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47-43-01: 47-43-01 CESSNA: (Was Mandatory Note 12 of AD-768-5.) Applies to 120 and 140 Aircraft Serial Numbers Up to and Including 11842.
Compliance required prior to January 1, 1948.
Reroute the lower end of the primer line located on the left side of the firewall and rotate the strainer fitting so that it points downward and to the left at an angle of 60 degrees to the horizontal. Slip approximately 6 inches of vinylite tubing over the upper and lower ends of this primer line and install a shield around this line between the two pieces of vinylite tubing. This will preclude the possibility of fuel coming in contact with the left exhaust manifold in the event of a failure in this primer line.
(Cessna Service Letter No. 34, dated March 24, 1947, covers this same subject.)
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47-07-01: 47-07-01 BELLANCA: (Was Mandatory Note 4 of AD-773-5.) Applies to Models 14-13, 14-13-2 All Serial Numbers Up to and Including 1200.
To be accomplished not later than next periodic inspection.
Check fuel selector valve handle for proper indexing on valve by setting handle in L-ON and in R-ON position, by disconnecting the fuel line, and by blowing through line when there should be free passage of air. After tank positions have been set, the valve handle and shank should be permanently marked to identify the index position. Attach handle positively to shank by drilling through one side of the handle and halfway through the shank with a drill of number 53 size and inserting a pin of 1/16-inch diameter drill rod.
(Bellanca Service Bulletin No. 4 covers inspection of the valve handle installation.)
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62-10-01: 62-10-01 BELL: Amdt. 428 Part 507 Federal Register April 21, 1962. Applies to All Model 47J Helicopters Serial Numbers 1420 Through 1802 With Rollpins P/N 49-040-187-1750 or Clevis Pins P/N MS 20392-2-49 Installed In Elevator Spar, and With End Rib P/N 47-267- 453-1 (0.025-Inch Thick) Installed; Except Those Helicopters That Have Bell Kit No. 47-3746-1 or 47-3746-2 Installed.
Compliance required as indicated.
Numerous reports have been received of fatigue cracking of the tubular spar of both the right and left elevator at the Rollpin hole at B.L. 7.0, and fatigue cracking of the inboard rib of the elevators. To preclude failure of the elevator, the following shall be accomplished:
(a) Within 25 hours' time in service after the effective date of this AD:
(1) Remove the elevators from the tail boom in accordance with the Bell Maintenance Manual.
(2) Clean the area around the Rollpin hole and remove any zinc chromate putty from any plugged hole in the tubular spar at B.L. 7.0 for both right and left elevators.
(3) Inspect for cracks in the tubular spar of both elevators at the Rollpin hole at B.L. 7.0 using a 5-power or higher magnifying glass.
(4) Inspect the inboard rib for cracks using a 5-power or higher magnifying glass.
(b) If cracks are found in the tubular spar modify the elevator with Bell Helicopter Kit No. 47-3746-1 or 47-3746-2, "Improved Design Synchronized Elevator," or FAA engineering approved equivalent prior to further flight.
(c) If no cracks are found in the tubular spar, install clevis pin in accordance with subparagraphs (1) through (4) or Bell Service Letter No. 56 and reinspect in accordance with subparagraph (5).
(1) Position coupling assembly P/N 47-267-483-1 on elevators and line drill through Rollpin holes with a "D" (0.2460-inch diameter) drill. Remove sharp edges from holes. Install MS 20392-3-49 clevis pins, AN 960-4162 washers, and AN 381-3-6 cotter pins. A finger tight slip fitof the clevis pins is desired, approximately 0.0005 inch loose.
(2) Reinstall the elevator on the helicopter, shim as required to prevent preload or end play at bearings.
(3) Check clearance between skin and end of clevis pins. Trim skin, if necessary, to obtain clearance.
(4) Rerig elevator in accordance with the Bell Maintenance Manual.
