Results
49-28-01: 49-28-01 NAVION: Applies to All Airplanes Equipped With Product Techniques, Inc., Propeller Spinners. To be accomplished as specified below: In order to preclude the possibility of continued use of any unsatisfactory spinners which may be in existence, the following steps must be taken: 1. Spinners not previously installed on an aircraft must be inspected before installation for the thickness of the bulkhead. This may be accomplished by measuring the thickness of the bulkhead at the perimeter; measurements should be taken about 1/2 inch from the edge for best accuracy. The manufacturing process used in fabricating the bulkheads reduces the gage of the material about 0.012 inch at the perimeter. Bulkheads measuring less than 0.045 inch at this point must be replaced with heavier gage steel bulkheads. 2. Spinners which have partially completed the inspections previously required by AD 49-01-01 may continue in service until the next 10-hour inspection at which time thebulkhead should be gaged as specified above and replace if under 0.045-inch thickness. 3. Spinners having satisfactorily passed the five inspection periods required by AD 49-01-01 must be inspected for bulkhead thickness, as specified above, within the next 50 hours of flight, and replaced if under 0.045-inch thickness. It has been determined that bulkheads of 0.058-inch and 0.064-inch cadmium plated steel are satisfactory and do not require the periodic inspections previously specified in AD 49-01-01. However, bulkheads made of thinner gage steel, or of aluminum, are unsatisfactory. This supersedes AD 49-01-01.
46-40-02: 46-40-02 FAIRCHILD: (Was Mandatory Note 2 of AD-667-2; 3 of AD-633-2; 3 of AD- 535-2; 3 of AD-564-2; 3 of AD-600-2.) Applicable to Models 22 and 24 Series Aircraft Serial Numbers 3300 to 3319, Inclusive, and 3350 to 3358, Inclusive, Except Seaplanes. Compliance required at next 100-hour inspection. Unless previously accomplished, reinforce the oleo tubes at the piston end by installing sleeve. Fairchild part SK-1636. Sleeve may be attached by using two AN 435-5-4 rivets or by welding through four 5/16-inch holes drilled in the piston tube. All damaged parts should be replaced. Fairchild Service Bulletin 40-5, revised July 3, 1945, and 44-6-C dated May 8, 1944, cover this same subject.)
2008-16-18: We are adopting a new airworthiness directive (AD) for the products listed above. This AD results from mandatory continuing airworthiness information (MCAI) issued by an aviation authority of another country to identify and correct an unsafe condition on an aviation product. The MCAI describes the unsafe condition as: Recently an RB211 HP turbine disc has been found with a crack which had propagated further than expected from the risk model that was used to establish the original inspection. We are issuing this AD to detect cracks that could cause the high pressure (HP) turbine disc to fail and result in uncontained failure of the engine.
71-19-01: 71-19-01 SCHLEICHER: Amendment 39-1283. Applies to Model K8 Gliders. Compliance is required within the next 25 hours' time in service after the effective date of this AD unless already accomplished. To prevent failure of the airbrake bellcrank inboard bearing bracket to main fuselage attachment welds, modify the bracket attachment in accordance with Schleicher Modification No. 10, dated October 25, 1962 or an FAA-approved equivalent. This amendment becomes effective September 7, 1971.
46-30-01: 46-30-01 LUSCOMBE: (Was Mandatory Note 1 of AD-694-4.) Applies Only to Model 8A Serial Numbers 2201 to 2614 Inclusive; 2616 to 2632 Inclusive; 2635, 2637, 2639, 2642, and 2645. Compliance required prior to completion of next 10 hours of operation. Replace the adjustment screw now installed in the lower end of the control stick horn, which is located beneath the floor boards with an AN 520-10-44 screw in order to prevent interference with the lower fuselage skin on the forward flange of the landing gear bulkhead. (Luscombe Service Bulletin No. 2-46 covers this same subject.)
