Wide-speed-range aerodynamic design remains one of the critical bottlenecks in the development of horizontal takeoff and landing aerospace planes. These vehicles operate across an exceptionally broad flight envelope, encountering significant variations in both dynamic pressure and atmospheric conditions. The disparity in required lift across low-speed and high-speed regimes leads to conflicting demands on the lifting surface sizing, introducing considerable challenges in achieving an aerodynamically balanced configuration throughout the entire mission profile. Specifically, satisfying the lift requirements for low-speed takeoff typically results in excessive lifting surface area under high-speed conditions, causing a marked deviation from the optimal lift-to-drag ratio and limiting the vehicle’s aerodynamic efficiency. To address this challenge, we first analyze the lift matching requirements associated with wide-speed-range operations of horizontal takeoff aerospace planes. A global aerodynamic optimization framework is then developed, explicitly incorporating lift matching constraints into the configuration design process. The proposed methodology is applied to the wing planform and airfoil shape optimization of the Sanger aerospace plane carrier aircraft. Under the constraint of meeting low-speed takeoff lift requirements, the optimized wing achieves an improvement of 20.00% in the available lift-to-drag ratio under supersonic conditions and 8.12% under hypersonic conditions, effectively mitigating the aerodynamic efficiency degradation at high speeds. Finally, the optimization framework is extended to account for fuselage-wing aerodynamic interference effects, and its applicability to full-vehicle configuration aerodynamic optimization is demonstrated.

For the numerical simulation of near-/mid-field sonic boom propagation of a supersonic civil aircraft, the conventional Computational Fluid Dynamics (CFD) using a time-marching method is computationally intensive, while the acoustic propagation method based on solving the augmented Burgers equation has difficulty in simulating velocity and density variations of three-dimensional flows. To address these problems, a high-fidelity numerical simulation method of using a space-marching approach is proposed, and an in-house code “SMFlow3D” is developed. The simulation procedure consists of three steps. First, a structured Mach-aligned grid of a conical shape is generated. Second, circumferential distributions of density, velocity, and pressure variables are extracted at one-body length distance from a supersonic aircraft, and are interpolated to the grid nodes as initial conditions. Third, within a finite difference method framework, three-dimensional steady governing equations in a curvilinear coordinate system are solved along the post-Mach cone direction using a third-order Runge-Kutta method. The developed method was applied to simulate the near-/mid-field sonic boom propagation for the JAXA Wing Body (JWB) model, which was proposed in the 2nd Sonic Boom Prediction Workshop. The predicted sonic-boom waveforms are compared with the results obtained from conventional CFD method and acoustic propagation method based on the augmented Burgers equation, validating the accuracy and effectiveness of the developed method. Furthermore, near-field simulation for a Delta Wing Body model demonstrates that, compared to conventional CFD method, the developed method reduces computational time by approximately 97.3%, while improving the precision of shock capturing. The developed method can provide support for rapid prediction of sonic boom and analysis of shock system evolution.

To meet the challenge of drag reduction for next-generation supersonic transport aircraft, increasing attention has been focused on Natural Laminar Flow (NLF) technology. However, the highly swept wings and high-Reynolds-number conditions of such aircraft dramatically amplify Crossflow (CF) instabilities inside boundary layers, making it difficult to maintain a large laminar flow region. To explore novel NLF designs on supersonic wings, this article investigates the mechanisms underlying the attenuation of Tollmien–Schlichting (TS) and CF instabilities by modifying pressure distributions. The evolution of TS and CF instabilities are evaluated under typical pressure distributions with different leading-edge flow acceleration region lengths, pressure coefficient slopes and pressure coefficient deviations. The results show that shortening the leading-edge flow acceleration region and using a flat pressure distribution are favorable for suppressing CF instabilities, and keeping a balance of disturbance growth between positive and negative wave angles is favorable for attenuating TS instabilities. Based on the uncovered mechanisms, a strategy of supersonic NLF design is proposed. Examination of the proposed strategy at a 60° sweep angle and Ma = 2 presents potential to exceed the conventional NLF limit and achieve a transition Reynolds number of 17.6 million, which can provide guidance for NLF design on supersonic highly swept wings.

