## 32441 Federal Register / Vol. 80, No. 110 / Tuesday, June 9, 2015 / Rules and Regulations

The woven composite fan blades are a novel and unusual design feature that requires additional airworthiness standards for type certification of the LEAP–1A and –1C engine models.

Discussion of Comments

A notice of proposed special conditions, No. 33–14–02–SC, for the CFM LEAP–1A and—1C engine models was published in the Federal Register on Friday, November 14, 2014 (79 FR 68137). No comments were received and the special conditions are adopted as proposed.

Applicability

As discussed above, these special conditions are applicable to the LEAP –1A and –1C engine models. Should CFM apply at a later date for a change to the type certificate to include another model on the same type certificate incorporating the same novel or unusual design feature, the special conditions would apply to that model as well.

Conclusion

This action affects only certain novel or unusual design features on LEAP–1A and –1C engine models. It is not a rule of general applicability and applies only to CFM, who requested FAA approval of this engine feature.

List of Subjects in 14 CFR Part 33

Aircraft, Engines, Aviation Safety, Reporting and recordkeeping requirements.

The authority citation for these special conditions is as follows:

Authority: 49 U.S.C. 106(g), 40113, 44701,

44702, 44704.


### The Special Conditions

Accordingly, pursuant to the authority delegated to me by the Administrator, the following special conditions are issued as part of the type
certification basis for CFM LEAP–1A and –1C engine models.

1. Part 33, Requirements

In addition to the airworthiness standards in 14 CFR part 33, effective February 1, 1965, with Amendments 33–1 through 33–32 applicable to the LEAP–1A and –1C engine models, CFM
will:

(a) Conduct an engine fan blade
containment test with the fan blade failing at the inner annulus flow path
line instead of at the outermost retention groove.

(b) Substantiate by test and analysis, or other methods acceptable to the FAA,
that a fan disk and fan blade retention system with minimum material properties can withstand, without failure, a centrifugal load equal to two
times the maximum load the retention system could experience within approved engine operating limitations. The fan blade retention system includes the portion of the fan blade from the inner annulus flow path line inward to the blade dovetail, the blade retention components, and the fan disk and fan
blade attachment features.

(c) Using a procedure approved by the FAA, establish an operating limitation that specifies the maximum allowable number of start-stop stress cycles for the fan blade retention system. The life
evaluation must include the combined effects of high-cycle and low-cycle fatigue. If the operating limitation is less than 100,000 cycles, that limitation must be specified in Chapter 5 of the Engine Manual Airworthiness Limitations Section. The procedure
used to establish the maximum allowable number of start-stop stress cycles for the fan blade retention system will incorporate the integrity requirements in paragraphs (c)(1), (c)(2), and (c)(3) of these special conditions for the fan blade retention system.

(1) An engineering plan, which
establishes and maintains that the combinations of loads, material properties, environmental influences, and operating conditions, including the effects of parts influencing these parameters, are well known or predictable through validated analysis, test, or service experience.

(2) A manufacturing plan that identifies the specific manufacturing constraints necessary to consistently produce the fan blade retention system with the attributes required by the engineering plan.

(3) A service management plan that defines in-service processes for maintenance and repair of the fan blade retention system, which will maintain attributes consistent with those required by the engineering plan.

(d) Substantiate by test and analysis, or other methods acceptable to the FAA, that the blade design below the inner annulus flow path line provides multiple load paths and crack arresting features that prevent delamination or crack propagation to blade failure during the life of the blade.

(e) Substantiate that during the service life of the engine, the total probability of an individual blade retention system failure resulting from all possible causes, as defined in § 33.75, will be extremely improbable with a cumulative calculated probability
of failure of less than 10

¥ 9 

### per engine flight hour.

(f) Substantiate by test or analysis that not only will the engine continue to meet the requirements of § 33.75 following a lightning strike on the composite fan blade structure, but that the lightning strike will not cause damage to the fan blades that would prevent continued safe operation of the affected engine.

(g) Account for the effects of inservice deterioration, manufacturing variations, minimum material properties, and environmental effects during the tests and analyses required by paragraphs (a), (b), (c), (d), (e), and (f) of these special conditions.

(h) Propose fleet leader monitoring and field sampling programs that will monitor the effects of engine fan blade usage and fan blade retention system integrity.

