Design Of Aircraft Structures Research Articles

Experimental Tests and Numerical Analysis of Aerodynamic Properties of the Composite-Made Jet-Propelled Aerial Target.

The design of an aircraft's internal structure, and therefore the appropriate choice of material type, is a direct function of the performed tasks and the magnitude and type of the acting loads. The design of a durable aircraft structure with appropriate stiffness and lightness requires knowledge of the loads that will be applied to the structure. Therefore, this paper presents the results of an aerodynamic experimental test and numerical analysis of a newly designed jet-propelled aerial target. The experimental tests were carried out in a low-speed wind tunnel for a wide range of angles of attack and sideslips. Moreover, they were performed for various configurations of the airplane model. In addition, the results of the experimental test were supplemented with the results of the numerical analysis performed using computational fluid dynamics methods. During numerical analysis, specialized software based on solving partial differential equations using the Finite Volumes Method was used. This article presents the methodology of the conducted research. The results of the aerodynamic analysis are presented in the form of diagrams showing the aerodynamic force and moment components as a function of the angle of attack and sideslip. In addition, qualitative results of the flow around the plane have been presented. The results obtained prove that the adopted methods are sufficient to solve these types of problem. The aerial system was positively verified during the qualification tests of the system at the Polish Air Force training range and finally received the type certificate.

Optimization of composite aeronautical components by Re-designing with double-double laminates

This paper introduces a novel approach to optimize aircraft structural design for enhanced performance and environmental sustainability. Traditional constraints on composite laminate stacking sequences, such as the symmetrical and balanced nature of classical quasi-isotropic laminates, often lead to over-dimensioning of aeronautical components. To overcome these limitations, the adoption of a new class of laminates known as Double-Double laminates, originally introduced by Prof. S.W. Tsai, is proposed. Unlike traditional laminates, Double-Double laminates offer flexibility in layer orientations, small building blocks, easy homogenization, and tapering capability.This study presents various case studies of increasing complexity to assess the effectiveness of the proposed methodology. First, a multicopter frame is redesigned to reveal the methodology's ability to reduce component weight while ensuring strength requirements under typical service loads. Then, the “DD Automated Design Tool”, a newly developed Finite Element Method tool interfacing with the Lamsearch engine, is introduced. This tool facilitates the rapid identification of optimal composite double-double laminate layups based on material properties, applied loads, and boundary conditions. By iteratively interacting between the Finite Element Method software and the Lamsearch engine, the tool efficiently identifies the best laminate angles to meet specific strength requirements.To validate the weight-saving capability of the proposed methodology, the optimizations on a composite fuselage section and a real composite wing-box under realistic loading conditions have been carried out. Both cases resulted in significant weight reductions compared to initial configurations, demonstrating the effectiveness of Double-Double laminates in achieving structural efficiency and meeting performance demands.

Experimental study on the influence of welding structure details of TC4 titanium alloy under thermo-vibration coupling environment on vibration fatigue life

The widespread application of titanium alloy materials in aerospace structural design makes them susceptible to thermo-vibration coupling fatigue failure in harsh vibration and thermal environments. Vibration fatigue issues significantly affect the safety and operational reliability of aircraft. In order to obtain optimal design details for titanium alloy welding structures under thermo-vibration coupling conditions, experimental specimens with different structural details of TC4 titanium alloy welding structures were designed to simulate actual structural designs. The structural details included skin with U-shaped supports at different angles, skin with L-shaped supports, and skin with T-shaped supports. A thermo-vibration coupling vibration fatigue test system was established, and vibration fatigue tests were conducted at a temperature of 150°C. All structural fatigue crack failures occurred at the welds. The experimental results were statistically analysed, and using the Chauvinet criterion, abnormal data points were removed. The influence of structural details on vibration fatigue life was determined through F-tests. It was found that for the skin with a U-shaped support structure, a support angle of 90° resulted in a longer vibration fatigue life. When comparing skin with T-shaped support and skin with L-shaped support structures, the T-shaped support structure exhibited a longer vibration fatigue life. Among the three structural types, the skin with a T-shaped support structure has the longest vibration fatigue life. These experimental conclusions serve as important references for the design of thin-walled structures in aircraft structural design.

