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[Federal Register Volume 77, Number 20 (Tuesday, January 31, 2012)] [Rules and Regulations] [Pages 4650-4653] From the Federal Register Online via the Government Printing Office [www.gpo.gov] [FR Doc No: 2012-1953] –––––––––––––––––––––––––––––––––– DEPARTMENT OF TRANSPORTATION Federal Aviation Administration 14 CFR Part 39 [Docket No. FAA-2010-0068; Directorate Identifier 2010-NE-05-AD; Amendment 39-16930; AD 2012-02-07] RIN 2120-AA64 Airworthiness Directives; General Electric Company Turbofan Engines AGENCY: Federal Aviation Administration (FAA), DOT. ACTION: Final rule. –––––––––––––––––––––––––––––––––– SUMMARY: We are superseding two existing airworthiness directives (ADs) for General Electric Company (GE) CF6-45 and CF6-50 series turbofan engines with certain low-pressure turbine (LPT) rotor stage 3 disks installed. The existing ADs currently require inspections of high-pressure turbine (HPT) and LPT rotors, engine checks, and vibration surveys. This new AD retains the requirements of the two ADs being superseded, adds an optional LPT rotor stage 3 disk removal after a failed HPT blade borescope inspection (BSI) or a failed engine core vibration survey, establishes a new lower life limit for the affected LPT rotor stage 3 disks, and requires removing these disks from service at times determined by a drawdown plan. This AD was prompted by the determination that a new lower life limit for the LPT rotor stage 3 disks is necessary. We are issuing this AD to prevent critical life-limited rotating engine part failure, which could result in an uncontained engine failure and damage to the airplane. DATES: This AD is effective March 6, 2012. The Director of the Federal Register approved the incorporation by reference of a certain publication listed in this AD as of February 22, 2011 (76 FR 6323, February 4, 2011). ADDRESSES: For service information identified in this AD, contact General Electric Company, GE-Aviation, Room 285, 1 Neumann Way, Cincinnati, OH 45215, phone: (513) 552-3272; email: geae.aoc@ge.com. You may review copies of the referenced service information at the FAA, Engine & Propeller Directorate, 12 New England Executive Park, Burlington, MA. For information on the availability of this material at the FAA, call (781) 238-7125.

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Examining the AD Docket You may examine the AD docket on the Internet at http://www.regulations.gov; 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 Document 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: Tomasz Rakowski, Aerospace Engineer, Engine Certification Office, FAA, Engine & Propeller Directorate, 12 New England Executive Park, Burlington, MA 01803; phone: (781) 238-7735; fax: (781) 238-7199; email: tomasz.rakowski@faa.gov. SUPPLEMENTARY INFORMATION: Discussion We issued a notice of proposed rulemaking (NPRM) to amend 14 CFR part 39 to supersede AD 2011-02-07, Amendment 39-16580 (76 FR 6323, February 4, 2011) and AD 2011-18-01, Amendment 39-16783 (76 FR 52213, August 22, 2011). Those ADs apply to the specified products. The NPRM published in the Federal Register on October 19, 2011 (76 FR 64844). That NPRM proposed to retain the requirements of AD 2011-02-07 and AD 2011-18-01, except that reporting to the FAA would no longer be required and there would be an optional LPT rotor stage 3 disk removal after a failed HPT blade BSI or a failed engine core vibration survey. That NPRM also proposed to establish a new lower life limit for the LPT rotor stage 3 disk part numbers listed in Table 1 of the proposed AD, and proposed to require removing these disks from service at times determined by a drawdown plan. Comments We gave the public the opportunity to participate in developing this AD. The following presents the comments received on the proposal and the FAA's response to each comment. Support for the NPRM as Written One commenter, The Boeing Company, supports the NPRM (76 FR 64844, October 19, 2011) as written. Request To Allow Credit for Vibration Surveys Performed in a Test Cell One commenter, MTU Maintenance Hannover GmbH, requested that we add a paragraph that allows credit for performing vibration surveys in a test cell, as meeting the AD vibration survey requirements. We agree. We added paragraph (k)(8) to the AD, which states ''Vibration surveys carried out in an engine test cell as part of an engine manual performance run fulfill the vibration survey requirements of paragraphs (k)(2) through (k)(3) of this AD.''

