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98-07-09: This amendment adopts a new airworthiness directive (AD), applicable to Bell Helicopter Textron, Inc. (BHTI) Model 47B, 47B-3, 47D, 47D-1, 47G, 47G-2, 47G-2A, 47G-2A-1, 47G-3, 47G-3B, 47G-3B-1, 47G-3B-2, 47G-3B-2A, 47G-4, 47G-4A, 47G-5, 47G-5A, 47H-1, 47J, 47J-2, 47J-2A, and 47K helicopters, that requires installing a safety washer kit designed to preclude separation of the stabilizer bar damper link (damper link) if the damper link rod end bushing (bushing) loosens and exits the damper link rod end. This amendment is prompted by two reported incidents in which the bushings loosened and exited the damper link rod ends, allowing the damper link to slide over the retention bolt and separate from the stabilizer bar (in the first incident), and from the hydraulic damper (in the second incident). The actions specified by this AD are intended to prevent failure of the damper link assembly, which can result in degraded control response and subsequent loss of control of the helicopter.
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2021-22-11: The FAA is adopting a new airworthiness directive (AD) for certain MD Helicopters Inc. (MDHI), Model 369D, 369E, 369F, 369FF, 369H, 369HE, 369HM, 369HS, 500N, and 600N helicopters. This AD was prompted by a report of a spiral crack in the pilot-to-copilot tail rotor torque tube (torque tube). This AD requires a one-time visual and recurring borescope inspections of the torque tube and depending on the results, removing the torque tube from service. The FAA is issuing this AD to address the unsafe condition on these products.
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93-23-02: 93-23-02 JETSTREAM AIRCRAFT LIMITED: Amendment 39-8736. Docket 93-NM-180-AD.
Applicability: Model 4101 airplanes; as listed in Jetstream Aircraft Limited Alert Service Bulletin J41-A52-021, dated August 24, 1993; certificated in any category.
Compliance: Required as indicated, unless accomplished previously.
To prevent the inability of passengers to evacuate the airplane through the main entrance door in an emergency, accomplish the following:
(a) Within 30 days after the effective date of this AD, perform a detailed visual inspection to detect damage of the electrical loom in the main entrance door in accordance with Jetstream Aircraft Limited Alert Service Bulletin J41-A52-021, dated August 24, 1993.
(b) If any damage is found during the inspection required by paragraph (a) of this AD, prior to further flight: Replace the loom with a new loom; modify the electrical loom routing and support by re-routing the loom and support clear of the speedlock solenoid and the inner skin of the door; perform an operational test of the main entrance door warning system; and perform a functional test of the main entrance door, examining the clearance between the speedlock solenoid and the door inner skin to ensure that the electrical loom cannot become trapped between those parts; in accordance with Jetstream Aircraft Limited Alert Service Bulletin J41-A52-021, dated August 24, 1993.
(c) If no damage is found during the inspection required by paragraph (a) of this AD, prior to further flight, modify the electrical loom routing and support by re-routing the loom and support clear of the speedlock solenoid and the inner skin of the door; perform an operational test of the main entrance door warning system; and perform a functional test of the main entrance door, examining the clearance between the speedlock solenoid and the door inner skin to ensure that the electrical loom cannot become trapped between those parts; in accordance with Jetstream Aircraft Limited Alert Service Bulletin J41-A52-021, dated August 24, 1993.
(d) An alternative method of compliance or adjustment of the compliance time that provides an acceptable level of safety may be used if approved by the Manager, Standardization Branch, ANM-113, FAA, Transport Airplane Directorate. Operators shall submit their requests through an appropriate FAA Principal Maintenance Inspector, who may add comments and then send it to the Manager, Standardization Branch, ANM-113.
NOTE: Information concerning the existence of approved alternative methods of compliance with this AD, if any, may be obtained from the Standardization Branch, ANM-113.
(e) Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate the airplane to a location where the requirements of this AD can be accomplished.
(f) The inspection, replacement, modification, and tests shall be done in accordance with Jetstream Aircraft Limited Alert Service Bulletin J41-A52-021, dated August 24, 1993. This incorporation by reference was approved by the Director of the Federal Register in accordance with 5 U.S.C. 552(a) and 1 CFR Part 51. Copies may be obtained from Jetstream Aircraft, Inc., P.O. Box 16029, Dulles International Airport, Washington, DC 20041-6029. Copies may be inspected at the FAA, Transport Airplane Directorate, 1601 Lind Avenue, SW., Renton, Washington; or at the Office of the Federal Register, 800 North Capitol Street, NW., suite 700, Washington, DC.
