2019 年 2 月 第 37 卷第 1 期 西 北 工 业 大 学 学 报 Journal of Northwestern Polytechnical University Feb. Vol.37 2019No.1 https: / / doi.org / 10.1051 / jnwpu / 20193710167客机均采用简单襟翼及 定轴旋转机构,以方便实现巡航阶段改变机翼的后 缘弯度。 然而,简单襟翼的效率普遍会低于传统的 富勒襟翼。 因此为了保证起飞着陆阶段较好的气动 性能,需要下偏扰流板,以保证良好的缝道参数,提 高增升装置效率 [3•4] 。 当飞机接近地面飞行时,如起飞、着陆阶段,迫 使被其扰动的气流方向向平行于地面的方向改变, 作用在飞机上的气动力也发生改变,这种现象被称 作地面效应 [5] 。 文献[6•7] 指出地面效应对干净构 型和增升构型的影响规律不同。 例如普通干净构型 的地效飞行器在近地飞行时,升力增加,可以获得较 大的升阻比,从而提高巡航效率。 而文献[ 8] 的研 究表明地面效应使得多段翼型的升力系数减小,这 将直接影响飞机的起降性能。 因此,研究地面效应 对带扰流板下偏的多段翼型气动性能的影响,对于 我国发展远程宽体客机具有重要的实际意义。 目前,国内外学者对单段翼型地面效应已经进 行了大量的研究,而对多段翼型的研究工作较少。 Hsiun 等 [9] 通过数值模拟研究了弯度和厚度对二维 翼型地面效应的影响,发现下翼面与地面之间的气 流通道决定了翼型的气动特性;Ahmed 等 [7] 对迎角 0°~ 10°范围内 NACA4412 翼型的地面效应进行了 实验研究,实验结果表明随着离地高度逐渐减小,阻 力逐渐增大,但升力系数变化趋势取决于随迎角变 化的翼型下表面与地面之间的气流通道形状;屈秋 林等 [10] 对 NACA4412 翼型在大范围迎角下的地面 效应进行了数值模拟,结果表明升力系数变化趋势
Angle of attack has impact on transonic flow filed and aerodynamic force, but most of current researches on flutter use zero angle hypothesis, which has no consideration about angle of attack. Therefore, we use unsteady Reynold Averaged Navier-Stokes (RANS) equation and structural dynamic equation to establish the time domain aeroelastic analysis method. The solution in time domain is the fourth-order implicit Adams linear multi-step method which is based on prediction-correction method. The numerical simulations were used to analyze the transonic flutter boundary of Isogai Case A Model which was based on zero angle condition and nonzero angle respectively. The simulation results show that the reduced flutter speed decreases as the preset angle of attack decreases between 0.73 and 0.76, which shows a 12.5% decrease of the flutter speed at the farthest. Nonzero angle makes the transonic dip weaker and wider than fully turbulent flow. Changing in angle of attack of 6°, the flutter speed in the deepest position of transonic dip has increased by 124% compared to the flutter speed of 0°. Therefore, when flutter characters of airfoil is analyzed, the effects of the initial angle of attack must be taken into account in order to analyze flutter boundary correctly. In other words, increasing the angle of attack offers a way to control the system in terms of delaying flutter.
The internal blown flap was numerically simulated. Firstly, a parameterization method was developed, which can properly describe the shape of the internal blown flap according to such geometrical parameters as flap chord length, flap deflection, height of blowing slot and its position. Then the reliability of the numerical simulation was validated through comparing the pressure distribution of the CC020-010EJ fundamental generic circulation control airfoil with the computational results and available experiment results. The effects of the geometrical parameters on the aerodynamic performance of the internal blown flap was investigated. The investigation results show that the lift coefficient increases with the increase of flap chord length and flap deflection angle and with the decrease of height of blowing slot and its front position. Lastly, a method of optimal design of the geometrical parameters of the internal blown flap was developed. The design variables include flap chord length, flap deflection, height of blowing slot and its position. The optimal design is based on maximum lift coefficient, the angle of attack of 5 degrees and the design constraint of stall angle of attack of less than 9 degrees. The optimization results show that the optimal design method can apparently raise the lift coefficient of an internal blown flap up to 1.7.
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