二元高超声速进气道动态攻角特性风洞实验

为了研究攻角导致的来流定常/非定常变化对二元高超声速进气道性能的影响,在马赫3.85条件下,对一个设计马赫数为6.0的二元高超声速进气道进行了动态攻角变化的风洞实验。实验中,攻角在0°～8.2°连续变化,动态频率最大达到9.3Hz。实验研究表明:攻角在0°～8.2°连续变化过程中,进气道始终保持着起动状态;攻角增加使进气道出口增压比逐渐上升,流量系数、出口总压恢复系数和出口马赫数逐渐降低;在保持起动状态的攻角连续变化过程中,振荡的二元高超声速进气道其性能变化趋势与准稳态时相似,攻角的动态变化未对进气道性能产生显著影响。     

火箭在地面风作用下的振动载荷识别

将地面风作用下的火箭作为一个非均质变截面梁,研究了非轴对称火箭的任意风向风振载荷识别问题。根据结构风振的特性,引入等效阻力系数和升力系数,利用风载荷作用下火箭的振动加速度或弯矩响应识别等效阻力和升力系数以及归一化随机载荷的功率谱。给出了等效阻力和升力系数与以往风载试验中使用的非定常阻力和升力弯矩系数的关系。     

火星盘缝带伞超声速风洞试验结果分析

为了研究盘缝带(DGB)伞在超声速条件下的阻力特性、摆动角以及伞绳载荷分布的不均匀性,在FD 12风洞中开展了盘缝带伞的阻力特性风洞试验,并通过安装柔性传感器测量了伞绳所受载荷。试验结果表明,在Ma1.50到Ma2.50来流条件下,盘缝带伞的阻力系数随着马赫数的增大先增加后减小,在Ma1.75时达到最大,为0.60,而最大摆动角则是先减小后增大,在Ma2.00时摆动角最小,为7.4°,传感器的安装对降落伞风洞试验结果影响很小。通过柔性传感器的标定和误差分析,获取了伞绳载荷数据,结果显示不同伞绳所受拉力的比值可达到1.98。且传感器数据与风洞天平的测量降落伞总载荷结果吻合,进一步验证了柔性传感器结果的正确性。     

Mesosphere/lower thermosphere wind regime parameters using a newly installed SKiYMET meteor radar at Kazan (56°N, 49°E)

New meteor radar (MR) horizontal wind data obtained during 2015–2018 at Kazan (56°N, 49°E) are presented. The measurements were carried out with a state-of-the-art SKiYMET meteor radar. Monthly mean vertical profiles of zonal and meridional components of the prevailing wind speeds, also amplitudes and phases of the components of diurnal (DT) and semidiurnal tide (SDT) winds are displayed as contour plots for a mean calendar year over the four recent years and compared with distributions of these parameters provided by the previous multiyear (1986–2002) meteor radar (MR) measurements at Kazan and by the recent HWM07 empirical model. The analysis shows that the SKiYMET zonal and meridional prevailing wind speeds are generally in good agreement, sharing the same seasonal features, with the earlier MR seasonal winds. Comparisons with the HWM07 model are not favourable: eastward solstitial cells as modelled are significantly larger, >30?m/s compared to 15–20?m/s. Also, reversal lines are too variable with height, and the positions of modelled cells (positive and negative) are unlike those of either MRs at Kazan or other MLT radars. Both MR systems provide the large SDT amplitudes, approximately 30?m/s and vertical wavelengths, approximately 55?km, for both components at middle latitudes in winter. They also show the well known strong SDT September feature (heights 85–100?km, the vertical wavelength ～55–60?km), and the weak summer SDT for 80–91?km. HWM07 shows unrealistic amplitudes and phases above 90?km by height and month: minimal amplitudes in equinoxes and no September feature.The weak DT of middle to high latitudes provide similar amplitude and phase structures from both MRs, 1986–2002 and 2015–2017: largest amplitudes (10–12 or 8–10?m/s) for the evanescent meridional tide in summer, peaking in late July; weakest (0–2, 2–4?m/s) at 80 to 92–96?km, when the tide is vertically propagating (January, February, November, December) with a vertical wavelength near 40?km. Again, HWM07 differs in amplitude and phase structures: showing peak amplitudes in equinoxes: April, 15?m/s at 88?km; October, 21?m/s at 89?km.Coupling of the MR wind parameters with the ERA5 wind parameters is studied for a case in 2016. It is shown that the prevailing winds and DT amplitudes and phases of both datasets can be simply linked together, but that the ERA5 SDT amplitudes are significantly underestimated at the top model levels of the ERA5 reanalysis project.     

Robust LPV modeling and control of aircraft flying through wind disturbance

This article deals with the disturbance attenuation control of aircraft flying through wind shear via Linear Parameter Varying(LPV) modeling and control method. A Flight Dynamics Model(FDM) with wind shear effects considered was established in wind coordinate system. An LPV FDM was built up based on function substitution whose decomposing function was optimized by Genetic Algorithm(GA). The wind disturbance was explicitly included in the system matrix of LPV FDM. Taking wind disturbance as external uncertainties, robust LPV control method with the LPV FDM was put forward. Based on ride quality and flight safety requirements in wind disturbance, longitudinal and lateral output feedback robust LPV controllers were designed respectively,in which the scheduling flight states in LPV model were actually dependent parameters in LPV control. The results indicate that LPV FDM can reflect the instantaneous dynamics of nonlinear system especially at the boundary of aerodynamic envelope. Furthermore, the LPV FDM also can approach nonlinear FDM's response in wind disturbance special flight. Compared with a parameter-invariant LQR controller designed with a small-disturbance FDM, the LPV controllers show preferable robustness and stability for disturbance attenuation.     

