鸟体姿态对结构抗鸟撞性能的影响

低于现行标准规定能量的大量鸟撞事故中,航空结构仍然发生实质性破坏的情况,说明只考虑鸟体的质量和速度不足以保证飞机安全。本文中针对弹性平板、雷达罩及机翼前缘等飞机典型结构,开展了不同姿态鸟体的鸟撞分析研究。分析结果发现,鸟体姿态对结构的抗鸟撞性能有比较显著的影响,不同的结构特点反映的响应规律也不同:对吸能结构,姿态角越大,吸收的能量越多,被保护的结构就越安全;而对承力结构,姿态角越大,高应力区域越大,结构就越危险。因此,在结构的抗鸟撞安全性评估中,除了完成特定姿态鸟体的鸟撞实验,针对危险工况还应通过数值分析评估不同鸟体姿态的结构撞击响应,进一步确保结构的抗鸟撞能力。     

范德华力对磁头/硬盘系统气体润滑特性的影响

通过数值方法分析范德华力对磁头承载力的影响以及随飞行姿态的变化规律.结果表明,范德华力主要作用在磁头滑块尾部的凸台上,当飞行高度低于10 nm后,范德华力使得滑块的承载能力明显降低,从而导致飞行高度下降.范德华力主要与飞行高度和俯仰角有关,随着飞行高度和俯仰角减小,范德华力使承载能力和俯仰转矩明显减小.范德华力对侧翻角的影响不敏感,随着侧翻角的增加,范德华力使滑块的承载能力和侧翻转矩略微减小.在磁头设计时,通过增加俯仰角和减小尾部凸台面积可以减小范德华力的影响.     

基于分布式合成双射流的飞行器无舵面三轴姿态控制飞行试验

将自主可控的合成双射流激励器集成于常规布局飞行器中, 进行了三轴无舵面控制飞行试验, 验证了分布式合成双射流对飞行器巡航时的无舵面姿态调控能力. 对合成双射流激励器进行改进, 设计了分布式三轴姿态控制合成双射流激励器, 滚转环量控制激励器分别安装于两侧机翼翼尖处后缘, 射流出口靠近压力面; 偏航反向合成双射流控制激励器分别安装于靠近两侧机翼翼尖20%弦长处, 上、下沿展向均匀布置; 俯仰环量控制激励器安装于V尾下的平尾后缘, 射流出口靠近压力面. 针对巡航速度为30 m/s的飞行器, 进行了三轴姿态控制飞行试验, 结果表明: 分布式合成双射流实现了飞行器巡航时的三轴无舵面姿态操控; 横航向控制存在耦合, 滚转环量控制激励器实现了飞行器的双向滚转操控, 能产生的最大滚转角速度达16.87°/s, 偏航反向合成双射流控制激励器实现了飞行器的双向偏航操控, 能产生的最大偏航角速度达9.09°/s; 俯仰环量控制激励器实现了飞行器的纵向控制, 能产生的最大俯仰角速度达7.68°/s.      

低成本车载GPS/INS组合导航姿态角更新算法

低成本INS系统的元件误差严重影响INS导航精度.针对车载系统,提出一种低成本车载GPS/INS组合导航姿态角更新算法.首先在GPS/INS组合导航Kalman滤波方程基础上,给出两种姿态角更新的观测方程.然后给出利用GPS测速确定航向角的原理,并且对低成本车载INS系统的俯仰角和翻滚角进行了分析,指出由INS随机误差造成的俯仰角和翻滚角误差比其本身量值要大,建议令俯仰角和翻滚角数值保持不变.利用实测算例确定了不同速度下的航向角精度,并且验证了该算法的有效性,以及相对于基于位置、速度组合的Kalman滤波,导航精度有明显提高.     

模糊神经网络预测在卫星姿态测量中的应用

以满足对地观测卫星测姿精度为目标,将由惯性基准、红外地平仪和太阳敏感器测姿过程视为典型的建模问题,讨论了基于自适应神经网络的模糊推理系统(ANFIS)的卫星姿态预测。仿真结果表明,ANFIS预测能够满足卫星姿态测量精度的要求,具有较强的容错性,同时该方法可将俯仰、横滚和偏航三个姿态分离建模,有利于提高卫星姿态测量的可靠性,为卫星姿态测量信息处理提供了一种新的方法。     