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Request To Add a Requirement for Raw Exhaust Gas Temperature (EGT) Trend Data Point Exceedance One commenter, Evergreen International Airlines, requested that we add a requirement that two consecutive raw EGT trend data point exceedances must be confirmed by a corresponding shift of other engine parameters to trigger the HPT blade BSI. We partially agree. We agree that EGT system error should not force a BSI of turbine blades. But we disagree with troubleshooting the EGT raw data points once the EGT system error was ruled out. We added paragraph (o)(4) to the AD to state that, for the purposes of this AD, a raw EGT trend data point above the smoothed average is a confirmed temperature reading over the rolling average of EGT readings that is not a result of EGT system error. We also rearranged the wording in paragraph (iv) in Table 2 of the AD for clarification. Correction to Engine Model CF6-50-E2D Since we issued the NPRM (76 FR 67844, October 19, 2011), we discovered that, in applicability paragraph (c), engine model CF6-50-E2D was incorrect. We corrected it to read CF6-50E2B in the AD. Conclusion We reviewed the relevant data, considered the comments received, and determined that air safety and the public interest require adopting the AD with the changes described previously. We have determined that these minor changes: Are consistent with the intent that was proposed in the NPRM (76 FR 64844, October 19, 2011) for correcting the unsafe condition; and Do not add any additional burden upon the public than was already proposed in the NPRM. We also determined that these changes will not increase the economic burden on any operator or increase the scope of the AD. Costs of Compliance We estimate that this AD will affect 387 CF6-45 and CF6-50 series turbofan engines installed on airplanes of U.S. registry. We also estimate that it will take about 8 work-hours to perform the HPT blade inspection, 6 work-hours to perform a vibration survey, 4 work-hours to perform an ultrasonic inspection, 2 work-hours to perform an EGT resistance check, 1 work-hour to perform an EGT thermocouple inspection, and 7 work-hours to clean and perform an fluorescent-penetrant inspection of the LPT rotor stage 3 disk for each engine. The average labor rate is $85 per work-hour. The cost estimate for the work just described was covered in the two ADs we are superseding. For this AD, we estimate that a replacement LPT rotor stage 3 disk prorated part cost is $75,000. Based on these figures, we estimate the total cost of this AD to U.S. operators to be $29,025,000. 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

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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 We have determined that 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] The FAA amends § 39.13 by removing airworthiness directive (AD) 2011-02-07, Amendment 39-16580 (76 FR 6323, February 4, 2011) and AD 2011-18-01, Amendment 39-16783 (76 FR 52213, August 22, 2011), and adding the following new AD:

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FAA Aviation Safety

AIRWORTHINESS DIRECTIVE

www.faa.gov/aircraft/safety/alerts/ www.gpoaccess.gov/fr/advanced.html

2012-02-07 General Electric Company: Amendment 39-16930; Docket No. FAA-2010-0068; Directorate Identifier 2010-NE-05-AD. (a) Effective Date This airworthiness directive (AD) is effective March 6, 2012. (b) Affected ADs This AD supersedes AD 2011-02-07, Amendment 39-16580 (76 FR 6323, February 4, 2011) and AD 2011-18-01, Amendment 39-16783 (76 FR 52213, August 22, 2011). (c) Applicability This AD applies to General Electric Company (GE) CF6-45A, CF6-45A2, CF6-50A, CF6-50C, CF6-50CA, CF6-50C1, CF6-50C2, CF6-50C2B, CF6-50C2D, CF6-50E, CF6-50E1, CF6-50E2, and CF6-50E2B turbofan engines, including engines marked on the engine data plate as CF6-50C2-F and CF6-50C2-R, with any of the low-pressure turbine (LPT) rotor stage 3 disk part numbers listed in Table 1 of this AD installed.

Table 1–Applicable LPT Rotor Stage 3 Disk Part Numbers

9061M23P06 9061M23P07 9061M23P08 9061M23P09 9224M75P01

9061M23P10 1473M90P01 1473M90P02 1473M90P03 1473M90P04

9061M23P12 9061M23P14 9061M23P15 9061M23P16 1479M75P01

1479M75P02 1479M75P03 1479M75P04 1479M75P05 1479M75P06

1479M75P07 1479M75P08 1479M75P09 1479M75P11 1479M75P13

1479M75P14 N/A N/A N/A N/A

(d) Unsafe Condition This AD was prompted by the determination that a new lower life limit for the LPT rotor stage 3 disks listed in Table 1 of this AD is necessary. We are issuing this AD to prevent critical life-limited rotating engine part failure, which could result in an uncontained engine failure and damage to the airplane. (e) Compliance Comply with this AD within the compliance times specified, unless already done.

