Parametric study of an elliptical fuselage made of a sandwich composite structure

We size the shell thickness of a pressurised elliptical fuselage and analyse the weight gains or savings compared to a circular fuselage. Three fuselage construction cases are analysed: a monolithic construction, a symmetric sandwich with two facesheets of equal thicknesses separated by a lightweight core, and an unsymmetric sandwich with two facesheets of different thicknesses. We develop a proper dimensionless analysis comparing the proposed elliptical fuselage with a circular one for two different scenarios: equivalent cross-section areas and enclosing a similar rectangular box. We apply a semi-analytical thin shell theory to compute the loading due to pressurisation around the circumference of the fuselage cross-section. We select the facesheet thicknesses above and below the sandwich core at every location to minimise weight. The goal is to compute the structural weight gain or penalty that is incurred by replacing a circular fuselage with an elliptical one to resist internal pressurisation in function of the scenario, eccentricity, fuselage diameter, and sandwich core thickness. We find that an elliptical cross-section incurs a significant penalty in terms of necessary facesheet thickness even with an optimised unsymmetric sandwich construction. This penalty can be minimised by keeping the eccentricity low, the loading intensity low and the core thick.     

Aeroelastic response of a coupled rotor/fuselage system in hovering and forward flight

Summary The aeroelastic response analysis of a coupled rotor/fuselage system is approached by iterative solution of the blade aeroelastic response in the non-inertial reference frame fixed on the hub, and the periodic response of the fuselage in the inertial reference frame. A model of the coupled system hinged with the flap and lag hinges, the pitching bearing which may not coincide with the hinges, and the sweeping-blade configuration is established. The moderate-deflection beam theory and the two-dimensional quasi-steady aerodynamic model are employed to model the aeroelastic blade, all the kinetic and inertial factors are taken into account in a unified manner. A five-nodes, 15-DOFs pre-twisted nonuniform beam element is developed for the discretization of the blade, three rigid-body-motion DOFs are introduced for the motion of the hinges and the bearing. The Hamilton's principle is employed to evaluate the equation of motion of the blade. The derived nonlinear ordinary differential equations with time-dependent periodic coefficients are solved by a modified quasi-linearization method, which is developed for the higher DOF periodic system. The resulting periodic forces and moments exerted on the fuselage by all the blades are evaluated every time, when the converged nonlinear periodic response of the blade is obtained under the consideration of the equilibrium of the blades. The fuselage structure is simplified to be a beam structure, the governing equation is established in the inertial reference frame and a two-nodes beam element is used to discretize the flexible fuselage. The periodic response of the fuselage is solved by a simple shooting method. The iteration of the rotor/fuselage response is continued, until the aeroelastic responses of the blade and the fuselage converge simultaneously. Both the hovering and the forward flight states can be considered. The results of a computed numerical example by the developed program are presented to verify in practice the economy of the modeling as well as the reliability and efficiency of the corresponding solving methods. Received 4 May 1998; accepted 11 August 1998     

开式压力机床身有限元分析与结构优选

分析了开式压力机床身的结构特点,然后介绍了应用有限元法对七式压力机床身进行结构分析时的力学模型简化方法,并结合若干工程实例,对开式压力机床身进行身进行了有限元分析与结构优选,取得了既减轻床身重量、又提高强度和刚度的显效果,为结构的合理设计与改进提供了可靠的理论依据。     

THE STRENGTHENING OF AERO-TRANSPORT WITH LARGE OPENING

为了对运输类飞机机身大开口结构进行加强,满足刚度连续变形协调的设计要求,本文对机身大开口结构和完整机身结构的刚度进行了深入研究,首先简化了计算模型,对刚度进行了计算,提出了刚度比的定义,得出刚度比与大开口角度、机身半径、蒙皮厚度以及边梁面积之间的关系表达式,得到运输类飞机机身大开口结构加强的原则和方法,在型号上成功得到了应用,用于指导初期的结构设计.     