多自由度仿生推进系统的CPG控制

为了提高航行稳定性和机动性而设计的四尾鳍组合推进水下航行器,尾鳍运动自由度众多且相互耦合,稳定且快速的控制方案对提高航行器的整体性能至关重要。本文根据尾鳍运动特点,建立了中枢模式发生器(CPG)模型,协调控制8个驱动舵机,实现巡游、倒退、偏航、俯仰等各种航行状态下尾鳍的组合运动;通过陀螺仪监测航行器的偏航角与俯仰角,形成反馈信号引入CPG模型,对尾鳍运动进行反馈控制,进一步提高了航行稳定性。     

Attitude control by inertia wheels of a tethered interferometric SAR for topographic mapping

Various attitude control laws of a two masses tethered space system by using inertia wheels are analysed in this paper. The authors propose a zero-momentum system, consisting of the integration of three independent reaction wheels on board each platform. The control laws are derived by using a simplified attitude model and are then implemented in a three-dimensional numerical simulator, which simultaneously integrates the orbital and attitude equations. Final results show that pitch and yaw angular differences between the two platforms can be kept within values adequate for microwave remote sensing applications (10^–3 deg); in addition roll high frequency oscillations are damped.Based on a paper presented at the 11th National Congress of the Italian Association of Aeronanties and Astronautics, October 1991.     

Nonlinear dynamics of a satellite with deployable solar panel arrays

The multibody dynamics of a satellite in circular orbit, modeled as a central body with two hinge-connected deployable solar panel arrays, is investigated. Typically, the solar panel arrays are deployed in orbit using preloaded torsional springs at the hinges in a near symmetrical accordion manner, to minimize the shock loads at the hinges. There are five degrees of freedom of the interconnected rigid bodies, composed of coupled attitude motions (pitch, yaw and roll) of the central body plus relative rotations of the solar panel arrays. The dynamical equations of motion of the satellite system are derived using Kane's equations. These are then used to investigate the dynamic behavior of the system during solar panel deployment via the 7-8th-order Runge-Kutta integration algorithms and results are compared with approximate analytical solutions. Chaotic attitude motions of the completely deployed satellite in circular orbit under the influence of the gravity-gradient torques are subsequently investigated analytically using Melnikov's method and confirmed via numerical integration. The Hamiltonian equations in terms of Deprit's variables are used to facilitate the analysis.     

Decentralized coordinated attitude control for satellite formation flying via the state-dependent Riccati equation technique

The goal of the present study is to develop a decentralized coordinated attitude control algorithm for satellite formation flying. To handle the non-linearity of the dynamic system, the problems of absolute and relative attitude dynamics are formulated for the state-dependent Riccati equation (SDRE) technique. The SDRE technique is for the first time utilized as a non-linear controller of the relative attitude control problem for satellite formation flying, and then the results are compared to those from linear control methods, mainly the PD and LQR controllers. The stability region for the SDRE-controlled system was obtained using a numerical method. This estimated stability region demonstrates that the SDRE controller developed in the present paper guarantees the globally asymptotic stability for both the absolute and relative attitude controls. Moreover, in order to complement a non-selective control strategy for relative attitude error in formation flying, a selective control strategy is suggested. This strategy guarantees not only a reduction in unnecessary calculation, but also the mission-failure safety of the attitude control algorithm for satellite formation. The attitude control algorithm of the formation flying was tested in the orbital-reference coordinate system for the sake of applying the control algorithms to Earth-observing missions. The simulation results illustrate that the attitude control algorithm based on the SDRE technique can robustly drive the attitude errors to converge to zero.     

四旋翼飞行器建模及位置跟踪控制(英文)

相对其他无人飞行器平台,四旋翼飞行器有其独特的优势,因而受到广泛的关注。位置跟踪控制对四旋翼飞行器的应用非常重要。在阐述四旋翼飞行器的飞行原理和操控机制的基础上,研究了其动力学模型,并提出了一种简化的数学模型。四旋翼飞行器是欠驱动耦合系统,为了实现系统解耦并得到清晰的控制回路,设计了多回路PID控制方案,其控制目标是位置和偏航角,而姿态角和横滚角由位置误差调节。最后,通过仿真验证了控制方法的有效性。     

微小型航姿测量系统及其数据融合方法

以小型无人机航姿测量系统的微小型化为背景,利用MEMS惯性测量元件研制了一种低成本微型航姿测量系统.针对MEMS器件用于载体航姿测量时精度低、易发散的问题,提出一种计算量小、实时性强的加速度信息、磁场强度信息、陀螺信息的融合方法.采用卡尔曼滤波器对系统的俯仰角、滚转角和航向角的误差进行最优估计;设计数据融合的判别准则,并根据判据的判断结果调整卡尔曼滤波器中的量测信息,使系统可用于小型无人机的定高自主飞行.实验结果表明,系统输出航姿的更新频率可达100Hz,航姿测量误差小于0.6°,航姿标准差小于0.09°;将其应用于某小型固定翼飞行器的飞行控制系统中进行自主飞行实验,完成了预定的飞行任务.     