(5) Reinspect in accordance with (a)(1) through (a)(3) within each succeeding 50 hours' time in service until Bell Helicopter Kit No. 47-3746-1 or 47-3746-2, "Improved Design Synchronized Elevator", or FAA approved equivalent is installed.
(d) If cracks are found in the inboard rib, repair the elevator as specified below, or modify with Bell Helicopter Kit No. 47-3746-1 or 47-3746-2, or FAA engineering approved equivalent prior to further flight.
(1) Remove the inboard rib by drilling out the rivets and remove the Bell P/N 47-267-404-7 shoulder from the rib by drilling out the rivets.
(2) Add a doubler of 0.032 thickness, or a new rib of 0.032 thickness, material aluminum alloy 2024-0, or a Bell rib P/N 47-267-453-7 (one required per elevator).
(3) Rivet Bell P/N 47-267-404-1 shoulder to the old rib and new doubler or the new rib. Use the rivet pattern in the shoulder with AN 470-AD3 or -4 rivets.
(4) Install the rib assembly, using the rivet pattern in the elevator skin with MS 20600 AD4 or -5 rivets.
(e) If no cracks are found in the inboard rib:
(1) Reinstall the elevator on the helicopter in accordance with Bell Maintenance Manual.
(2) Reinspect rib for cracks in accordance with (a)(4) within each succeeding 50 hours' time in service until Bell Helicopter Kit No. 47-3746-1 or 47-3746-2, "Improved Design Synchronized Elevator", or FAA engineering approved equivalent is installed.
(f) Upon request of the operator, an FAA maintenance inspector subject to prior approval of the Chief, Engineering and Manufacturing Branch, Southwest Region, may adjust the repetitive inspection intervals specified in this AD to permit compliance at an established inspection period of the operator if the request contains substantiating data to justify the increase for such operator.
(Bell Service Bulletin No. 135 SB dated July 27, 1961, covers this same subject. Bell's Service Letter No. 56 covers an acceptable fix for paragraphs (c)(1) through (c)(4) of this AD.)
This directive effective May 22, 1962.
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2008-19-11: We are adopting a new airworthiness directive (AD) for the products listed above. This AD results from mandatory continuing airworthiness information (MCAI) provided by the European Aviation Safety Agency (EASA) to identify and correct an unsafe condition on Turbomeca S.A. Arrius 2B1, 2B1A, 2B2, and 2K1 turboshaft engines. The MCAI describes the unsafe condition as:
A short circuit of some tantalum capacitors inside certain electronic control (EEC) units may lead to an in-flight shutdown on one of the two engines resulting from:
--Direct activation of the overspeed electronic protection;
--Non-direct activation of the electronic overspeed protection by lowering the threshold,
--Spurious activation of the starting sequence; or
--Loss of power control with no freeze of the fuel-metering valve.
We are issuing this AD to prevent in-flight engine shutdowns and possible forced autorotation landing or accident.
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91-15-25: 91-15-25 GENERAL ELECTRIC COMPANY (GE): Amendment 39-7090. Docket No. 90- ANE-34.
Applicability: GE CF6-80A series and CF6-80C2 series engines, installed on, but not limited to, Airbus A300 and A310 and Boeing 747 and 767 aircraft.
Compliance: Required as indicated, unless already accomplished.
To prevent an uncontained engine failure, accomplish the following:
(a) Inspect high pressure compressor rotor (HPCR) stages 11-14 spool-shafts for vane to spool rubs within 500 cycles in service (CIS) after the effective date of this AD, or prior to accumulating 8000 cycles since new, whichever occurs later, according to the following:
(1) Inspect CF6-80A series engines, Serial Numbers (S/N) 580-101 through 580-319, and S/N 585-101 through 585-222, installed with an HPCR stages 11-14 spool-shaft, Part Number (P/N) 9225M37G11, 9225M37G14, 9225M37G16, 9225M37G19, 9225M37G20, or 9225M37G21, in accordance with the Accomplishment Instructions in GE CF6-80A Service Bulletin (SB) 72-459, Revision 2, dated June 14, 1989.