(g) This amendment becomes effective on December 7, 1993.
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51-24-02: 51-24-02 MARTIN: Applies to All Model 202 Aircraft.
Compliance required at the 3,000-hour period following the modification of the fin attachments per Martin 202 Service Bulletin No. 99 and every 3,000 hours thereafter.
To insure that the fin-to-fuselage attachments contain no structural defects, and to reduce the possibility of fretting corrosion, accomplish the following inspections and shim installations:
(1) Determine that the three fin pins (P/N 2021A11549, 2021A14243 and 2021A14244) are not worn to less than 0.292 inch in width, and are securely attached to the fin.
(2) Inspect the three fuselage receptacles (P/N 2021U27375 and 2021U27415 left and right) for the three fin pin attachments. Replace any receptacle whose slot is greater than 0.357 inch in width.
(3) Using a 4- to 6-power glass, inspect the fin forward attach plate (P/N 2021D26541), the fuselage chords (P/N 2021D25001 and 25002), and the four reinforcing straps (P/N 2021U43293, 43294 and 43295left and right) for evidence of cracks or fretting corrosion.
(4) Cracks or fretting corrosion found in item (3) should be removed with complete crack removal verified by satisfactory etch inspection. Reworked surfaces should be repfinished to a polished finish. All bolt holes should be chamfered and polished to an 0.020-inch radius. If a crack is found in a hole in any reinforcing strap, the strap must be replaced.
(5) Etch inspect the fin rear chords (P/N 2021D14234), using a 10- to 15-power glass, to detect cracks in the chords. Cracks and any surface roughness must be removed. If cracks are found in a splice hole or any splice holes have been distorted, such that the hole bushings are no longer a press fit, the holes are to be reamed oversize and oversize bushing installed. (If the hole had a crack, re-etch inspect the area to insure crack removal.) If the diameter of a hole exceeds 0.843 inch, that fin chord must be replaced.
(6) Minimum thicknesses of refinishedstructural members are:
Part
Number
Minimum Thickness (inch)
Fin attach plate
2021U43641
0.177
Fin rear chord
2021D14233 & 4
0.312
Fuselage attach plate
2021D26541
0.050
Fuselage Chord
2021D25001 & 2
0.250
Reinforcing strap
2021U43293, 4 & 5
0.187
(7) When reinstalling fin, insert shims of 1310 clear 01 vinyl sheet (or equivalent), press polished both sides, to fit between the faying service of the fin, fuselage and reinforcing straps at the fin rear spar attachment. Insert similar shims between the faying surfaces of the fuselage and fin attach plates at the fin front spar attachment.
Upon request of the operator, an FAA maintenance inspector, subject to 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 suchoperator.
(Martin 202 Service Bulletin No. 182 covers this same subject.)
Revised August 19, 1961.
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99-01-09: This document publishes in the Federal Register an amendment adopting Airworthiness Directive (AD) 99-01-09 which was sent previously to all known U.S. owners and operators of Sikorsky Aircraft Corporation (Sikorsky) Model S-76C helicopters by individual letters. This AD requires, before further flight, installing a placard in the cockpit adjacent to the fuel quantity gauge that states "No flight operations to be conducted with less than 250 lbs. fuel in each tank." This AD must be placed in the Operating Limitations section of the Rotorcraft Flight Manual. This AD also requires, within 50 hours time-in-service (TIS) or 30 calendar days, whichever occurs first, defueling, engine starting, and if necessary, inspecting fuel supply lines. This amendment is prompted by an in-flight engine flame-out that occurred on October 27, 1998. The actions specified by this AD are intended to prevent air from getting into a fuel supply line when there is less than 250 lbs. of fuel in either fuel tank, engine flame-out, and a subsequent forced landing.
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92-03-05: 92-03-05 SAAB-SCANIA: Amendment 39-8160. Docket No. 91-NM-148-AD.
Applicability: Models SAAB SF340A and SAAB 340B series airplanes; as listed in SAAB Service Bulletin 340-33-030, Revision 2, dated September 27, 1991; certificated in any category.
Compliance: Required as indicated, unless accomplished previously.
To prevent premature failure of the emergency lights after an emergency landing, accomplish the following:
(a) Within 120 days after the effective date of this AD, modify the exit and dome light assemblies, in accordance with SAAB Service Bulletin 340-33-030, Revision 2, dated September 27, 1991.
