CENTER FOR EMERGENCY MEDICINE OF WESTERN PA INC

Operated by CENTER FOR EMERGENCY MEDICINE OF WESTERN (per NTSB report)

DENTS AND LOCALIZED DELAMINATION WERE FOUND ON THE UNDERSIDE SKIN OF THE CABIN FLOOR, MEASURING UP TO 89 MM × 914 MM (3.5 IN × 36 IN) AND 50 MM × 610 MM (2 IN × 24 IN). REPAIRED IN ACCORDANCE WITH AIRBUS HELICOPTERS RDAS RDAS EC135 53 2026 2040A, ISSUE DATE 11 JUN 2026. THE REPAIR INCLUDED REMOVAL OF DAMAGED SKIN, INSPECTION AND REPAIR OF HONEYCOMB CORE AS REQUIRED, AND INSTALLATION OF AN ALUMINUM UNDERSIDE SKIN DOUBLER BONDED AND CURED PER APPROVED DATA. SURFACE PROTECTION AND FINISH WERE RESTORED. REPAIR INSPECTED BY VISUAL AND TAP TEST AND FOUND SATISFACTORY.

DAMAGE WAS NOTED TO THE RH KEEL BEAM CAP (P/N: L533M1010203) HORIZONTAL FLANGE IN THE FORM OF SEVERAL AREAS OF CORROSION. THERE ARE REPORTED TO BE 3EA AREAS OF CORROSION IN THE RH KEEL BEAM CAP. A DETAILED DESCRIPTION OF THE DAMAGE AND APPROVED REPAIR PROCEDURES ARE PROVIDED IN AIRBUS REPAIR DESIGN APPROVAL SHEET RDAS-EC135-53-2026-1925, ISSUE A.

POST REMOVING DAMAGED TEFLON LINER ON BLADE. DISCOVERED CRACK ON WEDGE ON WEDGE OF UPPER TRAILING BUSH HOLE TO OUTER ROVING LOOP.

DURING 500-HOUR INSPECTION, A DENT WAS DISCOVERED ON THE R/H SIDE OF THE TRANSMISSION DECK (P/N: L533M1017295). THE MAXIMUM DEPTH IS APPROXIMATELY 0.022" WITH A MAXIMUM DIAMETER OF 0.8. AN APPROXIMATE LOCATION OF THE DAMAGE IS AT STA. 4600MM, W.L. 2465MM, AND R.B.L. 50MM. A DETAILED DESCRIPTION OF THE DAMAGE REPAIR AND APPROVED REPAIR PROCEDURES ARE PROVIDED IN AIR METHODS ENGINEERING AUTHORIZATION 23-5300-EA-5655, INITIAL REVISION DATED 18 MAY 2026 AND IS SUBSTANTIATED BY FAA FORM 8110-3,

REAR POSITION LIGHT SCREWS LOST TORQUE WITH SUBSEQUENT LOSS OF POSITION LIGHT ASSY. INFLIGHT.

MAIN ROTOR COLLECTIVE ACTUATOR P/N 1646A0000-01 , S/N 0018 COMPONENT TOTAL TIME 26,400.00 HOURS , TSO 0.0 WAS INSTALLED AT STAT MED EVAC FIXED BASE OPERATION ON AIRCRAFT S/N 1042 AT ACTT : 7602.5. PRIOR TO COMPLETION OF AIRCRAFT MAINTENANCE AND MAINTENANCE GROUND RUNS THE STAT RECORDS DEPARTMENT DISCOVERED THAT THE ACTUATORS HISTORICAL RECORDS SHOWED NO EVIDENCE OF THE AIRBUS MANDATED SERVICE LIFE LIMIT OF 20,000 HOURS BEING COMPLIED WITH. AIRBUS TECHNICAL SERVICES WERE NOTIFIED. AIRBUS WAS UNABLE TO CONFIRM IF THE HOUSING HAD EVER BEEN REPLACED. THE ACTUATOR WAS OVERHAULED AT LIEBHERR - AEROSPACE SALINE, INC AND ARRIVED WITH A VALID FAA FORM 8130-3. THIS COMPONENT WAS ORDERED FROM AIRBUS AND WAS SHIPPED TO STAT FROM AIRBUS USA. STAT MED EVAC REMOVED THE DISCREPANT ACTUATOR PRIOR TO ANY AIRCRAFT STARTS OR OPERATIONS.

