Emergency Procedures for the Cessna 404 Titan

function and the emergency beep trim was selected on this engine. At 1117 hrs the aircraft was approximately 5 miles from the Aberdeenshire coast at 6,000 feet when the first officer, who was the handling pilot, drew the commander's attention to the No 1 engine torque which was slowly rising from the matched cruise power setting of 55%. The commander attempted briefly to prevent the torque increasing further by using the No 1 engine emergency beep trim system, however this was not successful and because the torque continued to rise he closed down the No 1 engine when the torque had reached 75%. Until the point of engine shut down the oil temperatures and pressures had remained normal and only the ‘ENG 1 CHIP DET' caption was illumin- ated. Captions illuminated indicating primary and auxiliary transmission lubrication failure and the 'ROTOR BRAKE ON' warning light also illuminated. An immediate descent was initiated and a radio call made to Aberdeen although a formal distress call was not transmitted. The cabin attendant was on the flight deck at this time and the commander told him to prepare the passengers for a possible ditching. When he opened the flight deck door he saw that the cabin was filling with a light coloured smoke and he immediately drew this to the attention of the commander before going into the cabin. During the subsequent investigation passengers reported that before the smoke began to appear there was a loud screeching sound from the rear of the aircraft terminating with an explosion or snapping noise. Some also reported a general increase in the level of vibration. 2 1.2 Having briefed the passengers from his normal position at the front of the cabin the cabin attendant walked slowly down the aisle to confirm that his instructions had been carried out. When he was adjacent to rows 10 and 11 at the rear of the aircraft he could feel heat from the area of the port cabin roof forward of the rear bulkhead. Meanwhile the commander had selected the cabin ventilation to 'RAM VENT', which was effective in clearing the smoke, and he then turned his attention to the gearbox temperature and pressure gauges. The temperature gauge which was selected to the 'SCAN' position, showed a rapid rise in temperature indication to the maximum gauge temperature. When the gearboxes were selected individually however, normal temperatures were shown. The pressure gauge indicated zero regardless of the position of the selector switch. The commander initially intended to make an emergency landing on the beach but when the cabin attendant reported that smoke was clearing he elected to divert to Longside for a single engined landing on the 2,000 feet runway. He called Longside at approximately 1121 hrs and landed there at 1126 hrs. During the final stages of the approach the cabin attendant reported that some smoke was again coming into the rear of the cabin. When the aircraft had halted on the runway the passengers were evacuated through the forward exit and the remaining engine was shut down using the emergency 'T' handle. The rotor brake was applied normally and after the rotor had come to a stop the crew inspected the aircraft to ensure that there was no persisting fire. Injuries to persons Injuries Fatal Serious Minor/None 1.3 Damage to aircraft 1.4 Crew Passengers Others 3 32 Substantial fire damage in the area of the No 1 engine transmission gearbox. Other damage There was no other damage. 1.5 Personnel information Commander: Licence: Helicopter type ratings: Male, aged 41 years Airline Transport Pilot's Licence (Helicopters) valid until 24 March 1987 Alouette II Sikorsky S61N Boeing Vertol Chinook 3 Instrument rating: Medical Certificate: Certificate of test: Flying experience: dated 27 January 1983 valid until 28 February 1983 Boeing Vertol Chinook 27 January 1983 Total all types: 5,520 hours Total helicopter: 5,460 hours Total BV234: 860 hours Total last 30 days: Total last 7 days: 45 hours 14 hours Duty time: Co-pilot: Licence: Helicopter type ratings: Instrument rating: Medical Certificate: Certificate of test: Flying experience: Off duty 20 February 1983 On duty 0600 hrs 21 February until 1500 hrs on 21 February Male, aged 26 years Airline Transport Pilot's Licence (Helicopters) valid until 24 January 1993 Bell 47 Bell 206 Sikorsky S61N Boeing Vertol Chinook None Valid until 31 March 1983 Total all types: Total helicopter: Boeing Vertol Chinook 7 February 1983 1,845 hours 1,515 hours Total BV234: 377 hours Total last 30 days: 54 hours Total last 7 days: 13 hours Duty time: 1.6 Aircraft information: 1.6.1 General information Type: Manufacturer: Operator: Airframe Serial Number: Off duty 20 February 1983 On duty 0600 hrs 21 February until 1500 hrs on 21 February Boeing Vertol 234LR Boeing Vertol Company, Philadelphia, USA British Airways Helicopters MJ 004 Date of Manufacture: Certificate of Airworthiness: Certificate of Registration: Certificate of Maintenance: Next Check Due: Total Airframe Hours: 1981 Transport Category (Passenger) valid until 28 May 1983 G-BWFC in the name of British Airways Helicopters Validated on 31 October 1982 at 1,934.20 flying hours A4 Inspection (25 hour check) due at 2,323.50 hours 2,325.05 hours No 1 Engine Transmission Gearbox Manufacturer: Boeing Vertol No 1 Gearbox Part Number: Engine Transmission 234D6200-1 No 1 Engine Transmission Gearbox Serial Number: All-3520 Total Engine Transmission Gearbox Hours .1.6.2 Aircraft weight and centre of gravity 1,241 hours since new The helicopter's weight at take-off was calculated to be 17,798 kg and its centre of gravity was within limits. The maximum approved gross weight for this helicopter is 22,000 kg. 1.6.3 Technical Log 1.7 Examination of the helicopter's technical log did not reveal any defects which were related to the incident. Meteorological information The area was dominated by a large anticyclone centered just to the east of Scotland. There was a layer of stratocumulus the base of which was at around 3,000 feet and the top at 5,500 feet. The visibility was generally in excess of 20 kilometres. 