点光源反射镜法向标校技术

针对基于反射式点光源进行在轨辐射定标过程中反射镜法向标校建模不够完善的问题,提出基于反射镜与相机几何模型的反射镜法向标校及矢量控制算法.通过解算模型求解相机与反射镜间的几何误差,建立了太阳图像质心坐标与反射镜法向之间的关系,可实现多点自动化标校反射镜法向,提高镜法向标校及系统指向精度.实验结果表明,利用解算后的几何模型反解不同时刻质心坐标进行多点反射镜法向标校,相机观测太阳像素角分辨率标准误差分别为:X轴方向0.02165°、Y轴方向0.01982°,综合角分辨率误差为0.02936°,优于太阳观测器对反射镜法向标校精度.实现了相机观测太阳取代人工借助太阳观测器观测太阳的自动化镜法向标校,扩展了标校灵活度,系统综合指向精度优于0.1°,为固定实验场联网自动化集中控制不同能级梯度的点光源阵列在轨辐射定标和调制传递函数检测奠定基础.     

钝头机体用FADS-α的求解算法及精度研究

针对钝头机体用嵌入式大气数据传感(flush air data sensing,FADS)系统的4类攻角求解算法及算法的求解精度进行研究.针对典型的15°钝头体外形,在Mach数Ma=2.04,3.02,5.01,攻角α=-5°～30°,侧滑角β=0°的条件下,首先基于势流理论及修正的Newton流理论建立了钝头机体用...     

基于准直光束的高精度滚转角测量方法研究

对于滚转角测量精度低并且难于测量的问题,提出了一种基于准直光束光斑位置变化的高精度滚转角测量方法。当被测物转动时,CCD上2光斑位置随之改变,2个光斑中心连线斜率亦改变。斜率变化由被测物俯仰、偏摆、滚转运动引起。在测量系统中基于自准直原理测量偏摆角和俯仰角,运用相关的算法,消除由偏摆和俯仰运动引起的滚转角误差,从而实现滚转角的精确测量。同时运用Zemax建立系统仿真模型,进行了滚转角的仿真实验测量。将仿真实验结果输入到滚转角解算模型中解算,结果表明:在0～1 800范围内滚转角的解算值与Zemax的设定值完全一致,由此验证了测量方法的可行性及正确性。     

温度和电场作用下保偏光纤对方位传递的影响

针对基于保偏光纤的非通视方位传递系统中温度与电场引起的方位传递误差,从非通视方位传递系统的工作原理出发,分别推导了基于温度、电场作用保偏光纤的方位角解算模型,并仿真分析了保偏光纤长度、热膨胀系数及电场振幅等因素对方位传递精度的影响.实验结果表明,当温度为20℃时,方位角解算误差对温度不敏感;当温度高于20℃时,方位角解算误差容易受温度、保偏光纤长度的影响,可通过调整保偏光纤的热膨胀系数、泊松比控制方位传递误差.电场大小和电场作用长度共同影响方位角解算误差,当作用角度为45°时,电场振幅不会产生方位角解算误差.该研究结果对保偏光纤非通视方位传递系统的环境适应性研究及系统测量精度的提高具有一定的指导意义.     

基于径向基网络的炭黑填充胶生热率模型重构

在分析炭黑填充胶生热率随温度、频率变化趋势基础上,构建了生热率的径向基网络模型(RBF网络),并对RBF网络进行模拟,将RBF网络与反向传播网络(BP网络)进行比较,分析误差。模拟结果表明RBF网络精度最高,探讨了将人工神经网络理论引入到炭黑/橡胶复合材料热物性参数研究领域的可行性。     

脉冲激光探测方位角磁电检测技术

针对常规弹药脉冲激光周向探测系统无法精确获取目标方位信息的问题,设计了基于磁电检测的单光束激光引信全向方位角探测方案.对磁电检测系统进行建模,建立圆柱形永磁体转动磁场模型,推导出磁阻传感器所测位置磁场的解析式,验证所测磁场为一正弦磁场信号.依据此正弦信号,设计了上升沿阈值周期检测算法,并运用FPGA与TDC-GP21对激光回波出现时间与电机转速信号周期进行高准确度时间间隔测量实现方位角的解算.依据方案设计原理样机并编写上位机程序,进行方位角探测实验.实验结果表明:磁电检测系统采用多重屏蔽方法,能有效抑制电磁干扰;并能实时监测电机转速,实现方位角解算,方位角解算误差在±2°以内.满足激光引信方位角测量的高准确度、抗干扰能力强等要求.     

圆光栅测角系统误差分析与修正

为了提高测角系统的测量精度，基于误差的特性和表现形式，从谐波角度对测角系统中的刻划误差、光电信号误差、偏心误差、倾斜误差和变形导致的误差建立相应的数学模型，并对偏心误差和倾斜误差进行仿真分析。搭建了金属圆光栅测角系统，对系统中的误差进行了分析和修正。实验结果表明:测角系统测角误差为134.2″，修正后误差为32.23″，圆光栅测角系统误差分析为后续的误差补偿提供了理论基础。     

Ritchey角精度对平面镜检测的影响的分析与验证

为进一步提高Ritchey-Common法的检测精度,分析了实验中Ritchey角精度对整体检测结果的影响。通过仿真模拟,分析并确定出最佳Ritchey角测试范围在20°～50°之间,此时面形误差检测结果精度可达0.01λ(λ=0.6328μm)。仿真过程中模拟Ritchey角存在误差时对检测结果的影响,当Ritchey角误差控制在±1°时,拟合结果与原始面形的残差降至0.0007λ,能够满足测试要求。针对Ritchey角测量存在误差的问题,利用测得系统光瞳面的图像压缩比例来计算Ritchey角大小,此方法的计算误差可控制在0.2°以内。实验中选择3个角度来检测,在数据处理时将测得数据两两组合进行解算。29.6°&47.8°组合拟合结果与Zygo干涉仪直接检测结果的残差的峰谷(PV)值为0.068λ、均方根(RMS)值为0.0105λ,证明Ritchey角的选择及其计算精度对检测整体精度具有一定影响。     

基于矢量波像差理论的两反系统装配失调解算方法

以矢量波像差理论中的三级像差理论为基础,提出了一种两反系统装配失调量解算方法。该方法仅采用轴上视场波前像差系数建立失调量解算模型,然后基于球差系数解算间隔误差,并基于彗差和像散系数解算偏心和倾斜误差,大幅提高了解算精度和效率。以某一两反光学系统为例,利用光学设计软件Zemax进行模拟装调,系统轴上视场失调像差系数均减小到10~(-7)数量级,失调误差均校正到10~(-5)数量级,达到了良好的装调效果。最后,利用该失调量解算模型指导两反光学系统的装调,失调量解算精度及装调精度达到了使用要求。Zemax模拟装调结果和实际装调效果均证明了所提方法的正确性。     

降低DOAS系统探测限的新型反演算法研究

差分吸收光谱法(DOAS)是一种高灵敏测量大气痕量气体成分含量的有效的光学遥感方法,该方法基于最小二乘拟合模型,利用获得的痕量气体的差分吸收光学密度与标准的吸收截面进行拟合,反演待测气体的浓度.建立了基于径向基(RBF)神经网络的痕量气体浓度反演的新模型,对网络的隐层参数采用改进最近邻聚类学习算法训练,对输出层权值的训练采用梯度下降算法,使得网络收敛快,能更好地实时、在线反演测量光谱.并针对DOAS技术的特点,把拟合残差输入网络集中训练,使得RBF网络在反演真实痕量气体吸收时,效果更佳.实验结果表明该新型反演方法提高了DOAS系统的反演精度,降低了DOAS系统的探测限.     

飞行器大攻角非定常气动特性神经网络建模

大攻角气动特性预测与气动建模是新型飞行器提升飞行性能的重要内容.以轴对称导弹简化模型为研究对象，首先采用计算流体力学方法，对70°大攻角状态的非定常气动特性进行数值模拟，计算方法基于RANS的N-S方程，湍流模型采用SA模型，对流场采用有限体积法离散，无黏项采用Roe通量差分分裂格式，黏性项采用中心差分，时间推进采用LU-SGS格式的双时间步法.飞行器运动模式采用强迫振荡的方式，对5种不同振荡频率进行了非定常数值计算，并记录每一内迭代周期最终的气动力和力矩数值.其次，以CFD预测结果作为气动建模的样本，采用动导数模型、多项式模型等传统方法，进行气动建模，并分析其有效性和精度.最后采用神经网络方法对大攻角非定常气动力进行建模，并和动导数模型、多项式模型进行精度对比.结果表明，基于神经网络的人工智能气动建模方法具有较高的精度和适应性.该方法为飞行器大攻角非定常非线性气动建模，大攻角飞行稳定性分析与控制提供理论参考.      

Numerical investigation on properties of attack angle for an opposing jet thermal protection system

The three-dimensional Navier-Stokes equation and the k-ε viscous model are used to simulate the attack angle characteristics of a hemisphere nose-tip with an opposing jet thermal protection system in supersonic flow conditions.The numerical method is validated by the relevant experiment.The flow field parameters,aerodynamic forces,and surface heat flux distributions for attack angles of 0°,2°,5°,7°,and 10° are obtained.The detailed numerical results show that the cruise attack angle has a great influence on the flow field parameters,aerodynamic force,and surface heat flux distribution of the supersonic vehicle nose-tip with an opposing jet thermal protection system.When the attack angle reaches 10°,the heat flux on the windward generatrix is close to the maximal heat flux on the wall surface of the nose-tip without thermal protection system,thus the thermal protection has failed.     

超声速流场条件下激光辐照耦合效应数值模拟

激光辐照结构物包含复杂的多物理场耦合问题,其存在流、热、固多种机制的耦合效应。结合计算流体力学(CFD)和有限元方法,对超声速条件下的激光辐照平板问题进行了热流固耦合分析。采用CFD方法得到平板附近流场分布,利用有限元方法计算平板的温度分布,并将二者结合起来实现流体和固体间的数据交互。理论分析确定了流场效应的最主要影响参数为来流马赫数与攻角。对于不同马赫数,激光区域在6 Ma条件下存在温度的谷值,小于等于6 Ma条件下主要体现为冷却效应,而6 Ma以上主要体现为气动加热效应。攻角增大会导致激光区流体质量流量的增加,使冷却效应更加明显。最后综合分析了流场气动加热和冷却两种效应的产生机制。     

钝锥再入飞行器黏性干扰效应建模

针对典型的钝锥外形, 采用统一气体动理学格式(UGKS)模拟了高度70~110 km下不同Mach数和攻角的流场, 进行了流场特性的分析, 并基于黏性干扰的理论成果, 将气动力特性与第3黏性干扰参数、攻角和Mach数等参数进行关联, 建立了气动力系数的黏性干扰模型, 给出了模型预测结果的相关性分析和准确性评估。经初步测试, 该模型预测结果与UGKS直接模拟结果具有良好的一致性, 对工程应用快速获取高空气动特性具有重要意义。      

Estimation of energetic efficiency of heat supply in front of the aircraft at supersonic accelerated flight. Part 1. Mathematical models

Fuel economy at boost trajectory of the aerospace plane was estimated during energy supply to the free stream. Initial and final flight velocities were specified. The model of a gliding flight above cold air in an infinite isobaric thermal wake was used. The fuel consumption rates were compared at optimal trajectory. The calculations were carried out using a combined power plant consisting of ramjet and liquid-propellant engine. An exergy model was built in the first part of the paper to estimate the ramjet thrust and specific impulse. A quadratic dependence on aerodynamic lift was used to estimate the aerodynamic drag of aircraft. The energy for flow heating was obtained at the expense of an equivalent reduction of the exergy of combustion products. The dependencies were obtained for increasing the range coefficient of cruise flight for different Mach numbers. The second part of the paper presents a mathematical model for the boost interval of the aircraft flight trajectory and the computational results for the reduction of fuel consumption at the boost trajectory for a given value of the energy supplied in front of the aircraft.     

