MAVERICK HELICOPTER INC

Operated by Maverick Helicopters Inc. (per NTSB report)

AFTER COMPLETION OF A .6 HOUR FLIGHT AND WHILE HOVER TAXIING TO HELIPAD, THE CENTRAL WARNING PANEL ENGINE CHIP LIGHT ILLUMINATED. PILOT IMMEDIATELY DIVERTED TO THE CLOSEST ALTERNATE DESIGNATED HELIPAD AND LANDED, UNDER NORMAL ENGINE POWER AND A NORMAL AIRCRAFT SHUTDOWN PERFORMED. AN EXAMINATION OF THE ENGINE CHIP DETECTORS REVEALED THAT THE METAL CHIPS GENERATED, PREDOMINATELY ON THE MODULE 5 AND ELECTRIC DETECTOR, EXCEEDED THE ALLOWABLE LIMITS OF THE MM. THE ENGINE WAS REMOVED. TRACED THE LIKELY SOURCE OF THE ENGINE CHIPS TO THE MODULE 4 REAR BEARING.

DURING ROUTINE INSPECTION, A HYDRAULIC FLUID LEAK WAS NOTED ON THE BELT DRIVEN HYDRAULIC PUMP DRIVE AND A PREVIOUSLY REPAIRED YELLOW TAGGED PUMP ASSEMBLY INSTALLED. THE REPLACEMENT PUMP DRIVE SHAFT LIP SEAL HAD BEEN REPLACED AS PART OF THE REPAIR. A NORMAL GROUND RUN AND LEAK CHECK WAS PERFORMED WITH NO DISCREPANCY NOTED. DURING THE OPERATIONAL CHECK FLIGHT THE CREW NOTED THE ILLUMINATION OF THE HYDRAULIC PRESSURE LOSS SEGMENT ON THE CAUTION WARNING PANEL AND PERFORMED AN IMMEDIATE PRECAUTIONARY LANDING. FOLLOWING AIRCRAFT SHUTDOWN AN EXAMINATION OF THE DUAL HYDRAULIC SYSTEM WAS PERFORMED. IT WAS EVIDENT THAT THE BELT DRIVEN HYDRAULIC PUMP INTERNAL “O” RING SEALING OR DRIVE SHAFT LIP SEAL HAD FAILED DURING OPERATION AND THAT THE BELT DRIVEN PORTION OF THE DUAL HYDRAULIC SYSTEM HAD SUFFERED A TOTAL LOSS OF HYDRAULIC FLUID. THE SECONDARY PORTION OF THE DUAL HYDRAULIC SYSTEM, THE MAIN GEAR BOX DRIVEN PUMP AND IT’S FLUID RESERVOIR WERE FOUND IN SERVICEABLE AND FULLY SERVICED CONDITION. ALL FLUID RESIDUE WAS CLEANED AS REQUIRED. A SERVICEABLE REPLACEMENT BELT DRIVEN HYDRAULIC PUMP WAS INSTALLED AND THE HYDRAULIC FLUID RESERVOIR SERVICED. GROUND RUN AND LEAK CHECK WAS SATISFACTORY AND THE AIRCRAFT OPERATIONAL CHECK FLIGHT COMPLETED WITH THE AIRCRAFT RETURNED TO SERVICE. THE DEFECTIVE PUMP WAS RETURNED TO THE SHOP FOR DETERMINATION OF LEAK FAILURE ROOT CAUSE. UP TO THIS DATE SEVERAL DOZEN ROUTINE DRIVE SHAFT LIP SEAL REPLACEMENTS WERE COMPLETED PER THE COMPONENT MAINTENANCE MANUAL PREVIOUSLY ON THIS MODEL HYDRAULIC PUMP IN OUR SHOP WITH ONLY MINOR LEAKAGE NOTED DUE TO SHAFT SEAL RUNNING FACE DETERIORATION. THESE PUMPS, WHEN INSTALLED AND FOUND TO BE LEAKING AFTER SEAL REPLACEMENT, ARE ROUTINELY DISCARDED AS BEYOND ECONOMICAL REPAIR. NO LEAK OF THIS MAGNITUDE HAS OCCURRED PREVIOUSLY IN OUR EXPERIENCE WITH THESE PUMPS.

DURING CRUISE FLIGHT THE CENTRAL WARNING PANEL, ENGINE CHIP, SEGMENT ILLUMINATED. PILOT PROCEEDED TO PERFORM A NORMAL POWERED PRECAUTIONARY LANDING. MAINTENANCE PERSONNEL DISPATCHED TO THE ACFT DETERMINED METAL FOUND ON THE AIRCRAFT'S ENGINE ELECTRIC CHIP DETECTOR EXCEEDED MFG LIMITATIONS FOR AIRWORTHINESS. AIRCRAFT TO BE TRANSPORTED TO MAINTENANCE BASE FOR ENGINE REPLACEMENT AND ENGINE RETURN TO MFG FOR EVALUATION AND REPAIR.

