 and subsequent.

B. REASON

One operator reported cracks in two aircraft on the outboard side of the left wing lower surface splice joint at BL 65. The cracks were in panels two and three, generally through and along chordwise attachment holes. The cracks were found as a result of a fuel leak and a fuel stain. The aircraft with the fuel leak had a six

- inch crack in panel two and a completely sev ered panel three. Aircraft time was 41,098 flight hours. The other aircraft with 39,424 flight hours had a four

- inch crack in panel two. Check of the Lockheed house tear

- down aircraft with approximately 37,000 flight hours revealed small multiple cracks in panels two and three in both left and right wings.

Since release of the basic bulletin, other cracks up to 4 inches long have been found by means of visual, X

- ray, eddy current and ultrasonic inspections. The 4

- inch crack was found on an aircraft wit h 33,000 flight hours. Crack propagation analysis indicates this crack probably initiated at under 30,000 flight hour s.

Panel cracks in BL 65 lower chordwise splice joint may reduce the structural integrity of the wing and if allowed to remain unrepaired could precipitate or aggravate fuel leakage and cause cracking of adjacent panels. Cracks in adjacent wing panels at BL 65 splice joint area may jeopardize fa il

- safe capability of the wing.

On January 14, 2011, the Widespread Fatigue Damage rule of 14 CFR Part 26 became effective. In compliance with this rule, Lockheed Martin established a Limit of Validity for L

- 188 airplanes. The maintenance actions defined in this Service Bulletin are necessary to ensure that the airplane will be free from widespread fatigue damage up to the Limit of Validity published in the L

- 188 Maintenance Manual, Chapter 4 Airworthiness Limitations Section, Paragraph 13. The inspection s o f the BL 65 wing root lower joint specified herein are to begin at th e I nspection S tart P oint (ISP) and the structural modification specified herein is to be accomplished prior to the Structural M odification P oint ( SMP )

.

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This change defines an ISP for the initial inspection for widespread fatigue damage in the BL 65 wing root joint. V isual, X

- ray, ultrasonic and bolt hole eddy current inspections are recommended to detect possible c racks in the left and right wing lower surfaces, on the inboard and outboard sides of the BL 65 splice joint. Repeat inspection requirements are included.

This change also defines a SMP for a required modification to protect against widespread fatigue d amage in the BL 65 wing root joint

. The modification includes oversize and replacement of fasteners in the BL 65 wing root lower joint to skin panel attachments.

Repair drawings have been prepared for some locations where cracks have been reported and for adjacent areas.

D. SAFETY INTENT

The safety intent of this service bulletin is to prevent loss of residual strength of the BL 65 wing root lower joint due to multiple site damage in the chordwise fastener rows attaching the wing lower skin panel s to the rib lower cap

. Repetitive inspections beginning at the ISP and modification of the joint prior to the SMP will ensure structural integrity of the wing root joint

.

E

. COMPLIANCE

NOTE
Inspection and modification times contained herein are based on typical commercial transport aircraft usage. These time intervals should be adjusted to account for individual aircraft usage that is more sever e than typical commercial usage

.

(1) Inspection Start Point (ISP):

The initial inspection for widespread fatig ue damage in the BL 65 wing root joint is to be accomplished at or prior to the ISP shown in Table 1, per S ection 2.A. of this service bulletin

. The se inspections are to be repeated at the interval shown in Table 1 unti l the SMP is reached. Any cracks found in the joint shall be repaired before further flight

.

(2) Structural Modification Point (SMP):

The required modification to protect against widespread fatigue damage in the BL 65 wing root joint is to be accomplished at or prior to the
SMP shown in Table 1, per S ection 2. C . of this service bulletin

. The modification includes oversize and replacement of fasteners in the
BL 65 wing root lower joint rib cap to skin panel attachments

. Any cracks found in the joint shall be repaired before further flight

.

