Browse Topic: Failure modes and effects analysis (FMEA)
As per Committee/Henry E. Harschburger recommendations
Electric aviation is advancing rapidly, with aircraft from manufacturers like Joby and Archer well on their way to certification, aircraft electrification will continue and begin to apply to larger aircraft. To support larger electrified rotorcraft, rotors will need to grow if disc-loading and hover efficiency are to be maintained. A consequence of this is the need to reduce rotor speed to maintain an acceptable acoustic signature, especially for operation in urban environments. Most current applications utilize radial flux motors, sometimes with a reduction gearbox. Gearboxes can improve overall propulsion system power density by enabling higher motor speeds but are generally not preferred as they introduce additional potential failure modes and maintenance schedules. In this paper a holistic approach is used to understand the trade-offs between rotor and motor and their consequences on propulsion system power density.
Dufour Aerospace designs and manufactures an automated tilt-wing aircraft for critical cargo delivery missions. Emphasizing operational efficiency, the platform integrates path generation and tracking techniques tailored for the unique dynamics of tilt-wing flight and builds upon the existing lower level control. While there exist a myriad of methods for high-level aircraft automation ranging from PID to MPC, they often require a trade-off between complexity and the capability to handle non-linear dynamics of the system they are controlling. Hence, a lightweight, deterministic geometric path generation approach using clothoid-based transitions between three waypoints and a robust SO(3)- based path tracking controller adapted for tilt-wing dynamics are presented. Additionally, a high-level automation framework is introduced that includes failure mode handling for GNSS loss and communication breakdowns. This system ensures mission continuity and operational safety while supporting flexible mission planning. The methods are validated through extensive flight testing on both small and large-scale aircraft. The latter prove the scalability, safety, and reliability of the presented solution for tilt-wing aircraft automation and enhance the aircraft's capabilities in real-world emergency response and complex operational scenarios.
ABSTRACT The use of active inceptor systems allows for control of the aircraft even during mechanical failures within the control inceptor. For the specific case of isometric failure, whereby the inceptor 'freezes' in position, a virtual force displacement model is used to continue to provide control input. Testing on DLR's experimental helicopter (ACT/FHS) has shown the potential to encounter pilot-induced oscillation (PIO) tendencies when flying using this mode. This paper presents results from a simulation campaign undertaken to determine whether PIOs could be exposed through this use of control and/or the resultant severity. The results show that control limiters cause severe PIOs during the isometric failure. Unacceptable failure characteristics were found for six different vehicle configurations and PIOs were exposed by all four pilots. PIO incipience was predicted through the use of offline tools. In future, it is recommended that specific PIO investigations are undertaken during the evaluation of active inceptor failure modes.
Health and Usage Monitoring Systems installed on modern rotorcraft can be used for substantiating component life extensions leveraging actual loads and usage information, reducing uncertainty from assumed loads and usage. One challenge inhibiting fielding such extensions is lack of clearly defined methodology to robustly determine the criticality of the HUMS application. This paper proposes a semi-quantifiable methodology for determining criticality. It is applicable to both loads- and usage-based approaches for part-number-level life adjustments. The approach relies on baseline component design data and is independent of HUMS data to eliminate the possibility of a HUMS errors affecting the bounding criticality assigned to the application. It takes into account the specifics of the component, including the SN working curve, relevant component failure mode severity, and validated compensating provisions that either reduces the likelihood of the component failure mode and/or provides significant "hidden" conservatism not explicitly considered in the assigned uncompensated component failure severity level.
Rolling element bearing failures form one of rotating equipment's most critical failure modes. Vibration analysis has been successfully used for bearing fault detection and diagnostics but does not estimate the spall length of the bearing. An estimate of the spall length would provide insight into the degrading reliability of a drivetrain as the fault propagates. This would improve the timeliness of scheduling a maintenance action. In this paper, a synthetic tachometer signal is generated from the bearing fault itself. It is synchronous to the rolling element, allowing for a time-domain representation of waveform using the time-synchronous average. From this, an estimate of the length of the bearing fault can be determined.