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(f) Borescope Inspections (BSI) of High-Pressure Turbine (HPT) Rotor Stage 1 and Stage 2 Blades For the BSIs required by paragraphs (f)(1), (f)(2), and (f)(3) of this AD, inspect the blades from the forward and aft directions. Inspect all areas of the blade airfoil. Your inspection must include blade leading and trailing edges and their convex and concave airfoil surfaces. Inspect for signs of impact, cracking, burning, damage, or distress. (1) Perform an initial BSI of the HPT rotor stage 1 and stage 2 blades within 10 cycles after the effective date of this AD. (2) Thereafter, repeat the BSI of the HPT rotor stage 1 and stage 2 blades within every 75 cycles since last inspection (CSLI). (3) Borescope-inspect the HPT rotor stage 1 and stage 2 blades within the cycle limits after the engine has experienced any of the events specified in Table 2 of this AD. (4) Remove any engine from service before further flight if the engine fails any of the BSIs required by this AD.

Table 2–Conditional BSI Criteria

If the engine has experienced: Then borescope-inspect:

(i) An exhaust gas temperature (EGT) above redline. Within 10 cycles.

(ii) A shift in the smoothed EGT trending data that exceeds 18 °F (10 °C), but is less than or equal to 36 °F (20 °C).

Within 10 cycles.

(iii) A shift in the smoothed EGT trending data that exceeds 36 °F (20 °C) Before further flight.

(iv) Two consecutive raw EGT trend data points that exceed 18 °F (10 °C), but is less than or equal to 36 °F (20 °C), above the smoothed average.

Within 10 cycles.

(v) Two consecutive raw EGT trend data points that exceed 36 °F (20 °C) above the smoothed average

Before further flight.

(g) Actions Required for Engines With Damaged HPT Rotor Blades For those engines that fail any BSI requirements of this AD, before returning the engine to service: (1) Remove the LPT rotor stage 3 disk from service; or (2) Perform a fluorescent-penetrant inspection (FPI) of the inner diameter surface forward cone body (forward spacer arm) of the LPT rotor stage 3 disk as specified in paragraphs (l)(1)(i) through (l)(1)(iii) of this AD. (h) EGT Thermocouple Probe Inspections (1) Inspect the EGT thermocouple probe for damage within 50 cycles after the effective date of this AD or before accumulating 750 CSLI, whichever occurs later. (2) Thereafter, re-inspect the EGT thermocouple probe for damage within every 750 CSLI. (3) If any EGT thermocouple probe shows wear through the thermocouple guide sleeve, remove and replace the EGT thermocouple probe before further flight, and ensure the turbine mid-frame liner does not contact the EGT thermocouple probe.

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(i) EGT System Resistance Check Inspections (1) Perform an EGT system resistance check within 50 cycles from the effective date of this AD or before accumulating 750 cycles since the last resistance check on the EGT system, whichever occurs later. (2) Thereafter, repeat the EGT system resistance check within every 750 cycles since the last resistance check. (3) Remove and replace, or repair any EGT system component that fails the resistance system check before further flight. (j) Ultrasonic Inspection (UI) of the LPT Rotor Stage 3 Disk Forward Spacer Arm Within 75 cycles after the effective date of this AD, perform a UI of the forward spacer arm of the LPT rotor stage 3 disk. Use Appendix A of GE Service Bulletin (SB) No. CF6-50 S/B 72-1312, Revision 1, dated October 18, 2010, paragraph 4. except for paragraph 4.(12), to do the UI. (k) Engine Core Vibration Survey (1) Within 75 cycles after the effective date of this AD, perform an initial engine core vibration survey. (2) Use about a one-minute acceleration and a one-minute deceleration of the engine between ground idle and 84% N2 (about 8,250 rpm) to perform the engine core vibration survey. (3) Use a spectral/trim balance analyzer or equivalent to measure the N2 rotor vibration. (4) If the vibration level is above 5 mils Double Amplitude then, before further flight, remove the engine from service. (5) For those engines that fail any engine core vibration survey requirements of this AD, then before returning the engine to service: (i) Remove the LPT rotor stage 3 disk from service; or (ii) Perform an FPI of the inner diameter surface forward spacer arm of the LPT rotor stage 3 disk as specified in paragraphs (l)(1)(i) through (l)(1)(iii) of this AD. (6) Thereafter, within every 350 cycles since the last engine core vibration survey, perform the engine core vibration survey as required in paragraphs (k)(1) through (k)(5) of this AD. (7) If the engine has experienced any vibration reported by maintenance or flight crew that is suspected to be caused by the engine core (N2), perform the engine core vibration survey as required in paragraphs (k)(1) through (k)(5) of this AD within 10 cycles after the report. (8) Vibration surveys carried out in an engine test cell as part of an engine manual performance run fulfill the vibration survey requirements of paragraphs (k)(2) through (k)(3) of this AD. (l) Initial and Repetitive FPI of LPT Rotor Stage 3 Disks (1) At the next shop visit after the effective date of this AD: (i) Clean the LPT rotor stage 3 disk forward spacer arm, including the use of a wet-abrasive blast, to eliminate residual or background fluorescence. (ii) Perform an FPI of the LPT rotor stage 3 disk forward spacer arm for cracks and for a band of fluorescence. Include all areas of the disk forward spacer arm and the inner diameter surface forward spacer arm of the LPT rotor stage 3 disk. (iii) Remove the disk from service before further flight if a crack or a band of fluorescence is present. (2) Thereafter, clean and perform an FPI of the LPT rotor stage 3 disk forward spacer arm, as specified in paragraphs (l)(1)(i) through (l)(1)(iii) of this AD, at each engine shop visit that occurs after 1,000 cycles since the last FPI of the LPT rotor stage 3 disk forward spacer arm.