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Table 1

- Inspection and Modification of BL65 Lower Joint

BL 65 Lower Joint Configuration Inspection Start Point & Structural Modification Point (3)

ISP (4) SMP

Initial Repeat

Basic Wing

- unrepaired areas 19,000 FH 2,000 FH 29,000 FH

Wing areas covered by the 842383 or 842387 type repair

Wing areas covered by the 842394 preventive & repair installation

NOTES: 1. All figures indicate flight hours on aircraft.

2. Confirm crack indications by BHEC before structure disassembly. Eddy current hole probe inspection requires attachment removal. Extreme care must be exercised so that damage is not created during removal and reinstallation of fasteners. 3 . Applicable to lower surface panels only, both inboard and outboard of BL 65 splice. 4

. I nspection of non

- repaired areas is required.

F

. APPROVAL

The EFFECTIVITY information of Section 1, Paragraph A. and the ISP/SMP of Section 1, Paragraph E, a nd t he technical content of the ACCOMPLISHMENT INSTRUCTIONS of Section 2 of this service bulletin are
FAA

- approved.

G

. MAN POWER

(1) Less than one man

- hour is required to visually inspect each affected aircraft for fuel leaks.

(2) Approximately two man

- hours are required to drain inboard fuel tanks and sumps prior to X

- ray inspection.

(3) Approximately 20 man

- hours are required to X

- ray inspect each affected aircraft with film placed outside of wing.

(4) Approximately 40 man

- hours are required to purge and enter the fuel tank and X

- ray inspect with film placed inside of wing.

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(5) Approximately 60 man

- hours are required to bolt hole eddy current inspect each affected aircraft.

(6) Approximately 20 man

- hours are required to ultrasonic inspect each affected aircraft.

(7) Approximately 300 to 500 man

- hours are required to repair a typical crack area on either wing (after draining and purging the affected fuel tank and shoring aircraft).

(8) Approximately 400 man

- hours are required to accomplish the SMP modi fication on each affected aircraft.

H

. MATERIAL AND TOOLING
All parts and material and tooling required to accomplish this change are not supplied, and, therefore, should be obtained and/or fabricated by the operator.

I

. WEIGHT AND BALANCE
The change in aircraft weight and moment resulting from repairs should be computed by the operator based upon extent of repair accomplished.

J

. REFERENCES
(1) SIL

- 111, Supplemental Instructions: Airworthiness Limitations Section (ALS) and Widespread Fatigue Damage (WFD) Instructions for Continued Airworthiness (ICA) for L

- 188 Type Aircraft, Models 188A and 188C, October 19 , 2012.

( 2 ) Electra Maintenance Manual

( 3 ) Electra Structural Repair Manual

( 4 ) Electra Supplemental Inspection Document, LR 29428

( 5 ) The following drawings have been supplied with earlier issues of this bulletin. They contain information for fabricating inspection calibration “ test specimens ” and for repair of cracks in various areas.