The Eagle Flight Research Center (EFRC) at Embry-Riddle Aeronautical University (ERAU) is investigating the handling qualities of partial and full rotor failure modes of a multi-rotor vehicle testbed employing Distributed Electric Propulsion (DEP) systems intended for Advanced Air Mobility (AAM) vehicles. In order to pave the way for commercial operations, the AAM industry requires a deeper understanding of the handling characteristics and the vehicle's dynamics and controllability under rotor failure conditions. The objective of the research performed at the EFRC centered around designing and testing different thrust and moment control allocation methods for an electric Vertical Take-Off and Landing (eVTOL) vehicle, in addition to assessing their performance in both nominal and failure modes of operation. This paper focuses on analyzing the predicted handling qualities for a full-scale quadrotor testbed vehicle with RPM, collective, and cyclic blade pitch control allocation. The study uses flight-test data to develop a bare airframe dynamic model to optimize the controller gains, which are used to evaluate the predicted handling qualities though development of vehicle specific Handling Qualities Task Elements (HQTE) maneuvers. The evaluation is performed using an in-house built flight simulator together with the Cooper-Harper Rating Scale (CHRS) for Handling Qualities test evaluation. The handling qualities were evaluated and compared for different control strategies and a method for quantitative evaluation of handling qualities is presented.
Emerging vertical flight concepts being proffered for solutions to the Future Vertical Lift (FVL) mission set such as compound high speed rotorcraft can be designed with multiple, coupled control effectors thus creating redundant systems in one or two more axes to generate control forces and moments which allow for a range of trim states. In the FVL mission area future rotorcraft will be asked to fly into high threat environments where potential failure modes can be encountered due to enemy fire or mechanical failure causing reduction of the safe flight envelope. Fault detection creates options to increase the survivability of the crew and passengers allowing an emergency flight envelope to be proposed. One of the more serious potential failures due to enemy fire is a loss of yaw control. Faults in yaw control can be detected in a compound rotorcraft with a vectored thrust ducted propeller (VTDP) or similar anti-torque thruster. An online Kalman filter (KF) for a dimensional yaw moment coeff icient model will be used to estimate vehicle yaw coeff icients. Deviation from the nominal coefficients will be monitored based on the KF statistics in the case of both rudder and tail rotor failure at 60, 40, and 20 ft/s in forward flight. Both frozen zero rudder and ganged sector faults as well as failed tail rotor faults were successfully detected at all airspeeds except the failed tail rotor at 60 ft/s. For the yaw control faults considered, post fault excitation appears airspeed dependent. An online KF estimator for yaw control fault detection could successfully be integrated into the design of a compound rotorcraft with VTDP thereby increasing system safety.
Gear design changes impact on gear crack propagation trajectory is investigated through numerical study. General purpose linear elastic fracture mechanics software, FRANC2D and FRANC3D, are used to simulate 2D and 3D gear crack propagation. FRANC can model non-planner, arbitrary shape crack surface for crack tip stress distributions, stress intensity factors, and crack propagation analyses. Maximum tensile stress and NASGRO4 fatigue crack growth models are employed to predict crack propagation direction and life. Three-dimensional idler gear crack propagation simulation shows the predicted crack trajectory is close to the field observation. Various 2D models are simulated to investigate the crack trajectory impact factors and design strategies to prevent gear rim failure. As shown in previous studies, the initial crack position and orientation play pivot role to control gear failure mode - tooth or rim. For a fixed crack position, this study shows the ratio between bend stress and centrifugal stress dominates gear fracture mode. The less centrifugal stress, the crack more likely to break tooth, while lower bend stress more likely lead to break rim. To prevent rim failure through increasing the rim thickness results in a significant weight penalty. The larger the gear, more the weight penalty. Based on the simulation results, the recommended design strategy is to evaluate gear rim failure risk during the gear train layout phase. It is difficult to be improved at the individual gear design phase.
Items per page:
50
1 – 50 of 1141