Aerospace Science and Technology最新文献

Transonic flows at high Reynolds numbers can lead to high dynamic pressures and, consequently, to aerostructural deflections of aircraft structures. This study aims to develop and validate a high-fidelity static aeroelastic analysis environment that is efficient and that can be used in an industrial setting. The aerodynamics is represented by numerical solutions of the Reynolds-averaged Navier-Stokes equations with appropriate turbulence closures. The load transfer process uses finite element shape functions in order to distribute the aerodynamic loads into the structural discretization. The structural analysis employs a modal basis approach, and a wingtip deflection convergence study is performed to find an adequate modal basis size. Radial basis functions are used for the fluid mesh displacement, and the influence of the support radius is evaluated to determine the optimal values relative to the wing mean aerodynamic chord. The capability is tested using the static aeroelastic benchmarks of the High Reynolds Aerostructural Dynamics Project (HIRENASD) and NASA's Common Research Model (CRM). The static aeroelastic results demonstrate robustness and consistency for the aerodynamic coefficients, pressure distributions, and structural deflection predictions at different normalized dynamic pressure values and grid refinement levels.

The variable cycle high-pressure compression system (VCHCS) consists of a core driven fan stage (CDFS) and a high-pressure compressor (HPC), which are crucial components for controlling the variable cycle engine's bypass ratio. However, the double bypass (DB) matching mechanism of the VCHCS remains insufficiently understood. This study focuses on investigating the DB matching characteristics of the VCHCS under different representative internal bypass conditions. The findings demonstrate that as the internal bypass conditions transition from near choke to near stall, the external bypass stability margin of the VCHCS decreases progressively, and the external bypass characteristic curve of the HPC exhibits an approximate “counter-clockwise rotation” pattern. Furthermore, the operating state of the CDFS and the redistribution of internal and external bypass flow rate, collectively determine the matching state of the compressor system. Building upon the aforementioned summarization regarding the impact of internal bypass conditions on the matching characteristics of the VCHCS, this research integrates flow field analysis to explicate the mechanism of stall under typical operating conditions. The investigation reveals that, when the internal bypass condition under design point and the external bypass condition approaches the stall boundary, an excessive accumulation of low-energy fluid occurs at the external bypass outlet guide vane, subsequently triggers the occurrence of stall. Specifically, when the internal bypass conditions under the near stall point, the predominant limitation on further elevation or reduction of external bypass back pressure lies in the extensive blockage of a massive low-energy fluid on the suction surface of the HPC second stage stator.

The paper investigates the performance improvement issue for reentry vehicles under uncertainties from the perspective of probability. The disturbance estimation-triggered control (DETC) proves to achieve transient performance increase compared with the standard disturbance-observer control methods, and the presented approach further exploits the probability-oriented transient performance improvement based on the collaborative and adaptive Bayesian optimization (CABO) technique, which constructs the main contribution of the paper. Based on the attitude dynamics of reentry vehicles, the DETC method is first introduced to guarantee the tracking stability and robustness against the uncertainties including the aerodynamic perturbation and wind effects. Meanwhile, the performance improvement is analyzed theoretically. Then, by virtue of the CABO algorithm, the CABO-based DETC is presented by combining the performance and probability indexes. Finally, the simulation results verify the effectiveness of the proposed control scheme and parameters influence is also discussed.

Space debris generally exhibits complex tumbling motions, and its direct capture may cause damage to space manipulator and base spacecraft. Current detumbling strategies typically require continuous contact collisions with the target and do not take into account actuator limitations. Thus, this paper presents an adaptive optimal concurrent control of space free-floating multi-fingered robot (SFMR) for multipoint repeated contact detumbling of non-cooperative targets. Firstly, the multipoint repeated intermittent contact model between the SFMR and the target is established. Further, an optimal admittance control that considers manipulator actuator limits and target motion bounds is formulated to generate a compliant detumbling trajectory. Through transforming the state and input inequality constraints into extended dynamical subsystems and saturation functions, respectively, the optimal control problem (OCP) is transformed into a readily solvable equality constraint problem. Moreover, an enhanced nonsingular terminal sliding mode (ENTSM) control with radial basis function neural network (RBFNN) compensation is presented in the presence of lumped uncertainties and external disturbances, which achieves rapid finite-time convergence and low-chattering. The simulation results show that the proposed method can reduce the velocity of the space target effectively without causing base spacecraft interference while achieving accurate trajectory tracking.

