• 한국소음진동공학회논문집
• /
• 제15권7호
• /
• pp.813-820
• /
• 2005

A deployable missile control fin has some structural nonlinearities because of the worn or loose hinges and the manufacturing tolerance. The structural nonlinearity cannot be eliminated completely, and exerts significant effects on the static and dynamic characteristics of the control fin. Thus, It is important to establish the accurate deployable missile control fin model. In the present study, the nonlinear dynamic model of 4he deployable missile control fin is developed using a substructure synthesis method. The deployable missile control fin can be subdivided Into two substructures represented by linear dynamic models and a nonlinear hinge with structural nonlinearities. The nonlinear hinge model is established by using a system identification method, and the substructure modes are improved using the Frequency Response Method. A substructure synthesis method Is expanded to couple the substructure models and the nonlinear hinge model, and the nonlinear dynamic model of the fin is developed. Finally, the established nonlinear dynamic model of the deployable missile control fin is verified by dynamic tests. The established model is In good agreement with test results, showing that the present approach is useful in aeroelastic stability analyses such as time-domain nonlinear flutter analysis.

• 대한전기학회:학술대회논문집
• /
• 대한전기학회 2006년도 심포지엄 논문집 정보 및 제어부문
• /
• pp.219-221
• /
• 2006

The recently developed autopilot controller can make the input-output (I/O) dynamic characteristics of the nonlinear missile dynamics linear and independent of flight conditions. However, significant fin actuator dynamics can degenerate the I/O dynamic performance of the overall system. In this paper, we propose a singular perturbation-like method to compensate the effect of significant fin actuator dynamics in nonlinear missile control. The proposed compensation method does not require the time derivatives of fin angles but can maintain the linear I/O dynamic characteristics provided by the recently developed autopilot controller.

• 제어로봇시스템학회:학술대회논문집
• /
• 제어로봇시스템학회 2003년도 ICCAS
• /
• pp.537-541
• /
• 2003

In order to control the missiles by aerodynamics, control surfaces sometime called fins are used. Deflection angles of these fins are the right control variables of the aerodynamics, but aerodynamicists prefer to use analytic variables called aileron, elevator and rudder instead of these physical variables, because these three analytic variables dominantly influence on the roll, pitch and yaw channels of the missile maneuver, respectively, and each can be assumed a linear combination of four fin deflection angles. On that basis, roll, pitch and yaw autopilots for controlling the attitudes or lateral acceleration of the missile are designed, and as a consequence outputs of each autopilot are aileron, elevator and rudder commands, respectively. In the existing fin control scheme for the typical tail-fin controlled cruciform missiles, firstly these outputs are distributed to four fin defection commands, and after that four fins are actuated by fin controllers so that their deflections follow the commands. This paper shows that performance of such control schemes can be degraded significantly when fin actuators have certain physical constraints such as slew rate, voltage or current limit, uncertainty of actuator dynamics, and so on, and propose a new control scheme which alleviates such problems. This scheme can be widely applied to various fin actuation systems. But in this paper, for convenience, tail-fin controlled cruciform missile is taken as an example, and it is shown that a proposed control scheme gives better performance than the existing one.

• 제어로봇시스템학회논문지
• /
• 제10권7호
• /
• pp.658-668
• /
• 2004

This paper is concerned with a control allocation strategy using the dynamic inversion and the pseudo inverse control which generates the nominal control input trajectories. In addition, an autopilot design method is proposed by using time-varying control technique which is time-varying version of the pole placement of linear time-invariant system for an agile missile with aerodynamic fin and thrust vectoring control. The control allocation proposed in this paper is capable of extracting the maximum performance by combining each control effector, aerodynamic fin and thrust vectoring control. The adopted time-varying control technique for the autopilot design enhances the robustness of the tracking performance for a reference command. The main results are validated through the nonlinear simulations with aerodynamic data.

