Numerical modeling of the formation of aerosol particles in jet engine plumes

The formation of liquid sulfate aerosols in the isobaric axisymmetric plume of a subsonic aircraft is modeled numerically. The specific features of the appearance and evolution of sulfate aerosols attributable to 2D effects, such as the parameter nonuniformity in the initial section (at the nozzle outlet), mixing of the hot engine jet with the cold air stream, and the transverse turbulent diffusion of aerosol particles and gas mixture components. The equations of gas dynamics for a turbulent axisymmetric jet, the equations of chemical kinetics, the equations for the liquid fractions (water and sulfuric acid), and the relations for the binary nucleation, condensation growth and coagulation of aerosol particles are used. The distributions of the parameters determining the formation of the aerosol phase in the exhaust plume of a B-747 aircraft are obtained and the geometry of the nucleation zone in this plume is determined.     

Transverse Elastic–Friction Vibrations of a Running Aircraft

An analytical model of the plane-parallel running of an aircraft under crosswind load is proposed. The compliance of the internal constraints of the aircraft and the nonlinearity of the side force on the landing gear wheels are taken into account. The model is intended to study the transverse elastic friction vibrations of the aircraft hull and landing gears. The steady motion of the aircraft is considered. The dependence of the lateral friction force on the slip angle is assumed to have the form of a sinusoidal segment. The Routh–Hurwitz instability conditions are derived in an explicit form     

CALCULATION OF AIRCRAFT DYNAMIC LOADS ON SHIP-DECK AND TEST VERIFICATION

舰载飞机的舰面动载荷是结构的临界设计载荷。针对常规方法下载荷计算的设计已不能满足舰载飞机非常规工况(例如弹射拦阻、自由飞钩住、偏心着舰等)载荷计算的设计要求的问题,研究和建立了一套完整的全机动载荷计算的设计方法和流程,并以某型飞机全机落震试验为例,通过与试验数据的对比分析,验证了该设计方法和流程的合理性和可行性。     

Identification of dynamic properties of plate-like structures by using a continuum model

The common approach currently used in the aircraft structural analysis is the finite element method. NASA's research in computational structures technology (CST) is helping to develop the finite element analysis to a new stage, although the significant limitations still exist. The elements used in the finite element method are usually void of dynamics. The consequence is that hundreds and thousands of elements are needed to represent large flexible aircraft structures in order to acquire analytical accuracy. To avoid the large dimensionality the current practice is to reduce the order of the model for structural system identification and control synthesis. This approximation, however, can lead to system instability due to the dynamics which are ignored.In contrast, distributed parameter modeling seems to offer a viable alternative to the finite element approach for modeling large flexible aerospace structures. Distributed parameter models have the advantage of improved accuracy, reduced number of modal parameters, and the avoidance of modal order reduction. Most of the effort on the continuum modeling so far is contributed to the beam-like structures which are composed of beams, tethers and rigid bodies. For the aircraft structural analysis, however, another important type of structural elements is plate. The principle of the monocoque or semi-monocoque type of aircraft construction is fundamentally the use of a thin-walled tube to carry compression, tension, shear, and bending. It is necessary, therefore, to expand the continuum modeling methodology to the plate-like structures to satisfy the requirement in the aircraft structural analysis, especially for the monocoque structures.This paper has developed a continuum modeling algorithm for the identification of dynamic properties of plate-like structures. A closed-form solution of the Timoshenko plate equation consistent with the maximum likelihood estimator has been derived. The closed-form expressions of the gradient functions have thereby been resulted from the solution of the partial differential equation. The proposed distributed parameter model involves far fewer unknown parameters than independent modal characteristics for finite element models. Illustration of this approach is given by a computer simulation which shows that the estimated results by using continuum model are reasonably accurate compared with the theoretical results.     

Mechanical phenomena in ground run of an aircraft with near-critical slip angles

The paper is concerned with mechanical phenomena accompanying ground run of aircraft with near-critical slip angles. Such phenomena (reduced effectiveness of directional control, self-excited frictional vibrations of a landing gear, internal resonance of an aileron, transverse oscillations of the legs of a multileg landing gear) are due to the nonlinear dependence of the side force on the slip angles __________ Translated from Prikladnaya Mekhanika, Vol. 42, No. 6, pp. 128–135, June 2006.     

