Aviation Investigation Report A05O0115

The Bell Textron 212 helicopter (registration C-GNHX, serial number 30983) was being ferried from Bolton, Ontario, to Richmond, British Columbia. The recently purchased helicopter was being flown by the company’s chief pilot with two passengers on board. At 1220 eastern daylight time, the helicopter was at an altitude of 1500 feet above sea level with an airspeed of 100 knots, when there was a series of loud bangs immediately followed by severe airframe vibrations. The pilot had difficulty controlling the helicopter for the next 10 to 15 seconds.

The pilot immediately lowered the collective, pulled back on the cyclic control and brought the engine throttles to idle. He regained control of the helicopter, but the banging and vibrations continued. Every time one of the advancing main rotor blades came forward, it would climb off track abnormally. The vibrations and banging became more severe as the flight continued. The pilot proceeded toward a large ploughed field for an emergency landing. As the airspeed decreased, the helicopter became more controllable, and a successful landing was carried out. There were no injuries to the occupants. The helicopter was substantially damaged from the in flight vibrations.

Other Factual Information

The weather for the flight was good visual meteorological conditions and was not considered to be a factor in this occurrence.

Records indicate that the aircraft was maintained in accordance with existing regulations and approved procedures. The pilot was certified and qualified for the flight.

After landing, a post-flight inspection revealed that one of the main rotor blades had sustained damage. A small section of skin near the blade tip, aft of the spar doubler, on the lower surface of the rotor blade had debonded. The skin was raised and curled, but had not separated from the blade (see Photo 1). The debonded skin measured 25 inches by 2 inches between stations 263 and 288. It was later discovered that several of the main rotor head components and transmission had been damaged by the severe vibrations encountered during the flight.

The damaged main rotor blade (part number 212-015-501-115, serial number A-3257) had accumulated 3251 hours of flight time since new. The total service life for the blade was 4000 hours. A review of the blade’s service records indicated that it was manufactured by Bell Helicopter and entered service in December 1996.

In early 2005, the same blade had been damaged while the helicopter was parked in a hangar. The blade was then shipped to an authorized rotor blade repair shop.

While paint was being stripped from the rotor blade in preparation for repair, deep corrosion pitting was discovered on the lower skin surface between stations 243 and 262, just inboard of where the debonding later occurred on the 10 June 2005 flight (see Photo 1). Because the pitting pattern exceeded the allowable limits, the repair shop proposed a repair procedure to Bell Helicopter and received approval. The repair procedure included removing the damaged skin and replacing it with a bonded external doubler. The trailing edge trim tab was also replaced.

The skin–to–inner core bonding procedure required using a bladder and heater blanket tool. This tool ensures proper curing of the adhesive by applying heat and pressure to the area being repaired. This type of repair is performed regularly to repair damaged rotor blades. The bladder and heater blanket tool that was used covered the rotor blade from its tip to a point inboard of the repair area, which included the area where the debonding took place on the occurrence flight. The repair process called for the temperature to be controlled and monitored during the entire cure cycle. After the repair was completed, the blade was inspected by tap hammer in the repair area and all the way to the tip. The blade was then returned and installed on C GNHX. As part of the investigation, records of the repair procedure were reviewed by the TSB Engineering Laboratory and Bell Helicopter, and it was verified that the procedure was performed in accordance with the standard recommended procedures.

Following the repair using the bladder and heater blanket tool, the blade was in service for approximately four flight hours before the lower skin debonded on the occurrence flight at the spar doubler between stations 263 and 288.

The debonded section between stations 263 and 288 was examined. There was a cohesive bond of the adhesive to the skin substrate. However, the original bonding adhesive used during the manufacturing process was not uniformly adhered to the spar doubler; only small remnants of the adhesive remained (see Photo 2). The total amount of adhesion could not be accurately quantified because the adhesive on the skin and on the spar doubler had eroded. This erosion of the adhesive was caused by the airflow and environmental elements entering the debond area in the time frame between the skin debonding and the completion of the emergency landing.