(2) Inspect CF6-80C2 series engines, S/N 690-101 through 690-181, S/N 695-101 through 695-150, and S/N 705-101 through 705-112, installed with an HPCR stages 11- 14 spool-shafts, P/N 9380M30G07, 9380M30G08, 9380M30G09, 9380M30G10, or 1531M21G01, in accordance with the Accomplishment Instructions in GE CF6-80C2 SB 72-130, Revision 2, dated October 18, 1989.
(3) Inspect in accordance with the applicable requirements of paragraph (a)(1) or (a)(2) of this AD, HPCR stages 11-14 spool-shafts which were installed in a high pressure compressor (HPC) at an engine shop visit, with stages 10 through 13 vane radiiless than the values indicated in Table 1 of this AD for CF6-80A service engines, and Table 2 of this AD for CF6-80C2 series engines.
(b) The inspection requirements of paragraph (a) of this AD are not applicable to HPCR stages 11-14 spool-shafts visually inspected at the piece-part level, determined not to have vane to spool rub damage,and were installed in an HPC with stages 10 through 13 vane radii greater than or equal to the values indicated in Table 1 of this AD for CF6-80A series engines, and Table 2 of this AD for CF6-80C2 series engines.
(c) Remove from service within 500 CIS after the effective date of this AD or prior to accumulating 8,000 cycles since new, whichever occurs later, HPCR stages 11-14 spool-shafts with vane to spool rub damage.
NOTE: CF6-80A SB 72-460 and CF6-80C2 SB 72-131 introduce an FAA approved rework procedure to increase the FAA approved life limit for HPCR stages 11-14 spool-shafts with vane to spool rub damage.
(d) Aircraft may be ferried in accordance with the provisions of FAR 21.197 and 21.199 to a base where the AD can be accomplished.
(e) Upon submission of substantiating data by an owner or operator through an FAA Inspector (maintenance, avionics or operations, as appropriate), an alternate method of compliance with the requirements of this AD or adjustments to thecompliance times specified in this AD may be approved by the Manager, Engine Certification Office, Engine and Propeller Directorate, Aircraft Certification Service, FAA, 12 New England Executive Park, Burlington, Massachusetts 01803-5299.
(f) The inspections shall be done in accordance with the following GE documents:
DOCUMENT
PAGE NO.
ISSUE/REVISION
DATE
GE CF6-80A
2-9
Original
11/25/86
SB 72-459
1
Rev. 2
6/14/89
Total Pages:
9
GE CF6-80C2
2-9
Original
11/25/86
SB 72-130
1
Rev. 2
10/18/89
Total Pages: 9
This incorporation by reference was approved by the Director of the Federal Register in accordance with 5 U.S.C. 552(a) and 1 CFR Part 51. Copies may be obtained from General Electric Aircraft Engines, CF6 Distribution Clerk, Room 132, 111 Merchant Street, Cincinnati, Ohio 45246. Copies may be inspected at the FAA, New England Region, Office of the Assistant Chief Counsel, Room 311, 12 New England Executive Park, Burlington, Massachusetts, or at the Office of the Federal Register, 1100 L Street, NW, Room 8401, Washington, D.C.
TABLE 1
CF6-80A HPC Stator Vane Minimum Radii Vane
Stage
Position
Minimum Vane Radius (in)
Stage 10 Vanes
1-6, 35-80
11.416
7-8, 33, 34
11.417
9, 10, 31, 32
11.418
11, 12, 29, 30
11.419
13-16, 25-28
11.420
17-24 (or 1-80 if round grind)
11.421
Stage 11 Vanes
1-6, 35-80
11.623
7, 8, 33, 34
11.624
9, 10, 31, 32
11.625
11, 12, 29, 30
11.626
13, 16, 25-28
11.627
17-24 (or 1-80 11.628 if round grind)
Stage 12 Vanes with Liners
1-10, 31-80
11.769
11, 12, 29, 30
11.770
13-15, 26-28
11.771
16-25 (or 1-80 if round grind)
11.772
Stage 12 Vanes without Liners
1-10, 31-80
11.786
11, 12, 29, 30
11.787
13-15, 26-28
11.788
16-25 (or 1-80 if round grind)
11.789
Stage 13 Vanes
1-10, 31-80
11.903
11, 12, 29, 30
11.904
13-15, 26-28
11.905
16-25 (or 1-80if round grind)
11.906
NOTES: (1) Vane radius is measured from vane tip at vane centerline to stator case centerline.
