Fuselage Design Research Articles

An analytical solution to “pillowing effect” for composite fuselages

Abstract “Pillowing effect”, introduced by inner cabin air pressure on stringer-frame stiffened fuselage cylinder, is a critical loading case for civil aircraft fuselage. The “pillowing effect” causes out-of-plane bending and shear to fuselage panels, which is a rather huge load for such membrane structures. The fuselage’s longitudinal load and bending load interact nonlinearly with each other in a so called “beam-column” formation, invalidating the superposition principle for linear elasticity. An analytical solution to the axisymmetric “beam-column” model is given in this study to account for the “pillowing effect” of orthotropic composite fuselages. This formula can be employed to perform stress checks against “pillowing effect” during fuselage design.

Numerical two-dimensional optimization of a cylindrical fuselage for bioinspired unmanned aerial vehicle based on non-dominated sorting genetic algorithm-II

To achieve an aerodynamically efficient design of fuselage for bioinspired unmanned aerial vehicles (UAV), the development of a systematic method is of paramount need. This work presents a multiobjective optimization study for shape parameterization of a 2D fuselage of foldable flapping wing UAV in low Reynolds number (2e05) flight conditions. The novelty of the research work lies in integrating S1223 airfoil characteristics for predicting the efficient shape design of the fuselage. This paper offers four sections; (i) the validation of the ANSYS Fluent model with experiment, (ii) the exploration of design variables using Design of Experiment (Sparse Grid Initialization), (iii) the implementation of response surface method (Genetic Aggregation) to know about dimensional sensitivity among independent and dependent variables and (iv) the application of multiobjective optimization method i.e. NSGA II to optimize the drag and lift coefficient. To identify the superiority, a comparative study between the original and optimized fuselage is presented considering many parameters like pressure contours, velocity contours, pressure coefficients, and streamlines representations. It is evident from various results that the 2D optimum shape significantly minimizes the drag coefficient and increases the lift coefficient. This work also makes room for 3D shape optimization, which helps in prototype fabrication for real-time flight conditions.

Computational investigation of the aerodynamic performance of an optimised alternative fuselage shape

PurposeThe purpose of this study is to re-evaluation fuselage design when the main wing’s has the ability to fulfill stability requirements without the need for a tailplane. The aerodynamic requirements of the fuselage usually involve a trade-off between reducing drag and providing enough length for positioning the empennage to ensure stability. However, if the main wing can fulfill the stability requirements without the need for a tailplane, then the fuselage design requirements can be re-evaluated. The optimisation of the fuselage can then include reducing drag and also providing a component of lift amongst other potential new requirements.Design/methodology/approachA careful investigation of parameterisation and trade-off optimisation methods to create such fuselage shapes was performed. The A320 Neo aircraft is optimised using a parameterised 3D fuselage model constructed with a modified PARSEC method and the SHERPA optimisation strategy, which was validated through three case studies. The geometry adjustments in relation to the specific flow phenomena are considered for the three optimal designs to investigate the influencing factors that should be considered for further optimisation.FindingsThe top three aerodynamic designs show a distinctive characteristic in the low aspect ratio thick wing-like aftbody that has pressure drag penalties, and the aftbody camber increased surface area notably improved the fuselage’s lift characteristics.Originality/valueThis work contributes to the development of a novel set of design requirements for a fuselage, free from the constraints imposed by stability requirements. By gaining insights into the flow phenomena that influence geometric designs when a lift requirement is introduced to the fuselage, we can understand how the fuselage configuration was optimised. This research lays the groundwork for identifying innovative design criteria that could extend into the integration of propulsion of the aftbody.

Optimization of an aircraft fuselage assembly to minimize radiated sound power

During the structural design of an aircraft fuselage, the acoustic performance is generally not considered. However, vibrations transmitted through the fuselage generate noise leading to passenger discomfort. Currently, the main source of noise reduction comes from the addition of damping material to the fuselage. This work investigates the use of various design optimization techniques to reduce radiated sound power by changing the fuselage structure’s geometry which consists of a skin panel, frames, and stringers. Design optimization tools, such as topology, size, and shape optimization were used to determine an optimal design for each design variable that minimized radiated sound power in the fuselage assembly. The design variables that were studied included skin panel thickness and design, frame and stringer spacing, and frame and stringer cross section. Equivalent radiated power was used as an objective function for numerical optimizations within Altair OptiStruct as an indirect method for minimizing radiated sound power numerically. To understand the source of the sound power improvements, a physics analysis was conducted for each design variable that compared sound power, equivalent radiated power, and radiation efficiency.

