Aerospace Science and Technology最新文献

英文中文

Fault-tolerant target tracking control for the USV-UAV platform via the visual-based guidance and fault-tolerant control

IF 5 1区工程技术 Q1 ENGINEERING, AEROSPACE

Aerospace Science and Technology

Pub Date : 2025-04-22 DOI: 10.1016/j.ast.2025.110230

Guoqing Zhang, Chuanjie Lin, Jiqiang Li, Wenjun Zhang, Xianku Zhang

This paper investigates a new cooperative trajectory tracking framework for the USV-UAV platform that uses a visual-based inversion guidance principle and sensor fault-tolerant control mechanism in the presence of external disturbances. This provides a new strategy for the platform independent of traditional navigation sensor. In the visual guidance module, the reference path of the USV-UAV would be calculated by utilization of the mapping technique according to the sampled images of the target vehicle obtained by an UAV. Further, the desired guidance signals are provided on basis of the fixed position relative to the target vessel. Associate with the developed guidance signal, a robust adaptive fault-tolerant control algorithm is designed to execute a tracking and monitoring mission of the unsupervised vehicles, where the constant and time-varying attitude sensor faults can be addressed by an application of the adaptive observer technique. Besides, the robust neural damping and dynamic surface control techniques are also introduced for tackling the problems of the model uncertainties, external disturbances and computational burden. Through the Lyapunov theorem, the semi-global uniformly ultimately bounded (SGUUB) stability property is proved. The advantages and the effectiveness of the proposed algorithm are evaluated using the numerical simulations.

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引用次数: 0

Air system optimization coupled with electric supercharger matching of a two-stroke aircraft engine based on machine learning and NSGA-III

IF 5 1区工程技术 Q1 ENGINEERING, AEROSPACE

Aerospace Science and Technology

Pub Date : 2025-04-21 DOI: 10.1016/j.ast.2025.110234

Lingfeng Zhong , Qianfan Xin , Rui Liu , Raihanul Islam , Md Saiful Islam , Yufeng Chen

To increase the flight altitude of small unmanned aerial vehicles (UAVs), achieving power recovery at high altitudes through supercharging is crucial. A spark ignition two-stroke engine model was developed in GT-POWER and calibrated using experimental data. Engine performance with an electric supercharger was analyzed. The coupling mechanism of the intercooler in the intake system, the exhaust resonance pipe in the exhaust system, and the electrically supercharged two-stroke engine were studied. The results demonstrated that the supercharged two-stroke engine with the intercooler successfully maintained pwer at alitudes between 1,000 and 6,000 m without degradation, despite the decrease in power-to-weight ratio due to the intercooler. The nondominated sorting genetic algorithm-III (NSGA-III) and machine learning were used to optimize geometry parameters of the exhaust resonance pipe. Brake specific fuel consumption (BSFC), compressor power, and exhaust temperature were selected as optimization objectives. The Pareto solution set revealed significant tradeoff relationships among the objectives. In the Pareto solution set, the optimal values of BSFC, compressor power consumption, and exhaust gas temperature are 389.6 [g/(kW·h)], 1.31 kW, and 604.1 K, respectively. The optimized exhaust resonance pipe can maintain engine and compressor performance in a wide altitude range, allowing the engine to maintain power at higher altitudes.

为了提高小型无人驾驶飞行器（UAV）的飞行高度，通过增压实现高空动力恢复至关重要。在 GT-POWER 中开发了火花点火二冲程发动机模型，并使用实验数据进行了校准。分析了使用电动增压器的发动机性能。研究了进气系统中的中冷器、排气系统中的排气共振管和电动增压二冲程发动机的耦合机制。结果表明，带有中冷器的增压二冲程发动机在海拔 1,000 至 6,000 米之间成功地保持了功率，尽管中冷器导致功率重量比下降，但功率没有下降。非支配排序遗传算法-III（NSGA-III）和机器学习被用来优化排气共振管的几何参数。制动比油耗（BSFC）、压缩机功率和排气温度被选为优化目标。帕累托解决方案集揭示了各目标之间的重要权衡关系。在帕累托解集中，BSFC、压缩机功率消耗和排气温度的最优值分别为 389.6 [g/(kW-h)]、1.31 kW 和 604.1 K。优化后的排气共振管可在较宽的海拔范围内保持发动机和压气机的性能，使发动机在较高海拔地区也能保持动力。

