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  wing surface
The variation of the boundary conditions on the wing surface with time and coordinates may be arbitrary.
      
Numerical solution of the problem of supersonic gas flow past an arbitrary delta wing surface in the compression region
      
The velocity field generated by wing vibrations propagating along an elastic wing surface with finite velocity is studied.
      
The Newton method [4] was used to calculate the pressure distribution over the wing surface.
      
A numerical investigation is carried out and some results of calculating the unsteady viscous shock layer equations for various forms of the time dependence of the injection velocity and wing surface temperature are presented.
      
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  aerofoil
Laminar boundary layer in a flow past an aerofoil with a circular cavity
      
The paper describes a numerical study of a method of preventing the separation of a laminar boundary layer from the forward section of a symmetric aerofoil, the flow past which does not separate at zero angle of incidence.
      
In order to increase the maximum angle of incidence at which the flow has still not separated, a circular cavity (vortex cell) located almost completely inside the aerofoil is introduced on the segment vulnerable to separation.
      
A study of 3D aerodynamic design for a transonic compressor blading optimized by the locations of aerofoil maximum thickness and
      
An aerofoil above which is built the artificial cavity low pressure region is called "cavitating airfoil".
      
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  a wing
Therefore, the method makes it possible to examine the aperiodic motion of a wing as a rigid body, consider any wing deformations, analyze the wing entry into a gust, study the effect of a weak shock wave on the wing, etc.
      
The entry of a wing into the zone of a sharp-edged gust is considered in the linear formulation.
      
Problems of unsteady flow about a wing are examined in [5, 6].
      
Calculation of aerodynamic characteristics of a wing profile with discharge of a controllable jet into the external flow
      
A wing profile of infinite span, whose lower surface is replaced by a system of guide vanes, is placed in a flow of an ideal incompressible fluid.
      
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An investigation Was conducted to determine the effects of geometrical variables on longitudinal and lateral aerodynamic characteristics of parawings at low speed. The variables under study included variations in the shape of the trailing-edge,. the rib-strips, the apex angles, the strakes, the chord of the outer part of the wing, the shape of the keel, the slackness ratio, and the dihedral ange.The results of the study indicate that, for single keel two lobed parawing, the shape of the trailing-edge has a significant...

An investigation Was conducted to determine the effects of geometrical variables on longitudinal and lateral aerodynamic characteristics of parawings at low speed. The variables under study included variations in the shape of the trailing-edge,. the rib-strips, the apex angles, the strakes, the chord of the outer part of the wing, the shape of the keel, the slackness ratio, and the dihedral ange.The results of the study indicate that, for single keel two lobed parawing, the shape of the trailing-edge has a significant effect on the aerodynamic characteristics, the maximum lift-drag ratio with inward curved trailing edges can be much higher than that with straight trailing edges. Adding ribs to the flexible wing surface results in a decrease in drag and increase in lift-drag ratio . For parawings of higher apex angles , incresing the apex angle results in higher lift-drag ratio, and adding a small strake can increase the maximum lift coefficient and improve the stall characteristics. The lift characteristics can also be improved using a curved keel with proper curvature .Increasing slackness ratio exhibits a stabilizing increment in the directional stability Incresing the dihedral angle exhibits a stabilizing increment in the lateral stability, but contributes a destabilizing effect to the directional stability. within a certain range, the variation of the dihedral angle plays an inverse role on the lateral as well as the directional stability as that of the slackness ratio .