(i) Mark each fan blade legibly and permanently with a part number and a
serial number.

Issued in Burlington, Massachusetts, on June 1, 2015.
Ann C. Mollica, Acting Manager, Engine and Propeller Directorate, Aircraft Certification Service.

[FR Doc. 2015–14084 Filed 6–8–15; 8:45 am]

**BILLING CODE 4910–13–P**



### DEPARTMENT OF TRANSPORTATION

**Federal Aviation Administration**

**14 CFR Part 39**

**[Docket No. FAA–2014–0227; Directorate Identifier 2013–NM–211–AD; Amendment**  **39–18165; AD 2015–11–02]**

**RIN 2120–AA64**



### Airworthiness Directives; Lockheed **Martin Corporation/Lockheed Martin Aeronautics Company Airplanes** 

**AGENCY** 

### : Federal Aviation Administration (FAA), DOT.

**ACTION** 

### : Final rule.

**SUMMARY** 

### : We are superseding Airworthiness Directive (AD) 95–26–11 for all Lockheed Martin Corporation/ Lockheed Martin Aeronautics Company Model L–1011 series airplanes. AD 95– 26–11 required repetitive inspections to detect cracking of the fittings that attach the aft pressure bulkhead to the fuselage stringers, repetitive inspections to detect cracking of the fittings and of the splice tab of the aft pressure bulkhead, and corrective actions if necessary. This new AD requires repetitive inspections to detect cracking of the fittings that attach the aft pressure bulkhead to the fuselage stringers, repetitive inspections to detect cracking of the fittings and of the splice tab of the aft pressure bulkhead, repetitive inspections for cracking of


## 32442 Federal Register / Vol. 80, No. 110 / Tuesday, June 9, 2015 / Rules and Regulations

certain aft fuselage skin panels, astructural modification, a postmodification inspection program, and corrective actions if necessary. This AD was prompted by a determination that the fittings at stringer attachments to the upper region of the aft pressure bulkhead are subject to widespread fatigue damage (WFD). We are issuing this AD to prevent simultaneous failure of multiple stringer end fittings through fatigue cracking at the aft pressure bulkhead, which could lead to rapid decompression of the airplane.

**DATES** 

### : This AD is effective July 14,

2015.

The Director of the Federal Register approved the incorporation by reference of a certain publication listed in this AD as of July 14, 2015.

The Director of the Federal Register approved the incorporation by reference of a certain other publication listed in this AD as of January 11, 1996 (60 FR 668702, December 27, 1995).

**ADDRESSES** 

### : For service information identified in this AD, contact Lockheed Martin Corporation/Lockheed Martin Aeronautics Company, L1011 Technical Support Center, Dept. 6A4M, Zone 0579, 86 South Cobb Drive, Marietta, GA 30063–0579; telephone 770–494– 5444; fax 770–494–5445; email L1011.support@lmco.com; Internet http://www.lockheedmartin.com/ams/
tools/TechPubs.html. You may view this
referenced service information at the FAA, Transport Airplane Directorate, 1601 Lind Avenue SW., Renton, WA. For information on the availability of this material at the FAA, call 425–227–
1221. It is also available on the Internet

at http://www.regulations.gov by searching for and locating Docket No.

FAA–2014–0227.

Examining the AD Docket

You may examine the AD docket on
the Internet at http:// www.regulations.gov by searching for and locating Docket No. FAA–2014– 0227; or in person at the Docket Management Facility between 9 a.m. and 5 p.m., Monday through Friday, except Federal holidays. The AD docket contains this AD, the regulatory evaluation, any comments received, and
other information. The address for the Docket Office (phone: 800–647–5527) is Docket Management Facility, U.S. Department of Transportation, Docket Operations, M–30, West Building Ground Floor, Room W12–140, 1200 New Jersey Avenue SE., Washington,

DC 20590.

**FOR FURTHER INFORMATION CONTACT** 

### : Carl Gray, Aerospace Engineer, Airframe Branch, ACE–117A, FAA, Atlanta Aircraft Certification Office (ACO), 1701 Columbia Avenue, College Park, GA 30337; phone: 404–474–5554; fax: 404– 474–5605; email: c arl.w.gray@faa.gov.

