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2017-19-17: We are superseding Airworthiness Directive (AD) 2016-17-02, which applied to certain Dassault Aviation Model FALCON 900EX and FALCON 2000EX airplanes. AD 2016-17-02 required revising the airplane flight manual (AFM) to include procedures to follow when an airplane is operating in icing conditions. AD 2016-17-02 also provided optional actions after which the AFM revision may be removed from the AFM. Since we issued AD 2016-17-02, we have determined additional actions are necessary to address the identified unsafe condition. This new AD retains the requirement of AD 2016-17-02, and also requires a detailed inspection of the wing anti-ice system ducting (anti-ice pipes) for the presence of a diaphragm, and replacement of ducting or re- identification of the ducting part marking. We are issuing this AD to address the unsafe condition on these products.
58-01-06: 58-01-06 PIPER: Applies to Model PA-23 Aircraft, Serial Numbers 23-1 to 23-1219 Inclusive. Compliance required as indicated. Due to the installation of the front stabilizer-to-fuselage attachment fitting P/N 17093-00, on additional aircraft to those covered by AD 57-13-09 and since special inspections are not required when the redesigned fitting P/N 17093-03 is installed, this supersedes the portions of AD 57-13-09 concerning this fitting and revision issued on Card No. 57-22. Inspect visually for cracks, the front stabilizer fitting, P/N 17093-00 every 100 hours until replaced with the redesigned fitting P/N 17093-03. Fittings found cracked must be replaced. (Piper Service Bulletin No. 160 dated October 7, 1957, covers the same subject.)
58-01-07: 58-01-07 PIPER: Applies to All J-3 Series and J-5 Series Aircraft. Compliance required by February 1, 1958. To preclude the possibility of failures of the fork end of the turnbuckles in the control system, the following inspection and rework is necessary. Failures of the fork end of the turnbuckles have occurred in the area covered by the safety wire. This results from binding caused by the attaching bolt being drawn up too tightly on the fork end of the turnbuckle. Inspect the turnbuckle to horn attachment at the elevators, rudder and ailerons to determine that an AN 23-12 clevis bolt is installed with one AN 960-10 washer under the nut. This assembly should swivel freely.
61-03-03: 61-03-03 HARTZELL: Amdt. 247 Part 507 Federal Register February 2, 1961. Applies to All HC-93Z30-2D and HC-B3Z30-2D Propellers Installed on Pratt and Whitney R-985 Engines. (These May Be Found in Such Aircraft As Beech 18 Series, Grumman G-21A, and Lockheed 12A.) Compliance required as indicated. Due to failure or cracking of several B-1803 cylinders in the threaded area, resulting in the loss of engine oil and control of the propeller, the following shall be accomplished, unless the replacement required in paragraph (b) has already been made: (a) Check for oil leaks in the propeller hub within the next 25 hours' time in service and every 25 hours' time in service thereafter until the replacement required in paragraph (b) is accomplished. It is not necessary to remove the spinner for this inspection. If an oil leak is discovered replace the cylinder as provided in paragraph (b) before further flight. (b) Unless already accomplished, replace cylinder B-1803 andcollar 834-7 with cylinder B-1803-1 and collar 834-7A at the next propeller overhaul or within the next 400 hours of time in service, whichever occurs first. The hub model dash number is to be restamped -2E in place of the present -2D. (Hartzell Bulletin No. 73 dated April 18, 1960, and Bulletin No. 73 amendment dated September 8, 1960, cover this subject.) This directive effective February 13, 1961.
2017-19-20: We are adopting a new airworthiness directive (AD) for certain General Electric Company (GE) CT7-8A and CT7-9B model turboshaft engines. This AD was prompted by reports from the manufacturer that the high-pressure compressor (HPC) impeller installed on these engines may have suffered from material degradation during the manufacturing process. This AD requires removal of the affected HPC impellers. We are issuing this AD to address the unsafe condition on these products.
57-10-02: 57-10-02 PIPER: Applies to Model PA-23 Aircraft Serial Numbers 23-1 to 23-1391 Inclusive. Compliance required as soon as possible but not later than June 15, 1957. It has been reported that the flare on the oil pressure gage line has cracked or broken where it attaches to the connector fitting on the aft side of the firewall resulting in loss of oil pressure. Inspect both right and left oil pressure line flares to determine whether or not they are normal and also to determine that the lines in the area of the flares are not cracked or broken. Lines that are cracked or broken or have defective flares should be cut off and reflared. Care should be taken that there is no line strain against the fitting or the retainer block and that the line into the fitting is straight. If the line is too short for this repair, it should be cut off and spliced using a connector and flexible hose, Piper P/N 17766-07 or equivalent. (Piper Service Bulletin No. 152A covers this subject.)
