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The existing MDOE methods based on the parametric model cannot meet the tunnel test requirements of advanced vehicle because of strong nonlinear aerodynamic behaviors. It is necessary to develop the MDOE method based on the non-parametric model instead. The formal design of experiments method of MDOE for wind tunnel tests is researched in this paper. The applicability of Latin Hypercube Sampling and Uniform Design in wind tunnel tests is compared through virtual experiments. And several optimizing methods based on Uniform Design are developed, which can improve the quality of design results to meet the requirement of wind tunnel tests. Studies show that, Uniform Design is more robust and uniform than Latin Hypercube Sampling for the MDOE method based on the non-parametric model, the optimization including boundary complement, density adjustment of sampling points and repetition points design can utilize the priori information to the design, the formal design of experiments method developed in the paper requires less sampling points than the OFAT, which are only 66.7% in the demonstration, and samples the intensive non-linear changing law adequately and exactly. The studies on the formal design of experiments method of MDOE in wind tunnel tests laid the foundation for further study of the MDOE method based on the non-parametric model. [ABSTRACT FROM AUTHOR]

longitudinal and lateral aerodynamic characteristics of the low aspect ratio flying wing calibration model are investigated in a low speed wind tunnel. Normal force measuring experiment is conducted to gain the longitudinal aerodynamic characteristics and yaw control characteristics, and the PIV test is also conducted to investigate the flow mechanism of the low aspect ratio flying wing. The results indicate that the leading-edge separation vortex appears on the wing's spine surface when the attack angle is at 6 degree. The vortex intensity increases and the vortex core shifts to the symmetric plane of flying wing with the increase of attack angle. Increasing the attack angle further, the vortex core becomes unsteady and begins to oscillate, finally break entirely. The broken position shifts from the ending edge to the leading edge. Yaw control characteristics of low aspect ratio flying wing is also studied in this paper. The results indicate that the flying wing is static stabile at the test attack angle. When the attack angle is less than 6 degree, it is controllable in yaw direction. And when attack angle is more than 6 degree, the yaw control ability decreases because the control surface may lays in the wake region of broken leading-edge vortex. [ABSTRACT FROM AUTHOR]

In conventional force measure wind tunnel test, the balance can get the aerodynamic force of the aircraft and also the directly moment. It is while the pressure center is usually considered as the intersection point between aerocraft's longitudinal axis and its aerodynamic force, which is not the virtually real value of the pressure center of the aircraft. How to obtain the pressure center of re-entry aerocraft accurately with the general method in wind tunnel test remains a difficult problem. In this paper, the position of re-entry vehicle's pressure center was obtained through the technique of attack angle continuous change and real-time data acquisition. Also the pressure center of re-entry module change rule with the change of attack angle was obtained. Experiments demonstrated the feasibility of the proposed method. With the increasing of negatives attack angle, the contribution of axial force to pitch moment becomes great, while the influence of normal force is just the contrary, which results in thickcreased gap between the conventional measured value and the real value of vehicle's center of pressure. The vertical position of re-entry aircraft's load center is relatively small in the test. The final CA-Cm curve shows that the axial force should not be neglected or simplified in wind tunnel test. [ABSTRACT FROM AUTHOR]

Changing the model attitude with the coupled angle of attack and roll is an important mode for the lateral test. Although the lateral flexible angle is offered in this kind of test, the data of lateral flexible angle is often neglected, which leads to the inaccuracy of the relation between the aerodynamic loads and the model attitude. Lateral flexible angle is generated from the lateral force and the yawing moment, and the effect of the lateral flexible angle will be notable when the values of the lateral force and the yawing moment are large in the test, for example, the yaw steerage test. The effect of the lateral flexible angle is analyzed, offers a data processing method of the lateral flexible angle. Calculative and experimental results show that the method can increase the accuracy of the application data. [ABSTRACT FROM AUTHOR]

Both spike and forward-facing jet can reduce the shock wave drag acted on a blunt head hypersonic vehicle significantly. However, they are not be adopted widely because of some shortages. The combination of spike and forward-facing jet takes advantages of both of them, while more complex aerodynamical phenomena and parameter optimization issue are involved. In this paper, the parametric optimization of combination of spike and forward-facing jet is carried out by response surface methodology. The objective of the optimization is to minimize the drag and the jet flux. The optimized parameters are the length of the spike, total pressure of the jet and the diameter of the jet out let. The response surface model is involved to feedback the relationship between the response and the designing parameters. IV-optimal is used to design the sample points, of which the response values are achieved through numerical simulation. The desirability value is used to estimate the benefits. The principal conclusions are as following: The function of drag and design parameters is supposed to be 2nd or 3rd order non-linear polynomial, and the coupling effects among design parameters are evident, especially between the diameter of jet outlet and the total pressure of jet. Response surface model speculated from sample points simulations reveals the relationship between the drag and the design parameters, from which the drag prediction interval of arbitrary parameters combination in design space can be obtained. The optimal parameters of minimum drag and jet mass flow are acquired, where the drag is predicted in a good precision. The study shows that response surface methodology in parameters optimization has high reliability and practicality with less computation requirement, and has good prospect of application in aircraft design. [ABSTRACT FROM AUTHOR]