**SUPPLEMENTARY INFORMATION** 

### :

Discussion

We issued a supplemental notice of proposed rulemaking (SNPRM) to amend 14 CFR part 39 to supersede AD 95–26–11, Amendment 39–9469 (60 FR 66870, December 27, 1995). AD 95–26– 11 applied to all Lockheed Martin Corporation/Lockheed Martin Aeronautics Company Model L–1011 series airplanes. The SNPRM published
in the Federal Register on November 17, 2014 (79 FR 68377). We preceded the SNPRM with a notice of proposed rulemaking (NPRM) that published in
the Federal Register on April 14, 2014 (79 FR 20819). The NPRM proposed to require repetitive inspections to detect cracking of the fittings that attach the aft pressure bulkhead to the fuselage stringers, repetitive inspections to detect cracking of the fittings and of the splice tab of the aft pressure bulkhead, repetitive inspections for cracking of certain aft fuselage skin panels, astructural modification, a postmodification inspection program, and corrective actions if necessary. The NPRM was prompted by a determination that the fittings at stringer attachments to the upper region of the aft pressure bulkhead are subject to WFD. The SNPRM proposed to reduce the post-structural modification repetitive inspection interval. We are issuing this AD to prevent simultaneous failure of multiple stringer end fittings through fatigue cracking at the aft pressure bulkhead, which could lead to rapid decompression of the airplane.

Comments

We gave the public the opportunity to participate in developing this AD. We
received no comments on the SNPRM (79 FR 68377, November 17, 2014) or on
the determination of the cost to the public.

Conclusion

We reviewed the relevant data and determined that air safety and the public interest require adopting this AD as proposed except for minor editorial changes. We have determined that these minor changes:

• Are consistent with the intent that was proposed in the SNPRM (79 FR 68377, November 17, 2014) for correcting the unsafe condition; and

• Do not add any additional burden upon the public than was already proposed in the SNPRM (79 FR 68377, November 17, 2014).

We also determined that these changes will not increase the economic burden on any operator or increase the scope of this AD.

Related Service Information Under 1

CFR Part 51

We reviewed Lockheed Service Bulletin 093–53–105, Revision 3, dated May 31, 2013. The service information describes procedures for inspections for cracking of the stringer end fittings at the aft pressure bulkhead, corrective actions, and a modification that includes replacement of the stringer end fittings of certain stringers with new fittings. This service information is reasonably available because the interested parties have access to it through their normal course of business or by the means identified in the

**ADDRESSES** 

### section of this AD.

Costs of Compliance

We estimate that this AD affects 26 airplanes of U.S. registry.

We estimate the following costs to comply with this AD:



### E STIMATED 

### COSTS

Action Labor cost Parts cost Cost per product Cost on U.S. operators

Inspections [actions retained

from AD 95–26–11, Amendment 39–9469 (60 FR 66870, December 27, 1995)]. 23 work-hours × $85 perhour = $1,955 per inspection cycle. $0 $1,955 per inspection cycle $50,830 per inspection cycle.

Inspections and modification

[new action]. 185 work-hours × $85 per hour = $15,725. 6,750 $22,475 ................................ $584,350.


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We estimate the following costs to do any necessary replacements that would be required based on the results of the inspections. We have no way of determining the number of aircraft that might need these replacements:



### O N 

### -CONDITION 

### COSTS

Action Labor cost Parts cost Cost per product

Replacement of one fitting ...................................... 16 work-hour × $85 per hour = $1,360 .................. $250 $1,610



### We have received no definitive data that would enable us to provide cost
estimates for the other on-condition actions specified in this AD.

Authority for This Rulemaking

Title 49 of the United States Code specifies the FAA’s authority to issue rules on aviation safety. Subtitle I, section 106, describes the authority of
the FAA Administrator. Subtitle VII: Aviation Programs, describes in more detail the scope of the Agency’s authority.

We are issuing this rulemaking under the authority described in Subtitle VII, Part A, Subpart III, Section 44701: ‘‘General requirements.’’ Under that section, Congress charges the FAA with promoting safe flight of civil aircraft in air commerce by prescribing regulations for practices, methods, and procedures the Administrator finds necessary for safety in air commerce. This regulation is within the scope of that authority
because it addresses an unsafe condition that is likely to exist or develop on products identified in this rulemaking

action.

