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97-15-07: This document publishes in the Federal Register an amendment adopting Airworthiness Directive (AD) 97-15-07, which was sent previously to all known U.S. owners and operators of Aeromot-Industria Mecanico Metalurgica Ltda. (Aeromot) Model AMT-200 powered sailplanes. This AD requires immediately inspecting, using non-destructive testing (NDT) methods, the forward horizontal stabilizer front bolt, P/N 53451, for defects (scratches, damaged threads, or surface cracks, etc.), and replacing the bolt immediately if found defective or at a certain time period if not found defective. This AD was the result of a failure of the forward horizontal stabilizer bolt, part number (P/N) 53451, on one of the affected powered sailplanes. This failure was caused by a low cycle fatigue crack that was induced by over torquing the bolt. The actions specified by this AD are intended to prevent failure of the forward horizontal stabilizer bolt, which could result in separation of the horizontal stabilizer from the powered sailplane and consequent loss of control.
58-05-03: 58-05-03 VICKERS: Applies to All Viscount 700 Series Aircraft. Compliance required as indicated. In accordance with the British Air Registration Board's list of "Essential Modifications and Inspections", compliance with the following Vickers-Armstrong corrective measures is considered mandatory. The FAA concurs and considers compliance therewith mandatory. Compliance provisions are detailed in the Vickers publications referenced in each item. A. MODIFICATION ELEVATOR AND RUDDER CONTROL LOCKS. To insure freedom of movement over the complete range of travel of the gust lock lever, install special shoulder greaser bolts in accordance with Mod. D.1421 to prevent over-tightening of the nut securing the lock lever. Also, on elevator gust lock unit P/N A.1949-1 chamfer the sides of the slotted recess in accordance with TNS No. 64, issue 2. Compliance required by April 15, 1958. B. CONNECTION OF A.C. PHASE TRANSFORMER. Mod. D.1677 revises the electrical circuit, placing transformers on the input side of the circuit breakers to prevent feed back through the primary of A.C. phases transformers. (This is further to Mod. D.1513.) Compliance required by April 15, 1958. C. REDISTRIBUTION OF BUS-BAR SUPPLIES TO PITOT HEAD HEATERS. By supplying the heaters from different bus-bars, complete loss of pitot-head deicing due to failure of one bus-bar is prevented. Mod. D.2019 covers this subject. Compliance required by September 15, 1958. D. INSPECTION AND MODIFICATION OF ELEVATOR ANTI-BALANCE TAB MECHANISM. To prevent over-center jamming of the elevator anti-balance tab mechanism, PTL No. 160 prescribes precautionary checks of the tab mechanism maximum travel as well as modifying the link. (PTL No. 160 and Mod. D.2239 cover this subject.) Compliance required by April 15, 1958. E. IMPROVE THE LOCKING OF THE FORWARD ELEVATOR AND AILERON LEVER GROUP AT COCKPIT STATION 53 AND AILERON LEVER GROUP AT FUSELAGE STATION 462.68. To prevent lost movement in the aileron control system and possible loosening of the aileron bellcrank level, Mod. D.2091 introduces new fittings and attachment means. Compliance with PTL No. 141 or Mod. D.2091 required by April 15, 1958. F. INCREASE CUT-OUT IN PILOTS FLOOR FOR ELEVATOR CONTROL CLEARANCE. Mod. D.2120 increases size of cutaway in pilots floor area to eliminate fouling of elevator rod attachment bolt when incorrectly assembled. Compliance with PTL No. 140 or Mod. D.2120 required by July 15, 1958. G. FAIRLEADS FOR ELEVATOR TRIM TAB CHAIN AT FRAME STATION 783. In order to preclude jamming of the elevator trim tab chain where it passes through the frame Station 783, a fairlead Vickers drawing No. 70152, P/N's 1205, 1207, and 1209, should be installed at the cutaway in the frame in accordance with Vickers Mod. D.1602. Compliance required by April 15, 1958. H. INSPECTION OF TYPE S.1, S.2 AND S.4 RELAYS. To prevent shorting of relays due to loosening ofthe securing screws on the armature springs PTL No. 147 specifies inspection for security and locking of the shakeproof lock washers and to further lock by the application of either Glyptal Varnish CS.184 or Bakelite Varnish V.130. Compliance required by April 15, 1959.
