# Selection Methodology of an Electric Actuator for
## 1 Introduction
A \"local electro-hydraulic generation system\" (LEHGS) is being used as backup for the purpose of steering and nose landing gear actuation in Airbus A380, but landing gear actuation on this aircraft still has numerous hydraulic power sources [6]. Small aircraft or Unmanned Ariel Vehicles (UAVs) usually rely on electrically operated flight control systems instead of conventional Flight control. In such a scenario, landing gear would be the only component requiring hydraulic system for its operation if not shifted to an electrically operated actuation system, which is far from optimized aircraft design philosophy [7]. Hence, the present research work aims to select a landing gear retraction mechanism for nose landing gear that is based on electric drive, which meets all of the design requirements. In recent years, there is a trend of increased use of the electric actuation system in the aerospace, which is known as power-by-wire (PBW) Electric Aircraft actuation. There are different approaches of PBW technologies and applications, which are extended to landing gear, flight control, actuation system, and Thrust Vector Control (TVC) [8]. The requirement and functional need of actuation in landing gears are different, which includes the rising and lowering the landing gears, wheel braking, and nose-wheel steering [9]. The primary limitation toward electrically actuated landing system lies in aircraft safety. Decades of working experience with hydraulics in automobiles and aircraft has made it reliable and safe for operations. Owing to the use of conventional actuation systems for quite a long time, a lot of industrial innovations have been made in the development of these actuation systems. Moreover, a lot of technical expertise is also available for the maintenance of these systems. Despite these advantages, peculiar risks are always associated with the use of hydraulic as a power source. The risk of fire from the inflammable fluids used in the system and their general messiness bring a lot of criticism to the use of hydraulic in the aircraft system. Moreover, the desire to lessen aircraft airframe noise and weight has shifted the balance in favor of the electrical actuation mechanism. They are designed to run more efficiently at extreme temperature ranges when compared to the hydraulic, which requires specific temperature ranges for effective working. The use of electrical actuators will additionally give the capability to lessen the maintenance and operational costs. The Electro-Mechanical Actuator (EMA) mostly consists of \"Mechanical Drive Assembly\" (i.e., an actuator module, which includes mechanical components and a servomotor), which helps in converting the electrical energy to mechanical energy, and an \"Electronic Control Unit (ECU)\", which is required for control signals and power. Figure 1 shows the basic layout of an EMA. The power control box and electronic control box, combined, make up the ECU of the actuator, while the motor, gearbox, and screw mechanism make up the mechanical drive assembly of actuator. Electrical motor is the fundamental part of electro-mechanical actuator that coverts the electrical energy to mechanical energy. The motor selection usually depends upon the available power supply. Permanent Magnet Synchronous Motor (PMSM) is the best option for the system in which AC power source is directly available [8]. It is used to control the direction and speed of winding by the electric switch, which, in turn, is controlled by electronic power devices. The servo motor have compatibility of high speed with high power density, and also offers high reliability with good heat dissipation which is the most important requirement of the flight control system. Although, there are many types of motors that are suitable, but four common motors options are as follows:  
* the Switched Reluctance (SR) motor;
* the Brushless DC motor;
* permanent Magnet Synchronous Motor (PMSM); and,
* brushed DC motor.