Regulatory Findings

This AD will not have federalism implications under Executive Order
13132. This AD will not have a substantial direct effect on the States, on the relationship between the national government and the States, or on the distribution of power and responsibilities among the various levels of government.

For the reasons discussed above, I certify that this AD:

(1) Is not a ‘‘significant regulatory action’’ under Executive Order 12866,

(2) Is not a ‘‘significant rule’’ under DOT Regulatory Policies and Procedures (44 FR 11034, February 26, 1979),

(3) Will not affect intrastate aviation in Alaska, and

(4) Will not have a significant economic impact, positive or negative, on a substantial number of small entities under the criteria of the Regulatory Flexibility Act.

List of Subjects in 14 CFR Part 39

Air transportation, Aircraft, Aviation safety, Incorporation by reference, Safety. Adoption of the Amendment
Accordingly, under the authority delegated to me by the Administrator, the FAA amends 14 CFR part 39 as
follows:

**PART 39—AIRWORTHINESS**
**DIRECTIVES**

■ 

### 1. The authority citation for part 39
continues to read as follows:

Authority: 49 U.S.C. 106(g), 40113, 44701.

**§ 39.13 [Amended]**

■ 

### 2. The FAA amends § 39.13 by removing Airworthiness Directive (AD) 95–26–11, Amendment 39–9469 (60 FR 66870, December 27, 1995), and adding the following new AD:

2015–11–02 Lockheed Martin Corporation/ Lockheed Martin Aeronautics Company: Amendment 39–18165 ; Docket No. FAA–2014–0227; Directorate Identifier

2013–NM–211–AD.

(a) Effective Date

This AD is effective July 14, 2015.

(b) Affected ADs

This AD replaces AD 95–26–11, Amendment 39–9469 (60 FR 66870, December 27, 1995).

(c) Applicability

This AD applies to all Lockheed Martin Corporation/Lockheed Martin Aeronautics Company Model L–1011–385–1, L–1011– 385–1–14, L–1011–385–1–15, and L–1011– 385–3 airplanes, certificated in any category.

(d) Subject

Air Transport Association (ATA) of America Code 53, Fuselage.

(e) Unsafe Condition

This AD was prompted by a determination that the fittings at stringer attachments to the upper region of the aft pressure bulkhead are subject to widespread fatigue damage (WFD). We are issuing this AD to prevent simultaneous failure of multiple stringer end fittings through fatigue cracking at the aft pressure bulkhead, which could lead to rapid decompression of the airplane.

(f) Compliance

Comply with this AD within the compliance times specified, unless already
done.

(g) Retained Detailed Visual Inspection

This paragraph restates the requirements of paragraph (a) of AD 95–26–11, Amendment 39–9469 (60 FR 66870, December 27, 1995), with no changes. Perform a detailed visual inspection to detect cracking of the fittings that attach the aft pressure bulkhead to the fuselage stringers (hereinafter referred to as ‘‘fittings’’) at stringers 1 through 10 (right side) and at stringers 56 through 64 (left side), at the later of the times specified in either paragraph (g)(1) or (g)(2) of this AD.

(1) Prior to the accumulation of 20,000 total flight cycles; or

(2) Within the next 25 flight cycles or 10 days after September 28, 1995 (the effective date of AD 95–18–52, Amendment 39–9366 (60 FR 47465, September 13, 1995)),
whichever occurs earlier.

(h) Retained Corrective Action for Cracked Fitting

This paragraph restates the requirements of paragraph (c) of AD 95–26–11, Amendment 39–9469 (60 FR 66870, December 27, 1995), with no changes. If any cracked fitting is detected during the inspection required by paragraph (g) of this AD: Before further flight, accomplish the requirements of paragraphs (h)(1) and (h)(2) of this AD.

(1) Replace the cracked fitting with a new fitting, or with a serviceable fitting on which a detailed visual inspection has been performed previously to detect cracking and
that has been found to be free of cracks.

(2) Perform a detailed visual inspection to detect cracking in the radius at the lower end of the vertical leg of the bulkhead T-shaped frame between the stringer locations on either side of the stringer having the cracked fitting. If any cracked T-shaped frame is detected: Before further flight, repair in accordance with a method approved by the Manager, Atlanta Aircraft Certification Office (ACO), FAA.

