The proposal outlined a design process coupled with a full evaluation and testing phase for a fully electric replacement for the current mechanical-hydraulic rotor brake system that exists on most large rotorcraft in production currently allow regenerative capture of the inertial energy stored in the aircraft drivetrain during deceleration.
The concept offered further potential alongside other similar projects to downsize and in the long term even remove the hydraulic circuit that exists on aircraft resulting such that the benefits of electrical systems becomes significant. Furthermore, once on the aircraft, installation of the equipment provides potential for several performance and system benefits as there would exist coupled to the rotor a bi-directional power flow controller within an independent energy store. Possible systems would include fast rotor start, auto-gyro assist for safe landing in the event of primary power failure and extra power on-board electrical power generation for redundancy during flight.
This report highlighted the technical developments made in the field of an integrated, optimised electric regenerative braking system for the main rotor of a medium-sized helicopter.
The majority of rotorcraft use mechanical heat dissipating braking systems similar to those used on modern cars in order to decelerate the main rotor which stores considerable energy. The result is that the entire main rotor kinetic energy is lost as heat to the surrounding environment every time that the helicopter lands. In a medium sized aircraft, the energy stored in the rotor when it is at full speed is equivalent to roughly 2MJ of energy. Electrical machines can be used to decelerate the main rotor of the aircraft regeneratively reducing the proportion of energy wasted upon every landing. At the time of writing, no electrical system of significant power exists on medium sized aircraft and so this provides an opportunity to explore not only the specific technologies that could be used but also understand the reliability and certification issues as well as opportunities for closed integration possible with these new, modern systems.
The work contained in this project has provided a first step in providing design frameworks and understanding and demonstrating solutions for the use of significant electric power based systems in a rotorcraft.
A design process in which an optimised electric machine and power electronic converter is required that integrates tightly with the surrounding platform was developed and assessed. Tools for predicting performance of the target electric drive were developed and validated. The benefits and shortcomings of using such tools were presented and discussed. Technical highlights in terms of the electric drive design were presented and their performance analysed.
A static braking system for rotor holding, not dissimilar in nature to the existing hydraulic/mechanical systems used currently on aircraft was conceptualised considering the existing system's shortcomings. One of the principal factors addressed was that of the brake coming on in flight. This was tackled from a conceptual and reliability perspective. The concepts were benchmarked and the most promising taken forward for prototyping and characterisation. The final design was presented and its performance evaluated. New data for materials used in the braking process were characterised and a novel framework for the design of such a platform presented.
Tools for assessing the impact of integration of the electric drive system into the existing drivetrain had been developed and used to rationalise the design where appropriate or highlight areas for further investigation. Dynamic analyses demonstrated that removing the need to rigidly constraint/mount the electric machine part of the electric braking system on both sides would mitigate excessive vibration in the machine from the main drivetrain. System level simulations had been developed that allow some scenarios to be played out without the need for the entire toolchain developed as part of the design process of this project. This allowed scenarios such as electric start of the main rotor to be investigated and improved the case for switch from a hydraulic system to an electric equivalent as part of a wider transition to electric systems.
Whilst the technology developed was not considered safety critical in terms of its availability, it does lie on the primary mechanical driveline for the rotorcraft. Hence in order to demonstrate its suitability in this respect a full reliability and fault analysis was undertaken. The tools developed are new in terms of electric power systems on rotorcraft and provided a capability not only to assess this but other electric platforms being considered for rotorcraft.
Finally, the prototype systems for each of the developed technologies were presented and characterised. The electric machine met its primary requirement of providing a flexible non-contact braking capability and is as designed. The machine fit into the target volume, however the system’s ability to generate power during flight is seen to be less than expected. This was thought to be as a result of a mechanical interface in the machine that has provided extra resistance in the thermal path that is between the primary heat source and the main heat sink. Whilst this does not compromise the ability of the machine to generate torque and power it does reduce the ability of the machine to dissipate that heat via the integrated gearbox mounting and would need to be addressed prior to adoption on an aircraft.