A large-scale low-speed wind tunnel turbofan aircraft dynamic simulation test system is established to accurately obtain the dynamic influence of turbofan engine intake and exhaust on the aircraft, providing technical support for the refined development of military and civil turbofan aircraft in China. The technology of internal air bridge balance is one of the core technologies in the test system. The turbine powered simulator (TPS) needs to be driven by high-pressure gas. The stiffness of the high-pressure gas supply pipeline and the high-pressure gas flowing through the fixed end to the measuring end of the balance will bring great measurement interference to the balance, and then affect the force measurement accuracy of wind tunnel test. It is necessary to weaken the rigidity of the high-pressure air supply pipeline and eliminate the interference of air supply. Eliminating or weakening the interference load caused by the air supply pipeline and high-pressure air supply flowing through the pipeline on the balance force measurement is a necessary prerequisite to ensure that high-precision test data can be obtained. In this study, a new type of air bridge disturbance elimination unit structure is proposed. The technical research on the design mechanism and calibration correction of the influence of air bridge stiffness,air supply pressure, air supply temperature and air supply flow on the force measurement interference of the balance is carried out. The final test results show that the corresponding technical indicators of the internal air bridge balance meet the expectations and can meet the needs of turbofan aircraft dynamic simulation wind tunnel test.

Inspired by the bionic principle of bird covert feathers on the upper wings, this paper designs four types of flexible sawteeth with different thicknesses, and the sawteeth are installed at different positions on the upper-surface of the NACA 0018 two-dimensional airfoil at an angle-of-attack of 15°, and the velocity profiles in the wake flow is measured by the hot-wire anemometer. The Reynolds number based on chord length of 300 mm is 5.1×10⁵, and the mean velocity profiles and root-mean-square velocities show that the thinner flexible sawteeth can eliminate the leading-edge shear layer when installed near the leading-edge of the airfoil, and the power spectral density results show that the turbulent fluctuations are suppressed. On the contrary, the thicker flexible sawteeth are able to control the leading-edge flow separation when installed near the trailing-edge of the airfoil, and the high-speed fluid is induced to adhere to the upper-surface of the airfoil.

Upper atmosphere vehicles, operating at altitudes of 100~300 km, have demonstrated significant application potential in fields such as remote sensing reconnaissance, emergency communications, and space exploration, owing to their high ground observation accuracy and low communication latency. These advantages have positioned the upper atmosphere vehicle as a frontier equipment in both commercial spaceflight and national defense security, while simultaneously triggering global competition for upper atmosphere orbit resources. This paper discusses the core value and far-reaching significance of developing upper atmosphere vehicles, and reviews the domestic and international research trends and development status in this field. Faced with the complex and dynamically changing space environment at upper atmosphere orbit altitudes, the paper adopts an overall design perspective to identify and present four challenges faced by upper atmosphere vehicles:force system equilibrium，energy balance,thermal/erosion protection, and system coordination. It also highlights potential solutions and directions for technological breakthroughs. The paper aims to outline research approaches, accelerate the construction of a comprehensive technical system, and facilitate the transition of upper atmosphere orbit space from scientific exploration to engineering applications.

This paper provides technical support for the engineering application of hybir laminar flow control (HLFC)technology in drag reduction for general aircraft. A numerical investigation of HLFC technology for drag reduction on the vertical tail of the AG300 business aircraft is presented. The simulations are based on the Reynolds-Averaged Navier-Stokes (RANS) equations, employing the k-ω SST turbulence model coupled with the γ-transition model. The transition prediction method was validated using experimental data of the infinite swept NLF(2)-0415 airfoil, demonstrating high accuracy and reliability. This paper evaluates the drag reduction effects of uniform suction, non-uniform suction distributions, and reduced leading-edge sweep angle under suction conditions. The results indicate that HLFC can significantly reduce vertical tail drag, with further drag reduction achieved by decreasing the leading-edge sweep angle.However, to maintain flow stability, the suction intensity must be kept below a critical threshold.

A high-performance diffuser section was designed based on polynomial function centerline and area ratio flow channel modeling method. The lip smoothing method based on the combination of the fuselage and the diffuser section,and the unequal interpolation of the B-spline of the incision, is used to obtain the basic configuration of the submerged inlet.The basic configuration design parameters is adjusted to obtain inlet configurations with different inlet area and position.Subsequently the inlet preliminary optimal configuration is obtained based on CFD numerical simulation evaluation.Whereafter fine design and numerical simulation evaluation of the inlet were carried out to obtain the final optimal configuration, which achieved a total pressure recovery coefficient of 0.9137 at Mach number of 0.85 with the matching point of inlet and engine reached. The maximum absolute value of the distortion coefficient DC60 under all specified states was 0.4448, meeting the requirements of the intake duct design indicators. The research results could provide certain technical references for the design of submerged inlet of various aircraft.