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引用次数: 0

Uncertainty effect of leading edge fouling on aerodynamic performance of compressor cascades

IF 5 1区工程技术 Q1 ENGINEERING, AEROSPACE

Aerospace Science and Technology

Pub Date : 2025-04-21 DOI: 10.1016/j.ast.2025.110235

Panpan Tu , Guang Yang , Limin Gao , Tantao Liu , Song Yang

To investigate the uncertainty effects of leading edge fouling (Fouling-LE) on aerodynamic performance, uncertainty models are developed by integrating fouling thickness distribution with sparse grid techniques. The results indicate that Fouling-LE uncertainty reduces the available range of negative incidence by 2.5°, increases the likelihood of aerodynamic performance degradation (pressure ratio π and loss ω), raises the probability of a lower airflow turning angle, and introduces a significant risk of flow instability across all incidences. Specifically, the probability of Fouling-LE causing π to exceed the stable operating range is 16.14 % and 10.08 %, while the probability of ω exceeding the stable operating range is 9.88 % and 16.25 %, at the incidence of i=-2.5° and 7°, respectively. The influence of Fouling-LE on the flow field progressively extends downstream, altering the aerodynamic loading on the blade surfaces and inducing a forward shift in the transition position.

引用次数: 0

Investigation on the effect of geometric configuration on the flow field morphology and propulsive performance of oblique detonation combustor 几何构造对斜爆燃器流场形态和推进性能影响的研究

IF 5 1区工程技术 Q1 ENGINEERING, AEROSPACE

Aerospace Science and Technology

Pub Date : 2025-04-20 DOI: 10.1016/j.ast.2025.110227

Haoran Yan, Xin Han, Tao Zhang, Chongguang Shi, Yancheng You

This study conducts numerical simulations of Oblique Detonation Wave (ODW) flow fields in confined spaces, focusing on how variations in the combustor’s initial wedge angle and expansion angle affect flow field morphology and propulsion performance. A new initiation mechanism induced by small wedge angles has been identified, enabling the transition from shock-induced combustion to oblique detonation combustion. A simplified model was developed to predict the position and height of the Mach Stem (MS) in the detonation flow field, providing insights into the regulation of the transition from normal detonation combustion to oblique detonation combustion. The study confirms that the incident wave’s intensity and the separation zone's geometric parameters are the two critical factors governing MS dynamics. Furthermore, the downstream parameter variations associated with shock-induced combustion, oblique detonation combustion, and normal detonation combustion are analyzed, explaining the relationship between propulsion performance and flow field structure. Adjusting wall curvature to reduce detonation wave overdrive is found to improve the combustor’s propulsion performance significantly. This research establishes a clear link between detonation flow fields and combustor geometry, offering valuable insights into the dynamic regulation of Oblique Detonation Engines (ODEs) under near-realistic operating conditions.

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引用次数: 0

Adaptive attitude tracking control for reusable launch vehicle with actuator and sensor faults

IF 5 1区工程技术 Q1 ENGINEERING, AEROSPACE

Aerospace Science and Technology

Pub Date : 2025-04-17 DOI: 10.1016/j.ast.2025.110226

Wenting Li, Jinsong Zhao, Chen Sun, Shanwei Su, Yan Lin

This paper investigates an adaptive fault-tolerant control (FTC) for a reusable launch vehicle (RLV) during its reentry phase. Four grid fins are used as system actuators and to avoid feedback errors, sensor redundancy is employed. A novel sensor fusion function for each state variable is constructed, which is differentiable and consists of weighting functions of the redundant sensors. The weighting functions can reflect whether the sensors are normal and in particular, for a faulty sensor, the value of its weighting function can automatically decrease. Consequently, the impact of the fault can be eliminated. By incorporating the sensor fusion functions into the dynamics model of the RLV, an adaptive controller is proposed, which can not only withstand stuck and any finite number of partial loss of effectiveness (PLOE) of actuators but also guarantee the tracking performance. Simulation results are presented to illustrate the effectiveness of the proposed scheme.