(2) Vane positions are numbered clockwise, aft looking forward, starting with No. 1 at the left hand horizontal split line upper stator case.
(3) These revised minimum vane radii were incorporated into the CF6-80A Engine Manual, GEK 72501, in Revision 23.
TABLE 2
CF6-80C2 HPC Stator Vane Minimum Radii Vane
Stage
Position
Minimum Vane Radius (in)
Stage 10 Vanes
1-6, 35-80
11.409
7, 34
11.410
8, 9, 32, 33
11.411
10, 31
11.412
11, 12, 29, 30
11.413
13, 14, 27, 28
11.414
15, 16, 25, 26
11.415
17-24 (or 1-80 if round grind)
11.416
Stage 11 Vanes
1-6, 35-80
11.617
7, 34
11.618
8, 9, 32, 33
11.619
10, 31
11.620
11, 12, 29, 30
11.621
13, 14, 27, 28
11.622
15, 16, 25, 26
11.623
17-24 (or 1-80 if round grind)
11.624
Stage 12 Vanes
1-10,32-84
11.780
11, 31
11.781
12, 13, 29, 30
11.782
14-16, 26-28
11.783
17-25 (or 1-84 if round grind)
11.784
Stage 13 Vanes
1-10, 31-80
11.906
11, 12, 29, 30
11.907
13, 14, 27, 28
11.908
15, 16, 25, 26
11.909
17-24 (or 1-80 if round grind)
11.910
NOTES: (1) Vane radius is measured from vane tip at vane centerline to stator case centerline.
(2) Vane positions are numbered clockwise, aft looking forward, starting with No. 1 at the left hand horizontal split line upper stator case.
(3) These revised minimum vane radii were incorporated into the CF6-80C2 Engine Manual, GEK 92451, in Revision 7.
This amendment (39-7090, AD 91-15-25) becomes effective on September 6, 1991.
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62-08-03: 62-08-03 BEECH: Amdt. 421 Part 507 Federal Register April 17, 1962 as amended by Amendment 39-1019. Applies to Models 35-33, 35-A33 and 35-B33, Serial Numbers prior to CD-803, except CD-745 and CD-789; Models 35, A35, B35, C35, D35, E35, F35, G35, H35, J35, K35, M35, N35, and P35, Serial Numbers prior to D-6952; Model 50, Serial Numbers H-1 through H-11; Models B50 and C50, Serial Numbers CH-12 through CH-360; Models D50, D50A, D50B, D50C and D50E, Serial Numbers prior to DH-327, except DH-323 and DH-324; Model E50, Serial Numbers EH-1 through EH-70; Model F50, Serial Numbers FH-71 through FH-96 except FH-94; Model G50, Serial Numbers GH-94 and GH-97 through GH-119; Model H50, Serial Numbers HH-120 through HH-149; Model J50, Serial Numbers JH-150 through JH- 163; Models 95-55 and 95-A55, Serial Numbers prior to TC-276 except TC-235, TC-245, TC- 266, TC-273 and TC-274; Model 65, Serial Numbers prior to LC-141 except LC-125; Models 95, B95 and B95A, Serial Numbers prior to TD-499 airplanes.
Compliance required within 100 hours' time in service after the effective date of this amendment unless already accomplished.