Aircraft Structural Design and Life-Cycle Assessment through Digital Twins

Numerical modeling tools are essential in aircraft structural design, yet they face challenges in accurately reflecting real-world behavior due to factors like material properties scatter and manufacturing-induced deviations. This article addresses the potential impact of digital twins on overcoming these limitations and enhancing model reliability through advanced updating techniques based on machine learning. Digital twins, which are virtual replicas of physical systems, offer a promising solution by integrating sensor data, operational inputs, and historical records. Machine learning techniques enable the calibration and validation of models, combining experimental inputs with simulations through continuous updating processes that refine digital twins, improving their accuracy in predicting structural behavior and performance throughout an aircraft’s life cycle. These refined models enable real-time monitoring and precise damage assessment, supporting decision making in diverse contexts. By integrating sensor data and updating techniques, digital twins contribute to improved design and maintenance operations by providing valuable insights into structural health, safety, and reliability. Ultimately, this approach leads to more efficient and safer aviation operations, demonstrating the potential of digital twins to revolutionize aircraft structural analysis and design. This article explores various advancements and methodologies applicable to structural assessment, leveraging machine learning tools. These include the utilization of physics-informed neural networks, which enable the handling of diverse uncertainties. Such approaches empower a more informed and adaptive strategy, contributing to the assurance of structural integrity and safety in aircraft structures throughout their operational life.

Evaluation of measurement uncertainty of drop hammer impact test for composite materials

Future aircraft structures are usually dominated by composites, and one of the key defects and forms of damage to composite structures is the impact damage caused by low-energy (especially low-speed) foreign matter. This kind of low-energy impact test is usually completed by the drop hammer testing machine, so accurately assessing the impact test of the falling hammer completed in accordance with the standard test method, and the reliability of the permissible value of the impact performance of the objective reaction composite is of great significance to the future aircraft structure design. Taking a standard composite test piece as an example, considering the key factors in the whole process of impact test, a mathematical model is constructed, and the measurement uncertainty of different configuration composite test pieces of different configurations is calculated according to the test data, which provides a basis for aircraft structural design.

Model-based approach for the automatic inclusion of production considerations in the conceptual design of aircraft structures

Including production considerations in the early design stages of aircraft structures is challenging. Production information is mostly known by experts and rarely formally documented such that it can be effectively used during the design process. Producibility is mostly considered after completing the design, resulting in increased cost and development time due to the late discovery of production issues. This paper presents a new model, called the Manufacturing Information Model (MIM), which supports the automatic inclusion of production considerations into the design process. The MIM provides a single source of truth and a generic structure to capture and organize production-related information in a product system. Furthermore, it provides compatibility analyses to automatically warn for or exclude infeasible designs. Analysis tools use the information stored within the MIM to calculate the mass, costs, and production rate of the product. To show the functionalities of the MIM, it has been applied to the conceptual design of a wing box at a Tier 1 company. This use case shows how the MIM supports trade-off decisions, as it allows for the identification of trends and the ranking of different manufacturing concepts. Overall, the MIM provides a structured and formal approach to include production information in the conceptual design, improving the decision-making process.

Structural design and material comparison for aircraft wing box beam panel

This paper aims to present a comprehensive investigation to obtain the structural calculations needed to design a rigid panel of aluminum alloy for the wing box beam of an ATR 72–500 aircraft. For this design process, several types of materials, including composites like CFRP, are considered so it is possible to compare the actual existing part made of aluminum to them, thus checking the advantages these new materials offer. The research presents an introduction to structural design and provides a study of the relevant literature. The aircraft's principal characteristics and performance abilities were collected so that structural loads can be computed. Research used several methods, a design using conventional methods, applying the theory of elasticity is performed using the Theory of Farrar, allowing us to obtain an analytical solution to the problem, followed by checking the obtained results using Ansys FEM software combined with the parts being designed with CATIA. Furthermore, this same panel is calculated using composite materials instead of conventional aluminum, allowing us to compare both solutions. This research shed light on the intricate process of aircraft structural design, materials selection, and calculation methodologies, highlighting the ongoing pursuit of new and advanced materials. This paper makes clear that using composite materials presents several advantages over traditional ones, allowing for lighter, safer, more fuel-efficient, and more sustainable aircraft. The use of composite materials in the construction of airplane structures is driven by many factors. The results show that the chosen composite materials reduce weight, are durable, have low maintenance requirements, reduce noise, enhance fuel economy, and are resistant to corrosion.