引用次数: 0

Drag and overall aeroheating reduction of a dual-disk-dual-jet in rarefied hypersonic flow in near space

IF 5 1区工程技术 Q1 ENGINEERING, AEROSPACE

Aerospace Science and Technology

Pub Date : 2025-04-16 DOI: 10.1016/j.ast.2025.110229

Chen Ye, Shuzhou Fang, Zijian Ni, Tianwei Liu, Shenxing Luo

Based on the direct simulation Monte Carlo method (DSMC), this paper studies the near-space hypersonic rarefied flowfield around the dual-disk with and without jet. The research extends the dual-disk-dual-jet configuration to near space and finds that only increasing the number of disks will no longer be a good solution for reducing drag in hypersonic rarefied flow. The study proposes a comprehensive evaluation of blunt body and forebody shock control devices for aeroheating reduction. A key finding is that an opposing jet on the first disk is crucial for the aeroheating protection of the spike-disk itself, while eliminating or weakening stagnation points on disks. The research further demonstrates that allocating the fixed jet mass flow rate to the opposing and lateral jet is an optimal strategy in hypersonic rarefied flow. This paper also shows that the jet only needs a lower mass flow rate in rarefied flow. This study investigates the performance of a dual-disk-dual-jet model in hypersonic rarefied flow in near space. Compared to the dual-disk without jet (base model), in the studied parameter range, when the pressure ratio is 0.08, the total mass flow rate of the jet is 0.0427kg/s, the drag and aeroheating reduction of the blunt body in the dual-disk-dual-jet model are reduced by 64.3 % and 86.8 %, respectively.

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引用次数: 0

Performance analysis of a rotating detonation turbine engine considering operating condition limitations

IF 5 1区工程技术 Q1 ENGINEERING, AEROSPACE

Aerospace Science and Technology

Pub Date : 2025-04-15 DOI: 10.1016/j.ast.2025.110221

Ying Wang , Fengbo Wen , Liangjun Su , Jiajun Han , Zhiyuan Zhao , Yuxi Luo

This paper presents the construction of a thermodynamic model specifically suited for a rotating detonation turbine engine. It takes into account the actual operating conditions of an aircraft engine, including limitations on the turbine inlet temperature. The study investigates the thermodynamic cycle performance, pressure gain characteristics, energy-saving properties, and economy of the rotating detonation turbine engine. The results show that when the compressor pressure ratio remains consistent, the indicators of the rotating detonation turbine engine are better, and the advantages are obvious at low pressure ratio and high turbine inlet temperature. Under the working conditions selected in this paper, the thermal efficiency can be increased by up to 11.5% and the fuel consumption can be reduced by up to 12.8%. When the total pressure ratio of the engine is consistent, the compressor power consumption of the rotating detonation turbine engine is significantly reduced, which can reduce the number of compressor and turbine stages and improve the engine thrust-to-weight ratio. The influence of varying flight conditions on engine performance was also examined, offering guidance for parameter selection across different scenarios. Combined with the numerical calculation results, the engine's thermodynamic model was characterized using the actual rotating detonation process, and the performance of the rotating detonation turbine engine was evaluated more accurately. The results show that the rotating detonation process experienced by the working fluid in the rotating detonation combustor is in good agreement with the Humphrey model. The cycle curve obtained by two-dimensional numerical calculation is basically consistent with the ideal model. The difference in thermal efficiency calculated by the two methods is 0.3% at the minimum and no more than 10% at the maximum.

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引用次数: 0

Aerodynamic and stealth integrated design of hypersonic vehicle based on discrete adjoint method 基于离散积分法的高超音速飞行器空气动力和隐身综合设计

IF 5 1区工程技术 Q1 ENGINEERING, AEROSPACE

Aerospace Science and Technology

Pub Date : 2025-04-15 DOI: 10.1016/j.ast.2025.110222

Meng Zhang , Lu Xia , Ke Zhao , Longlong Shi , Ruibin Zhang , Yanyang Zhu , Jun Deng

