• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.50 no.6
• /
• pp.393-400
• /
• 2022

The Divert and Attitude Control System(DACS) used in high-altitude engagements is expensive and complex. In this paper, we design a high-altitude terminal guidance and control loop of guided-missile equipped with a Thrust Vector Control(TVC) that is less expensive and simpler than DACS. The proposed system utilizes a quaternion feedback control technique to track the thrust attitude command converted from the acceleration command of true proportional navigation guidance. The performance analysis of the proposed terminal guidance and control loop is conducted through engagement simulations against ballistic targets at a high altitude.

• Journal of the Korean Society of Propulsion Engineers
• /
• v.19 no.2
• /
• pp.9-16
• /
• 2015

The control law for simultaneous pressure and thrust control of solid DACS(Divert Attitude Control System) is suggested. To regulate the two variables effectively, the control structure of sequential loop closer is applied to the system considering the physical characteristics of each variable and the weighted pseudo-inverse method is suggested to allocate effective command for indeterminate system. Also, the pressure guidance law for safe and high acceleration is applied to the homing stage to verify the effectiveness of the command distribution.

• Journal of the Korean Society of Propulsion Engineers
• /
• v.21 no.5
• /
• pp.61-70
• /
• 2017

Robust thrust distribution method of solid DACS is researched. For the case of the system which has higher number of actuation nozzles than the degree of freedom of thrust to be controlled, the robust thrust allocation law which accommodate the abnormal operation is suggested. Assuming the situation that some nozzles are uncontrollable, the error between nozzle throat area command and response can be calculated. The error is used for realtime reshaping of weighting matrix. From the weighting effect, the nozzle which operated abnormally has low responsibility for the command then, the thrust error is reduced. The suggested algorithm is verified by the simulation of abnormal operation condition of DCS and ACS nozzle respectively.

• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.32 no.6
• /
• pp.72-80
• /
• 2004

In this paper, the design procedure of a guidance algorithm in the boosting phase of missiles with free-flight after thrust cut-off is introduced. The purpose of the guidance is to achieve a required velocity vector at the thrust cut-off. Trajectory optimizations for four cost functions are performed to investigate implementable trajectories in the pitch plane. It is observed from the optimization results that high angle of attack maneuver in the beginning of the flight are required to satisfy the constraints. The proposed guidance algorithm consists of the pitch program to produce open-loop pitch attitude command and the yaw attitude command generator to nullify the velocity to go. The pitch program utilizes the pitch attitude histories obtained from the trajectory optimization.

• Journal of Institute of Control, Robotics and Systems
• /
• v.20 no.11
• /
• pp.1103-1111
• /
• 2014

The results of control law design for a tilt-rotor unmanned aerial vehicle that has a nacelle mounted wing extension (WE) are presented in this paper. It consists of a control surface mixer, stability and control augmentation system (SCAS), hold mode for altitude / speed / heading, and a guidance mode for preprogram and point navigation which includes automatic take-off and landing. The conversion corridor and the control moments derivatives between the original tilt-rotor and its variant of the nacelle mounted WE were compared to show the effectiveness of the WE. The nacelle conversion of the original tilt-rotor starts when the airspeed is greater than 30 km/h but its WE variant starts at 0 km/h in order to reduce the drag caused by the high incidence angle of the WE. The stability margins of the inner loop are presented with the optimization approach. The outer loops for the hold mode are designed with trial and error methods with linear and nonlinear simulation. The main control parameter for altitude control of the helicopter mode is thrust command and it is transferred to the pitch attitude command in airplane mode. Otherwise, the control parameter for the speed of the helicopter mode is the pitch attitude command and it is transferred to the thrust command in airplane mode. Therefore the speed and altitude hold mode are coupled to each other and are engaged at the same time when an internal pilot engages any of the altitude or speed hold modes. The nonlinear simulation results of the guidance control for the preprogrammed mode and point navigation are also presented including automatic take-off and landing in order to prove the full control law.

• Advances in aircraft and spacecraft science
• /
• v.10 no.1
• /
• pp.1-18
• /
• 2023

This paper investigates the integrated control of an air-breathing hypersonic vehicle considering the safety of propulsion system under acceleration. First, the vehicle/engine coupling model that contains a control-oriented vehicle model and a quasi-one-dimensional dual-mode scramjet model is established. Next, the coupling process of the integrated control system is introduced in detail. Based on the coupling model, the integrated control framework is studied and an integrated control system including acceleration command generator, vehicle attitude control loop and engine multivariable control loop is discussed. Then, the effectiveness and superiority of the integrated control system are verified through the comparison of normal case and limiting case of an air-breathing hypersonic scramjet coupling model. Finally, the main results show that under normal acceleration case and limiting acceleration case, the integrated control system can track the altitude and speed of the vehicle extremely well and adjust the angle deflection of elevator to offset the thrust moment to maintain the attitude stability of the vehicle, while assigning the two-stage fuel equivalent ratio to meet the thrust performance and safety margin of the engine. Meanwhile, the high-acceleration requirement of the air-breathing hypersonic vehicle makes the propulsion system operating closer to the extreme dangerous conditions. The above contents demonstrate that considering the propulsion system safety will make integrated control system more real and meaningful.

• Journal of Institute of Control, Robotics and Systems
• /
• v.15 no.2
• /
• pp.240-248
• /
• 2009

Modem version of supersonic jet fighter aircraft must have been guaranteed appropriate controllability and stability in HAoA(High Angle of Attack). The HAoA flight control law have two parts, one is control law of departure prevention and the other is control law of departure recovery support. The control laws of departure prevention for advanced jet trainer consist of HAoA limiter, roll command limiter and rudder fader. The control laws of departure recovery support are consist of yaw-rate limiter and MPO(Manual Pitch Override) mode. The guideline of pitch rocking using MPO mode is simple, but operating skill of pitch rocking is very difficult by the pilot with inexperience of departure situation. Therefore, automatic deep stall recovery system is necessary. The system called the "Automatic Pitch Rocker System" or APRS, provided a pilot initiated automatic maneuver capable of an aircraft recoveries in situations of deep stall, speed and altitude. This paper addresses the design and validation for APRS to recovery of an deep stall without manual pitch rocking by the pilot. Also, this system is designed to recovery of speed, attitude and altitude after deep stall recovery using ATCS (Automatic Thrust Control System) and autopilot. Finally, this system is verified by real-time pilot evaluation using HQS (Handling Quality Simulator).

• Journal of the Korean Society of Propulsion Engineers
• /
• v.19 no.3
• /
• pp.54-64
• /
• 2015

This paper deals with an acceleration controller design for a kill vehicle equipped with a divert and attitude control system (DACS). In the proposed method, the attitude control system (ACS) is used to produce the thrust command to nullify angle-of-attack. For the angle-of-attack control, a nonlinear angle-of-attack controller is proposed based on the feedback linearization methodology. Since the flight path angle is identical to the attitude angle under the condition of zero angle-of-attack, the divert control system (DCS) can directly produce the lateral acceleration which is demanded from the guidance loop. In the proposed method, we can minimize the aerodynamic uncertainty due to the propulsive force. Additionally, we can simplify the operation logic of DCS and ACS. In this paper, nonlinear simulations are performed to show the performance of the proposed method.