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(m) Removal of LPT Rotor Stage 3 Disks Remove LPT rotor stage 3 disks listed in Table 1 from service as follows: (1) For disks that have fewer than 3,200 flight cycles since new (CSN) on the effective date of this AD, remove the disk from service before exceeding 6,200 CSN. (2) For disks that have 3,200 CSN or more on the effective date of this AD, do the following: (i) If the engine has a shop visit before the disk exceeds 6,200 CSN, remove the disk from service before exceeding 6,200 CSN. (ii) If the engine does not have a shop visit before the disk exceeds 6,200 CSN, remove the disk from service at the next shop visit after 6,200 CSN, not to exceed 3,000 cycles from the effective date of this AD. (n) Installation Prohibition (1) After the effective date of this AD, do not install or reinstall in any engine any LPT rotor stage 3 disk that exceeds the new life limit of 6,200 CSN. (2) Remove from service any LPT rotor stage 3 disk that is installed or re-installed after the effective date of this AD, before the disk exceeds the new life limit of 6,200 CSN. (o) Definitions (1) For the purposes of this AD, an EGT above redline is a confirmed over-temperature indication that is not a result of EGT system error. (2) For the purposes of this AD, a shift in the smoothed EGT trending data is a shift in a rolling average of EGT readings that can be confirmed by a corresponding shift in the trending of fuel flow or fan speed/core speed (N1/N2) relationship. You can find further guidance about evaluating EGT trend data in GE Company Service Rep Tip 373 ''Guidelines For Parameter Trend Monitoring.'' (3) For the purposes of this AD, an engine shop visit is the induction of an engine into the shop after the effective date of this AD, where the separation of a major engine flange occurs; except the following maintenance actions, or any combination, are not considered engine shop visits: (i) Induction of an engine into a shop solely for removal of the compressor top or bottom case for airfoil maintenance or variable stator vane bushing replacement. (ii) Induction of an engine into a shop solely for removal or replacement of the stage 1 fan disk. (iii) Induction of an engine into a shop solely for replacement of the turbine rear frame. (iv) Induction of an engine into a shop solely for replacement of the accessory gearbox or transfer gearbox, or both. (v) Induction of an engine into a shop solely for replacement of the fan forward case. (4) For the purposes of this AD, a raw EGT trend data point above the smoothed average is a confirmed temperature reading over the rolling average of EGT readings that is not a result of EGT system error. (p) Previous Credit (1) A BSI performed before the effective date of this AD using AD 2010-06-15, Amendment 39-16240 (75 FR 12661, March 17, 2010) or AD 2010-12-10, Amendment 39-16331 (75 FR 32649, June 9, 2010) or AD 2011-02-07, Amendment 39-16580 (76 FR 6323, February 4, 2011) within the last 75 cycles, satisfies the initial BSI requirement in paragraph (f)(1) of this AD. (2) A UI performed before the effective date of this AD using AD 2011-02-07, Amendment 39-16580 (76 FR 6323, February 4, 2011) or GE SB No. CF6-50 S/B 72-1312, dated August 9, 2010 or GE SB No. CF6-50 S/B 72-1312 Revision 1, dated October 18, 2010, satisfies the inspection requirement in paragraph (j) of this AD.