DAMAGE WAS NOTED TO THE R/H KEEL BEAM CAP (P/N: L533M1010203) IN THE FORM OF 2 EA SPOTS OF CORROSION LOCATED BENEATH THE EDGE OF THE COCKPIT FLOOR. THE CORROSION AT LOCATION 1 MEASURED AT APPROXIMATELY 1.0” X 0.50” AND 0.020” DEEP. CORROSION AT LOCATION 2 MEASURED AT APPROXIMATELY 0.70” X 0.50” AND 0.027” DEEP.

DAMAGE WAS FOUND TO THE FRAME 5 (P/N: L533M1088219/L533M1105219) IN THE FORM OF CRACKS. THERE ARE FOUR CRACKS TO BOTH THE L/H AND R/H FRAME NEAR THE UPPER AND LOWER MOST POINTS WHERE THE MEDICAL INTERIOR BEAM ATTACHES. THE LONGEST CRACK MEASURED TO BE APPROXIMATELY 1.625" ON THE FWD L/H SIDE OF THE FRAME 5 UPPER MOUNT LOCATION OF THE MEDICAL WALL. CRACKS WERE ALSO DISCOVERED ON BOTH L/H AND R/H FRAME 5 ANGLES (P/N: L533M1088218/L533M1105218) AT THE AFT INBOARD LOWER ATTACHMENT POINT OF THE MEDICAL STC. THE L/H ANGLE CRACKED MEASURED APPROXIMATELY 1.05", WITH THE R/H CRACK MEASURING 1.0" IN LENGTH.

DAMAGE WAS FOUND TO THE FRAME 5 (P/N: L533M1088219/L533M1105219) IN THE FORM OF CRACKS. THERE ARE FOUR CRACKS TO BOTH THE L/H AND R/H FRAME NEAR THE UPPER AND LOWER MOST POINTS WHERE THE MEDICAL INTERIOR BEAM ATTACHES. THE LONGEST CRACK MEASURED TO BE APPROXIMATELY 1.625" ON THE FWD L/H SIDE OF THE FRAME 5 UPPER MOUNT LOCATION OF THE MEDICAL WALL. CRACKS WERE ALSO DISCOVERED ON BOTH L/H AND R/H FRAME 5 ANGLES (P/N: L533M1088218/L533M1105218) AT THE AFT INBOARD LOWER ATTACHMENT POINT OF THE MEDICAL STC. THE L/H ANGLE CRACKED MEASURED APPROXIMATELY 1.05", WITH THE R/H CRACK MEASURING 1.0" IN LENGTH.

DAMAGE WAS FOUND TO THE FRAME 5 (P/N: L533M1088219/L533M1105219) IN THE FORM OF CRACKS. THERE ARE FOUR CRACKS TO BOTH THE L/H AND R/H FRAME NEAR THE UPPER AND LOWER MOST POINTS WHERE THE MEDICAL INTERIOR BEAM ATTACHES. THE LONGEST CRACK MEASURED TO BE APPROXIMATELY 1.625" ON THE FWD L/H SIDE OF THE FRAME 5 UPPER MOUNT LOCATION OF THE MEDICAL WALL. CRACKS WERE ALSO DISCOVERED ON BOTH L/H AND R/H FRAME 5 ANGLES (P/N: L533M1088218/L533M1105218) AT THE AFT INBOARD LOWER ATTACHMENT POINT OF THE MEDICAL STC. THE L/H ANGLE CRACKED MEASURED APPROXIMATELY 1.05", WITH THE R/H CRACK MEASURING 1.0" IN LENGTH

BATTERY DISCHARGE, BATTERY TEMP, BUS TIE #1 AND #2 ILLUMINATED IN FLIGHT.

DURING COMPLIANCE WITH EASB EC135-67A-039 PITTING WAS DISCOVERED ON CONNECTING ROD OF THE LATERAL HYDRAULIC ACTUATOR.

DURING COMPLIANCE WITH EASB EC135-67A-039 PITTING WAS DISCOVERED ON CONNECTING ROD OF THE LATERAL HYDRAULIC ACTUATOR.