A new approach for correlating Paris parameters with fatigue life of aircraft fuselage panels

This work presents a numerical technique that establishes a relationship between the C and m parameters of the Paris Law and the number of fatigue life cycles. This relationship is particularly important for determining the material's ability to withstand a specified number of cycles in an aircraft fuselage design. The methodology is based on a multiscale problem and comprises two stages: the macro model, which focuses on internal stresses and critical point location, and the micro model, which considers the critical fatigue life cycle number (n) across a "grid" of C and m parameters, resulting in the optimal curve N(C,m). To validate the adopted methodology, the academic program BemCracker2D has been used to calculate the internal stress fields, simulate cracks, and estimate fatigue life. Three case studies were conducted, and the results indicated that the range of C and m values obtained depends on the configuration of the macro model, the physical parameters of the material, and the defined number of cycles in the design. Ultimately, this computational technique allows for generalization to any fuselage damage analysis model, providing C and m Paris parameter data to mitigate fatigue damage.

Study of the influence of turbulent flows on the aerodynamic performance of fixed-wing aircraft type UAV MALE

A study of the air behavior in the near-wake region of the NACA0012 airfoil infinite wing model at low Reynolds numbers was conducted. The study consisted of comparing the results obtained from experiments with those from numerical simulations. The numerical results of the velocity profiles in the near-wake region showed good agreement with those provided by the PIV system for most of the simulations and those in the literature. In addition, boundary layer separation and reattachment forming a laminar separation bubble along the airfoil was captured using simulations in Ansys Fluent and compared with a literature database to better confirm the simulation results. The UAVs have a complex cross-section shape of the fuselage. The second part of this paper explores the aerodynamic consequences of its shape, at the operational flight ceiling, in a comparative approach for three different fuselage designs.

Parallel multi-objective Bayesian optimization approaches based on multi-fidelity surrogate modeling

Aerospace product design optimizations, such as micro-aerial vehicle fuselage design, often involve multiple objectives. Multi-objective Bayesian optimization (MOBO) is an efficient approach in solving problems concerning multiple conflicting objectives. Conventional MOBO approaches are often limited by sequential optimizations, which can only add one sample in each iteration. The parallel computing strategy is a desirable way to accelerate the optimization process by adding a batch of samples in one iteration. Unfortunately, existing parallel MOBO approaches are constrained to handling single-fidelity data, thereby missing out on the potential benefits of utilizing auxiliary low-fidelity data. To further improve the optimization efficiency of the parallel MOBO approach, this paper proposes two parallel MOBO approaches based on multi-fidelity surrogate modeling. Specially, the cheap auxiliary low-fidelity data can be utilized to improve the performance of the parallel MOBO approach by multi-fidelity surrogate modeling. The updating points and fidelity levels are determined by a modified hypervolume excepted improvement function, and two parallel computing strategies are developed for multi-point sampling. Additionally, a constraint handling strategy is developed for problems with constraints by adopting the probability of feasibility functions. The proposed approaches are demonstrated through numerical benchmark examples and two real-world applications involving the multi-objective optimizations of a micro-aerial vehicle fuselage and a metamaterial vibration isolator. Results in the real-world applications show that the proposed approaches significantly improve the optimization efficiency with faster convergence speed and exhibit superior overall performance compared with the state-of-the-art MOBO methods.

Design and development of an experimental bench concept for testing of fuselage stiffened panels, using a virtual testing methodology

PurposeFuselage structures are subjected to combinations of axial, bending, shear and differential pressure loads. The validation of advanced metallic and composite fuselage designs against such loads is based on the full-scale testing of the fuselage barrel, which, however, is highly demanding from a time and cost viewpoint. This paper aims to assist in scaling-down the experimentation to the stiffened panel level which presents the opportunity to validate state-of-the-art designs at higher rates than previously attainable.Design/methodology/approachDevelopment of a methodology to successfully design tests at the stiffened panel level and realize them using advanced, complex and adaptable test-rigs that are capable of introducing independently a set of distinct load types (e.g. internal overpressure, tension, shear) while applying appropriate boundary conditions at the edges of the stiffened panel.FindingsA baseline test-rig configuration was developed after extensive parametric modelling studies at the stiffened panel level. The realization of the loading and boundary conditions on the test-rig was facilitated through innovative supporting and loading system set-ups.Originality/valueThe proposed test bench is novel and compared to the conventional counterparts more viable from an economic and manufacturing point of view. It leads to panel responses, which are as close as possible to those of the fuselage barrel in-flight and can be used for the execution of static or fatigue tests on metallic and thermoplastic curved integrally stiffened full-scale panels, representative of a business jet fuselage.