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(3) An engine core vibration survey performed before the effective date of this AD using AD 2011-02-07, Amendment 39-16580 (76 FR 6323, February 4, 2011) or GE SB No. CF6-50 S/B 72-1313, dated August 9, 2010 or GE SB No. CF6-50 S/B 72-1313 Revision 1, dated October 18, 2010, within the last 350 cycles, satisfies the initial survey requirement in paragraphs (k)(1) through (k)(5) of this AD. (4) An FPI of the LPT rotor stage 3 disk forward spacer arm performed before the effective date of this AD using AD 2011-18-01, Amendment 39-16783 (75 FR 52213, August 22, 2011), within the last 1,000 flight cycles of the LPT rotor stage 3 disk, satisfies the initial inspection requirements in paragraphs (l)(1)(i) through (l)(1)(iii) of this AD. (q) Alternative Methods of Compliance (AMOCs) (1) AMOCs previously approved for AD 2010-06-15, Amendment 39-16240 (75 FR 12661, March 17, 2010) are not approved for this AD. However, AMOCs previously approved for AD 2010-12-10, Amendment 39-16331 (75 FR 32649, June 9, 2010), AD 2011-02-07, Amendment 39-16580 (76 FR 6323, February 4, 2011), or AD 2011-18-01, Amendment 39-16783 (76 FR 52213, August 22, 2011) are approved for this AD. (2) The Manager, Engine Certification Office, may approve alternative methods of compliance for this AD. Use the procedures found in 14 CFR 39.19 to make your request. (r) Related Information (1) For more information about this AD, contact Tomasz Rakowski, Aerospace Engineer, Engine Certification Office, FAA, Engine & Propeller Directorate, 12 New England Executive Park, Burlington, MA 01803; phone: (781) 238-7735; fax: (781) 238-7199; email: tomasz.rakowski@faa.gov. (2) For service information identified in this AD, contact General Electric Company, GE-Aviation, Room 285, 1 Neumann Way, Cincinnati, OH 45215, phone: (513) 552-3272; email: geae.aoc@ge.com. You may review copies of the referenced service information at the FAA, Engine & Propeller Directorate, 12 New England Executive Park, Burlington, MA. For information on the availability of this material at the FAA, call (781) 238-7125. (s) Material Incorporated by Reference (1) You must use the following service information to do the UIs required by this AD, unless the AD specifies otherwise. The Director of the Federal Register approved the incorporation by reference (IBR) under 5 U.S.C. 552(a) and 1 CFR part 51 of the following service information on July 22, 2011: General Electric Company Service Bulletin No. CF6-50 S/B 72-1312 Revision 1, dated October 18, 2010. (2) For service information identified in this AD, contact General Electric Company, GE-Aviation, Room 285, 1 Neumann Way, Cincinnati, OH 45215, phone: (513) 552-3272; email: geae.aoc@ge.com. (3) You may review copies of the service information at the FAA, Engine & Propeller Directorate, 12 New England Executive Park, Burlington, MA. For information on the availability of this material at the FAA, call (781) 238-7125. (4) You may also review copies of the 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/ibr_locations.html.

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Issued in Burlington, Massachusetts, on January 20, 2012. Peter A. White, Manager, Engine & Propeller Directorate, Aircraft Certification Service.

ENGINEERING NOTICES

ACLF 056 Issue 2 15 May 2000

ISSUE: 001/12 DATE: 19/01/2012

Title: All Fleets - Honeywell EGPWS Troubleshooting

1. Background : Recently an increased number of Enhanced Ground Proximity Warning System (EGPWS) Computer replacements have resulted in many of these units being tested by Honeywell and found to have no faults. This engineering notice was drafted to inform maintenance engineers about additional procedures that should be used when troubleshooting the EGPWS.