DURING CRUISE FLIGHT PILOT OBSERVED AN ENG CHIP LIGHT INDICATION ON THE CENTRAL WARNING PANEL & IAW FLIGHT MANUAL MANDATED PROCEDURES COMPLETED A ROUTINE LANDING AT THE FIRST SUITABLE LANDING SITE. THE HELICOPTER LANDED ON A SECONDARY DIRT ROAD. UPON EXAMINATION BY A CERTIFICATED AIRFRAME & POWER PLANT TECHNICIAN THE CHIP PLUG METAL DEBRIS WAS FOUND SIGNIFICANT ENOUGH TO BE OF AN UN-AIRWORTHY NATURE & THE ACFT GNDED. ENG WAS REPLACED WITH A NEWLY OVERHAULED ENG ASSEMBLY & THE ACFT RETURNED TO SERVICE. ENG SHOP WAS ABLE TO TRACE THE METAL GENERATION TO A BEARING OF THE POWER TURBINE MODULE 4 & MODULE RETURNED FOR REPAIR.

DURING CRUISE, . PILOT DETECTED AN ELECTRICAL ODOR, SLIGHT & DEVELOPED TO A STRONG ODOR WITH SOME MODERATE SMOKE DEVELOPMENT. FOLLOWED FLIGHT MANUAL EMERGENCY PROCEDURES, PILOT USED EMERGENCY ELECTRICAL SYS BREAKAWAY WIRED CUTOUT SWITCH TO CUT OFF ALL ELECTRICAL POWER TO ACFT. RESIDUAL SMOKE AND ODOR CLEARED. ELECTED TO LAND. EXAMINATION OF THE CONSOLE MOUNTED 30 ALPHA SWITCH PANEL CONTROL SWITCH, IDENTIFIED AS SMOKE SOURCE. DETERMINED THE ADJACENT STROBE CONTROL SWITCH HAD VISIBLE HEAT DISCOLORATION AND A SINGLE BASE PIN FOR THE SWITCH FOUND HEAT WELDED TO IT’S 30 ALPHA PCB BOARD SOCKET. A SERVICEABLE 30 ALPHA PANEL WITH SERVICEABLE SWITCHES WAS INSTALLED AND THE ACFT RETURNED TO THE MX BASE FOLLOWING A SATISFACTORY OPS CHECK. DETAIL INSPECTION OF THE DEFECTIVE SWITCH INTERNAL PCB, ONCE SWITCH WAS REMOVED, LESS THE WELDED PIN STUCK IN ITS SOCKET, MADE IT APPEAR TO BE THE LIKELY SOURCE OF THE FAILURE WITH SUBSEQUENT HEAT DAMAGE TO THE 30 ALPHA PANEL CIRCUIT BOARD.

ON SHORT FINAL TO LANDING, ACFT CAUTION WARNING PANEL ENGINE CHIP LIGHT ILLUMINATED. ENGINE WAS SHUTDOWN NORMALLY, INSPECTION OF THE CHIP DETECTORS, METAL WAS FOUND THAT APPEARED TO BE MORE THAN TYPICAL. CLEANED THE CHIP PLUGS, A 10 MINUTE PENALTY GROUND RUN WAS PERFORMED. RECHECKED CHIP PLUGS ON SHUTDOWN REVEALED NO UNUSUAL METAL GENERATION. FOLLOWING A SEVERAL MINUTE HOVER CHECK, WITH NO FURTHER CHIP LIGHT ILLUMINATION, ACFT DEPARTED FOR THE 10 MINUTE FLIGHT LEG TO OUR REMOTE FUELING STATION. ON SHORT FINAL, ENGINE CHIP LIGHT ILLUMINATED AGAIN. A NORMAL LANDING AND ENGINE SHUTDOWN WAS PERFORMED. ON FURTHER INSPECTION OF CHIP DETECTORS, DETERMINED THAT THE METAL CHIPS GENERATED EXCEEDED THE ALLOWABLE MM SPECS. ENGINE REMOVED AND MODULES RETURNED TO VENDOR. TEARDOWN INSPECTION BY THE VENDOR, DETERMINED THAT THE AFT BEARING OF THE POWER TURBINE/MODULE 4 HAD FAILED. METAL GENERATED, ALSO CONTAMINATED THE MODULE 5.

HYDRAULIC PRESSURE LIGHT ILLUMINATED IN FLIGHT. RETURNED TO BASE. DURING INSPECTION, DETERMINED THE DRIVE BELT HAD DISENGAGED FROM THE HYRAULIC PUMP DRIVE PULLEY. EXAMINATION REVEALED THAT THERE HAD BEEN A LOSS OF INTERFERENCE FIT BETWEEN THE PULLY AND THE PULLEY'S BEARING INNER RACE WHICH ALLOWED THE PULLY TO OPERATE WITHOUT BEARING ACTION. DRIVE PULLY TO BEARING INNER RACE CIRCLIP CAME UNDONE ALLOWING ENOUGH PLAY IN THE ASSY THAT THE PULLEY WOBBLED ENOUGH TO THROW THE BELT. BELT FOUND UNDAMAGED. SERVICEABLE DRIVE PULLEY ASSY WITH PUMP WAS INSTALLED AND RETURNED TO SERVICE.

