Aircraft Wing Design Research Articles

Compliant mechanism design of combined aircraft wing for stable separation – CORRIGENDUM

Compliant mechanism design of combined aircraft wing for stable separation – CORRIGENDUM

Investigation of Hybrid Laminar Flow Control Capabilities from the Flight Envelope Perspective

The present work conducts a multifidelity aerodynamic design of a medium-range commercial aircraft wing featuring hybrid laminar flow control (HLFC) technology. The design was performed to investigate the capabilities of the HLFC, considering the entire design flight envelope, low-speed performance requirements, and various potential mission scenarios. A combination of medium-fidelity aerodynamic design tools for laminar flow control using MSES-COCO-LILO, CFD RANS for high-lift cases, the quasi-three-dimensional for database generations, and SUAVE for mission analysis were used to design and assess aircraft’s performance with the HLFC wing. Studies of the design objective demonstrated a conflict between friction and compressibility drag that strongly affects the design objective function and the wing performance at various off-design conditions. A strong dependence of HLFC on flight conditions was observed to indicate technology performance limitations and a tradeoff between airplane emissions, range, and costs. Finally, a comparison of medium- and low-fidelity performance results showed deviations in the design fuel burn of 4%, a minimum possible harmonic range difference of 3%, and direct operating costs of 2%. The outcome suggests the need for better compressibility and parasite drag models for early design stages to accurately model the HLFC technology without introducing complicated methods at the first sizing steps.

Design Optimization of Cessna 172 Wing With Biomimetic Design Approach

While the aircraft is moving in the opposite direction of the flow in the air, the wind resistance and the moment effect due to this resistance negatively affect the flight performance. In design studies, it is aimed to increase aerodynamic performance by minimizing these two negative factors. In this study, the effect of wing cross section and three-dimensional airfoil on aerodynamic performance is investigated numerically. Within the scope of this study, to reach the intended design, biomimetic design approach was used and new-wing designs were created to mimic the bird species’ wings, which have the highest aerodynamic performance in nature. Based on the literature search for two-dimensional wing section selection, it was seen that the most preferred sections were identified and compared with the wing section of the Cessna 172 aircraft (NACA2412) in flow analysis. In the flow analysis conducted in the XFLR5 program, the aerodynamic performances of the wing sections at Reynolds value and angle of attack were investigated. According to this analysis, the aerodynamic efficiency of the NACA2412 section was higher than that of the other sections. In the three-dimensional flow analysis, biomimetic wing designs and the wing of the Cessna 172 aircraft were examined in the XFLR5 program at a cruising speed and angle of attack. It was observed that the aerodynamic efficiency of the wing design, which is inspired by the albatross, is higher than the other designs. Owing to the flow analysis, the albatross wing design provided 6.26% improvement in the lift coefficient, 15.73% in drag coefficient and 15.16% improvement in the glide ratio compared to the Cessna 172 aircraft wing design. For structural analysis, the pressure values obtained from the flow analysis results were used as the load distribution on the wing. In the designs created using the same material, it was observed that the weight of the wing inspired by the albatross was 34.156% less weight and 50.902% less deformation value obtained compared to Cessna 172 aircraft.

Compliant mechanism design of combined aircraft wing for stable separation

Abstract Stable separation is a crucial condition that must be met in order for combined aircraft to successfully engage in cooperative flight. In order to achieve the desired fast and controlled separation, this paper proposes a novel design for a torque-driven compliant separation mechanism. By taking into account the compliance characteristics of a sinusoidal acceleration function curve, a mechanical model for the separation mechanism is developed. By utilising the Coulomb friction law, an accurate determination of the aerodynamic load distribution under various conditions is achieved. Subsequently, the relationship between the unlocking moment and the aerodynamic load is derived based on these findings. Through the utilisation of the finite element method, a model of the separation mechanism is generated. To ensure the safety and reliability of the compliant separation mechanism, the mechanical properties of the structural materials are thoroughly analysed under the maximum aerodynamic load. Subsequently, the separation mechanism structure is constructed and subjected to testing in order to showcase the compliance characteristics. In addition, this paper conducts a simulation to analyse the impact of flight speed and angle-of-attack on the separation process. By doing so, the optimal conditions for separation are determined. The methods and findings presented in this study have the potential to contribute valuable insights to the design of combined aircraft.

Effects of structural geometric nonlinearities on the transonic aeroelastic characteristics of wing

The structural geometric nonlinearity has a great influence on the transonic aeroelastic characteristics of aircraft wings, which is an important problem for the aerodynamic and structural design of wing. In this paper, a CFD/CSD coupling method based on the high precision in-house solver is first established, and the reliability of the CFD and the CFD/CSD coupling method is then verified. Subsequently, the effects of geometric nonlinearity on the transonic aeroelastic characteristics of wing under different dynamic pressures are studied. The results show that: (1) At low dynamic pressure conditions, for geometric linear wing structure, limit-cycle oscillation occurs at various angles of attack, and the limit-cycle oscillation amplitude decreases with the increases of angle of attack. At low angle of attack, the wing deformation characteristics are dominated by first bending and first torsion, and the first torsion weaken as the angle of attack increases. For the geometric nonlinear wing structure, the limit-cycle oscillation occurs only at an attack angle of 0° As the angle of attack increases, the structural displacement eventually exhibits static deformation characteristics. (2) Under high dynamic pressure conditions: the displacement of geometric linear wing structure presents a divergent trend. For geometric nonlinear wing structure, the limit-cycle oscillation amplitude decreases first and then increases with the increase of the angle of attack. At low angle of attack, the deformation characteristics of the wing are dominated by first bending and first torsion. With the increase of the angle of attack, the first torsion weakens, while the second bending strengthen. With the further increase of the angle of attack, the second bending weaken and become co-dominated by the first bending and first torsion. It can be seen from the above results that the aeroelastic characteristics of both the geometric linear structure and the geometric nonlinear structure wings are significantly different. These findings provide valuable insights for the meticulous design of aircraft wings.

Spectral Properties of Mimetic Operators for Robust Fluid–Structure Interaction in the Design of Aircraft Wings

This paper presents a comprehensive study on the spectral properties of mimetic finite-difference operators and their application in the robust fluid–structure interaction (FSI) analysis of aircraft wings under uncertain operating conditions. By delving into the eigenvalue behavior of mimetic Laplacian operators and extending the analysis to stochastic settings, we develop a novel stochastic mimetic framework tailored for addressing uncertainties inherent in the fluid dynamics and structural mechanics of aircraft wings. The framework integrates random matrix theory with mimetic discretization methods, enabling the incorporation of uncertainties in fluid properties, structural parameters, and coupling conditions at the fluid–structure interface. Through spectral and localization analysis of the coupled stochastic mimetic operator, we assess the system’s stability, sensitivity to perturbations, and computational efficiency. Our results highlight the potential of the stochastic mimetic approach for enhancing reliability and robustness in the design of aircraft wings, paving the way for optimization algorithms that integrate uncertainties directly into the design process. Our findings reveal a significant impact of stochastic perturbations on the spectral radius and eigenfunction localization, indicating heightened system sensitivity. The introduction of randomized singular value decomposition (RSVD) within our framework not only enhances computational efficiency but also preserves accuracy in low-rank approximations, which is critical for handling large-scale systems. Moreover, Monte Carlo simulations validate the robustness of our stochastic mimetic framework, showcasing its efficacy in capturing the nuanced dynamics of FSI under uncertainty. This study contributes to the fields of numerical methods and aerospace engineering by offering a rigorous and scalable approach for conducting uncertainty-aware FSI analysis, which is crucial for the development of safer and more efficient aircraft.

The Amount of Circulation when Forming the Aerodynamic Appearance of a Main Aircraft

The work proposed by the authors describes the mechanism for the formation of induced drag for a wing of finite span, as well as a possible option for taking into account the circulation value as a control parameter in the process of preliminary design of a mainline aircraft wing equipped with additional aerodynamic surfaces (wing tips, winglets).Providing the least frictional resistance is achieved by the smoothness of the wing surface in combination with laminar profiles (with the greatest length of the laminar section of the boundary layer), when critical values of the M number are reached, supersonic zones appear on the wing, therefore it is necessary to delay the occurrence of the wave crisis as much as possible - apply the sliding effect for the swept wing of the final scope. A reduction in inductive drag can be achieved by using a high aspect ratio wing or various types of wing tips, which make it possible to control the amount of circulation and localize the process of pressure flow near the end part of a wing of finite span by increasing the effective span. The study of the flow around an aerodynamic model of a mainline aircraft formed the basis of the conceptual model and mathematical support that implements the methodology for selecting the composition of rational design parameters. The identified experimental data and functional dependencies made it possible to form a vector of parameters for a formal aerodynamic model, connecting the geometry of the aircraft model (shape) and its aerodynamic characteristics through dimensionless coefficients for various operating characteristics. Analysis of the results and subsequent processing in order to synthesize a design solution on a validation basis involves the use of a formalized multi-parameter approach, i.e. machine (computer) processing is required. The implementation of a multi-parameter approach in application software in the C++ programming language will allow us to explore and establish connections between parameters and obtain a new version of the design alternative using a smaller amount of data.