Detection of cracks in aircraft structures using digital image processing technology

Full-scale aircraft structural fatigue testing exposes structural weaknesses by simulating the use of the aircraft in service and monitoring the development and expansion of fatigue cracks to guide the optimisation of aircraft structural design, the determination of service life and the development of maintenance and repair programmes. If crack damage is not detected in time, the uncontrolled expansion of the crack may lead to structural damage and prolong the test cycle. Therefore, it is crucial to achieve accurate detection of early cracks in fatigue testing of aircraft structures. Digital image technology is a new technology supported by a number of technical and theoretical achievements. In practical application, it relies on the support of computer functions, by converting image information into two-dimensional digital signals, thus realising various operations. In this paper, digital image processing techniques are used to deepen the backbone feature extraction network structure and improve the prediction scale of small targets, thus improving the performance of the model in terms of feature extraction and small defect detection. Experimental results show that the method can achieve accurate detection of small cracks, providing an effective technical approach for the automatic detection of cracks in aircraft structural fatigue tests. The technique of detecting cracks in aircraft structures using digital image processing also provides a new idea for research into improving aircraft endurance testing, which can be further developed with the aid of image technology.

Structural Optimization of Horizontal Stabilizer Tailing Edge Hinge Rib

Aircraft structural design requires lightweight design on the premise of meeting various performances. The number of hinge ribs in the aircraft trailing edge compartment is large, so it is very important to optimize the structure to achieve lightweight design. In this paper, the topology optimization algorithm is used to optimize the hinge ribs of the trailing edge cabin, and the secondary design is carried out according to the topology optimization results to achieve lightweight. The optimized rear edge cabin hinge rib can meet the requirements of strength, rigidity, process, manufacturing, etc., and the weight is greatly reduced compared with before optimization.

Experimental and numerical research on the analysis methods for buckling and post-buckling of inclined stiffened panel under shear load

The purpose of this paper is to investigate the buckling and post-buckling behavior of inclined stiffened composite panels under shear load. Two types of stiffened composite panels were manufactured and tested. The buckling load, ultimate load and other required structural properties of the inclined stiffened composite panels were obtained. The test results show that the failure load of the two types of panels are 5.79 times and 11.2 times of the buckling load, respectively, indicating that such inclined stiffened panels have a large post-buckling capacity. The damage modes of the panel are tearing around the skin near the clamps and debonding of the stringers from the skin. The combination of Hashin failure criterion, Cohesive model and nonlinear connection of clamp was used to conduct finite element modeling, which effectively predicts the buckling, post-buckling and damage modes of the structure. The buckling and post-buckling behavior of stiffened panels with different stinger configurations were discussed based on finite element analysis. The results show that when the initial buckling occurs between the stringers, the buckling and failure load of the stiffened panel become larger as the equivalent skin thickness between the stringers become larger. The research in this paper can provide reference for aircraft structural design.

Impact analysis of compression preloaded honeycomb sandwich structures

Honeycomb cores made of phenol resin impregnated paper are widely used in sandwich structures of aircraft. Owing to the rather weak core material, this kind of sandwich is prone to a range of damages as a result of impact loading which may accidentally occur during operation. When employed in aircraft these structures have to be damage tolerant. This means barely visible damages have to be considered in the structural design process. In the standard design process for damage tolerant aircraft structures especially the effect of impacts has to be considered but is mostly investigated on unloaded sandwich components. However, aircraft structures are not completely stress free, even when the aircraft is in parking position. At least the structural weight will cause a permanent static preload. In the literature there are only few investigations dealing with the impact behaviour of preloaded honeycomb sandwich structures. In general, studies on composite structures have shown that tension preloads lead to higher structural impact resistance, while prestressing in compression reduces mechanical properties such as impact strength and stiffness of the structure. For this reason, a compression preload is considered as the more critical load case. In order to evaluate the effect of static compression preloads an experimental study on five different honeycomb sandwich configurations was conducted. The experimental investigations showed that even low preloads cause a higher structural damage area, which results in a lower residual strength of the structure. The results indicated that it is recommendable to consider structural compression preloads in the aircraft structural design process.