Experimental and numerical analysis of composite fuselage panel impacted by tire burst debris under airworthiness clause

This study focuses on an experimental and numerical simulation investigation on the dynamic response of the composite fuselage panel under the impact of the tire burst debris. According to the parameters specified in the airworthiness clause, the T800/M21C composite material is selected as the research object to carry out the tire debris impact test, and the consistent results are obtained by employing the applicable mesoscopic model of the tire debris, progressive failure model of composite, the interface model and considering the strain rate dependence of the composite material in numerical simulation. The results show that under the airworthiness condition, the peak strain rate of the fuselage panel is about 16 s−1 during the impact of tire debris. The damage modes of the fuselage panel are debonding of the stringers from the skin and local matrix tensile damage. The Puck failure criterion, which considers the angle of the fracture surface of the matrix, is proved to be more suitable for the prediction of matrix damage than the Hashin failure criterion. The developed progressive damage failure impact model can be used in structural analysis and optimization design of aircraft fuselage, providing reference data for airworthiness verification.

Should We Care about how Birds Fly?

Bird flight has intrigued human minds ever since the Stone Age. Paintings in Lascaux cave in France are considered the oldest representation of birds in flight made by early ancestors of humans. The ability to fly has been interpreted as a divine power by some cultures which proceeded to include wings in drawings and sculptures of their deities. In the Renaissance period, Polymaths like Leonardo Da Vinci studied birds and their motion in detail. His codex of birds mentions the objective of this study as to develop a human-powered flying machine. Even, the Wright brothers used warping wings in their flight tests mimicking the flight patterns of birds. The wing and fuselage design of fixed-wing aircraft from the earliest versions to modern jet-powered airliners is directly inspired by the bio-mechanics of large birds. However, with the success of fixed-wing aircraft, the flapping wing concepts saw declining interest as a viable design option for air vehicles. During this time, the aerodynamics of bird wings was only studied in fundamental science to answer the questions on their aerodynamic performance. However, the popularization of Unmanned Arial Vehicles (UAVs) renewed the interest in bird-inspired air vehicles with flapping wings as a possible design.

Structural comparison of vertical and horizontal layout of carrying arms of rotary-wing UAV with finite element analysis

In this study, numerical analysis of the fuselage of a rotary wing unmanned aerial vehicle was conducted. A fuselage that is resistant to the loads on the fuselage and has maximum lightness has been designed. In this context, the fuselage design was conducted based on the loads that the aircraft's fuselage would be exposed to during landing and take-off, and a three-dimensional modeling was created. Numerical analyzes were carried out using the designed solid model finite element method. It has been observed that the obtained data can meet the loads on the airframe without any breakage in the specified configuration. The findings obtained at the end of the study were supported by graphics

Design and Aerodynamic Analysis of Fixed-Wing Vertical Take-Off Landing (FW-VTOL) UAV

This study aims to produce a Fixed Wing-Vertical Takeoff Landing (FW-VTOL) design that has good aerodynamic characteristics on its wing and fuselage. The design process begins with determining the Design Requirement and Objective (DRO). According to DRO, the aircraft is designed to be able to operate at speeds of 14 m/s and an altitude of 120 m (cruise phase) above ground level. The Maximum Takeoff Weight (MTOW) of the planned is estimated using previous research data. Based on calculations, for it to have a flight time of 30 minutes, the MTOW of the FW-VTOL is 5.2 kg. The fuselage and wing of the FW-VTOL design were carried out by referring to the MTOW of 5.2 kg. Three candidate airfoils were proposed, and their aerodynamic characteristics were determined using XFLR5 software. The aerodynamic characteristics of the three proposed fuselage candidates are also calculated. At an angle of attack of 30, the Eppler 66 airfoil produces the most lift coefficient and the lift-to-drag ratio of 0.8775 and 94.152, respectively. While the NACA 4412 and S9000 airfoils create lift coefficients of 0.7963 and 0.7963, respectively. From the calculation of aerodynamic characteristics, the first, second, and third fuselage candidates produce drag coefficient values of 0.1397, 0.1576, and 0.1368, respectively. The aerodynamic characteristics of the airfoil and fuselage are then input into the decision matrix table. As a result, the selected airfoil candidate is Eppler 66, while we choose the fuselage design is the third fuselage candidate.

