This study explores the aerodynamic behavior of Mars rotor systems (single/coaxial configurations) through experimental tests and computational simulations. Detailed analysis of a thin blade rotor system revealed pitch-angle-dependent flow mechanisms. At phi = 15.8 degrees , a stable leading-edge separation bubble forms on the upper surface, which is penetrated and split by shock waves in the tip region (r/R approximate to 0.7-0.9) with increasing rotation speed, creating shock separation-bubble interaction. For phi = 19.8 degrees , mid-span shock-vortex interactions (r/R = 0.5-0.75) occur with different phase vortex shedding phenomena, demonstrating the intricate dynamics of compressible low-Reynolds flows. Systematic evaluation of pitch angles (15.8 degrees-19.8 degrees) elucidates critical performance trade-offs for Martian hovering rotors. Additionally, reduced rotor spacing ratio (H/D = 0.09) amplifies nonlinear aerodynamic interactions, lowering thrust-power ratios due to altered effective angles of attack. Doubling the rotor spacing (H/D = 0.18) leads to an approximate doubling of the thrust, enabling the system to achieve the required thrust of 7.42 N at 2103 RPM, as opposed to 3124 RPM for H/D = 0.09. Such improvements effectively decrease blade interference and reduce the extent of leading-edge separation bubbles. This enhances system efficiency, elevating the Figure of Merit to 0.602 and narrowing the efficiency gap with single rotors from 32.6 % to 7.1 %. Optimal spacing balances thrust performance and rotational energy consumption.

The aerial-aquatic propeller is a core component of unmanned aerial-aquatic vehicles. Aiming at the problem that the efficiency of aerial-aquatic propellers in air and water cannot be balanced, this paper proposes a design method for aerial-aquatic propellers based on the principle of segmented design. The blade is divided into an underwater thrust design zone, a water-air transition zone, and an air pulling force design zone. The underwater thrust design zone is designed according to the criterion of the optimal lift-drag ratio underwater to ensure the underwater propulsion efficiency. The water-air transition zone considers the overall smooth transition from the underwater propulsion section to the air pulling force section. The air pulling force design zone is specially designed for air pulling force requirements. It does not generate effective thrust under the underwater cruise condition and only contributes to induced drag. This zonal design significantly reduces the underwater resistance of the propeller, enabling it to maintain efficient operation in both air and water. Through numerical simulation, the performance differences between the aerial-aquatic propeller designed in this paper and traditional air propellers are systematically compared. The results show that: (1) In the air hovering condition: The hovering efficiency of the optimized propeller is 3% better than that of traditional 9-inch and 10-inch propeller, reaching 57%, however 3% lower than that of traditional 12-inch propeller; (2) In the underwater condition: The propulsion efficiency of the optimized propeller is significantly increased by 11.5% compared with that of the traditional 12-inch propellers, reaching 59.5%, effectively extending the underwater endurance of the unmanned aerial-underwater vehicles. This study verifies the effectiveness of the proposed segmented design strategy in solving the medium adaptability contradiction of aerial-aquatic propeller and provides a reliable theoretical basis and technical reference for the design of aerial-aquatic propulsion systems.

Low Reynolds number flows are prevalent in the aerospace domain, posing significant challenges for aerodynamic analysis and optimization. Previous studies have primarily employed single-mapping neural network models, which can predict drag values from airfoil modeling parameters but lack the capability for inverse reconstruction of airfoil shapes based on specified drag values. To address this limitation, we propose an invertible neural network model that establishes a bidirectional mapping between airfoil modeling parameters and drag values at low Reynolds numbers, thereby enabling inverse design functionality. As a first step, we implemented a zero-padding extension method based on the FrEIA framework to resolve the dimensionality mismatch between airfoil parameter vectors and drag values. We then demonstrated the model's efficacy in predicting drag coefficients from given airfoil modeling parameters with high accuracy. Furthermore, leveraging the invertibility of the network, we achieved inverse prediction and optimization of airfoil modeling parameters by adjusting target drag values. Experimental validation confirms the model’s effectiveness in both forward drag prediction and inverse airfoil shape optimization, underscoring its potential for aerodynamic design applications.