Flow characteristics of supersonic gas passing through a circular micro-channel under different inflow conditions

Gas flow in a micro-channel usually has a high Knudsen number. The predominant predictive tool for such a microflow is the direct simulation Monte Carlo(DSMC) method, which is used in this paper to investigate primary flow properties of supersonic gas in a circular micro-channel for different inflow conditions, such as free stream at different altitudes, with different incoming Mach numbers, and with different angles of attack. Simulation results indicate that the altitude and free stream incoming Mach number have a significant effect on the whole micro-channel flow field, whereas the angle of attack mainly affects the entrance part of micro-channel flow field. The fundamental mechanism behind the simulation results is also presented. With the increase of altitude, thr free stream would be partly prevented from entering into micro-channel.Meanwhile, the gas flow in micro-channel is decelerated, and the increase in the angle of attack also decelerates the gas flow. In contrast, gas flow in micro-channel is accelerated as free stream incoming Mach number increases. A noteworthy finding is that the rarefaction effects can become very dominant when the free stream incoming Mach number is low. In other words, a free stream with a larger incoming velocity is able to reduce the influence of the rarefaction effects on gas flow in the micro-channel.     

Aerodynamic characteristics of passenger and transport aircraft models at harmonic oscillations on the pitch angle for high angles of attack

A systematic analysis has been performed for many-years experimental data obtained in the wind tunnel T-203 (SibNIA) for testing the models of passenger and transport aircraft for the case of harmonic oscillation at the pitch angle for low subsonic velocities. The key features of behavior of aerodynamic derivatives coefficients and dependencies of current values of normal force coefficient and longitudinal moment coefficient on the angle of attack have been demonstrated for the stalling modes of streamlining. It was demonstrated that at near-critical angles of attack, we have a strong dependency of aerodynamic derivatives of pitch moment on the normalized oscillation frequency for the range of natural values; this makes the traditional mathematical model of aerodynamic loads (uses the aerodynamic derivatives at fixed frequencies of oscillation) unfit for the considered scope of experimental tasks.     

Nonlinear Modeling Study of Aerodynamic Characteristics of an X38-like Vehicle at Strong Viscous Interaction Regions

Strong viscous interaction and multiple flow regimes exist when vehicles fly at high altitude and high Mach number conditions. The Navier–Stokes(NS) solver is no longer applicable in the above situation. Instead, the direct simulation Monte Carlo (DSMC) method or Boltzmann model equation solvers are usually needed. However, they are computationally more expensive than the NS solver. Therefore, it is of great engineering value to establish the aerodynamic prediction model of vehicles at high altitude and high Mach number conditions. In this paper, the hypersonic aerodynamic characteristics of an X38-like vehicle in typical conditions from 70 km to 110 km are simulated using the unified gas kinetic scheme (UGKS), which is applicable for all flow regimes. The contributions of pressure and viscous stress on the force coefficients are analyzed. The viscous interaction parameters, Mach number, and angle of attack are used as independent variables, and the difference between the force coefficients calculated by UGKS and the Euler solver is used as a dependent variable to establish a nonlinear viscous interaction model between them in the range of 70–110 km. The evaluation of the model is completed using the correlation coefficient and the relative orthogonal distance. The conventional viscous interaction effect and rarefied effect are both taken into account in the model. The model can be used to quickly obtain the hypersonic aerodynamic characteristics of X38-like vehicle in a wide range, which is meaningful for engineering design.     

CFD在高速飞行器热气弹问题中的应用

采用气动力/热/结构耦合的方法对高速细长体飞行器结构热静气动弹性问题进行了研究.为保证耦合计算精度，达到准确预测热气动弹性特性的能力，气动力和气动热计算采用CFD数值模拟方法，热应力和热变形计算采用有限元方法并通过热考核试验验证.以该简单细长体飞行器模型为研究对象，对其热静气动弹性特性进行了计算与分析，计算结果表明:CFD/CSD耦合可准确模拟热气弹问题，且气动加热造成结构温升不均衡是结构变形的主导因素，力热耦合静气弹变形与单纯受力分析变形形式不同，对飞行器气动特性影响规律不同.准确预测飞行器热气动弹性特性对飞行器结构设计十分必要.