Aerodynamic and stealth characteristic are important indicators that affect the combat capability of hypersonic weapons. In order to coordinate the contradiction between the two disciplines and improve the penetration ability of hypersonic vehicle, the multidisciplinary optimization design based on discrete adjoint method was adopted to optimize the shape of hypersonic vehicle. Firstly, an integrated design platform for aerodynamic and stealth was built by coupling the numerical method of aerodynamics and stealth, the gradient calculation method based on adjoint equation, the free-form deformation (FFD) parameterization method and the RBF-TFI dynamic grid technology. Secondly, taking the HTV-2 aircraft as an example, the aerodynamic shape of the hypersonic vehicle was optimized. After aerodynamic optimization, the lift-to-drag ratio was increased by 11.85%, and the aerodynamic characteristics were greatly improved. While, the stealth characteristics in the main threat area changed little. Finally, the aerodynamic and stealth integrated optimization design of HTV-2 hypersonic vehicle was carried out. The results shown that after the aerodynamic and stealth integrated optimization, the lift-to-drag ratio was increased by 7.82%. At the same time, the average radar cross section (RCS) of head direction ± 60 ° angle domain in the yaw plane was decreased by 40.21%, and the average RCS head direction ± 30 ° angle domain in the pitch plane was decreased by 30.47%. It was verified that the design system based on discrete adjoint method established in this study could realize the aerodynamic and stealth integrated optimization design of hypersonic vehicle under hundred-dimensional design variables and constraints, so that the aerodynamic and stealth characteristics of the aircraft were substantially improved after optimization.

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引用次数: 0

Measurement of the second-mode's propagation velocity in the hypersonic boundary layers of a flared cone with dual-frame focusing schlieren and optical flow velocimetry 利用双框架聚焦分裂和光流测速仪测量扩口锥高超音速边界层中的第二模式传播速度

IF 5 1区工程技术 Q1 ENGINEERING, AEROSPACE

Aerospace Science and Technology

Pub Date : 2025-04-15 DOI: 10.1016/j.ast.2025.110216

Xiaolong Wang , Bo Zhou , Zhixin Zhao , Puyuan Wu , Cunbiao Lee

Understanding the transition mechanism in the hypersonic boundary layer is essential for hypersonic vehicles, where the second-mode wave plays a key role in the process. An experiment utilizing dual-frame focusing schlieren and optical flow velocimetry is performed in the Mach 6 wind tunnel at Peking University. Detailed two-dimensional velocity field, frequency, and density fluctuation intensity of the second-mode waves in the near-wall region are analyzed and discussed. The experiment observes the density gradient's velocity fluctuating in the horizontal direction within the sonic line, while its mean velocity remains uniform. The measurement of the second-mode frequency meets the results by PCB® and FLDI but with a much lower sampling frequency.

引用次数: 0

Lateral jet-rarefied hypersonic freestream interaction over a three-dimensional cone

IF 5 1区工程技术 Q1 ENGINEERING, AEROSPACE

Aerospace Science and Technology

Pub Date : 2025-04-15 DOI: 10.1016/j.ast.2025.110220

Guang Zhao , Chengwen Zhong , Sha Liu , Jianfeng Chen , Hao Jin , Congshan Zhuo

A conserved discrete unified gas kinetic scheme is adopted to analyze how the incoming flow Knudsen number, freestream Mach number, and jet Mach number affect the flow-field characteristics and aerodynamic forces of a three-dimensional cone lateral jet model, thereby accounting for the 3D effect's contribution to jet interaction. Moreover, the differences between the constant-momentum-ratio and constant-pressure-ratio cases are compared and analyzed when either the freestream or jet velocity is changed. The results of this study accurately replicate the flow-field properties in near-continuum conditions, showing good agreement with those in classical literature. The evolution trend of the flow field under the influence of the rarefied-gas effect is also presented. The findings reveal that: 1) the three-dimensional effect markedly reduces the additional force/moment; 2) the principle, previously found in two-dimensional jet-model studies, that the height of the barrel shock remains unchanged with the momentum ratio when the freestream Mach number varies, still applies in the three-dimensional model; 3) when the momentum ratio or pressure ratio is kept constant while increasing the freestream Mach number, the local peak pressure value in the interference region has a linear relationship with the stagnation-point value, though their slopes differ; 4) maintaining a constant momentum ratio while varying the jet Mach number can also make the height of the barrel shock consistent. This research will provide valuable references for the application of jet-control devices in flight vehicles across various flow regimes.

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