DURING A ROUTINE REVENUE FLIGHT, PILOT NOTED THAT THE ACFT VEMD HAD A YELLOW BAR WARNING UNDER THE VOLTAGE INDICATION, INDICATING LOW VOLTAGE OF 24 VOLTS. ON RESET, AN INDICATION OF 250 AMPS WITH THE RED BAR WARNING. FOUND SHORTED GENERATOR WINDINGS AND AN ELECTRICAL ODOR.

ACFT EXPERIENCED A FORCED LANDING. ON THE FIRST SCHEDULED STRIP TOUR AT DUSK, THE PILOT SUDDENLY NOTED SOME UNUSUAL MODERATE LATERAL VIBRATIONS IN THE AIRFRAME. PILOT IMMEDIATELY DIVERTED BACK TO THE AIRPORT AND JUST SHORT OF FLARING THE ACFT INTO THE HOVER/TAXI FLIGHT MODE, HAD BOTH RED AND AMBER GOVERNOR LIGHTS ILLUMINATE ON THE CAUTION WARNING PANEL. PILOT EXECUTED AN IMMEDIATE AND UNEVENTFUL RUN ON LANDING. ACFT TOWED TO THE HANGER. PILOT DID INDICATE THAT THE ENGINE NG SPEED WAS FLUCTUATING DURING RETURN, BUT WAS UNSURE IF THAT WAS THE CAUSE OR THE EFFECT OF THE AIRFRAME LATERAL VIBRATIONS. FAILURE CODE INDICATED ON THE VEMD WAS CODE 129 INDICATING A STEPPER MOTOR OR RESOLVER FAILURE. FROM PREVIOUS INCIDENTS WHERE NG INSTABILITY WAS NOTED, SUSPECT HMU, WHICH WAS OF OLDER STYLE DIAPHRAGM, WAS EXPERIENCING A PARTIAL DIAPHRAGM RUPTURE. HMU HAS BEEN RETURNED TO THE MFG FOR EVALUATION AND REPAIR. EXAMINATION OF THE ACFT REVEALED NO DAMAGE AND SUBSEQUENT TO HMU REPLACEMENT AND A NORMAL CHECK FLIGHT RESULT THE ACFT HAS BEEN RETURNED TO SERVICE.

ACFT PREPARED TO LAUNCH FOR REVENUE FLT, AN ENGINE CHIP LIGHT ILLUMINATED. ENGINE SHUTDOWN NORMALLY. CHECK OF ELECTRICAL CHIP DETECTOR REVEALED A FEW SMALL METAL SLIVERS BUT SUBSEQUENT INSP REVEALED A SIGNIFICANT NUMBER OF METAL CHIPS ON THE MODULE 5 NON ELECTRIC CHIP DETECTOR. MODULE 1 NON ELECTRIC CHIP DETECTOR WAS FOUND CLEAN. DISCOVERED THE METAL GENERATION WAS NOT A ROUTINE PROBLEM BUT RATHER THE RESULT OF PARTS COMING LOOSE IN THE MODULE & CAUSING FOD TO OCCUR. . IT IS KNOWN THAT A PLUG HAD COME LOOSE FROM THE PRE TU 156 INTERMEDIATE PINION GEAR CENTRAL BORE AND A RETAINING BOLT HEAD SHEARED OFF ALLOWING 2 WASHERS ALONG WITH THE BOLT HEAD TO DROP INTO THE GEARBOX.

DURING PREFLIGHT INSP, A CRACK WAS DISCOVERED RADIATING FROM THE INSIDE LOWER AFT RT CORNER OF THE TAIL BOOM TAIL CONE CUTOUT FOR THE HORIZ STABILIZER. AN OVER LYING DOUBLER WAS NOT CRACKED AND OBSCURED THE CRACK FROM EASY VIEWING. WITH THE HORIZ STABILIZER REMOVED THE CRACK WAS EASILY VIEWED FROM THE LT SIDE OF THE TAIL BOOM LOOKING ACROSS THE TAIL BOOM INTERIOR. AN APPROVED REPAIR FROM THE MFG IS BEING APPLIED THAT HAS AN EXTENSION ON THE DOUBLER THAT REINFORCES THE TAIL BOOM CUTOUT INSIDE CORNER. THE DOUBLER HAS BEEN REMOVED AND THE TAIL CONE CRACK LENGTH MEASURED JUST OVER 1.5 INCH IN LENGTH. A STOP DRILL HAS BEEN APPLIED TO THE CRACK AND THE DOUBLER MFG AND INSTALLED IAW APPROVED DATA. NEWLY DELIVERED TAIL BOOMS NOW HAVE DOUBLERS EXTENDING BOTH ABOVE AND BELOW THE HORIZONTAL STABILIZER CUT OUTS TO HELP PREVENT CRACK DEVELOPMENT.