49-47-02: 49-47-02 BELL: Applies to All Models 47B, 47B3, 47D, and Model 47D1 Serial Numbers 145 Through 164. Compliance as indicated. As a result of recent failures occurring to the tail rotor drive system the following inspections and replacements must be accomplished: 1. Not later than the next 25 hours of operation and at each 50-hour period thereafter the following inspection procedure must be accomplished until the tail rotor drive shaft parts as described below are incorporated: (a) Remove and disassemble tail boom extension for complete inspection. (b) Visually inspect bearings for wear, cracks, chips, and brinelling. (c) Inspect surfaces of shaft for dents, cuts, and signs of fatigue. (d) Conduct a magnaflux inspection of the tail boom extension drive shaft, examining particularly the shaft surface for approximately 3 inches of length from each end of the shaft. (e) Magnaflux 47-644-C14 sleeve inspecting for cracks at thread roots closestto flange. 2. To compensate for shaft misalignment and flight distortion of the shaft assembly, the following modifications shall be accomplished not later than February 15, 1950: (a) Installation of a redesigned extension drive shaft P/N 47-644-126 and 47- 644-181. (b) Installation of a spline coupling in tail rotor drive shaft forward of the universal joint P/N 47-644-177 and 47-644-130. (Bell Service Bulletin No. 69, dated November 11, 1949, covers this subject.)
2008-17-17: We are adopting a new airworthiness directive (AD) for all Boeing Model 747 airplanes. This AD requires a one-time inspection of certain fuselage skins at section 41 to find any external doublers that cover the inspection areas and to identify the external doublers that end on a stringer and those that do not, and related investigative and corrective actions if necessary. This AD results from reports of cracks found at fastener locations in the fuselage skins at section 41. We are issuing this AD to detect and correct fuselage skin cracks at fastener locations along the skin-to-stringer attachments, which could join together and become large and consequently result in rapid decompression of the cabin.
46-38-01: 46-38-01 GRUMMAN: (Was Mandatory Note 4 of AD-654-1.) Applies to G-21 and G- 21A Aircraft. Compliance required prior to November 1, 1946. Reroute to vacuum system tubing in the engine compartments to prevent raw fuel from lodging in the check valve, thus deteriorating the valve and possible seeping through. The flap system storage tank should be inspected and if fuel is found it should be thoroughly flushed. The check valves in the vacuum system are to be removed and inspected for wear and deterioration. Presence of fuel in the flap system storage tank is sufficient cause for the replacement of the check valves. (Grumman Service Bulletin No. 19 dated July 18, 1946, covers this same subject.)
48-06-05: 48-06-05 DOUGLAS: Applies to DC-4 and C-54 Aircraft. \n\n\tCompliance required by November 1, 1948. \n\n\tTo prevent failure of the aileron hinge eyebolts and the rudder trim tab and elevator trim tab hinges and control horns, the following changes must be made: \n\n\t1.\tOn airplanes having aileron assemblies P/N 5166075, -1, -500, -501, or -503, replace No. 3 hinge eyebolt P/N 1166013 with new eyebolt P/N 1357234. Airplanes having aileron assemblies P/N 5078609, -1, -500, -501, -502, or -503, do not require this hinge replacement. \n\n\t2.\tOn elevator and rudder trim tabs replace all hinges and control horns with new parts made of steel. \n\n\t3.\tReplace the standard AN bolts in elevator trim tab and rudder trim tab hinges and control horns with new special close tolerance high strength bolts using new AN 310 type nuts, AN 960 washers and AN 380 cotter pins. \n\n\tInspection should be accomplished in accordance with Note 51-9-2 except in the case of aircraft in which the standard AN rudder and elevator hinge bolts have not been replaced with new close tolerance high strength bolts using new AN 310 type nuts, AN 960 washers and AN 380 cotter pins. In such cases, inspection should be accomplished at intervals not to exceed 1,000 hours. \n\n\t(Douglas Service Bulletin DC-4 No. 83 covers the above rework.)
49-33-01: 49-33-01 DOUGLAS: Applies to All Model DC-6 Airplanes Equipped With Hamilton Standard Model 43D60/6851A-O Props. \n\n\tCompliance required not later than September 10, 1949. \n\n\tAs a precautionary measure against fatigue failure, Hamilton Standard Model 43D60/6851A-O propeller blades may not be operated in excess of 3,500 hours unless the taper-bores have been remachined and reshotpeened by the manufacturer. In addition, any blades of this model which have sustained damage sufficient to change their face or edge alignment shall be withdrawn from service pending final results of manufacturer's test program.
