本文選題：火星 + 大氣進(jìn)入��； 參考：《北京理工大學(xué)》2015年碩士論文

【摘要】：未來的火星探測任務(wù),要求探測器具有在特定區(qū)域精確著陸的能力。大氣進(jìn)入段是火星著陸過程環(huán)境最惡劣的階段,探測器的制導(dǎo)與控制不僅直接影響開傘點(diǎn)的精度,對探測器的最終著陸精度乃至任務(wù)的成敗也將產(chǎn)生重要影響。本論文以火星精確著陸探測任務(wù)為背景,針對火星大氣進(jìn)入段的探測器制導(dǎo)與姿態(tài)控制問題,研究了大氣進(jìn)入段軌跡優(yōu)化方法以及制導(dǎo)與姿態(tài)控制方法,并進(jìn)行了數(shù)值仿真與分析。首先,對火星的動(dòng)力學(xué)環(huán)境參數(shù)進(jìn)行了分析;給出了火星大氣進(jìn)入段相關(guān)坐標(biāo)系及轉(zhuǎn)換關(guān)系,進(jìn)而推導(dǎo)并建立了火星大氣進(jìn)入段的軌道動(dòng)力學(xué)和姿態(tài)動(dòng)力學(xué)模型,為進(jìn)入段軌跡優(yōu)化與制導(dǎo)控制方法研究提供基礎(chǔ)。其次,考慮進(jìn)入段軌道動(dòng)力學(xué)模型,給出了火星大氣進(jìn)入段受到的各類約束,并通過約束規(guī)范化改進(jìn)Gauss偽譜法,對火星進(jìn)入段軌跡進(jìn)行優(yōu)化設(shè)計(jì)。以傾側(cè)角為控制量,求解滿足各約束條件的最優(yōu)參考軌跡,然后,采用阿波羅式的ETPC攝動(dòng)制導(dǎo)律,對優(yōu)化的軌跡進(jìn)行了跟蹤控制仿真,并對仿真結(jié)果進(jìn)行了分析。然后,針對采用質(zhì)心偏置的方式來獲得有限升力的低升阻比火星探測著陸器,考慮火星進(jìn)入大氣環(huán)境的復(fù)雜多變性,采用能量作為變量,以減小開傘點(diǎn)誤差和增加開傘高度為目標(biāo),研究了一種基于參數(shù)修正PID反饋-自適應(yīng)增益的預(yù)測跟蹤火星大氣進(jìn)入制導(dǎo)律,分別對該方法的預(yù)測、縱向和橫向制導(dǎo)律進(jìn)行了設(shè)計(jì),并通過數(shù)值仿真驗(yàn)證了方法的有效性。最后,針對火星大氣進(jìn)入段姿態(tài)控制問題,給出了姿態(tài)控制相平面姿態(tài)分析方法。研究并給出了基于相平面分析的俯仰/偏航通道PD/D穩(wěn)定控制律和滾轉(zhuǎn)通道時(shí)間-燃料最優(yōu)機(jī)動(dòng)控制律,仿真分析驗(yàn)證了方法的有效性。
[Abstract]:Future missions to Mars will require the spacecraft to land accurately in a particular area. The atmospheric entry section is the worst stage of the Martian landing environment. The guidance and control of the spacecraft not only directly affect the accuracy of the parachute opening point, but also have an important impact on the final landing accuracy of the spacecraft and the success or failure of the mission. In this paper, based on the precise landing mission of Mars, aiming at the guidance and attitude control of the Mars atmosphere, the trajectory optimization method and the guidance and attitude control method of the atmospheric entry section are studied. Numerical simulation and analysis are carried out. First of all, the dynamic environment parameters of Mars are analyzed, and the relative coordinate system and the transformation relation are given, and then the orbital dynamics and attitude dynamics model of Mars atmospheric entry section are derived and established. It provides the foundation for the research of trajectory optimization and guidance control. Secondly, considering the orbital dynamics model of the entry phase, the various constraints on the entry segment of the Mars atmosphere are given, and the trajectory of the entry segment of Mars is optimized by the improved Gauss pseudospectral method of constraint normalization. The optimal reference trajectory satisfying each constraint condition is solved with the tilting angle as the control quantity. Then, the tracking control simulation of the optimized trajectory is carried out by using the ETPC perturbation guidance law of Apollo, and the simulation results are analyzed. Then, considering the complex variability of Mars entering the atmospheric environment, considering the complex variability of Mars entering the atmospheric environment, the energy is used as a variable for the low lift Mars probe lander, which uses the centroid bias to obtain a finite lift. In order to reduce the error of opening point and increase the height of the parachute, a predictive tracking guidance law based on parameter-modified PID feedback and adaptive gain for Mars atmospheric entry guidance is studied. Longitudinal and transverse guidance laws are designed, and the effectiveness of the method is verified by numerical simulation. Finally, aiming at the problem of attitude control in Martian atmosphere, a phase plane attitude analysis method for attitude control is presented. The PD/D stability control law and the time-fuel optimal maneuvering control law for pitch / yaw channel based on phase plane analysis are studied and presented. The simulation results show that the method is effective.
【學(xué)位授予單位】：北京理工大學(xué)
【學(xué)位級別】：碩士
【學(xué)位授予年份】：2015
【分類號】：V448.2

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本文編號：1804087