추진-활공 방식 유도탄의 적응 슬라이딩 모드 기반 조종루프 설계

이용우,김유단,문관영,전병을 한국항공우주학회 2014 한국항공우주학회 학술발표회 논문집 Vol.2014 No.4

추진-활공 구간이 존재하는 유도탄은 급격한 속도변화 및 공력 불확실성이 존재하며, 유도탄의 성능을 보장하기 위해서는 가속 및 감속에 따른 파라미터 변화와 모델 불확실성에 강건한 조종루프가 요구된다. 본 연구에서는 비선형 제어기법인 적응 슬라이딩 모드 기법을 사용하여 공력 핀으로 조종되는 추진-활공 방식 STT(Skid-to-turn) 유도탄의 조종루프를 설계하였다. 이를 위해 유도탄의 속도변화 및 동역학을 수학적으로 모델링하고, 슬라이딩 모드 제어기법을 사용하여 가속도 조종루프를 설계하였으며, 모델 불확실성을 보상하기 위한 적응규칙을 고려하였다. 모델링의 타당성과 설계된 조종루프의 성능을 검증하기 위하여 불확실성을 고려한 수치 시뮬레이션을 수행하였다. A missile during boosting and gliding phase experiences rapid velocity transition and aerodynamic uncertainties. To maintain the performance of the missile, it is required to design a robust autopilot with respect to parameter variation due to acceleration/deceleration as well as to model uncertainties. In this study, an adaptive sliding mode missile autopilot is designed for a fin controlled skid-to-turn missile under boosting and gliding phase. Mathematical modeling of the velocity variation and missile dynamics is performed, and sliding mode based autopilot with adaptation rule is designed. Numerical simulation is carried out to demonstrate the effectiveness of the modeling and performance of the proposed autopilot.

Roll-Pitch-Yaw Integrated H<SUB>∞</SUB> Controller Synthesis for High Angle-of-Attack Missiles

Byung-Hun Choi,Seon-Hyeok Kang,H. Jin Kim,Dae-Yeon Won,Youn-hwan Kim,Byung-Eul Jun,Jin-ik Lee 한국항공우주학회 2008 International Journal of Aeronautical and Space Sc Vol.9 No.1

In this work, we explore the feasibility of roll-pitch-yaw integrated autopilots for high angle-of-attack missiles. An investigation of the aerodynamic characteristics of a surface-to-air missile is presented, which reveals the strong effects of cross coupling between the longitudinal and lateral dynamics. Robust control techniques based on H∞ synthesis are employed to design roll-pitch-yaw integrated autopilots. The performance of the proposed roll-pitch-yaw integrated controller is tested in high-fidelity nonlinear five-degree-of-freedom simulations accounting for kinematic cross-coupling effects between the lateral and longitudinal channels. Against nonlinearity and cross-coupling effects of the missile dynamics, the integrated controller demonstrates superior performance when compared with the controller designed in a decoupled manner.

SDRE 기법을 이용한 STT 유도탄의 자동조종장치 설계

홍유경,김유단,문관영,전병을 한국항공우주학회 2013 한국항공우주학회 학술발표회 논문집 Vol.2013 No.4

본 논문에서는 STT(Skid To Turn) 유도탄의 발사초기 추진단계에서 급격한 속도변화에 의해 제어성능이 저하되는 문제를 해결하기 위해 SDRE 기법을 이용하여 속도변화를 고려한 자동조종장치를 설계한다. 이중루프 설계법을 기반으로 내부루프와 외부루프를 분리하여 SDC(State Dependent Coefficient) 형태로 정리하고 각 시스템에 대해 최적 제어기를 설계한다. 설계된 제어기의 성능을 검증하기 위하여 수치 시뮬레이션을 수행하였으며, 그 결과 파라미터 변화에도 제어성능에 영향이 없이 자세각 명령을 추종할 수 있는 유도탄의 자동조종장치를 설계할 수 있었다. In this paper, to overcome the degradation of STT(Skid To Turn) missile in boost phase, velocity variation is considered and SDRE technique is utilized to design the autopilot. Based on two loop design method, the inner-loop and outer-loop are independently formulated as SDC(State Dependent Coefficient) form and optimal controller is designed. Numerical simulation is conducted to verify the performance of the designed autopilot, and the simulation result shows that the autopilot is not affected by variations of parameter and makes the missile follow commanded attitudes in the boost phase.