• 제어로봇시스템학회논문지
• /
• 제2권2호
• /
• pp.102-107
• /
• 1996

It is difficult to analyze the high speed rolling missile with the generally used missile body fixed coordinates. In this study, we formulate the dynamic equations of the high speed rolling missile with the principal axis of inertia, and make the analytical model of one axis steering missile using pitch/yaw symmetry and complex summation method. With this model we analyze the control characteristics and propose the design considerations of high speed rolling missile with one axis control fin using PNG law in conjuntion with a seeker signal.

• 제어로봇시스템학회:학술대회논문집
• /
• 제어로봇시스템학회 2003년도 ICCAS
• /
• pp.525-530
• /
• 2003

This paper is concerned with a control allocation strategy using the dynamic inversion which generates the nominal control input trajectories, and autopilot design using the time-varying control technique which is time-varying version of pole placement of linear time-invariant system for an agile missile with aerodynamic fin and thrust vectoring control. Dynamic inversion can decide the amount of the deflection of each control effector, aerodynamic fin and thrust vectoring control, to extract the maximum performance by combining the action of them. Time-varying control technique for autopilot design enhance the robustness of the tracking performance for a reference command. Nonlinear simulations demonstrates the dynamic inversion provides the effective nominal control input trajectories to achieve the angle of attack command, and time-varying control technique exhibits good robustness for a wide range of angle of attack.

• 한국군사과학기술학회지
• /
• 제6권3호
• /
• pp.122-130
• /
• 2003

This paper is concerned with a control allocation strategy using the dynamic inversion and the pseudo inverse control which generates the nominal control input trajectories, and autopilot design using time-varying control technique which is time-varying version of pole placement of linear time-invariant system for an agile missile with aerodynamic fin and thrust vectoring control. Control allocation of this paper is capable of extracting the maximum performance from each control effector, aerodynamic fin and thrust vectoring control, by combining the action of them. Time-varying control technique for autopilot design enhance the robustness of the tracking performance for a reference command. The main results are validated through the nonlinear simulation.

• 안보군사학연구
• /
• 통권1호
• /
• pp.311-343
• /
• 2003

In this paper, we propose a generalized disturbance torque suppression control scheme of servo motors for missile fin actuators. Our controller consists of both a model based feed-forward controller and a stabilizing feedback controller. The feed-forward controller is designed such that the output of nominal plant tracks perfectly the reference position command with a desired dynamic characteristics. The feedback controller stabilizes the overall closed loop system. Furthermore, the feedback controller contains a free function that can be chosen arbitrary. The free function can be designed so as to achieve both the suppression of disturbances and the robustness to model uncertainties. In order to illuminate the superior performance of our control scheme to the conventional ones, we present some simulation results.

• 한국항공우주학회지
• /
• 제30권7호
• /
• pp.29-35
• /
• 2002

접는 미사일 조종날개의 공력탄성학적 특성을 조사하였다. 접는 미사일 조종날개는 2차원 익형 모델로 가정하였다. 초음속 DPM을 이용하여 초음속 비정상 공기력을 계산하였으며, 최소 상태 변수 근사법을 이용하여 비정상 공기력을 근사화하였다. 선형 및 비선형 플러터 해석을 위해 근궤적법과 시간적분법을 사용하였다. 비선형 플러터 해석을 위해 전개부의 힌지는 비대칭 이선형 스프링으로 가정하였으며, 기술함수를 이용하여 선형화하였다. 플러터 해석으로부터, 비선형 파라미터가 공력탄성학적 특성에 미치는 영향을 조사하였다.

• 한국소음진동공학회논문집
• /
• 제12권10호
• /
• pp.808-815
• /
• 2002

The nonlinear characteristics for hinge of a deployable missile control fin are investigated experimentally. The nonlinearity is caused by a worn or loose hinge and manufacturing tolerance and cannot be eliminated completely. The structural nonlinearity has an effect on the static and dynamic characteristics of the control fin. Therefore, it is necessary to establish the accurate nonlinear model for the hinge of the control fin. In the present study the existence of nonlinearities in the hinge is confirmed from the frequency response experiments such as tip random excitation and base sine sweep. Using the system identification method. especially, ″Force-state Mapping Technique″, the types of nonlinearities are identified and the nonlinear hinge model of the control fin is established.