The rotating detonation gas turbine boasts several advantageous characteristics, including self-pressurization, minimal entropy increase, high thermal efficiency, and a high thrust-to-weight ratio. These features make it a promising candidate for the next generation of aerospace power devices. This study investigates the flow characteristics and performance attributes of the supersonic turbine stage coupled with a rotating detonation chamber based on different rotational speeds, blade ratio of rotor and stator, and hub radius, and uses DMD (Dynamic Mode Decomposition) to analyze the main modes of the supersonic turbine. The research identifies distinct wave modes in both the stator and rotor blades of the supersonic turbine. In the stator blades, the modes include waveless contact modes, opposite side λ-wave oblique shock wave modes, and derived modes from opposite side λ-wave oblique shock waves. In the rotor blades, the modes encompass non-separation modes (waveless modes, asymmetric dual λ-wave modes, oblique cut wave modes, and single-side λ-wave modes on the pressure surface), leading-edge stator excitation single separation bubble modes, saddle-type separation modes, and dual separation bubble modes. The efficiency and power of the supersonic turbine stage increase with the rotational speed. The turbine efficiency is highest when the blade ratio is 10:7. Under multi-wave mode conditions, detonation waves demonstrate adaptive self-regulation abilities, ultimately achieving equilibrium in speed, detonation wave height, and spacing. The dominant modes in the stator blades and combustion chamber are governed by detonation waves and slip lines, while in rotor blades, the primary mode is the flow-directional vortex mode governed by detonation waves and slip lines at a frequency lower than their propagation frequency. This research holds certain reference significance for understanding the internal flow characteristics of rotating detonation gas turbines, as well as discussing the patterns of efficiency, power, and load characteristics, thereby offering valuable insights for the practical application of rotating detonation gas turbines.

Due to its physical characteristics of large viscosity and non-volatilization, aviation kerosene has a poor atomization effect in aeroengine cylinder, and is inappropriate for aviation piston engine unless high pressure injection is used. Air-assisted injection can effectively solve this problem. In the paper, the CFD software is used to establish a 3-D numerical model of the air-assisted injector, and the influence of the nozzle shape on the airflow movement and interaction between gas and liquid is investigated. The accuracy of the simulation model is confirmed by the comparison of the simulation results with the spray morphology and penetration of the spray experiments in the constant volume bomb. Based on this, three nozzle shape models are established to simulate the air-assisted spray flow field of aviation kerosene RP-3 under various ambient back pressures. The influence of nozzle shape on the flow state of compressed air and spray characteristics is compared and analyzed. The results show that when the back pressure is 0.09 MPa, the oblique shock waves can be observed near both large and small circular arc nozzle exits, and the attenuation degree of airflow velocity by the oblique shock wave is relatively small. The stronger interaction between the gas-liquid two phases is beneficial to fuel atomization. Moreover, the normal shock wave appears in the conical nozzle where the injected nitrogen has less kinetic energy. Several large-scale vortices are generated in the near field of the spray, which promotes the mixing of fuel and surrounding nitrogen. Therefore, for aviation kerosene which is difficult to atomize, a large and small circular arc shaped nozzle with strong atomizing ability should be used. When the mixture is required to be evenly distributed in the combustion chamber, the conical nozzle should be preferred.

In this study, a cost-effective high-fidelity six-degree of freedom module augmented with an onboard controller is developed. The unmanned aerial vehicle (UAV) under consideration is a wing-alone configuration with a cropped delta planform and reflex airfoil cross-section designed to perform manoeuvres at high angles of attack. The six-degree of freedom dynamics for the aforementioned configuration is developed based on Newtonian mechanics incorporating linear and nonlinear aerodynamic models to expand flight simulations in corresponding regimes. Aerodynamic stability and control parameters for the simulations are obtained from full-scale wind tunnel measurements and flight tests. It is observed that the linear aerodynamic model is valid in angle of attack (α) domain $-5°\le \alpha \le 10°$, followed by nonlinearity up to $\alpha \approx 21°$, post which lift coefficient remains almost constant with angle of attack till 45°. The simulation model is validated with the corresponding flight data obtained from flight tests. A user-friendly graphical interface with real-time animation of UAV has been developed for the aforementioned configuration by integrating MATLAB and Flightgear environment with real-time control input from the hardware to enable the pilot for visual and tactile feedback, respectively. A controller based on total energy control is designed for velocity and altitude hold modes, while an attitude hold controller is designed for pitch and roll angle tracking to enable controller-in-the-loop remote piloting simulations. In order to enhance practical applicability, the designed module has been tested under various simulated wind conditions. It is noticed from the results that the steady-state error in attitude remains below 2% with a maximum overshoot of 5° from reference. The total energy controller achieves a zero steady-state error with a maximum overshoot of 2 m/s for velocity and 6 m for altitude, respectively. The developed module enhances remote piloting training, flight response assessment, and control algorithm testing for cropped delta planform UAVs.

