The sensor-based incremental nonlinear dynamic inversion (INDI) using angular acceleration measured by inertial measurement unit (IMU) sensor is a very robust control method on various model uncertainties when the aircraft maneuvers with moderate angle-of-attack (AoA) and high gravity in transonic speed flight conditions. However, the measured angular acceleration has time delay characteristics due to actuator and aircraft dynamics, IMU sensor dynamics, differential angular rate and structural coupling filter (SCF) and so on. These characteristics of angular acceleration feedback reduce dramatically the stability margin of the control system. In this paper, we propose the synchronization filter design method of the control surface feedback path for improving stability margin, based on the proposed hybrid INDI control method using error between the angular acceleration measured from IMU sensor and the angular acceleration calculated from on-board model (OBM) and control surface feedback. To evaluate the proposed control method, we perform the frequency-domain linear analysis and the time-domain simulation. As a result of the evaluation, synchronization method of control surface feedback not only improves the stability margin characteristics of the control system but also eliminates the structural coupling in low frequency range by designing the control surface command feedback using actuator command which is the output of flight control computer (FLCC).
The highly maneuverable fighter aircraft is exposed to unexpected pitch motion such as over-Nz (normal acceleration) and Nz-drop characteristics in transonic and supersonic flight conditions with moderate angle of attack. These characteristics not only degrade flying qualities by destabilizing the aircraft but also threaten flight safety by increasing the structural load. This article proposes an additional augmentation control in the incremental nonlinear dynamic inversion structure, which feeds back the error of pitch angular acceleration to mitigate unexpected pitch motion in slow down turn maneuver. We evaluate the stability, flying qualities, and robustness of the proposed control system by performing the frequency-domain linear analysis and the time-domain numerical simulations based on the mathematical model of advanced trainer aircraft. As a result of the evaluation, the additional augmentation control further improves flying qualities and deceleration performance of the aircraft by decreasing over-Nz and Nz-drop characteristics in high-Nz maneuvering in the transonic flight condition as well as ensuring the stability and robustness of the control system against the major uncertainty factors of the aircraft system compared to the existing transonic pitching moment compensation (TPMC) control in which the predefined scheduling for Nz feedback is used.
The fighter aircraft uses several different loading configurations for air-to-surface and air-to-air combat missions. To maintain wings-level flight with an asymmetric weapon configuration, a pilot controls a roll trim system. However, it is difficult to apply an accurate roll trim input for wings-level flight in the actual flight under disturbance. The inaccurate roll trim input degrades the performance of the roll autopilot system. In this paper, to solve this problem, an integrator was additionally designed in the command part of the roll autopilot system. The initial transient response was improved by scheduling the limiter to restrict the roll attitude error. As a result of the evaluation of the simulation for the designed flight control law, the roll attitude following performance was found to be improved in the autopilot system operation under the inaccurate roll trim condition.
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