Scaling procedures of cabin noise generated by turbulent boundary layer excitation

This paper presents a new method for measuring the cabin noise of a structure in a wind tunnel. A method for scaling the cabin sound was derived to obtain the cabin noise of a structure, and the derivation of the scaling procedure was based on a theoretical hypothesis regarding the cabin noise prediction for a scaled model in a wind tunnel. A frequency offset was generated because of the error introduced by model manufacture and installation, and a proposed modal test method was used to eliminate the frequency offset. Both a full-scale structure and scaled structure were measured in the wind tunnel tests. The cabin noise of the full-scale model was compared with the results obtained using the scaling procedure based on the scaled model. The comparisons of the measurement results indicate that the scaling procedures developed in this paper are effective for vibro-acoustic predictions in wind tunnels. Moreover, background noise tended to affect the results of the cabin sound for the wind tunnel test, and thus background noise should be prevented through specific design efforts.     

Effect of an end plate on surface pressure distributions of two swept wings

A series of wind tunnel tests was conducted to examine how an end plate affects the pressure distributions of two wings with leading edge(LE) sweep angles of 23° and 40°. All the experiments were carried out at a midchord Reynolds number of 8×10~5, covering an angle of attack(AOA) range from -2° to 14°. Static pressure distribution measurements were acquired over the upper surfaces of the wings along three chordwise rows and one spanwise direction at the wing quarter-chord line. The results of the tests confirm that at a particular AOA, increasing the sweep angle causes a noticeable decrease in the upper-surface suction pressure. Furthermore, as the sweep angle increases, the development of a laminar separation bubble near the LEs of the wings takes place at higher AOAs. On the other hand, spanwise pressure measurements show that increasing the wing sweep angle results in forming a stronger vortex on the quarter-chord line which has lower sensitivity to AOA variation and remains substantially attached to the wing surface for higher AOAs than that can be achieved in the case of a lower sweep angle. In addition, data obtained indicate that installing an end plate further reinforces the spanwise flow over the wing surface, thus affecting the pressure distribution.     

Hypersonic flutter and flutter suppression system of a wind tunnel model

The aeroelastic behavior of a thin flat rudder model was numerically simulated and experimentally investigated in a hypersonic wind tunnel. In particular, a flutter suppression system taking advantage of collision within small gaps was proposed and a novel system for the flutter simulation of the whole nonlinear aeroelastic system including the flutter suppression system was developed. First, the critical flutter dynamic pressure of the rudder without the flutter suppression system was calculated with different methods. Then, the whole nonlinear aeroelastic system, including theflutter suppression system, was simulated to design the gap size. Finally, the flutter suppression system was experimentally validated in a hypersonic wind tunnel operating at Mach number 5. The typical phenomenon of Limit Cycle Oscillation(LCO) was observed, avoiding the structural failure of the model and the consistency between numerical and experimental results was demonstrated.The proposed suppression system can improve the design and reusability of test models of hypersonic flutter experiments.     

高超声速气动加热关联方法的适应性分析

针对高超声速飞行器热环境预测上有着重要应用的关联方法缺乏构造机制和适应性分析的问题，从驻点热流的Kemp-Riddell公式出发，推导了驻点斯坦顿数和雷诺数之间的经典关联关系，从理论上分析了上述关系的适应性。结果显示斯坦顿数和雷诺数关联关系受马赫数的制约，飞行条件和试验条件马赫数变化范围较大的情况下，更适合采用雷诺数和马赫数的组合变量开展关联分析。当试验来流静焓较低的情况下，过低静焓对热环境的关联特性有不利影响，需要提高试验来流马赫数，远离临界马赫数确保地面数据同飞行条件之间的关联性。在国内4座风洞上开展的风洞试验和DSMC/CFD的数值仿真分析工作，校验了上述结论在连续流区域的正确性。     

Validation of numerical prediction of dynamic derivatives: The DLR-F12 and the Transcruiser test cases

The dynamic derivatives are widely used in linear aerodynamic models in order to determine the flying qualities of an aircraft: the ability to predict them reliably, quickly and sufficiently early in the design process is vital in order to avoid late and costly component redesigns. This paper describes experimental and computational research dealing with the determination of dynamic derivatives carried out within the FP6 European project SimSAC. Numerical and experimental results are compared for two aircraft configurations: a generic civil transport aircraft, wing-fuselage-tail configuration called the DLR-F12 and a generic Transonic CRuiser, which is a canard configuration. Static and dynamic wind tunnel tests have been carried out for both configurations and are briefly described within this paper. The data generated for both the DLR-F12 and TCR configurations include force and pressure coefficients obtained during small amplitude pitch, roll and yaw oscillations while the data for the TCR configuration also include large amplitude oscillations, in order to investigate the dynamic effects on nonlinear aerodynamic characteristics. In addition, dynamic derivatives have been determined for both configurations with a large panel of tools, from linear aerodynamic (Vortex Lattice Methods) to CFD. This work confirms that an increase in fidelity level enables the dynamic derivatives to be calculated more accurately. Linear aerodynamics tools are shown to give satisfactory results but are very sensitive to the geometry/mesh input data. Although all the quasi-steady CFD approaches give comparable results (robustness) for steady dynamic derivatives, they do not allow the prediction of unsteady components for the dynamic derivatives (angular derivatives with respect to time): this can be done with either a fully unsteady approach i.e. with a time-marching scheme or with frequency domain solvers, both of which provide comparable results for the DLR-F12 test case. As far as the canard configuration is concerned, strong limitations for the linear aerodynamic tools are observed. A key aspect of this work are the acceleration techniques developed for CFD methods, which allow the computational time to be dramatically reduced while providing comparable results.