Fixed Wing

BEECH 1900D
SDR # 20081016011
 
Rudder Torque Tube – Failure

SDR submitted:

During a functional check, the Aircraft Maintenance Engineer found that the L/H torque tube was broken just inboard of the left rudder pedal block. Disassembly revealed that the rudder torque tube failure started at the taper pin hole that is used to secure the rudder pedal assembly block to the shaft. Further inspection revealed a secondary crack at about 90 degrees. A third crack was found parallel to the initial crack. The failure appears to be from an overload condition.

It is difficult to detect a failure of this type as the initial failure takes place in an area not visible unless the pedal assembly block is dismantled. Once the crack gets large enough to compromise the torsional ability of the shaft, complete failure would likely occur. The entire rudder and steering system was inspected to determine if other related systems were a factor in this event. No fault found.

Preventive actions taken; a one time check of the rudder pedal assembly for the next routine inspection for the 1900 fleet.

Transport Canada Comments:
Because the pedal block is not disassembled unless necessary, it is recommended that this area be closely examined for defects around the pedal block/torque tube interface. A physical check for any “rotational play” of the torque tube may give an indication of impending failure. Hammer and wrench in X formation indicating end of article

Rudder Torque Tube – Failure

BOMBARDIER, CL600 2C10 (RJ700)
SDR # 20081122004
 
Pitch Trim Actuator - Upper Attachment Loose Bolts

SDR submitted:

During a routine maintenance check in the translating nut area of the horizontal stab trim actuator (HSTA), two bolts were found loose. These bolts, which are positioned at the upper L/H & lower R/H bolt location, connect the upper half of the pitch trim actuator assembly (Item 33) to the aeroplanes structure. See pictorial below. (Reference CRJ700/900 AMM Figure 27-42-01, Items 24 & 35)

Transport Canada Comments:
The Type Certificate Holder (TCH) has taken corrective action by revising the AMM TR 20-022 regarding the general tasks for torquing fasteners. Additionally, a revision was made to the AMM TR 27-0345 and TR 27-0346 to address the installation of the trim actuator. Manual Change Requirement (MCR) Chapter 20 will now require paint strip marks on the bolt and the nut. The TCH now advises replacement of the nuts after each use.

The 2000 Flight Hour oil level check on the HSTA may be an opportune time to do a precautionary one-time nut torque check. Hammer and wrench in X formation indicating end of article

Pitch Trim Actuator - Upper Attachment Loose Bolts | STRUCTURE | POSITIONING TOOL

CESSNA 172N
SDR # 20090209003
 
Assymetrical (Split) Flap Condition – Flap Rollers

SDR submitted:

Shortly after take-off, the R/H flap retract flap cable failed. This was visually confirmed by the pilot. The consequences of the cable failure caused the L/H flap (slave flap) to fall into the fully extended position with the R/H flap (master flap) proceeding to go to the selected flap retracted position. Fortunately, the pilot was able to gain control of the aeroplane.

Initial investigation revealed that the R/H cable was completely severed in the area where the cable rolls over the central flap pulley cluster located in the middle of the cabin behind the headliner (FS 65.33). The subject flap cable was also found frayed in several other areas along the complete flap cable run.

Further inspection found that the L/H aft outboard flap roller (Part Number 0523921) failed due to a “flat spot” on the roller and would not roll thus locking the flap during the retraction cycle. With the slave flap locked in the flap track, the flap motor continued to tension the flap retract cable until the cable failed. The flap motor and transmission can create extremely high torque/tension, which exceeds the tensile strength of the flap cable, if the flap is in a “jammed” condition. The root cause of this event appears to be due to the failure of the flap roller with the flap cable failure as a subsequent event.

FAA AD 80-06-03 and the recommendations of the manufacturer defined in Cessna SEB95-3 had been complied with by this operator.

Transport Canada Comments:
Transport Canada Civil Aviation wishes to advise owners/operators to closely inspect the flap rollers for cracks or flat spots. Premature fraying of flap cables may also be an indication of an impending flap roller failure. Hammer and wrench in X formation indicating end of article

Assymetrical (Split) Flap Condition – Flap Rollers

CESSNA 208B
SDR # 20090122010
 
Upper Cargo Door

SDR submitted:

Enroute to destination at 9000 (AGL) feet altitude, the upper half of the rear cargo door opened in flight with a door warning light illumination. The pilot lowered the flaps and reduced the airspeed to 100 knots as per Airplane Flight Manual (AFM) and carried out an uneventful landing at the nearest airfield. The crew visually inspected the cargo door and the door locking mechanism but found no apparent defects or damage. The crew then secured the door and continued on to their final destination.