2012-02-02: We are superseding an existing airworthiness directive (AD) for certain Cessna Aircraft Company (Cessna) Models 172R and 172S airplanes. That AD currently requires you to inspect the fuel return line assembly for chafing; replace the fuel return line assembly if chafing is found; and inspect the clearance between the fuel return line assembly and both the right steering tube assembly and the airplane structure, adjusting as necessary. Since we issued that AD, we have received a field report of a fuel return line chafing incident on a Cessna Model 172 airplane with a serial number (S/N) that was not included in the AD. This AD retains the actions of the current AD and adds S/Ns to the Applicability section of the AD. Chafing of the fuel return line assembly could lead to fire. This new AD requires the actions of the current AD and adds S/Ns to the Applicability section of the AD. We are issuing this AD to correct the unsafe condition on these products.
65-09-04: 65-09-04 LOCKHEED: Amdt. 39-59 Part 39 Federal Register April 28, 1965. Applies to Models 188A and 188C Series Aircraft. Compliance required as indicated. There have been a number of cracks in the wing upper surface in the area of the main landing gear rib forgings, some of which have extended out from previous crack repairs and from reinforcements. (a) Within the next 700 landings after the effective date of this AD, unless already accomplished within the last 700 landings prior to the effective date of this AD, accomplished the following or an equivalent approved by the Chief, Aircraft Engineering Division, FAA Western Region: (1) Inspect the external surface of the upper wing planks visually for cracks in the area bounded by the front spar-to-plank joint and the rear spar-to-plank joint between Wing Station 221 and the nacelle outboard skate angle, and between Wing Station 155 and the nacelle inboard skate angle; (2) Inspect for cracks by dye penetrant technique, that portion of the external surface of the upper wing plank area located 2.5 inches forward and 2.5 inches aft of the No. 3-to- No. 4 plank joint between Wing Station 221 and the nacelle outboard skate angle and between Wing Station 155 and the nacelle inboard skate angle, which is not hidden by the existing 0.25 inch reinforcing doubler; and (3) Inspect the tang of the No. 4 plank at the No. 3-to-No. 4 plank joint between Wing Station 221 and the nacelle outboard skate angle and between Wing Station 155 and the nacelle inboard skate angle, for vertical cracks at the tang radius, by either of the following: (i) Externally by the ultrasonic technique described in Lockheed Service Bulletin 88/SB-620, Section 2.A. (2)(b), pages 5 through 7, or later FAA-approved revision. Visually reinspect internally if cracks are indicated by the ultrasonic inspection. (ii) Visual internal inspection. (b) Repair any cracks found during the inspections required by this AD before further flight, in accordance with the Lockheed Electra Structural Repair Manual or an equivalent approved by the Chief, Aircraft Engineering Division, FAA Western Region, except that the aircraft may be flown in accordance with the provisions of FAR 21.197 to a base where the repair can be made. NOTE: Regional approval required by (b) may be facilitated by obtaining prior approval of a Structural DER. (c) Reinspect internally by visual means or by X-ray technique any area repaired in accordance with (b) within 1,400 landings after inspection in accordance with (a), and at intervals thereafter not to exceed 1,400 landings from the last inspection. Upon the request of the operator, an FAA maintenance inspector, subject to prior approval of the Chief, Aircraft Engineering Division, FAA Western Region, may authorize alternative inspection techniques and intervals for specified repaired areas, provided the request is accompanied by technical data describing the length, location,and disposition of the crack, and the details of the repair. (d) Reinspect in accordance with (a) all areas found free of cracks at intervals not to exceed 1,400 landings from the last inspection. (e) When subsequent reinspection shows evidence of growth of a crack repaired in accordance with (b), remove the repair, and repair the engine cracked area in accordance with (b) before further flight except that the aircraft may be flown in accordance with the provisions of FAR 21.197 to a base where the repair can be made. (f) The repetitive inspections required by (c) for repaired areas and (d) for areas free from cracks may