Failure Mode Investigation and Redundancy Management of an Electromechanical Control Actuator for Launch Vehicle Application

Abstract

Electromechanical actuators (EMAs) are increasingly being used in aircrafts under the More Electric Aircraft initiatives. New satellite launch vehicles are being configured with EMA for the thrust vector control and mixture ratio control systems. Analysis of failure modes and their effects in the design stage of a new system is an essential exercise to be done for safety–critical applications of aircraft and human-rated launch vehicle. The thrust vector control system used in the terminal stage of a satellite launch vehicle which does not have any fault tolerance, referred hereafter as no redundancy actuator (NRA), is considered for the case study. At first, a thorough study is conducted on the failures encountered in the NRA and the corrective actions to be taken to avoid these failures in the new design are identified. Improvement in fault tolerance is attempted in two phases. In the first phase, the redundancy features are incorporated in the electrical systems of the actuator such as electric motors and sensors which form part of the closed loop system, leading to the electrical redundancy actuator (ERA). The next logical step is to use a high redundancy actuator (HRA) having fault tolerance to stuck and loose failures in its moving elements such as ball bearings, ball screws and motors. In this paper, failure mode and effects analysis (FMEA) of the existing and the two proposed configurations is carried out. The HRA is designed following the failure mode avoidance technique through four strategies, which are relaxing a constraint limit, using physics of incipient failure, having different operating modes and exploiting dependencies among failure modes. It is concluded that the HRA configuration eliminates a number of single failure points present in NRA and ERA, and it is found to have a better fault tolerance to the failures in its electrical and mechanical elements.