60-03-08: 60-03-08 PIPER: Amdt. 92 Part 507 Federal Register January 26, 1960. Applies to PA- 23 and PA-23 "160" Aircraft, Serial Numbers 23-1 and Up. Compliance required as indicated. As a result of several instances of fuel system cable malfunctioning and fuel valve leakage, the following inspections shall be accomplished: (a) Ascertain that fuel is contained in all fuel cells. With the master switch on, energize the electric fuel pumps. While the fuel selector and crossfeed valve levers are cycled through their positions several times, observe through appropriate access openings the main valve (P/N 18598), auxiliary valve (P/N 17920) if installed, and the crossfeed valve (P/N 492044) in the cabin control pedestal. If leakage is observed, effect repairs. (Refer also to Piper Service Manual paragraphs 9-17 and 9-18.) (b) Check rigging and clamping of control cable and rigging of valve control linkage to eliminate bowing and to insure proper valve operation.Compliance with item (a) required by March 15, 1960, and every 50 hours of time in service thereafter on aircraft Serial Numbers 23-1 through 23-1695 until improved valves (P/N 492050 main and crossfeed, and 492051 auxiliary) are installed. After the improved valves are installed, the inspection period may be increased to every 100 hours of time in service. Aircraft Serial Numbers 23-1696 and up incorporated the new valves; therefore, these aircraft having less than 100 hours of time in service on March 15, 1960, may be operated until they have 100 hours, at which time the initial inspection shall be conducted, and then shall be inspected every 100 hours thereafter. Compliance with item (b) required by March 15, 1960, and every 50 hours of time in service thereafter on aircraft Serial Numbers 23-1 through 23-289 until a more rigid control cable (P/N 18815 for main fuel system and P/N 18816 for auxiliary fuel system and idler bellcrank P/N 18782) is installed. After the new cableis installed, the inspection period may be increased to every 100 hours of time in service Aircraft Serial Numbers 23-290 and up incorporated the new cables; therefore, these aircraft having less than 100 hours of time in service on March 15, 1960, may be operated until they have 100 hours, at which time the initial inspection shall be conducted, and then shall be inspected every 100 hours thereafter. (c) Not later than 25 hours' time in service after October 19, 1960, aircraft Serial Numbers 23-1 through 23-1695 which have improved valves (P/N 492050 main and crossfeed, and P/N 492051 auxiliary) installed must be inspected to ascertain that P/N 492051 was used for auxiliary and main shutoff valves on aircraft with auxiliary fuel systems, and P/N 492050 was used for main shutoff valves on aircraft without auxiliary tanks and for all crossfeed valves in accordance with Piper Service Letter 322A and Piper Service Manual paragraphs 9-17 and 9-18. If incorrect combinations are found, valves must be installed properly prior to further flight. (The above fuel valves are also identified by Scott P/N 4240-2 for P/N 492050 valve and P/N 4240-3 for P/N 492051 valve. (Piper Service Letters Nos. 322A, and 286 cover the same subject.) Revised October 19, 1960. Revised March 21, 1961.
61-08-02: 61-08-02 RUPERT: Amendment 275 Part 507 Federal Register April 15, 1961 as amended by Amendment 39-2751. Applies to All Aircraft Equipped With Rupert Models 50, N50, 65, 80, and S-2194 Safety Belts, except those that contain Cam Clasps marked Cummings & Sander, Division of American Safety Equipment Corporation, P/N M 4090. \n\n\tCompliance required within the next 75 hours of time in service after the effective date of this AD. \n\n\tRecurring instances have been reported wherein Rupert belt assemblies have slipped under low tension loads. Accordingly, it has been determined by test the Rupert safety belt assemblies manufactured under Technical Standard Order C22 standards are deficient and must either be replaced with belt assemblies that conform to TSO-C22 standards or be reworked by accomplishing the following: \n\n\t(a) Remove all of the three bar slide length adjustors and replace them with new Rupert R-60 adjustors. Replacement adjustors will have thicknesses greater than 0.120. \n\n\t(b) Measure the distance between the crests of the outermost serrations on the tongue of the buckle. (See Figure 1.) If this distance measures 5/16 inch or less the buckle shall be reworked by replacing the tongue with one in which this distance is at least 3/8 inch. \n\n\t(Rupert Service Bulletin No. 101 covers this same subject.) \n\n\tAmendment 275 Part 507 Federal Register April 15, 1961 became effective May 16, 1961. \n\n\tThis Amendment 39-2751 becomes effective October 26, 1976.