On the non-linear vibrations of a circular membrane

Free axisymmetric vibrations of a stretched circular membrane are studied using a membrane theory consisting of a pair of non-linear partial differential equations coupled between the transverse and radial displacements of the membrane. A systematic perturbation method, in which the amplitude of the transverse displacement is taken as the perturbation parameter, is used to obtain periodic solutions of the non-linear equations. The initial membrane strain enters the problem as a parameter which is allowed to vary over a range of values. A case of self-resonance is encountered when the initial membrane strain approaches some critical values. This self-resonance case is also treated through an appropriate modification of the perturbation method.     

Similarity solutions for power-law and exponentially stretching plate motion with cross flow

Laminar boundary layers generated by power-law plate stretching with cross flows are studied. Only the stretching solutions of Banks [10] are considered, those being bounded by exponentially stretched plates. In one case the cross flow is generated by a uniform transverse stream far above the stretching plate or a wall moving with uniform transverse velocity. Two other cases deal with cross flows generated by transverse shearing motions of the surface. Possible two parameter solutions appear, but here we present two one-parameter families of cross flow solutions generated by transverse plate shearing motion. Streamwise and transverse shear stresses and velocity profiles are displayed in graphical form.     

考虑气动参数扰动的再入轨迹快速优化方法

针对传统再入轨迹优化方法收敛速度慢、对初值敏感程度高等的局限性,提出了一种基于序列凸优化的再入轨迹快速求解方法.该方法以倾侧角的变化率作为控制量,改进了现有凸化策略,考虑到抑制数值优化过程中由于数值离散方式带来的锯齿化现象,采用 B 样条曲线离散控制量,同时为避免算法在初始猜想值附近出现伪不可行的问题,增加额外虚拟控制量,通过一种"回溯直线"搜索的方法,提高算法的稳定性、快速性和寻优结果的光滑性.为研究飞行器再入过程中的气动参数扰动问题,采用采样点少、易于实现,计算效率高的广义混沌多项式理论研究方法,建立了基于广义混沌多项式和凸优化相结合的再入轨迹鲁棒优化模型,该模型在优化过程中考虑气动参数扰动对寻优结果的影响作用,避免了传统轨迹与制导律的复杂迭代设计环节,可有效降低优化轨迹对气动参数扰动的敏感程度,在气动参数不确定条件的干扰下,依然可以保证飞行器顺利安全的完成飞行任务.最后,以美国某可重复使用飞行器的再入任务为例,验证了基于序列凸优化的再入轨迹优化方法的快速性以及鲁棒优化模型对气动参数扰动的抗干扰性能力,表明了该方法具有一定的工程应用性.      

FAST OPTIMIZATION METHOD OF REENTRY TRAJECTORY CONSIDERING AERODYNAMIC PARAMETER PERTURBATION

针对传统再入轨迹优化方法收敛速度慢、对初值敏感程度高等的局限性,提出了一种基于序列凸优化的再入轨迹快速求解方法.该方法以倾侧角的变化率作为控制量,改进了现有凸化策略,考虑到抑制数值优化过程中由于数值离散方式带来的锯齿化现象,采用 B 样条曲线离散控制量,同时为避免算法在初始猜想值附近出现伪不可行的问题,增加额外虚拟控制量,通过一种"回溯直线"搜索的方法,提高算法的稳定性、快速性和寻优结果的光滑性.为研究飞行器再入过程中的气动参数扰动问题,采用采样点少、易于实现,计算效率高的广义混沌多项式理论研究方法,建立了基于广义混沌多项式和凸优化相结合的再入轨迹鲁棒优化模型,该模型在优化过程中考虑气动参数扰动对寻优结果的影响作用,避免了传统轨迹与制导律的复杂迭代设计环节,可有效降低优化轨迹对气动参数扰动的敏感程度,在气动参数不确定条件的干扰下,依然可以保证飞行器顺利安全的完成飞行任务.最后,以美国某可重复使用飞行器的再入任务为例,验证了基于序列凸优化的再入轨迹优化方法的快速性以及鲁棒优化模型对气动参数扰动的抗干扰性能力,表明了该方法具有一定的工程应用性.     