Fatigue life estimation of aircraft structural component using FE approach and validation through analytical and experimental methods

Aircraft structural design requirements encompass safe-life and damage-tolerant design, analysis, and verification through testing to arrive at the overall service life. The fatigue phenomenon is most critical and local; hence fatigue studies are necessary to conduct first at the coupon level before attempting at the component or full-scale level. The present work demonstrates the development of fatigue life estimation procedures such as high cycle fatigue (HCF) life and fatigue crack growth (FCG) life on specimens as per ASTM standards.The HCF life is predicted for the Aluminum (Al) 2024-T3 ASTM E466 specimen subjected to constant amplitude loading with a stress ratio (R) of 0.1 based on Basquin's method. The computational 2D finite element (FE) model is developed to predict HCF life using a safe-life approach through Nastran Embedded Fatigue (NEF). The effect of fatigue strength modifying factors affecting fatigue life is also explored. Similarly, the FCG life is predicted for the Al 2024-T3 ASTM E647 specimen subjected to constant amplitude loading with a stress ratio of 0.1 using Walker's equation. The computational 3D FE models with a single edge crack in each are developed to predict FCG life using FRANC3D. The FE-based HCF and FCG life prediction procedures demonstrated in this work are verified by comparing FE results with analytical and experimental ones. Therefore, these FE-based methodologies for HCF and FCG life prediction can be adopted at the feature and structural component levels. The developed computational methods reduce the experimental effort, cost, and time involved in the overall fatigue design of the aircraft structures.

Vibration Response Law of Aircraft Taxiing under Random Roughness Excitation

Understanding the vibration response patterns of aircraft taxiing under runway roughness excitation is crucial for aircraft design and runway performance evaluation. In this paper, we establish pavement roughness models for both asphalt and concrete surfaces, taking into account their unique structural characteristics. We construct a six-degrees-of-freedom aircraft model using multi-rigid-body system dynamics theory and employ the pseudo-excitation method to examine the influence of pavement roughness types on the steady vibration response of aircraft taxiing at constant speeds. Furthermore, we analyze the non-stationary vibration response patterns of aircraft during takeoff and landing taxiing using the method of instantaneous frequency response function with space frequency. Lastly, we explore the effect of stochastic structural parameters on aircraft vibration response using the Monte Carlo method. Our findings reveal that the roughness power spectrum differs between asphalt and concrete pavements, and the established roughness models in this paper demonstrate a strong fit (R2 > 0.95). The type of pavement roughness has a relatively minor impact on the power spectral density distribution of the aircraft vibration response, suggesting that the same roughness model can be used for both asphalt and concrete pavements when high accuracy is not required. The power spectral density distribution of aircraft vibration response varies across different motion attitudes, with the vibration response during landing being significantly larger than that during takeoff. Among the aircraft structural parameters, the randomness of the sprung mass has the most substantial effect on the aircraft vibration response, potentially causing the variation coefficient of dynamic load on the front landing gear to exceed 0.11. Tire stiffness comes next, which can lead to the variation coefficient of dynamic load on the main landing gear reaching 0.07. The results have a guiding role in optimizing aircraft structure design and ensuring pavement performance.

An overview of mechanical properties and failure mechanism of FRP laminates with hole/cutout

AbstractAerospace and aircraft companies follow a pyramidal structure of testing, ranging from specimen tests to full‐scale structure tests, for aircraft structural design. At the base level, open hole strength evaluation is one of the qualification procedures for composite components. The holes are made in the structural components to accommodate bolts and rivets in order to assemble the parts into a single unit. Holes in the component are also used to fulfill functional requirements, such as the passage of wire bundles. FRP components are sensitive to these holes/cutouts as substantial decrement in the strength is observed in several experiments. The degradation in the strength of the component is due to stress concentration around the hole. This review aims to comprehend the effect of the hole on the mechanical properties of FRP laminates; the emphasis is mainly on CFRP laminates and experimental results. An overview of tensile, compressive, flexural and post buckling properties of laminates with the hole is presented to aid in the optimal design of the component. Finally, the failure mechanism during loading is discussed and described through a graphical presentation.