OPTIMIZATION OF FUSELAGE OPENING ZONE OF CIVIL AIRCRAFT

在民用飞机机身上通常布置较多舱门开口,导致机身结构刚度发生急剧变化。为了保证飞机结构刚度以及载荷的传递,本文对机身开口结构的刚度及强度进行了深入研究,明确了开口结构刚度的影响因素,并对开口角度大小及加强结构尺寸进行了优化。以上研究对机身开口结构设计及加强提供了方向和方法,可以用于飞机机身开口结构初步设计阶段。     

Nonlinear aeroelastic coupled trim and stability analysis of rotor-fuselage

Based on the Hamilton principle and the moderate deflection beam theory, discretizing the helicopter blade into a number of beam elements with 15 degrees of freedora, and using a quasi-steady aero-model, a nonlinear coupled rotor/fuselage equation is established. A periodic solution of blades and fuselage is obtained through aeroelastic coupled trim using the temporal finite element method （TEM）. The Peters dynamic inflow model is used for vehicle stability. A program for computation is developed, which produces the blade responses, hub loads, and rotor pitch controls. The correlation between the analytical results and related literature is good. The converged solution simultaneously satisfies the blade and the vehicle equilibrium equations.     

Sliding mesh algorithm for CFD analysis of helicopter rotor–fuselage aerodynamics

The study of rotor–fuselage interactional aerodynamics is central to the design and performance analysis of helicopters. However, regardless of its significance, rotor–fuselage aerodynamics has so far been addressed by very few authors. This is mainly due to the difficulties associated with both experimental and computational techniques when such complex configurations, rich in flow physics, are considered. In view of the above, the objective of this study is to develop computational tools suitable for rotor–fuselage engineering analysis based on computational fluid dynamics (CFD). To account for the relative motion between the fuselage and the rotor blades, the concept of sliding meshes is introduced. A sliding surface forms a boundary between a CFD mesh around the fuselage and a rotor‐fixed CFD mesh which rotates to account for the movement of the rotor. The sliding surface allows communication between meshes. Meshes adjacent to the sliding surface do not necessarily have matching nodes or even the same number of cell faces. This poses a problem of interpolation, which should not introduce numerical artefacts in the solution and should have minimal effects on the overall solution quality. As an additional objective, the employed sliding mesh algorithms should have small CPU overhead. The sliding mesh methods developed for this work are demonstrated for both simple and complex cases with emphasis placed on the presentation of the inner workings of the developed algorithms. Copyright © 2008 John Wiley & Sons, Ltd.     

鸭式旋翼/机翼飞机悬停及小速度前飞气动干扰实验研究

鸭式旋翼/机翼飞机是一种新概念可垂直起降高速飞行器,为了解该飞机在悬停及小速度前飞时的全机气动干扰特性,在南京航空航天大学开口风洞中进行了飞机全机气动力实验,实验采用多台测力天平分别测量主机翼和机身的气动力.结果表明,悬停时受主机翼高速旋转产生的下洗尾流影响,机身产生了较大的法向力和低头力矩;前飞时下洗尾流对机身的法向力和俯仰力矩有比较严重的干扰,对滚转力矩和偏航力矩干扰较小,对侧向力有一定影响.实验结果为飞机的飞行动力学特性研究以及控制律设计提供了参考.＃     

An unstructured grid Reynolds stress model for separated turbulent flow simulations

Most of the turbulence models in the literature contain simplified assumptions which make them computationally inexpensive but of limited accuracy for the solution of separated turbulent flows. Dramatic improvements in computer processing speed and parallel processing make it possible to use more complete models, such as Reynolds Stress Models, for separated turbulent flow simulations, which is the focus of this work. The Reynolds Stress Model consists of coupling the Reynolds transport equations with the Favre–Reynolds averaged Navier–Stokes equations, which results in a system of 12 coupled non-linear partial differential equations. The solutions are obtained by running the PUMA_RSM computational fluid dynamics code on unstructured meshes. The equations are solved all the way to the wall without using any wall functions. Results for high Reynolds number flow around a 6:1 prolate spheroid and a Bell 214ST fuselage are presented. For the prolate spheroid basic flow features such as cross-flow separation are simulated. Predictions of circumferential locations of cross flow separation points are in good agreement with the experiment. A grid refinement study is performed to improve the computations. The fine mesh solution predicted locations of primary and secondary separation points with errors of roughly 2° and 0°, respectively. Flow simulations around an isolated Bell 214ST helicopter fuselage were also performed. Predicted pressure and drag force correlate well with the wind tunnel data, with a less than 10% deviation from the experiment. Drag predictions also show relative speed of Reynolds Stress Model compared to Large Eddy Simulation to compute time averaged quantities. For numerical solutions parallel processing is applied with the MPI communication standard. The code used in this study is run on Beowulf clusters. The parallel performance of the code PUMA_RSM is analysed and presented.     