2017-19-14: We are adopting a new airworthiness directive (AD) for certain Dassault Aviation Model FALCON 900EX airplanes. This AD was prompted by a determination that new or more restrictive maintenance requirements and/or airworthiness limitations are necessary. This AD requires revising the maintenance or inspection program, as applicable, to incorporate new or more restrictive maintenance requirements and/or airworthiness limitations. We are issuing this AD to address the unsafe condition on these products.
56-26-03:
79-10-09: 79-10-09 HUGHES HELICOPTERS: Amendment 39-3466. Applies to Hughes Model 369D Helicopters certificated in all categories, except Serial Numbers 0409 and subsequent. Compliance required as indicated. To prevent failure of the tail rotor pitch control assembly, which could result in loss of tail rotor control, accomplish the following: (a) Before further flight after the effective date of this AD, check the 369D21803-3 locknut and tang washer P/N MS172209 or HS1551S238 of the tail rotor pitch control assembly P/N 369D21800 as follows: Pull back the non-rotating boot P/N 369D21806 from the tail rotor pitch control assembly P/N 369D21800 and check the locknut and tang washer by hand for looseness. (1) If the locknut or tang washer is loose, remove and replace the tail rotor pitch control assembly with a serviceable assembly with a white dot on the locknut, before further flight. (2) If either the locknut or tang washer is not loose, repeat the check forlooseness of the locknut or tang washer before each engine start-up. See (c) below for termination action. The checks required by this AD may be performed by the pilot. NOTE 1: For the requirements regarding the listing of compliance and method of compliance with this AD in the helicopter's permanent maintenance record, see FAR 91.173. (b) For the operators that have previously complied with Hughes Service Information Notice DN-37, dated December 1, 1978, the following applies: (1) If the tail rotor pitch control assembly was repaired due to looseness of the locknut or tang washer, repeat the checks in (a) above and remove and replace the tail rotor pitch control assembly within 150 hours time in service after the receipt of the airmail letter dated April 9, 1979. (2) For the tail rotor pitch control assemblies for which records of repair are not known, see (b)(1) above. (3) For the tail rotor pitch control assemblies which were repaired per DN- 37 with no looseness of the locknut or tang washer, the requirements of this AD do not apply. NOTE 2: The pitch control assemblies which have loose locknut or tang washers must be overhauled per the Hughes Model 369D Helicopter Basic Handbook of Maintenance Instructions and DN-37, dated December 1, 1978, before return to service. (c) For those assemblies which do not have loose tang washers or locknuts, per paragraph (a)(2) above, within 100 hours of time in service from the receipt of the airmail letter dated April 9, 1979, accomplish DN-37. (d) Equivalent checks and repairs may be used when approved by the Chief, Aircraft Engineering Division, FAA Western Region. This amendment becomes effective May 17, 1979, as to all persons except those persons to whom it was made immediately effective by the airmail letter dated April 9, 1979, which contained this amendment.
80-19-11 R1: 80-19-11 R1 GATES LEARJET: Amendment 39-3932 as amended by Amendment 39-5039. Applies to the following models and serial number airplanes, unless noted: MODEL SERIAL NUMBERS 23 23-003 through 23-099 24, 24A 24-100 through 24-180 24B, 24B-A 24-181 through 24-217 24-219 through 24-229 24C, 24D, 24D-A 24-218, 24-230 through 24-328 24E, 24F, 24F-A 24-329 and subsequent 25, 25A 25-003 through 25-060 25B, 25C 25-061, 25-067 through 25-201, 25-204, 25-205 25D, 25F 25-206 and subsequent 28, 29 28-001 and subsequent 29-001 and subsequent 35, 36, 35A, 36A 35-001 and subsequent 36-001 and subsequent COMPLIANCE: Required as indicated, unless previously accomplished. To assure that the crew is provided additional instructions for the safe operation of the airplane and that the airplane's automatic flight control and stall warning systems are properly adjusted, accomplish the following: A) Before further flight, insert the following information in the FAA Approved Airplane Flight Manual and operate the airplane in accordance with these insertions: 1. In Section 1, LIMITATIONS, adjacent to AIRSPEED LIMITS, MAXIMUM OPERATING SPEED VMO/MMO: a. Delete any procedures relative to exceeding VMO or MMO. b. Add the following limitation: WARNING: Do not extend the spoilers, or operate with the spoilers deployed, at speeds above VMO/MMO due to the significant nose down pitching moment associated with spoiler deployment. 2. In Section 1, LIMITATIONS, add a new limitation: TRIM SYSTEMS a. To assure proper trim systems operations, the BEFORE STARTING ENGINES trim system checks must be successfully completed before each flight. WARNING: Failure to conduct a complete pitch trim preflight check prior to each flight increases the probability of an undetected system failure. An additional single failure in the trim system could result in a runaway. In certain critical flight conditions an unrestrained runaway could result in high speeds, severe buffet, wing roll off, loads in excess of structural limit and extremely high forces necessary for recovery. b. Pitch trim system runaway training that actually involves running the trim in flight to simulate malfunctions is prohibited. 3. In Section 1, LIMITATIONS, adjacent to STALL WARNING SYSTEM, add the following: On Models 23, 24, 24A, 24B, 24B-A, 24D, 24D-A, 25, 25A, 25B, and 25C with unmodified wings, and the same models with Howard/Raisbeck Mark II wings: WARNING: Do not intentionally fly the airplane slower than initial stall warning (shaker) onset. 4. In Section 1, LIMITATIONS, adjacent to YAW DAMPER: a. Delete any references to disengaging the yaw damper before landing, or landing with the yaw damper engaged. b. Add the following yaw damper requirements: On landing, the following yaw damper disengage procedures shall apply: (1) The airplane shall be configured forlanding at least 500 ft. AGL for normal landing: (2) The yaw damper shall be disengaged during the landing flare. CAUTION: If landings are attempted in turbulent air conditions with the yaw damper OFF, the airplane may exhibit undesirable lateral-directional (Dutch-Roll) characteristics. These characteristics are improved as the wing/tip fuel is consumed. The pilot shall observe the NOTE relative to turbulence contained in the BEFORE LANDING section of Section II of the Airplane Flight Manual and increase airspeed as required. 5. In Section II, NORMAL OPERATION PROCEDURES, adjacent to BEFORE STARTING ENGINES Procedures: a. Delete current preflight procedures on all trim systems. b. Add the following new trim system preflight checks: NOTE: Some early Model 23, 24 airplanes incorporate a cutoff button that interrupts pitch, roll and yaw axes. (1) Pitch Trim Selector Switch--EMER (or SEC). (2) Operate EMERGENCY (or SEC) pitch trim switch NOSE UP and NOSE DOWN and check for stabilizer movement. Stabilizer movement will be approximately one-half of the rate of primary trim. (3) Either Control Wheel Trim Switch - Operate NOSE UP and NOSE DOWN. Trim motion shall not occur. (4) Pitch Trim Selector Switch - OFF. (5) Actuate pilot's and copilot's Control Wheel Trim and Trim Arming Switches (if applicable) and pedestal EMERGENCY (or SEC) Pitch Trim Switch. Trim motion shall not occur. (6) Pitch Trim Selector Switch - NORM (or PRI). (7) EMERGENCY (or SEC) Pitch Trim Switch - Operate NOSE UP and NOSE DOWN. Trim motion shall not occur. NOTE: On all Model 23 airplanes and Model 24 (Serial Number 24-100 through 24-169) airplanes, except for those incorporating Accessory Kit AAK70-3, trim motion will occur. (8) Pilot's Control Wheel Trim Switch - Without depressing arming button (if applicable), move switch to LWD, RWD, NOSE UP, and NOSE DOWN; trim motion shall not occur. Depress arming button (if applicable); trim motion shall not occur. Then depress arming button (if applicable) and move switch to LWD, RWD, NOSE UP and NOSE DOWN; trim motion shall occur. (9) Repeat Step (8) for Copilot's Control Wheel Trim Switch. (10) Trim by positioning Copilot's Control Wheel Trim Switch in one direction; then trim in opposite direction using the Pilot's Control Wheel Trim Switch. Pilot's trim shall override the Copilot's trim. Repeat for all lateral and pitch trim positions. (11) Pilot's Control Wheel Trim Switch - NOSE UP. While trimming, depress Control Wheel Master Switch (if applicable) or Cutoff Button (if applicable); trim motion shall stop when the Control Wheel Master Switch is held. Repeat procedure for NOSE DN condition; trim motion shall stop. Repeat procedure for LWD & RWD lateral trim on airplanes equipped with Cutoff Button. (The procedures in this paragraph are not applicable to Model 25, S.N. 25-003 through 25-205 and Model 24, S.N. 24-170 through 24-328, except those airplanes modified by AAK76-4A.) (12) Repeat Step (11) using copilot's Control Wheel Trim Switch, and Control Wheel Master Switch (if applicable), or Cutoff Button (if applicable). (13) YAW TRIM Switch - Operate each half separately (if installed); trim motion shall not occur. (14) YAW TRIM Switch - Operate both halves simultaneously; trim motion shall occur. On aircraft with Cutoff Button, check that the Cutoff Button stops the trim. (15) Trim - Set all axes for takeoff. 