The design of low sonic boom high efficiency aerodynamic layout is one of the most important key technologies for supersonic civil aircraft. Based on the inverse design method of sonic boom minimization theory, this paper proposes an advanced supersonic civil aircraft low sonic boom aerodynamic layout using the aft-body design based on wave system beneficial interference and the mixed credibility anti-design method based on the parameterized near-field overpressure signal, and analyzes the effect of each step of sonic boom reduction and the whole sonic boom blanket. Then, the effects of flying altitude, Mach number and wing swept angle on the sonic boom loudness of the aerodynamic layout are studied. Finally, the effects of flying altitude and Mach number on the aerodynamic characteristics of the layout are examined using CFD numerical simulation method. The results show that the anti-design method based on the sonic boom minimization theory can reduce the ground perceived loudness in decibel by 6.54 PLdB, while the aft-body design method based on the wave system beneficial interference and the hybrid reliability inverse design method of parameterized near-field overpressure signal can further reduce it by 0.97 PLdB and 4.04 PLdB, respectively. Reasonable match of the flying height and Mach number can effectively reduce the ground perceived loudness in decibel. Reasonable selection of the flying height and cruise Mach number can effectively improve cruise efficiency. This study has certain engineering value for both aerodynamic layout design and aircraft conceptual design of supersonic civil aircraft.

Accurate prediction of sonic boom is one of key challenges for the design of a low-boom supersonic aircraft. For most of available far-field prediction methods, the effect of atmospheric turbulence appearing in the planetary boundary layer cannot be considered, which results in remarkable inaccuracy of predicting ground-level sonic boom waveform. Although some efforts have been made to overcome the shortcoming, the turbulence effects are not yet well described so far. This article proposes an improved method by extending the two-dimensional Heterogeneous One-Way Approximation for the Resolution of Diffraction (HOWARD) equation to account for the axial and transverse convections of wind fluctuation as well as the effect of temperature fluctuation. The proposed method is validated by comparing the predictions with the flight-test data of JAXA D-SEND#1 LBM, which shows that the result of the proposed method is in better agreement with the flight-test data than that of the method without considering atmospheric turbulence effects. Then, distortion mechanism of sonic boom waveforms caused by atmospheric turbulence is analyzed by using the proposed method. It is indicated that the effect of turbulent convection makes uniform sonic-boom wavefronts irregular, which creates the condition of diffraction effect to perturb waveforms. Finally, the proposed method is applied to investigate the behavior of two types of waveforms given by the sonic boom minimization theory. Results show that a far-field waveform with a weaker initial shock is more beneficial for low-boom design of a supersonic aircraft.

Accurate prediction of tip vortices is crucial for predicting the hovering performance of a helicopter rotor. A new high-order scheme (we call it WENO-K) proposed by our research group is employed to minimize numerical dissipation and extended to numerical simulation of unsteady compressible viscous flows dominated by tip vortices over hovering rotors. WENO-K is referred to as an adaptively optimized WENO scheme with Gauss-Kriging reconstruction, and its advantage is to reduce dissipation in smooth regions of flow while preserving high-resolution around discontinuities. Here WENO-K scheme is adopted to reconstruct left and right state values within the Roe Riemann solver updating the inviscid fluxes on a structured dynamic overset grid. To minimize the accuracy loss for high-order reconstruction on artificial boundaries of overset grid, a method of multilayer fringes is proposed to carry out interpolation between background grid and blade grid. Massively parallel computing considering automatic load balance on averagely partitioned overset grid is developed to reduce the wall-clock time of an unsteady simulation. Numerical results for Caradonna-Tung (C-T) rotor in hover at the conditions of subsonic and transonic tip Mach numbers show that the thrust coefficient error for the result of WENO-K scheme is no more than 3%. Compared with WENO-JS scheme, WENO-K scheme achieves about 40% improvement on accuracy of predicting rotor thrust with only 4.1% extra computational cost. More importantly, WENO-K scheme can capture more sophisticated unsteady flow structures and resolve tip vortices to a larger wake age with an increment of about 270° compared to WENO-JS scheme.