61-06-04: 61-06-04 DOUGLAS: Amdt. 267 Part 507 Federal Register March 18, 1961. Applies to All DC-8 Series Aircraft With 1,200 or More Hours' Time in Service on Clevis, P/N 2619862, In Rudder Hydraulic System. \n\n\tCompliance required as indicated. \n\n\tAs a result of several reported instances of failed clevis, P/N 2619862, on the rudder lockout cylinder, the following must be accomplished. \n\n\t(1)\tUnless already accomplished within the last 20 hours' time in service, each clevis, P/N 2619862, in service which has not been inspected per (2) must be visually inspected for cracks prior to the next flight. The bolt attaching the clevis to the link to the gripper arm need not be removed for this inspection. "Cracked clevis must be replaced with a new clevis, P/N 2619862 or P/N 2772031 (Kit A of Douglas Service Bulletin 27-100) or FAA approved equivalent, prior to further flight." \n\n\t(2)\tUnless already accomplished within the last 140 hours' time in service, each clevis, P/N 2619862, whichremains in service following the inspection prescribed in (1) must be inspected with dye penetrant or magnetic particle method or equivalent within the next 20 hours' time in service. The bolt attaching the clevis to the link to the gripper arm must be removed for this inspection. "Cracked clevis must be replaced with a new clevis, P/N 2619862 or P/N 2772031 (Kit A of Douglas Service Bulletin 27-100) or FAA approved equivalent, prior to further flight." Clevis, P/N 2619862, retained in service must be reinspected at intervals not to exceed 160 hours' time in service. After clevis, P/N 2619862, is replaced with a redesigned clevis, P/N 2772031, inspections of replaced parts may then be made at normal inspection periods. \n\n\t(3)\tUnless already accomplished, a rudder creep rate check and necessary adjustment of the valve rod must be accomplished within the next 100 hours' time in service. "The adjustment and creep rate check is to be accomplished in accordance with Supplement No. 1dated February 27, 1961, to Douglas Alert Service Bulletin A27-100. Installation of new connecting rod assembly P/N 4772143-1 is optional (Kit B or Kit C of Douglas Service Bulletin 27-100). \n\n\t"(4)\tWithin the next 20 hours' time in service, unless already accomplished, install a placard adjacent to the rudder reversion light to read: 'WHEN LIGHT COMES ON SHUT OFF RUDDER HYDRAULIC POWER IMMEDIATELY,' or install warning placard in accordance with Kit D or Kit E of Douglas Service Bulletin 27-100. \n\n\t"(Supplement No. 1 dated February 27, 1961, to Douglas Alert Service Bulletin A27-100 and Service Bulletin 27-100 dated May 12, 1961, covers this subject.)" \n\n\tThis directive becomes effective upon publication in the Federal Register for all persons except those to whom it was made effective immediately by individual telegrams dated March 3, 1961. (Material enclosed by quotation marks effective August 4, 1961.)
94-24-07: This amendment adopts a new airworthiness directive (AD), applicable to certain Airbus Model A320 series airplanes, that requires modification of the brake steering control unit (BSCU). This amendment is prompted by reports that the BSCU on these airplanes allowed a 90-degree rotation of the nose gear after landing, which resulted in significant damage to the wheels. The actions specified by this AD are intended to prevent failure of the nose gear tires and wheels and the loss of directional control of the airplane while it is on the ground.
95-14-03: This amendment adopts a new airworthiness directive (AD), applicable to all British Aerospace Model BAC 1-11-200 and -400 series airplanes, that requires repetitive radiographic inspections to detect corrosion of the center torque shaft of the wing spoiler, and replacement, if necessary. This amendment is prompted by a report of the wing spoiler failing to retract fully after deployment, which caused the wing to drop significantly. Subsequent investigation revealed that the torque shaft assembly of the wing spoiler had failed due to severe corrosion. The actions specified by this AD are intended to prevent such failures, which can result in an adverse effect on controllability of the airplane.
2008-16-17: We are adopting a new airworthiness directive (AD) for the PZL Swidnik S. A. (PZL) Model W-3A helicopters. This AD results from mandatory continuing airworthiness information (MCAI) issued by the European Aviation Safety Agency (EASA), which is the Technical Agent for the Member States of the European Community. The MCAI states: "In PZL W-3A helicopter S/N 37.07.05, and previously also in the PZL W-3AS model helicopters, leakage was found in the pipe 37.59.006.00.00 installed in the pressure line of hydraulic system 2, in the part between the hydraulic block and the ground hydraulic unit panel. The hydraulic system in the part between hydraulic blocks and the ground hydraulic unit panel is used only during periodical inspections, for the performance of which it is required to use the hydraulic power unit. This condition, if not corrected, could result in a fire hazard.'' The actions specified in this AD are intended to prevent this unsafe condition.