The Martian atmosphere presents a unique aerodynamic challenge for aerial vehicles, characterized by ultralow Reynolds numbers (∼104) and compressible subsonic Mach numbers—a regime where viscous effects and shock-induced phenomena are tightly coupled. This study numerically investigates the drag-divergence behavior of the NACA 0012-34 airfoil under representative Martian conditions using two-dimensional Reynolds-Averaged Navier–Stokes (URANS) simulations with the Transition SST (γ–Reθ) model. Validation against existing Mars-relevant wind-tunnel data at Rec ∼=1.1×104 shows reasonable agreement with experimental trends. Systematic simulations at Rec∼=1.9×104 across Mach 0.2–0.8 indicates a more rapid increase in drag beyond approximately M ≈ 0.6–0.7, rather than a sharply defined divergence point. Results further show that increasing lift coefficient promotes earlier drag rise, whereas modest Reynolds number increases within this low-Re regime primarily affect drag magnitude. These findings quantify the compressibility limits for airfoil performance in Martian conditions and support aerodynamic design considerations for future Mars aircraft and rotorcraft.

IntroductionTo address the mass-law limitation inherent in conventional passive vibration isolators, this study introduces a novel micro-scale hollow-scatterer cylindrical lattice architecture and develops a two-stage optimization framework for achieving broadband low-frequency vibration attenuation.MethodsIn the first stage, a locally resonant unit cell is engineered, comprising a lead-based hollow scatterer, a viscoelastic rubber coating, and a thermosetting resin connector. A Bloch-theory-informed two-degree-of-freedom (2-DOF) reduced-order model is formulated to provide a mechanistic and computationally efficient prediction of the first locally resonant bandgap. Leveraging parametric sensitivity analysis and the Box-Behnken response surface methodology, seven critical geometric parameters are systematically optimized. In the second stage, curvature is deliberately introduced as an additional design degree of freedom by transforming the planar connector into a single-curvature cylindrical configuration. Further parametric investigation and experimental validation are conducted to characterize the attenuation performance.ResultsThe resulting optimized unit cell achieves a 32.7% reduction in the lower bandgap cutoff frequency (to 595.2 Hz) and a 64.9% expansion of the absolute bandgap width (to 762.8 Hz) with model prediction errors consistently below 10%. The curvature-induced geometric stiffening effect shifts the upper bandgap boundary mode from pure out-of-plane bending to a hybrid bending-membrane coupling mode, thereby broadening the bandgap to 1,418.16 Hz. Experimental validation using vibration testing shows reasonable agreement with numerical predictions in terms of attenuation onset and relative bandwidth trend.DiscussionCollectively, this work establishes a synergistic design paradigm integrating data-driven parametric modeling with curvature-mediated structural tailoring, providing a robust, scalable methodology for the rational design of lightweight, high-performance low-frequency vibration isolation systems.

The Martian atmosphere is characterized by a low density and low speed of sound, which result in the low Reynolds number compressible flows. In this regime, conventional airfoils perform poorly due to the boundary layer separation and the formation of shack wave. The current paper investigates the hovering performance and the structure analysis of a Martian rotor blade built with a triangular airfoil using numerical analysis. The airfoil, with a thickness-chord ratio of t / c = 5% at 30% chord, has been shown through experiments to exhibit non-linear lift enhancement due to the roll-up vortex caused by the sharp leading edge at high angles of attack. The designed blade has a pitch axis of 40% chord, close to the airfoil center of gravity. In order to evaluate the blade thickness distribution along the radial station, Carbon Fiber, due to its high strength-to-weight ratio is applied to the blade. It is found that the main source of stress is inertia force rather than aerodynamic loads and that the blade is structurally safe. Finally, the blade reaches a Figure of Merit of FM = 0.73 at the collective pitch angle of 8 deg and the minimum tensile and compressive factor of safety of 2.90 and 1.74 respectively.

Low Reynolds number flows are characterized by boundary layer separation due to the effect of viscous forces. The separated flow may reattach through the exchange between the molecules resulting in Laminar Separation Bubble LSB. LSB, known for its detrimental effect on the performance is sensitive to the airfoil geometry. The current paper describes the use of trip boundary on the suction surface of SD7003 airfoil as flow control technique to enhance the airfoil performance at and using numerical analysis. Due to the complexity of finding the trip size and location, Response Surface Methodology RSM is used to obtain the optimum combinations between the height, width and the position of the turbulator. Thus, the design variables are the height [0.2mm,0.6mm], width [140mm, 200mm] and the position of the trip away from the leading edge [10%c, 25%c]. Using the Desirability approach with Nelder-Mead simplex algorithm, the trip location is found to be effective downstream the separation location of the untripped airfoil. At the optimum design parameters, the length of LSB has decreased by 2.32% and the airfoil performance increased by 6%.

