Search

Due to its layout, there are difficulties in realizing heading attitude control of a flying-wing unmanned aerial vehicle. In this paper, a reverse jet control scheme has been designed: (1) to replace the resistance rudders that are used for the yaw control of a conventional flying-wing unmanned aerial vehicle, (2) to assist and optimize heading attitude control, eliminate the adverse effects of the control surface and enhance stealth performance, and (3) to promote the use of rudderless flight for flying-wing unmanned aerial vehicles. To explore the control mechanism and the flow field of the reverse jet scheme, three-dimensional numerical simulations and low-speed wind tunnel experiments were carried out. First, the numerical simulations evaluated the feasibility and effectiveness of the reverse jet control scheme and explored and optimized the excitation parameters for the scheme. Then the forces were measured in a wind tunnel, and particle image velocimetry experiments were carried out. A reverse jet actuator was independently designed to verify the results of the numerical simulation. The results show that when the reverse jet excitation is applied, the jet obstructs the mainstream, destroys the flow field at the excitation position, and causes early separation of the flow, which increases the pressure drag on the wings and produces a control effect. The control effect mainly depends on the separation degree of the leeward surface. The larger the jet momentum coefficient is, the smaller the jet angle is, and the closer the excitation position is to the leading edge, the greater the separation degree of the leeward surface is, the better the heading attitude control effect is.

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.

To explore the control efficiency of circulation flow control technology, a circulation control actuator with an independent gas source has been designed and applied in roll attitude control of a small unmanned aerial vehicle. The circulation control devices are arranged at the two ends of the wing on an unmanned aerial vehicle scale model, the changes in aerodynamic force and aerodynamic moment caused by turning on the actuator are measured in a wind tunnel, and the flow field characteristics are analysed using particle image velocimetry technology. The flight control effect of the roll attitude is verified via a flight test. Experimental and flight test results show that the control of roll attitude can be achieved by turning on the circulation control actuator on one side, and the maximum efficiency that the circulation control generates is equivalent to 8° aileron deflection with production of a favorable yaw moment to achieve a coordinated turn. The circulation control actuator can increase lift and reduce drag when opened on both sides simultaneously. The maximum lift-to-drag ratio of the UAV increased from 5 to 9, and this approach can also suppress flow separation and delay stall at high angles of attack. The aileron or trailing edge flaps can be replaced with circulation control actuators, and the circulation control technology can also be applied to aerodynamic performance improvement and flight control in other types of aircraft.

To study wing-body rock, various tests were used for the canard-configuration models in the wind tunnel, including free rolling, disturbance on free rolling, static moment measurements, dynamic derivative measurements, smoke wire method and particle image velocimetry. The models also have a chine forebody, strake wings, main wings and a vertical tail. In the tests, the canard deflection angles are 0° and 20° relative to the axis of the body. First, the roll angle time histories were obtained by free rolling motion and disturbed motion; the results demonstrate that both canard-configuration models have nonzero mean roll angles. In particular, the existence of the other larger nonzero mean roll angles on the same side is a curious aspect at the critical pitch angle, such as at θ = 37° for the model with canard deflection of 20°. Next, static moment tests and dynamic derivative tests demonstrate that although the roll angle time histories are non-limit cycle motions, they are irregularly self-excited oscillation under the influence of multi-vortex structures. Finally, the constituent parts of structures over wing bodies and the flow mechanism of abruptly jumping from an equilibrium roll angle to the other on the same roll side were revealed by the smoke wire and particle image velocimetry measurements.

The surface dielectric barrier discharge plasma actuator driven by nanosecond pulses is recognized as an effective fluid actuator for flow separation control. The operation condition of nanosecond dielectric barrier discharge actuators for separated flow control still requires further study, particularly prioritizing the improvement of the effectiveness and reducing energy consumption in flow separation control implementation. In this study, experiments are conducted using a two-dimensional NASA SC(2)-0712 airfoil in a wind tunnel with a Reynolds number of 0.5 × 106 (25 m/s). The pressure measurement experiments show that the location of actuators affects the efficiency of separation control. Particle image velocimetry results indicate that the most efficient location of the actuator is upstream of the separation point and near the original point of the separated shear layer. Meanwhile, the particle image velocimetry results show the vorticity attaches to the airfoil wall after discharge, which suggests that the reattachment is due to the generation of large-scale vortices. These present structures result in the mixing of the shear layer with the main flow thereby delaying separation and reattaching a separated flow. This study shows the most efficient location related to the separation point. Furthermore, it indicates the reattachment of flow is attributed to the motion of vortexes coherent structure.

A cruciform-finned slender body can excite a limit-cycle rolling oscillation at high angles of attack. To suppress the unwanted motions, flow control approaches should be employed if the aerodynamic control surfaces lose control efficiency at high angles of attack. As a promising technology, the ns-dielectric barrier discharge plasma actuator has been successfully used in high-speed and high-Reynolds-number flow control applications. The present work employs a π-type ns-dielectric barrier discharge plasma actuator and vortex generators to suppress self-excited rolling oscillation at α = 50°. The free-to-roll tests show that the plasma actuator ignited at F+ ∼ 1.5 and that the vortex generators can suppress rolling oscillation. The flow patterns from particle image velocimetry measurement at different cross-sections and rolling angles suggest that the vorticity decrease of the leeward vortices may be the control mechanism for the plasma actuator. For the vortex generators, evident modification of the flow field structure can be observed due to the vortices generated from the vortex generators, which decreases the rolling moment induced by the asymmetry vortices to suppress the self-excited rolling oscillation.

Next-generation aircraft and missile are required to have extreme maneuverability, which should maintain stability at high angle of attack. However, the unsteady flow field surrounding the air vehicle would affect the aerodynamics loads and induce unwanted nonlinear motions, of which the rolling motions are usually generated. In order to investigate the unsteady rolling characteristics of a cruciform-finned slender body, the free-to-roll and force measurement tests including particle image velocimetry measurement have been conducted comprehensively. Different types of rolling motion, trimming at equilibrium positions, self-excited rolling oscillation, and self-excited spinning are observed during the free-to-roll experiment depending on the different angles of attack. The force measurement results show that the rolling motions are related to both the static and dynamic stabilities of the rolling moment at balance points. The stabilities of the rolling moment would change as the angle of attack for the model increases. At last, the flow field results from particle image velocimetry measurement indicate that the unsteady rolling motions may be induced by the interaction between the asymmetry vortices and strake wings, fins.

A new dynamic test mechanism has been developed in the unsteady low-speed wind tunnel at Nanjing University of Aeronautics and Astronautics to enhance the abilities of experiments in the unsteady flow cases. The structure of the new test mechanism is described in detail in this paper and the equations used to control the performance of test model’s motions are also listed here. Then a wing-body model was tested in the new test system for the first time, including the forced oscillation tests and free-to-roll test at high angle of attack region. The test results are also discussed.