Application of Structural Health Monitoring over a Critical Helicopter Fuselage Component

The helicopter design is a challenging experience for fatigue concern as it is subjected to a very wide range of low- and high-frequency load cycles per flight, very much more than a fixed wing aircraft. Moreover, thinking of the various and harsh environments where the helicopter could operate, also corrosion and low velocity impacts could generate further crack nucleation and propagation into the fuselage. Health and Usage Monitoring Systems (HUMS) has received considerable attention from the helicopter community in recent years with the declared aim to increase flight safety and mission reliability, extend duration of life-limited components and of course reduce inspection and maintenance activities. In particular, Structural Health Monitoring (SHM) seems capable to help in reducing the maintenance and operational costs, which is about 25 per cent of the direct operating cost of the helicopter, thus playing an important role especially in the case of the ageing helicopters. In fact, the damage tolerant design approach makes the fatigue resistance evaluation not only a safety issue but also a maintenance related concern. In effect, thanks to the continuous evaluation of the current structural health of the helicopter through a SHM system, it could be possible to set a Condition Based Maintenance, which means substituting a component according to its current structural condition instead of relying just on the design assumptions. The approach could bring to a maximization of both the machine availability and reliability, thus conjugating safety with economy. Strictly connected to a damage tolerant design, a sensor network is thus needed in order to monitor the structural health of the machine and the recent improvements in non-destructive techniques for crack detection are making the concept more affordable from both the technological and the economical points of view. The aim of the present work is to explain a novel method to apply the SHM concepts on a critical zone of a helicopter fuselage, passing through the creation of a complete FE Model of the fuselage, either in healthy and damaged situations, and considering the different stress distributions caused by a progressive crack in the most critical areas. This would represent the key step for the extraction of information from the sensor data, thus allowing to distinguish between the undamaged and damaged structures. The helicopter tail structure is presented herein as a good candidate for the application and testing of the SHM system. The main reason is the criticality of the region, where the torque generated by tail rotor to balance the rotation induced by the main rotor is undergone. In particular, the attention will be focused on some simplified reinforced panels, well suited to indicate the general behaviour of the entire structure and particularly adapt for the safe and early application on board of the machine.