MIMO conditional integrator control for unmanned airlaunch

2012 American Control Conference (ACC), 2012

A satellite launching procedure known as airlaunch is modeled and simulated at the staging phase, where a reusable unmanned aerial vehicle airlaunches a second (rocket) stage. This allows to study the effects of the split phase on the stability of the airlaunch system. A robust conditional integrator controller is designed with the objective of stabilizing the system during and after the airlaunch. The controller is indeed able to assure system stability for rather large disturbances. Performance of the proposed control algorithm is illustrated through simulations.

2012 IEEE 51st IEEE Conference on Decision and Control (CDC), 2012

A Multiple Input Multiple Output (MIMO) controller based on the dynamic feedback linearization technique is designed for the robust global stabilization of a new satellite launching strategy called (unmanned) airlaunch. This strategy consists in using a two-stages launching system. The first stage is composed of an airplane (manned or unmanned) that carries a rocket launcher which constitute the subsequent stages. The control objective is to stabilize the aircraft in the launch phase. It is developed and is applied to the full multi-input multioutput model of the aircraft. The considered model is highly nonlinear, mostly as a consequence of possible large angle of attack, sideslip and roll angle. Finally, the present work illustrates through simulations the good performance of the proposed control algorithm.

A Multiple Input Multiple Output (MIMO) controller based on the conditional servo-compensator technique is designed for the robust stabilization of a new satellite launching strategy called (unmanned) airlaunch. This strategy consists in using a two-stages launching system. The first stage is composed of an airplane (manned or unmanned) that carries a rocket launcher which constitute the subsequent stages. The control objective is to stabilize the aircraft in the launch phase. It is developed separately for two nonlinear motion modes of the model, the longitudinal and lateral modes, and is applied to the full multi-input multi-output model of the aircraft. The controller is indeed able to assure system stability for rather large disturbances. Performance of the proposed control algorithm is illustrated through simulations.

2012

A satellite launching procedure known as airlaunch is modeled and simulated at the staging phase, where a reusable unmanned aerial vehicle air-launches a second (rocket) stage. Airlaunch is described by the variations in mass, inertia and aerodynamic coefficients of airlaunch system and their effects on the stability of the airlaunch system at the stage separation. In this paper, this procedure may imply on large variation in angle of attack, sideslip and roll angle of the airlaunch system caused by perturbations on aerodynamic forces and moments during the launch phase. In order to stand these perturbations, it is proposed an LQR controller obtained by optimal and robust control theory. This controller is designed with the objective of stabilizing the system after the airlaunch. Performance of the proposed control algorithm is illustrated through computer simulations.

2011

This work develops a Multi-Input Multi-Output (MIMO) Conditional Integrator (CI) controller to a class of MIMO nonlinear systems, in the case of asymptotically constant references, motivated by airspace applications. These results are then applied to an aircraft control case and are compared by computer simulations with previous results from the literature that represent the starting point of the current paper. The obtained controller allows in a first step finite time semiglobal stability to a residual region, followed by exponential stability since entering this region.

"This paper investigates the use of an L 1 adaptive controller direct approach to solve the attitude control problem of a launch vehicle (LV) during its atmospheric phase of flight. One of the most important difficulties in designing a controller for launch vehicles (LVs) is the widely changing system parameters during launch. Aerospace systems such as aircraft or missiles are subject to environmental and dynamical uncertainties. These uncertainties can alter the performance and stability of these systems. Unknown variations in thrust and atmospheric properties, eccentricities of nozzles, and other unknown conditions cause changes in a system. The L1 adaptive controller ensures uniformly bound transient and asymptotic tracking for the system’s signals – input and output – simultaneously. This adaptive control technique quickly compensates for large changes in the LV dynamics. The effect of feedback gain selection and robustness of this approach against system uncertainties and actuator disturbances are also discussed. The adaptive control method is then simulated with representative LV longitudinal motion. The effectiveness of the proposed control schemes is demonstrated through hardware-in-the-loop simulation."

AIAA Guidance, Navigation and Control Conference and Exhibit, 2008

This paper describes the fundamental principles of launch vehicle flight control analysis and design. In particular, the classical concept of "drift-minimum" and "load-minimum" control principles is reexamined and its performance and stability robustness with respect to modeling uncertainties and a gimbal angle constraint is discussed. It is shown that an additional feedback of angle-of-attack or lateral acceleration can significantly improve the overall performance and robustness, especially in the presence of unexpected large wind disturbance. Non-minimum-phase structural filtering of "unstably interacting" bending modes of large flexible launch vehicles is also shown to be effective and robust.

Journal of Aerospace Technology and Management, 2024

Understanding of various aerodynamic factors involved in flight trajectories is fundamental to design launch vehicles. First and foremost, computer simulation is an efficient way of predicting its behavior in the movement across the atmosphere. Considering that the available Brazilian version of Analysis, Simulation and Trajectory Optimization Software for Space Applications (Astos) does not simulate a controlled vehicle in six degrees of freedom (DoF), the aim of this article is to complement the Astos outcomes, particularly evaluating the trajectory of a controlled launch vehicle from liftoff to orbit injection, considering the model of rigid body dynamics with a six DoF. This approach carried out with an in-house developed simulator called Scott that simulated a multistage launcher with three flight configurations. In the Scott computer program, a launcher was modeled with differential equations in six DoF, coupled axes attitude control system, and aerodynamic coefficients that changed as a function of Mach number. These features improved the results generated by Astos software for the same configurations and the same initial conditions. Additionally, the results provided by Scott were close to actual vehicle in terms of attitude change and Mach number reached.

Advances in Aerospace Guidance, Navigation and Control, 2013

The necessity of high maneuverability and vertical launching require thrust vector control additional to aerodynamic control. That hybrid usage of aerodynamic and thrust vectoring controls effectively increases the agility of the missile against air defense threats. This requirement and the rapidly changing dynamics of this type of missiles renders the guidance and control design critical. However, the findings suggest that classical guidance and control design approaches are still valuable to apply and can have successful performance within the effective flight envelope. It is very rare that a study concerns from detailed dynamics and analysis of the dynamics covering flight mission and algorithms. In this study, together with the modeling of the agile dynamics of a vertical launch surface to air missile and the corresponding thrust forces and moments depending on linear supersonic theory, the application of the flight control algorithms are presented. Two classic linear autopilot structures are studied. During autopilot design process, an additional term related to short period dynamics of boost phase is proposed and the drastic effect of this term is shown. In addition to control algorithms, guidance algorithms are also defined to fulfill the mission of the missile. Body pursuit algorithm is applied for rapid turnover maneuver and midcourse guidance. Proportional navigation guidance is chosen for terminal phase. In addition, an alternative maneuvering technique is proposed to reduce further side slip angle during vertical flight.