At home base, maintenance personnel inspected the cargo door and the upper hinge assembly but found no defects. The inside paneling of the door was removed for a detailed inspection of the door latching mechanism but no faults were found. It was then concluded that the door plunger assembly Part Number (P/N) K5SN may not have been fully engaged in its detent, which allowed the outer door handle P/N 2617094-16 to rotate and come open during flight.

Transport Canada Comments:
The Cessna 208 series aeroplanes can have several different configurations of passenger/cargo/freight doors. Always take the time to ensure that all external doors are secured prior to flight. It appears that human factors may have been a factor in this event. Hammer and wrench in X formation indicating end of article

CESSNA U206G
SDR # 20081219003
 
Horizontal Stabilizer Attachment Doubler

SDR submitted:

During a 100 hour inspection, a crack was discovered at the Fuselage Station (FS) 209 bulkhead on the right stabilizer attachment doubler. Further investigation revealed a similar crack on the left side. Maintenance personnel removed the L/H and R/H bulkhead and stabilizer attachment doublers and found another crack at the second rivet hole located on the R/H stabilizer attachment doubler. An additional crack was also found in the upper right corner radius of the bulkhead. The failed parts have been sent to Cessna.

Cessna Service Bulletin (SB) SEB88-3 provides for a modification that incorporates additional structural members on the horizontal stabilizer attachment bulkhead (FS 209). This particular aeroplane had already complied with the SEB.

Cessna – Continuing Airworthiness Inspection Program (CAP) 53-40-04 “Fuselage Vertical Fin Attachments” recommends a visual and dye penetrant inspection of the bulkhead at 12 000 hours and repetitive inspections at each 2000 hours thereafter. The operator stated that the Cessna CAP inspections are not part of the operators’ maintenance schedule approval, but perhaps should be implemented by all operators.

Transport Canada Comments:
Be advised that Cessna Service Kits SK 210-125 “reinforcement doublers to stab attach fittings” and SK 210-126 “improved stabilizer attach fittings on the front spar” are available from Cessna. Transport Canada Civil Aviation (TCCA) recommends that owners and operators comply with all of the manufacturer’s maintenance publications and instructions.

TCCA has raised these problems with the Federal Aviation Administration who are currently investigating with the type certificate holder, Cessna. Hammer and wrench in X formation indicating end of article

DHC 2 (BEAVER)
SDR # 20080611005
 
Wing Angle Modification

SDR submitted:

While carrying out the annual inspection of a DHC 2 aeroplane that had the Wing Angle Modification installed, some movement was found at the horizontal stabilizer forward attachment points. The attachment brackets were removed for inspection and it was found that the fuselage fitting was ”wearing” into the horizontal stabilizer attachment fitting (Part Number BSFS-159/160/197). It was later confirmed that a fretting/corrosion event had occurred on the Tailplane Brackets.

The STC holder has revised the Supplemental Type Certificate (STC) and introduced a new bushing (.125 to .15) for the Tailplane Bracket. The new bushing length is now the same dimension as the tailplane bracket bushing. This should limit any further fretting and/or corrosion events. Additionally, the STC holder has issued Stolarius Information Bulletin STOL-IB#01 dated 20 January 2009 addressing this specific issue and the method of inspection.

Transport Canada Comments:
Numerous Beaver operators have installed this modification; therefore it is recommended that operators carefully inspect the tailplane attachment brackets as per the above-mentioned Stolarius Information Bulletin. Additionally, an opportune time to perform the inspection would be at the AN174H-12A attachment bolt replacement or during the AD CF-1991-42R1 Front Spar Centre Web Cracks Inspection. Hammer and wrench in X formation indicating end of article

DIAMOND DA20 C1
SDR # 20100727004
 
Rudder Support Bracket - Cracks

SDR submitted:

Several operators have reported via SDRs, cracks on the lower rudder hinge support area (rudder tower); especially at the corners in the weld area. The lower part of the rudder is attached to the rudder lower mounting support bracket. Additionally, a service history review has revealed numerous other foreign and domestic defect reports of this nature.

Diamond Aircraft Limited has investigated the above concerns and confirmed that even after the failure of one of the welds; the support tower is more that capable of supporting the full limit load. Nevertheless, Diamond will soon introduce an improved design that encompasses the addition of gussets at the four corners of the rudder support tower. The addition of gussets will provide added support and will soon be incorporated onto new production aeroplane.