ІНЖЕНЕРНА МЕТОДИКА РЕАЛІЗАЦІЇ КВАЛІМЕТРІЇ У СИСТЕМІ МЕТРОЛОГІЧНОГО ЗАБЕЗПЕЧЕННЯ ЕТАПІВ ПРОЕКТУВАННЯ АВІАКОНСТРУКЦІЙ ІЗ ПОЛІМЕРНИХ КОМПОЗИЦІЙНИХ МАТЕРІАЛІВ

In connection with the permanent growth worldwide, including in the domestic aircraft industry, of the use of polymer composite materials in the structures of civil aircraft, the problem of analyzing the effectiveness of the use of composites at all stages of the life cycle of the corresponding structures becomes relevant. An engineering method for the implementation of qualimetry in the system of metrological support for the main stages of designing structures made of polymer composite materials in civil aircraft of the transport category, based on the use of integrated quality indicators and expert assessment of the characteristics of the main design stages, is proposed and justified. General key tables for the synthesis of significance levels of the main stages of designing composite structures for expert evaluation of the effectiveness of operations have been developed. The main (defining) principles of integral indicators of quality of the components of design levels are established. The reliance of experts on the main features of the quality indicators of the design stages ensures the minimization of errors in establishing the significance ranks of these stages. The decomposition of the classifier of the type "Design of aircraft structures from polymer composite materials" was carried out, which systematizes the analysis and synthesis of the main steps for implementing the proposed engineering methodology for qualimetric support for the design of aircraft structures from polymer composite materials, the sequence and implementation by an expert group.A criterion for the effectiveness of an engineering methodology for qualimetric support for the implementation of design stages has been implemented in the form of an additive model of the ratio of the effect of each design stage, expressed as the sum of quality indicators to the total cost of funds for its provision, taking into account the coefficient of significance of the indicators. A hypothetical example of the implementation of the proposed engineering methodology for analyzing the effectiveness at the initial stages of its development at an enterprise that predicts a promising increase in the volume of use of composites in domestic aircraft of the transport category and their modifications is given. The proposed engineering method for the implementation of qualimetry in the system of metrological support for the stages of designing composite structures of domestic aircraft, after practical testing and refinement in a specific production environment, can become an effective tool for increasing the effectiveness of predicting the growth in the volume of use of polymer composite materials in the domestic aircraft industry.

A novel reduced order model for drone impacts with aircraft structures

Small unmanned aerial vehicles are a threat for manned aviation. Their increased use by hobby pilots and within the commercial sector has been accompanied by an increase in incidents involving manned aircraft. The problem is that current aircraft structures are designed to resist collisions with birds. They are not designed to withstand drone impacts. The composition of drones differs significantly from previously known load cases. Drones consist of several components with various materials. This means, that there is no analytic model to determine the impact force of such drone strikes with aircraft structures. Within this work, a novel reduced order model for drone impacts is developed. It is validated with high velocity impact test data and explicit finite element simulations. The impact of fragmenting components of the drone are modelled with the aircraft impact model. The impact of non-fragmenting components like motors are described with a spring-mass model. The results show that the approach of superimposing a spring-mass model with the aircraft impact model leads to good results. In case of a rigid target only minor deviations occur within the validity range of the model. Damage and degradation of the target is not included in the model what leads to larger deviations in case of an impact with deformable structures. Nevertheless, the model is very well suited for rapid load estimation and can qualitatively reproduce contact force curves. It can be used for preliminary design of aircraft structures without conducting time and cost intensive tests and simulations.