Aircraft Wing Design for Extended Hybrid Laminar Flow Control

Laminar flow offers significant potential for increasing the energy-efficiency of future transport aircraft. The German Cluster of Excellence SE2A is developing a new approach for hybrid laminar flow control. The concept aims to maintain laminar flow up to 80% of the chord length by integrating suction panels at the rear part of the wing, which consist of a thin suction skin and a supporting core structure. This study examines effects of various suction panel configurations on wing mass and load transfer for an all-electric short-range aircraft. Suction panel material, as well as thickness and relative density of the suction panel core are modified in meaningful boundaries. Suction panels made from Ti6Al4V offer the most robust design resulting in a significant increase in wing mass. For the studied configurations, they represent up to 33.8% of the mass of the wingbox. In contrast, panels made from Nylon11CF or PU1000 do not significantly increase the wing mass. However, the use of these materials raises questions about their robustness under operational conditions. The results demonstrate that the choice of material strongly influences the load path within the wing structure. Ti6Al4V suction panels provide sufficient mechanical properties to significantly contribute to load transfer and buckling stiffness. Locally, the share of load transfer attributed to the suction panel exceeds 50%. In contrast, compliant materials such as Nylon11CF or PU1000 are inherently decoupled from load transfer. Unlike the thickness of the suction skin, the relative density of the core structure strongly affects the wrinkling stiffness. However, wrinkling failure did not appear critical for the examined suction panel configurations. In the present study, the mechanical properties of Ti6Al4V cannot fully be exploited. Therefore, compliant suction panels made from Nylon11CF are preferred in order to achieve a lightweight solution, provided that they meet operational requirements.

Static analysis of shear-deformable aircraft wings using a multilayered functionally graded material model

This study presents a comprehensive investigation of the static analysis of a shear-deformable aircraft wing made of functionally graded materials (FGMs). A finite element-based multilayered FGM model is employed for this purpose. In order to ensure the safe structural design of the aircraft wing, it is crucial to analyze the transverse displacement of the wing under various loading conditions. The presented model employs a power-law distribution based on the rule of mixture to derive the effective material properties of the FGM wing. The analysis focuses on three standard FGM aircraft wing profiles, namely NACA 0009, NACA 2424, and NACA 4415, which are representative of commonly used geometries in aircraft design. Additionally, the study explores the effects of volume fraction index, loading conditions, boundary conditions, and aspect ratio on the static analysis of the shear-deformable FGM aircraft wing. These parameters play a significant role in shaping the static behavior of the wings and offer valuable insights into the design of FGM aircraft wings.

A co-rotational formulation for quasi-steady aerodynamic nonlinear analysis of frame structures

The design of structures submitted to aerodynamic loads usually requires the development of specific computational models considering fluid-structure interactions. Models using structural frame elements are developed in several relevant applications such as, the design of advanced aircraft wings, wind turbine blades or power transmission lines. In the case of flexible frame structures submitted to fluid flows, the computation of inertial and aerodynamic forces for large displacements and rotations is a challenging task. In this article, we present a novel formulation for the efficient computation of aerodynamic forces in frame structures, coupling the co-rotational framework with the quasi-steady theory. A numerical procedure is provided considering a tangent matrix for the aerodynamic forces. This formulation is implemented in the open-source library ONSAS, allowing users to reproduce the results or solve other frame nonlinear dynamic problems. The proposed formulation and its implementation are validated through the resolution of four numerical examples. The formulation and the numerical procedure proposed efficiently provide accurate solutions for these challenging problems with large displacements and rotations.