Drawing No. Title

*841640 "B" Change Notes

- General & MLG Rib Shimming Instructions

842383 Repair Instl

- BL 65 Outboard, Lower Panel No. 1 thru No. 6

- 88/SK

- 707

842384 Finger Plate

- BL 65 Outboard, Lower

- 88/SK

- 707

842385 Plate

- BL 65 Outboard, Lower

- 88/SK

- 707

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842386 Fitting

- Lower Left Wing Outboard BL 65

- 88/SK

- 707

842387 Repair Instl

- BL 65 Inboard, Lower Panel No. 1 thru No. 9

- 88/SK

- 707

842392 Fitting

- Lower, Inboard BL 65

- 88/SK

- 707

842393 Fitting

- Lower, Underlying Tang, Inboard
BL 65

- 88/SK

- 707

** 842394 A Change Repair Instl

- BL 65 Inboard & (2 sheets) Outboard Panel 1,2, & 3 Lower Surface

- 88/SK

- 707

** 842395 A Change Plate

– Wing BL 65 Lower Splice Inner

- 88/SK

- 707

842396 Plate

- BL 65 Lower

- 88/SK

- 707

842397 Cap

– BL 65 Lower Rib

- 88/SK

- 707

842398 Fitting

- BL 65 Cap Splice

- 88/SK

- 707

842429 Test Specimen

- Wing, Lower Panel,
BL 65 Area Outer Section

842430 Test Specimen

- Wing, Lower Panel,
BL 65 Area Center Section

**842472 Angle

– BL 65 Rib Lower Cap

- Repair

– 88/SK

- 707

**842473 Doubler

– BL 65 Rib Lower Cap

– Outboard

– 88/SK

- 707

* Previously furnished with 88/SB

- 670B, dated 9

- 3

- 70

** Previously furnished with 88/SB

- 707B, dated 10

- 22

- 82 All others furnished with 88/SB

- 707A, dated 11

- 12

- 81

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. CORRESPONDENCE

Please reference this service bulletin number and direct all correspondence pertaining to this service bulletin, except purchase orders, to:

LOCKHEED MARTIN AERONAUTICS COMPANY Enterprise Operations Center 86 South Cobb Drive Department 3E1M, Zone 0591 Marie tta, GA 30063

- 0591, USA

Telephone: (770) 494

- 9131
FAX: (770) 494

- 9122

Email: electra.support@lmco.com

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. ACCOMPLISHMENT INSTRUCTIONS

NOTE : This service bulletin is based on as

- delivered condition of aircraft by Lockheed Martin to original purchaser. In addition, Lockheed Martin has taken into account other known Lockheed Martin service bulletins and/or modifications mandated by the FAA . Operators accomplishing this inspection and modification must ac count for variations in aircraft configuration due to incorporation of service bulletin and/or Airworthiness Directive modifications and repairs.

A. INSPECTION

(1) Visually inspect each wing lower surface at BL 65 splice joint for signs of fuel leakage , as feasible during pre

- /post

- flight inspections and A/B/C checks in accordance with the operator’s current maintenance program. Refer to Maintenance Manual Chapter 28

- 2

- 1, Integral Fuel Tank Sealing, pages 1 9 and 20

.

( 2 ) Primary X

- Ray Inspection

X

- ray inspect both inboard and outboard sides of the lower s plice joint , front to rear spar , as described in the Electra Structural Repair Manual (SRM) Chapter 51

- 2

- 8, Figure 1, Sheet 1, Item 2, (lower joint only) except tube angle should be five degrees maximum. Use AA and M film externally , double loaded, with lead screens. Increase Milliamp seconds (MAS) to 450 to obtain film density of 3.0 to 3.5 in viewing area. Tube must be perpendicular to lower surface con tour while maintaining a 36

- inch focal distance as shown in Figure 1.

NOTE: It will be necessary to drain inboard fuel tanks and sumps prior to making radiographs.

NOTE: Confirm X

- ray indications with eddy current inspection per 2.A.(5) prior to repair.

( 3 ) Alternate X

- Ray Inspection

As an alternate to paragraph 2.A.( 2 ), the following procedure may be substituted. X

- ray inspect by placing film inside the w ing

over BL 65 lower splice fasteners and locating the X

- ray tube below the wing surface. Both inboard and outboard sides of BL 65 lower splice joint are to be X

- rayed from the front to rear spar. Inboard fuel tanks must be dry.

(a) The inboard side of the splice is to be X

- rayed by placing the tube inboard of BL 65. The outboard side of the spli ce is to be X

- rayed by placing the tube outboard of BL 65.

(b) The tube must be perpendicular within 5 degrees to the lower surface contour while maintaining a 36

- inch focal distance.

(c) Place the film inside the wing over splice fasteners. Use M Film Day

- Pak with paper folded back so that the film is abutted a gainst the BL 65 rib tee vertical leg

.

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- 14, Revision C Page 8 (d) Use 150 KV and 300 Milliamp Seconds (MAS) to obtain 3.0 to 3.5 film density in the viewing area.

NOTE: Confirm X

- ray indications with eddy current inspection per 2.A.(5) prior to repair.