(i) Retained Repetitive Fitting Inspections

This paragraph restates the requirements of paragraph (d) of AD 95–26–11, Amendment 39–9469 (60 FR 66870, December 27, 1995), with no changes. Repeat the inspections and other necessary actions required by paragraphs (g) and (h) of this AD at intervals not to exceed 1,800 flight cycles or 3,000 flight hours, whichever occurs earlier, until paragraph (j) of this AD is accomplished.

(j) Retained Eddy Current Surface Scan (ECSS) Inspections, and Related Investigative and Corrective Actions

This paragraph restates the requirements of paragraph (e) of AD 95–26–11, Amendment 39–9469 (60 FR 66870, December 27, 1995), with revised compliance times specified in paragraph (k) of this AD, exclusion of an ECSS inspection for certain airplanes, and new service information. Except as provided by paragraph (l) of this AD: At the applicable time specified in paragraph (k)(1) of this AD, accomplish the requirements of paragraphs


## 32444 Federal Register / Vol. 80, No. 110 / Tuesday, June 9, 2015 / Rules and Regulations

(j)(1) and (j)(2) of this AD. Repeat the ECSS inspections thereafter at the compliance time specified in paragraph (k)(2) of this AD. Accomplishment of the ECSS inspection constitutes terminating action for the repetitive inspection requirements of paragraph (i) of this AD.

(1) Perform an ECSS inspection to detect cracking of the fittings at stringers 1 through 14 (right side) and at stringers 52 through 64 (left side), in accordance with the Accomplishment Instructions of Lockheed L– 1011 Service Bulletin 093–53–105, Revision 1, dated November 17, 1995; or Lockheed Service Bulletin 093–53–105, Revision 3, dated May 31, 2013; except for airplanes with a large (47-inch-wide) aft passenger door, an ECSS inspection of stringers 12, 13, 53, and 54 is not required by this paragraph. Except as provided by paragraph (m) of this AD, if any cracking is detected, prior to further flight, replace the fitting with a new fitting without pilot holes, rework the fitting, and perform various follow-on actions (i.e., bolt hole eddy current (BHEC), ECSS, and borescope inspections; and repair) of the inner and outer tee caps, in accordance with the Accomplishment Instructions of
Lockheed L–1011 Service Bulletin 093–53– 105, Revision 1, dated November 17, 1995; or Lockheed Service Bulletin 093–53–105, Revision 3, dated May 31, 2013, except as required by paragraph (p) of this AD. As of the effective date of this AD, use only Lockheed Service Bulletin 093–53–105, Revision 3, dated May 31, 2013, for accomplishing the actions required by this paragraph.

(2) Perform an ECSS inspection to detect cracking of the lower (or inner) surface of the upper bonded splice tab of the bulkhead assembly at stringers 1 through 14 (right side) and at stringers 52 through 64 (left side), in accordance with the Accomplishment
Instructions of Lockheed L–1011 Service Bulletin 093–53–105, Revision 1, dated November 17, 1995; or Lockheed Service Bulletin 093–53–105, Revision 3, dated May 31, 2013. As of the effective date of this AD, use only Lockheed Service Bulletin 093–53– 105, Revision 3, dated May 31, 2013, for accomplishing the actions required by this paragraph.

(i) Except as provided by paragraph (m) of this AD, if any cracking is detected at the upper bonded splice tab, repair in accordance with a method approved by the Manager, Atlanta ACO, FAA.

(ii) Except as provided by paragraph (m) of this AD, if any cracking is detected at a fastener, prior to further flight, perform a BHEC inspection to detect cracking of the forward flange of the inner tee cap, in accordance with the Accomplishment
Instructions of Lockheed L–1011 Service Bulletin 093–53–105, Revision 1, dated November 17, 1995; or Lockheed Service Bulletin 093–53–105, Revision 3, dated May 31, 2013. If any cracking is detected, prior to further flight, repair in accordance with the Accomplishment Instructions of Lockheed L– 1011 Service Bulletin 093–53–105, Revision 1, dated November 17, 1995; or Lockheed Service Bulletin 093–53–105, Revision 3, dated May 31, 2013, except as required by paragraph (p) of this AD. As of the effective date of this AD, use only Lockheed Service Bulletin 093–53–105, Revision 3, dated May 31, 2013, for accomplishing the actions required by this paragraph.