DURING INSP A CRACK WAS DISCOVERED RADIATING FROM THE INSIDE LOWER AFT RT CORNER OF THE TAILBOOM TAIL CONE CUTOUT FOR THE HORIZ STABILIZER. AN OVER LYING SKIN DOUBLER WAS NOT CRACKED AND OBSCURED THE CRACK FROM EASY VIEWING. WITH THE HORIZ STABILIZER REMOVED THE CRACK WAS EASILY VIEWED FROM THE LT SIDE OF THE TAILBOOM LOOKING ACROSS THE TAILBOOM INTERIOR. AN APPROVED REPAIR FROM THE MFG IS BEING IMPLEMENTED THAT HAS A LOWER EXTENSION ON THE DOUBLER TO HELP REINFORCE THE TAIL CONE OPENING AND STOP DRILLS THE CRACK. NEWLY DELIVERED TAILBOOMS NOW HAVE DOUBLERS THAT EXTEND BOTH ABOVE AND BELOW THE HORIZ STABILIZER TAIL CONE CUTOUT TO PREVENT CRACK DEVELOPMENT.

DURING PERIODIC INSP, A CRACK WAS DISCOVERED IN THE INSIDE CORNER OF THE TAIL BOOM TAIL CONE DOUBLER. THIS DOUBLER WAS A MFG PART THAT WAS INTENDED AS A REPAIR OF A CRACK IN THE TAIL CONE SHEET METAL PREVIOUSLY REPORTED UNDER UNIQUE CONTROL NUMBER M7KA20101210 NR4545. THE DOUBLER WAS MFG AND INSTALLED IAW MFG APPROVED REPAIR DOCUMENTATION SPECIFIC TO THIS ACFT. AS THE ORIGINAL .75 INCH CRACK IN THE TAIL CONE ENDED AT A RIVET BORE HOLE NO STOP DRILL HAD BEEN PERFORMED. AFTER REMOVAL OF THE CRACKED DOUBLER IT WAS DETERMINED THAT THE UNDER LYING TAIL CONE CRACK HAD CONTINUED PAST THE RIVET BORE HOLE AN ADDITIONAL .75 INCH WHERE IT HAD PREVIOUSLY TERMINATED MIRRORING THE CRACK IN THE DOUBLER. THE REMOVED DOUBLER HAS BEEN RETURNED TO THE MFG ENGINEERING DEPARTMENT FOR ASSESSMENT OF THE REPAIR FAILURE AS TO THE LIKELY CAUSE WHETHER ENGINEERING OR WORKMANSHIP OR BOTH. THE CRACKS STARTING POINTS ARE LOCATED IN THE RADIUS OF THE TAIL BOOM CUT OUT LOWER AFT RT CORNER. IT IS OF NOTE THAT THE RADIUS IS VERY SMALL AND THE DOUBLER MFG INSTRUCTIONS HAD NO RADIUS DIMENSION MINIMUM NOTED. THE DOUBLER INSIDE CORNER WAS FINISHED TO MATCH THE EXISTING TAIL CONE CUT OUT PROFILE AT INSTALLATION. CURRENTLY AWAIT A NEW APPROVED REPAIR FROM MFG, AN ADDITIONAL .75 INCH.

DURING ROUTINE INSP, A VISIBLE RADIATING CRACK IN THE PAINT AROUND THE HEAD OF A COUNTERSUNK RIVET ON THE TAILBOOM CONE TO FENESTRON RING JUNCTION WAS NOTED. FOLLOWING CHEMICAL STRIPPING OF THE PAINT IT WAS APPARENT VISUALLY THAT THE SUSPECT AREA OF THE RING JUNCTION WAS CRACKED. FURTHER 3 ADDITIONAL COUNTERSUNK RIVETS ALSO HAD CRACKS RADIATING CENTERED THROUGH THE RIVET BORES THAT WERE NOT APPARENT BEFORE THE PAINT BEING STRIPPED. ALL 4 POSITIONS WERE LOCATED ADJACENT TO ONE ANOTHER IN A STAGGERED PATTERN AND LOCATED APPROX 1" DIAGONALLY FROM ONE ANOTHER. 2 RIVETS ARE WITHIN .3750" OF THE RING FRAME FLANGE EDGE AND EACH HAVE SINGLE CRACKS RADIATING DIRECTLY TO THE FLANGE EDGE FROM THE RIVET CENTERS. THE OTHER 2 RIVETS ARE LOCATED 1" FROM THE FLANGE EDGE AND EACH HAS 2 CRACKS ALIGNED WITH ONE ANOTHER THROUGH THEIR COUNTERSINK BORE CENTERS THAT TERMINATE ON THE FLANGE SURFACE WITH OVERALL LENGTHS NOT LONGER THAN ONE AND A QUARTER INCH. A SINGLE RIVET WITH CRACKS RADIATING IS THE NORMAL FIND FOR THIS TAIL BOOM TO FENESTRON JUNCTION RING FITTING DEFECT. THERE WAS NO EVIDENCE OF RIVETS WORKING AT CRACK LOCATIONS. A REPLACEMENT TAIL BOOM HAS BEEN ORDERED AS NO FIELD REPAIR IS CURRENTLY APPROVED BY THE MFG. THE MFG IS FULLY AWARE OF CRACKING PROBLEMS IN THIS LOCATION AND HAS NEW APPROVED PRE-EMPTIVE REPAIRS AND DESIGN CHANGES BEING IMPLEMENTED AT THIS TIME.