A scanning electron microscope examination of the debonded skin sample indicated that the adhesive, while remaining firmly attached to the skin, had replicated sanding marks from the spar along the adhesive to spar interface, clearly showing that most of the adhesive had bonded well to the skin, but not to the spar. There was little indication of any adhesive remaining attached to the spar surface. The sanding marks on the spar surface are indicative of the standard procedure used for surface preparation before applying the adhesive. Following this initial examination, preparation for repairing the blade was begun by cleaning the spar surface. However, during this process, the surface was altered, and bonding evidence was removed, preventing further analysis of the bond to the spar surface.

Other sections of the blade were examined to determine the overall adhesive nature of the bond between the skin and spar surface. Further samples of skin were peeled from different sections of the blade inboard of the area damaged on the occurrence flight. The area where the previous repair had been completed showed sporadic cohesive separation on the skin side. All of the other samples removed indicated even adhesion between the skin and spar surfaces.

During the examination of other sections of the blade along the spar doubler, two areas were discovered where the honeycomb inner core was crushed and had separated from the rear of the spar. The two areas were located between station 211 and station 224 and between station 263 and station 287. A layer of foam adhesive rests between the core and the spar (see Photo 3). In the two damaged areas, the adhesive foam was present at the spar, but there was no contact between the adhesive foam and the core. Bell Helicopter’s Engineering Department determined that, even though there was a non-effective bond between sections of the honeycomb core and the spar, the skin–to–spar bondline would not have experienced static or fatigue failure during the service life of the rotor blade.

The available records were examined to determine if this blade had incurred any additional damage during its service life. There was no indication of any additional damage.

Data from the FDR indicated that the aircraft did not exceed any limitations prescribed in the AFM, yet propeller operating disturbances occurred. Although the static engine pitch angle limitation was unknowingly exceeded and might have contributed to a simultaneous temporary loss of engine oil pressure in both engines, the resolution of acceleration vectors showed that the effect of the pitch angle was equivalent to a steady-state pitch-up angle within limitations during the manoeuvre.

Low oil pressure conditions are known to occur during negative g manoeuvres with the original PRV but the low oil pressure condition did not last long enough to affect propeller operation. Data suggest that the modified PRV contributed to the propeller speed disturbance by prolonging the low oil pressure conditions. This result has been replicated in subsequent flight testing involving push-overs to about -0.10 g from positive pitch angles of about 23° and demonstrates that propeller operating disturbances can be anticipated any time that negative g conditions are encountered. This previously unknown consequence is the subject of discussions between the manufacturer and TC.

Section 2.2.5 of the AFM, Manoeuvring Limit Load Factors, indicates that the allowable load factors limit the permissible bank angle in turns and limit the severity of pull-up and push-over manoeuvres. Bank angles were not an issue in this incident, and it is true that load factor limits would restrict the severity of a pull-up or push-over manoeuvre. While the flight test load factor targets were within the limits specified in the AFM, the AFM is not produced as a flight test design guide. Certification flight testing programs could involve manoeuvres considered to be extreme for the category of aircraft. The Cascade Aerospace Inc. flight test engineering group did not know exactly what pitch angle would be achieved during the planned series of manoeuvres. Therefore, it would be prudent for engineering groups to fully research the limits of systems, such as lubrication and fuel systems, when designing such flight testing programs and incorporate such information in the flight test cards applicable to each flight.

The airframe manufacturer concluded that the pitch-up attitude achieved was extreme and is not likely to be encountered in normal operations for this type of aircraft. However, a query of another DHC-8 operator (not a Q400 model) indicated that high positive pitch angles (30° or more) could be achieved in flight training manoeuvres. Therefore, it is not beyond reason for load factors of less than 0 g to be induced during a training exercise or actual traffic or terrain avoidance manoeuvre. Without knowledge of the associated consequences, an already abnormal manoeuvre could involve much higher risks if an unsuspecting flight crew encountered a single or double loss of engine oil pressure and subsequent propeller overspeed condition(s).