(k) New Revised Compliance Times for Paragraph (j) of This AD

(1) Do the initial inspections required by paragraph (j) of this AD at the earlier of the times specified in paragraphs (k)(1)(i) and (k)(1)(ii) of this AD.

(i) Prior to the accumulation of 20,000 total flight cycles, or within 30 days after January 11, 1996 (the effective date of date of AD 95– 26–11, Amendment 39–9469 (60 FR 66870, December 27, 1995)), whichever occurs later.

(ii) At the later of the times specified in paragraphs (k)(1)(ii)(A) and (k)(1)(ii)(B) of
this AD.

(A) Before the accumulation of 13,875 total flight cycles.

(B) Within 365 days or 1,000 flight cycles after the effective date of this AD, whichever
occurs first.

(2) Repeat the inspections specified in paragraph (j) of this AD within 2,500 flight cycles after accomplishing the most recent inspection required by paragraph (j) of this AD, and repeat the inspection thereafter at intervals not to exceed 1,750 flight cycles.

(l) Retained Inspection Deferral for Paragraph (j) of This AD

This paragraph restates the requirements of paragraph (f) of AD 95–26–11, Amendment 39–9469 (60 FR 66870, December 27, 1995). Accomplishment of the initial ECSS inspections required by paragraph (j) of this AD may be deferred to a date within 120 days after January 11, 1996 (the effective date of date of AD 95–26–11), provided that, in the interim, a visual inspection as specified in paragraph (g) of this AD is accomplished within 30 days after January 11, 1996 (the effective date of date of AD 95–26–11), and repeated thereafter at intervals not to exceed 50 flight cycles. Once the ECSS inspections begin, the visual inspections may be
terminated.

(m) Retained Inspection Deferral With Revised Compliance Time and New Deferral

This paragraph restates the requirements of paragraph (g) of AD 95–26–11, Amendment 39–9469 (60 FR 66870, December 27, 1995), with a revised compliance time, service information, and a new deferred action. As of the effective date of this AD, the deferral specified in paragraphs (m)(1) and (m)(2) of this AD cannot be done. If cracking was found before the effective date of this AD, the deferral specified in paragraphs (m)(1) and (m)(2) of this AD may be done. (1) If two or more adjacent fittings on both sides of the cracked fittings or bonded splice tabs/
fasteners are determined to be free of cracks by the ECSS inspection required by paragraphs (j)(1) and (j)(2) of this AD, repeat the ECSS inspection of the adjacent fittings thereafter at intervals not to exceed 600 flight cycles until the cracked fittings or splice tabs/fasteners are replaced or repaired, in accordance with the Accomplishment
Instructions of Lockheed L–1011 Service Bulletin 093–53–105, Revision 1, dated November 17, 1995; or Lockheed Service Bulletin 093–53–105, Revision 3, dated May 31, 2013. At the applicable time specified in paragraphs (m)(1)(i) and (m)(1)(ii) of this AD: Replace the cracked fitting and/or splice tab/ fasteners, in accordance with the Accomplishment Instructions of Lockheed L–1011 Service Bulletin 093–53–105, Revision 1, dated November 17, 1995; or Lockheed Service Bulletin 093–53–105, Revision 3, dated May 31, 2013. As of the effective date of this AD, use only Lockheed Service Bulletin 093–53–105, Revision 3, dated May 31, 2013, for accomplishing the actions required by this paragraph.

(i) For any crack found before the effective date of this AD: Within 2,500 flight cycles after finding the crack.

(ii) For any crack found on or after the effective date of this AD: Before further flight after finding the crack.

(2) If two or more adjacent fittings on both sides of the cracked fittings or bonded splice
tabs/fasteners are determined to be free of cracks by the ECSS inspection required by paragraphs (j)(1) and (j)(2) of this AD, the follow-on inspection (i.e., BHEC, ECSS, and borescope inspections) of the inner and outer tee caps required by paragraph (j)(1) of this AD may also be deferred until the cracked fittings are replaced as required by paragraph (m)(1) of this AD, but no later than before the accumulation of 20,800 total flight cycles.