be discontinued if the upper wing planks are inspected, repaired if necessary, and reinforced in accordance with the following: (1) Inspect for cracks as specified in (a) the areas not hidden by the 0.25 inch reinforcing doubler and not previously repaired; (2) Inspect for cracks as specified in (c) repaired areas; (3) Inspect for cracks asspecified in (a) the area hidden by the 0.25 inch reinforcing doubler with the doubler removed; (4) Repair in accordance with (b) any cracks found; and (5) Incorporate the reinforcements described by Lockheed Drawings 841314, 841315, SED/64-9010 and SED/64-9011 in accordance with the accomplishment instructions, Section 2.A. through 2.L. of Lockheed Service Bulletin 88/SB-619, or later FAA-approved revision, or equivalent reinforcements approved by the Chief, Aircraft Engineering Division, FAA Western Region. (g) For the purpose of complying with this AD, subject to acceptance by the assigned FAA maintenance inspector, the number of landings may be determined by dividing each aircraft's hours' time in service by the operator's fleet average time from takeoff to landing for the aircraft type. (h) Upon request of the operator, an FAA maintenance inspector, subject to prior approval of the Chief, Aircraft Engineering Division, FAA Western 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. This supersedes AD 64-25-02. This directive effective April 28, 1965.
97-16-04: This amendment supersedes an existing airworthiness directive (AD), applicable to certain Saab Model SAAB SF340A and SAAB 340B series airplanes, that currently requires inspections to detect improper connections of the wire harness installation to the cartridges of the fire extinguishers in the engine nacelles, correction of any discrepancy, and modification of the wiring. This amendment adds a revised modification of that wiring, which, if accomplished, would terminate the inspections currently required by the existing AD. This amendment is prompted by reports indicating that, due to the removal of a certain clamp during maintenance, these fire extinguisher cartridges still could be connected incorrectly after the modification required by the existing AD has been accomplished. The actions specified by this AD are intended to prevent incorrect wiring of the cartridges, which would result in inability of the fire extinguishers to jointly discharge extinguishing agent into anacelle in the event of an engine fire.
97-15-17: This amendment adopts a new airworthiness directive (AD), applicable to certain Saab Model SAAB 2000 series airplanes, that requires replacing the Abex alternating current (AC) electric motor with a new modified Abex AC electric motor having an improved fan. This amendment is prompted by reports indicating that the integrated hydraulic package (IHP) unit stopped functioning during flight because the fan on the AC electric motor came into contact with the housing of the motor due to inadequate clearance. The actions specified by this AD are intended to prevent loss of IHP function, which, if combined with other hydraulic system failures, could result in reduced controllability of the airplane.
76-20-07: 76-20-07 BELLANCA: Amendment 39-2738. Applies to Bellanca CH, CH-300 Pacemaker, 300-W Pacemaker, CH-400 Skyrocket, E Pacemaker, F. Skyrocket, 14-9, 14-9L, 14- 12F-3, 14-13, 14-13-2, 14-3-3, 14-13-3W, 31-42 Pacemaker Aircraft, certificated in all categories. To prevent failure of forward or aft wing spars due to wood decay caused by water collecting in the wing, accomplish the following: Prior to the next flight after receipt of this AD unless already accomplished, and at each annual inspection thereafter, perform the following or an approved equivalent inspection: (a) Using a blunt tool (e.g. screwdriver handle) tap along the entire length of the upper and lower skin directly over the front and rear spars starting inboard of the root rib. A decayed area will emit a sound that has a discernible difference in quality in comparison to an undecayed section. Fabric covered wings can be inspected by depressing the fabric over the spar prior to tapping. A suspected area mustbe inspected visually through available skin access holes, or cut outs in the skin. The front face of the rear spar and the front spar must also be checked by inserting an awl through the drain holes on the lower surface of the wing and tapping. (b) Using a light and mirror, inspect the interior of the wings for moisture, water stains, pooled dust or dirt which may indicate previously collected water, wood discoloration, woodchecks, delamination of surfaces and corrosion of metallic surfaces. The inspection will include among other areas: 1. Inboard of root rib including the ends of front and rear spar and spar attachment and the inside surface of top and bottom skin and root rib. 2. Forward face of front spar visible through gap between fuselage and lower wing surface. 