Full Electric Helicopter Anti-Torque

On the way to complete electric flight, the electrification of helicopter subsystems is an essential milestone. This paper discusses the design of an electric helicopter anti-torque system, which uses Kopter's AW09 helicopter as a platform and shall be tested in ground tests. Analysis of state of the art anti-torque devices for helicopters has helped to identify concepts, which are suitable to be combined with electric propulsion and actuation. Engineering models are used to estimate the power benefits of varied tail rotor RPM, enlarged and steerable vertical stabilizers and drag reducing devices, which cover the rotor in forward flight. In connection with operational benefits viewed from the OEMs perspective, an architecture is proposed which consists of an electric driven shrouded tail rotor, an electric pitch actuation system and additional aerodynamic surfaces, like a steerable vertical stabilizer and a drag optimized tail rotor cover. The systems were developed according to the results of a safety analysis to meet the requirements of CS-27. The electric tail rotor drive is designed with an internal level of redundancy that allows to compensate for subsystem failures

Development of an integrated electrical swashplateless primary and individual blade control system

It has been shown that a helicopter rotor in forward flight can greatly benefit from more complex pitch control schemes than used today. Thus, control system add-ons have been realized which provide those missing degrees of freedom, usually referred to as Individual Blade Control (IBC). If such systems are added to a conventional control system, however, additional means of actuation have to be placed in the rotating frame, which require their own information and power transfer links. Instead of implementing additional separate hardware it might be worthwhile to consider an integrated actuation system designed to fulfill both the primary control as well as the IBC requirements. This paper discusses the following three aspects: (a) what are the performance requirements of such an integrated control system which is able to fulfill both functions, (b) what actuation technology is best suited for this application and (c) how can the reliability required for a primary flight control system be realized with a minimum degree of redundancy and thus at a competitive system weight? It is shown that by using modern customized electromechanic actuators and by choosing an appropriate system architecture this type of integrated control system becomes feasible at a very competitive system weight. This paper presents the basic principles, discusses the multi-disciplinary investigations and describes the approach to demonstrate its feasibility through dedicated bench tests