In recent years, wind-dispersed plant seeds (e.g., dandelion and salsify seeds) have attracted considerable attention in the bionics field due to their high-porosity disc structures, which hold potential value for the development of new bionic aircraft. Most existing studies rely on numerical simulations and particle image velocimetry (PIV) technology in traditional wind tunnels to indirectly analyze flow fields and aerodynamic forces. However, these methods are costly, and traditional horizontal wind tunnels fail to meet the experimental requirements for vertical airflow. Therefore, it is particularly necessary to develop a low-cost, efficient, and accurate aerodynamic force measurement method.This study focuses on investigating the aerodynamic force measurement technology for high-porosity disc structures. A vertical flow field test platform was designed and built, and a high-precision, high-sensitivity box-type six-component balance was developed. To avoid the interference of vertical airflow on the measurement accuracy of the balance and the flow field quality, a cascade structure with annular guide vanes was designed by drawing on wind tunnel corner flow guiding technology, which guides the vertically upward airflow to the horizontal direction.The influence of the guide vanes on airflow deflection was analyzed using the computational fluid dynamics (CFD) method. The results show that under a specific installation angle, the airflow deflection is uniform and tends to be horizontal, exerting little influence on the flow field of the test platform. To verify the accuracy of the platform, a 14 mm dandelion seed model was scaled up by 18 times, and numerical simulations and wind tunnel experiments were conducted on simplified models with different porosities. The drag force data showed good agreement, indicating that the platform has high measurement accuracy within the considered error range.

The purpose of this study is to explore the influence of the internal channel configuration of supersonic inlet on stable subcritical condition and buzz, the numerical simulations of inlet schemes with the same external-compression configuration and different internal channel configurations are carried out. The study object is a two-dimensional external compression inlet with a shock-on-lip Mach number of 3.0. Six internal channel configurations are studied at freestream Mach number of 2.0. The working process of each inlet from stable subcritical condition to little buzz and big buzz is obtained. The results show that the internal channel configuration has an important impact on the subcritical characteristics of inlet. The scheme with short internal channel and small outlet height has broad stable subcritical range and big buzz which seems to be triggered by Dailey criterion. The scheme with long internal channel and large outlet height has narrow stable subcritical range, and its buzzes include little buzz and big buzz. The little buzz is independent of the common buzz triggering criterion (Ferri criterion and Dailey criterion). A little buzz formation mechanism of terminal shock/cavity coupled self-excited oscillation is proposed. The study is helpful to increase the understanding of buzz formation mechanism.

During the flight of high -speed aircraft, complex shock-shock interference phenomena occur, leading to a sharp increase in the heat flux in the wall area of the aircraft, which poses a huge challenge to the aircraft’s thermal protection system. To address the extreme aerodynamic heating problem caused by shock-shock interference, reverse jet is applied to the inlet lip of an X-51-like aircraft. A numerical simulation method is used to study the influence of the reverse jet at the inlet lip on the flow field structure and heat flux under the interference of the forebody shock. By changing the angle of attack of the incoming flow, the interference types of the forebody shock on the lip shock at different angles of attack and the influence laws on the lip flow field and heat flux distribution are obtained. Furthermore, the heat -reduction characteristics of reverse jets with different total pressure ratios under types III, IV, and V shock interferences are investigated.The results show that the change in the angle of attack will change the shock interference type at the lip. When the angle of attack changes from 3.8° to 4.0°, the interference type changes from type III to type IV. The reverse jet can convert type III shock interference into type-II-like and type IV shock interference into type-III-like, reducing the lip heat flux. For type V shock interference, the reverse jet will convert it into a more severe type-IV-like, resulting in an increase in the peak wall heat flux. The heat-reduction laws of the reverse jet at the inlet lip of the X-51-like aircraft at different angles of attack are obtained, and the action mechanism of the reverse jet under different shock interference types is clarified,which can provide support for the thermal protection design of the same type of aircraft.In addition, the research results show that the introduction of the reverse jet at the lip will reduce the total pressure loss of the inlet. Compared with the case without jet, the total pressure recovery coefficient is increased by 49%.At the same time, it increases the mass flow rate,which is beneficial to improving the performance of the inlet.