(n) New Repetitive Borescope Inspections of Certain End Fittings and Corrective Actions

For airplanes with a large (47-inch-wide) aft passenger door: At the later of the times specified in paragraphs (n)(1) and (n)(2) of this AD, do a borescope inspection for cracking of the stringer end fittings at stringer locations 12, 13, 53, and 54; and do all applicable related investigative and corrective actions; in accordance with the Accomplishment Instructions of Lockheed Service Bulletin 093–53–105, Revision 3, dated May 31, 2013, except as specified in paragraph (p) of this AD. Do all applicable related investigative and corrective actions before further flight. Repeat the inspection of the stringer end fittings thereafter at intervals not to exceed 1,750 flight cycles until the actions required by paragraph (q) of this AD
have been done.

(1) Before the accumulation of 13,875 total flight cycles.

(2) Within 365 days or 1,000 flight cycles after the effective date of this AD, whichever
occurs earlier.

(o) New Repetitive Borescope Inspections of Fuselage Skin Panels

For airplanes with a large (47-inch-wide) aft passenger door: At the later of the times specified in paragraphs (o)(1) and (o)(2) of this AD, do an ECSS inspection for cracking of the left and right aft fuselage skin panels; and do all applicable related investigative and corrective actions; in accordance with the Accomplishment Instructions of Lockheed Service Bulletin 093–53–105, Revision 3, dated May 31, 2013, except as specified in paragraph (p) of this AD. Do all applicable related investigative and corrective actions before further flight. Repeat the inspection of the aft fuselage skin panels thereafter at intervals not to exceed 1,750 flight cycles until the modification required by paragraph (q) of this AD is done.


## 32445 Federal Register / Vol. 80, No. 110 / Tuesday, June 9, 2015 / Rules and Regulations

(1) Before the accumulation of 13,875 total flight cycles.

(2) Within 365 days or 1,000 flight cycles after the effective date of this AD, whichever
occurs first.

(p) New Service Information Exception

If any cracking is found during any inspection required by this AD, and Lockheed Service Bulletin 093–53–105, Revision 3, dated May 31, 2013, specifies contacting Lockheed for appropriate action: Before further flight, repair the cracking in accordance with a method approved by the Manager, Atlanta ACO, FAA. As of the effective date of this AD, for a repair method to be approved by the Manager, Atlanta ACO, as required by this paragraph, the Manager’s approval letter must specifically refer to this

AD.

(q) New Pre-Structural Modification Inspections and Structural Modification

Before the accumulation of 20,800 total flight cycles: Do the applicable actions specified in paragraphs (q)(1) and (q)(2) of
this AD.

(1) Perform pre-structural modification inspections by doing the actions required by paragraphs (j), (n), and (o) of this AD.

(2) Perform a structural modification of the aft pressure bulkhead by removing and replacing all stringer end fittings with new or refurbished fittings at stringers 1 through 14, and 52 through 64, in accordance with the Accomplishment Instructions of Lockheed Service Bulletin 093–53–105, Revision 3, dated May 31, 2013.

(r) New Post-Structural Modification Repetitive Inspections

Within 13,875 flight cycles after performing the actions required by paragraph (q)(2) of this AD: Do the actions specified in paragraphs (j), (n), and (o) of this AD, and repeat thereafter at intervals not to exceed 1,750 flight cycles.

(s) No Reporting Requirement

Although Lockheed Service Bulletin 093– 53–105, Revision 3, dated May 31, 2013, referenced in this AD specifies to submit certain information to the manufacturer, this AD does not include that requirement.

(t) Alternative Methods of Compliance (AMOCs)

(1) The Manager, Atlanta ACO, FAA, has the authority to approve AMOCs for this AD, if requested using the procedures found in 14 CFR 39.19. In accordance with 14 CFR 39.19, send your request to your principal inspector or local Flight Standards District Office, as appropriate. If sending information directly to the manager of the ACO, send it to the attention of the person identified in paragraph (u) of this AD.

(2) Before using any approved AMOC, notify your appropriate principal inspector, or lacking a principal inspector, the manager of the local flight standards district office/ certificate holding district office.

(u) Related Information

For more information about this AD, contact Carl Gray, Aerospace Engineer, Airframe Branch, ACE–117A, FAA, Atlanta Aircraft Certification Office (ACO), 1701 Columbia Avenue, College Park, GA 30337; phone: 404–474–5554; fax: 404–474–5605;

email: carl.w.gray@faa.gov.