Missile Autopilot Design for Agile Turn Control During Boost-Phase

Ryu, Sun-Mee,Won, Dae-Yeon,Lee, Chang-Hun,Tahk, Min-Jea The Korean Society for Aeronautical and Space Scie 2011 International Journal of Aeronautical and Space Sc Vol.12 No.4

This paper presents the air-to-air missile autopilot design for a $180^{\circ}$ heading reversal maneuver during boost-phase. The missile's dynamics are linearized at a set of operating points for which angle of attack controllers are designed to cover an extended flight envelope. Then, angle of attack controllers are designed for this set of points, utilizing a pole-placement approach. The controllers' gains in the proposed configuration are computed from aerodynamic coefficients and design parameters in order to satisfy designer-chosen criteria. These design parameters are the closed-loop frequency, damping ratio, and time constant; these represent the characteristics of the control system. To cope with highly nonlinear and rapidly time varying dynamics during boost-phase, the global gain-scheduled controller is obtained by interpolating the controllers' gains over variations of the angle of attack, Mach number, and center of gravity. Simulation results show that the proposed autopilot design provides satisfactory performance and possesses good [ed: or "sufficient" or "excellent"] capabilities.

Missile Autopilot Design for Agile Turn Control During Boost-Phase

류선미,원대연,이창훈,탁민제 한국항공우주학회 2011 International Journal of Aeronautical and Space Sc Vol.12 No.4

This paper presents the air-to-air missile autopilot design for a 180° heading reversal maneuver during boost-phase. The missile’s dynamics are linearized at a set of operating points for which angle of attack controllers are designed to cover an extended flight envelope. Then, angle of attack controllers are designed for this set of points, utilizing a pole-placement approach. The controllers’ gains in the proposed configuration are computed from aerodynamic coefficients and design parameters in order to satisfy designer-chosen criteria. These design parameters are the closed-loop frequency, damping ratio, and time constant;these represent the characteristics of the control system. To cope with highly nonlinear and rapidly time varying dynamics during boost-phase, the global gain-scheduled controller is obtained by interpolating the controllers’ gains over variations of the angle of attack, Mach number, and center of gravity. Simulation results show that the proposed autopilot design provides satisfactory performance and possesses good [ed: or “sufficient” or “excellent”] capabilities.

Missile Autopilot Design for Agile Turn Control During Boost-Phase

Sun-Mee Ryu,Dae-Yeon Won,Chang-Hun Lee,Min-Jea Tahk 한국항공우주학회 2011 International Journal of Aeronautical and Space Sc Vol.12 No.4

This paper presents the air-to-air missile autopilot design for a 180° heading reversal maneuver during boost-phase. The missile’s dynamics are linearized at a set of operating points for which angle of attack controllers are designed to cover an extended flight envelope. Then, angle of attack controllers are designed for this set of points, utilizing a pole-placement approach. The controllers’ gains in the proposed configuration are computed from aerodynamic coefficients and design parameters in order to satisfy designer-chosen criteria. These design parameters are the closed-loop frequency, damping ratio, and time constant; these represent the characteristics of the control system. To cope with highly nonlinear and rapidly time varying dynamics during boost-phase, the global gain-scheduled controller is obtained by interpolating the controllers’ gains over variations of the angle of attack, Mach number, and center of gravity. Simulation results show that the proposed autopilot design provides satisfactory performance and possesses good [ed: or “sufficient” or “excellent”] capabilities.

SDRE 기반 유도탄 조종루프의 효율 및 성능 개선 연구

이재호,이용우,홍유경,김유단,문관영,전병을 한국항공우주학회 2014 한국항공우주학회 학술발표회 논문집 Vol.2014 No.4

SDRE(State Dependent Riccati Equation) 기법은 선형 최적제어기법을 비선형 시스템에 적용할 수 있도록 변형된 제어기법이다. SDRE는 설계과정이 직관적인 장점이 있어 다양한 비선형 시스템의 제어기 설계에 적용되고 있다. 그러나 매 시점마다 Riccati 방정식의 해를 구해야 하는 특성 상 계산에 소요되는 시간이 길고, 상태변수와 관련이 없는 독립변수를 고려하는 방식에 따라 성능 차이가 발생하여 유도탄 조종루프와 같이 빠른 동역학적 특성을 지니는 시스템에는 실시간으로 제어입력을 제공하기 어렵다. 본 연구에서는 이러한 단점을 보완하기 위하여 시스템 차수를 낮추어 Riccati 방정식 자체의 복잡도를 감소시키고, 해를 구하는 주기를 조절하여 계산 효율성을 증대시키기 위한 방안을 제안하였으며 독립변수를 처리하는 적절한 방식을 제안하였다. 제안한 기법의 성능을 검증하기 위해 시뮬레이션을 수행하였다. With its intuitive design process and widely applicable structure resembling linear quadratic regulator (LQR), the SDRE(State Dependent Riccati Equation) method has been widely applied to various control system design. However SDRE require the solution of the Riccati equation for every time step, which consumes considerable amount of computation. Also the performance of the controllers varies as the treatment on the state-independent variables. The drawbacks not only degenerate its performance and computational efficiency, but also cause the time-delay on control input generation. Therefore it is not easy to implement SDRE to the relatively fast system such as missile autopilot. In this study, the methods for improving efficiency and performance of the SDRE based missile autopilot are proposed. Numerical simulations are carried out with a nonlinear missile model to verify proposed method.