Evolutionary generative design of supercritical airfoils: an automated approach driven by small data

Supercritical airfoils are critical components in the design of commercial wide-body aircraft wings due to their ability to enhance aerodynamic performance in transonic flow regimes. However, traditional design methods for supercritical airfoils can be time-consuming and require significant manual effort, not to mention the high cost associated with computational fluid dynamics analysis. To address these challenges, this paper introduces a highly automated approach for supercritical airfoil design, called Evolutionary Generative Design (EvoGD). The EvoGD approach is based on the framework of Evolutionary Computation and employs a series of sophisticated data-driven generative models incorporated with physical information to iteratively refine initial airfoil shapes, resulting in improved aerodynamic performances and reduced constraint violations. Moreover, to speed up the evaluation of the generated airfoils, a series of accurate and efficient data-driven predictors are utilized. The efficacy of the EvoGD approach was demonstrated through experiments on a dataset of 501 supercritical airfoils, including one baseline design and 500 randomly perturbed airfoils. On average, the generated airfoils showed improved performance in terms of buffet lift coefficient, cruise lift-to-drag ratio, and thickness by 5%, 4%, and 1%, respectively. The best generated airfoil outperformed the baseline design in terms of critical buffet lift coefficient and cruise lift-to-drag ratio by 7.1% and 6.4%, respectively. The entire design process was completed in less than an hour on a personal computer, highlighting the high efficiency and scalability of the EvoGD approach.

Topology Optimization of Aircraft Wing

The objective of this paper is to obtain an optimal concept design of aircraft wing. Topology optimization is used to yield an optimized shape and material distribution for a set of loads and constraints within a given design space. Initially a solid wing was considered as design space and meshed using hexahedral elements to derive the optimal location of spars and ribs such that the structure has minimum compliance. Load data includes flight, fuel and engine loads. The wing was constrained in all degrees of freedom at the root. It was shown that the results can be used to derive an appropriate configuration similar to the existing design practices. A new geometry with spars, ribs, hinge locations and lug attachments was next considered. The skin was meshed with shell elements. Hinge loads were also included. The optimization gave a box section for front spar, channel section for rear spar, truss pattern for ribs and V shaped member for hinges. Using this global material distribution a component level topology optimization was performed. The concept design is realized in to CAD and then a composite free-size optimization is performed to determine material distribution and ply drop regions etc. Finally a composite size and shape optimization is done.

Fuel sloshing-induced effects on the dynamic response of a scaled research wing demonstrator

Fuel sloshing inside aircraft wing tanks is not currently considered as a means of passive loads alleviation due to the lack of maturity of modelling tools and overall understanding of vertical sloshing. In this work, a scaled research wing demonstrator is presented, designed for the study of fuel sloshing inside an aircraft wing-representative model. An experimental campaign was conducted to examine the impact of fuel sloshing on the vertical dynamics of a scaled aircraft wing demonstrator, investigating the influence of excitation amplitude, filling level, baffle geometry, spanwise liquid position, and dihedral angle. It was found that the sloshing of the liquid in the fuel tank causes significant energy dissipation following step release excitation, maximized in the 40-60% filling range depending on the dihedral angle. The sloshing-induced damping also varies with the amplitude of excitation, with maximum added damping ratio of 0.038 achieved earlier in the motion at larger excitation amplitudes. The spanwise distribution of the liquid was found to have the strongest effect on the sloshing-induced damping, with the baffles' main impact being through the compartmentalization of the tank. Solid baffles that maintain favorable liquid distribution can increase the damping effect, while perforated baffles are not as effective. This work indicates that by considering the fuel dynamics rather than the fuel equivalent dry mass during the aircraft wing design, the net damping may be increased, leading to lighter wing structure designs.

Probabilistic approach to safety factor evaluation for aircraft wing design

Probabilistic approach to safety factor evaluation for aircraft wing design

Multidisciplinary design of aircraft wing for multiregime purposes

In this paper optimization process of the geometrical design of the wing is described. The goal is to minimize the aerodynamic drag at a requested lift in the cruise regime and thus reduce the fuel consumption. The second goal is to ensure the requested landing regime of the aircraft. Both cases are solved under the wing structural and weight considerations. As a fast design computational tool, a lifting line theory is used. It is moreover supplemented by non-linear aerodynamic airfoil characteristics to provide more accurate results of aerodynamic drag of the wing. These airfoil data were calculated with the CFD and enables to take significant accuracy between low and high-fidelity method which is verified in previous work. The list of requirements is set for aircraft for 40 passengers with a range of 2,500 km at a speed of 438 km/h. As the baseline the L-610 aircraft is selected. Optimized wing geometry in verified in cruise and landing regime/configuration with CFD solver. From the calculation of the weight of the wing and CFD results, the flight performance is calculated for the desired flight mission.