1.8 Aids to navigation 1.8.1 On the ground Radar Very High Frequency Omni Range (VOR) 5 1.8.2 1.9 1.10 1.11 1.12 Distance Measuring Equipment (DME) Medium Frequency Non-directional Beacon (NDB) Decca In the air The aircraft was equipped with: VOR/ILS - twin installation DME Automatic Direction Finder (ADF) Decca Weather radar Communications The aircraft was in contact with Aberdeen Radar and immediately after the incident called them saying "WE'RE HAVING A PROBLEM WITH OUR ENGINE AND REQUEST DESCENT TO FLY BELOW CLOUD WE'RE HEADING TOWARDS SCOTSTOWN HEAD". Radar cleared them to descend to 1,500 feet and gave them their position as 12 miles northwest of Scotstown Head. At 1121 hrs when the aircraft was at 2,500 feet below cloud the aircraft commander decided to divert to Longside and called that airfield for clearance and requested fire cover for the landing. Aerodrome and ground facilities Longside was an unlicensed heliport with a 2,000 feet tarmac runway operated by North Scottish Helicopters. A fire truck was available which carried a selection of hand operated fire extinguishers and in addition a 20 gallon lightwater extinguisher and a 50 kg BCF extinguisher. Additional fire cover was requested from the Grampian Fire Service at Peterhead and this was available some 9 minutes after the aircraft had landed. Flight recorders The helicopter was not equipped with a flight data recorder or a cockpit voice recorder, nor were these required to be fitted. Wreckage and impact information 1.12.1 Examination of the aircraft When the AIB investigation was initiated at Longside airfield on 23 February 1983, the No 1 engine and transmission gearbox along with the combining gearbox and fire damaged wiring looms had been removed from the helicopter. The photographs, included at Appendix 1, were taken shortly after the helicopter landed at Longside and show the areas of damage. Examination of the helicopter revealed that severe fire damage had affected the No 1 engine transmission gearbox fairing and associated stub fairing with lesser damage to the engine intake screen and lip fairings. The stub fairing had been disrupted by failure of the transmission cross-shaft between the engine transmission gearbox and the combining gearbox (refer Appendix 2). The mid section of this shaft was missing leaving a short section still attached to the transmission gearbox flexible coupling and some 15 cm of shaft attached to the combining gearbox input coupling. What remained of the cross-shaft was fire affected, particularly around the outboard failure where the fractured area was splayed radially and exhibited a rotational rub-type witness mark on the fractured surface. Evidence of fire entrainment inboard from the engine transmission gearbox output shaft seal was observed. There was marked flame erosion of the aluminium alloy rib at the root end of the stub fairing and molten aluminium alloy spatter and thermal discolouration on the two adjacent steel flying control rods. The protective finish on the left rear side of the combining gearbox was blistered and two lubrication hoses in the same area showed evidence of fire induced surface craze cracking. Electrical wiring grommets at the base of frame 482 were damaged and the left mounting structure for the combining gearbox was heat discoloured and distorted. The combining gearbox filler cap assembly had been found in the 'open' condition and the associated hinge had cracked. The internal filter-screen 'cylinder' was partially split longitudinally due to break-away of part of the soldered seam. Inspection of the left engine transmission gearbox main lubrication system pressure filter showed it to be heavily contaminated with fine steel and bronze/aluminium alloy particles, although the by-pass indicator button was found in the normal position, as were those associated with two other pressure filters for the combining gearbox and starboard engine transmission gearbox. The smaller auxiliary system pressure filter for the No 1 engine transmission gearbox was similarly contaminated. The No 1 engine transmis- sion gearbox scavenge line filter 'screen' had trapped some six sections of bronze bearing cage and in addition a 3.5 cm long section of inner race land (or 'rail') from a large bearing had fallen out of the scavenge hose upon disconnection. No evidence of oil level was apparent in any of the three reservoir sight-glasses, although this is normal for Nos 1 and 2 reservoirs due to drain-back into their respective engine transmission gearboxes. Small oil samples were taken from all three reservoirs (ie, the No 1 engine transmission gearbox, combining gearbox and No 2 engine transmission gearbox). In addition, approximately one litre of oil was drained from the No 1 engine transmission gearbox and some two litres from the No 2 engine transmission gearbox. External examination of the No 1 engine transmission gearbox assembly at British Airways Helicopters, Aberdeen, showed it to be generally blackened by fire, with a particularly noticeable heat-band effect around the housing in a position outboard of the input pinion main roller bearing. There was some localised blistering around the breather hose union on the upper forward area of the housing and the hose connection was found to be loose by half a turn. The small diameter auxiliary pressure hose exhibited a white 'frosty' appearance, caused by the Teflon inner hose exuding through the steel braided outer sheath, due to heat absorption. 