TAILBOOM CONE SKIN CRACK WAS DISCOVERED DURING ROUTINE PERIODIC INSPECTION IN THE LOWER AFT RT CORNER OF THE CUTOUT FOR THE HORIZ STAB. THE CRACK WAS CONSEALED UNDERNEATH THE TAIL CONE SKIN DOUBLER AND WAS OBSERVED BY VIEWING FROM THE LT SIDE OF THE HORIZONTAL STABILIZER CUTOUT ACROSS THE TAILBOOM CONE INTERIOR WITH THE STABILIZER REMOVED. THIS PROBLEM IS KNOWN TO THE ACFT MFG WHO HAS PROVIDED AN APPROVED REPAIR SCHEME THAT IMPROVES ON THE TAIL CONE DOUBLER OVERLAP IN THIS AREA WITH STOP DRILLING OF THE CRACK. A MAJOR REPAIR 337 IS SUBMITTED FOR THIS REPAIR WHEN COMPLETED.

DURING ROUTINE INSPECTION, A VISIBLE CRACK IN THE PAINT AROUND THE HEAD OF A COUNTERSUNK RIVET ON THE TAIL BOOM CONE TO FENESTRON RING JUNCTION WAS NOTED. FOLLOWING THE STRIPPING OF THE PAINT THROUGH CHEMICAL MEANS AND APPLICATION OF DYE PENETRANT INSPECTION MEASURES AT THE SUSPECT CRACK LOCATION IT WAS DETERMINED THAT A 1 INCH CRACK CENTERED TROUGH THE COUNTERSUNK RIVET BORE EXISTED ON THE RING FLANGE FOR THE FENESTRON ATTACHMENT. THIS DEFECT HAS NO APPROVED FIELD REPAIR AND A REPLACEMENT SERVICEABLE TAILBOOM HAS BEEN ORDERED FROM THE MFG. THE MFG IS FULLY AWARE OF CRACKING PROBLEMS IN THIS LOCATION AND HAS NEW APPROVED REPAIRS AND DESIGN CHANGES BEING IMPLEMENTED AT THIS TIME.

PILOT WAS JUST ENTERING CRUISE FLIGHT ABEAM THE AIRPORT WHEN AN EXISTING PREVIOUSLY REPAIRED STOP DRILLED AND PATCHED CRACK BEGAN PROPAGATING RAPIDLY BEYOND THE STOP DRILL AND BEYOND THE PATCH. A LARGE CRACK FORMED IN A QUESTION MARK SHAPE INTO THE CENTER OF THE PLEXIGLASS WINDSHIELD WITH THE CIRCULAR MOST CRACKED PORTION NEARLY BASKETBALL SIZE. PILOT IMMEDIATELY REDUCED AIRSPEED AND REQUESTED CLEARANCE FROM THE AIRPORT AUTHORITIES TO LAND ON AN AIRPORT RAMP AREA AS DIRECTED. CLEARANCE WAS GRANTED AND A ROUTINE LANDING PERFORMED. IT WAS AN UNSEASONABLY COLD DAY WHICH IS THOUGHT TO BE THE MAJOR CONTRIBUTING FACTOR IN THE CRACK CONTINUING. THE CRACK WAS FOUND TO BE WITHIN THE MAINTENANCE DOCUMENTATIONS LIMITATION FOR REPAIR BY STOP DRILL AND ADHESIVE PATCH APPLICATION WHICH HAD BEEN PERFORMED SOME MONTHS EARLIER. A NEARLY IDENTICAL CRACK AND REPAIR EXISTED ON THE RT WINDSCREEN WHICH REMAINED INTACT DURING THE INCIDENT.

DURING ROUTINE PERIODIC INSPECTION A VISIBLE CRACK IN THE PAINT AROUND A COUNTERSUNK RIVET ON THE TAIL BOOM CONE TO FENESTRON RING JUNCTION WAS NOTED. FOLLOWING THE STRIPPING OF THE PAINT AND THE COMPLETION OF A DYE PENETRANT INSPECTION IT WAS DETERMINED THAT THE MOUNTING FLANGE FOR THE FENESTRON ATTACHMENT HAD A 1.5 INCH CRACK CENTERED THROUGH THE COUNTERSUNK RIVET BORE AT THE LOCATION OF THE CRACKED PAINT. THIS DEFECT HAS NO APPROVED FIELD REPAIR AND A REPLACEMENT TAILBOOM HAS BEEN ORDERED.