The original certification flight test demonstrations conducted by Bombardier did not result in any abnormalities of propeller operation with the original PRV installation in the engines. Since the introduction of the modified PRV (SB 35038, Revision 3), this in-flight incident and subsequent test cell results demonstrate that the loss of engine oil pressure was prolonged by the modified PRV and resulted in the propeller underspeed when the blades were driven in the coarse pitch direction. This effect of the PRV modification had not been foreseen. When this event occurred, the balance of the system performed as designed as evidenced by the recovery of the No. 1 engine/propeller to its prior operating state. Because the engines were operating at less than 50 per cent torque, had it not been for the increase in torque due to the advancement of the No. 2 power lever - which completed the conditions required for activation of the AUPC - the No. 2 engine/propeller system likely would have recovered to its prior operating state as well. Had the engines been operating at more than 50 per cent torque, both propellers could have been forced to operate on the overspeed governors. This is not a critical event, and the aircraft could have continued flight to an aerodrome under this condition.

Following an intentional in-flight shutdown and subsequent in-flight re-start of the No. 2 engine, the crew found that the propeller could not be unfeathered. None of the aircraft manuals or training material, nor the FlightSafety International pilot training course inform flight crews that, when the PEC detects a failure condition that requires propeller feathering, the ability to subsequently unfeather the propeller will be lost, even in the case of a greater emergency. Just such a greater emergency occurred with this same aircraft about 30 flight hours later (TSB report A05P0137) when the No. 1 engine suddenly shut down in flight without warning.

Findings as to causes and contributing factors

1. During the flight test manoeuvre, both aircraft engines developed a loss of oil pressure when exposed to a high pitch-up angle along with a brief and small negative g condition. When exposed to low oil pressure conditions, the modified engine oil pressure regulating valve allowed the low oil pressure condition to exist long enough to affect propeller operation.
   
2. The extended low oil pressure condition resulted in propeller speed fluctuations, and caused both propellers to enter an underspeed condition. This fulfilled one of three requirements for the subsequent No. 2 propeller overspeed condition.
   
3. While the No. 2 propeller was in the underspeed condition, advancement of the No. 2 power lever resulted in the engine torque exceeding 50 per cent. This fulfilled the second requirement for the subsequent No. 2 propeller overspeed condition. The occurrence of these two conditions simultaneously for more than one second completed all of the requirements for activation of the No. 2 automatic underspeed protection circuit (AUPC) and resulted in the No. 2 propeller overspeed condition.

Findings as to risk

1. Any load factors below 0 g, such as during a training exercise or an actual traffic or terrain avoidance manoeuvre, could involve much higher risks if an unsuspecting flight crew became distracted by a second unexpected event such as a single or double loss of engine oil pressure and subsequent propeller overspeed condition(s).
   
2. None of the aircraft manuals or training material, nor the FlightSafety International pilot training course, inform flight crews that, when the propeller electronic control (PEC) detects a failure condition that requires propeller feathering, the ability to subsequently unfeather the propeller will be lost, even in the case of a greater emergency.

Other Finding

1. Information was not recovered from the cockpit voice recorder (CVR) since it was overwritten while external electrical power was applied to the aircraft following the incident flight.

Safety action taken

Following this incident, Transport Canada (TC) cancelled the flight permit until it could be shown that adequate inspections had determined that the aircraft was safe to operate. A new flight permit was issued on 10 June 2005 for the balance of the flight test program.

Cascade Aerospace Inc. instituted a self-imposed restriction limiting the balance of the flight test program to an intentional manoeuvring minimum limit of +0.5 g as opposed to the previous self-imposed limit of 0 g.

On 21 July 2005, the TSB Pacific regional office distributed two occurrence bulletins. One addressed the issue of the availability of information regarding the engine pitch-up limitation. The other addressed the issue of the availability of information to flight crews regarding the inability to unfeather a propeller subsequent to a fault condition that requires a propeller to be feathered.

This report concludes the Transportation Safety Board's investigation into this occurrence. Consequently, the Board authorized the release of this report on 02 August 2006.