直升机气动弹性力学发展现状(续)

Ⅲ.单片桨叶气动弹性问题的求解1.桨叶离散化方法求解旋翼桨叶气动弹性力学问题的第一步是将连续桨叶离散化,即把一个具有无限多个自由度的连续参数系统离散化为具有有限个自由度的离散系统。常用方法有三种: 1)整体模态方法在直升机旋翼气动弹性力学中,以往用得多的离散方法是整体模态法,或称为整体伽辽金方法。此方法的实质是利用桨叶自由振动振型是线性独立      

直升机旋翼／机身耦合系统的气弹响应分析（二）方程的求解

在文（一）得到的直升机旋翼系统运动方程的基础上,运用动态子结构的方法对旋翼系统和机身作为相对独立的部分进行分析,通过界布的力平衡和几何协调将子系统耦合系统整体系统,用分割－迭代法求解直至旋翼系统和机身的响应同时收敛到精度要求,并研制相应的计算程序,给出了工程算例。     

Towards consistent hybrid overset mesh methods for rotorcraft CFD

The overset mesh method chimera is popular within the rotorcraft research community, because the use of multiple, non‐matching grids make the CFD simulations of bodies in relative motion much simpler. Consequently, the relative motion between the helicopter blades and fuselage can be accurately accounted for. In this paper, the method for treating overset grids within CFD codes is presented. It is compatible with multi‐block, structured‐grid solvers. The proposed method is based on hierarchy of overset, non‐matching grids, whose cells are automatically identified as computational or non‐computational and localised with respect to all grids they overlap with. The efficiency of the method relies on the hierarchical, multi‐step approach, for the overset mesh localisation and the use of a tree search. Because of the high efficiency of the algorithm, the search for overlapping cells can be carried out on‐the‐fly, during time‐marching of the unsteady, implicit CFD solver. In addition, the algorithm is suitable for parallel execution. The method has been demonstrated for several flows, ranging from simple aerofoils to rotor‐body interaction. The paper presents and demonstrates the method and shows that it has a low CPU overhead. It also highlights the limitations of the method and suggests remedies for improvement. Copyright © 2013 John Wiley & Sons, Ltd.     

Axial crack propagation and arrest in a pressurized cylinder: An experimental-numerical analysis

The feasibility of using a previously developed crack-kinking criterion to predict crack arrest at a tear strap in a pressurized fuselage was studied with instrumented axial rupture tests of 21 models of an idealized fuselage. A rapidly propagating axial crack, which was initiated from a precrack, kinked immediately upon extension and propagated diagonally until it turned circumferentially and propagated along the tear straps. An elastodynamic finite element analysis of the rupturing model fuselage yielded the mixed-mode stress intensity factors,K _I andK _II , and the remote stress component, σ _OX . This numerical procedure was also used to predict the crack trajectories in full-scale fuselage rupture tests. All numerical results agreed well with their measured counterparts regardless of size.     

轮胎破片冲击机身的非等比例相似模型研究

为降低机身结构抗冲击性能的实验成本,利用相似理论建立机身的非等比例缩放模型,开展模型实验是行之有效的方法。基于量纲分析的方法,建立Johnson-Cook线性应变率函数的修正关系;鉴于生产制造技术的限制,考虑扭曲厚度的非等比例机身模型对相似性行为的影响,采用指数函数法建立了非等比例模型的相似修正关系。通过对比实验中破片冲击过程的变形形态、靶板的应变时间历程曲线和最终变形轮廓,验证了数值模型的有效性。此外,分析了破片偏航姿态、机身材料、厚度和质量等因素对机身结构抗冲击性能的影响。结果表明:(1) 150 m/s的冲击速度下,破片冲击角度90o和着靶角度180o是最严苛的冲击条件。综合多种因素,分析认为3.5 mm厚的钛合金为机身结构的最佳选择,并以此作为全尺寸原型验证相似模型;另外,提出了一种可以快速获取缩比模型的设计方法。(2)应变率效应对轮胎破片冲击机身结构的影响并不显著,等比例缩放模型与原型结果吻合较好。(3)厚度扭曲的非等比例模型能够有效地预测原型结构的变形行为;虽然,在时间尺度上,模型与原型存在一定的偏差;但是,在空间尺度上,非等比例相似模型能够有效地修正扭曲厚度造成中心最大挠度的预测误差,修正后的最大误差不超过5.1%,这表明该方法能够有效地指导机身结构的相似模型设计。     