3. Forward face of rear spar to root rib, visible through gap between fuselage and lower wing surface. 4. Aft face of rear spar, visible through lightening holes or cut-outs after flaps are removed. For airplanes not equipped with flaps a skin cut-out on the bottom wing may be used. 5. Flap stops for airplanes so equipped. 6. Forward face of front spar in the vicinity of landing light for airplanes so equipped. 7. The plywood wing skin interior for airplanes so equipped. 8. Interior of wing at trailing edge where fabric terminates including the area around the flaps for airplanes so equipped. 9. Inspect for restriction in all drain holes in wing bottom surface. 10. Inspect the seal (e.g. fabric fillet, channel rubber with compatible sealant etc.) at the intersection of the fuselage and top wing skin leading edge fairing. 11. Inspect for serviceability of the seal (e.g. doped fabric fillet, silicone rubber sealant, etc.) between the fuel tank scupper and wing skin. 12. The spar butt ends. (c) On the exterior surfaces of the wing, inspect for any damage that may allow water entry such as cracks or breaks in the paint,bubbles, discoloration, boils, soft spots or other evidence of fabric deterioration or fabric delamination from wood skin. Include in the inspection the wing root area, top and bottom surfaces of front and rear spars inboard of landing gear, wing walk area for airplanes so equipped and the top and bottom of the fuel tank area. (d) If any defects set forth in paragraph (b) and (c) are detected, repairs must be accomplished in accordance with FAA-approved standard practice AC 43-13-1A or an equivalent repair approved by the Chief, Engineering and Manufacturing Branch, FAA, Eastern Region, prior to further flight, except that the airplane may be flown in accordance with FAR 21.197 to a base where the repair can be performed. (e) All skins that are cut out in accordance with paragraph (a) must be repaired in accordance with paragraph (d). This amendment is effective October 8, 1976 and was effective upon receipt for all recipients of the airmail directive of July 14, 1976.
2011-24-04: We are adopting a new airworthiness directive (AD) for certain Model DC-10-10, DC-10-10F, and MD-10-10F airplanes. This AD was prompted by reports of three instances of fuel leaks in the lower cap splice of the wing rear spar at station Xors=409. Investigation revealed the fuel leak was due to a crack in the lower cap. If not corrected, this condition could result in fuel leaks or cracking of the lower wing skin and structure, causing possible inability of the structure to sustain the limit load and adversely affecting the structural integrity of the airplane. This AD requires repetitive inspections for cracking on the lower cap of the rear spar of the left and right wings between stations Xors=417 and the outboard edge of the lower cap splice of the wing rear spar at station Xors=400; temporary and permanent repairs if necessary; and repetitive inspections of repaired areas, and corrective actions if necessary. We are issuing this AD to correct the unsafe condition on these products.
2012-02-07: 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 theairplane.
97-15-14: This amendment adopts a new airworthiness directive (AD) that applies to Industrie Aeronautiche e Meccaniche Rinaldo Piaggio S.p.A. (Piaggio) Model P-180 airplanes. This action requires inspecting for cracks around the vertical pin and the torque tube bottom flange of the rudder, and the fasteners that connect the torque tube to the bottom flange (torque tube bottom flange assembly). If cracks are not found, repetitively inspect until cracks are visible. If cracks are evident, this action requires modifying the rudder torque tube bottom flange assembly by replacing the cracked part with a part of improved design, which terminates the repetitive inspection. This AD is the result of several reports of fatigue cracks around the pin that vertically supports the rudder axle. The actions specified by this AD are intended to prevent fatigue cracks in the rudder torque tube bottom flange, which could result in loss of rudder control and possible loss of the airplane.