本文研究了伞翼几何参数变化对纵向与横侧气动特性的影响。在实验中改变的因素包括伞翼后缘形状,翼面上的肋条,伞翼顶角,边条小翼,外翼弦长,龙骨形状,伞翼的张开比和上反角。 研究结果表明,对于单龙骨双叶伞翼,后缘形状改变对气动特性有很大影响,采用向内弯曲的后缘与直后缘的情况相比,能使伞翼的最大升阻比提高很多。在翼面上加肋条,使伞翼的阻力减小、升阻比增大。对于顶角比较大的翼面,增加顶角将获得更大的升阻比。在这种翼面上加边条小翼,可使最大升力系数提高并且改善失速特性。采用适当弯曲的龙骨也使升力特性得到改进。 增加伞翼的张开比使航向静稳定度增大。增加伞翼的上反角使横向静稳定度增大,航向静稳定度减小。在一定范围内,改变上反角对横向和航向稳定性的影响与张开比的作用相反。

In this paper, a simple but effective modeling of the turbulence response to a sudden adverse pressure gradient has been incorporated into a boundary layer auxiliary equation. The modification attempts to reproduce the lag process of the response, so that the predicted accuracy of boundary layer parameters (displacement thickness, momentum thickness, shape factor and skin friction coefficient) in shock-wave boundary layer interaction and airfoil trailing edge regions at transonic speed is improved. The modification...

In this paper, a simple but effective modeling of the turbulence response to a sudden adverse pressure gradient has been incorporated into a boundary layer auxiliary equation. The modification attempts to reproduce the lag process of the response, so that the predicted accuracy of boundary layer parameters (displacement thickness, momentum thickness, shape factor and skin friction coefficient) in shock-wave boundary layer interaction and airfoil trailing edge regions at transonic speed is improved. The modification also improves the preceding parameters accuracy at high subsonic speeds, in particular, in airfoil trailing-edge region. This computational method is simple and convenient, and only a little computational time is needed. If about 120 interpolation points are used alone the surface of the airfoil, the computation time is about 25, minutes on a Felix c-125 computer.

本文注意到翼面的压力突跃在湍流附面层中的传递过程对附面层流动历程所产生的明显影响。文中提出了对平衡状态湍流附面层速度亏损型方程的改进形式以体现这种影响,力求准确地模拟附面层对压力突跃响应的滞后过程,从而改善跨音速翼面激波与附面层相互干扰区域及翼型后缘区的附面层参数的计算精度。计算结果表明,经改进的速度型方程的适用范围扩大到翼面存在较强激波的情况。干扰区内参数的计算精度达到很令人满意的程度。改进也使高亚音速翼型,尤其是后缘区的参数精度也有了提高。该方法计算简便,需用机时少。

A new method is proposed for calculating supersonic unsteady aerodynamic forces in this paper.It combines the piston theory for unsteady flow with the conical flow theory for steady flow.Its essential feature lies in that the interaction of points on a wing neglected in the piston theory is approximated by conical flow theory.This method was applied to flutter calculations of 29 cases for wings with 10 different types of plan forms.The results have been compared with flutter test data obtained in wind tunnels...

A new method is proposed for calculating supersonic unsteady aerodynamic forces in this paper.It combines the piston theory for unsteady flow with the conical flow theory for steady flow.Its essential feature lies in that the interaction of points on a wing neglected in the piston theory is approximated by conical flow theory.This method was applied to flutter calculations of 29 cases for wings with 10 different types of plan forms.The results have been compared with flutter test data obtained in wind tunnels in satisfactory agreement.The accuracy of flutter analysis improved by this method is much higher than that obtained by the piston theory.It is shown that this method is feasible and reliable for supersonic flutter analysis.It provides the advantages of briefer algorithms,higher numerical accuracy,less computing time and easier programming.

本文提出了一种计算超音速非定常气动力的新方法,它将超音速非定常的活塞理论和定常的锥形流理论结合起来,使得被活塞理论所忽略的翼面上各点间的相互影响用锥形流理论来近似计入。文中运用该方法对十个不同平面形状的机翼进行了二十九种状态的颤振计算,并和风洞颤振实验结果作了比较。比较的结果是令人满意的。特别是和用活塞理论所做的颤振分析相比,计算精度有了明显提高。这表明本方法用于超音速颤振分析是确实可行的,它具有方法简洁,计算精度高,计算时间少,程序编制容易的优点。

 
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