The present work investigates the application of vortex generators for separation control in axisymmetric isolator flows. Reynolds-Averaged Navier-Stokes computations were performed to simulate a Mach 2.5 flow past a half-isolator geometry with a ${20}^{\circ }$ axisymmteric compression ramp based on the experiments of Funderburk and Narayanaswamy at North Carolina State University. Single vortex generator and a vortex generator array placed upstream of the shock-induced separation region were investigated. Near wall streamlines, surface pressure contours and density contours on crosswise planes were compared with experiments for flow control using a single vortex generator. Results indicate that the present computations are able to capture the wake and shock structures, and also predict reduction in the streamwise extent of flow separation downstream of the device trailing edge. Finally, the effects of the shape/orientation (forward facing/backward facing) of a single vortex generator, and the design of an array of three vortex generators (with devices of similar and mixed orientations) on the flow separation were also investigated. Contours of near-surface streamwise velocity showed that device orientation had a strong effect on separation control, which is attributed to the differences in the primary streamwise vortices shed in the two cases. Further, the overall reduction in the footprint of separated flow was determined to be most with the use of an array of vortex generators with mixed orientations.

This study presents an innovative approach to enhancing the performance of perovskite solar cells through the integration of a functionally graded triply periodic minimal surface (FG-TPMS) layer. The research focuses on the mechanical and vibrational characteristics of doubly curved panels embedded with three distinct iterations of the FG-TPMS model: the primitive, gyroid, and wrapped package graph (IWP). By employing higher-order shear deformation theory (HSDT), the analysis accounts for the complex geometrical and material gradations within the FG-TPMS structures. An advanced analytical method utilizing trigonometric functions is developed to accurately predict the natural frequencies and mode shapes of these novel composite structures. In order to assess the vibrations of TPMS-reinforced perovskite solar cells surrounded by an elastic foundation, this work proposes the implementation of a novel Support Vector Machine (SVM)-deep neural network (DNN)-Genetic Algorithm (GA) employing mathematical modeling datasets. Using the SVM-DNN-GA algorithm, predicted accuracy is improved. In order to simulate and forecast the vibrational behavior of the reinforced solar cells, the integrated methodology makes use of the advantages of each technique. The results indicate that the integration of FG-TPMS layers significantly enhances the mechanical stability of the perovskite solar cells. The application of HSDT reveals detailed insights into the dynamic responses of the doubly curved panels, highlighting the potential for fine-tuning their vibrational characteristics to further improve solar cell performance. This research underscores the potential of FG-TPMS structures in advancing solar cell technology, providing a foundation for future studies to explore the integration of complex geometries and material gradations in photovoltaic applications.

To further improve the performance of the afterburner, this study proposed a new scheme for the integrated afterburner with a strut flame stabilizer and a mixer. A numerical study was carried out to examine the cold performance of this scheme at different flight altitudes, inlet Mach numbers, and bypass ratios. The results showed that the integrated afterburner had a good flow field distribution with four low-speed recirculation zones formed at appropriate locations. The obstruction effect of the strut, airflow mixing, and vortex shedding were the main factors affecting the total pressure recovery performance. The total pressure recovery coefficient decreased with the increase in inlet Mach number, bypass ratio, and flight altitude. Nevertheless, the integrated afterburner maintained good total pressure recovery performance with a total pressure recovery coefficient greater than 0.965. The cold air at the outlet of the mixer on both sides of the strut formed a recirculation zone at the tail end of the strut, thereby improving the thermal mixing performance of the integrated afterburner. The thermal mixing efficiency increased with the bypass ratio and flight altitude, while it decreased with increasing inlet Mach number, but it was still higher than 0.80.