To alleviate the locking problem in the ANCF beam elements, sufficient transverse gradient vectors are incorporated in the cross section to enrich the distribution of transverse strain along the cross section of the beam. Building upon this novel concept, this paper utilizes Pascal trigonometric polynomial to determine the position interpolation field of beam elements, and the distribution of transverse gradient vectors along the beam section is clarified through the collocation of boundary points and Chebyshev interpolation nodes, and then a series of locking-free beam models, based on the absolute nodal coordinate formulation, are developed. Additionally, it reveals the inherent mechanical mechanism of higher-order beam models in alleviating locking through strict mathematical analysis. Furthermore, to demonstrate the effectiveness of the new elements, six numerical simulation examples are designed, namely, three static examples and three dynamic examples, which involve small deformation statics, large deformation statics, small-scale elastic deformation, large-scale elastic deformation problems. Finally, the simulation results of the first four order beam models, Patel–Shabana model, and ECM approach are compared and analyzed in detail. The results indicate that the proposed higher-order beam models have high accuracy and can effectively eliminate the unnecessary influence caused by locking in complex mechanical problems, involving statics and dynamics problems.

In this paper, a novel locking-free finite beam element is proposed utilizing the absolute nodal coordinate formulation. By incorporating a gradient vector along the transverse direction at the boundary points, the linear interpolation of the gradient vector field within this element is achieved. Consequently, the problem of constant transverse strain distribution, which is observed in the Omar–Shabana beam element, is effectively addressed. Building upon this concept, this study further extends the proposed element from two dimensions to three dimensions. Additionally, it analyzes and compares the locking alleviation mechanism of the newly developed element and Patel–Shabana beam element. The analysis aims to provide insights into the factors contributing to the locking alleviation of different absolute nodal coordinate formulation (ANCF) elements. Furthermore, to demonstrate the effectiveness of the new element, six numerical simulation examples are designed, comprising three and three dynamic examples. These examples encompass small deformation statics, large deformation statics, small-scale motion, large-scale motion, and rotational motion problems. Finally, the results indicate that the proposed ANCF beam element can effectively alleviate the locking problem. The element exhibits robust adaptability, rationality, and effectiveness when subjected to complex mechanical characteristics by comparing the numerical results of this element with the classical Omar–Shabana, high-order Shen, and Patel–Shabana elements.

为研究弯折波现象和分析弯折波对舰载机拦阻过程的影响,基于Euler–Bernoulli非线性梁理论,建立了考虑弯折波的拦阻索离散模型.对拦阻过程进行数值仿真,仿真结果表明,弯折波在整个拦阻过程均存在,使得拦阻钩载荷出现波动,且这种波动在拦阻初期最为剧烈,之后逐渐减弱.弯折波的存在明显缩短了拦停时间和拦停位移,还使得拦阻索拉力的最大值更大.此外,奇数重弯折波与偶数重弯折波在传播方向、拉力角和弯折波平均波速等方面均表现出明显的不同.

在弹射起飞准备阶段,舰载机轴线并不能准确地与弹射轨道重合,产生一定的偏心距.为了分析舰载机不同初始偏心距下的弹射起飞的动态特性,以某轻型舰载机为例,建立了舰载机弹射起飞六自由度动力学模型,仿真分析了不同初始偏心距对舰载机姿态运动、弹射杆受力、起落架受力的影响.计算结果表明,初始偏心距对舰载机姿态中的偏航运动影响最为显著,随着初始偏心距的增大,偏航幅值增大,侧向力增大,主起落架垂向受力差值增大,但初始偏心距对舰载机的俯仰运动的影响较小.

拦阻着舰过程中,舰载机着舰距离短、载荷大,有必要确定起落架的真实动态响应.针对拦阻着舰的复杂性,以某型舰载机为研究对象,通过LMS.Virtual.Lab Motion建立了该舰载机拦阻着舰动力学模型.考虑飞机在自由飞行状态下,以不同舰上结合速度以及不同下沉速度挂索着舰的情况.通过仿真分析,挂索时的下沉速度越大,舰上结合速度越小,主起落架受载越大,前起落架则呈现相反的情况.分析结果可为舰载机拦阻着舰过程研究以及起落架结构设计提供参考.

针对具有渐进最优性的快速扩展随机树(RRT*)算法在面对高维、复杂环境时所表现出的寻路效率低、收敛速度缓慢的问题,在RRT*的基础上,提出一种基于采样空间自调整的渐进最优快速扩展随机树(AS-RRT*)无人机(UAV)航迹规划算法.该算法可以自适应调整采样空间,进而引导树更为高效地生长,而这些主要通过有偏采样、节点筛选和节点学习这三种策略来实现.首先,在采样空间中定义向光和背光区域来进行有偏采样,而向光和背光区域的概率权重由当前扩展失败率决定,从而保证算法在搜索初始航迹时同时具有探索性和方向性;然后,在完成初始航迹的搜索后,算法就开始周期性地筛选节点,高质量的节点作为学习样本来产生新的抽样分布,质量最低的节点在算法达到最大节点数量后被新节点替代.在多种不同类型的环境下进行了对比仿真实验,结果表明所提算法在一定程度上改善了采样算法固有的随机性,而且相较于传统的RRT*算法,该算法在相同环境里使用了更少的寻路时间,在相同时间里生成了更低代价的航迹,且在三维空间里的改进更为明显.

硬式飞艇升空前其内置氦气囊充气比例（饱和度）小于100%,氦气囊是非饱和的。氦气囊的非饱和形态影响飞艇内置氦气囊的布局形式和重心位置。为了研究硬式飞艇氦气囊的非饱和形态变化规律,首先将采用控制体积法获得的氦气囊的饱和形态作为初始状态,利用任意拉格朗日–欧拉方法模拟氦气囊的泄气过程,获得了气囊和流场之间的动态关系、不同饱和度的气囊形态及内压等主要参数。然后进行了地面氦气囊泄气试验,并将数值计算结果与试验结果进行了对比。结果表明,泄气初始阶段囊体的外形变化较小,气囊中上部维持饱和形态不变,气囊底部及其两侧形状最先变化;随着饱和度的降低,囊体底部被挤压的区域逐渐向上移动,囊内压力逐渐降低。该研究可为飞艇氦气囊设计及氦气囊的布局形式提供参考。

In order to study the movement after damping and the attitude of floating of the helicopter with emergency floatation bags. An arbitrary Lagrangian Eulerian (ALE) solver and a penalty coupling method were used for predicting the fluid-structure interaction forces. The finite element model contains Lagrangian shell elements for a helicopter with emergency air-bags and Eulerian solid elements for the water and air. Based on the Navier-Stokes equations, the simulation of three-dimensional numerical wave tank was built by ALE algorithm, adopt the method of numerical simulation, setting up the model with coupling of the helicopter and wave tank, obtain the moving posture after helicopter damping, the results show that the helicopter's roll angle between 0° ～ 14° before overturning, the configuration of the helicopter with airbag has good floating characteristics, which is conducive to the development of water rescue.

针对翼伞系统在实际飞行过程中受到环境因素及系统误差的影响而导致翼伞系统落点位置发生变化的问题,采用蒙特卡洛方法对翼伞系统在飞行过程中受到的多种随机干扰因素进行了分析,给出了各干扰因素的处理方法,建立了翼伞系统九自由度蒙特卡洛落点仿真模型,得到了各干扰因素对翼伞系统落点分布的影响规律。仿真结果表明,回收物质心位置偏差对翼伞系统落点分布影响较大,侧风干扰对翼伞系统在侧风方向落点分布有较大影响。有关结论可以为回收落区划分及回收物搜寻提供理论依据。