( 4 ) Alternate Ultrasonic Inspection

Ultrasonic inspection techniques may be used as an alternate to
X

- ray inspection. The following ultrasonic inspection technique has been established for the detection of cracks that lie in a chord

- wise direction at, or between the 0.250

- inch diameter hole s in the wing lower panel surface, outboard and inboard of BL 65. This technique comprises two inspections

- one for center section panel holes and one for out er panel holes . Prior to performing this inspection, an area approximately 4 inches outboard and 4 inches inboard of BL 65 must be thoroughly cleaned and dried.

The following equipment and procedures have been successfully applied to a similar structure. See Figure 2 for all details.

(a) Calibration standard: A portion of L

- 188 aluminum wing panel or a standard made to exact drawing dimensions containing at least two 0.250

- inch diameter holes. Make two elox notches (0.250

- inch long x 0.100

- inch deep) on the top edge of one hole and on th e lower edge of another hole in opposite positions ( Figure 2, V iew A). It is recommended that one standard be made for the outer panel inspection and a second standard be made for the center section panel inspection. See drawings 842429 and 842430,
respec tively.

(b) Test equipment: Sonic Mark IV ultrasonic scope, or equivalent.

(c) Transducer: Krautkramer

- Branson MWB 70° 5mhz (shear wave), or equivalent.

(d) Couplant: Lubriplate, or equivalent.

(e) Control settings:

Range 2.0" Fine sweep 800 Delay 0

- 4 Reject 3 Damping 16 Coarse gain 60 db Fine gain 10 db Frequency 2 mhz Terminal T/R

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- O uter P anel H ole, exterior crack check

1 Using couplant, place transducer on the exterior surface of the outer panel reference standard approximately 1/4

- inch back from the 0.25

- inch splice plate radius and note relative responses from initial pulse, 0.250

- inch dia meter hole and end of standard. Adjust baseline and sweep controls in order to place initial pulse at left of scope display, 0.250

- inch diame ter hole response in center of display and response from end of standard at right side of display ( Figure 2, V iew B).

2 Locate 0.250

- inch diameter hole (E, V iew A) and optimize response. Adjust gain to obtain a 95% response. Move transducer to obtain resp onse from notch on exterior surface and optimize signal.

3 Apply couplant and place transducer on aircraft wing panel approximately 1/4

- inch outboard from wing lower splice joint (A, V iews C, D). Locate 0.250

- inch diameter hole and optimize response, increasing gain to obtain a 95% signal. Inspection is performed by scrubbing transducer forward and aft of each fastener hole of BL 65 wing splice joint. Scrubbing area is from splice joint to a point 3/8

- inch outboard. Crack response on scope display will be at right of 0.250

- inch diameter hole response.

NOTE: If it becomes difficult to maintain 95% signal response from fastener holes, it may be necessary to remove or thin paint in local a reas.

(g) Procedure

- O uter P anel H ole, interior crack check

1 Apply couplant on outer panel reference standard as above and locate 0.250

- inch diameter hole (I, V iew A) and optimize response. Adjust gain to obtain a 95% signal. Move transducer to obtain response from notch on interior surface and optimize signal by moving transducer approximately 5/8

- inch back from splice joint.

2 Apply couplant and place transducer on aircraft wing panel approximately 1/4

- in c h outboard from wing lower panel splice joint. Locate 0.250

- inch diameter hole and optimize response, increasing gain to obtain a 95% signal. Move transducer approximately 5/8

- inch back from splice joint (B, V iews C, D) and perform inspection by scrubbing transducer forward and aft of each fastener hole of BL 65 wing splice joint. Scrubbing area is from 1/2

- inch to 3/4

- inch from splice joint. Crack response on scope display will be at right of 0.250

- inch diameter hole response.

CAUTION: Gain changes must b e made carefully in order to

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- inch diameter hole response and response from end of panel. Careful scope display interpretation is necessary . Verification of crack response by eddy current bolt hole inspection is recommended.