DURING HOVER OPERATION PILOT HEARD A LOUD SQUEALING NOISE AND A NR 1 ENGINE CHIP INDICATION. PILOT LANDED IMMEDIATELY AND DURING POST FLIGHT WALK AROUND FOUND NR 1 ENGINE POWER TURBINE TO HAVE DAMAGE.

DURING A POST ENGINE MAINTENANCE GROUND RUN, THE PILOT COULD NOT MOVE THE CYCLIC IN THE AFT DIRECTION DURING THE HYDRAULIC SYSTEM CHECK. INSPECTION OF THE MAIN ROTOR CONTROL SYSTEM REVEALED THAT THE LONGITUDINAL ACTUATOR WAS AT THE FULL AFT TRAVEL LIMIT. WITH HYDRAULIC POWER REMOVED AND NO CYCLIC INPUT THE AFFECTED ACTUATOR WOULD MOVE TO THE CENTERED POSITION.

AUTOPILOT FAILED PREFLIGHT TEST, AUTOPILOT ME LED. DURING TROUBLESHOOTING OF AUTOPILOT SYSTEM BY MAINTENANCE THE AIRCRAFT EXPERIENCED UNUSUAL VIBRATIONS/OSCILLATIONS IN THE HOVER CHECK. FOUND FAULTY ELECTRO HYDRAULIC ACTUATOR

DURING FLIGHT THE OVER SPEED CAUTION ILLUMINATED FOLLOWED BY CHIP A CHIP CAUTION FOR THE NR 2 ENG. THE PILOT TRIED TO BURN THE CHIP & WAS UNSUCCESSFUL. THE PILOT LANDED SAFELY AT THE NEAREST AIRPORT & DURING SHUTDOWN OF THE ENG THERE WAS A LOUD BANG. DISCOVERED METALLIC PARTICLES ON BOTH NON ELEC CHIP PLUGS. N1=12740.7 & N2=4718.

NUMBER 2 ENG DHIP LIGHT THAT WOULD NOT CLEAR AFTER ON FUZZ BURN APPLICATION. UPON INSPECTION, THE REAR BEARING CHIP PLUG WAS FOUND TO HAVE SIGNIFICANT BUILDUP OF METALLIC MATERIAL, N1=3478 N2=9543.

DURING A PERIODICAL INSPECTION, A 7 INCH CRACK WAS DISCOVERED ON THE LT SIDE OF THE TAIL BOOM. THE CRACK WAS MOSTLY LOCATED UNDERNEATH THE LT HORIZONTAL FITTING.

DURING POST MX FLIGHT CHECKS, THE PILOT SAID SMALL INPUTS TO THE COLLECTIVE RESULTED IN LARGE MOVEMENTS TO THE ACFT. THE ENG TORQUES WERE ALSO DIFFICULT TO STABILIZE. AFTER LANDING, SHUTDOWN & INSPECTION BY MX, A LAG OF COLLECTIVE INPUT TO THE MOVEMENT OF THE MAIN ROTOR BLADE WAS NOTICED WHILE OPERATING THE HYDS ON THE GND.

CRACKS IN 5 TAIL ROTOR STATORS WERE FOUND DURING AN INSPECTION.

DURING FLIGHT THE PILOT HEARD A NOISE FROM THE ENG WITH AN ASSOCIATED ACFT YAW. THE TORQUE OF THE ENG FELL BELOW THE OPPOSING ENG FOLLOWED BY A CHIP CAUTION ON THE CAD. THE ENG THEN TORQUE INDICATION CAME BACK TO MATCH THE OPPOSING ENG. UPON POST FLIGHT WALK AROUND THE PILOT NOTICED OIL COMING FROM THE TURBINE AREA OF THE ENG, N1=3429, N2=9101.

ENGINE FLAMEOUT DURING TRAINING FLIGHT WHILE NR 1 ENG WAS IN MANUAL MODE.

DURING A PREFLIGHT INSPECTION, DAMAGE WAS DISCOVERED ON THE MAIN ROTOR DRIVE LINK. IT WAS DETERMINED THAT THE DAMAGE WAS MOST LIKELY CAUSED BY THE LOWER SWASH PLATE HOSE CLAMP, DUE TO THE FACT THAT THE CLAMP WAS MISSING.

