Design Of Aircraft Structures Research Articles

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.

The relationship between surface roughness and fatigue crack growth rate in AA7050-T7451 subjected to periodic underloads

The accurate prediction of fatigue crack growth rates is necessary to assist with optimising the design, maintenance, and availability of aircraft structures. Roughness induced crack closure is considered by many researchers as an important factor that can impact these growth rates. It has also been suggested that crack closure may be the governing factor behind the accelerative effects of underloads. As such, the focus of this study is on better understanding what correlation there is between the known accelerative effects of underloads, and the local fracture surface roughness of fatigue cracks. Here, specially designed loading spectra are used to quantify the effects of underloads on local fatigue crack growth. Photogrammetry is then used to quantify surface roughness local to the application of the underloads, in order to explore any potential relationship between observed crack growth acceleration and local crack surface roughness. In addition, a general fractographic analysis of the fatigue fracture surface was conducted to understand what correlation there may be between any observed features and the fatigue crack growth rate. It is found that there is little correlation between small-scale surface roughness and the observed acceleration from underloads. However, these is some evidence that large out-of-plane features, such as river patterns, correspond with changes in crack growth rate.

Damage assessment of aircraft wing subjected to blast wave with finite element method and artificial neural network tool

Damage assessment of the wing under blast wave is essential to the vulnerability reduction design of aircraft. This paper introduces a critical relative distance prediction method of aircraft wing damage based on the back-propagation artificial neural network (BP-ANN), which is trained by finite element simulation results. Moreover, the finite element method (FEM) for wing blast damage simulation has been validated by ground explosion tests and further used for damage mode determination and damage characteristics analysis. The analysis results indicate that the wing is more likely to be damaged when the root is struck from vertical directions than others for a small charge. With the increase of TNT equivalent charge, the main damage mode of the wing gradually changes from the local skin tearing to overall structural deformation and the overpressure threshold of wing damage decreases rapidly. Compared to the FEM-based damage assessment, the BP-ANN-based method can predict the wing damage under a random blast wave with an average relative error of 4.78%. The proposed method and conclusions can be used as a reference for damage assessment under blast wave and low-vulnerability design of aircraft structures.

Analysis of the openings of aircraft composite laminated wing beam web under shear load

Abstract In this paper, experiments and finite element methods are used to study the openings of the composite wing spar web of a civil aircraft. First, according to the design principles of composite laminates, determine the layup sequence and layup ratio of the spar, and then lay the beam webs and the beam edge strips integrally; then, calculate the equivalent of the spar by using the classic laminate theory Material properties, the method to calculate the equivalent elastic modulus of composite laminates of different thicknesses is derived; then, the influence of opening shape and opening angle on the strength of composite wing spar web under shear load is analyzed. Finally, the reinforcement design is carried out on the opening position. This research method not only solves specific engineering design problems, but also proposes a set of practical, cost-effective and efficient design methods for composite material opening shapes, which has a certain generality. Its design ideas and methods provide technical support for civil aircraft structural design.

Experiment and analysis of mechanical properties of carbon fiber composite laminates under impact compression

Abstract Composite materials are more and more widely used in aircraft structural design, and impact damage is the most serious defect/damage form of aviation composite structures in service. Therefore, it is of great significance to study the impact resistance of composites with different materials and structural forms for aircraft structural design. In this article, the damage of reinforced fiber/resin composite laminates with unidirectional carbon fiber prepreg (UIN23100), carbon fiber woven fabric (W-3021FF), and carbon fiber-aramid fiber blended fabric (W-38211) under the same impact energy were studied. The impact damage area was scanned and analyzed by ultrasonic C-scan, and the impact resistance of the three laminates were obtained. Finally, the compression bearing capacity of carbon fiber laminates after impact was tested, and the compression failure mode of each material has been analyzed. The results show that the impact resistance of carbon fiber unidirectional prepreg was stronger than that of carbon fiber woven fabric; the compression bearing capacity of carbon fiber laminate after impact was higher than that of carbon fiber unidirectional prepreg; the post impact bearing capacity of W-38211 was higher than that of W-3021FF.

Residual stress and long fatigue crack growth behaviour of laser powder bed fused Ti6Al4V: Role of build direction

Laser powder bed fused (LPBF) Ti6Al4V alloy has a great prospect to manufacture complex aircraft structures. The anisotropic fatigue crack growth behaviour and the residual stress effects are two main concerns on the damage tolerance design of aircraft structures, and should be studied further. In this paper, the compact tension (CT) specimens in four off-axis build directions (15°, 30°, 60° and 75°) were designed and manufactured. Residual stresses of the as-built and the stress relieved samples were measured. Fatigue crack paths and fracture surfaces were analyzed and compared. It was shown that the residual stress along the loading direction in the 15° as-built sample was about 3.5 times larger than that perpendicular to the loading direction, while that of the 75° as-built sample was opposite. The crack deflection angle of the 60° as-built sample was the largest, and it was about 22.4°. Fatigue crack growth rate in the as-built sample fluctuated. After post-treatment, the anisotropy of residual stress state and fatigue crack growth behaviour decreased.