(h) Procedure

- C enter S ection P anel H ole, exterior crack check

1 Using couplant, place transducer on the exterior surface of the center section panel reference standard as in paragraph 2.A.(5)f.1,( Figure 2, V iew B).

2 Locate 0.250

- inch diameter hole (E, V iew A) as in paragraph 2.A.(5)f.2.

3 Apply couplant and place transducer on aircraft center section panel approximately 1/4

- inch inboard from wing lower splice joint as in paragraph 2.A.(5)f.3. (C, V iews C, D). Inspection is performed by scrubbing transducer forward and aft of each fastener hole of BL 65 wing splice joint. Scrubbing is from point 1/4

- inch to 3/4

- inch inboard. Crack response on scope display wil l be at right of 0.250

- inch diameter hole response.

(i) Procedure

- C enter S ection P anel H ole, interior crack check

1 Apply couplant on center section panel reference standard as in paragraph 2.A.(5)g.1.

2 Move transducer to obtain response from notch on interior surface and optimize signal by moving transducer toward splice joint.

3 Apply couplant and place transducer on center section panel approximately 1/4

- inch inboard as in paragraph 2.A.(5)g.2, (D, V iews C, D). Inspection is performed as above in paragraph 2.A.(5).h.3. Scrubbing area is from edge of wing splice joint to a point 3/8

- inch inboard. Crack response on scope display will be at right of 0.250

- inch diameter hole response.

CAUTION: Gain changes must be made carefully in order to

minimize effects of exterior paint. No sealant has been applied to interior of center section. No tack rivets have been installed inboard of BL 65 splice joint. Verification of crack response by eddy current bolt hole inspection is recommended.

NOTE: Confirm ultrasonic indications with BHEC per 2.A.(5) prior to repair.

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(5) Bolthole Eddy Current Inspection

( a ) P reparation of part : R emove s e al a nt from open f a s ten e r hol e s with a n y lon br i stled brush of t h e appro p riate d i ameter and an approved alcohol or solvent.

( b ) Inspection : Perform equipment calibration and hole inspection per procedure in Electra S tructural Repair Manual, 51

- 2

- 8 Paragraph 6 Eddy Current Bolthole Procedure, or equivalent inspection procedure.

NOTE : Construct calibration standard(s) similar to that shown in SRM, 51

- 2

- 8, Figure 2 or SID , Section V, Inspection Techniques . The standard should represent all existing hole diameters and material stack

- u ps within the inspection area. Cut e lox notches in the layer s representing the wing ski n plank and the rib lower cap in separate holes

.

( 6 ) M ark and R eport D amage

Re p ort any d amage d etec t ed by th ese inspecti o n pr o cedure s to L o ck h eed M a rtin A ero n auti c s Com p any. Provide the followi n g in f or m ation of all d a mage det e ct e d:

( a ) Aircraft S erial Numb e r
( b ) Flight Ho ur s
( c ) Part numb e r and d e script i on
( d ) Locat i on and s p e cific faste n er h o le locati o n
( e ) Len g th, d e pth and ori e nta t ion of crack

Any cracks found in the joint shall be repaired before further flight

.

NOTE : Prior to reporting any detected damage to Lockheed Martin Aeronautics Company , or attempting to repair any detected damage , it is recommended that a borescope inspection (side view lens) of the hole bores be performed to determine if the reported damage is due to fatigue crackin g, embedded metal at faying surfaces or mechanical damage (gouging, reamer/drill marks, elongation) within the hole bore.

B. REPAIR

(1) Repair applicable areas in accordance with drawing 842383, 842387 or 842394 A Change

.

NOTE: 842394 A repair is the preferred method as it covers a greater area of the wing

.

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(2) Whenever a repair is made using 842383 or 842387 drawing repair, bolt hole e ddy current inspect all affected splice holes in inboard and outboard panels, P/N 804957 rib tee lower flanges, and P/N 807688 splice plate before repair installation and after hole rework (see paragraph 2.B.(3) below).

