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**88/SB**

**- 7**  **22**

**FUSELAGE**

**– MODIFICATION**  **OF CIRCUMFERENTIAL SKIN AND STRINGER SPLICE AT** 
**F.S. 695**

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

On January 14, 2011, the Widespread Fatigue Damage rule of 14 CFR Part 26 became effective. In compliance with § 26.21 of this rule, Lockheed Martin established a Limit of Validity (LOV) 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 L OV published in the Electra Maintenance Manual, Chapter 4
Airworthiness Limit ations Section, Paragraph 13. A Structural
Modification Point ( SMP ) is defined for some circumferential stringer splice s at FS 695

.

C. DESCRIPTION

Structural Modification P oint ( SMP ) is defined to protect against widespread fatigue damage in the circumferential fuselage s plice at

FS 695

. Some splice fastener hole s in stringers 1

- 7 and 66

- 72 are modified to eliminate possible stringer cracks emanating from the splice fasteners.

D

. SAFETY INTENT

The safety intent of this service bulletin is to prevent loss of residual strength of the circumferential splice at FS 695

. M odification of the stringer splices prior to the SMP will ensure structural integrity of the FS 695 splice

.

E

. COMPLIANCE

The required modification to protect against widespread fatigue damage
in the FS 695 circumferential splice is to be accomplished at or prior

to the SMP of 38,20 0 flight hours. Any cracks found in the stringer splices during the modification shall be repaired before further flight.

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. APPROVAL

The EFFECTIVITY information of Section 1

. A. and the modification time
of Section 1. E

. and the technical content of the ACCOMPLISHMENT INSTRUCTIONS of Section 2 of this service bulletin are FAA approved.

G

. MANPOWER

Since t he work described in this service bulletin has not been performed at a Lockheed Martin facility, the estimated man

- power has not been validated and should not be used in making fixed price quotes for accomplishment of the work. The following estimates are provided only for planning purposes.

Gain Access and Restore aircraft 2 manhours Modify stringer splices 8 manh ours Inspect stringer splices 8 manhours
18 manhours

H

. MATERIAL AND TOOLING

Parts, materials and tooling listed in Section 3, in addition to replacement fasteners, paint, and other material are to be supplied/fabricated by the aircraft owner/operator.

I. WEIGHT AND BALANCE

Not affected

.

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 a nd 188C, September 27, 2012.

( 2 ) Electra Maintenance Manual

( 3 ) Electra Structural Repair Manual

(4) Electra Supplemental Inspection Document, LR 29428

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

E mail: 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.

A. MODIFICA TION

Perform the following modifica tion prior to the SMP specified in
Section 1.E. as follows:

(1) Jacking the aircraft in accordance with Maintenance Manual Chapter 7 as follows:

( a) Remove any excess weight from the fuselage aft of FS 695.

(b ) Jack the aircraft using either wing jacks or MLG jacks in conjunction with a nose landing gear jack. Do not use the nose jack fitting on the fuselage. Landing gear strut s should be lock ed to prevent possible strut deflection during the course of this modification

.

( c ) After the aircraft is jacked, s hore the aft fuselage by installing a contoured support board under the FS 1117 fuselage frame. The support board should be thick enough to span the FS 1117 frame cap width such that load will be dissipat ed in the structure without deformation

. Pad the shoring to eliminate the possibility of scratching or denting the structure. Reference SRM, 51

- 2

- 2, Section 3
and 51

- 2

- 3, Section 4. Apply a moderate amount of upward force to snug the contour board to th e FS 1117 frame. The aft fuselage may deflect slight upward, but the pitch attitude of the aircraft should not be changed.

( 2 ) Remove interior fuselage panels, insulation batting and
extraneous material to gain access to stringer splices on the
forward and aft sides of the FS 695 splice from stringer 7 to stringer 66.

( 3 ) Visually i nspect the forward and aft ends of the stringer splices
of s tringers 1

- 7 and 66

- 72 as shown in Figure 1 , Section B

- B for general condition, cracking, and loose or distressed fasteners.

( 4 ) At stringers 1

- 7 and 66

- 72

, remove the four (4) rivets common to the stringer and splice member at the forward and aft ends of the splice as shown in Figure 1, Section B

- B.

( 5 ) Perform bolt hole eddy current inspections in each of the four
(4) fastener hole s per Electra Supplemental Inspection Document, Section V, or Electra SRM 51

- 2

- 8, or an equivalent inspection procedure. Crack indications should be confirmed by an alternate inspection method.

(6) Holes without crack indications shall be reamed to the maximum permissible hole diameter of t he same size rivet per Table 1.

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- 1 4 Page 4 ( 7 ) Holes with crack indications may be reamed to the maximum permissible hole diameter of the next larger size rivet per Table 1. Re

- inspect hole per 2.A.(5) after reaming. If crack indication remains after reaming, r epair the crack ed s tringer per Electra SRM 53

- 2

- 2, as applicable, prior to further flight

.

( 8 ) Measure hole diameter and install replacement fastener per limits
shown in Table 1. If original hole is larger than max permissible diameter, step up to next rivet size and type. Maintain a minimum 3D center to center spacing and 1.75ED.

Table 1 Fastener Replacements

Rivet
Size Original
Rivet
Type Specification
Hole
Diameter Max Permissible

Hole Dia. Replacement
Rivet
Type 1/8 AN470AD

- 4 0.126

- 0.131 0.138 MS20470AD

- 4
5/32 n/a 0.158

- 0.162 0.171 MS20470AD

- 5
3/16 AN470DD

- 6 0.189

- 0.194 0.198 MS20470DD

- 6
1/4 n/a 0.252

- 0.259 0.263 MS20470DD

- 8

( 9 ) If a severed stringer is found during the inspection

, additional inspect ion of the fuselage skin is required. Perform external Eddy Current Surface inspection of the skin at the skin

- to

stringer attachment s immediately forward and aft of the stringer break as shown in Figure 1, View A

- A. Perform Eddy Curren t inspection per Electra Supplemental Inspection Document, Section

V

, or Electra SRM 51

- 2

- 8 Paragraph 7 , or equivalent inspection method. Confirm skin crack with dye penetrant per SID Section V

or SRM 51

- 2

- 8 Paragraph 1 or equivalent method. R epair skin per
SRM 53

- 2

- 1 and repair stringer per SRM 53

- 2

- 2, as applicable

, prior to further flight.

( 10 ) Re p ort any d amage d etec t ed by th ese inspecti o n s to L o ck h eed M a rtin

A ero n auti c s Com p any per paragraph 1.K

. 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

( 1 1 ) After modification, restore aircraft interior, remove shoring and dejack the aircraft.

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- 1 4 Page 5 Figure 1 FS 695 Circumferential Splice Inspection Locations

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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.