Aeropropulsive Performance Analysis of Axisymmetric Fuselage Bodies for Boundary-Layer Ingestion Applications

Boundary-layer ingestion (BLI) has been proposed as one of the novel airframe–engine integration technologies to reduce aircraft fuel consumption. The current numerical analysis involves the evaluation of the effect of fuselage design on the power consumption of a boundary-layer ingesting propulsor modeled as an actuator volume without nacelle. An axisymmetric fuselage model is used as a canonical case to study BLI in transonic flight conditions. The flowfield is investigated through the power balance and the exergy analysis methods. Results show that the fuselage geometry and flight conditions only have a minor effect on the BLI power saving benefit when compared to the effect on the drag power of the fuselage. This indicates that, for the range of fuselage geometries and flight conditions studied, the isolated fuselage drag can be used for a qualitative performance assessment of different fuselage designs even for BLI configurations. Also, the power saving results obtained based on the power balance and the exergy analysis methods show similar qualitative trends for the fuselage geometries and flight conditions considered. Furthermore, the BLI propulsor has a negligible effect on the upstream anergy generation rate. Turbulence and temperature gradients within the flow are the important reasons for the deterioration of the BLI propulsor performance as expected.

Investigation of Wing, Fuselage and Tail Design Parameters in Boeing and Airbus Aircraft

The study aims to examine the wing, fuselage and tail design parameters of Airbus and Boeing aircraft and to reveal the design criteria for the sizing of large commercial jets. Tin the study, the design and sizing data of 11 Airbus models and 14 Boeing models were compiled. The number of abreast seats, maximum seats, fuselage length, fuselage width, fineness ratio, tail section length, cockpit length and cargo compartment length are the main parameters studied about the fuselage design. Wing area, wingspan, aspect ratio, mean aerodynamic chord (MAC), taper ratio, dihedral angle, quarter chord sweep angle, and winglet lengths are the main parameters studied about the wing design. The area, span, aspect ratio, taper ratio, and quarter chord sweep angle of vertical and horizontal tail are the main parameters studied about the tail design. As a result of the study, exceptional designs and average design criteria for Airbus and Boeing aircraft have been revealed with the help of charts presented. In addition, the obtained linear correlations reveal that the MAC has about 0.14 times the wingspan and about 1.168 times the average chord length.

Aircraft turnaround time estimation in early design phases: Simulation tools development and application to the case of box-wing architecture

This work deals with the problem of estimating the turnaround time in the early stages of aircraft design. The turnaround time has a significant impact in terms of marketability and value creation potential of an aircraft and, for this reason, it should be considered as an important driver of fuselage and cabin design decisions. Estimating the turnaround time during the early stages of aircraft design is therefore an essential task. This task becomes even more decisive when designers explore unconventional aircraft architectures or, in general, are still evaluating the fuselage design and its internal layout. In particular, it is of paramount importance to properly estimate the boarding and deboarding times, which contribute for up the 40% to the overall turnaround time. For this purpose, a tool, called SimBaD, has been developed and validated with publicly available data for existing aircraft of different classes. In order to demonstrate SimBaD capability of evaluating the influence of fuselage and cabin features on the turnaround time, its application to an unconventional box-wing aircraft architecture, known as PrandtlPlane, is presented as case study. Finally, considering standard scenarios provided by aircraft manufacturers, a comparison between the turnaround time of the PrandtlPlane and the turnaround time of a conventional competitor aircraft is presented.

Analysis of Fuselage Using Static and Thermal Interface for Different Joints

Fuselage forms the main structure of the aircraft that accommodates the passenger and cargo. The fuselage is continuously subjected to various loads during flight as well as after landing. The integrity of the fuselage-structure is very important for the safety of the aircraft. This project describes a conceptual design of fuselage structure by using CAD software. This project aims to design and analysis of fuselage using static and thermal interface for different joints (riveted, bonded, hybrid) loads for static and thermal stiffness. For this purpose, CAE simulations will be performed on a fuselage model using static and thermal conditions.