61-04-02: 61-04-02 DOUGLAS: Amdt. 250 Part 507 Federal Register February 10, 1961. Applies to All DC-6, DC-6A and DC-6B Aircraft; Fuselage Number 1 Up to and Including Fuselage Number 722, Having in Excess of 9,000 Hours' Time in Service. \n\n\tCompliance required as indicated. \n\n\tThere have been numerous cases reported of spar cap cracking on DC-6 Series aircraft. Cracking usually occurs in spar cap tangs in the area of the Station 60 attachments and progresses chordwise. In addition, service experience has shown that the temporary repair of the above service difficulties per Douglas Rework Drawing 5611387 does not have the service life originally anticipated. As a result of this service experience, the upper and lower, front and center spar caps in the area of wing Station 60, with special attention to the spar cap tangs between wing Stations 55 and 65, must be inspected for cracks as follows: \n\n\t(a)\tThe upper and lower, front and center spar caps must be inspected within the next 450 hours' time in service unless already accomplished. Aircraft inspected prior to issuance of this AD must also comply with the repetitive inspections rework and/or repairs specified in (b), (c), (d), and (e). \n\n\t(b)\tThe upper front and center spar caps on all DC-6, DC-6A, and DC-6B aircraft, Fuselage Nos. 1 through 722, must be reinspected at intervals not to exceed 1,600 hours' time in service. \n\n\t(c)\tThe lower front and center spar caps must be reinspected as follows: \n\n\t\t(1)\tModel DC-6 aircraft, Fuselage Nos. 1 through 172, which have not been reworked in accordance with DC-6 Service Bulletin Nos. 569 and 724, at intervals not to exceed 1,600 hours' time in service. \n\n\t\t(2)\tModel DC-6 aircraft, Fuselage Nos. 1 through 172, which have been reworked in accordance with DC-6 Service Bulletin Nos. 569 and 724, at intervals not to exceed 3,250 hours' time in service. \n\n\t\t(3)\tModel DC-6, DC-6A and DC-6B aircraft, Fuselage Nos. 174 through 722, at intervals not to exceed 3,250hours' time in service. \n\n\t(d)\tIf cracks are found, FAA approved permanent rework or temporary repair as recommended by the manufacturer or FAA approved equivalent is required prior to further flight except ferry flight in accordance with the provisions of CAR 1.76. Temporary repairs may be made per Douglas Rework Drawing 5611387, or FAA approved equivalent, providing crack limitations as established on this drawing have not been exceeded. \n\n\t(e)\tAircraft incorporating a temporary repair must be reinspected at intervals not to exceed 750 hours' time in service pending the accomplishment of the FAA approved manufacturer's recommended permanent rework or FAA approved equivalent. Such rework or equivalent must be accomplished within 4,200 hours' time in service after incorporating the temporary repair. \n\n\t(f)\tThe inspections required by this AD may be discontinued for any area reworked in accordance with FAA approved permanent repair instructions. \n\n\t(Douglas Alert Service Bulletin A-678 revised June 3, 1960, covers this subject.) \n\n\tThis supersedes AD 60-15-01. \n\n\tThis directive effective March 14, 1961.