(b) An alternative method of compliance or adjustment of the compliance time, which provides an acceptable level of safety, may be used when approved by the Manager, Standardization Branch, ANM- 113, FAA, Transport Airplane Directorate.
NOTE: The request should be forwarded through an FAA Principal Maintenance Inspector, who may concur or comment andthen send it to the Manager, Standardization Branch, ANM-113.
(c) Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate the airplane to a location where the requirements of this AD can be accomplished.
(d) The modification required by this AD shall be done in accordance with SAAB Service Bulletin 340-33-030, Revision 2, dated September 27, 1991, which contains the following list of effective pages:
Page Number
Revision Level
Date
1-2, 4
2
September 27, 1991
6-7
1
April 29, 1991
3, 5, 8-10
(original)
(undated)
This incorporation by reference was approved by the Director of the Federal Register in accordance with 5 U.S.C. 552(a) and 1 CFR Part 51. Copies may be obtained from SAAB-SCANIA AB, Product Support, S-581.88, Linkoping, Sweden. Copies may be inspected at the FAA, Transport Airplane Directorate, 1601 Lind Avenue S.W., Renton, Washington; or at the Office of the Federal Register, 1100 L Street N.W., Room 8401,Washington, D.C.
(e) This amendment (39-8160, AD 92-03-05) becomes effective on March 9, 1992.
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93-24-03: 93-24-03 BEECH AIRCRAFT CORPORATION: Amendment 39-8752; Docket No. 93-CE- 22-AD. Supersedes AD 92-15-06, Amendment 39-8300 which superseded AD 91-23-07, Amendment 39-8076.
Applicability: The following Beech model and serial numbered airplanes, certificated in any category:
MODELS
SERIAL NUMBERS
35-33, 35-A33, 35-B33,
35-C33, E33, F33, and G33
CD-1 through CD-1304
35-C33A, E33A, and F33A
CE-1 through CE-1425
E33C and F33C
CJ-1 through CJ-179
36 and A36
E-1 through E-2518
A36TC and B36TC
EA-1 through EA-500
Compliance: Required as indicated after the effective date of this AD, unless already accomplished (compliance with superseded AD 92-15-06 or superseded AD 91-23-07).
To prevent separation of the rudder from the airplane caused by cracks in the forward rudder spar, accomplish the following:
(a) Upon the accumulation of 1,000 hours time-in-service (TIS) or within the next 100 hours TIS, whichever occurs later, inspect the rudder forward spar for cracks in accordance with the instructions in Beech Service Bulletin (SB) No. 2333, Revision 1, dated November 1991.
(b) If no cracks are found, accomplish one of the following:
(1) Reinspect the rudder forward spar for cracks in accordance with the instructions in Beech SB No. 2333, Revision 1, dated November 1991, at intervals not to exceed 500 hours TIS until either paragraph (b)(2), (b)(3), or (b)(4) of this AD is accomplished;
(2) Install Kit No. 33-6001-1 S in accordance with Beech SB No. 2333, Revision 1, dated November 1991;
(3) Install a Spacecraft Machine Products (SMP) rudder spar upper-hinge reinforcement bracket in accordance with Supplemental Type Certificate (STC) SA4899NM; or
(4) Replace the rudder assembly with either part number 33-630000-137, - 139, -141, -167, or -169, as applicable, in accordance with the instructions in Beech SB No. 2333, Revision 1, dated November 1991.
(c) If cracks are found, prior to further flight, accomplish one of the following:
(1) Replace the rudder assembly with either part number 33-630000-137, - 139, -141, -167, or -169, as applicable, in accordance with the instructions in Beech SB No. 2333, Revision 1, dated November 1991;
(2) Install Kit No. 33-6001-1 S in accordance with Beech SB No. 2333, Revision 1, dated November 1991; or
(3) If the cracks are found in the area of the upper hinge, the middle hinge, or both the upper and middle hinge as specified in Beech SB No. 2333, Revision 1, dated November 1991, then stop drill the cracks and install an SMP upper-hinge reinforcement bracket in accordance with STC SA4899NM. For cracks in the middle hinge, install the upper-hinge reinforcement bracket and also install an SMP rudder spar middle-hinge reinforcement bracket in accordance with STC SA5870NM.