(v) Material Incorporated by Reference

(1) The Director of the Federal Register approved the incorporation by reference (IBR) of the service information listed in this paragraph under 5 U.S.C. 552(a) and 1 CFR
part 51.

(2) You must use this service information as applicable to do the actions required by this AD, unless the AD specifies otherwise.

(3) The following service information was approved for IBR on July 14, 2015.

(i) Lockheed Service Bulletin 093–53–105, Revision 3, dated May 31, 2013 (The date of May 15, 2013, on page 1 of Lockheed Service Bulletin 093–53–105, Revision 3, dated May 31, 2013, is incorrect and should be May 31, 2013).

(ii) Reserved. (4) The following service information was approved for IBR on January 11, 1996 (60 FR 66870, December 27, 1995).

(i) Lockheed L–1011 Service Bulletin 093– 53–105, Revision 1, dated November 17,

1995.

(ii) Reserved. (5) For Lockheed service information identified in this AD, contact Lockheed Martin Corporation/Lockheed Martin Aeronautics Company, L1011 Technical Support Center, Dept. 6A4M, Zone 0579, 86 South Cobb Drive, Marietta, GA 30063–0579; telephone 770–494–5444; fax 770–494–5445;

email L1011.support@lmco.com; Internet http://www.lockheedmartin.com/ams/tools/

TechPubs.html.

(6) You may view this service information at FAA, Transport Airplane Directorate, 1601 Lind Avenue SW., Renton, WA. For information on the availability of this material at the FAA, call 425–227–1221.

(7) You may view this service information that is incorporated by reference at the

National Archives and Records Administration (NARA). For information on the availability of this material at NARA, call
202–741–6030, or go to: http://www.archives.gov/federal-register/cfr/ibrlocations.html.

Issued in Renton, Washington, on May 18,

2015.
John P. Piccola, Jr.,
Acting Manager, Transport Airplane Directorate, Aircraft Certification Service.

[FR Doc. 2015–13325 Filed 6–8–15; 8:45 am]

**BILLING CODE 4910–13–P** 

### DEPARTMENT OF TRANSPORTATION

**Federal Aviation Administration**

**14 CFR Part 39**

**[Docket No. FAA–2015–1937; Directorate Identifier 2014–SW–067–AD; Amendment**  **39–18171; AD 2015–11–08]**

**RIN 2120–AA64**



### Airworthiness Directives; Agusta **S.p.A. (Agusta) Helicopters**

**AGENCY** 

### : Federal Aviation Administration (FAA), DOT.

**ACTION** 

### : Final rule; request for

comments.

**SUMMARY** 

### : We are superseding airworthiness directive (AD) 2014–02– 08 for Agusta Model A109C, A109S, A109K2, A109E, and AW109SP helicopters. AD 2014–02–08 required inspecting the lock wires securing the tail rotor (T/R) duplex bearing locking nut (locking nut) to determine whether any lock wires are missing or damaged.
This AD retains some of the requirements of AD 2014–02–08 but removes the terminating action, expands the applicability, and adds a daily pilot check. This AD was prompted by reports of loosening T/R locking nuts. These actions are intended to prevent failure of the T/R and subsequent loss of control of the helicopter.

**DATES** 

### : This AD becomes effective June

24, 2015.

We must receive comments on this AD by August 10, 2015.

**ADDRESSES** 

### : You may send comments by any of the following methods:

• Federal eRulemaking Docket: Go to http://www.regulations.gov. Follow the online instructions for sending your comments electronically.

• Fax: 202–493–2251.

• Mail: Send comments to the U.S. Department of Transportation, Docket Operations, M–30, West Building Ground Floor, Room W12–140, 1200 New Jersey Avenue SE., Washington,

DC 20590–0001.

• Hand Delivery: Deliver to the
‘‘Mail’’ address between 9 a.m. and 5 p.m., Monday through Friday, except Federal holidays.

Examining the AD Docket

You may examine the AD docket on
the Internet at http:// www.regulations.gov or in person at the Docket Operations Office between 9 a.m. and 5 p.m., Monday through Friday, except Federal holidays. The AD docket contains this AD, the European Aviation Safety Agency (EASA) AD, the economic evaluation, any comments