A 2-D airfoil shape optimization in transonic low-Reynolds number regime is conducted. A Navier–Stokes flow solver with a transition model (k-ω SST γ-Re θ ) is used to evaluate the fitness function. Single-point and multi-point formulations of the optimization results are compared. In addition, the effects of Mach number and angles of attack on aerodynamic characteristics of the optimized airfoils are investigated under low Reynolds number (Re = 17,000) and high-subsonic-flow ([Formula: see text]) conditions. The results show that the corresponding drag divergence Mach number curves of the conventional airfoil present almost a parallel shifting at the entire Mach number range. By contrast, the unconventional airfoil starts showing a significant drag reduction when Mach number is greater than 0.75. Besides, the maximum lift-to-drag ratio is highly influenced by the Mach number because of the formation, movement, type, and strength of a shock wave. In addition, the distinguishing difference in the conclusion between two airfoils is that the lift fluctuation of the conventional airfoil amplifies with the increase of the Mach number. However, the unconventional airfoil shows an opposite trend.

In recent years, there has been rapid development in electric aircraft, particularly electric vertical takeoff and landing (eVTOL) aircraft, as part of efforts to promote green aviation. During the conceptual design stage, it is crucial to select appropriate values for key parameters and conduct sensitivity analysis on these parameters. This study focuses on an electric tilt-rotor aircraft and proposes a performance analysis method for electric aircraft while developing a general design tool specifically for this type of aircraft. Subsequently, the impact of wing incidence angle, sweep angle, span, propeller solidity, battery-specific energy, and battery mass on range, maximum takeoff weight, and hover power are analyzed. The results show that the battery mass, wingspan, and wingtip chord length have great effects on the maximum takeoff weight; among these, battery mass had the greatest influence. In terms of range, the battery energy density has a great positive effect on range, while the increase in wing angle of incidence, wingtip chord length and battery mass have some negative effects on range.

Minimum distance to a solid wall is a primary parameter in turbulence models and overset grid assembly for computational fluid dynamics. In present work, a parallel advancing front method is proposed based on partitioned unstructured grids for the sake of further efficient wall distance computation. In order to overcome the inherent problem of conventional advancing front method in parallel environment, a novel "advancing twice and rippling once" strategy is developed to compute wall distance efficiently and further recover the accuracy. Significantly, the established framework is of modular nature to be easily extended to overset grid system for complex multi-body configurations. The performance of the developed techniques is evidenced by comparison with the existing alternative ways in terms of computing efficiency and accuracy. Subsequently, further case studies are performed to examine its capability to deal with complex engineering applications such as a full transportation, a wing-store configuration, a helicopter and a F-16 fighter with onboard payloads. Results show that the proposed advancing front methodology is in practice able to solve extremely complex geometries with both convex and concave shapes with high efficiency and robustness.

火星的稀薄大气环境迫使无人机在亚临界雷诺数范围工作,低雷诺数层流分离问题给无人机气动性能带来极其不利的影响.同时,火星大气的声速较低,使无人机运行的马赫数更高,压缩效应增强并可能产生激波.为研究火星环境下翼型局部振动的流动控制作用,采用基于动网格的数值方法对非定常流场进行模拟.选取NACA5605低雷诺数薄翼型,雷诺数为1.5×104,马赫数为0.43和0.63.时均流场和时均气动力系数结果显示:翼型局部振动能够明显减少时均分离区的大小,起到增升减阻的作用.非定常流场表明流动控制机理在于振动产生的涡流运动抑制了翼型尾缘附近的层流分离.研究了不同振幅、频率和振动位置下的流动控制效果.最佳参数下,马赫数为0.43时升阻比最多提高24.7％,马赫数为0.63时升阻比最多提高52％.