6. In Section III, EMERGENCY PROCEDURES, add a new PITCH UPSET (NOSE-UP or NOSE-DOWN) Emergency Procedure: A nose-up pitch axis malfunction or nose-up pitch trim system runaway can result in extremely high pitch attitudes, heavy airframe buffet, and require control forces in excess of 75 pounds for recovery. A nose-down pitch axis malfunction, nose-down pitch trim system runaway, or nose-down overspeed can result in extremely high airspeeds and require control forces in excess of 75 pounds for recovery. WARNING: Do not extend spoilers on any nose-down pitch upset at any speed due to significant nose-down pitching moment associated with spoiler deployment. NOTE: Control pressures may be heavy. Copilot assistance is recommended with this procedure. IMMEDIATELY: a. Attitude Control - As required to maintain aircraft control. - If in nose-up attitude, roll into bank or maintain existing bank until the aircraft nose passes through the horizon. - If in nose-down attitude, level the wings before pulling the nose up. b. Thrust levers - As required. (If in nose-down attitude, immediately reduce thrust levers to IDLE position.) c. Control Wheel Master Switch or Cutoff Button - Depress and hold until step g. is accomplished. d. PITCH TRIM Selector Switch - OFF. e. STALL WARNING Switches - OFF. WARNING: On any speed excursions beyond MMO, the elevator control must be smoothly and steadily applied to prevent encountering excessive aileron activity and airframe buffet. Beyond .85 M1, a 1.5 g pull-up may be sufficient to excite aileron activity and the g level must be limited to that required to maintain lateral control. AFTER AIRCRAFT CONTROL IS REGAINED: f. Spoilers - Check retracted. g. Autopilot's Pitch Circuit Breaker - Pull. h. If control force continues, select other trim system and retrim the aircraft. i. Isolate malfunctioning system by switching systems ON one at a time. Pause between activating each system to determine the defective system. 7. In Section IV, PERFORMANCE DATA, adjacent to the appropriate takeoff charts, add the following: Increase all Chart V1, VR and V2 speeds by: a. Model 23, 24, 24A, 24B, 24B-A, 24D, 24D-A, 25, 25A, 25B, 25C, with unmodified wings, plus 5 KNOTS Indicated Airspeed. b. Model 23, 24, 24A, 24B, 24B-A, 24D, 24D-A, 25, 25A, 25B, 25C, with Howard/Raisbeck Mark II wings, plus 5 KNOTS Indicated Airspeed. (Increase applies to FLAP 10 and FLAP 20 charts, and is not applicable to FLAP 10 OVERSPEED chart.) 8. In Section IV, PERFORMANCE DATA, adjacent to each TAKEOFF DISTANCE CHART, add the following: Increase all chart takeoff distances by: a. Model 23, 24, 24A, 24B, 24B-A, 24D, 24D-A, 25, 25A, 25B, 25C, with unmodified wings, plus 10 percent. b. Model 23, 24, 24A, 24B, 24B-A, 24D, 24D-A, 25, 25A, 25B, 25C, with Howard/Raisbeck Mark II wings, plus 10 percent. (Increase applies to FLAP 10 and FLAP 20 charts, and is not applicable to FLAP 10 OVERSPEED chart.) 9. In Section IV, PERFORMANCE DATA, adjacent to each TAKEOFF WEIGHT LIMITS chart, add the following: a. Reduce the Limiting Weight-Brake Energy takeoff weights for Model 23, 24, 24A, 24B, 24B-A, 24D, 24D-A, 25, 25A, 25B, 25C, with unmodified wings, 500 lbs. b. Reduce the FLAP 10 and FLAP 20 takeoff weight limits for Model 23, 24, 24A, 24B, 24B-A, 24D, 24D-A, 25, 25A, 25B, 25C, with Howard/Raisbeck Mark II wings, 500 lbs, if the airplane is at climb limited gross weight and if takeoff weight is above 14,500 lbs. For takeoff weights above 14,000 lbs. and below 14,500 lbs., reduce the weight to 14,000 lbs. Takeoff weight reduction not applicable to FLAP 10 OVERSPEED. 10. In Section IV PERFORMANCE DATA, adjacent to LANDING APPROACH SPEEDS chart, add the following: Increase all chart Landing Approach Speeds by: a. Model 23, 24, 24A, with unmodified wings, plus 8 KNOTS Indicated Airspeed. b. Model 23, 24, 24A with ECR 736 (CJ610-6 engines and increased gross weight), and Model 24B, 24B-A, 24D, 24D-A, with unmodified wings, plus 4 KNOTS Indicated Airspeed. c. Model 25, 25A, with unmodified wings, plus 3 KNOTS Indicated Airspeed. d. Model 25, 25A, with unmodified wings with ECR 936 (AAK 70-5), plus 5 KNOTS Indicated Airspeed. e. Model 25B, 25C, with unmodified wings, plus 5 KNOTS Indicated Airspeed. f. Model 23, 24, 24A, 24B, 24B-A, 24D, 24D-A, 25, 25A, 25B, 25C, with Howard/Raisbeck Mark II wings, plus 5 KNOTS Indicated Airspeed. 