93-20-05: 93-20-05 AYRES CORPORATION: Amendment 39-8713; Docket No. 93-CE-30-AD. Applicability: The following model and serial number airplanes, certificated in any category: Models Serial Numbers S2R 5000 through 5099, 1380R, and 1416R through 2583R S2R-R1340 R1340-001 through R1340-030 (with or without DC suffix) S2R-R3S R3S-001 through R3S-011 (with or without DC suffix) S2R-R1820 R1820-001 through R1820-035 (with or without DC suffix) S2R-T11 T11-001 through T11-005 (with or without DC suffix) S2R-T15 T15-001 through T15-029 (with or without DC suffix); and T27-001 through T27-029 and T-27-031 (with or without DC suffix) S2R-T34 6000 through 6049, T34-001 through T34-180, T34-190, T34-191 and T34-192 (with or without DC suffix); T36-001 through T36-180 (with or without DC suffix); and T41-001 through T41-180 (with or without DC suffix) S2R-T45 T45-001 through T45-003 (with or without DC suffix) S2R-T65 T65-001 (with or without DC suffix) S2R-HG-T65 T65-002 through T65-010 (with or without DC suffix) S2RG6 G6-101 through G6-112 S2R-G10 G10-101 Compliance: Required as indicated, unless already accomplished. NOTE 1: The compliance times specified in this AD take precedence over those referenced in Ayres Service Bulletin (SB) No. SB-AG-32, dated February 12, 1993. To prevent structural damage to the vertical tail caused by a damaged vertical tail attachment bracket, which could result in loss of control of the airplane, accomplish the following: (a) Within the next 50 hours time-in-service after the effective date of this AD, inspect the bracket that attaches the vertical tail front spar to the horizontal stabilizer for damage (cracks, broken lugs or bolts, or elongated holes) in accordance with the ACCOMPLISHMENT INSTRUCTIONS: I. Inspection, section of Ayres SB No. SB-AG-32, dated February 12, 1993. (b) If any damage is found to the bracket during the inspection specified in paragraph (a) of this AD, prior to further flight, replace the bracket with an aluminum bracket, part number (P/N) 40301T007, and install a new close out plate, P/N 40309T003, in accordance with the ACCOMPLISHMENT INSTRUCTIONS: II. Repair, section of Ayres SB No. SB-AG-32, dated February 12, 1993. (c) Within the next 100 hours TIS after the effective date of this AD, unless already accomplished in accordance with paragraph (b) of this AD, replace the bracket with an aluminum bracket, part number (P/N) 40301T007, and install a new close out plate, P/N 40309T003, in accordance with the ACCOMPLISHMENT INSTRUCTIONS: II. Repair, section of Ayres SB No. SB-AG-32, dated February 12, 1993. (d) The replacement required by paragraph (c) of this AD may be accomplished instead of the inspection specified in paragraph (a) of this AD provided it is accomplished at or prior to the 50-hour TIS compliance time. (e) Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate the airplane to a location where the requirements of this AD can be accomplished. (f) An alternative method of compliance or adjustment of the initial or repetitive compliance times that provides an equivalent level of safety may be approved by the Manager, Atlanta Aircraft Certification Office, 1669 Phoenix Parkway, Suite 210C, Atlanta, Georgia 30349. The request shall be forwarded through an appropriate FAA Maintenance Inspector, who may add comments and then send it to the Manager, Atlanta Aircraft Certification Office. NOTE 2: Information concerning the existence of approved alternative methods of compliance with this AD, if any, may be obtained from the Atlanta Aircraft Certification Office. (g) The inspection and replacement required by this AD shall be done in accordance with Ayres Service Bulletin No. SB-AG-32, dated February 12, 1993. 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 the Ayres Corporation, P.O. Box 3090, Albany, Georgia 31708. Copies may be inspected at the FAA, Central Region, Office of the Assistant Chief Counsel, Room 1558, 601 E. 12th Street, Kansas City, Missouri, or at the Office of the Federal Register, 800 North Capitol Street, NW., suite 700, Washington, DC. (h) This amendment (39-8713) becomes effective on December 3, 1993.
60-11-09: 60-11-09 SUD AVIATION: Amdt. 154 Part 507 Federal Register May 18, 1960. Applies to All Alouette II SE 3130 Helicopters Equipped With Tail Rotor Blade Model Numbers 34.40.000 and 34.60.000. Compliance required each five hours of time in service. (a) Visually inspect the upper and lower blade surfaces to determine that the blade cuff at the attachment bolts and the skin around the entire reinforcement plate area are free from cracks. (b) Check the end of the reinforcement plate for bonding separation by exerting light thumb pressure on the blade immediately outboard of the plate. (c) If evidence of cracking or bonding separation is found blades must be replaced prior to further flight. (d) All blade numbers 34.40.000 and 34.60.000 must be retired at 2,500 hours of service time. (Sud Alouette Helicopter Service Bulletin No. 34.11.138B covers the same subject in Part D.) Revised February 28, 1962.