The shapes of pin-fins play an essential role in enhancing overall and local forced convective heat removal. This study presents a new design optimization scheme based on the free shape deformation method, which gives rise to nonuniform pin-fin shapes depending on local flow structures interacting with each local pin-fin. The grid of pin-fins is embedded in a parameter space formed by basic function and control points. By moving the position of the control points, the shapes of the pin-fins and the neighborhood grid are deformed during optimization. The optimization objective function is the heat transfer coefficient of the pin-fin arrays at a limited pressure drop to ensure cycle efficiency. In addition, to reduce the computational resources, the active learning surrogate model for constraint global optimization is adopted, which is combined with computationally expensive computational fluid dynamics simulations. To exemplify the effectiveness of the proposed method on optimizing the pin-fin shape, two-dimensional pin-fins in inline arrangements are considered. Finally, an improvement in the heat transfer of 33.4% is achieved as compared to the baseline. Therefore, the proposed method of pin-fin optimization is effective in improving thermal performance.

In the present study, an efficient overset grid method by means of parallel implicit hole-cutting is proposed for the sake of simulating unsteady flows in aerospace engineering involving multiple bodies in relative movement. In view of the degraded computational efficiency and robustness for conventional overset grid assembly, several innovative techniques are developed within the overset grid assembly process, viz., a bookkeeping alternative digital tree method to speed up the donor-cell searching, a fast parallel advancing front algorithm to accelerate the wall-distance calculation and a message-passing strategy with efficient information communication and lower storage expenditure within distributed computational architecture. The contribution of the developed techniques is evidenced by comparison with the existing alternative ways in terms of computing efficiency. Subsequently, the overset grid method is embedded into an in-house programed URANS solver to examine its capability in predicting the flow field of complex applications such as helicopter, store separation and component deploying. Results show that the developed overset grid methodology is, in practice, able to resolve the aerodynamic characteristics of complex aerospace engineering with a high-fidelity flow topology and accuracy.

A computational study has been conducted on various airfoils to simulate flows at low Reynolds numbers 17,000 and 21,000 with Mach number changes from 0.25 to 0.85 to provide understanding and guidance for Mars rotory wing designs.The computational fluid dynamics tool used in this study is a Reynolds-averaged Navier-Stokes solver with a transition model (k-ω SST γ-Reθ).The airfoils investigated in this study include NACA airfoils (4, 5, and 6% camber), UltraThin airfoils, and thin cambered plates (3% camber, but various maximum camber locations).Airfoils were examined for lift and drag performance as well as surface pressure and flow field characteristics.The influence of Reynolds and Mach number effects on the flow past airfoils was analyzed and significant impact on flow separation and subsequent wake patterns was demonstrated.In general, the Mach number shows a significant impact on the flow past airfoils, including of flow separation, trailing-edge wake patterns, and shock wave types.A stretched trailing-edge separation pattern is clearly observed from NACA and Ultra-thin airfoils.In addition, NACA airfoil shows a trailing edge separation, and a shock wave starts to appear at moderate to high incidences.By contrast, UT airfoil shows a leading-edge separation as incidence increases.As angles of attack rise at moderate, an "A-type" shock appears on the up-airfoil's surface, which alters the response of the outer flow pressure to displacement surface perturbations, including the influence on the growth, curvature, and even unsteadiness of the separated shear layer.

This paper extends our previous work on the isogeometric dynamic buckling analysis of thin-shell structures to the trimmed and multipatch situation where features such as cutouts and stiffeners can be easily incorporated. To be specific, a modified generalized-[Formula: see text] time integration scheme combined with a geometric nonlinear isogeometric Kirchhoff–Love shell element is used to simulate the complex buckling and postbuckling behaviors of thin-shell structures. The developed method can damp properly high-frequency contents while maintaining second-order accuracy in the dynamic buckling analysis. For the integration of arbitrary-shaped trimmed elements, a geometrically exact blending function method is developed to improve the efficiency of the dynamic shell buckling analysis. To deal with multipatch geometries, a penalty-based weak coupling approach is developed, where coupled patches with nonconforming trimmed interfaces or even with prescribed angles, such as stiffeners, can be analyzed. We demonstrate the accuracy, stability, and flexibility of the proposed framework with several numerical examples. In particular, the influences of “free” and “partially free” control points, penalty factor, trimming, as well as different modeling strategies on the dynamic solutions of shell structures are investigated.

