This paper details the process involved in the numerical optimisation of a helicopter engine inlet electrothermal ice protection system. Although the process was developed using a production aircraft, it is demonstrated here using a generic intake and flight conditions, due to confidentiality of the actual design. The process includes adherence to the overall system design objectives (maximum power demand), including tolerances required to account for an industrial system (aircraft voltage variation, manufacturing tolerances). The numerical optimisation was performed using a combination of 2D and 3D methods to define the required heated area, power density, locations and settings for temperature control sensors. The use of 2D design tools allows a rapid iteration process to be performed, leading to the possibility of a higher level of optimisation within the allowable time-frame compared to the use of full 3D methods. Different flight settings (different Angle of Attack and Angle of Sideslip) were analysed, for two different droplet diameters (MVD values of 20 microns and 40 microns, coherent with the requirements of the CS-29 Appendix C icing envelope). An iterative transient thermal analysis was then performed, over the required temperature range of 0°C to -30°C, to optimise the heater placement and heater powers, assessing the predicted level of ice accretion in all cases. Assessment in dry conditions in the range of +5°C to -30°C was also performed in order to ensure that structural over-temperature did not occur. The positions of embedded temperature sensors were defined as part of the analysis, including the temperature control values which would feed into the controller. The process was deemed to be successful as the final design was successfully tested in flight and in an icing wind tunnel facility.