2008-16-11: We are adopting a new airworthiness directive (AD) for certain McDonnell Douglas Model DC-8-61, DC-8-61F, DC-8-63, DC-8-63F, DC-8-71F, and DC-8-73F airplanes. For certain airplanes, this AD requires non- destructive testing (NDT) to detect cracks of the door jamb corners of the forward and aft service doors, and doing applicable related investigative and corrective actions. For certain other airplanes, this AD requires inspecting and repairing if necessary or replacing previously repaired door jamb corners with an applicable repair. This AD results from reports of numerous cases of cracks in the skin at the door jamb corners of the forward and aft service doors. We are issuing this AD to detect and correct fatigue cracking of door jamb corners of the forward and aft service doors, which could adversely affect the structural integrity of the airplane.
60-08-03: 60-08-03 DOUGLAS: Amdt. 133 Part 507 Federal Register April 14, 1960. Applies to DC-8 Series Aircraft Serial Numbers 45281 to 45290 Inclusive, 45408 to 45413 Inclusive, 45422, 45423, 45588 to 45594 Inclusive. \n\n\tCompliance required as indicated. \n\n\t(a) Within the next 50 hours' time in service, unless already accomplished, visually inspect the upper and lower wing rib caps, P/N 5615316-1 (left hand) and -2 (right hand) and P/N 5615317-1 (left hand) and -2 (right hand) respectively, at wing station XRS 139.0 for any evidence of cracks. Use at least a 10-power magnifying glass or equivalent. If any doubt exists, utilize dye penetrant or other inspection methods for verification. Aircraft with cracks are not to be returned to service until the damaged parts are repaired in accordance with (c) or (f) or replaced in accordance with (d). \n\n\t(b) Parts which show no evidence of cracks shall be reinspected in accordance with (a) at periods not to exceed 200 hours' time in service until the provisions of (e) are accomplished. \n\n\t(c) Parts which are found to be cracked and which are not replaced per (d) nor permanently repaired per (f) are authorized a reinspection period not to exceed 2,500 hours' time in service, provided: \n\n\t\t(1) The crack in the vertical tang does not exceed a length of ten inches and does not terminate closer than 1/16 inch from the heavy section of the part. \n\n\t\t(2) A stop hole 1/4- to 1/2-inch diameter is drilled in the extreme end of the crack, or the attachment hole in which the crack terminates is enlarged to 1/2-inch diameter. \n\n\t\t(3) A minimum of twenty bulkhead web to cap lock bolts immediately forward of the rear spar are removed, the gap between the web and cap vertical tang is accurately measured, 7075-T6 shims, tapered as necessary in both directions, are installed to fill the gap, and the original type of bulkhead web to cap attachment are reinstalled. \n\n\t\t(4) No fuel or pay load is in the airplane during subsequent jacking operations in which the jack point at the bulkhead in question is utilized. \n\n\t(d) Parts found to be cracked beyond the limits of (c)(1) must be replaced prior to further flight. The replacement parts are subject to the 200 hour time in service inspection limitation of (b) unless, during installation, the gap measurement and shimming provisions of (c)(3) are accomplished. When properly shimmed during installation, the new parts will not be subject to any further special inspections or subsequent airplane jacking weight limitation. \n\n\t(e) Parts inspected and found to have no cracks will no longer be subject to special inspections or airplane jacking weight limitation after gap measurement, shimming and reattachment provisions of (c)(3) have been accomplished. \n\n\t(f) As an alternative to the repair specified in paragraph (c), cracked parts may be repaired per the FAA approved permanent repair recommended by the manufacturer providing crack limitations specified in(c)(1) have not been exceeded and cracks are processed per (c)(2). The special inspections and jacking procedures required by this AD may be discontinued for any part repaired in accordance with this paragraph. \n\n\t(Douglas DC-8 Service Bulletin 57-7 revised June 30 and July 25, 1960, covers this subject.) \n\n\tThis supersedes AD 60-07-04. \n\n\tRevised October 29, 1960.
2008-16-04: We are adopting a new airworthiness directive (AD) for BHTC Model 222, 222B, and 222U helicopters. This AD results from mandatory continuing airworthiness information (MCAI) issued by an aviation authority of another country to identify and correct an unsafe condition on an aviation product. The aviation authority of Canada, with which we have a bilateral agreement, states in the MCAI: "It has been determined that the existing rigging procedures for the tail rotor pitch change mechanism have to be changed due to possibility of parts interference.'' The cumulative effect of individual part tolerances resulting in the total assemblage of those parts being out of tolerance could result in the tail rotor yoke striking another part other than the flapping stop (parts interference) cited in the MCAI. Also, the misalignment of the tail rotor counterweight bellcrank may result in higher tail rotor pedal forces and a higher pilot workload after failure of the No. 1 hydraulic system. Both parts interference and the misaligned counterweight bellcrank create an unsafe condition. We are issuing this AD to require actions that are intended to address these unsafe conditions.
