There is an obvious aerodynamic interference problem that occurs for a quad tilt rotor in near-ground hovering or in the conversion operating condition. This paper presents an aerodynamic interference test of the quad tilt rotor in a wind tunnel. A 1:35 scale model of the quad tilt rotor is used in this test. To substitute for the ground, a moveable platform is designed in a low-speed open-loop wind tunnel to simulate different flight altitudes of the quad tilt rotor in hovering or forward flight. A rod six-component force balance is used to measure the loads on the aircraft, and the flow field below the airframe is captured using particle image velocimetry. The experimental results show that the ground effect is significant when the hover height above the ground is less than the rotor diameter of the quad tilt rotor aircraft, and the maximum upload of the airframe is approximately 12% of the total vertical thrust with the appearance of obvious fountain flow. During the conversion operating condition, the upload of the airframe is reduced compared with that in the hovering state, which is affected by rotor wake and incoming flow. The aerodynamic interference test results of the quad tilt rotor aircraft have important reference value in power system selection, control system design, and carrying capacity improvement with the advantage of ground effect.

Large-slenderness-ratio (LSR) aircraft exhibit more severe lateral–directional coupling compared with other aircraft, which poses a significant challenge to their flight safety, especially during high-speed maneuvers. Reliable attitude decoupling control is, therefore, essential for LSR aircraft. In this study, a novel control framework that combines incremental nonlinear dynamic inversion (INDI) and extended state observer (ESO) is proposed for supersonic roll maneuver control of LSR aircraft. The ESO is used to estimate the angular acceleration on the basis of an onboard mathematical model. The acceleration estimator based on ESO achieves superior noise reduction compared with the complementary filter (CF) and reduces the onboard model requirement without significantly sacrificing estimation accuracy. Monte Carlo (MC) simulations and frequency–domain analysis demonstrate the effectiveness and robustness of the proposed controller. The sensitivity of parameter uncertainties is also investigated, revealing that the natural frequency of the actuator is the most critical parameter affecting robustness. Finally, flight tests validate the effectiveness of the proposed control structure.

In this study, we investigated the effects of surface sliding discharge of a dielectric barrier discharge (SL-DBD) plasma actuator on the boundary layer flow using the time-resolved particle image velocimetry (PIV) technique. According to the PIV measurements, the peak-shaped mean jet flow is induced by the SL-DBD plasma actuator in quiescent air. Under the freestream of U_∞ = 6 m/s on the flat plate, the results show that the streamwise velocity fluctuation induced by SL-DBD plasma is increased pronouncedly at the downstream of the actuator. Meanwhile, the mean flow velocity increases obviously in the downstream of the actuator. The instantaneous results show the formation of the vorticity at the bottom of the boundary layer after the plasma excitation. In addition, the vorticity has the behavior to lift up, possibly because of the induced peak-shaped jet.

The sliding dielectric barrier discharge (SL-DBD) has attracted attention due to its ability to suppress flow separation. This paper investigated the effect of SL-DBD on the flow field in the flat plate boundary layer by time-resolved particle image velocimetry. We obtained the finite-time Lyapunov exponent field and the Lagrangian coherent structures of the flow field through the velocity field. The results show the effect of SL-DBD has spatial differences, and SL-DBD will produce a “strong–weak–strong” spatial effect on the flow field. The directionality of the flow structure induced by SL-DBD is the main reason for the difference. SL-DBD will induce a large number of vortex structures in the local downstream area of the electrode. The vortex structure enhances the mixing and squeezing effects between the upper and lower flow fields. The upper flow field moves the lower flow field downward by about 0.1 mm through the squeezing effect. In addition, at the downstream region adjacent to the electrode, the oblique momentum injection of the SL-DBD is the dominant effect. At one electrode distance from the electrode, the SL-DBD induces a large number of vortex structures. However, when the distance from the electrode is twice the electrode spacing, the number of vortex structures decreases, and the structure becomes larger. The results show spatial differences in the perturbation of the flow field by SL-DBD, which cannot be ignored when SL-DBD is used to suppress flow separation.

In this paper, an integrated design method for guidance and control laws for a short takeoff and vertical landing (STOVL) unmanned aerial vehicle (UAV) during decelerating transition and vertical landing phases under thrust vector control is studied. Based on the principle of time-scale separation, the outer-loop guidance law adopts an implicit dynamic inverse method to provide attainable overload vector control instructions for decelerating transition and vertical landing phases. Meanwhile, the inner-loop control law adopts an improved eigen-structure assignment method, which tracks guidance instructions and maintains a stable attitude. The attitude nozzles are joined to the attitude control with the dropping of dynamic pressure, thereby assisting aerodynamic surfaces to stabilize attitudes. Subsequently, the feasibility and effect of the control methods for a STOVL UAV during decelerating transition and vertical landing phases are verified by a 6-DOF flight simulation platform. The simulation results demonstrate that the integrated design method is beneficial for a quick evaluation of the overall scheme.

Aiming at the interception problem of hypersonic target end phase, the flight interception simulation based on semi-active radar guidance is firstly established, and a warhead based on fragment damage is designed. Secondly, research proposes an effective method of target damage and verifies the interception effect by simulation. AUTODYN was used to establish the 3D finite element model based on the nonlinear flight dynamics model, to simulate the damage process and fragmentation dispersion. In this paper, the probability of success of missile interception is estimated by using conditional probability and the distribution of fragment field combined with the miss distance in the interception process and the fragment explosive dispersion law of the steel ball, so as to ensure the destruction of the terminal hypersonic incoming weapon within the safe range.