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Prediction of fatigue crack propagation in flat specimens and thinwalled structural members of the skin of the aircraft wing in


Our approach is verified by test calculations on an aircraft wing with two responses, namely, the lift and drag coefficient, and two variables, namely, the angle of attack and the Mach number.


This methodology is illustrated in the design of actuators for delaying flow separation on the leading and trailing edge devices of a multielement highlift system of a typical commercialaircraft wing at takeoff conditions.


In a collaborative research project, aircraft wing leading edge structures with a glassbased FML skin have been designed, built, and subjected to bird strike tests that have been modelled with finite element analysis.


Modelling of Bird Strike on an Aircraft Wing Leading Edge Made from Fibre Metal Laminates  Part 2: Modelling of Impact with SPH

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 A general formulation for oscillatory subsonic potential flows around threedimensional bodies of various configuration and its application to the calculation of dynamic stability derivatives of the aircraft are presented. By applying the Green function method, we obtained an integrodifferential equation relating the perturbation velocity potential to its normal derivatives on the surface of the body. In order to solve this equation, the surface of the body and its wave are divided into small quadrilateral... A general formulation for oscillatory subsonic potential flows around threedimensional bodies of various configuration and its application to the calculation of dynamic stability derivatives of the aircraft are presented. By applying the Green function method, we obtained an integrodifferential equation relating the perturbation velocity potential to its normal derivatives on the surface of the body. In order to solve this equation, the surface of the body and its wave are divided into small quadrilateral elements. The unknown φ and its derivatives are assumed to be constant within each element. Thus the integrodifferential equation reduces to a set of differentialdelay equations in time. This set of equations can be used as the basis of a general method for the fully unsteady flow calculation. For oscillatory subsonic potential flow, this set of equations further reduces to a set of linear algebraic equations which is solved numerically to yield the values of φ; at the centroid of each element. The pressure coefficient is evaluated by the finite difference method. The lift and the moment coefficients are determined by numerical integration of the pressure coefficient. The dynamic stability derivatives are obtained from the imaginary parts of the lift and the moment coefficients.The formulations in this paper are embedded into a general computer program. Several typical numerical results have been obtained by means of this program. Figure 2 shows the distribution of lift coefficient CL along the middle section for a rectangular wing oscillating in pitch with λ =2, τ =0.001, M∞ = 0, K = 2 .The result is identical to the original calculation by Merino. Figure 3 shows the distribution of pressure coefficient Cp for a harmonically oscillating spheroid witha/b= 8, M∞=0.5, K=2 . The result is in good agreement with the analytical solution of wave equation.Figures 5 , 6, 7 show the distributions of lift coefficient CL at various stations of an aircraft (wingbodytail combination) oscillating in pitch with M∞ = 0.6, K 0.005, 0.01. Vable 2 shows the dynamic stability derivatives CLa, Cma of the aircraft. The. results are in good agreement with the experimental data.  本文介绍处理不同外形三维物体亚音速振荡绕流一种统一的方法。本方法的主要特点是采用有限元素法直接解由格林定理导出的物面速势积分微分方程以求得物面的速势分布,然后再用有限差分法对速势进行微分求物面的压力分布。 由于本方法理论上比较严格,适用于复杂外形物体绕流的计算,所得结果又比较准确,因此近几年来在国外得到越来越广泛的应用。在本文中,采用了与有关文献相同的基本方程,但在气动影响系数的计算上略有不同,本文并将这一方法应用于飞行器动导数的计算,所得的结果与实验结果符合。  When the upper and/or lower panels of the loading box of the aircraft wing or tail are made of the fiber reinforced composite laminates, they can frequently be simplified as a nonmoment plate. This paper introduces an optimum design (i. e. minimum weight design) procedure of the laminated plate on static failure strength condition. The mathematical tool used in the procedure is Lagrangian Multiplier method, and the static failure strength condition is adopted as HillTsai criteria or Norris criteria.The... When the upper and/or lower panels of the loading box of the aircraft wing or tail are made of the fiber reinforced composite laminates, they can frequently be simplified as a nonmoment plate. This paper introduces an optimum design (i. e. minimum weight design) procedure of the laminated plate on static failure strength condition. The mathematical tool used in the procedure is Lagrangian Multiplier method, and the static failure strength condition is adopted as HillTsai criteria or Norris criteria.The formulae for optimum design have been not only derived, but also reformed to be convenient for the computer. How to establish the ultimate strength of laminates is discussed in detail. An example, illustrating the solution procedure and how to select the optimum scheme of lamination design, is presented in the paper. Some technical problems are briefly discussed in the last part.At the stage of the preliminary structural design, this procedure can be considered as an engineering method of lamination optimum design for the loading panel of laminated composite, which works under tension or compression (assuming that the buckling failure would not occur).  飞机翼面结构受力盒段的上下壁板若采用纤维增强复合材料叠层板则经常可简化为无矩叠层板;本文介绍这种叠层板按静力破坏条件的一种铺层优化设计方法。本法所用的数学工具为拉格朗日乘子法,采用的静力破坏条件为HillTsai判据或Norris判据。对于铺层优化设计要用的公式进行了推导并且为了便于应用电子计算机对所得的公式进行了相应的改造。 关于叠层板极限强度的确定,文中也作了较仔细的探讨。为了具体说明本法的解题步骤以及如何选出铺层设计的优化方案举出了一个数例。文章的最后对几个技术问题进行了简要的讨论。 在结构打样设计阶段,对于翼面结构中处于拉伸或压缩受载(在不发生屈曲失稳破坏的条件下)情况下的复合材料壁板,本法可作为其铺层优化设计的一种工程方法。  A new method for calculating threedimensional compressible laminar. and turbulent boundary layers on practical wings is described in this paper. In the method, a nonorthogonal coordinate system proposed by Cebeci is used, and a secondary coordinate transformation is suggested to overcome difficulties arised when negative transverse velocities appear locally. Several computational results for Lockheed C5A aircraft wing and F8 supercritical wing are given in the paper, and comparisons with other... A new method for calculating threedimensional compressible laminar. and turbulent boundary layers on practical wings is described in this paper. In the method, a nonorthogonal coordinate system proposed by Cebeci is used, and a secondary coordinate transformation is suggested to overcome difficulties arised when negative transverse velocities appear locally. Several computational results for Lockheed C5A aircraft wing and F8 supercritical wing are given in the paper, and comparisons with other methods and with experimental data are made as far as possible.  本文给出一种计算实际机翼可压缩层流紊流三维边界层的新方法。方法使用了Cebeci提出的一种非正交坐标系,并采用一种二次变换以克服在某处横流发生反号时计算遇到的困难。对Lockheed C5A机翼和F8超临界机翼给出了一系列计算结果,并同其它方法的结果以及试验数据作了比较。   << 更多相关文摘 
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