(d) If a modification or replacement has been accomplished in accordance with either paragraph (b)(2), (b)(3), (b)(4), (c)(1), (c)(2), or (c)(3) of this AD, then no repetitive inspections are required by this AD.
(e) Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate the airplane to a location where the requirements of this AD can be accomplished.
(f) An alternative method of compliance or adjustment of the initial or repetitive compliance times that provides an equivalent level of safety may be approved by the Manager, Wichita Aircraft Certification Office, FAA, 1801 Airport Road, Mid-Continent Airport, Wichita, Kansas 67209. The request shall be forwarded through an appropriate FAA Maintenance Inspector, who may add comments and then send it to the Manager, Wichita Aircraft Certification Office.
NOTE: Information concerning the existence of approved alternative methods of compliance with this AD, if any, may be obtained from the Wichita Aircraft Certification Office.
(g) The inspections, installations, or replacements required by this AD shall be done in accordance withBeech Service Bulletin No. 2333, Revision 1, dated November 1991. This incorporation by reference was previously approved by the Director of the Federal Register in accordance with 5 U.S.C. 552(a) and 1 CFR Part 51 on August 22, 1992. Copies may be obtained from Beech Aircraft Corporation, P.O. Box 85, Wichita, Kansas 67201-0085. Copies may be inspected at the FAA, Central Region, Office of the Assistant Chief Counsel, Room 1558, 601 E. 12th Street, Kansas City, Missouri, or at the Office of the Federal Register, 800 North Capitol Street, NW., suite 700, Washington, DC.
(h) This amendment (39-8752) supersedes AD 92-15-06, Amendment 39-8300 which superseded AD 91-23-07, Amendment 39-8076.
(i) This amendment (39-8752) becomes effective on January 21, 1994.
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85-23-01: 85-23-01 LOCKHEED-CALIFORNIA COMPANY: Amendment 39-5155. Applies to Lockheed Model L-1011-385 series airplanes, certificated in any category. Compliance required as indicated, unless previously accomplished. To detect inadequate holding torque on slat asymmetry brakes installed on both LH and RH wings and to take corrective action, accomplish the following:
A. On aircraft having P/N 672157-115 or prior part number brakes:
(1) Conduct a functional test of each P/N 672157-115 or prior part number slat control system brake assembly within 300 flight hours after the effective date of this AD, or 3000 flight hours from the last test, whichever is later. Thereafter, conduct the same test at intervals not to exceed 3000 flight hours from the last test.
(2) If the left and right slat asymmetry brake holding torque measures 1500 inch-pounds or more, but less than 2400 inch-pounds, the airplane may continue in service.
(3) If the left or right slat asymmetry brake holding torque measures less than 1500 inch-pounds, but more than 1300 inch-pounds, the aircraft may be flown in that condition for the time necessary to replace the brake assembly, but not more than 500 flight hours from the time the deficient brake was found.
(4) If the left or right slat asymmetry brake holding torque measures less than 1300 inch-pounds, prior to further flight, replace the brake assembly with one having a holding torque of 1700 inch-pounds or more, but less than 2400 inch-pounds.
B. The following alternate procedure may be applied to P/N 672157-115 or prior part number brakes:
(1) Conduct a functional test within 300 flight hours after the effective date of this AD, or 4000 flight hours from the last test, whichever is later, if the last test substantiated that the brake holding torque was at least 1700 inch-pounds. Thereafter, conduct the same test at intervals not to exceed 4000 flight hours from the last test.
(2) If the left and right asymmetry brake holding torque measures 1700 inch- pounds or more, but less than 2400 inch-pounds, the airplane may continue in service.
(3) If the left or right slat asymmetry brake holding torque measures less than 1700 inch-pounds, but more than 1300 inch-pounds, the aircraft may be flown in that condition for the time necessary to replace the brake assembly, but not more than 500 flight hours from the time the deficient brake was found.
(4) If the left or right slat asymmetry brake holding torque measures less than 1300 inch-pounds, prior to further flight, replace the brake assembly with one having a holding torque of 1700 inch-pounds or more, but less than 2400 inch-pounds.
C. Replace P/N 672157-113 or prior part number brakes with P/N 672157-115 or later part number brakes within one year after the effective date of this AD.
D. On aircraft having P/N 671157-117 (with "lip seal" instead of "O-ring" seal) or later part number brakes:
(1) Conduct a functional test of each P/N 672157-117 or later part number brake assembly within 300 flight hours after the effective date of this AD, or 4000 flight hours from the last test, whichever is later. Thereafter, conduct the same test at intervals not to exceed 4000 flight hours from the last test.