Transport Canada Comments:
This area is currently inspected each 100 hours using a 10-powered magnifying visual aid. Transport Canada Civil Aviation recommends that maintenance personnel inspect at each opportunity and to also ensure that rudder cable tensions are within the manufacturers’ tolerance. Hammer and wrench in X formation indicating end of article

EMBRAER, EMB 135LR
SDR # 20110120004
 
Emergency Door Obstruction

SDR submitted:

During a standard maintenance visit, the left hand over-wing emergency exit door was found to be obstructed and unable to be opened.

Further investigation revealed that a cabin lining attaching screw was incorrectly installed, causing interference with the emergency door frame structure, making it unable to be opened.

The screw was correctly oriented and installed, enabling correct door operation.

Transport Canada Comments:
It is unknown as to how long this condition existed or how it occurred.

Transport Canada Civil Aviation would like to emphasize the importance to all maintenance personnel for correct operations of all emergency door exits.

Whenever work is done on or around an emergency exit door, it is essential that an operational test of the door be carried-out to confirm no obstructions exist. Hammer and wrench in X formation indicating end of article

Emergency Door Obstruction | Point of contact

LEARJET, 35A
SDR # 20101108004
 
Cracked Nose Landing Gear Bracket

SDR submitted:

During an inspection of the nose landing gear, a crack was detected extending from the grease nipple on a bracket assembly. The bracket provided the interconnection of the nose gear retraction/extension actuator to the nose landing gear strut assembly.

The bracket was replaced in accordance with the applicable Aircraft Maintenance Manual (AMM) procedures and the aeroplane was made serviceable.

Transport Canada Comments:
The Illustrated Parts Catalog (IPC) provides the requirement to upgrade to a different part number replacement bracket where the grease nipple is threaded into the bracket assembly instead of being pressed in, as was the case for the failed part.

Transport Canada Civil Aviation would like to advise all operators to be aware of this possible crack propagation scenario and of the available improved replacement part. Hammer and wrench in X formation indicating end of article

Cracked Nose Landing Gear Bracket | Crack propagation

PILATUS - SW, PC 12 47E
SDR # 20101025001
 
Oil Cooler Door Cracks

SDR submitted:

During a 100 hour inspection, it was discovered that the Thermostatic Oil Cooler Door Panel was cracked on both ends including a sheared rivet on the R/H end.

A new door was installed and rigged per Pilatus instructions and the aeroplane made serviceable.

Transport Canada Comments:
Due to the high occurrence rate for this type of fault, Pilatus Continuing Airworthiness has been advised.

The cause of the failures is presently unknown but it is suspected that it may be due to a lack of material robustness.

Transport Canada Civil Aviation would like to advise all operators of this possible condition. Hammer and wrench in X formation indicating end of article

Oil Cooler Door Cracks

PIPER PA-46 SERIES
SDR # 20100809010
 
Engine Mount & Nose Landing Gear Interface - Cracks/Corrosion

SDR submitted:

There have been many SDR reports of engine mount cracks being found at the area of the engine mount/nose landing gear attachment interface. The PA 46 series aeroplane has two styles of engine mounts; the old style engine mount having a two-piece welded foot and the newer style engine mount have a one-piece machined foot (no weld).

A domestic repair station has received and repaired over 30 cracked engine mounts, with almost all cracks/corrosion being found on the older style (two-piece mount) foot. Additionally, the FAA has knowledge of over 49 reports of damaged engine mounts that can potentially result in the collapse of the nose landing gear.

Piper has recently published Service Bulletin (SB) 1103D (dated 2 February 2011) to add more PA46 (Malibu/Mirage/Matrix) models. Please note that SB 1103C states that any cracked engine mount must be changed before further flight.

Additionally, Piper SB 1154C (January 2008) reduces the engine mount inspection interval from 100 Hours to 50 Hours while introducing the newer style one-piece engine mount that replaces the older style mount.

Additionally, Piper SB 1103B (November 2003) recommends that the nose gear actuator mounting bolt be inspected for sufficient thread engagement; if the engine mount has been replaced because of cracks.

Piper Service Letter (SL) 1001 was issued some years ago (December 1987) to exam the cluster weld area of the engine mount near the R/H landing gear mount point.

Transport Canada Comments:
Engine mount cracks typically occur when the landing gear is subjected to excessive loads such as hard landings, operating on rough ground surfaces, excessive speed turns during taxi and/or towing operations. Engine mount welds are particularly vulnerable to damage by cracking, largely because of the vibrations to which they are subjected.

The FAA have also issued SAIB CE-09-13 (February 2009) on the subject of certain PA 46 series aeroplane having cracks in the engine mount where both the nose landing gear trunnion and the nose landing gear attachment area.

Transport Canada Civil Aviation recommends that owners/operators comply with both the Piper Service Information and the FAA SAIB. Hammer and wrench in X formation indicating end of article

Date modified: