質(zhì)量矩飛行器制導(dǎo)控制問(wèn)題研究
發(fā)布時(shí)間：2018-12-14 13:23
【摘要】：導(dǎo)彈機(jī)動(dòng)控制方式從本質(zhì)上講均通過(guò)調(diào)節(jié)控制力矩實(shí)現(xiàn),具體實(shí)現(xiàn)形式包括兩類:調(diào)節(jié)控制力和控制力臂。傳統(tǒng)舵面控制及噴氣推力控制均屬于調(diào)節(jié)控制力的范疇,但舵面控制難以解決高速機(jī)動(dòng)過(guò)程中的氣動(dòng)燒蝕問(wèn)題,而噴氣推力控制則受限于攜帶燃料有限且導(dǎo)致固液耦合等缺點(diǎn)。質(zhì)量矩控制則屬于調(diào)節(jié)控制力臂的范疇,因其執(zhí)行機(jī)構(gòu)在彈體內(nèi)部避免了舵面控制氣動(dòng)燒蝕等問(wèn)題,又因其無(wú)需攜帶額外燃料解決了噴氣推力控制攜帶燃料有限及固液耦合等缺點(diǎn)。鑒于此質(zhì)量矩控制較傳統(tǒng)控制方式優(yōu)勢(shì)明顯且應(yīng)用前景廣泛,但其獨(dú)特的控制模式增加了導(dǎo)彈空間運(yùn)動(dòng)復(fù)雜程度,為制導(dǎo)控制設(shè)計(jì)帶來(lái)諸多新的挑戰(zhàn)。本文以質(zhì)量矩飛行器再入機(jī)動(dòng)精確打擊為背景,旨在研究雙滑塊/差動(dòng)副翼側(cè)滑轉(zhuǎn)彎(skid-to-turn,STT)質(zhì)量矩飛行器動(dòng)力學(xué)建模、制導(dǎo)控制及其一體化設(shè)計(jì)問(wèn)題,從而為質(zhì)量矩導(dǎo)彈技術(shù)的發(fā)展提供理論支撐。首先,基于多剛體系統(tǒng)建模方法建立了雙滑塊/差動(dòng)副翼STT質(zhì)量矩飛行器完整空間運(yùn)動(dòng)模型。充分考慮質(zhì)量矩導(dǎo)彈的運(yùn)動(dòng)學(xué)耦合、慣性耦合及氣動(dòng)慣性交叉耦合等因素,基于動(dòng)量定理建立了彈體坐標(biāo)系下彈體質(zhì)心平動(dòng)動(dòng)力學(xué)方程,基于動(dòng)量矩定理建立了彈體坐標(biāo)系下系統(tǒng)繞彈體質(zhì)心轉(zhuǎn)動(dòng)動(dòng)力學(xué)方程,同時(shí)建立了再入坐標(biāo)系下彈體質(zhì)心平動(dòng)運(yùn)動(dòng)學(xué)方程及繞彈體質(zhì)心轉(zhuǎn)動(dòng)運(yùn)動(dòng)學(xué)方程。這些方程完整地描述了質(zhì)量矩飛行器空間運(yùn)動(dòng)機(jī)理,揭示了滑塊運(yùn)動(dòng)與彈體質(zhì)心平動(dòng)及繞彈體質(zhì)心轉(zhuǎn)動(dòng)的內(nèi)在聯(lián)系。與傳統(tǒng)舵面控制及噴氣控制相比,質(zhì)量矩飛行器獨(dú)特的控制機(jī)理,即滑塊運(yùn)動(dòng)與彈體運(yùn)動(dòng)之間的內(nèi)在耦合聯(lián)系,使得質(zhì)量矩飛行器空間運(yùn)動(dòng)模型較為復(fù)雜。其次,針對(duì)地面固定目標(biāo)精確打擊問(wèn)題,分別提出了有/無(wú)終端角度約束的有限-r收斂制導(dǎo)律1。1)以彈目距離為參變量描述導(dǎo)彈與目標(biāo)的相對(duì)運(yùn)動(dòng)關(guān)系,建立了新的制導(dǎo)模型。該制導(dǎo)模型包括兩個(gè)微分方程,分別描述了視線俯沖運(yùn)動(dòng)及視線轉(zhuǎn)彎運(yùn)動(dòng),并且視線俯沖運(yùn)動(dòng)微分方程單獨(dú)解耦。基于該模型的制導(dǎo)律設(shè)計(jì)既保證了精確性又簡(jiǎn)化了制導(dǎo)律設(shè)計(jì)過(guò)程。2)基于該制導(dǎo)模型,提出了具有干擾抑制的有限-r收斂制導(dǎo)律,給出了過(guò)載形式的制導(dǎo)指令。與傳統(tǒng)制導(dǎo)律相比,該制導(dǎo)律理論上保證了視線旋轉(zhuǎn)角速率在彈目距離減小至期望值之前收斂為零。仿真驗(yàn)證了該制導(dǎo)律的正確性。3)基于該制導(dǎo)模型,提出了具有終端角度約束的有限-r收斂制導(dǎo)律,給出了過(guò)載形式的制導(dǎo)指令。與傳統(tǒng)制導(dǎo)律相比,該制導(dǎo)律理論上保證了視線角偏差及視線旋轉(zhuǎn)角速率在彈目距離減小至期望值之前收斂為零。與比例導(dǎo)引(proportional navigation guidance,PNG)和最優(yōu)導(dǎo)引(optimal navigation guidance,ONG)的仿真比較結(jié)果驗(yàn)證了該制導(dǎo)律的優(yōu)越性。然后,針對(duì)質(zhì)量矩飛行器姿態(tài)控制問(wèn)題,分別提出了有/無(wú)控制輸入飽和的有限時(shí)間收斂控制律。1)在合理簡(jiǎn)化基礎(chǔ)上建立了質(zhì)量矩導(dǎo)彈俯仰、偏航和滾轉(zhuǎn)三通道獨(dú)立控制模型�；谠摽刂颇Ｐ�,分別設(shè)計(jì)了俯仰、偏航和滾轉(zhuǎn)通道有限時(shí)間收斂姿態(tài)控制律。與傳統(tǒng)質(zhì)量矩導(dǎo)彈姿態(tài)控制律相比,該控制律理論上保證了姿態(tài)角偏差及其變化率有限時(shí)間收斂于原點(diǎn)鄰域。2)通過(guò)引入擴(kuò)張狀態(tài)觀測(cè)器實(shí)現(xiàn)對(duì)擾動(dòng)邊界的自適應(yīng)估計(jì),進(jìn)而設(shè)計(jì)了俯仰、偏航和滾轉(zhuǎn)通道控制輸入飽和的有限時(shí)間收斂姿態(tài)控制律。與傳統(tǒng)質(zhì)量矩導(dǎo)彈姿態(tài)控制律相比,該控制律理論上保證了控制輸入飽和情況下姿態(tài)角偏差及其變化率有限時(shí)間收斂于原點(diǎn)鄰域。通過(guò)特征點(diǎn)仿真和全彈道仿真驗(yàn)證了所提有/無(wú)控制輸入飽和的有限時(shí)間收斂姿態(tài)控制律的正確性。最后,針對(duì)地面逃逸目標(biāo)精確打擊問(wèn)題,分別提出了三通道獨(dú)立/全狀態(tài)耦合制導(dǎo)控制一體化設(shè)計(jì)。1)建立了質(zhì)量矩飛行器俯仰、偏航和滾轉(zhuǎn)三通道獨(dú)立制導(dǎo)控制一體化模型�；隰敯糇赃m應(yīng)反演方法設(shè)計(jì)了一體化控制律,通過(guò)魯棒自適應(yīng)函數(shù)項(xiàng)實(shí)現(xiàn)對(duì)有界未知擾動(dòng)的自適應(yīng)補(bǔ)償,引入非線性跟蹤微分器避免了傳統(tǒng)反演控制的“計(jì)算膨脹”問(wèn)題,理論上證明了系統(tǒng)狀態(tài)跟蹤誤差及擾動(dòng)估計(jì)誤差指數(shù)收斂于原點(diǎn)鄰域。仿真驗(yàn)證了所提三通道獨(dú)立制導(dǎo)控制一體化模型及控制律的正確性。2)建立了質(zhì)量矩飛行器俯仰、偏航和滾裝全狀態(tài)耦合制導(dǎo)控制一體化模型�；谧赃m應(yīng)動(dòng)態(tài)面反演方法設(shè)計(jì)了一體化控制律,通過(guò)魯棒自適應(yīng)函數(shù)項(xiàng)實(shí)現(xiàn)對(duì)有界未知擾動(dòng)的自適應(yīng)補(bǔ)償,引入動(dòng)態(tài)面技術(shù)避免了傳統(tǒng)反演控制的“計(jì)算膨脹”問(wèn)題,理論上證明了系統(tǒng)狀態(tài)跟蹤誤差、邊界層誤差及擾動(dòng)估計(jì)誤差指數(shù)收斂于原點(diǎn)鄰域。仿真驗(yàn)證了所提全狀態(tài)耦合制導(dǎo)控制一體化模型及控制律的正確性。
[Abstract]:The maneuvering control mode of the missile is realized by adjusting the control moment in nature, and the specific implementation form includes two types: adjusting the control force and the control force arm. The control of the conventional rudder surface and the control of the jet thrust belong to the category of the control force, but the control of the rudder surface is difficult to solve the problem of the aerodynamic ablation in the high-speed maneuver, and the jet thrust control is limited by the disadvantages of the limited carrying of the fuel and the coupling of the solid solution. the control of the mass moment belongs to the category of adjusting the control force arm, so that the problem that the rudder surface is controlled by the rudder surface and the like is avoided in the body of the elastic body due to the actuator of the control force arm, and the defect that the jet thrust control carries the limited fuel and the solid solution coupling and the like is solved by the air jet thrust control without carrying the extra fuel. In view of the obvious advantage of this quality moment control and wide application prospect, its unique control mode increases the complexity of the missile space motion and brings many new challenges to the guidance control design. The purpose of this paper is to study the dynamic modeling, guidance control and integrated design of the mass moment of the two-block/ differential aileron sideslip-turn (STT), so as to provide the theoretical support for the development of the mass-moment missile technology. First, a complete space motion model of the two-block/ differential aileron STT mass moment aircraft is established based on the multi-rigid-body system modeling method. Taking full consideration of the factors such as the kinematic coupling, the inertia coupling and the pneumatic inertia cross-coupling of the mass moment missile, the kinetic equation of the elastic body motion in the body coordinate system is established based on the momentum theorem. Based on the momentum moment theorem, the dynamic equations of the body rotation of the system in the body coordinate system are established, and the kinematic equations of the core translation of the elastic body in the re-input coordinate system and the kinematic equations about the core rotation of the elastic body are established. These equations describe the mechanism of the space motion of the mass moment aircraft, and reveal the internal relation between the movement of the slide block and the movement of the elastic body and the rotation of the body of the elastic body. Compared with the conventional rudder surface control and the air-jet control, the unique control mechanism of the mass moment aircraft, that is, the inherent coupling between the movement of the slide block and the body motion, makes the space motion model of the mass moment aircraft more complex. Secondly, aiming at the problem of target fixed target, the finite-r convergent guidance law with/ without terminal angle constraint is proposed. The relative motion between the missile and the target is described by using the target distance as the reference variable, and a new guidance model is established. The guidance model includes two differential equations, respectively describing the line-of-sight subduction motion and the line-of-sight turning motion, and the line-of-sight subduction motion differential equation is separately decoupled. The guidance law design based on the model not only ensures the accuracy but also simplifies the guidance law design process. 2) Based on the guidance model, a limited-r convergence guidance law with interference suppression is proposed, and the guidance instruction in the form of overload is given. Compared with the traditional guidance law, the guidance law theoretically ensures that the rotation angle rate of the line of sight converges to zero before the target distance is reduced to the expected value. The validity of the guidance law is verified by the simulation. 3) Based on the guidance model, a limited-r convergence guidance law with terminal angle constraint is proposed, and the guidance command in the form of overload is given. Compared with the traditional guidance law, the guidance law theoretically ensures that the line-of-sight angle deviation and the line-of-sight rotation angle rate converge to zero before the target distance is reduced to the expected value. The superiority of the guidance law is verified by the simulation results of the proportional guidance (PNG) and the optimal navigation guide (ONG). In this paper, a finite-time convergence control law with/ without control input saturation is proposed for the attitude control of the mass moment aircraft, and a three-channel independent control model of the pitching, yaw and rolling of the mass moment missile is established on the basis of a reasonable simplification. Based on the control model, the attitude control law with limited time for pitching, yaw and rolling is designed. in comparison with that attitude control law of the traditional mass moment missile, the control law theoretically ensure that the attitude angle deviation and the rate of change of the attitude angle converge to the origin neighborhood. The yaw and roll channel controls the limited time-convergence attitude control law of the input saturation. Compared with the conventional attitude control law of the mass moment, the control law theory ensures that the attitude angle deviation and the change rate finite time in the control input saturation condition converge to the origin neighborhood. The correctness of the finite-time convergence attitude control law with/ without control input saturation is verified by the characteristic point simulation and the full-trajectory simulation. in that end, the integrated design of three-channel independent/ full-state coupled guidance control is put forward, and the integrated model of independent guidance control for the pitch, yaw and roll of the mass moment is established. The integral control law is designed based on the self-adaptive inversion method of the Rurod, the self-adaptive compensation of the unknown disturbance is realized through the self-adaptive function of the Rurod, the problem of the 鈥淐alculate expansion鈥，
本文編號(hào)：2378684
[Abstract]:The maneuvering control mode of the missile is realized by adjusting the control moment in nature, and the specific implementation form includes two types: adjusting the control force and the control force arm. The control of the conventional rudder surface and the control of the jet thrust belong to the category of the control force, but the control of the rudder surface is difficult to solve the problem of the aerodynamic ablation in the high-speed maneuver, and the jet thrust control is limited by the disadvantages of the limited carrying of the fuel and the coupling of the solid solution. the control of the mass moment belongs to the category of adjusting the control force arm, so that the problem that the rudder surface is controlled by the rudder surface and the like is avoided in the body of the elastic body due to the actuator of the control force arm, and the defect that the jet thrust control carries the limited fuel and the solid solution coupling and the like is solved by the air jet thrust control without carrying the extra fuel. In view of the obvious advantage of this quality moment control and wide application prospect, its unique control mode increases the complexity of the missile space motion and brings many new challenges to the guidance control design. The purpose of this paper is to study the dynamic modeling, guidance control and integrated design of the mass moment of the two-block/ differential aileron sideslip-turn (STT), so as to provide the theoretical support for the development of the mass-moment missile technology. First, a complete space motion model of the two-block/ differential aileron STT mass moment aircraft is established based on the multi-rigid-body system modeling method. Taking full consideration of the factors such as the kinematic coupling, the inertia coupling and the pneumatic inertia cross-coupling of the mass moment missile, the kinetic equation of the elastic body motion in the body coordinate system is established based on the momentum theorem. Based on the momentum moment theorem, the dynamic equations of the body rotation of the system in the body coordinate system are established, and the kinematic equations of the core translation of the elastic body in the re-input coordinate system and the kinematic equations about the core rotation of the elastic body are established. These equations describe the mechanism of the space motion of the mass moment aircraft, and reveal the internal relation between the movement of the slide block and the movement of the elastic body and the rotation of the body of the elastic body. Compared with the conventional rudder surface control and the air-jet control, the unique control mechanism of the mass moment aircraft, that is, the inherent coupling between the movement of the slide block and the body motion, makes the space motion model of the mass moment aircraft more complex. Secondly, aiming at the problem of target fixed target, the finite-r convergent guidance law with/ without terminal angle constraint is proposed. The relative motion between the missile and the target is described by using the target distance as the reference variable, and a new guidance model is established. The guidance model includes two differential equations, respectively describing the line-of-sight subduction motion and the line-of-sight turning motion, and the line-of-sight subduction motion differential equation is separately decoupled. The guidance law design based on the model not only ensures the accuracy but also simplifies the guidance law design process. 2) Based on the guidance model, a limited-r convergence guidance law with interference suppression is proposed, and the guidance instruction in the form of overload is given. Compared with the traditional guidance law, the guidance law theoretically ensures that the rotation angle rate of the line of sight converges to zero before the target distance is reduced to the expected value. The validity of the guidance law is verified by the simulation. 3) Based on the guidance model, a limited-r convergence guidance law with terminal angle constraint is proposed, and the guidance command in the form of overload is given. Compared with the traditional guidance law, the guidance law theoretically ensures that the line-of-sight angle deviation and the line-of-sight rotation angle rate converge to zero before the target distance is reduced to the expected value. The superiority of the guidance law is verified by the simulation results of the proportional guidance (PNG) and the optimal navigation guide (ONG). In this paper, a finite-time convergence control law with/ without control input saturation is proposed for the attitude control of the mass moment aircraft, and a three-channel independent control model of the pitching, yaw and rolling of the mass moment missile is established on the basis of a reasonable simplification. Based on the control model, the attitude control law with limited time for pitching, yaw and rolling is designed. in comparison with that attitude control law of the traditional mass moment missile, the control law theoretically ensure that the attitude angle deviation and the rate of change of the attitude angle converge to the origin neighborhood. The yaw and roll channel controls the limited time-convergence attitude control law of the input saturation. Compared with the conventional attitude control law of the mass moment, the control law theory ensures that the attitude angle deviation and the change rate finite time in the control input saturation condition converge to the origin neighborhood. The correctness of the finite-time convergence attitude control law with/ without control input saturation is verified by the characteristic point simulation and the full-trajectory simulation. in that end, the integrated design of three-channel independent/ full-state coupled guidance control is put forward, and the integrated model of independent guidance control for the pitch, yaw and roll of the mass moment is established. The integral control law is designed based on the self-adaptive inversion method of the Rurod, the self-adaptive compensation of the unknown disturbance is realized through the self-adaptive function of the Rurod, the problem of the 鈥淐alculate expansion鈥，
本文編號(hào)：2378684
本文鏈接：http://www.sikaile.net/shoufeilunwen/gckjbs/2378684.html
最近更新
教材專著