DURING A PERIODICAL INSPECTION A CRACK WAS FOUND IN THE 10L FITTING OF THE TAIL BOOM. THIS TAIL BOOM HAD SB-117-C-2-53-007 REINFORCEMENT KIT INSTALLED. THIS TAIL BOOM WAS ALSO REPAIRED 852.1 FH AGO FOR INCORRECT RIVETS INSTALLED IN THE TOP SKIN OF THE BOOM.

DURING COMPLIANCE WITH THE AMOC FOR AD 2013-12-06 IT WAS NOTICED THAT THE BOLTS THAT SECURE THE AIR CONDITIONER PULLEY TO THE ROTOR BRAKE DISK WERE NOT THE CORRECT LENGTH. AFTER FURTHER INVESTIGATION IT WAS DETERMINED THAT THE BOLTS WERE THE CORRECT LENGTH & THE HOLES IN THE PULLEY WERE NOT THREADED THE ENTIRE DEPTH OF THE HOLE. THIS PREVENTED THE BOLTS, WITH THE NEW SAFETY SHEETS INSTALLED, FROM PROPERLY SECURING THE PULLEY TO THE DISK.

DURING A PERIODICAL INSPECTION, NUMEROUS RIVETS THAT ATTACH THE TOP SKIN OF THE TAIL STRUCTURE, WE DISCOVERED TO BE BROKEN AND MISSING. AFTER FURTHER INSPECTION IS WAS DISCOVERED THAT THE INCORRECT DIAMETER RIVETS WERE INSTALLED IN ALL RIVET LOCATIONS ATTACHING THE TOP SKIN.

DURING APPROACH INTO A LANDING ZONE, THE NR 2 ENGINE SHUTDOWN. INDICATIONS ON THE CAD WERE A "FLI DEGRADE" AND A DELTA N1. THERE WAS NO TORQUE INDICATION ON THE VEMD.

DURING FLIGHT, THE PILOTS WERE COMPLAINING OF A VIBRATION IN THE MAIN ROTOR. THE MAIN ROTOR TRACK AND BALANCE WAS CHECKED AND NECESSARY ADJUSTMENTS WERE MADE TO BRING IT WITHIN ACCEPTABLE LIMITS. THE VIBRATION REMAINED DURING FLIGHT. AFTER FURTHER INSPECTION OF THE MAIN ROTOR CONTROLS IT WAS DISCOVERED THAT THE BOLT AND BUSHINGS THAT ATTACHED THE UPPER GUIDANCE UNIT TO THE AIRFRAME HAD BEEN WORN BEYOND ACCEPTABLE LIMITS.

DURING A ROUTINE AIRFRAME INSPECTION, A LONGITUDIONAL CRACK WAS DISCOVERED ON THE TOP SIDE OF A TAIL ROTOR STATOR.

MAIN ROTOR BLADE HAS A CRACK.

DURING A DAILY AIRWORTHINESS CHECK, A MECHANIC NOTICED AN UNUSUAL PAINT CRACK ON THE LT SIDE OF THE FITTING, AFT OF THE HORIZ STABILIZER. AFTER REMOVAL OF PAINT IN THE AREA AND A VISUAL INSPECTED, A CRACK WAS DISCOVERED.

DURING REPLACEMENT OF THE LONG TAIL ROTOR DRIVESHAFT BEARINGS, IT WAS DISCOVERED THAT ALL 3 OF THE SPLIT PINS THAT ATTACH THE ADAPTOR TO THE DRIVESHAFT WERE CRACKED.

SIGNIFICANT TORQUE SPLIT INDICATED ON VEMD FOLLOWED BY AN UNCOMMANDED ENGINE SHUTDOWN DURING CRUISE FLIGHT. A SUBSEQUENT BORESCOPE INSP OF NR 2 ENGINE, REVEALED A CRITICAL FAILURE OF THE COMPRESSOR TURBINE STATOR.

DURING HOVER OPERATION, THE PILOT NOTICED SMOKE COMING FROM THE BOOT AROUND THE CYCLIC. AFTER LANDING A VISUAL INSP IN THE AREA BELOW THE CYCLIC BOOT COVER REVEALED LOOSE RING TERMINALS (ON WIRES 2006N AND 2008N) FOR THE MOVEABLE SEARCH LIGHT.

