Results
81-02-09: 81-02-09 BOEING: Amendment 39-4024. Applies to all Boeing Model 727-100 and 727-100C series airplanes equipped with the aft airstair emergency extension system, except all-cargo configurations and those airplanes where the aft stair emergency extension system has been deactivated.\n \n\tCompliance is required as indicated. Accomplish the following:\n \n\tA.\tPrior to June 1, 1981, replace or modify the aft airstair emergency extension control handle in a manner approved by the Chief, Seattle Area Aircraft Certification Office, FAA Northwest Region. (Note: Accomplishment of Boeing Service Bulletin 727-52-120 dated March 21, 1980, or later FAA approved revisions has been approved as a means of compliance with the requirements of this AD.) \n\n\tB.\tUpon request of the operator, an FAA aviation safety inspector, subject to prior approval by the Chief, Seattle Area Aircraft Certification Office, FAA Northwest Region, may adjust the compliance date if the request contains substantiating data to justify the change. \n\n\tThis amendment becomes effective April 1, 1981.
2017-19-27: We are adopting a new airworthiness directive (AD) for certain Bombardier, Inc., Model DHC-8-401 and -402 airplanes. This AD was prompted by the discovery of cracking on two test spoiler power control unit (PCU) manifolds during testing by the manufacturer. This AD requires replacement of affected spoiler PCUs. We are issuing this AD to address the unsafe condition on these products.
59-26-02: 59-26-02 PIPER: Applies to PA-24 and PA-24 "250" Airplanes Serial Numbers 24-1 To 24-1373 Inclusive. Compliance required by January 15, 1960. To prevent clogging, the two fuel cell vent tubes which are located under the wings shall be modified in the following manner: Measure a distance of 1/2-inch down from the bottom of the wing skin along the forward side of each protruding vent tube. At this point, cut the tube off at a 45-degree angle to the bottom skin so that the end of the tube remains square. (Piper Immediate Action Service Bulletin No. 180 covers this subject.)
2017-19-15: We are adopting a new airworthiness directive (AD) for certain Technify Motors GmbH TAE 125-02 reciprocating engines. This AD requires replacement of the clutch with a dual mass flywheel. This AD was prompted by a loss of engine power in flight caused by oil leaking from the gearbox radial shaft sealing ring that contaminated the clutch. We are issuing this AD to correct the unsafe condition on these products.
59-25-05: 59-25-05 FORNEY (ERCOUPE): Applies to All (Ercoupe) Forney Aircraft With Serial Numbers Up to 3,335 Inclusive. Compliance required by December 31, 1959, and thereafter every 100 hours of operation or periodic inspection, whichever occurs first. Fatigue failures have continued to occur in the rudder main rib where the control horn is attached after installation of reinforcement plates. Therefore, it is required that a visual inspection be made of the area around the rudder control horn for excessive deflection of the horn, canning of rudder skin, or any other unusual peculiarity which would indicate main rudder rib damage. If damage is evident, rudder rib Erco P/N 415-240 12 L/R must be replaced with Forney P/N F-24015 L/R or equivalent. This inspection may be discontinued when the heavier gage rib is installed. (Forney Service Bulletin No. 105 covers this subject.) This supersedes AD 47-20-07.
59-13-02: 59-13-02 PIPER: Applies to Models PA-24 and PA-24 "250" Aircraft Serial Numbers 24-1 to 24-978 Inclusive and 24-980. Compliance required within the next 100 hours of operation or by October 1, 1959, whichever occurs first. Service experience indicates that cracks have developed in the aileron balance weight attachment bulkheads. These bulkheads are riveted to the front spar of the aileron and are the supports to which the balance weight arm is attached. To reduce the probability of failure of the aileron balance weight arm attachment install reinforced bulkheads on both ailerons except on Serial Number 24-980 replace the balance weight attachment bulkhead on right aileron only. (Piper Service Bulletin No. 173 also covers this subject and states "Service Kit, Part Number 734-233, is available from your nearest Piper distributor or dealer free of charge if the airframe serial number is included on the purchase order.")