2. Discussion : IAW Appendix B: “Troubleshooting Do’s and Do Not’s” from Honeywell’s EGPWS Line Maintenance Manual Do use the built-in self-test levels to find the exact nature of the fault. Do check the three diagnostic LED's on the front panel of the EGPWS and refer to the Troubleshooting Guide in Chapter 3 “Fault Isolation” of Honeywell’s EGPWS Line Maintenance Manual. - Green Computer OK LED means “don’t remove” - Amber External Fault LED means “don’t remove” - Red Computer means “possibly remove”

Do Not use cockpit INOP lamps solely for troubleshooting as cockpit INOP Lamps are merely a pilot’s indication of which EGPWS functions are inactive, they are not positive indications of internal EGPWS faults. - Experience shows External (non-EGPWS) faults are responsible for INOP illumination 99% of the

time. Do Not remove the EGPWC if front panel green Computer OK LED is illuminated For a hard fault the “Current Status” Bite (Level 2 Self-Test) should be used. For an intermittent fault the “Fault History” Bite (Level 4 Self-Test) should be used. - The EGPWC should only ever be replaced if either BITE confirms that there is an “Internal Fault”. - If the BITE confirms an “External Fault” then the BITE will guide the engineer to whatever part of

the system needs troubleshooting. Remember, the MkVII and MkVIII EGPWS have an integral GNSSU and both Internal and External Fault LEDs will be lit if there is a problem with the GNSSU.

Honeywell’s WinViews software program can also be used to assist with troubleshooting, however for most routine defects the simple BITE Test is more than adequate. Finally, an option exists to install a special one-time Flight History Download PCMCIA Card into the EGPWC and have the data sent to Honeywell for analysis. - The Flight History Download is primarily used to determine the cause of EGPWS alerts, only

recording data prior to and following an actual alert event. While this information may be useful in determining root cause of nuisance or valid alert warnings the data is not so useful for determining failures unless associated with the corresponding generation of a nuisance alert warning.

- See Chapter 4 of Honeywell’s EGPWS Line Maintenance Manual for further details.

ENGINEERING NOTICES

ACLF 056 Issue 2 15 May 2000

3. Additional Info : Honeywell EGPWS Line Maintenance Manual DWG NO. 060-4199-180 Rev. G. Dtd. 29 Mar 2010 http://www51.honeywell.com/aero/common/documents/egpws-documents/Maintenance-documents/060-4199-180.pdf Honeywell EGPWS Troublshooting Card: DWG NO. 060-4199-190 Rev. B. Dtd. 11 Feb 2005 http://www51.honeywell.com/aero/common/documents/EmTroubleshootingCard.pdf Honeywell Flight History PCMCIA Card Download Procedure DWG NO. 060-4199-115 Rev. C. http://www51.honeywell.com/aero/common/documents/EmFlightHistoryPCMCIA.pdf AMM Supplement: 34-18-11 WDM Supplement: 34-48-11 IPC Supplement: 34-48-11 4. Summary : A high rate of EGPWS units returned to Honeywell are found to have No Fault Found. Utilise Honeywell’s EGPWS Line Maintenance Manual and Troubleshooting Card to troubleshoot the EGPWS thoroughly. Be aware of Honeywell’s recommended Do’s and Do Not’s and the various options for BITE Tests and History downloads.

Prepared by: Authorised by:

David Reilly Éamon Stapleton

Technical Services Engineer Technical Services Manager

ENGINEERING NOTICES

ACLF 056 Issue 2 15 May 2000

ISSUE: 002/12 DATE: 20/03/2012

ATA 51 Structures: Repair Data Collection

Background: With the introduction of a requirement to assess all the repairs on aircraft, a review of repair

paperwork is being performed. In most circumstances, while the repair sign-off is acceptable under

Part 145, the information provided is not extensive enough should the repair need to be reviewed by the aircraft manufacturer.

Discussion: In order to ensure that the additional information is recorded for possible future reference a new

form has been created ref: ACLF.038 ‘Repair Data Form’.

The form has 3 main sections, ‘Damage Details’, ‘Repair Details’ and ‘Proximity with Adjacent Repairs’.

• ‘Damage Details’ – records the location of the damage, type of damage and the dimensions.

Note: it is acceptable to cross refer to any report (e.g. ATR SDR) already produced. • ‘Repair Details’ – records details of cut-outs if performed, fastener details, material details,

etc.

• ‘Proximity with Adjacent Repairs’ – records whether there are any nearby repairs.

This form will be supplied in a word format so that the MRO/line engineer is able to fill in all the details and increase size of fields etc. when required. Photographs of the repairs are highly desired

and can be pasted in field ‘i’ in section ‘Repair Details’. Note: If possible please provide a copy of the completed form to technical services in the word

format. The form can be printed and inserted with the repair sign-off also.