TAILBOOM SKIN CRACK WAS DISCOVERED DURING ROUTINE PERIODIC CLEANING AND INSPECTION FOLLOWING REMOVAL OF THE HORIZONTAL STABILIZER DURING UNRELATED DEFECTIVE PART REPLACEMENT. THIS CRACK WAS CONCEALED BENEATH A TAIL CONE DOUBLER WHEN EXTERNALLY VIEWED BUT EASILY SEEN WITH THE STABILIZER REMOVED AND VISIBLE WITH SOME DIFFICULTY IF THE STABILIZER REMAINS INSTALLED. THIS PROBLEM IS KNOWN BY THE MFG WHO HAS PROVIDED A REPAIR SCHEME THAT IMPROVES THE TAIL CONE DOUBLER OVERLAP IN THIS AREA AND STOP DRILLS ANY CRACKS. A MAJOR REPAIR 337 IS SUBMITTED FOR THIS REPAIR AFTER COMPLETION.

DURING REPAIR OF ADJACENT TAILCONE SKIN CRACKS A DETAILED VISUAL INSP OF THE UPPER RT HORIZONTAL STABILIZER SPAR REVEALED A CRACK RADIATING FROM A SINGLE STABILIZER ATTACHMENT FITTING, ATTACHMENT BORE AROUND THE SPAR'S FLANGE RADIUS AND ENDING AT A RIVET BORE UNDER THE TAIL CONE SKIN. OVERALL CRACK LENGTH 1.1 INCH.

TAILBOOM CONE SKIN AND SKIN DOUBLER CRACKS WERE DISCOVERED DURING ROUTINE 100 HOUR INTERVAL IMPS LOCATED ADJACENT TO THE RT, AFT, LWR HORIZONTAL STABILIZER'S TAILBOOM THROUGH MOUNT CUTOUT.

WINDSCREEN FOUND CRACKED DURING PREFLIGHT INSPECTION. A NEW WINDSCREEN WAS INSTALLED. SUGGEST THAT A FORM OF REINFORCEMENT BE ADDED TO THE WINDSCREEN EDGES OR AN ALTERNATE AND MORE FLEXIBLE ADHESIVE SYSTEM BE DEVELOPED TO ALLOW FOR PLEXIGLASS EXPANSION AND CONTRACTION WITH SEASONAL TEMPERATURE CHANGES. COLD WEATHER SEEMS TO BE A CONTRIBUTING FACTOR AS MOST CRACKS DEVELOP BELOW 40 DEGREES FAHRENHEIT. EXTRA CARE IN TRIMMING AND FINISHING OF THE WINDOWS EDGES BEFORE INSTALLATION IS ENCOURAGED.

DURING ROUTINE INSPECTION THE MOST AFT OTBD CORNER OF THE MOUNTING FLANGE ASSEMBLY WHICH INCLUDES THE UPPER LOCKING PIN MECHANISM WAS CRACKED. A NEW ASSEMBLY WAS INSTALLED.

GOVENOR LIGHT ILLUMINATES INTERMITTENTLY WITH THIS EMERGENCY BACKUP CONTROL. ACTUATOR INSTALLED.

DURING ROUTINE INSP, LATCH APPEARED LOOSE AND CLOSER INSPECTION REVEALED LOCKING LATCH HINGE PINS SOCKETS AND HOOK CATCH WORN EXCESSIVELY. INSTALLED NEW LATCHES AS REQUIRED.

DURING ROUTINE INSPECTION 1 OF 3 FREQUENCY ADAPTERS FOUND TO HAVE A CRACK IN THE ELASTOMER NEARING ITS MAXIMUM ALLOWABLE LIMITATION OF 10MM. A NEW KIT OF 3 MATCHED ADAPTERS WERE INSTALLED.

DURING ROUTINE INSPECTION LATCH APPEARED LOOSE AND CLOSER INSPECTION REVEALED LOCKING LATCH HINGE PINS SOCKETS AND HOOK CATCH WORN EXCESSIVELY. HINGE ATTACHMENT BORE HAD A 1` LONG CRACK RADIATING. INSTALLED NEW LATCHES AS REQUIRED.

DURING ROUTINE MX INSPECTION 1 OF THE 3 FRQUENCY ADAPTERS FOUND TO HAVE A CRACK IN THE ELASTOMER NEARING ITS MAXIMUM ALLOWABLE LIMITATION OF 10MM. A NEW KIT OF 3 MATCHED ADAPTERS WERE INSTALLED.

DURING ROUTINE 500 HOUR INSPECTION A CRACK WAS FOUND RADIATING FROM AN ATTACHMENT FITTING SCREW BORE TO THE EDGE OF THE SPAR. A NEW SPAR WAS MATCH DRILLED AND INSTALLED IAW EUROCOPTER REPAIR SCHEME.

HINGE ATTACHMENT BORE HAD A 1` LONG CRACK RADIATING.

DURING ROUTINE INSPECTION LATCH APPEARED LOOSE AND CLOSER INSPECTION REVEALED LOCKING LATCH HINGE PINS SOCKETS AND HOOK CATCH WORN EXCESSIVELY. INSTALLED NEW LATCHES AS REQUIRED.

