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In order to solve the problems of poor weight estimation accuracy and difficult layout design trade-offs in the concept demonstration and overall program definition stage of civil aircraft, Aircraft Layout, a civil aircraft overall program structure rapid design software, is used to establish a numerical calculation-based layout program rapid design simulation platform. Based on the forward fuselage structure, this paper gives the results of rapid modeling and finite element simulation and analysis of the forward fuselage isotropic section of a certain type of aircraft, which provides a reference for the optimization design of the structural components of the forward fuselage section of a certain type of aircraft in the conceptual scheme design stage.

To address the issue of large-scale noise clusters and small-scale scattered noise in the point cloud of aircraft wing-fuselage docking surfaces, this paper proposes a multi-scale hybrid filtering denoising method. By utilizing the k-means++ clustering algorithm and introducing a density weight factor along with a dynamic neighborhood radius mechanism, large-scale noise clusters are effectively segmented and removed. Additionally, adaptive radius filtering is employed to eliminate residual fine noise, achieving a hierarchical denoising process across different scales. Validation using actual measurement data from a certain aircraft wing-fuselage docking surface demonstrates that this method outperforms traditional statistical filtering and singleradius filtering in denoising effectiveness, providing a high-quality data foundation for subsequent assembly accuracy control.

Aircraft crash analysis and design methods are important aspects of aircraft design. A full-scale finite element model of the B737-200 fuselage section containing 6 frame sections with an auxiliary fuel tank is developed using the explicit finite element code, Pamcrash, to perform simulated analysis. The simulation result outputs the fuselage’s maximum deformation. Seat tracks’ acceleration responses are compared with the drop test to determine the method’s correctness. This method is then applied to simulate the crash of an airplane fuselage section with an auxiliary fuel tank in the cargo compartment at a velocity of 9.144m/s to provide reference for engineering applications.

This study aims to examine the rotor surface pressure distribution, the rotor’s effects on the wing and fuselage, and the surrounding flow field structure during the operation of a tilt-rotor Unmanned Aerial Vehicle (UAV). Using Fluent software, numerical simulations are performed to analyze the flow field around the rotor and the pressure variations across its surface, under three operational scenarios: vertical climb, rotor tilt, and high-speed level flight. The findings indicate that the rotor’s surface pressure is predominantly concentrated at the leading edge across all three conditions, although the pressure distribution varies along the blade surface. Additionally, the rotor’s influence on the wing and fuselage is more pronounced during the rotor tilt and high-speed level flight conditions.

Hybrid hydrogen fuel cell (HH2FC) aircraft is one of the viable technological pathways to reduce carbon emissions in the aviation industry. For a configuration with aft-fuselage propulsion, a novel method is established for the initial sizing of the HH2FC aircraft in a situation where empirical data is lacking. The carbon emission reduction capabilities and potential for reducing fuel and carbon tax costs can be preliminarily assessed by the proposed initial sizing method. A case study on a narrow-body aircraft indicates that, at the current technological level, the carbon emissions and cost of the HH2FC aircraft will increase to achieve the same payload and range as the conventional aircraft. However, with the improvement of hydrogen technologies, such as fuel cell power density and hydrogen tank gravimetric density, HH2FC aircraft will reduce 6.6% carbon emissions compared to conventional aircraft. As a result, both economic and environmental benefits can be achieved in the future.

This study focuses on the flight test design for helicopter emergency flotation systems, and conducts modeling and simulation analyses of post-inflation deployment flight dynamics characteristics. First, CFD software is employed to compute the aerodynamic characteristics of the fuselage in both the deployed and folded states of the flotation airbags, thereby obtaining the corresponding fuselage aerodynamic data. Subsequently, a flight dynamics model of a single-rotor with a tail rotor helicopter in the emergency flotation configuration was established in Simulink. The Simulink Linear Analysis Tool was then employed to simulate and analyze the equilibrium characteristics of the helicopter under various flight conditions after the flotation airbags were deployed. Finally, the influence of the emergency flotation configuration’s aerodynamic characteristics on the pitch inputs and attitude angles required for stable flight following airbag deployment was summarized.