Decomposition of Mathematical Programming Models for Aircraft Wing Design Facilitating the Use of Dynamic Programming Approach

Decomposition of Mathematical Programming Models for Aircraft Wing Design Facilitating the Use of Dynamic Programming Approach

Pre-buckling vibration and buckling analyses of composite skew plate: An analytical investigation

In this work, the analytical investigation for pre-buckling vibration and buckling analyses of a composite skew plate subjected to parabolically and linearly varying in-plane edge load are presented. When the composite skew plate is subjected to parabolic (non-uniform) in-plane edge loading, the pre-buckling stresses within the composite skew plate are not known a priori. To estimate the pre-buckling stresses within the composite skew plate for which the in-plane elasticity problem is solved by minimizing the membrane strain energy using the Ritz method. It is observed that the rate of diffusion of applied parabolic in-plane edge load within the skew plate to a state of uniform in-plane stress is faster in the case of isotropic material than composite material. Using estimated pre-buckling stresses, the total energy functional is derived from the total strain energy, potential energy, and kinetic energy. The total energy functional is reduced into sets of an ordinary differential equation and algebraic equation, respectively for pre-buckling vibration and buckling problems using the Ritz method in conjunction with BCOPs. The associated linear eigenvalue problems are solved to compute the pre-buckling vibration frequency and buckling load of the stressed skew plate. The outcome of the study may provide crucial inputs in the design of skewed bridge decks, ship structures, and aircraft wing design.

Design And Analysis of Aircraft Wing.

This research develops and analyses a general aviation plane. A preliminary representation of the eventual product is used to start the design process for an airplane. Based on a drawing, a design mission profile is utilized to determine the weight. A more advanced approach is used to estimate weight, which employs calculated performance criteria to produce a more exact weight estimate. The wing design has been demonstrated to be a feasible alternative for a similar general aviation aircraft. When traveling through air or other fluids, a wing is a type of fin that provides lift. Airfoils may be observed on the wings, which have a streamlined cross-section, as a result of this. The lift created by a wing is compared to the drag generated by it to determine its aerodynamic efficiency.

Effects of fiber properties on aerodynamic performance and structural sizing of composite aircraft wings

An advanced carbon fiber is expected to reduce the structural weight of a composite aircraft. Historically, various studies on composite aircraft designs have focused on the challenge to utilize the anisotropy of unidirectional laminates. However, there is still a research interest to consider the impact of changing the fibers for their positive influence on the structural weight and performances of a composite aircraft. In this study, the effects of fiber properties on the static aeroelastic design of aircraft wings were investigated using a multi-scale framework and two-way coupled aeroelastic analysis with structural sizing. The fibers analyzed were T700S, T800S, and T1100G, which have different stiffnesses and strengths. The design analysis was performed under two scenarios: fixing the angle of attack and fixing the lift while changing the angle of attack in the flow condition. The results from the former case indicate that the choice of fiber did not significantly affect the entire shape of the drag polar, while the highest lift coefficient was obtained with the stiffest fiber, T1100G, under the same angle of attack. The results from the latter case indicate that using T1100G can reduce the structural weight by 9.9% and 14% at the same lift compared to T800S and T700S, respectively, owing to its high buckling resistance and high tensile strength, rather than compressive strength.

Design and analysis of variable camber wing of propeller aircraft using the actuator disc method

AbstractVariable camber flap technology can adjust the spanwise circulation distribution, thereby reducing the induced drag. Therefore, the concept of variable camber flap is introduced into the design of propeller aircraft wing, and the design for drag reduction of propeller aircraft is carried out. The numerical simulation of the propeller aircraft is carried out by using the actuator disc method with non-uniform distribution of radial and circumferential loads. Through the unsteady simulation of a single propeller, the aerodynamic load on a periodic propeller is extracted as a boundary condition to the steady simulation of the full aircraft. The load extracted by the actuator disc is compared with the unsteady simulation result, which verifies the reliability of the method. The design for drag reduction at cruise and climb design conditions are respectively carried out with the variable camber flap technology. The variable camber cruise configuration is evaluated at both the begin and end cruise conditions. The results show that, after the flaps deflecting at a small angle according to the circulation distribution, the camber distribution of the wing is adjusted to make the circulation distribution closer to the elliptical circulation distribution. At the design cruise condition, the drag coefficient is reduced by 1.4 counts, and the lift-drag ratio increase by 0.1. At both begin and end cruise conditions, the drag coefficient decreases by 1 count, and the lift-drag ratio increases by 0.07. At the design climb condition, the drag coefficient decreases by 1 count, and the lift-to-drag ratio increases by 0.09.