11. In Section IV, PERFORMANCE DATA, adjacent to each LANDING DISTANCE CHART, add the following: Increase all chart landing distances by: a. Model 23, 24 and 24A, with unmodified wings, plus 10 percent. b. Model 23, 24, 24A with ECR 736 (CJ610-6 engines and increased gross weight) and Model 24B, 24B-A, 24D, 24D-A, with unmodified wings, plus 5 percent. c. Model 25, 25A, with unmodified wings, plus 4 percent. d. Model 25, 25A, with unmodified wings with ECR 936, (AAK70-5) plus 7 percent. e. Model 25B, 25C, with unmodified wings, plus 7 percent. f. Model 23, 24, 24A, 24B, 24B-A, 24D, 24D-A, 25, 25A, 25B, 25C, with Howard/Raisbeck Mark II wings, plus 7 percent. 12. In Section IV, PERFORMANCE DATA, adjacent to the LANDING WEIGHT LIMITS CHART, add the following: Reduce the LimitingWeight-Brake Energy landing weights as follows: a. Model 23, 24, 24A, with unmodified wings, 800 lbs. b. Model 23, 24, 24A, with ECR 736 (CJ610-6 engines and increased gross weight), and Model 24B, 24B-A, 24D, 24D-A with unmodified wings, 400 lbs. c. Model 25, 25A, with unmodified wings, 300 lbs. d. Model 25, 25A, with unmodified wings with ECR 936 (AAK70-5), 500 lbs. e. Model 25B, 25C, with unmodified wings, 500 lbs. f. Model 23, 24, 24A, 24B, 24B-A, 24D, 24D-A, 25, 25A, 25B, 25C, with Howard/Raisbeck Mark II wings, 500 lbs. NOTE: In order to comply with the requirements of paragraph A of this Airworthiness Directive, this AD, or a duplicate thereof, may be used as a temporary amendment to the Airplane Flight Manual and carried in the aircraft as part of the Airplane Flight Manual until replaced by the identical revisions to the Airplane Flight Manual provided by the manufacturer and approved by the FAA. The temporary Airplane Flight Manual Changes required by paragraph A) of this AD may be accomplished by the holder of at least a private pilot certificate issued under Part 61 of the Federal Aviation Regulations on any airplane owned or operated by that person who must make the prescribed entry in the Airplane Maintenance Records indicating compliance with paragraph A) of this AD. B) Except for the roll axis of the FC-200 autopilot installed on Model 35, 35A, 36 and 36A airplanes, within the next 75 flight hours, conduct the following inspections to assure capability of manually overriding the Automatic Flight Control Systems: 1. Energize the airplane electrical system by applying 28 VDC electrical power. 2. Roll Axis a. On airplanes equipped with FC-110 autopilot, remove the electrical power from the FC-110 Autopilot Computer. Open the computer and identify the Roll Calibration Board. On the Roll Calibration Board, temporarily install, in parallel with R18 (82 ohm) resistor, a 39 ohm, one watt resistor. Restore the electrical power and engage the Autopilot with the control wheel centered and verify that the roll slip clutch breakaway occurs by rotating the control wheel briskly (45 degrees per second) in both directions. If slippage is not verified, remove the capstan and adjust to proper torque per the appropriate Gates Learjet Service Manual. Return Autopilot Computer to original configuration and accomplish a functional check of the autopilot. 3. Yaw Axis a. Effective on all models: (1) Check and adjust the yaw capstan slip clutch torque (primary and secondary where applicable) in accordance with the appropriate Gates Learjet Service Manual. 4. Pitch Axis a. Effective on Models 24D, 24D-A, 24E, 24F, 24F-A, 25B, 25C, 25D, 25F, 28, 29, 35, 35A, 36 and 36A airplanes and airplanes incorporating Gates Learjet Kits AAK71-12 or AMK80-3 (torquers): (1) With the Autopilot disengaged, turn on both stall warning switches and move the control wheel forward and aft at a rapid rate (one second - stop to stop). Note the drag associated with control movement. Turn off the stall warning switches and repeat the rapid fore and aft movement. Note the decrease in drag, which is an indication that the electric disconnect clutch functions properly by disconnecting the drag of the pitch servo (torquer) from the control system. b. Effective on Models 23, 24, 24B, 24B-A, 24C, 25 and 25A airplanes except airplanes incorporating Gates Learjet Kits AAK71-12 or AMK80-3: (1) Check and adjust the pitch capstan slip clutch for proper torque in accordance with the appropriate Gates Learjet Service Manual. C) On airplane Models 35, 35A, 36 and 36A, within the next 150 flight hours conduct the following inspection of the FC-200 autopilot roll axis to assure capability of manually overriding that axis of Automatic Flight Control Systems: 1. Energize the airplane electrical system by applying 28 VDC electrical power. 