79-19-01 R2: 79-19-01 R2 BOEING: Amendment 39-3556 as amended by Amendment 39-4087 is further amended by Amendment 39-4486. Applies to all Boeing 720/720B, 707-300, 707-400, 707- 300B, and 707-300C series airplanes. \n\tA.\tAfter the effective date of this amendment, perform a low frequency eddy current inspection for cracks in the wing lower surface splice stringers in accordance with Boeing Service Bulletin 3226, Rev. 5, dated November 15, 1981, or later FAA approved revisions, or in a manner approved by the Manager, Seattle Aircraft Certification Office, FAA, Northwest Mountain Region. Inspections are to be made at the threshold times, within the prescribed initial interval and at repetitive intervals shown below: \n\tAt the inboard nacelle strut drag brace, the affected lower skin area covered by the fairing may be visually inspected for cracks and evidence of fuel leakage. If crack indications are noted in stringers, or skin cracks or fuel leakage are found at the diagonal brace fairing area, tank entry and inspection by high frequency eddy current of the wing splice stringers is required. \n\n\nAirplane\nThreshold\nInitial Inspection within\n\nRepetitive Interval \n\n\n\nunless accomplished \nwithin the last \n\n720/720B\n14,000 ldgs\n715 ldgs\n715 ldgs\n1,430 ldgs\n707-300/400\n21,000 ldgs\n1,675 ldgs\n1,675 ldgs\n3,350 ldgs\n707-300B\n19,000 ldgs\n1,425 ldgs\n1,425 ldgs\n2,850 ldgs\n707-300C\n17,000 ldgs\n725 ldgs\t\n725 ldgs\n1,450 ldgs \n707-300C\n\n17,000 ldgs\n1,425 ldgs\n1,425 ldgs\n2,850 ldgs \n\n(passenger only) \n\n\n\n\n\t\t\t\t\t\t\t\t\t\t \n\tB.\tIf cracks are found, repair prior to further revenue flight in accordance with a method approved by the Manager, Seattle Aircraft Certification Office, FAA, Northwest Mountain Region. \n\tC.\tFor the purpose of complying with this AD and subject to acceptance by the assigned FAA Maintenance Inspector, the number of landings may be determined by dividing each airplane's time-in-service by the operator's fleet average from takeoffto landing for the airplane type. \n\tD.\tUpon request of the operator, an FAA Maintenance Inspector, subject to prior approval of the Manager, Seattle Aircraft Certification Office, FAA, Northwest Mountain Region may adjust the inspection interval if the request contains substantiating data to justify the increase for that operator. \n\tE.\tAirplanes with cracked splice stringers may be flown in accordance with FAR 21.197 to a base where repairs can be performed. \n\tAll persons affected by this directive who have not already received these documents from the manufacturer may obtain copies upon request to the Boeing Commercial Airplane Company, P.O. Box 3707, Seattle, Washington 98124. These documents may also be examined at the FAA, Northwest Mountain Region, 9010 East Marginal Way South, Seattle, Washington. \n\n\tAmendment 39-3556 became effective September 18, 1979. \n\tAmendment 39-4087 became effective May 17, 1981. \n\tThis Amendment 39-4486 becomes effective November 15, 1982.
2017-19-19: We are adopting a new airworthiness directive (AD) for certain Rolls-Royce plc (RR) Trent XWB-75, Trent XWB-79, Trent XWB-79B, and Trent XWB-84, turbofan engines. This AD requires replacement of the low-pressure compressor (LPC) case support inboard pins. This AD was prompted by LPC case support inboard pins that may have reduced integrity due to incorrect heat treatment. We are issuing this AD to correct the unsafe condition on these products.
78-24-06 R1: 78-24-06 R1 BELL: Amendment 39-3358 as amended by Amendment 39-4032. Applies to Bell Models 206L and 206L-1 helicopters, certificated in all categories (Airworthiness Docket No. 78-ASW-53). Compliance required as indicated, unless previously accomplished. To prevent possible failure of the horizontal stabilizer, P/N 206-023-119, all dash numbers, accomplish the following: (a) Within the next 10 hours' time in service after the effective date of this AD, modify the left and right upper stabilizer supports, P/N 206-023-100-009 and -010, respectively, in accordance with Bell Helicopter Textron Service Bulletin 206L-78-3 dated October 23, 1978, or Bell Helicopter Textron Alert Service Bulletin 206L-80-16 dated November 17, 1980, or FAA approved equivalent, so that the critical area can be checked. (b) Before the first flight of each day after compliance with paragraph (a), visually check the stabilizer skin area exposed by the cutouts, in the upper stabilizer supports, forcracks. (1) If a crack is found, remove and replace the horizontal stabilizer before further flight. (2) If no cracks are found, continue the repetitive check specified above. (c) The checks required by this AD may be performed by the pilot. NOTE: For the requirements regarding the listing of compliance and method of compliance with this AD in the aircraft's permanent maintenance record, see FAR 91.173. The manufacturer's specifications and procedures identified and described in this directive are incorporated herein and made a part hereof pursuant to 5 U.S.C. 552(a)(1). All persons affected by this directive who have not already received these documents from the manufacturer may obtain copies upon request to the Service Manager, Bell Helicopter Textron, P.O. Box 482, Fort Worth, Texas 76101. These documents may also be examined at the Office of the Regional Counsel, Southwest Region, FAA, 4400 Blue Mound Road, Fort Worth, Texas, and at FAA Headquarters, 800 Independence Avenue, S.W., Washington, D.C. A historical file on this AD which includes the incorporated material in full is maintained by the FAA at its headquarters in Washington, D.C., and at the Southwest Regional Office in Fort Worth, Texas. Amendment 39-3358 became effective December 5, 1978. This amendment 39-4032 becomes effective May 11, 1981.
2017-19-26: We are superseding Airworthiness Directive (AD) 2008-12-04, which applied to certain The Boeing Company Model 737-600, -700, -700C, -800, and -900 series airplanes. AD 2008-12-04 required various repetitive inspections to detect cracks along the chem-milled steps of the fuselage skin, and to detect missing or loose fasteners in the area of a certain preventive modification or repairs; replacement of the time-limited repair with a permanent repair, if applicable; and applicable corrective actions which would end certain repetitive inspections. This AD reduces the post-modification inspection compliance times, limits installation of the preventive modification to airplanes with fewer than 30,000 total flight cycles, and adds repetitive inspections for modified airplanes. This AD was prompted by an evaluation by the design approval holder (DAH) that indicated that the upper skin panel at the chem-milled step above the lap joint is subject to widespread fatigue damage (WFD) if themodification was installed after 30,000 total flight cycles. We are issuing this AD to address the unsafe condition on these products.
58-22-03: 58-22-03 PIPER: Applies to All Model PA-23 Aircraft Equipped With Goodrich G-3-787 Main Wheel Assemblies. Compliance required as indicated. Failures of the Goodrich G-3-787 wheel assembly are being reported. These wheels may be continued in service subject to inspection as specified below, or replaced with Goodrich G-3-880 wheel assembly or Cleveland Aircraft Products wheel assembly Model 3060 or 3080A and brake assembly Model 3000-500 or another equivalent approved type wheel and brake combination. 1. Remove wheels and tires and inspect wheels at each one hundred hours of operation or at each tire change, whichever occurs first. 2. Inspect the flange area of both wheel halves by means of Dy-Check or Zyglo, whichever is available. Cracks may appear on either the inside or outside surface of the flange. 3. If cracks are present in the flange area, remove the defective wheel half from service and replace as indicated above. 4. To detect possible flange failures during preflight inspection, look for outward deformation of flanges. A wheel with a flange failure will appear to wobble when rotated. (Goodrich Service Bulletin No. 102, dated July 25, 1956, and Piper Service Letter No. 291, dated June 5, 1957, covers this same subject.)
60-21-02: 60-21-02 LOCKHEED: Amdt. 207 Part 507 Federal Register October 7, 1960. Applies to All Model 18 Aircraft Which Have Been Converted to the "Learstar" Configuration. Compliance required as indicated. Flight tests have disclosed that excessive temperature stratification at the carburetor air screen exists during operation with partial carburetor air preheat. This stratification results in erroneous readings on the cockpit carburetor air temperature gage and may, under icing conditions, cause ice formation on the cold portions of the air screen and in the carburetor. While full preheat is available if needed for ice elimination, only partial preheat should be used for continuous operation under certain temperature conditions in order to avoid exceeding the engine manufacturer's carburetor air temperature limit of 38 degrees C. The following action is required: (a) Effective November 15, 1960, Learstar aircraft shall be restricted against operation in known icing conditions until modifications to the air preheat system covered in paragraph (b) are accomplished. The following placard shall be posted in full view of the pilot: "OPERATION INTO KNOWN ICING CONDITIONS PROHIBITED." The limitations section of the FAA approved Airplane Flight Manual is hereby amended to incorporate this limitation. (b) The operating restriction into known icing conditions shall continue until modifications are accomplished to the carburetor air preheat system which will result in conservative C.A.T. indications for the prevention of ice formation on the carburetor screen and engine induction system at all preheat positions. Such modifications shall also permit operation with preheat under varying power and ambient temperature conditions without resulting in excessive C.A.T. A satisfactory modification to meet the requirements is covered in FAA approved PacAero Engineering Corporation Service Bulletin No. 14, dated August 26, 1960. An FAA approved airplane flight manualrevision setting forth recommended procedures for safe operation of the system will be supplied by Pac Aero with the modification kit. Any deviations from the modifications or procedures set forth in the service bulletin and airplane flight manual revision must be approved by FAA, Region Four, Engineering and Manufacturing Branch, Los Angeles, California. (c) Upon compliance with paragraph (b), the operating restriction set forth in paragraph (a) is cancelled.
79-17-01: 79-17-01 ROCKWELL INTERNATIONAL: Amendment 39-3525. Applies to Rockwell NA-265-60 airplanes which have been modified by Raisbeck Group Supplemental Type Certificate SA687NW. To prevent flutter caused by the accumulation of undrained water in the elevators, accomplish the following, unless already accomplished, within 30 days or 30 flight hours, whichever occurs first: A. Modify the elevators to provide water drainage provisions in accordance with Raisbeck Service Bulletin No. 8, dated July 13, 1979, or later FAA approved revisions. B. Using the procedure specified in Raisbeck Service Bulletin No. 8, dated July 13, 1979, or later FAA approved revisions, determine the static balance of the elevators and if required, rebalance them in accordance with instructions in that service bulletin. C. Equivalent modifications may be approved by the Chief, Engineering and Manufacturing Branch, Northwest Region. The manufacturer's specifications and procedures identified and described in this directive are incorporated herein and made a part hereof pursuant to 5 U.S.C. 552(a)(1). All persons affected by this directive who have not already received these documents from the manufacturer, may obtain copies upon request to The Raisbeck Group, 7777 Perimeter Road South, Boeing Field International, Seattle, Washington 98108. These documents may also be examined at FAA Northwest Region, 9010 East Marginal Way South, Seattle, Washington 98108. This amendment becomes effective August 15, 1979.
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.