7 1.12.2 The status of the Maintenance Inspection Panel in the ramp area at the rear of the helicopter had been recorded after the landing at Longside: the No 1 (left) engine transmission gearbox chip detector and the No 1 engine filter screen magnetic indicators had ‘tripped'. On the flight deck the circuit breakers for the engine transmission gearbox oil pressure warning, rotor brake warning, DC power warning and engine oil level warning had also tripped. Subsequent detailed examination 1.12.2.1 No 1 Engine and Transmission Gearbox The No 1 engine together with its transmission gearbox cross-shaft and the combin- ing gearbox were removed to the operators engineering base where they were examined in detail. Examination of the engine revealed no serious defects within the power plant. Some very light foreign object damage to the compressor rotor blade leading edges and stator trailing edges was apparent together with some acceptable light alloy 'spatter' on the stage 1 and 2 turbine blades. There was no damage to the free turbine. Strip examination of the No 1 engine transmission gearbox revealed that the input pinion main roller bearing Part No 114DS665-1 had failed, with complete break-up of the phosphor-bronze cage, inner race aft land and a 4 cm section broken away from the front land (refer Appendix 3). Excessive wear was evident on both races with the outer race exhibiting an approxi- mately 60° ‘arc' of break-up. All twelve rollers had a black heat-discolouration, with skid-induced 'flats' on their rolling surfaces. Some showed evidence of the rollers 'skidding' at 90 degrees to their normal axial orientation during the failure sequence. The lubrication plate located on the aft side of this bearing was badly worn on its forward face due to roller and bearing debris contact. Bearing cage and land debris had damaged many gear teeth as well as the output bevel ring gear baffle. The scavenge outlet at the base of the engine transmission gearbox housing was almost completely blocked with fine metallic debris, which had also accumulated on the adjacent chip/temperature detector assembly. All four lubrication hoses associated with the No 1 engine transmission gearbox (ie, main and auxiliary pressure, common scavenge return and 'breather') were found to have suffered total loss of their internal elastomer in the region of the outboard cross-shaft failure due to the effects of the fire in this area. Tests conducted to examine this feature showed that temperatures in the region of 600-650°C were required to produce this phenomenon. The failed input pinion main roller bearing was examined by the Royal Aircraft Establishment (RAE) Materials Department and Boeing Vertol, Philadelphia. 8 Neither laboratory could determine the primary cause of the bearing failure. All fractures examined were ascribed to over-stressing with the exception of a localised fracture evident on the forward land of the inner race which the RAE metallurgists thought due to low cycle fatigue caused by roller end radii contact and therefore secondary to some other problem which had allowed roller track divergence. RAE made the point that on low ductility steels of this type (AMS 6490, 'M50', Rockwell C60 minimum) fatigue is typically difficult to identify due to the lack of fine striation detail on the fractures. The only fatigue identified by Boeing was associated with the approximately 60° arc of surface break-up (spalling) present on the outer race, where the heat and mechanical deformation of the eventual failure had also induced re-hardening of the track surface layer. Oil samples taken from the No 1 engine transmission gearbox lubrication system were analysed by the Ministry of Defence Materials Quality Assurance Directorate. The analysis confirmed that the oil complied with the correct specification - ie, D.ENG R.D. 2497, AEROSHELL 555. Inspection of the Spectrometric Oil Analysis Programme (SOAP) results from G-BWFC before the incident showed them to have been generally satisfactory with no evidence of the impending bearing failure. 1.12.2.2 Other Components The No 1 engine transmission cross-shaft was examined at both RAE and Boeing Vertol Material Laboratories and found to exhibit evidence of excessive thermal exposure (1100°F minimum) at the outboard end and to a lesser extent within the inboard area. Both outboard and inboard failures were consistent with over-stressing. The output shaft sprag-clutch assembly could not be hand-rotated as found and was discoloured with a blue/brown 'shelac' appearance from over-heating in the presence of oil. Later examination by the manufacturer (Borg-Warner) indicated that it had been serviceable and showed evidence of excessive over-run wear on the sprags due to operating in a high temperature lubrication environment with fine metal particle contamination. The main lubrication pump cylinder was removed from the combining gearbox and inspected. No failure was observed. The auxiliary pumps and filter housing were removed from the front of the combining gearbox and similarly found to be mechanically intact. Removal of the rotor-brake, back-plate and aft bevel gear shaft from the combining gearbox and inspection of the associated bevel gears showed no evidence of any distress. The transmission oil pressure monitoring system sensors (main and auxiliary pressure switches and main system transducer) were checked after the incident and all operated satisfactorily. 1.12.2.3 Transmission Lubrication Monitoring System: (Refer Appendix 4) The transmission oil temperature monitoring system sensors, the combined chip/ temperature detector assembly and the associated temperature switch, were examined by their respective manufacturers. The detector had suffered damage to 9