Multi-Objective Reliability-Based Partial Topology Optimization of a Composite Aircraft Wing

Reliability-based partial topology optimization (RBPTO) is a new approach for aircraft structural design; however, it brings computational complexity and makes aeroelasticity analysis quite challenging. Therefore, the present study proposed the multi-objective reliability-based partial topology optimization of a composite aircraft wing using a fuzzy-based metaheuristic (MRBPTOFBMH) approach. The objective is to obtain an optimal layout including partial topology and sizing of the aircraft wing structure. Here, an optimal aeroelastic structure is designed by taking into account the uncertain nature of material properties and the permitted transverse displacement. To increase computational efficiency in the design process, a non-probabilistic approach called a possibilistic safety index-based design optimization (PSIBDO) with fuzzy uncertainties is proposed to quantify the uncertainties due to aeroelastic and structural constraints. Various optimum partial topological shapes and sizing of aircraft wing structures with various PSI values in the range of [0.001, 1.0] have been obtained in a single optimization run. These outcomes, including deterministic and reliable optimal aircraft wing structures, demonstrate the high effectiveness of the proposed MRBPTOFBMH technique to alleviate the complexity of unconventional aircraft wing structure design. The findings also reveal the ease in cooperation of the suggested technique with a high-performance multi-objective evolutionary algorithm (MOEA) and its application in real-world multi-objective design optimization (MODO) problems with the least computational requirements against the traditional method’s multiple runs. Furthermore, the proposed methodology can generate potential aircraft wing structures in a range of m = [89.38–127.84] kg, and flutter speed = [285.61–632.78] m/s, that adhere to all the constraints requirements.

A Multi-Dimensional Lagrange Multiplier Method to Identify the Load Distribution on 3D Special-Shaped Surface in the Strength Analysis of Aircraft Structure

In the process of aircraft structural design, the aerodynamic load and inertial load need to be distributed from single loading points to distributed finite element (FE) nodes before strength analysis. The most commonly used loading distribution method is a Multi-Point Arrangement (MPA) method, which introduces a one-dimensional Lagrange multiplier based on the principle of minimum deformation energy, and simplifies the special-shaped 3D surface in aircraft structure to a plane. However, the actual aircraft structure contains a large number of special-shaped surfaces, and the MPA method cannot accurately distribute the loads on these complex special-shaped surfaces, affecting the accuracy of strength analysis. This paper developed a new 3D load distribution method based on multi-dimensional Lagrange multipliers (MDLM), which can simultaneously achieve an efficient and accurate distribution of surface aerodynamic loads and inertial loads in all directions. Typical numerical cases showed that when an aircraft structure model is a plane, this MDLM method converges to the traditional MPA method. For 3D special-shaped surfaces, the average error of this MDLM method is 0.77–2.28%, which is significantly smaller than the average error of the traditional MPA method (3.30–7.40%).

Increasing the strength of the wing structureaircraft based on the use of metal-polymer composite material

In the modern aviation industry, the task of increasing the structural strength of the aircraft wing is being actualize. The author presents an engineering methodology for power calculation and design of composite aircraft structures based on the analysis of strength characteristics of metal and metal-polymer composite materials. The paper uses theoretical and empirical research methods, as well as the results of foreign and domestic scientific materials.

Quasi‐static fracture characterization of fiber metal laminate with crack resistance curve based on δ5

AbstractFiber metal laminate (FML) is a hybrid composite typical in aircraft industry. It is necessary to effectively characterize the quasi‐static fracture performance of FML, considering the importance of damage tolerance design of aircraft structures and the complicated fracture mechanism of FML. In present investigation, glass fiber reinforced aluminum laminates (Glare) with middle cracks were tested under quasi‐static tensile load. The experimental crack resistance curve (R curve) based on the special crack opening displacement at original crack tip (δ5) was used for revealing the superior crack resistance of Glare laminate to aluminum sheet, which expressed in crack extension aspect was the extended critical crack length. This was thought to be benefit from the contributions of glass fiber layers and fiber/metal delamination, by performing the progressive damage analysis. Since δ5 can embody the elastic–plastic response of the crack but without need to interactively correct for crack tip plasticity. The δ5‐R curve‐based macro and meso fracture models were proposed to accurately predict the residual strength and stable crack extension of Glare laminate, which gave further evidence for the effectiveness of δ5‐R curve in characterizing the quasi‐static fracture performance of FML.