A wing is an important part of the aircraft to improve aerodynamic performance. The current study is focused on an adaptive surrogate algorithm for airfoil aerodynamic optimization, which is based on a multi-output Gaussian process model. The conventional design method seriously relies on wind tunnel experiments and expensive computational simulations. The metamodels can significantly improve design efficiency and hence reduce the overall design costs. An active learning algorithm is proposed to improve the effectiveness of the multi-output Gaussian process model. The NSGA-II algorithm is adopted to obtain the optimal Pareto set with the optimization objectives of lift and drag coefficients for adaptive airfoil shapes. Besides, the Bezier curve and radial basis function are utilized in this study for airfoil mesh deformation. The results show that the airfoil shape can be obtained effectively by integrating the metamodel, active learning algorithm, and multi-objective optimization algorithm. The optimized results are of great engineering applications.

A numerical investigation on propeller-induced flow effects in tractor configurations on a Zimmerman wing-fuselage using the cambered thin airfoil is presented in this paper. The Reynolds number based on the mean aerodynamic chord was 1.3 × 105. Significant aerodynamic performance benefits could be found for a propeller in the tractor configuration. The numerical results showed that the propeller slipstream effect on the wings was highly dependent on the size of the propeller, and the major slipstream effect was working at 60% inboard wingspan, whereas less effects were observed towards the wingtip. The propeller slipstream increased the local angle of attack on the up-going blade side. This effect simultaneously augmented the section lift. The unsteady Reynolds-averaged Navier–Stokes (URANS) simulations helped to improve understanding of the interaction of the propeller wake and the wing-fuselage, which is an important aspect to guide the design of future efficient and controllable micro air vehicles. The results indicated that, in MAV designs, the slipstream from the propeller had a significant effect on the wing aerodynamics, regarding both performance and stability of the vehicle.

Model order reduction approach generates lower dimensional approximations to the original system while preserving model’s essential information and computational accuracy. For nonlinear structural dynamic problems, where the stiffness matrix is configuration dependent, an iterative solution procedure is inevitable and a revisit to all the elements is essential for updating the stiffness matrix. In this paper, the nonlinear dynamics of the planar curved beams and 3D cylindrical shells are studied based on the isogeometric analysis and their model order reductions are investigated based on the proper orthogonal decomposition and discrete empirical interpolation method (POD-DEIM). Numerical results show that IGA-based POD-DEIM method significantly improves the computational efficiency of the nonlinear dynamic analysis of the beam and shell structures.

Abstract In this paper, the permitted temperature value of the battery cell and DC‐DC converter is proposed. The flow and temperature field of the lithium‐ion batteries is obtained by the computational fluid dynamic method. Thus, the package structure of the battery pack is optimized based on four influencing factors. The results indicate that (1) setting a new inlet on the wall, I can improve ventilation and the inlet is better located below the waist of the battery pack. (2) Air inlet location close to the fan is easy to generate short air circuit, which leads to DC‐DC converter in poor condition of heat dissipation. (3) Adjusting the size of the air inlet mainly affects the temperature distribution of the cells but has little effect on the temperature of the DC‐DC converter. (4) Regulating the gap size can enhance the cell temperature uniformity. (5) The optimized battery pack structure is obtained, where the maximum cell surface temperature is 297.51 K, and the maximum surface temperature of the DC‐DC converter is 339.93 K. The above results provide an approach to exploring the optimal design method of lithium‐ion batteries for the container storage system with better thermal performance.

This study aims to develop a non-intrusive reduced-order aerodynamic modeling framework for agile motion across a broad operating range in low-Reynolds-number incompressible flows. Since most previous reduced-order models are only valid in the neighborhood of the operating point, large amounts of high-fidelity input-output data have to be generated to cover the interested operating range, which requires expensive computational resources. To tackle such an issue, a signal interpolation approach by combining the discrete empirical interpolation method (DEIM) with Kriging technique is proposed to approximate the aerodynamic responses to a prescribed linear ramp-step maneuver with remarkable accuracy, instead of performing computational fluid dynamics (CFD) simulation at each operating point. The interpolated input-output data can be used to establish low-dimensional state-space aerodynamic models, by identifying the stability derivatives first and subsequently capturing the remaining transient dynamics via the eigensystem realization algorithm (ERA). To demonstrate the proposed framework, unsteady lift coefficients of a two-dimensional flat plate pitching about the leading edge are investigated over the range with Reynolds numbers 100∼500 and base angles of attack 0∼10 deg. Numerical results show good agreement between the reduced-order models and high-fidelity CFD simulations.