Summary:

• Fill in ACLF.038 everytime a repair is performed

• Cross referencing to reports (e.g. ATR SDR, sketch) is acceptable. Note: please attach

referenced document to the form. • Please see attached a completed example for reference. If you have any queries about

completing fields etc please ask the base maintenance representative or technical services.

Prepared by: Authorised by:

Thomas Bambrick Éamon Stapleton

Technical Services Engineer Technical Services Manager

Repair Data Form

Page 1 of 3

Form: ACLF.038 Issue: 1 Date: 20/03/2012

Damage Details:

Indicate Concerned Structural Item(s) (Fuselage, Wing, Stabilizer, Rudder) along with Location and

Orientation (Frame, Stringer, Rib, Spar, Fuselage Station, Buttock Line, Inboard, Outboard, LH, RH, ..):

Internal Fuselage FR41 Outboard STGR 20 RH

Type of Damage

Dent Buckle Gouge Delamination

Crack X Nick Scratch Corrosion

Other

Note: Indicate with “x” the type of damage, if “Other” please provide description.

Dimensions of damage (mm)

Diameter Length Width Depth

1.5

Repair Details:

a. Indicate Cutout Dimensions (If Applicable); N/A

Length: Height (Width): Diameter (Other):

b. Indicate Blend-Out Dimensions (If Applicable); N/A

Material Removed: Diameter (Other):

c. Fastener Part Number(s):

d. Fastener Type(s):

Note: Indicate with “x”

Rivet X Bolt

Solid Blind

Countersunk Protruded

Other

e. Fastener Pitch(es): 20 mm

f. Repair Material Spec: 7075-T62

g. Repair Material Thickness: 1.4 mm

h. Repair Reference (SRM Fig, SRI etc)

ATR SRI DO/TS-5681/1014413 Issue B

i. Indicate Type of Repair (Doubler, Splice)

Splice

Registration: EI-SLJ Flight Hours: 31,979:09 MRO:

MSN: 324 Flight Cycles: 46,954 Date: 29-Oct-10

Work Order/TLP Ref: TLP 39405 Work Card / NRC:

Repair Data Form

Page 2 of 3

Form: ACLF.038 Issue: 1 Date: 20/03/2012

j. Repair Description: Attach sketch, rubbing, picture or drawing of the repair

Additional Comments

k: Is there a Repetitive Inspection Requirement Yes No X

Details:

l: Repair Weight (kg) 0.5 kg

Proximity with Adjacent Repairs:

a.: Are there any repairs located on the same or adjacent structure (i.e. within one Stringer Bay, Frame Bay, Rib

Bay or Longeron Bay)

Yes No X

STGR 20 RH

FR 41

RH AFT

Repair Data Form

Page 3 of 3

Form: ACLF.038 Issue: 1 Date: 20/03/2012

b. If Yes, Indicate the edge to edge distance between each repair, describe and give repair reference;

Please fax/email the completed form to Technical Services.

Fax: +353 (0)1 812 0939 Email: Alltechservices@aircontractors.com

ENGINEERING NOTICES

ACLF 056 Issue 2 15 May 2000

ISSUE: 003/12 DATE: 26/03/2012

Standard Practices: Acceptance of Standard Parts

Background: This engineering notice has been raised to highlight the publication of EASA SIB 2012-06: “Defective

Standard Hardware – MS21042, NAS1291 and LN9338 Self-Locking Nuts, and NAS626 Bolts”.

A copy of SIB 2012-06 is attached.

Discussion:

The SIB was issued as several manufacturers have received numerous reports of defective standard hardware installed on different areas their products. In particular, many self-locking nuts have been

found cracked, parallel to the nut axis, in some instances only a short time after installation. Broken

bolts have also been found.

The SIB recommends that standard hardware is subjected to a close visual inspection for surface irregularities, such as gouges or cracks, as illustrated in Figures 1 and 2 of the SIB, before being

installed on a product. Suspect parts should be quarantined until conformity to the manufacturing

standard is verified.

Personnel are reminded that as per the Air Contractors MM&OE section 2.2 “Upon receipt, items must be inspected for transit damage, proper certification, documentation and compliance with the

Purchase Order or Repair Order by an approved store person or certifying engineer.”

Summary:

• Perform a close visual inspection before fitting any standard parts, in particular the afore

mentioned parts listed within the attached SIB.

• Should any irregularities be found with the parts, quarantine the affected parts and provide

logistics with the details of the part when requesting a replacement.

