Aircraft Attitude Control via Angular Acceleration Feedback
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Solution Overview
Problem
Traditional PID control methods for multi-rotor aerial vehicles are complex, lengthy, and prone to system divergence and instability, with a strong dependence on inner loop calibration, and often fail to achieve optimal disturbance rejection performance.
Innovation Solution
A method and system that calculates aircraft configuration parameters based on physical characteristics, using a feedback control scheme to generate command signals for actuators, incorporating a mixer to combine results from pitch, roll, and yaw axis calculations, and optionally includes an angular acceleration loop for direct feedforward control.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Reliability
If conventional PID tuning methods are used for attitude control, then the control parameters can be adjusted, but the tuning process becomes complex and lengthy with system divergence and instability issues
Solution Approach 1:
The patent implements a feedback control mechanism where the actual attitude is continuously measured and compared with the desired attitude, and the control parameters are automatically adjusted based on the attitude error and its rate of change. This closed-loop feedback system eliminates the need for complex manual PID tuning while ensuring control stability through automatic adaptation.
Solution Approach 2:
The control system performs self-adjustment by automatically calculating optimal control parameters based on real-time attitude measurements and pre-stored aircraft configuration parameters. The system serves itself by eliminating the need for external tuning intervention, thereby simplifying the tuning process while maintaining reliability.
2Reliability
If conventional PID control with sequential tuning is used, then inner loop and outer loop parameters can be adjusted, but the process is lengthy and disturbance rejection performance is suboptimal
Solution Approach 1:
The patent pre-calculates and stores optimal control parameters based on aircraft configuration characteristics before flight. During operation, the system directly applies these pre-prepared parameters or makes minor real-time adjustments, eliminating the need for lengthy sequential tuning of inner and outer loops while ensuring optimal disturbance rejection performance from the outset.
Solution Approach 2:
The system dynamically changes control parameters based on the aircraft's actual attitude error and angular velocity measurements. By continuously adapting parameters according to real-time flight conditions rather than using fixed sequential tuning, the system achieves optimal disturbance rejection quickly without time-consuming adjustment processes.
3Reliability
If traditional control methods are used, then adjustments can be made after disturbances occur, but optimal disturbance rejection cannot be achieved under certain circumstances
Solution Approach 1:
The system uses real-time feedback from attitude sensors and angular velocity measurements to immediately detect disturbances and adjust control parameters accordingly. This continuous feedback mechanism enables the system to respond to disturbances as they occur rather than after they have already affected the aircraft, achieving optimal disturbance rejection with fast response speed.
Solution Approach 2:
The control parameters are made dynamic rather than static, allowing the system to adapt continuously to changing flight conditions and disturbances. By enabling parameters to change in real-time based on measured attitude error and angular velocity, the system achieves both optimal disturbance rejection performance and rapid response speed under varying circumstances.
Data Source
AI summary
A method for controlling an aircraft includes determining a target angular velocity of the aircraft based at least in part on a target attitude of the aircraft, determining a target angular acceleration of the aircraft based at least in part on the target angular velocity, and generating a command signal for at least one propulsion unit of the aircraft based at least in part on the target angular acceleration.