The airfoil is the foundation of an aircraft, and its characteristics have a significant impact on those of the aircraft. Conventional airfoil design mainly focuses on improving aerodynamic performance, while flying wing airfoil designs should also consider layout stability and stealth performance. The design requirements for an airfoil vary with its position on the flying wing layout aircraft based on corresponding spanwise flow field characteristics. By analyzing the spanwise flow characteristics of the flying wing, partition design models for flying wing airfoils were established in this study, and a series of flying wing airfoil designs that consider aerodynamics and aerodynamic/stealth were implemented. Then, the designed airfoils were configured on a three-dimensional X-47B layout for testing and verification. The results showed that the aerodynamic design and the aerodynamic/stealth design exhibited significant improvements in terms for aerodynamic and longitudinal trimming characteristics. However, the cruise drag performance of the aerodynamic/stealth design was slightly worse than that of the aerodynamic design, although the longitudinal moment trimming characteristics were basically the same. The stealth characteristics of the aerodynamic/stealth design had significant advantages, indicating that there were weak contradictions between the aerodynamic, stealth, and trimming requirements in the design of the flying wing.

Fluid thrust vectoring (FTV) control has obvious advantages in structural quality and stealth performance because of its fast response and light weight. However, improving FTV vector performance will cause a loss in engine performance due to the need to draw airflow from the engine. In order to alleviate the above problems and further improve the vector performance of FTV, a nozzle combined with throat skewing and shock vector control is proposed, and the secondary flow of the nozzle comes from the throat and is injected into the nozzle divergence section. The numerical results indicate that compared with the original configuration, the vector angle and vector efficiency of the new configuration are more linear with the nozzle pressure ratio (NPR), and the vector angle and vector efficiency are improved by 163% and 218%, respectively, while experiencing a maximum reduction in the thrust coefficient of 1.4%. Compared with the only bypass-type shock vector nozzle, the new configuration utilizes the diversion of the two jets to eliminate the reattachment of the separation bubble after the jet and its resulting abrupt change in vector performance, improving the performance while having good control characteristics. Additionally, a sensitivity analysis of the spacing between two jets is also carried out. The spacing between two jets should be increased to make the flow pass through two weaker shock waves to improve the vector performance while ensuring that the separation after the jet is no longer attached.

Diverterless supersonic inlets (DSI) is the predominant system for supersonic aircraft below Mach 2, offering superior boundary layer displacement and stability margins. However, conventional discrete adjoint optimization frameworks relying on the constant eddy viscosity (CEV) assumption introduce significant gradient errors in turbulent flow sensitivity analysis, limiting performance refinement of DSI configurations characterized by shock-boundary layer interactions and separation phenomena. This study addresses this challenge through a novel discrete adjoint framework integrating upwind schemes and full turbulence adjoint modeling, enabling accurate gradient computation while maintaining numerical stability for high-speed flows. The method is applied to optimize an integrated DSI system encompassing compression bump, diffuser duct, and cowl lip geometries. Gradient validation against finite-difference results demonstrates improved convergence efficiency and sensitivity accuracy compared to CEV-based approaches. At Mach 1.8, the optimized DSI achieves shock-on-lip (SOL) conditions with 4.0% total pressure recovery improvement (0.914) and 64.2% reduction in distortion index (0.273). Off-design performance at Mach 0.6–0.9 maintains exceptional recovery coefficients (0.982/0.973), demonstrating multi-point operational robustness. The design simultaneously minimizes internal flow separation, enhances boundary layer control, and balances supersonic pressure recovery with sub/transonic efficiency. This work advances adjoint-based aerodynamic optimization by resolving CEV limitations in capturing viscous-inviscid interactions, establishing a pathway for next-generation DSI development.

The design of the flying wing airfoil must consider aerodynamic stealth and trim requirements, with their coupling exacerbating the complexity of the design problem by introducing multiple local minimum points in design space. In addition, it would need a broad space with high dimensionality to obtain the ideal result, and expansion of design space could lead to more local minimum points, causing significant challenges to traditional optimization design methods. A Two-Stage Global–Local Constrained Optimization Method (TGLCOM) was proposed to address these issues. The parametric space was divided into a large-scale global space and a high-dimensional local space. A surrogate-based global constrained optimization method was applied in the large-scale global space, followed by a gradient-based algorithm in the high-dimensional local space to refine the design and obtain the global optima. The efficiency and robustness of the proposed TGLCOM were verified through the airfoil and flying wing layout aero/stealth design. The results indicated a minor conflict between the RCS drag and pitch moment performance. Moreover, the stealth design of the airfoil improved the stealth performance of the flying wing layout in both the yaw and pitch directions.

As a practical rounded trailing edge airfoil for coaxial rotors, DBLN-526 is a fore and aft symmetrical airfoil with two steps on its lower side. This airfoil has been used at the inboard section of the coaxial rotor system. As there are always two eddies behind the airfoil because of its rounded trailing edge, the interaction between the separation and transition should be considered. Thus, the unsteady Reynolds-averaged Navier-Stokes- (RANS-) based model was used to analyze the unsteady transition and separation features. Three different rounded trailing-edge airfoils were compared with DBLN-526. Power spectrum density analysis and calculations demonstrated that the lift coefficient fluctuation of DBLN-526 was smaller than that of other rounded airfoils with different angles. Further investigation indicated that the locations of transition for DBLN-526 can be fixed at a wide range of angles by the unique design on its lower side. Because of this settled transition location, the size of separation is decreased, and the position of separation is settled as well, which leads to a lower lift coefficient fluctuation. The turbulent kinetic energy after the transition was higher, which injected a lot of energy into the boundary layer, and the separation zone near the transition position was relatively smaller. This study provides an indication for controlling separation and reducing unsteady fluctuations for rounded trailing edge airfoils.