As a result of cracks in and one failure of the white plastic rams horn control wheel, accomplish the following:
(a) Within the next 5 hours' time in service after the effective date of this AD and thereafter at each 100 hours' time in service or twelve calendar months, whichever occurs first, visually inspect the white plastic rams horn control wheels for cracks. Give particular attention to the area on the forward side of the hub and in the area of the attachment pin.
(b) If cracks are found, replace the control wheel prior to further flight.
(c) The inspections required by this AD may be discontinued when the proper replacement control wheel as specified in Beech Service Bulletin dated March 1970, titled "Control Wheel, Clock and Map Light, Inspection of Plastic Rams Horn Type Control Wheels" has been installed or an FAA approved equivalent is installed.
(Beech Service Bulletin dated March 23, 1962, titled "Inspection of Plastic Rams Horn Type Control Wheels" covers this same subject.)
Amendment 421 effective April 30, 1962.
Revised February 16, 1965.
This Amendment (39-1019) becomes effective July 9, 1970.
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95-11-05: This amendment adopts a new airworthiness directive (AD) that is applicable to Societe Nationale Industrielle Aerospatiale and Eurocopter France (Eurocopter France) Model AS-355 E, F, F1, F2, and N helicopters. This action requires a check to ensure that the main gearbox (MGB) oil pressure warning light illuminates during each shutdown of the helicopter engine until the MGB oil pressure switch (switch) is removed and replaced. This amendment is prompted by a malfunction of the MGB switch. This condition, if not corrected, could result in failure to detect a loss of oil pressure, loss of the MGB, loss of power to the main rotor system, and subsequent loss of control of the helicopter.
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61-25-03: 61-25-03 DOUGLAS: Amdt. 375 Part 507 Federal Register December 9, 1961. Applies to All DC-8 Aircraft, Serial Numbers 45252-45289, 45291-45306, 45376-45393, 45408-45413, 45416-45419, 45421-45431, 45433-45437, 45442-45445, 45526, 45565-45570, 45588-45606, 45609-45614, 45616-45622, 45624-45627, 45636. \n\n\tCompliance required as indicated. \n\n\tTo remove from service certain wing flap inboard actuating cylinders which, because of design characteristics, have failed in service and caused airplane operating difficulties, to remove adverse corrosion from any of the wing flap actuating cylinders, and to incorporate increased protection against corrosion for all flap actuating cylinders, the following shall be accomplished: \n\n\t(a)\tRemove all wing flap inboard actuating cylinder assemblies Douglas P/N 3715408-5001 or 3764264-5001 identified as having been manufactured by the Clary Corporation as follows: \n\n\t\t(1)\tCylinder assemblies with more than 5,000 hours' time in service shall be removed within 375 hours' time in service after the effective date of this AD. \n\n\t\t(2)\tCylinder assemblies with less that 4,625 hours' time in service shall be removed within 750 hours' time in service after the effective date of the AD. \n\n\t\t(3)\tCylinder assemblies with total time in service between 4,625 and 5,000 hours shall be removed prior to reaching a total time in service of 5,375 hours. \n\n\t(b)\tReplace each cylinder assembly removed per (a), with an assembly which has been inspected for freedom from cracks and for proper heat treatment, modified as necessary and reidentified, all in accordance with the instructions in Paragraph 2, section entitled Kit D and/or Kit C, of Douglas DC-8 Service Bulletin No. 27-118, Revision No. 1 to Reissue No. 1 or later, or with an FAA approved equivalent new part. \n\n\t(c)\tAll inboard cylinder assemblies (Clary or otherwise) which have been replaced in accordance with (b), shall be identified with a single band of green paint near the aft end of the barrel or with other suitable markings which will permit ready determination that an acceptable cylinder assembly is installed in the airplane. \n\n\t(d)\tWithin 900 hours' time in service following the effective date of this AD, replace all wing flap actuating cylinder assemblies (6 per airplane) with cylinder assemblies in which the barrels have been inspected for evidence of internal corrosion, cleaned and treated or replaced as necessary, and sealed, all in accordance with the instructions in Paragraph 2, section entitled Kits A, B, C, D, or E, of Douglas DC-8 Service Bulletin No. 27-118, Revision No. 1 to Reissue No. 1 or later, or with FAA approved equivalent new parts. Each of these modified or replacement cylinder assemblies shall be marked with two bands of green paint near the aft end of the barrel, or with other acceptable identifying markings. In those cases where a single band of green paint has been used to identify compliance of inboard cylinder assemblies with the provisions of paragraph (b) and such cylinder assemblies are determined to be eligible for reinstallation in accordance with this paragraph, only one additional band of green paint shall be added. \n\n\t(Douglas DC-8 Service Bulletin No. 27-118 covers this same subject.) \n\n\tThis directive effective December 9, 1961.