(3) Bolt hole e ddy current indications that can be r emoved by reaming oversize ( 5/16

- inch diameter maximum ) do not need added repair. Use minimum diameter oversize (1/64

- inch or 1/32

- inch) required to remove indications. Do not ream oversize any 2 adjacent fasteners or more than 10% of the fasteners in the chordwise joint. When reaming, be sure external P/N 8076 8 8 splice plate is in place and located properly. Eddy current indications which are not removable in accordance with this p rocedure will require repair.

(4) Cracks or eddy current indications remaining in the P/N 804957 rib cap tee lower flanges that are not removed by reaming in accordance with paragraph 2.B.(3) above require replacement of that portion of the rib cap tee. Splice rib cap tee in accordance with SRM 57

- 2

- 3, Figure 4, Sheet 2 in structions.

(5) When 842394 repair/reinforcement is to be installed, bolt hole eddy current inspect all affected panel holes prior to installation of new parts. Back drill new parts from existing paired holes where required.

NOTE: Change A to drawing 842394 (specifically sheet 2) contains an alternate repair installation. In this 842394

- 121/

- 122 repair installation a portion of the BL 65 rib lower tee cap is removed in one piece, this eliminating the cutting off of the lower horizontal flange and maki ng a cut in the rib tee cap at the forward end as is done by the 842394

- 101/

- 102 repair installation. However, removal of P/N 807353

- 1 fitting at the front lower spar cap in the fuel tank area will be required for

access.

C. MODIFICATION

The BL 65 wing root lower joint shall be modified prior to the SMP given in Section 1.E. Table 1. The structural modification to the lower BL 65 joint involves removing original fasteners in the chordwise joint which attach the wing lower skin planks to t he root rib lower cap and splice plate , as shown in Figure 3 , and replacing with first oversize fasteners

.

(1) Stabilize the aircraft as outlined below using the guidance of Maintenance Manual Chapter 7, Lifting and Shoring.

(a) The total aircraft weight shall not exceed 95,700 lbs per Maintenance Manual Section 7

- 1

- 1, Aircraft Jacking Weight Limitations.

(b) Ensure that airplane C.G. limitations versus gross weight are complied with.

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- 14, Revision C Page 13 ( c ) Remove fuel from all tanks and purge a ccording to established procedures.

( d ) Use main landing gear jacks and the nose (fuselage) jack per Maintenance Manual Section 7

- 1

- 1, Figure 201 to jack the aircraft. Discharge main landing gear struts before jacking.

( e ) Shore the outer wings at WS 311 and WS 499 using contour boards.

( f ) If engines are installed during this modification, take the weight of each engine off the wing structure by use of an engine hoist , or equivalent method. Use a hoist dynamometer to make certain that the pulling force on the engine hoist does not exceed the total engine weight.
(g) Limit movement of mechanics on the wing and in the aircraft while the wing is shored and engines are hoisted.

(2) Gain access to the inner parts of the center and outer wings by established procedures.

( 3 ) The removal/replacement of fasteners in the BL 65 lower splice joint shall be phased such that only 1 out of every 3 fasteners in a chordwise row is removed at any given time. Do not simultaneously remove 2 adjacent fasteners either in th e c hordwise or spanwise directions.

(4) Inspect each fastener hole per Section 2. A .(5)

. Note any hole that has a crack indication

.

(5) Ream each fastener hole to first oversize per SRM 51

- 2

- 12 Attachment Installation Requirements . Re

- inspect any hole that had a crack indication per Section 2. A

.

(6) I nstall new 1/64” oversize fasteners of the same type as original fasteners, or approved substitute type for original fasteners , in all holes with no crack indication

.

( 7 ) Repair confirmed cracks per Secti on 2.B.

( 8 ) Touch up finish as required.

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- 14, Revision C Page 14 Figure 1. X

- Ray Tube Location

– Above Wing

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- 14, Revision C Page 15 Figure 2. Ultrasonic Inspection Techniques

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- 14, Revision C Page 17 3. MATERIAL INFORMATION

All p arts and material required to accomplish this change are not supplied and therefore, should be obtained and/or fabricated by the operator.