FAILURE CODE AT SHUTDOWN INDICATED FAULT WITH THE COLLECTIVE POTENTIOMETER. A NEW POTENTIOMETER WAS INSTALLED AND NO FURTHER DEFECT REPORTED. IT WAS RECENTLY DISCOVERED THAT THE ROUTING OF AIR CONDITIONING SYSTEM WIRING IN CONTACT WITH THE UNSHIELDED ELECTRICAL HARNESS FOR THE POTENTIOMETER WAS CAUSING SPURIOUS SIGNALS IN THE CONTROL DATA WHICH WERE INTERPRETED AS A PONTENTIOMETER FAILURE. DOZENS OF THESE PONTENTIOMETERS WERE REPLACED. REROUTING THE AIR CONDITIONING WIRING HAS ALMOST ENTIRELY ELIMINATED THE NEED TO REPLACE THESE POTENTIOMETERS. (K)

POWER CHECK WITH A HIGH TORQUE MARGIN AND OUT OF LIMIT T4 TEMP MARGIN WAS INDICATIVE OF TURBINE BLADE TIP DAMAGE VERIFYING DEFECT. ENGINE WAS REPLACED WITH A SERVICEABLE ASSY. THIS TURBINE RUB DEFECT HAS RECENTLY BECOME MORE PREVALENT ON THESE ENGINES. ANY CONTACT BETWEEN THE SEGMENT FACES AND THE HP TURBINE BLADE TIPS TRANSFERS METAL TO THE SEGMENTS WHICH BUILDS UP INTO RAISED DEPOSITS WHICH EFFECTIVELY MACHINE GROOVES IN THE BLADE TIPS LEADING TO THE LOSS OF THE TEMP T4 MARGIN WITH THE GAS BYPASS INCREASING THE POWER TURBINE TORQUE. IT IS NOT CLEAR IF THE SEGMENTS WARP INTO BLADE CONTACT OR THE BLADES STRETCH TO INITIATE THIS DEFECT. IT IS NOT CLEAR IF THE SEGMENTS WARP INTO BLADE CONTACT OR THE BLADES STRETCH TO INITIATE THIS DEFECT. (K)

DURING START ATTEMPT YELLOW AND AMBER GOVERNOR LIGHTS ILLUMINATED. FAULT CODE INDICATED START/STOP SWITCH AS DEFECTIVE. A NEW SWITCH WAS INSTALLED AND NO FURTHER DEFECT REPORTED. THIS IS A RARE FAILURE FOR THIS PART. (K)

DURING ROUTINE INSPECTION THE LT POSITION IGNITER WOULD NOT FIRE. SWAP OF BOX OUTPUT FEEDS TO THE IGNITERS VERIFIED THE BOX CIRCUIT WAS DEFECTIVE. THIS BOX GAVE GOOD SERVICE BUT SEVERAL OTHER HAVE HAD PROBLEMS WITH LOW SERVICE HOURS THESE BOXES ARE NON-REPAIRABLE. (K)

DURING ROUTINE 600 HOUR INSPECTION THE FUEL FILTER PREBLOCKAGE PRESSURE SENSOR WAS FOUND OUT OF TOLERANCE IAW MM DATA. A NEW SWITCH WAS INSTALLED. THIS IS A RARE FAILURE FOR THIS PART. (K)

POSITION LIGHTS FAILED TO ILLUMINATE WITH CIRCUIT BREAKER IN CLOSED POSITION AND NO APPARENT OPENS OR SHORTS IN THE WIRE HARNESS. INVESTIGATION REVEALED THAT THE PCB TRACE FOR THE POSITION LIGHT SWITCH POWER HAD MELTED MUCH AS A FUSE BLOWS. THIS IS NOT AN UNCOMMON PROBLEM FOR THIS PART. SUGGEST A CLOSE LOOK AT THE CURRENT PATH FOR THIS CIRCUIT TO SEE WHY THE CIRCUIT BREAKER PROTECTION FAILS TO PROTECT THE SWITCH PANEL PERHAPS A MORE ROBUST CIRCUIT TRACE IS REQUIRED ON THE PCB FOR THIS CIRCUIT. (K)

FOUND CRACKED DURING ROUTINE INSPECTION. THIS PART IS ALSO DAMAGED BY UV RAYS OVER TIME MAKING IT BRITTLE AND PRONE TO CHIPPING AND CRACKING. (K)

FOUND CRACKED DURING ROUTINE PERIODIC INSPECTION. INSTALLED NEW BRACKET AS REQUIRED. (K)