2. Check andadjust the roll capstan slip clutch for proper torque in accordance with the appropriate Gates Learjet Service Manual. D) Submit a written report of any out of tolerance roll, yaw, or pitch axis capstan slip torque to the Federal Aviation Administration (FAA), Aircraft Certification Program, Room 238, Terminal Building 2299, Mid-Continent Airport, Wichita, Kansas 67209. (Reporting approved by the Office of Management and Budget Order OMB No. 04-R0174.) E) To assure proper operation of the Stall Warning Accelerometer Unit, perform, within the next 25 flight hours, inspection of the Stall Warning Accelerometer in accordance with appropriate Gates Learjet Service Bulletin SB 23, 24, 25-301A, SB 28, 29-27-3A, or SB 35, 36-27-12A. Submit a written report on any discrepancy discovered during this inspection to FAA, Aircraft Certification Program, Room 238, Terminal Building 2299, Mid-Continent Airport, Wichita, Kansas 67209. (Reporting approved by Office of Management and Budget Order OMB No. 04-R0174.) NOTE: The owner/operator is responsible for submitting reports required by this AD. F) Airplanes may be flown in accordance with FAR 21.197 to a location where alterations and inspections required by this AD can be accomplished. G) Any equivalent method of compliance with this AD must be approved by the Manager, Aircraft Certification Program, FAA, Room 238, Terminal Building No. 2299, Mid-Continent Airport, Wichita, Kansas 67209. Telephone: (316) 942-4285. H) On or before October 1, 1986, accomplish the requirements of paragraphs 1. or 2., below, on Learjet Models 23, 24, 24A, 24B, 24B-A, 24D, 24D-A, 25, 25A, 25B, 25C, with unmodified wings, at an FAA certificated maintenance repair station, and insert in the appropriate sections of the Airplane Flight Manual (AFM) the permanent AFM revision pertaining to procedures and performance associated with Airplane Modification Kit (AMK) 83-4 or 84-5. The limitations and performance information required paragraphs A)3., A)7., A)8., A)9., A)10., A)11., and A)12. of this AD are superseded by the AFM revision included with these kits. 1. Incorporate AMK 83-4 to improve airplane handling qualities and aerodynamic stall characteristics, or 2. Incorporate AMK 84-5 to make the stall prevention system (pusher) operation consistent with the airplane performance and limitations. All persons affected by this proposal who have not already received these documents from the manufacturer may obtain copies upon request to the Gates Learjet Corporation, P.O. Box 7707, Wichita, Kansas 67277. This information may be examined at the FAA, Northwest Mountain Region, 17900 Pacific Highway South, Seattle, Washington, or at the FAA, Central Region, Room 100, 1801 Airport Road, Mid-Continent Airport, Wichita, Kansas. This AD supersedes the airmail letter AD on the same subject issued August 4, 1980, and identified as AD 80-16-06. Amendment 39-3932 became effective on October 9, 1980,to all persons except those to whom it has already been made effective by an airmail letter from the FAA dated September 4, 1980. This amendment 39-5039 becomes effective May 20, 1985.
54-19-01: 54-19-01 PIPER: Applies to Models PA-18A and PA-18A Restricted Category Aircraft With Dusting Venturi, Up to and Including Serial Number 18-3752. Compliance required not later than October 15, 1954. There have been several instances of excessive CO concentration in the cockpit when the dusting venturi is used. Such contamination has serious adverse effects upon pilot reaction. To prevent CO from entering the cockpit, a new trim plate should be installed and a new brake line cover plate should be placed over the brake line where the line enters the bottom of the fuselage. (Piper Service Letter No. 225, dated August 23, 1954, covers the same subject.)
2017-19-12: We are superseding Airworthiness Directive (AD) 2014-13-17, which applied to all Airbus Model A300 series airplanes; Airbus Model A300 B4-600, B4-600R, and F4-600R series airplanes, and Model A300 C4- 605R Variant F airplanes (collectively [[Page 43672]] called Model A300-600 series airplanes); and Airbus Model A310 series airplanes. AD 2014-13-17 required repetitive functional tests of the circuit breakers for the fuel pump power supply, and replacement of certain circuit breakers. This new AD requires installation of fuel pumps having a new standard, which terminates the repetitive functional tests. This AD was prompted by our determination that installation of a newly developed fuel pump standard will better address the unsafe condition. We are issuing this AD to address the unsafe condition on these products.
2017-19-03: We are adopting a new airworthiness directive (AD) for all Dassault Aviation Model MYSTERE-FALCON 900 airplanes. This AD was prompted by a determination that new or more restrictive maintenance requirements and/or airworthiness limitations are necessary. This AD requires revising the maintenance or inspection program, as applicable, to incorporate new or more restrictive maintenance requirements and/or airworthiness limitations. We are issuing this AD to address the unsafe condition on these products.
2017-18-15: We are adopting a new airworthiness directive (AD) for all Airbus Model A300 B4-600R and Model A300 F4-600R series airplanes; Model A300 B4-603, B4-622, and C4-605R Variant F airplanes; and Model A310-203, -221, -222, -304, -322, -324, and -325 airplanes. This AD was prompted by an evaluation by the design approval holder indicating that a section of the fuselage structure above the forward cargo door is subject to widespread fatigue damage (WFD). This AD requires an inspection for cracks of the fastener and tooling holes at certain locations and a check of the diameter of the holes, and repair or modification of the affected fuselage structure if necessary. We are issuing this AD to address the unsafe condition on these products.
77-13-07: 77-13-07 BEECH: Amendment 39-2932. Applies to Model 200 (Serial Numbers BB-2 thru BB-250) airplanes certificated in all categories. Compliance: Required as indicated, unless already accomplished. To assure that wings will carry the 50 feet/second gust design loads, within 50 hours' time in service after the effective date of this AD, accomplish the following: A) With aileron control system pinned in the neutral position rerig the droop of right and left aileron to zero degrees, with a tolerance of plus one-half degree up, minus zero degrees down in accordance with Beechcraft Service Instructions Number 0906 or later approved revisions. B) Any equivalent method of compliance with this AD must be approved by the Chief, Engineering and Manufacturing Branch, FAA, Central Region. This amendment becomes effective July 1, 1977.
55-13-01: 55-13-01 PIPER: Applies to Model PA-23, Serial Numbers 23-160 to 23-162 Inclusive, 23-166 to 23-168 Inclusive, 23-170, 23-172, 23-177, 23-193 to 23-219 Inclusive, 23-221 to 23- 224 Inclusive, 23-226 to 23-229 Inclusive, 23-231 to 23-234 Inclusive, and 23-237. Compliance required by August 1, 1955. It has been found that all the PA-23 Models listed above have installed two flexible fuel valve controls that have been found to flex excessively at the splice. The addition of an idler bellcrank should be accomplished by the use of Piper Kit No. 754 103 and the instructions included. Owners of PA-23's with auxiliary fuel tanks must not operate the auxiliary fuel system until this modification has been accomplished. (Piper Immediate Action Service Bulletin No. 139 applies to this malfunction.)
2017-18-19: We are adopting a new airworthiness directive (AD) for all Airbus Model A310-203, -204, -221, -222, -304, -322, -324, and -325 airplanes. This AD was prompted by reports of cracking in the drainage holes on the lower skin panel in the center wing box between frames (FR) 42 and FR46. This AD requires repetitive rotating probe inspections for cracking of the trellis boom drainage holes, the holes in the stringers bottom, and the holes of the inner pump, and corrective actions if necessary. We are issuing this AD to address the unsafe condition on these products.
2017-18-16: We are adopting a new airworthiness directive (AD) for certain The Boeing Company Model 737-700 and -700C series airplanes. This AD was prompted by a report that, for certain airplanes, the nose-up pitch trim limit and associated warning will allow the horizontal stabilizer position to be set outside acceptable limits for a mis-trimmed takeoff condition. This AD requires, depending on airplane configuration, replacing certain pitch trim light plates, relocating certain position warning horn switches, revising certain software, removing a certain placard, and doing related investigative and corrective actions if necessary. We are issuing this AD to address the unsafe condition on these products.
53-04-01: 53-04-01 PIPER: Applies to All Model PA-18 Aircraft Which Have Not Had the Control Stick Retention Device Modified to Incorporate a Through Bolt. Compliance required by April 1, 1953. In order to prevent the control stick inadvertently pulling out of the socket, continue the existing hole for the retention pin on the quick detachable control stick spring device through the control stick and socket and install an AN 3-14 or AN 3-14A through bolt. (Piper Service Letter No. 162 dated March 6, 1951, covers this same subject.)
54-02-02: 54-02-02 FEDERAL SKIS: Applies to All Universal (Stinson) 108 and Cessna 170, 170A and 170B Airplanes Equipped With Federal Models AWB-2500 and AWB-2500A Wheel Skis and Piper PA-20 Airplanes Equipped With Federal Model AWB-2100 Wheel Skis. Compliance required as soon as possible but not later than February 15, 1954. To preclude the possibility of the ski dropping down against the mechanical rigger and possible subsequent damage to the aircraft structure, the rigging arrangement must be revised in accordance with Federal Aircraft Works Drawing No. 11D1077.
77-13-04: 77-13-04 GRUMMAN-AMERICAN: Amendment 39-2931. Applies to G-164B, S/N's 138B, 142B, 177B and up, and G-164A, S/N's 1686, 1695 and up, certificated in all categories. Accomplish the following, unless previously accomplished: a) Before further flight, install a placard adjacent to the parking brake handle stating: "Warning - Parking Brake OFF Prior to Landing." b) Within the next 50 hours in service, install the alteration defined in Grumman American Aviation Corporation Service Bulletin No. 60, dated May 20, 1977, or an equivalent method approved by the Chief, Engineering and Manufacturing Branch, FAA Eastern Region. When this is accomplished, the warning placard may be removed. This Amendment becomes effective June 29, 1977.