МЕТОД ІНТЕГРОВАНОГО ПРОЕКТУВАННЯ НОСОВОЇ ЧАСТИНИ ФЮЗЕЛЯЖУ ЛІТАКА ТРАНСПОРТНОЇ КАТЕГОРІЇ

Regulatory and technical documentation, design features and methods of fuselage transport category aircraft design was performed and identified the need to update design methods and calculate the characteristics of the fuselage using parametric models and integrated design systems CAD / CAM / CAE / PLM. The method of integrated fuselage design of transport category aircraft is developed and theoretically substantiated. Within the framework of the proposed method, parametric models of master geometry, aerodynamic flow and mass-inertial characteristics of the fuselage were created, taking into account the design features of transport aircraft.The proposed method was used to study the influence of geometric parameters of fuselage nose section on aerodynamic and mass characteristics of the fuselage, showing the efficiency of work with parametric models.The choice of parameters of the fuselage nose section in preliminary and sketch design of a promising aircraft for local airlines is justified, which allowed to implement and test the suitability of the proposed method for in new competitive aircraft designing process.The use of the method for integrated fuselage design for local aircraft allowed to determine the rational configuration of the nose section of the fuselage and increase the fuel efficiency of the aircraft by 6.4%, reduce the aerodynamic drag of the fuselage by 10%, increase the viewing angle from the cockpit by 10%. and ensure compliance with current regulatory and technical documentation, as well as determine the mass-inertial characteristics of the fuselage and its parts and form a list of cockpit equipment that will meet flight safety requirements, taking into account the operating conditions and modifications of the aircraft.The configuration of the nose section of the aircraft fuselage for local airlines has been developed, which allows to fit modern requirements for cockpit equipment and layout, low fuselage impedance and high aerodynamic quality and fuel efficiency in cruising mode at speeds up to 850 km / h (M = 0, 8). As a result of verification using other methods and parameters of existing aircraft, the accuracy of the results obtained using the proposed method at the level of 5% was confirmed.

Design of a Four-Seat, General Aviation Electric Aircraft

Financial and environmental considerations continue to encourage aircraft manufacturers to consider alternate forms of aircraft propulsion. On the financial end, it is the continued rise in aviation fuel prices, as a result of an increasing demand for air travel, and the depletion of fossil fuel resources; on the environmental end, it is concerns related to air pollution and global warming. New aircraft designs are being proposed using electrical and hybrid propulsion systems, as a way of tackling both the financial and environmental challenges associated with the continued use of fossil fuels. While battery capabilities are evolving rapidly, the current state-of-the-art offers an energy density of ~ 250 Wh/kg. This is sufficient for small, general aviation electric airplanes, with a modest range no more than 200 km. This paper explores the possibility of a medium range (750 km) electric, four-seat, FAR-23 certifiable general aviation aircraft, assuming an energy density of 1500 Wh/kg, projected to be available in 2025. It presents the conceptual and preliminary design of such an aircraft, which includes weight and performance sizing, fuselage design, wing and high-lift system design, empennage design, landing gear design, weight and balance, stability and control analysis, drag polar estimation, environmental impact and final specifications. The results indicate that such an aircraft is indeed feasible, promising greener general aviation fleets around the world. Keywords: general aviation aircraft, electric aircraft, aircraft design

복합형 회전익 항공기 동체 설계를 위한 확장된 보 해석

복합형 회전익 항공기 동체 설계를 위한 확장된 보 해석

Crash simulation of fuselage section in the rebound process and the secondary-impact process

PurposeThis paper aims to investigate the rebound process and the secondary-impact process of the fuselage section that occurs in the actual crash events.Design/methodology/approachA full-scale three-dimensional finite element model of the fuselage section was developed to carry out the dynamic simulations. The rebound process was simulated by removing the impact surface at a certain point, while the secondary-impact process was simulated by striking the impact surface against the fuselage bottom after the first impact.FindingsFor the rebound process, the fuselage structure restores deformation due to the springback of the fuselage bottom, and it results in structural vibration of the fuselage section. For the secondary-impact process, the fuselage deformation is similar with that of the single impact process, indicating that the intermittent impact loading has little influence on the overall deformation of the fuselage section. The strut failure is the determining factor to the acceleration responses for both the rebound process and the secondary-impact process.Practical implicationsThe rebound process and the secondary-impact process, which is difficult to study by experiments, was investigated by finite element simulations. The structure deformations and acceleration responses were obtained, and they can provide guidance for the crashworthy design of fuselage structures.Originality/valueThis research first investigated the rebound process and the secondary-impact process of the fuselage section. The absence of the ground load and the secondary-impact was simulated by controlling the impact surface, which is a new simulating method and has not been used in the previous research.