1-27-81 No. 57-52

4-30-14, Revision C Page 1 of 17

LOCKHEED MARTIN ELECTRA SERVICE BULLETIN

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 severed 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 with 33,000 

flight hours. Crack propagation analysis indicates this crack probably 

initiated at under 30,000 flight hours.

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 fail-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 inspections of the 

BL 65 wing root lower joint specified herein are to begin at the 

Inspection Start Point (ISP) and the structural modification specified 

herein is to be accomplished prior to the Structural Modification Point 

(SMP).
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C. DESCRIPTION

This change defines an ISP for the initial inspection for widespread 

fatigue damage in the BL 65 wing root joint. Visual, X-ray, ultrasonic 

and bolt hole eddy current inspections are recommended to detect 

possible cracks 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 damage 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 panels 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 severe 

than typical commercial usage.

(1) Inspection Start Point (ISP):

The initial inspection for widespread fatigue damage in the BL 65 wing 

root joint is to be accomplished at or prior to the ISP shown in 

Table 1, per Section 2.A. of this service bulletin. These inspections 

are to be repeated at the interval shown in Table 1 until 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 Section 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. Inspection 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, and the technical content of the 

ACCOMPLISHMENT INSTRUCTIONS of Section 2 of this service bulletin are 

FAA-approved.

G. MANPOWER

(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 

modification 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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Drawing No. Title

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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K. 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

Marietta, GA 30063-0591, USA

Telephone: (770) 494-9131

FAX: (770) 494-9122

Email: electra.support@lmco.com
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2. 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 account 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 19 and 20.

(2) Primary X-Ray Inspection

X-ray inspect both inboard and outboard sides of the lower splice 

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 contour 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 wing 

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 splice 

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 against the BL 65 rib tee vertical leg.
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(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 holes 

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 outer 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 the lower edge of another hole 

in opposite positions (Figure 2, View 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, 

respectively.

(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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(f) Procedure - Outer Panel Hole, 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 diameter hole and end of standard. 

Adjust baseline and sweep controls in order to place 

initial pulse at left of scope display, 0.250-inch 

diameter hole response in center of display and 

response from end of standard at right side of 

display (Figure 2, View B).

2 Locate 0.250-inch diameter hole (E, View A) and 

optimize response. Adjust gain to obtain a 95% 

response. Move transducer to obtain response 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, Views 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 BL65 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 areas.

(g) Procedure - Outer Panel Hole, interior crack check

1 Apply couplant on outer panel reference standard as 

above and locate 0.250-inch diameter hole (I, View 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-inch 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, Views 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 be made carefully in order to 
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minimize effects of exterior paint and interior fuel 

tank sealant. Response from tack rivet may occur 

between 0.250-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 - Center Section Panel Hole, 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, View B).

2 Locate 0.250-inch diameter hole (E, View 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, Views C, D). Inspection is performed 

by scrubbing transducer forward and aft of each 

fastener hole of BL65 wing splice joint. Scrubbing 

is from point 1/4-inch to 3/4-inch inboard. Crack 

response on scope display will be at right of 

0.250-inch diameter hole response.

(i) Procedure - Center Section Panel Hole, 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, Views 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) Preparation of part: Remove sealant from open fastener 

holes with a nylon bristled brush of the appropriate 

diameter and an approved alcohol or solvent.

(b) Inspection: Perform equipment calibration and hole 

inspection per procedure in Electra Structural 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-ups within the inspection area. 

Cut elox notches in the layers representing the wing skin 

plank and the rib lower cap in separate holes.

(6) Mark and Report Damage

Report any damage detected by these inspection procedures to

Lockheed Martin Aeronautics Company. Provide the following 

information of all damage detected:

(a) Aircraft Serial Number

(b) Flight Hours

(c) Part number and description

(d) Location and specific fastener hole location

(e) Length, depth and orientation 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 cracking, 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: 842394A 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 eddy 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 eddy current indications that can be removed 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 807688 splice plate is 

in place and located properly. Eddy current indications which 

are not removable in accordance with this procedure 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 instructions.

(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 making 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 the 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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(c) Remove fuel from all tanks and purge according 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 the chordwise 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) Install 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 Section 2.B.

(8) Touch up finish as required.
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Figure 1. X-Ray Tube Location – Above Wing
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Figure 2. Ultrasonic Inspection Techniques
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Figure 3 BL 65 Lower Wing Root Joint
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3. MATERIAL INFORMATION

All parts and material required to accomplish this change are not supplied 

and therefore, should be obtained and/or fabricated by the operator.