A synthetic pid gain tuning framework for robust pitch attitude control of slender airframes

Abstract

The need to decentralize microsatellite launch missions, as a result of rapid expansion of space technology has led to the emergence of  slender (microsatellite) launch vehicles (SLV). However, navigation cost, in terms of onboard equipment, state estimation and control  algorithms presently prohibits microsatellite launch vehicle missions. Thus, to realize mission affordability of slender launch vehicles, this  work developed a Hardware In the Loop rig for optimizing hardware cost and maximizing the performance of implemented state estimation and control algorithms on slender launch vehicles navigation systems. Apriori, the National Agency for Space Research and  Development Agency scaled the characteristics of NASA's Ares I Rocket launcher to obtain a miniaturized slender launch vehicle. This  prototype is interfaced with MATLAB’s SIMULIINK environment to build an experimental rig for autopilot simulation to realize affordable  navigation systems on slender launch vehicles. In the feedback control loop of the simulated autopilot system, the proportional, integral  and derivative control gains of the simulated autopilot were initialized by classical control laws; this seamlessly transits to a smart fuzzy  logic based gain selection algorithm within the rise time of the system’s response. This smartly filters nonlinear structural vibration noise  from the state estimation system, as well as proactively selects the proportional, integral and derivative gains of the autopilot system.  Inference from the flight data sheet established rigorous coupling between structural and control hardware dynamics. Thus, to demonstrate structural interference cancellation, and improve on the autotuning ability of the semi- intelligent pitch attitude control  algorithm; a preplanned rocket trajectory of 700m altitude and 15 seconds flight duration was modelled for adaptive tracking such that  the desired control objectives are realised. In profile, the realized trajectory indicated that dynamic interaction between rocket structure  and control hardware was effectively attenuated. Inflight, the recorded maximum deviation from the referenced trajectory is 0.16% (overshoot). This transient error is mostly due to unmodelled wind induced structural excitation.