Measurements of Aerodynamic Performance of the Fuselage of a Hybrid Multi-Rotor Aircraft with Autorotation Capability

The design and optimization of aircraft structures need to be accompanied by the analysis of their aerodynamics. This paper presents numerical calculations and wind tunnel tests to determine the aerodynamic characteristics of the designed unmanned aircraft. The aim of the study was to analyze the aerodynamic performance of the designed hybrid unmanned aerial vehicle and its longitudinal stability. This unique design, i.e. a hybrid aircraft which is a combination of a gyrocopter and a multi-rotor aircraft has not been the object of research yet. The research object is a 1:1 scale model created by the rapid prototyping method. The research was a computational and experimental study. The ANSYS Fluent software was used for the calculations, and the computational mesh of the developed model consisted of 3.5 million tetrahedral elements. The numerical investigations were carried out using the created CFD model with a k-ω turbulence submodel. In addition, the obtained results made it possible to numerically analyze the forces acting on the individual components of the research object, which is a valuable extension of the wind tunnel tests. The experimental studies were performed in a closed-loop subsonic wind tunnel. Aerodynamic forces and moments were measured using a six-component force balance. The obtained results were compared to validate the developed numerical model. The research values describe the performance of the research object in terms of minimum drag force coefficient, maximum lift to drag ratio, and properties related to stability.

Numerical Simulation of the Stress-Strain State of a Hybrid Titanium-Polymer Composite Material for Optimizing the Design of Aircraft Structures

Work on the current topic of hybrid titanium-polymer composite materials (TPCM) is continuing. After checking the technological feasibility of two-stage processes for manufacturing parts of flat shapes and with a single curvature, identifying opportunities for improving the adhesive interaction, work is carried out to assess the possibility of varying design parameters, such as reinforcement schemes, combinations of components, etc.in order to obtain optimal properties of manufactured parts.

Experimental analysis and simulation of low-velocity impact damage of composite laminates

The capability to accurately analyse the response of multi-directional composite laminates during impact events is of high importance for the design of lightweight aircraft structures. In this work, both experimental and numerical analyses are performed covering a large design-space of laminates for all aspects from on-set of damage and barely visible impact damage up to clearly visible impact damage and full penetration of the laminates. The impact tests are simulated using a sophisticated three-dimensional continuum damage model, combined with an automated meso-scale model generation algorithm for ply-by-ply, material/fibre-aligned meshing of laminated composite coupons. To assess the accuracy of the predictions, an extensive validation test program of several configurations and impact energies has been performed, thus demonstrating that the simulations are capable of accurately predicting the damage and failure mechanisms under low-velocity impact loading. Not only the evolution of impact loads and energy dissipated are numerically analysed, but the competition of the dominant failure mechanisms from low impact energy and full penetration cases are also macroscopically replicated.

Methodology of aircraft structural design optimisation

Methodology of aircraft structural design optimisation

Retracted] Application and Optimization of Wing Structure Design of DF‐2 Light Sports Aircraft Based on Composite Material Characteristics

Compared with ordinary metal structures, advanced composite materials have the characteristics of high strength, high rigidity, and light weight. The use of composite materials in aircraft structures is currently a hot research topic. This research mainly discusses the optimization design of the composite wing structure of the DF‐2 light sports aircraft. This article takes the DF‐2 light sports aircraft planned to be produced by the company as the source. Based on its overall design basis, aerodynamic requirements, and the original wing structure design, according to the composite material aircraft structure design theory and method, the aircraft wing structure is carried out. Composite materials are materials with new properties that are composed of two or more materials with different properties at the macroscale by physical and chemical methods. Composite materials can be divided into functional composite materials and structural composite materials according to the nature of the application. Functional composites are materials with special functions, such as conductive composites, ablative materials, and frictional composites. At present, the main research is on structural composite materials, which are composed of two components: matrix material and reinforcing material. The new structural scheme design and structural strength analysis are designed to meet the structural strength requirements of the wing and the lightest weight. In this paper, according to the force transmission characteristics of different structural types of the wing, the characteristics of the load transmission are analyzed, and the shape parameters and load parameters of the wing structure design are used as initial conditions, and the quantitative analysis model of the wing structure is constructed according to the requirements of strength, stiffness, and stability. Through rapid mathematical modeling and analysis of the wing structure, the weight and efficiency of different configurations can be evaluated. Through the quantitative analysis model of the wing, the wing structure type can be quickly determined according to the wing parameters in the preliminary design, which makes the basis for the selection of the wing structure type. After optimization, the weight of the wing structure decreased from 0.966 kg to 0.803 kg, a decrease of 16.87%. The designability of composite materials is one of its major characteristics. By optimizing the layup angle, layup sequence, and dropout area, the performance indicators of the structure are finally improved. This research will promote the further development of the aerospace field.