目的 对一代伞兵战车缓冲气囊进行优化,设计出满足二代伞兵战车着陆安全的缓冲气囊.方法 基于现用的缓冲气囊,结合有限元仿真、试验设计,建立以气囊参数为设计变量、战车着陆冲击加速度峰值为目标函数的支持向量回归(SVR)模型.用遗传算法对现有气囊的宽度、高度、排气孔面积、排气孔爆破压强进行参数优化.结果 优化后的缓冲气囊系统将二代伞兵战车的冲击加速度峰值降低了73.4％,着陆减速过程变得平稳顺滑.结论 成功地对现有缓冲气囊的参数进行了优化改进,得到了适用于二代伞兵战车的缓冲气囊,并且为决策者提供了多种气囊设计方案.

介绍了自主设计和研制的科氏惯性力实验装置的构造、原理和特点,以及据此开设的科氏惯性力实验.该仪器的研制贴近教学实际,为理论力学课程开设科氏惯性力实验提供了必要的条件.十多年的教学实践表明,科氏惯性力实验内容精彩,呈现的力学现象奇妙有趣,能够加深学生对科氏惯性力概念的理解,收到了良好的教学效果.

Aiming at the task demand of China's future reentry and deep space exploration, the research on mechanically-deployed entry decelerator is carried out. Based on the structural characteristics of the mechanically-deployed entry decelerator, the numerical model of the fluid-solid coupling of the aerodynamic surface of the entry decelerator in the three-dimensional unsteady environment is established using the space-time conservation element and solution element method. The deformation and stress of the flexible surface are analyzed by comparing the material modulus of elasticity in numerical simulation. The following conclusions can be get through comparative analysis: A more visible vortex appears on the back of the decelerator, which inhibits the deformation of the flexible surface. The modulus of elasticity range of the flexible material, the distribution and size of the deformation and stress, and the influence of the thickness of the flexible surface on the deformation can be get by analysis. It meets the requirements when the elastic modulus of the flexible plane is about 50GPa. The maximum deformation position lies at the edge of the flexible surface. The middle of the flexible surface forms an inward depression, and its maximum variable shape is smaller than the amount of deformation of the edge. The effect of the thickness of flexible surface on the deformation is also discussed .The results of this paper provide the basis for the aerodynamic characteristics and aerodynamic heating analysis of the flexible surface, and also provides reference for the engineering development of the flexible surface.

Non-interference model is the baseline security model of information flow control.It ensures zero leakage of secret information,but its restrictiveness of security condition is too strong.Software system inevitably violates non-in-terference model and releases proper information for its requirement of function.In order to prevent attacker obtain ex-tra information from the channel of information release,the channel should be under control and trusted declassification policy and enforcement mechanisms should be established.Existing declassification policies are classified into WHAT, WHO,WHERE and WHEN dimensions,and existing enforcement mechanisms are classified into static enforcement, dynamic enforcement and secure multi-execution.The characteristics and deficiencies of these mechanisms were com-pared,the challenge of following study was discussed,and the direction of future study was out-looked.

In addition to the structure parameters of the airfoil profile, the angle and the length of leading edge cut are the main parameters that affect the aerodynamic performance of the ram-air parachute. In this paper, based on the fundamental airfoil profile of the large ram-air parachute, two important parameters, the angle and the length of leading edge cut, are changed and four kinds of airfoil profile with representative parameters of leading edge cut can be obtained in order to find the range of engineering applications on leading edge cut parameters more accurately. The numerical calculation method is adopted to study the aerodynamic performance of four airfoils and the results show that the pressure gradient of upper surface grows and the pressure gradient of lower surface decreases with the increase of the angle of the leading edge cut at negative angles of attack, while the angle of leading edge cut does not have significant effect on the pressure gradient of the upper and lower surface at positive angles of attack. With the increase of the angle of leading edge cut, the maximum lift-drag ratio of the airfoils also increases, but there will be a threshold. The influence of the leading edge cut parameters on the pitch moment coefficient is not significant. The results of this paper provide some reference value for design and optimization of ram-air parachute.