民机复合材料结构冲击后压缩设计值的确定方法

冲击后压缩设计许用值是复合材料飞机结构设计的一个非常重要的参数。本文通过分析常用的目视检测方法和飞机结构冲击能量统计结果,从凹坑深度和冲击能量截止值两方面定义了目视勉强可见损伤;根据民机设计兼顾安全性和经济性的特点,作者提出压缩设计许用值中的目视勉强可见损伤应综合考虑,慎重确定。文章还研究了复合材料层压板的抗冲击性能,从工程应用角度给出了结构冲击后压缩设计许用值的确定要点,即用小试样的冲击后压缩试验结果推导的基准系数和环境系数,对元件乃至典型结构件的冲击后压缩试验结果进行修正。     

Nonlinear dynamics of a cracked rotor in a maneuvering aircraft

The nonlinear dynamics of a cracked rotor system in an aircraft maneuvering with constant velocity or acceleration was investigated, The influence of the aircraft climbing angle on the cracked rotor system response is of particular interest and the results show that the climbing angle can markedly affect the parameter range for bifurcation, for quasiperiodic response and for chaotic response as well as for system stability. Aircraft acceleration is also shown to significantly affect the nonlinear behavior of the cracked rotor system, illustrating the possibility for on-line rotor crack fault diagnosis.     

用于鸟撞试验的仿真鸟弹研究

对用于飞机结构抗鸟撞试验的仿真鸟弹进行了研究,给出了配方,其基体材料为明胶和水.密度是仿真鸟弹的关键参数,通过工艺流程和添加调质材料对其进行控制.鸟弹外形尺寸和重量由模具保证,制作出了满足规范要求的1.8kg标准形态鸟弹.利用4块12mm厚的铝板,进行了3次真实鸟弹试验和1次仿真鸟弹的鸟撞对比试验.试验结果表明,仿真鸟弹具有足够的强度可承受发射过载.同时仿真鸟弹与真实鸟弹的动态变形模式以及结构动态应变响应时间历程基本一致.其结构动态应变响应最大值仅相差3.2％,而结构残余变形相差8.7％.上述结果证明了本文研制的仿真鸟弹可以在结构抗鸟撞试验研究中替代真实鸟弹.     

基于高阶变形理论的夹层板的弯曲与振动

考虑面板和夹芯的面内刚度和横向剪切刚度以及抗弯刚度,考虑了高阶剪切变形,根据横向剪应变分布情况给出横向剪切转角的位移函数,基于哈密尔顿原理,推导了基于高阶变形理论、适用于软、硬夹芯情况夹层板的基本方程。作为算例,以四边简支条件下的夹层板的弯曲与振动,在不同的面板与芯层的弹性模量比和厚度比下进行了计算,并与Reissner理论、Hoff理论以及邓宗白基于Reissner理论的修正模型的计算结果进行了对比。与前述理论与方法相比,本文方法考虑因素更为全面,对夹层板的适用范围更为广泛,计算结果更为精确。针对Nastran软件计算夹层板的振动问题,对其适用范围作了简要分析。     

折叠翼飞行器的动力学建模与稳定控制

折叠翼飞行器在变形过程中,其动力学模型呈现多刚体、多自由度和强非线性特点,同时气动力/力矩、压心、质心和转动惯量等参数也会大幅度变化,严重影响飞行稳定性. 由此,本论文将对飞行器的多刚体动力学建模与变形稳定控制进行研究.基于凯恩方法建立了折叠翼飞行器的多刚体动力学模型,并从中得到了变形所产生的附加力和力矩表达式.通过气动计算拟合出气动参数与折叠角之间的函数关系,由此分析了不同折叠角速度下飞行器的纵向动态特性, 结果表明,折叠翼飞行器变形过程中速度、高度和俯仰角均会发生变化,飞行器无法保持稳定飞行.为此提出了一种基于自抗扰理论的飞行器变形过程中的稳定控制方法.将折叠翼飞行器纵向非线性动力学模型中存在的非线性项、耦合项以及参数时变项都视为系统内外总扰动,利用扩张状态观测器对总扰动进行实时估计和补偿, 针对补偿后的系统设计PD控制器,实现了速度通道和高度通道的解耦控制.通过Lyapunov稳定性原理证明了系统的稳定性, 并进行数学仿真验证. 仿真结果表明,基于自抗扰理论设计的稳定控制器能够解决飞行器变形所带来的强非线性和参数时变等问题,保证飞行器的高精度稳定控制.      

DYNAMIC MODELING AND STABILITY CONTROL OF FOLDING WING AIRCRAFT 1)

折叠翼飞行器在变形过程中,其动力学模型呈现多刚体、多自由度和强非线性特点,同时气动力/力矩、压心、质心和转动惯量等参数也会大幅度变化,严重影响飞行稳定性. 由此,本论文将对飞行器的多刚体动力学建模与变形稳定控制进行研究.基于凯恩方法建立了折叠翼飞行器的多刚体动力学模型,并从中得到了变形所产生的附加力和力矩表达式.通过气动计算拟合出气动参数与折叠角之间的函数关系,由此分析了不同折叠角速度下飞行器的纵向动态特性, 结果表明,折叠翼飞行器变形过程中速度、高度和俯仰角均会发生变化,飞行器无法保持稳定飞行.为此提出了一种基于自抗扰理论的飞行器变形过程中的稳定控制方法.将折叠翼飞行器纵向非线性动力学模型中存在的非线性项、耦合项以及参数时变项都视为系统内外总扰动,利用扩张状态观测器对总扰动进行实时估计和补偿, 针对补偿后的系统设计PD控制器,实现了速度通道和高度通道的解耦控制.通过Lyapunov稳定性原理证明了系统的稳定性, 并进行数学仿真验证. 仿真结果表明,基于自抗扰理论设计的稳定控制器能够解决飞行器变形所带来的强非线性和参数时变等问题,保证飞行器的高精度稳定控制.     

Stress state of a transversely isotropic spherical shell with a rigid circular inclusion

Analytical expressions are derived for the stresses near a rigid circular inclusion in a transversely isotropic shallow spherical shell under uniform pressure. The form of solution depends on the range of the transverse shear compliance parameter. The influence of the relative radius of a rigid inclusion and transverse shear compliance on stress concentration is analyzed __________ Translated from Prikladnaya Mekhanika, Vol. 41, No. 12, pp. 67–73, December 2005.     

基于修正偶应力和高阶剪切变形理论的变截面微梁的自由振动

基于修正偶应力和高阶剪切理论建立了仅含有一个尺度参数的Reddy变截面微梁的自由振动模型,研究了变截面微梁自由振动问题的尺度效应和横向剪切变形对自振频率计算的影响。基于哈密顿原理推导了动力学方程与边界条件,并采用微分求积法求解了各种边界条件下的自振频率。算例结果表明,基于偶应力理论预测的变截面微梁的自振频率均大于经典梁理论的预测结果,即捕捉到了尺度效应。另外,梁的几何尺寸与尺度参数越接近,尺度效应就越明显,而梁的长细比越小,横向剪切变形对自振频率的影响就越明显。     

Eigenproblem of substructure chain and the expansion solution

In this paper, a new method for dealing with substructure chain is presented. Here, a substructure chain means a structure consisting of a number of identical substructures connected in series. This method is parallel to the expansion method in the structural analysis according to the transverse section. The method of shift invariance and that of initial parameter and transfer matrix, their mutual-relation and combined use, as well as the analysis of the eigenproblem on the transverse section, and the expansion method based on the eigenvectors are described in this paper. Project supported by National Natural Science Foundation of China.     

On the role of a balancing load in the problem of aerodynamic drag minimization in the supersonic flight regime

The problem of reducing the aerodynamic losses through balancing in the supersonic flight regime is considered. An analysis and a comparison of the drag components due to the aerodynamic surface deformation and the balancing weigh distribution is made with reference to the examples of a zero-thickness airfoil and a three-dimensional configuration of an aircraft with a wing of complicated planform. It is shown that minimum values of the aerodynamic drag are achieved as a result of complex optimization including the dead load mass as a varied parameter.     

Technical stability of nonlinear states of an elastic vehicle in vertical flight

Sufficient conditions of technical stability of nonlinear dynamic states of extended elastic flying systems in controlled longitudinal vertical flight are obtained. In these flying systems, the effect of variation of their crosssectional area, transverse strains, and oscillations is taken into account. The formulated criteria of technical stability depend on the basic parameters of the process controlled, namely, on the increment of the transverse load due to the curvature of the axis of the system and aerodynamic forces during vertical flight.