Methodology of aircraft structural design optimisation

Methodology of aircraft structural design optimisation

Algorithm for Compression Design Allowable Determination of Composite Laminates with Initial Delaminations

With the increasing demands for detailed design of composite aircraft structures, the method of covering all damages with low design allowables cannot meet the current requirements for aircraft structure design. Herein, this paper proposes a novel algorithm for design allowable determination of composite laminates by combining the damage distribution with damage factor model of design allowable, so as to provide different structures with more accurate design allowables based on their initial damages. For the composite laminates with initial delaminations, a model describing the effect of delamination size and depth position on the compression design allowable is developed and the compression design allowable of different aircraft structures are individually determined by employing abundant initial delamination statistics. Compared with the design allowable offered by the single-point method, the design allowable based on the initial damage can be increased by at least 5% to 20%, greatly improving the economic benefits of the aircraft structures and providing an important support for the damage tolerance design of the composite structures.

Editorial

Editorial

Structural Analysis and Optimization of Wings Subjected to Dynamic Loads

Dynamic loads arising from the Title 14 Code of Federal Regulations (14 CFR; “Aeronautics and Space,” Federal Aviation Administration, U.S. Dept. of Transportation, Subchapter C Aircraft, Part 25 Airworthiness Standards: Transport Category Airplanes, https://ecfr.federalregister.gov/current/title-14/chapter-I/subchapter-C/part-25 [retrieved 23 August 2021]) are critical drivers of aircraft design. The aircraft structural design process is an iterative one involving computationally expensive simulations. Existing methods simplify the dynamic loads to equivalent static ones and design the structure using shell-based analysis. Other literature approaches use beam models for structural sizing but make assumptions on the geometry, material distribution, or both. In this work, the use of the Variational Asymptotic Method is explored to 1) systematically reduce the three-dimensional (3-D) wingbox to a one-dimensional (1-D) model, 2) solve the 1-D model using a nonlinear beam theory, and 3) recover the stress field on the wingbox. An adjoint-driven gradient-based structural optimization framework is developed to size the structure for dynamic loads subjected to strength-based failure considerations. The proposed structural analysis method is applied to the wingbox structural analysis and sizing of a novel hybrid-electric propulsion aircraft: NASA’s Parallel Electric-Gas Architecture with Synergistic Utilization Scheme (PEGASUS) concept. Results in the paper show that the sizing of aircraft wings for 14 CFR specified maneuvers 2.5 and static maneuvers using the proposed approach produces a 6% difference in weight compared to a shell-based method but with a speedup of 7.8 times, and it is efficient at sizing the wingbox structure for dynamic loads generated by gust maneuvers.

Modelling and Finite Element Analysis of an Aircraft Wing using Composite Laminates

The use of finite element analysis is increasingly gaining popularity in the design and analysis of aircraft structures. Also, composite materials in the manufacturing and design of aircraft structures are increasing due to their exceptional properties of high stiffness to weight ratio. Applications include Composite structures are made up of laminates with different fibre orientations, which adversely affects the property of the composites formed. For this study, the three main parts of the aircraft wing are considered for designing: the spars, the ribs, and the aircraft’s skin. The spars and made up of aluminium alloy, and the ribs are made up of structural steel. Two different wing models are designed. For Model 1, the skin is assigned Titanium alloy as a material, and for Model 2, CFRP (Carbon Fibre Reinforced Polymer) with resin epoxy is used as the skin material. The ribs are made of structural steel and the spars of titanium alloy for both models. This paper aims to compare the materials mentioned above to find the optimum strength to weight ratio. It was found that, for a given uniform load on the bottom of the aircraft’s skin, a weight reduction of 2.37 % was observed in favour of Model 2, the deformation was reduced by 51 per cent, and the von-Mises stress was reduced by ≈ 85%. This paper’s results can be used in manufacturing structural components of UAV, especially the wings due to lighter weight and more flexibility and durability than standard titanium alloys and the significant role played by the orientation of the plies in deciding the strength of the composite.

