The flying-wing aircraft has excellent aerodynamic efficiency and stealth performance. However, due to the lack of tails, the flying-wing aircraft has a serious attitude control problem. In this paper, the effective flow control strategy of three-axis control is proposed by using continuous jets for a flapless flying-wing aircraft. The wind tunnel test of two kinds of flying-wing models, namely one flow control model and one mechanical control model, is conducted, and the control effect is analyzed and compared. By simultaneous blowing of the circulation control actuators inboard and differential blowing of the circulation control actuators outboard, the pitch and roll controls are achieved, respectively. It also has an effective control effect at very large angles of attack where the conventional control surface fails. A linear relationship is found between the increment of the controlled aerodynamic force/moment coefficient and the momentum coefficient for circulation control actuators. Moreover, to resolve the difficulty in yaw control, a novel wingtip jet is proposed based on the concept of the all-moving tip and compared with apex jet and circulation control jet. It is found that the wingtip jet is the most efficient actuator, followed by the simultaneous-blowing circulation control jet. Therefore, based on the research above, two optimized fluidic control configurations are proposed. One employs circulation control jet and wingtip jet, and the other is completely dependent on circulation control jet. Finally, the flow control mechanism of circulation control is discussed. Circulation control significantly accelerates the flow on the upper surface of the airfoil in attached flow and reduces the flow separation region in separated flow, leading to aerodynamic performance improvement. These results provide an important theoretic basis for the flapless flight control of flying-wing aircraft.

Flying wing configuration aircraft pursues excellent aerodynamic and stealth performance. However, traditional mechanical control surfaces can compromise its stealth profile and may lead to a decrease in control effectiveness at moderate angles of attack, potentially causing aircraft instability and loss of control. Active Flow Control (AFC) techniques effectively address the aforementioned drawbacks. In this study, a wind tunnel virtual flight test system, integrated with closed-loop active flow control, is constructed. The system is capable of simulating active flight attitude control of controlled model under both steady and unsteady incoming flow conditions. Utilizing this system, wind tunnel virtual flight tests of the flying wing configuration model with AFC are conducted, obtaining its three-axis attitude control characteristics under steady flow conditions. Results demonstrate that consistent and controlled pitch and roll moments can be generated through trailing edge circulation control, while required yaw moment can be generated through wingtip reverse jets, achieving stable three-axis attitude control of the flying wing aircraft. Particularly, when controlling longitudinal attitude of the model with pitch circulation control, the generated pitch moment is linearly correlated with jet momentum coefficient. Furthermore, a closed-loop control strategy for gust load alleviation based on model longitudinal attitude feedback is proposed. The stability enhancement ability of AFC for flying wing configuration model under gust disturbance is validated. Additionally, it is further discovered that the effectiveness of the stability enhancement control is jointly determined by the intensity of the jet applied for control and the phase relationship between the control signal and the gust disturbance.

Thrust vectoring technology plays an important role in improving the maneuverability of aircraft. In order to overcome the disadvantages of mechanical thrust vectoring nozzles, such as complications of structure and significant increases in weight and cost, fluidic thrust vectoring nozzles are proposed. Dual Throat fluidic thrust vectoring Nozzle (DTN) has received wide attention due to its excellent thrust vectoring efficiency and minimal thrust loss. In this study, three-dimensional unsteady numerical simulations of a single axisymmetric DTN are conducted first to analyze its dynamic response. Then the pitch and yaw control characteristics of DTN equipped on a flying-wing aircraft are investigated. It is found that the dynamic response will experience three stages: rapid-deflecting stage, oscillating stage, and steady stage. A complete recirculation zone forms at the end of the rapid-deflecting stage, which pushes the primary flow to attach to the wall opposite the secondary injection. Meanwhile, the exhaust flow is deflected. In terms of DTN’s application, the DTN equipped on the flying-wing aircraft is capable of providing effective pitch and yaw moments at all angles of attack and Mach numbers. In addition, continuous pitch and yaw moments can be obtained by adjusting the secondary mass flow ratios. The control moment is generated due to the asymmetrical pressure distribution of nozzle surface, which is mainly contributed by the pressure decrease on the secondary injection surface. Moreover, the DTN equipped on the flying-wing aircraft has a relatively high thrust vectoring efficiency of around 5°/% and a thrust coefficient of around 0.95 when nozzle pressure ratio equals 4. These results provide an important theoretical basis for the practical application of DTN.

The spinning missiles usually adopt oblique tails to generate the rolling moment. However, such a configuration suffers from the problem of poor stability of rotating speed, which affects the hitting accuracy. To improve the aerodynamic performance of spinning missiles, the traditional oblique tails are substituted with Gurney flaps on the trailing edge of the tails to generate rolling moment in the present study. Effect of the Gurney flap height on the aerodynamic characteristics of the missile is studied by performing numerical simulations, and also compared with the oblique tail model. The results show that the Gurney flaps can generate the rolling moment for the missile, which increases with the Gurney flap height. Compared to the oblique tail model, the Gurney flaps can provide a larger rolling moment for the missile with a relatively small impact on side force. The maneuvering efficiency of the Gurney flaps is slightly different at different Mach numbers. The rolling moment decreases with the increase of the angle of attack in the subsonic regime, while it remains nearly unchanged with the angle of attack variation in the supersonic regime. Meanwhile, the maneuvering efficiency of the Gurney flaps is lower in the supersonic regime compared to that in the subsonic regime. It is found that, the Gurney flaps induce asymmetric flow near the trailing edge of the missile tail, resulting in the occurrence of asymmetric pressure difference among the tails, thus leading to the rolling moment generation.