The airtight load is one of the main loads on civil aircraft fuselage, and the fuselage cross-section will tend to deform in a positive circle under the effect of the airtight load. Deformation varies for structures with different contour curvatures for the same airtight load and structural stiffness, especially in the contraction section of the fuselage. Therefore, the deformation of the structure can be controlled by optimizing the contour curvature without increasing the structure stiffness, resulting in a lighter overall weight for the same deformation, thus achieving weight reduction. In a civil aircraft concept design phase, structural deformation control was carried out by optimizing the nose contour, and the results showed that a structural weight reduction of approximately 4% was achieved with a maximum reduction of 23% in structural deformation.

To study the interference characteristics of the intermeshing-rotor fuselage, a numerical simulation method for the flow field of a crossed twin-rotor helicopter in a hovering state is established based on the lattice Boltzmann method. The effectiveness of the proposed numerical simulation method is first verified by the ROBIN helicopter rotor/fuselage interference example. Then, taking the simplified JZ-60 rotor and S-97 helicopter fuselage as the research objects, the aerodynamic interference characteristics of the intermeshing rotor and the fuselage are investigated, and the aerodynamic interference law between the two under the hovering state is obtained. The results show that the intermeshing rotor has periodic tension fluctuation, and the fuselage interference leads to about a 4% increase in the total tension of the rotor. The maximum hovering efficiency increases by about 5%, and the larger the tension coefficient is, the larger the fuselage gains on the hovering efficiency is. The fuselage generates periodic negative lift and lateral force under the influence of the rotor. The hovering efficiency gains of the rotor mainly originate from the increases of the blade lift coefficients in the vicinity of the azimuth angles of 0° and 180°. The rotor is in the range of 0.5° and 180°, which has the strongest interference effect on the fuselage near 0.6 R and -0.7 R.

A numerical model of a commercial fuselage section with dummies and seats under a crash environment was investigated to evaluate the dynamic responses of occupants. Firstly, the two dynamic tests about seats and occupants restraint system were performed to meet the requirements specified by CCAR-25, including the tension loads in straps and the head injury criterion, and gained a method for simulation on seats and occupants restraint system. Secondly, the detail fuselage FE model was built to compare different vertical velocities; the results showed that the greater the velocity is, the more violently the dummies moved, while the deformation of structures had no difference.

This research work focuses on fluctuating load spectrum generation and fatigue analysis of a wing‐fuselage attachment lug joint, in accordance with the requirements outlined in 14 CFR Part 25 for transport category aircraft. The study involves the generation of a fatigue load spectrum using an exceedance curve, followed by finite element analysis of a simple round‐ended lug. The exceedance curve is a statistical tool used to represent the probability of a variable exceeding a particular threshold over a specified time period. It is particularly useful in fatigue analysis as it helps quantify the likelihood of stress levels surpassing critical limits, aiding in the assessment of structural integrity and durability. The CAD model of the lug is created in CATIA, and the corresponding FE model is obtained using Altair HyperMesh. The investigation pertains to a turboprop aircraft weighing approximately 25,000 kg and accommodating 70–90 seats, with ATR 72 and Dash Q 400 being the prominent choices in this category within India. Mission Profiles for these aircraft are obtained from the Directorate General of Civil Aviation (DGCA) website, utilizing scheduled flight data. To ensure the structural integrity of the lug, a static analysis is performed for FE model convergence, leading to the determination of stress concentration factors. Subsequently, fatigue analysis is conducted using Nastran Embedded Fatigue (NEF), considering constant amplitude loading with stress ratios of −1 and 0.1. the fatigue life of the component is predicted based on Goodman and Gerber failure criteria. The analysis yields crucial insights into the fatigue life of the lug and its damage accumulation over time. The CAD model of the lug is created in CATIA, and the corresponding FE model is obtained using Altair HyperMesh. The investigation pertains to a turboprop aircraft weighing approximately 25,000 kg and accommodating 70–90 seats, with ATR 72 and Dash Q 400 being the prominent choices in this category within India.