2017-18-10: We are adopting a new airworthiness directive (AD) for certain Diamond Aircraft Industries GmbH Models DA 42, DA 42 M-NG, and DA 42 NG airplanes. 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 crack formation on the flap bell crank, which could cause the flap bell crank to fail. We are issuing this AD to require actions to address the unsafe condition on these products.
51-27-02: 51-27-02 BOEING: Applies to all Model B-377 Aircraft equipped with Goodrich H-3-626 and H-3-650 Wheel Assemblies. To be accomplished as indicated below. As a precautionary measure to preclude serious hazards which may result from fatigue cracks in Goodrich H-3-626 and H-3-650 wheel assemblies, all wheels shall be carefully inspected with at least a 4x magnifying glass at the periodic inspection following each 50th landing. These inspections shall be conducted until the wheels must be retired from service. The wheels must be retired at the time replacement wheels are made available. Due to the fact that a large quantity of B-377 wheel halves were delivered to the operators without being assigned a serial number, the following description will identify both the wheel to be replaced and also the replacement wheel. The wheels to be retired from service are those wheels which have a tapered spoke and no reinforcing beading around the inner side of the spoke cutout area. Replacement wheels are those wheels which have a reinforcing bead around the inner side of the spoke cut-out area and have either straight or tapered spokes. (Goodrich Service Bulletin No. 27, dated November 14, 1951, covers this identification problem in further detail.) Type I - Cracks Progressing Across a Spoke. The wheel should be rejected when there is more than one crack of this type to a spoke, when there is more than one crack in each spoke cutout, or when a crack is in excess of 1 inch in length. Type II - Cracks Progressing Radially Across the Brake Drum Mounting Flange. More than one crack of this type in any one spoke cutout will be cause for rejection. Type III - Cracks in the Tie Bolt Recess and the Junction of the Drum Mounting Area. This type crack is in a noncritical area and is cause for wheel rejection only when either of the following conditions occur: 1. The crack extends to the spoke cutout. 2. Developed cracks on either side of the recess progress to within 1 3/8 inches of each other. The above wheel crack limitations are based on the recommendations of the B. F. Goodrich Co. (Goodrich Service Bulletin No. 17 covers this subject and illustrates the three types of wheel cracks.)
2017-16-07: We are adopting a new airworthiness directive (AD) for all Airbus Model A330-200, A330-200 Freighter, A330-300, A340-500, and A340-600 series airplanes; and A340-313 airplanes. This AD was prompted by the discovery of Tartaric Sulfuric Anodizing (TSA)/Chromic Acid Anodizing (CAA) surface treatment in certain bulk cargo door frame holes of certain airplanes. This AD requires inspection of the fuselage bulk cargo door frames at specific locations, and corrective action if necessary. We are issuing this AD to address the unsafe condition on these products.
65-28-02: 65-28-02 FAIRCHILD: Amdt. 39-169 Part 39 Federal Register December 14, 1965. Applies to Model F-27 Series Airplanes Incorporating Upper Engine Mount Fittings P/N's 27-503148-31, -32, -41, -42, -51, -52, -61, -62, and Lower Engine Mount Fittings P/N's 27-503149-11, -12, -31, -32, -41, and -42. Compliance required as indicated. To detect cracks in the engine mount attachment fittings, accomplish the following: (a) Within the next 50 hours' time in service after the effective date of this AD, unless already accomplished within the 250 hours' time in service before the effective date of this AD, inspect the engine mount fittings in accordance with (d). (b) For engine mount fittings with 10,000 or more hours' time in service on the effective date of this AD, reinspect in accordance with (d) at intervals not to exceed 300 hours' time in service from the last inspection. (c) For engine mount fittings with less than 10,000 hours' time in service on the effective date of this AD, reinspect in accordance with (d) before the accumulation of 10,150 hours' time in service, unless accomplished after the accumulation of 9,850 hours' time in service, and thereafter at intervals not to exceed 300 hours' time in service from the last inspection. (d) Visually inspect for cracks each engine mount fitting (eight per airplane), on the forward side of the firewall, including all welds, using at least a 10-power glass or an FAA-approved equivalent. Clean all surfaces of each engine mount fitting prior to inspecting, (without removing fitting). (e) If a crack is found during the inspection specified in (d) before further flight, replace the engine mount fitting with a part of the same part number that has been inspected in accordance with (d), or an equivalent part approved by the Chief, Engineering and Manufacturing Branch, FAA Eastern Region. (f) Upon request of the operator, an FAA maintenance inspector, subject to prior approval of the Chief,Engineering and Manufacturing Branch, FAA Eastern 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 directive effective December 14, 1965.