Experimental study of a virtual allowables approach for the design of composite aircraft structures

Statistically derived allowables are representative for the behavior of a composite material system in a given structural context where data variability needs to be contemplated. In particular, they define the strength of the material as characterized by various coupon tests according to ASTM standard procedure subjected to the actual manufacturing process and layup stacking sequences. The current paper presents a detailed comparison between virtual allowables predictions obtained using MSC Digimat and the experimentally determined values of a certified material backed up by a 3x2x3 (batch/panel/specimen) stochastic approach. Tests generate the required data for building material model ab initio of carbon fibers reinforced epoxide resin in form of 2x2 twill woven fabric used for qualifying an innovative composite wing for the Next Generation Tilt Rotor (NGCTR). Numerical models of in-plane tension, compression and shear test methods have been assembled along with an associated material model allowing numerical predictions to be validated with the coupon level experimental results as final purpose. Standard statistical model has been adopted in order to include material and manufacturing process variabilities to define B-basis allowables. The calibrated statistical FEM method performed for different layup configurations have been used to fully characterize the mechanical behavior of the analyzed CFRP material and to predict performances for thicker laminates.

Eric Harold Mansfield. 24 May 1923—20 October 2016

Eric Mansfield was a structural engineer who worked at the Royal Aircraft Establishment Farnborough from 1943 to 1983 on a wide range of problems in the field of aircraft structures. His main working tool was the mathematical theory of elasticity, which he deployed with skill, ingenuity, perseverance and imagination to investigate many problems related to the design of aircraft structures during a period of development of higher flight performance and new materials. His studies included aircraft wings and fuselages, particularly in the presence of cut-outs and openings—for which he invented the ‘neutral hole’ concept—and problems involving large-deflection structural behaviour of plates, including thermal effects such as flexural and torsional buckling of fins and wings heated in supersonic flight. He also developed the novel and powerful tension field theory for understanding the post-buckling behaviour of thin webs in shear, and the analogous inextensional bending theory of thin wings. His mathematical models were confirmed by elegant experiments. Mansfield's book The bending and stretching of plates contains clear descriptions of many of his research investigations; it has become a classic text. All of Mansfield's writing is incisive and clear. He was widely regarded as a seminal figure in the field of applied structural mechanics in the second half of the twentieth century.

Computation of dimensional variations on the structural analysis of multi-cell aircraft box beams with python scripting

PurposePython codes are developed for the versatile structural analysis on a 3 spar multi-cell box beam by means of idealization approach.Design/methodology/approachShear flow distribution, stiffener loads, location of shear center and location of geometric center are computed via numpy module. Data visualization is performed by using Matplotlib module.FindingsPython scripts are developed for the structural analysis of multi-cell box beams in lieu of long hand solutions. In-house developed python codes are made available to be used with finite element analysis for verification purposes.Originality/valueThe use of python scripts for the structural analysis provides prompt visualization, especially once dimensional variations are concerned in the frame of aircraft structural design. The developed python scripts would serve as a practical tool that is widely applicable to various multi-cell wing boxes for stiffness purposes. This would be further extended to the structural integrity problems to cover the effect of gaps and/or cut-outs in shear flow distribution in box-beams.

Аэродинамический профиль в трансзвуковом потоке газа

The subject of investigation in this article is transonic flow. This is a condition in which local speeds of sound are appears on the wing surface, even at the subsonic speed of the nonturbulent flow. As a result, at a certain speed of the incoming flow, the flow regime around the aerodynamic surface will change sharply, which in turn changes the aerodynamic characteristics. Aerodynamic surfaces of the most transport category airplane experience transonic airflow during flight. The goal of the investigation is to study aerodynamic characteristics using numerical methods. The use of numerical methods in the design of aircraft structures is used more and more often to determine the optimal parameters for given operating conditions. This contributes to obtaining a more optimal and perfect design. In this article, we carried out a numerical analysis of the aerodynamic characteristics of airfoils in the transonic flow case using the CAE system CFD ANSYS. As a result of the research, the distributions of the pressure coefficients over the profile surface were obtained. The nature of the flow is obtained, which is similar to the picture of the pressure coefficients for transonic flow in the published sources of this topic. In the area of the middle of the profile, a shock-wave is observed. As a result, the flow around the airfoil changes, which contributes to a change in aerodynamic characteristics. The behavior of the aerodynamic drag and lift coefficients depending on the speed of the Mach number is considered. Also, the position of the center of pressure was analyzed at various velocities of the nonturbulent flow. The calculation was carried out at the cruising altitude of a medium-range aircraft of 11 km. For the calculations, we used the characteristics of the air temperature, the pressure of a given height from the table of the standard atmosphere.