PILOT COMPLAINED OF A VIBRATION THAT WAS MOST NOTICEABLE IN THE TAIL ROTOR PEDALS. UPON INSPECTION OF THE TAIL ROTOR A BLADE WAS DICOVERED TO HAVE EXCESSIVE LEAD-LAG MOVEMENT.

DURING THE BUILD UP OF THE MAIN GEAR BOX IT WAS DISCOVERED THAT THE ROTATING PORTION OF THE SWASH PLATE WOULD NOT ROTATE.

CRACKS FOUND EMANATING FROM LOWER 5 HI-LOK FASTENERS IN TITANIUM REINFORCEMENT LOCATED AT FRONT OF NR 1 ENGINE DECK. THIS RE-ENFORCEMENT IS ATTACHED TO THE AFT SIDE OF FRAME 5 IN THE AIRCRAFT FUSELAGE. MANUFACTURER HAS BEEN NOTIFIED AND REPAIR ORDER REQUESTED.

MAIN TRANSMISSION RECEIVED FROM MANUFACTURER (0 TSO). OPERATOR FOUND TWO BOLTS SAFETIED WITH ZERO TORQUE ON SWASH PLATE BALANCE WEIGHT. TORQUE RECOMMENDED 35.4 - 53.1 INCH POUNDS AS PER REM. BOTH OIL COOLER BLOWER FAN RETENTION NUTS FOUND SAFETIED WITH ZERO TORQUE APPLIED. TORQUE RECOMMENDED 425-513 INCH POUNDS AS PER AMM. AT INITIAL POST INSTALLATION GROUND RUN APPROXIMATELY 2 QUARTS OF OIL LEAKED FROM TAILROTOR OUTPUT FLANGE SEAL IN THE FIRST MINUTE OF OPERATION. AIRCRAFT WAS SHUT DOWN IMMEDIATELY .FURTHER INVESTIGATION REVEALED THAT THE TAIL ROTOR OUTPUT FLANGE CONNECTING PART DID NOT HAVE THE REQUIRED SEAL SLEEVE INSTALLED AT OVERHAUL. AFTER CONSULTATION WITH MANUFACTURER REPRESENTATIVES IT WAS DETERMINED THAT LOAD TESTING IS PERFORMED WITH A SLAVE TAIL ROTOR OUTPUT FLANGE INSTALLED. AFTER TESTING THE PRODUCTION FLANGE IS INSTALLED. APPARENTLY WHEN THE PRODUCTION FLANGE WAS INSTALLED AFTER LOAD CELL TESTING IT WAS INSTALLED WITHOUT THE SLEEVE THAT THE OUTPUT SEAL NORMALLY SEATS ON. SUBMITTER HAS BEEN ASSURED THAT THE MANUFACTURER IS CHANGING OVERHAUL PROCEDURES TO ENSURE THAT THIS DISCREPANCY DOES NOT OCCUR AGAIN. DISCREPANCIES WERE CORRECTED IAW MANUFACTURERS RECOMMENDATIONS.

A 2 INCH CRACK WAS DISCOVERED ON THE BULKHEAD SKIN THROUGH THE LIGHTNING HOLE ADJACENT TO FRAME 5 (FR5), LT SIDE NEAR THE NR 1 ENGINE FACEC COMPUTER. THE CRACK WAS DISCOVERED WHEN THE PELICAN RACK WAS LOWERED DURING INSPECTION.

A 4 INCH CRACK WAS DISCOVERED ON THE BULKHEAD SKIN ADJACENT TO FRAME 5 (FR5), RT SIDE NEAR THE NR 2 ENGINE FACEC COMPUTER. THE CRACK WAS DISCOVERED WHEN THE PELICAN RACK WAS LOWERED DURING INSPECTION.

DURING INITIAL CLIMB AFTER TAKEOFF, THE NR 1 ENGINE TORQUE DROPPED TO ZERO. THE PILOT INITIATED A SINGLE ENGINE LANDING WITH NO INCIDENT. INSP OF THE NR 1 ENG REVEALED THE SWIVEL FITTING ATTACHED TO THE COMPRESSOR MODULE TO BE LOOSE DESPITE ORIGINAL FACTORY SAFETY STILL BEING INSTALLED. THIS FITTING ENABLES THE P2 AIR LINE BETWEEN THE FUEL CONTROL AND COMPRESSOR MODULE TO BE AFFIXED TO THE COMPRESSOR HSG. THIS AIR SOURCE PROVIDES COMPRESSOR REF PRESSURE TO THE FUEL CONTROL FOR DECELERATION SCHEDULING. THE FITTING WAS REMOVED, INSPECTED, REINSTALLED AND SAFELY WIRED. GROUND RUN AND OPS CHECK FLIGHT WERE PERFORMED WITH NO FURTHER DEFECTS NOTED. ACFT WAS RETURNED TO SERVICE.