2008-02-04: We are adopting a new airworthiness directive (AD) that supersedes AD 2007-13-11, which applies to all Eclipse Aviation Corporation (Eclipse) Model EA500 airplanes. AD 2007-13-11 was prompted by reports of loss of primary airspeed indication due to freezing condensation within the pitot system. AD 2007-13-11 requires operational limitations consisting of operation only in day visual flight rules (VFR), allowing only a VFR flight plan, and maintaining operation with two pilots. Since we issued AD 2007-13-11, Eclipse developed a design modification to the pitot/angle-of-attack (AOA) system to eliminate the possibility of freezing condensation within the pitot/AOA system. Eclipse is incorporating this modification during production on Model EA500 airplanes starting with serial number (S/N) 000065. Consequently, this AD limits the applicability to airplanes under S/N 000065 and requires incorporating the modification. This AD also retains the operating limitations in AD 2007-13-11until the modification is incorporated. We are issuing this AD to prevent long- term reliance on special operating limitations when a design change exists that will eliminate the need for the operating limitations. Incorporating the modification will prevent loss of air pressure in the pitot system, which could cause erroneous AOA and airspeed information with consequent loss of control.
60-03-07: 60-03-07 PIPER: Amdt. 88 Part 507 Federal Register January 26, 1960. Applies to PA- 23 Aircraft Serial Numbers 23-747 To 23-1534 Inclusive. Compliance required not later than March 1, 1960. To preclude recurrence of short circuiting of the control circuits of the engine starter solenoids, electric cables Nos. P2B, P20A and P20B in the starter solenoid circuits shall be increased in size to 16 gage and a 15 ampere trip-free circuit breaker installed in the circuit at the point where lead P2B connects into the electric system. An appropriate placard must be installed adjacent to the new circuit breaker to provide adequate identification. (Piper Service Bulletin No. 175 dated June 9, 1959, covers this subject.)
57-22-01: 57-22-01 PIPER: Applies to All Models PA-16, PA-20 and PA-22 Aircraft. Compliance required as indicated. To preclude the possibility of inflight fires the following inspection and rework is necessary to eliminate combustible material and possible ignition sources from the area aft of the firewall, underneath the forward cabin floor. Access to this section may be gained by removing the metal panels or opening the fuselage side cowl panels rearward of the firewall underneath the aircraft as shown in Piper Service Bulletin No. 161a. The relative difficulty in gaining access to this area has probably contributed to poor maintenance. 1. On all PA-16, PA-20, and PA-22 aircraft, Serial Numbers 22-1 through 22-2699, the following inspection and rework is necessary prior to December 15, 1957. Remove and discard any sound-proofing material contaminated with engine or hydraulic oil. Where the plastic septum has separated from the fiberglass or shows signs of drying or cracking it should be removed in its entirety from the affected blanket. Uncontaminated fiberglass, from which the plastic septum has been removed, may be continued in service. Inspect electrical wiring for chafing of the insulation and replace any found in an unsatisfactory condition. Check for a reasonable clearance between hydraulic lines, electrical wires, control cables and fuel lines and rework as necessary. The sealing of the firewall on all affected aircraft must be inspected as described in Piper Service Bulletin No. 161a and when found deficient must be resealed in accordance with the manufacturer's service bulletin or accepted aeronautical practices. 2. On PA-22 aircraft Serial Numbers 22-2700 to 22-6194 inclusive, the procedure outlined in 1 should be followed within the next 100 hours of operation. 3. Periodic inspection should be made of the exhaust system in accordance with Piper Service Bulletin No. 161a pertaining to inspection of the exhaust stack gaskets,exhaust stacks, muffler assembly, and muffler tailpipe. 4. The sealing of the firewall on all PA-16, PA-20 and PA-22 aircraft must be inspected at 100-hour intervals in accordance with Piper Service Bulletin No. 161a. If found deficient, it must be resealed in accordance with the manufacturer's service bulletin or accepted aeronautical practices. 5. The 100-hour inspection requirement on all Model Piper PA-22 aircraft, Serial Numbers 22-1 to 22-6194 inclusive, can be eliminated if Piper Kit, P/N 754237 or equivalent, is installed.