According to the fixed landing and fixed point recovery of parafoil system's task,the theory of multiphase design is used to design the homing trajectory of parafoil system.In order to improve the accuracy of the landing point and reduce the energy consumption as the goal,the problem of trajectory planning is transformed to parameter optimizing on the basis of geometric relation of each phase trajectory,according to the initial height of the parafoil,different homing project are proposed,and using the left deflection controlled or the right deflection controlled respectively to plan the homing trajectory for the same initial position.The computational results show that the parafoil system use the left deflection controlled or the right controlled which can meet the requirements of the placement,but the energy consumed is different by the different operating modes,the trajectory with less energy consumption is selected as the optimal homing trajectory;The different homing project is selected with the different intial altitude of parafoil,for a large number of highly redundant and highly moderate initial points,the design method of the homing trajectory satisfies the requirements of parafoil homing accuracy,for the case where the initial height of the parafoil is low,through the design of homing project reduces the landing error,which is established a good foundation for parafoil's homing control.

For the need of design and homing project research of parafoil-payload systems,a method of dynamic modeling and simulation is presented.The flight dynamics of parafoil are calculated by the ADAMS software.According to the precise airdrop of Parafoil System′s task,the theory of multiphase design is used to plan the homing trajectory of parafoil system and a pid control system is set up to control the flight path of the parafoil system.The computational results show that,when the parafoil-payload system under the one side edge deflection controlled,the system′s turning performance will be affected,the flight path of the system is a spiral curve,and with the height of the system declining,the turning radius remains unchanged,the method of multiphase design is simple and is easy to implement,to meet the requirements of parafoil homing accuracy.

In order to research the security of rocket extra-cabin for internally gravity air-launch,the air-craft-rocket-parachute dynamic model is established based on multi-body dynamics software ADAMS;the aerodynamic forces are applied with dll subroutines. The influence factors and influence rules of the whole process safety of rocket extra-cabin are obtained by comparative analysis of the influences of launch pitch angle,wheel friction coefficients,speed of aircraft and pitch angle of the aircraft on the re-sults.

In order to study the lateral stability of the helicopter with emergency floatation bags,the move-ment after damping and the attitude of floating. Firstly,analyzing lateral static stability and dynamic stabil-ity of the helicopter with floatation bags through the calculation of the equal discharge of water theoretical method,and studied the influence of the different position of the center of gravity of the helicopter for sta-bility,it was concluded that when the aft center of gravity on limit position,the center of gravity is closer to the water surface,the helicopter has better stability. In helicopter damaged cases,the establishment of a helicopter with damaged compartment model, obtained the broken helicopter floating stability;finally through the damaged condition and the damaged helicopter floating time are analyzed.

降密策略的静态实施机制存在限制性过强的缺陷,基于虚拟机的动态监控机制不能完全适合Web和即时编译环境.为此,基于内联引用监控方法,实施了基于内容和地点维度的二维降密策略.提出了内联引用监控方法的程序变形规则,并证明了该方法的可靠性;根据该程序变形规则,将源程序进行变形重写,生成一个新的程序,它能脱离外部监控环境,实现自我监控.

基于金星探测的任务需求,美国提出了一种基于机械展开式结构的金星进入飞行器.为有效控制机械展开式飞行器的姿态,且得到其影响特征参数的稳定性限制范围,设计了移动质量块在初始位置往复正弦运动的控制规律;建立了单质量块的变质心进入飞行器系统控制模型,并利用CFD-FASTRAN软件计算了气动参数;对飞行器的变质心动力学模型进行简化;计算得到了质量块的移动幅度、质量比、飞行器的自旋频率特征参数的取值范围,保证了飞行器的姿态控制稳定性.