磁头/磁盘气体润滑特性计算与分析

提出了1种计算超薄气膜润滑轴承压力分布的有限差分法,在此基础上对某磁头快速或缓慢偏离平衡位置造成的压力变化及气体轴承刚度进行分析.结果表明:随着气膜特征高度降低,初始纵翻倾角减小,气体轴承刚度增大;垂直于磁盘方向的微小位移不会对磁头平衡造成威胁,但磁头纵翻造成的翻转力矩变化较复杂,需要加以控制;磁头飞行姿态的突变将引起明显的挤压效应,挤压效应的强弱与初始飞行姿态有关.     

火箭弹大动态单轴平台惯导系统姿态算法

火箭弹在飞行中常采用滚转稳定的控制方式,其滚转角速度的动态范围很大,因此实时、准确地测量滚转角速度和滚转姿态角成为制导火箭弹控制的关键问题。大动态单轴平台惯导系统将IMU安装在沿滚转方向的稳定平台上,通过伺服电机驱动单轴平台相对于弹体反旋,隔离滚转方向的大动态角速度,为IMU提供平稳的测试环境。介绍了大动态单轴平台惯导系统的组成和功能,搭建了样机,推导了惯导姿态解算的数学模型。经过120 s半实物仿真试验,系统俯仰姿态角误差<4°,偏航姿态角误差<3°,滚转姿态角误差<25°,结果验证了整体方案的可行性和姿态解算模型的正确性。为进一步提高姿态解算精度,搭建单轴平台组合导航系统,实现全部导航信息的高精度测量打下了基础。     

Spacecraft aerodynamics and trajectory simulation during aerobraking

This paper uses a direct simulation Monte Carlo （DSMC） approach to simulate rarefied aerodynamic characteristics during the aerobraking process of the NASA Mars Global Surveyor （MGS） spacecraft. The research focuses on the flowfield and aerodynamic characteristics distribution under various free stream densities. The vari- ation regularity of aerodynamic coefficients is analyzed. The paper also develops an aerodynamics-aeroheating-trajectory integrative simulation model to preliminarily calculate the aerobraking orbit transfer by combining the DSMC technique and the classical kinematics theory. The results show that the effect of the planetary atmospheric density, the spacecraft yaw, and the pitch attitudes on the spacecraft aerodynamics is significant. The numerical results are in good agreement with the existing results reported in the literature. The aerodynamics-aeroheating-trajectory integrative simulation model can simulate the orbit transfer in the complete aerobraking mission. The current results of the spacecraft trajectory show that the aerobraking maneuvers have good performance of attitude control.     

方位捷联惯导平台系统及其误差分析

根据导航系统对高精度的需要,在捷联惯导和平台惯导的基础上提出了一种方位捷联惯导平台,该平台取消了普通平台系统的方位环及其伺服回路,保留了俯仰环和横滚环及其对应的伺服回路,介绍了平台的组成及工作原理,并且推导出平台角误差方程、速度误差方程和位置误差方程.建立状态方程和量测方程后用Kalman滤波的方法对系统进行软件仿真,仿真结果表明,方位捷联惯导平台综合了捷联和平台的优点,具有平台角误差小、收敛速度快等特点.     

可重复使用运载器大姿态机动自抗扰控制

针对可重复使用运载器大俯仰角或偏航角转弯机动而产生的姿态角奇异的控制问题,提出了基于四元数的自抗扰控制方法。通过两级跟踪微分器从期望四元数中逐步得到三通道解耦的角加速度信号,然后利用扩张状态观测器观测模型中的不确定项,最终采用动态逆得到解耦的三通道发动机等效摆角或RCS(Reaction Control System)等控制信号,并设计了数字滤波器对弹性振动与液体晃动信号进行滤波处理。考虑到系统模型具有非线性、不确定性、11阶弹性振动、一阶液体晃动、风干扰和气动偏差等多种外部扰动条件,对可重复使用运载器从主动段到再入飞行段进行了非线性六自由度仿真分析。仿真结果表明,基于四元数的自抗扰姿态控制器具有快速、平稳、超调量小、抗干扰能力强、无系统抖振且控制参数较少的特点。     