混合专家系统及其在有限元模型化中的应用

本文结合传统专家系统和人工神经网络的优势建立了混合专家系统模型，并在此基础上建立了“飞机结构机身框有限元模型化专家系统”，能够自动地实现机身框一类结构的有限元模型化，本文采用人工神经网络解决结构类型的分类和识别、采用知识系统解决单元自动选取都取得了良好的效果。     

Fracture mechanics of plates and shells applied to fail-safe analysis of fuselage Part II: Computational results

In this paper, the problem of the fracture of a fuselage stiffened by longitudinal longerons and circumferential frames is analyzed by means of the finite element method. Our research is motivated by the fail-safety design concept of fuselage for civil aircraft. In this study, the total energy release rate are evaluated for five types of basic loading, namely, axial extension, pure bending, twisting, transverse shearing, and radial expansion due to internal pressure. The crack is located either at the mid-point or near the end of the fuselage. It extends in two bays with the stiffener at its center. The stiffener which bisects the crack is assumed to be broken at the location of the crack. Computational results indicate that the total energy release rate G_t increases with the increasing crack length. However, when the crack tip approaches the stiffener, the value of G_t decreases as a result of the reinforcement from the stiffener. For a crack near the end of the fuselage, as a result of boundary effect, the value of G_t is larger in comparison with the case of the crack at the mid-point of the fuselage. We also find that the effect of geometrical nonlinearity can reduce the value of G_t for the fuselage under axial tension or pure bending. For the fractured fuselage under pure bending, shell buckling can occur at the concave side of the fuselage prior to crack growth. The maximum tensile stress in the stiffener in front of the crack tip is also investigated.     

民机中央翼舱段适坠性仿真

建立了带中央翼的民机舱段的有限元模型,应用MSC.Dytran分析软件对民机中央翼舱段模型进行坠撞仿真,得到加速度响应和客舱舱体结构变形结果.结果表明,带中央翼舱段发生坠撞后,乘客的生存空间基本没有变化,撞击能量主要靠舱段下部的龙骨梁、框及蒙皮和中央翼壁板的破坏来吸收,地板导轨的峰值加速度比不带中央翼的典型舱段的峰值加速度要大许多.     

飞机槽型大开口结构刚度加强研究

为了飞机典型的方舱型机身大开口结构能够满足刚度设计要求,设计了一种对槽型大开口结构增加4个边梁进行刚度加强的设计方案,并构建了工程分析模型,与无开口槽型结构的弯曲刚度及扭转刚度进行了对比研究,得到了两种构型的刚度比,进一步得到了满足一定刚度指标下的边梁面积计算公式,提出了方舱型机身大开口结构刚度设计流程,可以用于方案阶段的飞机大开口结构加强设计。     

水陆两栖飞机波浪水面上降落耐波性数值分析

在规定的气象水文条件下,水陆两栖飞机起飞和降落的能力是决定其性能的重要因素,即耐波性能。采用ALE方法对流体域进行描述,运用基于微幅波理论的动边界数值造波方法模拟了不同波高和不同波长的动态海平面波浪,通过添加质量阻尼的消波方法抑制了固壁边界反射波对造波结果的影响,并采用罚函数耦合方法描述飞机与水体的耦合作用,研究了水陆两栖飞机在不同海情条件下波浪面上降落的纵摇运动、升沉运动以及底部压力等运动学和动力学特性,分析了水陆两栖飞机入水波浪的波长及波高对水陆两栖飞机耐波性能的影响,为飞机结构设计、水上降落操作规则制订及水陆两栖飞机耐波性物理水池试验提供参考。