Prepared by: Authorised by:

Thomas Bambrick Éamon Stapleton Technical Services Engineer Technical Services Manager

EASA SIB No: 2012-06

This is information only. Recommendations are not mandatory.

EASA Form 117 Page 1/3

EASA Safety Information Bulletin SIB No.: 2012-06

Issued: 22 March 2012

Subject: Defective Standard Hardware – MS21042, NAS1291 and LN9338 Self-Locking Nuts, and NAS626 Bolts

Ref. Publications: - CAA Israel AD 57-10-06-18, dated 27 July 2010. - Gulfstream Aerospace LP Service Bulletin (SB) 200-51-366,

dated 30 March 2010. - AgustaWestland Information Letter GEN-11-024, dated 20 July

2011, which can be downloaded here. - Bell Helicopter Textron Operation Safety Notice GEN-11-43

Revision A, dated 16 September 2011 - Robinson Helicopter Company Service Letters R22 SL-58, R44 SL-38 and R66 SL-01 (single document), dated 18 August 2011. - L-3 Communications, Aviation Recorder Alert Bulletins: FA2100CVDR SB 015, dated 10 January 2011, FA2100CVR SB 018, dated 14 January 2011 and FA2100FDR SB 020, dated 14 January 2011.

- CASA Airworthiness Bulletin 14-002, dated 12 October 2011 - CAA New Zealand Continuing Airworthiness Notice 14-001,

dated 2 December 2011.

Applicability: The Part Numbers of the affected hardware are MS21042, NAS1291 and LN9338 self-locking nuts, and NAS626 bolts. The identified defective hardware is known to have been installed on rotorcraft, aeroplanes, appliances and engines.

Description: Several manufacturers (see ref. publications above) have received numerous reports of defective standard hardware installed on different areas of their products. In particular, many self-locking nuts have been found cracked, parallel to the nut axis, in some instances only a short time after installation. Broken bolts have also been found.

At this time, apart from the specific case of the Gulfstream 200 and GALAXY (Israel Aircraft Industries) aeroplanes, insufficient evidence is available to determine whether an unsafe condition exists that would warrant the issuance of an AD under EC 1702/2003, Part 21A.3B.

EASA, in cooperation with other aviation authorities and a number of aircraft, engine and equipment manufacturers, is currently investigating the reported occurrences to determine whether (and if so, what) further action is necessary.

EASA SIB No: 2012-06

This is information only. Recommendations are not mandatory.

EASA Form 117 Page 2/3

Recommendation: Manufacturers, aircraft owners, operators and maintenance staff, are recommended to subject the standard hardware (as identified in the Applicability paragraph of this SIB) to a close visual inspection for surface irregularities, such as gouges or cracks, as illustrated in Figures 1 and 2 of this SIB, before being installed on a product. Suspect parts should be quarantined until conformity to the manufacturing standard is verified.

As not all non-conformities to the manufacturing standards can be visually detected, the importance of quality acceptance criteria for standard parts is also emphasized.

Occurrences of cracked nuts or bolts should be reported through the established channels, as well as to the Agency Occurrence Reporting System, here.

Of particular interest are the hardware Part Number, manufacturer (or markings), lot number, location on the product and time since install.

Contact: For further information contact the Safety Information Section,

Executive Directorate, EASA. E-mail: ADs@easa.europa.eu.

EASA SIB No: 2012-06

This is information only. Recommendations are not mandatory.

EASA Form 117 Page 3/3

Figure 1: Example of surface irregularities observed on a nut. Overall view.

Figure 2: Example of surface irregularities observed on a nut. Detail

ENGINEERING NOTICES

ACLF 056 Issue 2 15 May 2000

ISSUE: 005/12 DATE: 04/04/2012

Title A300 ATA 21: MEL 21-1 CORRECTION

The purpose of this engineering notice is to highlight and correct an error discovered in chapter 21-1 of the Air Contractors A300 MEL Revision 2.

Figure 1

Figure 1 above shows the current incorrect version for item 21-1 (1).

Figure 2

Figure 2 shows the corrected MEL for item 21-1 (1) which now indicates “All”. This correction will be incorporated in the next revision of the Air Contractors A300 MEL. In the meantime, please refer to this engineering notice for dispatch release.

Prepared by: Authorised by:

Brendan Smyth Éamon Stapleton DoE Technical Services Manager

Item 21 - Air Conditioning