Calibration of a four-hole pyramid probe and area traverse measurements in a short-duration transonic turbine cascade tunnel

A four-hole pyramid probe has been calibrated for use in a short-duration transonic turbine cascade tunnel. The probe is used to create area traverse maps of total and static pressure, and pitch and yaw angles of the flow downstream of a transonic annular cascade. This data is unusual in that it was acquired in a short-duration (5 s of run time) annular cascade blowdown tunnel. A four-hole pyramid probe was used which has a 2.5 mm section head, and has the side faces inclined at 60° to the flow to improve transonic performance. The probe was calibrated in an ejector driven, perforated wall transonic tunnel over the Mach number range 0.5–1.2, with pitch angles from -20° to + 20° and yaw angles from-23° to +23°. A computer driven automatic traversing mechanism and data collection system was used to acquire a large probe calibration matrix (～ 10,000 readings) of non dimensional pitch, yaw, Mach number, and total pressure calibration coefficients. A novel method was used to transform the probe calibration matrix of the raw coefficients into a probe application matrix of the physical flow variables (pitch, yaw, Mach number etc.). The probe application matrix is then used as a fast look-up table to process probe results. With negligible loss of accuracy, this method is faster by two orders of magnitude than the alternative of global interpolation on the raw probe calibration matrix. The blowdown tunnel (mean nozzle guide vane blade ring diameter 1.1 m) creates engine representative Reynolds numbers, transonic Mach numbers and high levels (≈ 13%) of inlet turbulence intensity. Contours of experimental measurements at three different engine relevant conditions and two axial positions have been obtained. An analysis of the data is presented which includes a necessary correction for the finite velocity of the probe. Such a correction is non trivial for the case of fast moving probes in compressible flow.     

A controllability perspective of dynamic soaring

Dynamic soaring is an exquisite flying technique to acquire energy from the atmospheric wind shear. In this study, a geometric nonlinear controllability analysis of an unmanned aerial vehicle (UAV) under dynamic soaring conditions is performed. To achieve such an objective, the state-of-the-art mathematical tools of nonlinear controllability are summarized and presented to an aeronautical engineering audience. The dynamic soaring optimal control problem is then formulated and solved numerically. The controllability of the UAV along the optimal soaring trajectory is analyzed. More importantly, the geometric nonlinear controllability characteristics of generic flight dynamics are analyzed in the presence and absence of wind shear to provide a controllability explanation for the role of wind shear in the physics of dynamic soaring flight. It is found that the wind shear is instrumental in ensuring controllability as it allows the UAV attitude controls (pitch and roll) to play the role of thrust in controlling the flight path angle. The presented analysis represents a controllability-based mathematical proof for the energetics of flight physics.     

基于对偶四元数的双星编队相对位姿测量

为了描述编队卫星中主从星的相对位置和姿态信息,提出了基于对偶四元数的编队卫星相对位姿测量算法。以双星编队飞行的位姿运动为主线,运用对偶四元数工具,充分发挥其能以最简洁的形式表示一般性刚体运动的优点,对卫星轨道和姿态进行分析并建立了对偶四元数位姿模型。同时设计类GPS测量技术来测量编队卫星的相对位置和姿态,该技术载波相位波长和伪码码元比GPS的更短,可获得更高精度的相对测量信号。由于状态方程和观测方程的非线性特征,使用UKF滤波来消除随机噪声对量测过程的干扰。实验结果表明,所设计的算法能够有效估计系统误差,卫星的位置误差和四元数误差收敛于零,验证了该算法的有效性。     

Direct neural discrete control of hypersonic flight vehicle

This paper investigates the discrete neural control for flight path angle and velocity of a generic hypersonic flight vehicle (HFV). First, strict-feedback form is set up for the attitude subsystem considering flight path angle, pitch angle, and pitch rate by altitude-flight path angle transformation. Secondly, the direct Neural Network?(NN) control is proposed for attitude subsystem via back-stepping scheme. The direct design is employed for system uncertainty approximation with less online tuned NN parameters and there is no need to know the information of the upper bound of control gain during the controller design. Thirdly, with error feedback and NN design, the semiglobal uniform ultimate boundedness (SGUUB) stability is guaranteed of the closed-loop system. Similar NN control is applied on velocity subsystem. Finally, the feasibility of the proposed controller is verified by a simulation example.