Investigation on Physical Meaning of Three-loop Autopilot

이창훈,Shaoming He,Ju-Hyeon Hong 제어·로봇·시스템학회 2020 International Journal of Control, Automation, and Vol.18 No.11

This paper aims to reveal the hidden physical meaning or the working principle of the three-loop autopilot. First, the minimal control structure that ensures the desired dynamic characteristics is analyzed. The results show that the three-loop autopilot has a redundant feedback loop from the minimal control structure standpoint. Based on this observation, its physical meaning is interpreted by utilizing the closed-loop tracking error dynamics in the presence of aerodynamic uncertainties. It turns out that the three-loop autopilot has the minimal control structure with an additional command which is based on the instantaneous direct model reference adaptive control. A clear mechanism of how the aerodynamic uncertainties are compensated is provided in the three-loop autopilot. Finally, numerical simulations are performed to validate our findings. The analysis results in this paper show the beauty of the three-loop autopilot: it can be considered as the most concise form to achieve desired dynamic characteristics and counteract model uncertainties. The potential importance of the results obtained is that they allow us toproperly design the three-loop autopilot by reflecting its physical meaning and have potential directions to improve the performance of the three-loop autopilot as well as the adaptive control.

바람축 롤각을 고려한 궤환선형화 기반 유도탄 조종루프 설계

이석원,김유단,이용우,홍유경,전병을,문관영 한국항공우주학회 2013 한국항공우주학회 학술발표회 논문집 Vol.2013 No.11

상태변수로 6 자유도 STT(Skid To Turn) 유도탄의 조종루프를 설계하기 위해서는 일반적으로 받음각, 옆미끄럼각과 자세각을 사용한다. 동체 자세각 대신 바람축 롤각을 사용하여 조종루프를 설계하면 직접 자세각을 조종할 때와는 달리 불필요한 동체회전이 발생할 수 있다. 따라서 자세각과 바람축 롤각 사이의 관계식을 고려하여 적절한 자세제어 명령을 생성하는 것이 필요하다. 본 연구에서는 동체축 롤각 안정화를 위하여 바람축 롤각 명령을 사용하는 궤환선형화 기반 STT 유도탄 조종루프를 설계하였다. 수치 시뮬레이션을 통해 설계된 바람축 롤각 명령과 기존의 조종루프의 성능을 비교하였다. Angle of attack, side slip angle, and bank angle are widely used as the states to design an autopilot of 6-DOF STT (Skid To Turn) missile autopilot. When aerodynamic bank angle is used to design the autopilot instead of the roll angle, an unnecessary roll maneuver may occur. Therefore, it is necessary to generate the appropriate aerodynamic bank command by using the relation between body angle and aerodynamic angle. In this study, feedback linearization based STT Missile autopilot is designed using aerodynamic bank angle. Numerical simulation is performed to verify the performance of the proposed controller.

파라미터 불확실성을 고려한 적응 슬라이딩 모드 제어기법 기반 유도탄 조종루프 설계

이용우,김유단,문관영,전병을 한국항공우주학회 2014 한국항공우주학회 학술발표회 논문집 Vol.2014 No.11

대부분의 추진-활공 방식 유도탄은 발사 시점부터 커다란 속도 변화가 발생한다. 이에 따라 공력 특성이 급격히 변화하고, 외란에 의해 다양한 파라미터 불확실성이 발생할 수 있는데, 이러한 요소들로 인한 유도탄 조종성능 저하를 줄이기 위해서는 조종루프를 설계하는 과정에서 이들을 적극적으로 고려해야한다. 본 연구에서는 속도변화 구간이 존재하는 추진-활공 방식 Skid-to-turn (STT) 유도탄의 롤-피치-요 통합 조종루프를 설계하였다. 비선형 제어기법으로 상태변수 불확실성에 강건하다고 알려진 슬라이딩 모드 기법을 사용하였고, 모델링 오차 및 공력특성 변화, 외란에 의한 불확실성을 고려하기 위하여 르야프노프 기반 적응규칙을 설계하였다. 제안된 조종루프의 성능을 검토하기 위하여 수치 시뮬레이션을 수행하였다. Guided missiles are generally operated in wide operating range accompanied by variations in aerodynamic characteristics and parameter uncertainties. Since the effects may affect the control performance of the autopilot, they should be considered in the design process of autopilot. In this study, a roll-pitch-yaw integrated autopilot is designed for a velocity-varying missile incorporating boosting and gliding phase. The sliding mode control scheme, which is robust with respect to parameter uncertainties, is applied. A two-loop structure is considered to deal with the non-minimum phase phenomena of the fin-controlled missile. In addition, Lyapunov method based adaptation rule is proposed to estimate the parameter uncertainties. Numerical simulation is performed to demonstrate the performance of the proposed controller.