DURING ROUTINE 500 HOUR INSPECTION THE PULLEY SPINDLE ON WHICH THE INNER RACE OF THE BEARING IS PRESSFIT WAS FOUND VERY WORN WITH MORE THAN .10 INCH OF THE PROPER SPINDLE DIAMETER WORN AWAY AND SOME CORRESPONDING WEAR ON THE BEARING INNER RACE AND WEAR ON THE HYDRAULIC PUMP DRIVE PULLEYS SPLINED COUPLING SLEEVE. THIS IS THE THIRD INSTANCE OF WHICH WE ARE AWARE OF THIS DEFECT OCCURRING. THE BEARING ITSELF HAD NO APPARENT ROUGHNESS WHEN TURNED BY HAND. SUSPECT SOME PULLEYS OR PERHAPS BEARINGS HAVE IMPROPER MACHINED DIAMETERS WHICH ALLOW THE INITIAL SLIPPAGE AND SUBSEQUENT CHATTERING DEVELOPS WHICH REMOVES MATERIAL OVER TIME. THIS HAPPENED BETWEEN INSPECTIONS. BELT TENSION MUST BE RELEASED TO DETECT THIS DEFECT AS THE BELT AND THE COUPLING SLEEVE HELP KEEP THE PULLEY ALIGNED DURING A CURSORY INSPECTION OR WHEN THE AIRCRAFT IS OPERATING. WE INSPECT FOR THIS EACH 100 HOURS.

ENGINE SMOKING FOLLOWING SHUTDOWN WAS THE OBSERVED SYMPTOM. CLOSER INSPECTION REVEALED TRACES OF OIL STREAKING FROM POWER TURBINE BLADE ROOTS. A SERVICEABLE LABYRINTH RING ASSEMBLY HAS BEEN INSTALLED. (K)

NAV LIGHT ASSY, FOUND INOPERATIVE DURING ROUTINE POST FLIGHT INSPECTION, INSTALLED NEW REPLACEMENT, OPERATIONAL CHECK WAS CORRECT. THIS PART HAS HAD VIBRATION RELATED FAILURES IN OUT FLEET ON SEVERAL OCCASIONS. (K)

FOUND SEVERAL HINGE SEGMENTS CRACKED ON ONE END OF THE HINGE LENGTH DURING ROUTINE INSPECTION. THIS IS FAIRLY COMMON DEFECT. (K)

FOUND PIN FALLING OUT DURING ROUTINE INSPECTION. (K)

FOUND CRACKED DURING ROUTINE INSPECTION. RECOMMEND THIS PART BE MFG OF A MORE DURABLE MATERIAL SUCH AS STAINLESS STEEL. THIS IS A FAIRLY COMMON DEFECT. (K)

FOUND LEAKING EXCESSIVELY BOTH STATICALLY AND IN OPERATION DURING ROUTINE INSPECTION. INSTALLED NEWLY OVERHAULED FREE WHEEL SHAFT ASSY WITH NO FURTHER LEAKAGE NOTED. TEAR DOWN OF REMOVED ASSY REVEALED SEAL LIP HAD BEEN DAMAGED DURING ASSY. AS THE SHAFT WAS REPORTED TO BE INSTALLED FOR TEST CELL ENGINE. RUNS IT SEEMS ODD THE LEAK WAS NOT DISCOVERED AT THE OVERHAUL FACILITY. THIS IS THE SECOND TIME A NEWLY OVERHAULED ENGINE WAS INSTALLED THAT HAD A DAMAGED SEAL WHEN RECEIVED FROM THE FACTORY. (K)

ENGINE FOUND LEAKING FROM POWER SHAFT HOUSING SEAL VICINITY DURING OPERATION. A NEWLY REPAIRED, SERVICEABLE HOUSING WITH A NEW MAGNETIC SEAL EMBODIED WAS INSTALLED WITH NO FURTHER LEAKAGE NOTED. THIS SEAL HAS LEAKAGE ISSUES FROM TIME TO TIME BUT WILL MAKE IT TILL OVERHAUL ALMOST HALF THE TIME. (K)

ENGINE SMOKING FOLLOWING REDUCTION OF POWER TO FLIGHT IDLE WAS THE OBSERVED SYMPTOM. CLOSER INSPECTION REVEALED TRACES OF OIL STREAKING FROM POWER TURBINE BLADE ROOTS. A SERVICEABLE LABYRINTH RING ASSEMBLY HAS BEEN INSTALLED. (K)

AMBER GOVERNOR LIGHT ILLUMINATED IN FLIGHT AND SUBSEQUENT TO FAILURE REPORT GENERATION THE COLLECTIVE POTENTIOMETER WAS INDICATED TO BE DEFECTIVE. A NEW PART WAS INSTALLED AND NO FURTHER DEFECT NOTED. (K)

PILOT REPORTS THAT HE HAD INTERMITTENT LOW ROTOR RPM WARNING SIGNALS WITH NORMAL ROTOR SPEED INDICATIONS. THE MAGNETIC PICKUP HAD A LOW ROTOR WARNING CIRCUIT COB. CONTINUITY PROBLEM. A NEW SENSOR WAS INSTALLED AND NO FURTHER DEFECT WAS REPORTED. SUGGEST A MORE ROBUST SENSOR BE DESIGNED AS THE COIL WIRES ARE HAIR THIN AND FAIL ROUTINELY. PERHAPS EVEN AN OPTIC SENSOR WOULD BE MORE RELIABLE. (K)