(2) If the left and right asymmetry brake holding torque measures 1500 inch- pounds or more, but less than 2400 inch-pounds, the airplane may continue in service.
(3) If the left or right slat asymmetry brake holding torque measures less than 1500 inch-pounds, but more than 1300 inch-pounds, the aircraft may be flown in that condition for the time necessary to replace the brake assembly, but not more than 500 flight hours from the time the deficient brake was found.
(4) If the left or right slat asymmetry brake holding torque measures less than 1300 inch-pounds, prior to further flight, replace the brake assembly with one having 1700 inch- pounds or more, but less than 2400 inch-poundsof holding torque.
E. Alternative means of compliance with this AD which provide an acceptable level of safety may be used when approved by the Manager, Los Angeles Aircraft Certification Office, FAA, Northwest Mountain Region.
F. Special flight permits may be issued in accordance with FAR 21.197 and 21.199 to operate airplanes to a base for accomplishment of tests or modifications required by this AD.
NOTE: Instructions to accomplish these tests and modifications are found in these Lockheed Service Bulletins:
(1) S/B 093-27-216, Revision 2, dated July 19, 1983, or later when approved by the Manager, Los Angeles Aircraft Certification Office, FAA, Northwest Mountain Region.
(2) S/B 093-27-269, dated July 19, 1983, or later when approved by the Manager, Los Angeles Aircraft Certification Office, FAA, Northwest Mountain Region.
(3) S/B 093-27-274, dated July 19, 1983, or later when approved by the Manager, Los Angeles Aircraft Certification Office, FAA,Northwest Mountain Region.
All persons affected by this directive who have not already received these documents from the manufacturer may obtain copies upon request to the Lockheed-California Company, P.O. Box 551, Burbank, California 91520, Attention: Commercial Support Contracts, Dept. 63-11, U-33, B- 1. These documents also may be examined at the FAA, Northwest Mountain Region, 17900 Pacific Highway South, Seattle, Washington, or at 4344 Donald Douglas Drive, Long Beach, California.
This amendment becomes effective November 24, 1985.
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2009-26-11: We are adopting a new airworthiness directive (AD) to supersede AD (AD) 2006-07-15, which applies to Thrush Aircraft, Inc. Model 600 S2D and S2R (S-2R) series airplanes (type certificate previously held by Quality Aerospace, Inc. and Ayres Corporation). AD 2006-07-15 currently requires repetitive inspections of the 1/4-inch and 5/16-inch bolt hole areas on the wing front lower spar caps for fatigue cracking; replacement or repair of any wing front lower spar cap where fatigue cracks are found; and reporting of any fatigue cracks found to the FAA. AD 2006-07-15 also puts the affected airplanes into groups for compliance time and applicability purposes. Since we issued AD 2006-07-15, FAA analysis reveals that inspections are not detecting all existing cracks and shows the incidences of undetected cracks will increase as the airplanes age. Consequently, this AD retains the actions of AD 2006-07-15 and imposes a life limit on the wing front lower spar caps that requires replacement of the wing front lower spar caps when the life limit is reached. This AD also changes the requirements and applicability of the groups discussed above and removes the ultrasonic inspection method. We are issuing this AD to prevent wing front lower spar cap failure caused by undetected fatigue cracks. Such failure could result in loss of a wing in flight.
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74-12-02: 74-12-02 BEECH: Amendment 39-1857. Applies to Models 36 and A36 (Serial Numbers E-1 through E-434); Models 58 and 58A (Serial Numbers TH-1 through TH-335); Models 60 and A60 (Serial Numbers P-4 through P-240) aircraft.
Compliance: Required as indicated, unless already accomplished.
To provide the required structural level of occupant protection during minor crash conditions for those aircraft equipped with optional, aft-facing seats, within the next 100 hours' time in service after the effective date of this AD, accomplish the following:
Install two improved seat back supports on each aft-facing seat in accordance with Beechcraft Service Instruction No. 0591-314, Rev. 1, or later approved revisions or any equivalent approved by the Chief, Engineering and Manufacturing Branch, FAA, Central Region.
NOTE: Seat back supports, Beech P/N 58-530193-1 are applicable to Models 36, A36, 58, 58A, and 60 aircraft, and seat back supports, Beech P/N 58-530193-3 are applicable to Model A60 aircraft.
This amendment becomes effective June 5, 1974.
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