INBD ENG MOUNT ON BOTH NR 1 AND NR 2 SIDES HAVE EXCESSIVE WEAR AFTER 600 HOURS OF FLIGHT TIME. MFG PN FOR THE RT SIDE IS 117-601101 AND THE PN FOR THE LT SIDE IS 117-601091.

FOUND 1 BRACKET OUT OF 4 THAT HOLD THE MAIN ROTOR HUB CAP ON TO THE MAIN ROTOR HEAD CRACKED. THIS BRACKET PN 117-13902. THE OLDER MODEL ACFT (A4,B1,B2 AND C1) HAD AN AD ATTACHED TO THIS PART YEARS AGO,THE BRACKET WAS BEEFED UP, THE PN CHANGED TO 117-13902 AND THE AD WAS NO LONGER APPLICABLE. THE CRACKS FOUND IN THE NEW THICKER BRACKETS ARE IN THE SAME PLACE THE AD ON THE OLD THINNER BRACKETS REQUIRED AN INSP.

AIR CONDITIONING DRIVE PULLEY BOLTS SHEARED AND THE PULLEY DETACHED FROM THE DRIVE. IT LOOKS LIKE THE BOLTS LOST TORQUE SLIGHTLY, BOLT HEADS WERE SAFETIED, THE THE PULLEY WORKED THE THE BOLTS UNTIL THEY SHEARED. I HAVE OPERATED 8500 HOURS ON 5 DIFFERENT EC-145S WITH THIS PULLY SYSTEM WITHOUT ANY PROBLEMS. I HAVE CHECKED MY OTHER 5 EC-145 HELICOPTERS FOR AIR-CONDITIONG PULLY ATTACHMENT AND THEY ARE ALL PERFECT. THIS SEEMS TO BE AN ISOLATED INCIDENT. --- AFTER FURTHER INVESTIGATION OF THE ATTACHING BOLTS, THE BOLT SHANKS THAT WERE STILL IN THE PULLEY WERE TIGHT HOWEVER FRETTING WAS FOUND ON THE PULLEY ATTACHING SURFACE. IT IS QUITE POSSIBLE THAT THE INSTALLER OF THIS STC (HELI-DYNE)INSTALLED THE WRONG LENGTH BOLTS.

AFTER A NORMAL SHUT DOWN NR 1 ENGINE RESTARTED ON ITS OWN AFTER ALL POWER AND SWITCHES WERE IN THE OFF POSITION. AFTER INSPECTING THE ENGINE, FOUND MANY POWER TURBINE BLADES MELTED. THIS ENGINE JUST CAME BACK 30.4 HOURS AGO AT OVERHAUL FACILITY. THE ENGINE IS A PW206B, I COULD NOT FIND THAT IN THE ABOVE SERIAL NUMBER SECTION

DURING A DAILY CHECK OF THE AIRCRAFT A GRINDING NOISE WAS DETECTED FROM THE TAIL ROTOR DRIVESHAFT COVER, AFTER AN INVESTIGATION THE GRINDING NOISE WAS FOUND TO BE COMING FROM THE 2 AFT TAIL ROTOR DRIVESHAFT HANGAR BEARINGS. THE BEARING PN: BB1B649616B, SN: N1899 AND N1828. THE BEARINGS WERE STAMPED (ITALY V) ON THE SIDE. THERE IS NO VISIBLE LOT NUMBER.

THE AIR-CONDITIONING COMPRESSOR SEIZED, THE DRIVE BELT CONTINUED TURNING, SMOKING UNTIL IT BROKE.

THE AIR CONDITIONING DRIVE BELT HAD CRACKS IN IT AT A BELT TOTAL TIME OF 359.3. THE LIFE LIMIT OF THIS BELT IS 1000 HOURS.

LEAKING FLUID FROM WEEP HOLE.