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2008-19-08: The FAA is superseding an existing airworthiness directive (AD), which applies to all Avions Marcel Dassault-Breguet Model Falcon 10 airplanes. That AD currently requires either revising the airplane flight manual and installing a placard in the flight deck to prohibit flight into known or forecasted icing conditions, or repetitively inspecting for delamination of the flexible hoses in the wing (slat) anti-icing system and performing corrective actions if necessary. The existing AD also requires replacement of the flexible hoses installed in the slat anti-icing systems, which ends the repetitive inspections. This new AD continues to require replacement of the flexible hoses installed in the slat anti-icing systems with new hoses, but at intervals defined in flight hours instead of flight cycles. This AD results from information we received from operators and the airplane manufacturer indicating that the repetitive interval for the required replacement deviated from the referenced service information. We are issuing this AD to prevent collapse of the flexible hoses in the slat anti-icing system, which could lead to insufficient anti-icing capability and, if icing is encountered in this situation, could result in reduced controllability of the airplane.
DATES: This AD becomes effective October 27, 2008.
On October 11, 2007 (72 FR 51161, September 6, 2007), the Director of the Federal Register approved the incorporation by reference of Dassault Service Bulletin F10-313, Revision 1, dated May 10, 2006.
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62-02-02: 62-02-02 BOEING AND DOUGLAS: Amdt. 385 Part 507 Federal Register January 16, 1962. Applies to All 707/720 Series and DC-8 Series Aircraft Equipped With Pratt and Whitney JT3D-3 Turbofan Engines. \n\n\tCompliance required as indicated. \n\n\t(a) For engines previously inspected by the procedure described in paragraph (c), reinspect in accordance with paragraph (c) every 225 hours' time in service thereafter. \n\n\t(b) For engines not previously inspected by the procedure described in paragraph (c) inspect in accordance with paragraph (c) as follows: \n\n\t\t(1) Inspect engines with 200 or more hours' time in service within the next 25 hours' time in service and every 225 hours time in service thereafter. \n\n\t\t(2) Inspect engines with less than 200 hours' time in service by the time 225 hours' time in service have been accumulated and every 225 hours' time in service thereafter. \n\n\t(c) Due to recent failures of P/N 393504 fourth stage compressor rotor disc, inspect the disc for cracks adjacent to the inside edge of the tie bolt circular rib in the disc web. Such cracks may progress along the inside of the rib and then toward the disc bore through the bore stiffening section. \n\n\tTo accomplish the inspection, remove the front accessory drive support assembly (N1 gearcase) and the front accessory drive main spur gear (N1 gearcase coupling). Using a strong light and borescope or similar optical device, visually inspect the fourth stage compressor rotor disc in the area noted above. If any cracking is found, the engine must be removed for disc replacement prior to further flight. \n\n\t(d) The requirement for main oil screen inspection per AD 61-24-01 does not apply when the No. 1 bearing compartment is exposed for this disc inspection. \n\n\t(Pratt and Whitney Aircraft telegraphic message of December 19, 1961, covers the same subject.) \n\n\tThis directive effective upon publication in the Federal Register for all persons except those to whom it was made effective immediately by telegram dated December 22, 1961. \n\n\tRevised October 12, 1962.
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