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我们将一个实验马赫数范围为0.4至0.9的高亚音速风洞经过扩建后成为一座跨音速风洞, 实验马赫数已扩大到0.3至1.2, 实验段已由半开式改为跨速音孔壁形式, 截面尺寸为530×760毫米, 孔壁开闭比为23%, 扩开全角为0.6°。跨音速实验段孔壁与槽壁的选择, 从消除亚音速洞壁效应来说, 两者相差无几, 从消除激波反射效应考虑, 孔壁明显的优于槽壁, 但孔壁扩开角宜小。孔径与壁厚之比及孔径与实验段高度之比, 对孔壁流动特性和对气流的扰动均有影响。驻室空间的大小对气流中心压力分布的均衡作用, 驻室出口面积与排气引射作用及扩压效率都有密切关系。在音速喷管下游设立一个有效的气流加速区域是建立跨音速流场最关键性的问题, 加速段的长度看来相当于一个实验段高度为宜。跨音速流场调整结果表明, 马赫数小于及等于1.0时在模型实验区域610毫米长度内, 轴向马赫数偏量(△M)/M小于±0.47%,马赫数梯度小于0.006/米。马赫数大于1.0及小于1.15时, (△M)/M小于±0.75%, 马赫数梯度小于0.006/米。马赫数大于1.15及小于1.20时, (△M)/M小于±1.57%, 马赫数梯度小于0.016/米。在实验马赫数范围内气流平均偏斜角Aα小于±0.05°, 因此认为已达到可用的跨音速流场标准[6]。风洞气流紊流度估计不大于0.1%, 总压损失不超过0.24大气压力, 最大流量为100公斤/秒, 以实验段高度530毫米为参考长度的风洞雷诺数及~e范围为3.7×10~8至8.2×10~9, 驻点为大气状况。跨音速标准模型实验结果表明, 纵向气动系数测量结果是可靠的, 同美国NACA空对空导弹外形的发射实验和飞机模型的风洞实验数据符合一致。阻力系数C_x偏差量一般在±0.001至0.004之内, 升力系数C_y差量一般在±0.005至0.03之内, 俯仰力矩系数M_z偏差量一般在±0.004至0.015之内, 重复性实验良好, 不重复性值均在±0.002至0.005之内, 因此该风洞可以提供模型实验使用。亚音速洞壁效应已基本上消除, 实验结果毋需进行洞壁干扰修正; 但激波反射效应的消除程度有待进一步研究和实验[7]。今后风洞跨音速性能仍应继续改进,希望马赫数大于1.15以后,流场均匀性(△M)/M不大于1%,最大实验马赫数达到1.22至1.25并消除洞壁激波反射效应,因此应对孔壁和加速段结构形式、扩压效率等给予注意。此外改善风洞天平测量系统的准确性和灵敏性问题随着流场问题的解决就显得更加迫切了。
We expanded a high-subsonic wind tunnel with an experimental Mach range of 0.4 to 0.9 into a transonic wind tunnel, and the experimental Mach number has been expanded to 0.3 to 1.2. The experimental section has been changed from semi-open to transonic Wall form, cross-sectional dimensions 530 × 760 mm, hole wall opening and closing ratio of 23%, expanding the full angle of 0.6 °. The choice of hole wall and groove wall in the transonic experiment section is almost the same from the elimination of the subsonic wall effect, which is obviously better than the groove wall in eliminating the shock reflection effect. However, the expansion angle of the hole wall Should be small. The ratio of the aperture to the wall thickness and the ratio of the aperture to the height of the experimental section affect both the flow characteristics of the wall and the disturbances to the airflow. The size of the resident space has an effect on the pressure distribution in the center of the airflow. The area of the outlet in the booth is closely related to the role of the exhaust gas injection and the efficiency of diffuser. Establishing an effective airflow acceleration region downstream of the sonic nozzle is the most critical issue for establishing a transonic flow field. The length of the acceleration section appears to be equivalent to the height of an experimental section. The transonic flow field adjustment results show that the Mach number of the axis is less than ± 0.47% and the Mach number gradient is less than 0.006 / m within the 610 mm length of the model experimental area when Mach number is less than or equal to 1.0. When the Mach number is greater than 1.0 and less than 1.15, (ΔM) / M is less than ± 0.75% and the Mach number gradient is less than 0.006 / meter. When the Mach number is greater than 1.15 and less than 1.20, (ΔM) / M is less than ± 1.57% and the Mach number gradient is less than 0.016 / meter. In the experimental Mach range, the mean deflection angle Aα is less than ± 0.05 °, so it is considered that the available transonic flow field standard [6] has been reached. The wind flow turbulence is estimated to be no more than 0.1%, the total pressure loss does not exceed 0.24 atm and the maximum flow rate is 100 kg / s. The Reynolds number of the wind tunnel with the experimental height of 530 mm and the ~ e range of 3.7 × 10 ~ 8 to 8.2 × 10 ~ 9, the stagnation point is the atmospheric condition. The experimental results of the transonic standard model show that the longitudinal aerodynamic coefficient measurement is reliable, consistent with the launching experiments of the NACA airfoils and the wind tunnel experimental data of the aircraft model. Resistance coefficient C_x deviation is generally within ± 0.001 to 0.004, the lift coefficient C_y difference is generally within ± 0.005 to 0.03, pitch torque coefficient M_z deviation is generally within ± 0.004 to 0.015, the repeatability test is good, not repeated The values are within ± 0.002 to 0.005, so the wind tunnel can provide model experiments. The subsonic wall effect has been basically eliminated, and the experimental results do not need to modify the wall interference. However, the elimination of the shock reflection effect needs further study and experiments [7]. In the future, the transonic performance of the wind tunnel should be further improved. When the Mach number is greater than 1.15, the flow field uniformity (ΔM) / M is no more than 1%, the maximum experimental Mach number reaches 1.22 to 1.25 and the shock reflection effect of the tunnel wall is eliminated, Therefore, to deal with the hole wall and acceleration section structure, diffuser efficiency pay attention. In addition, improving the accuracy and sensitivity of wind tunnel balance measurement system becomes more urgent with the solution of flow field problems.