Modern Aircraft Structures Research Articles

Development of on-board power systems and the concept of a more electric aircraft

The article reviews the main stages of development of aircraft power systems. The basic features of the power node for the Mi-24 combat helicopter, a multi-tasking aircraft, are presented F-22 and the Boeing B-787 passenger airplane. The results of the analysis of methods for managing the distribution of electricity on both conventional aircraft and the most modern aircraft structures (A-380, B-787) are presented. The basic structures of the aircraft's on-board power network are presented, divided into three periods: the period until the 1960s (with the dominance of the direct current network with a voltage of 28.5 V), the period until 2000 (with the dominance of the three-phase alternating current network with a voltage of 3 x 115 V / 200 V, 400 Hz) and the most recent period, after 2000 (dominated by a 270 V direct current network, obtained by converting the voltage from a three-phase alternating current generator 3x 115 V / 200 V, 400 Hz using a transformer-rectifier block). The concept of a more electric aircraft was presented as the main direction of current construction works for MEA concept (More Electric Aircraft).

Impact monitoring of large size complex metal structures based on sparse sensor array and transfer learning

During aircraft operations, the impact events experienced by the aircraft may cause damage to the structure, thus posing a safety hazard. Therefore, an accurate determination of where the impact occurred and the time history of the impact force can provide an important basis for assessing the condition of the aircraft. However, modern aircraft structures are often large and complex, and relying on dense arrays of sensors for monitoring adds additional weight to the aircraft and reduces the economics of aircraft operation. This paper proposes a region-to-point monitoring strategy. First, a Convolutional Neural Network (CNN) model with region localization capability is trained using the sparse sensor array acquisition data. Then, the weighted center algorithm is used to determine the specific location where the impact occurs, in which the added fuzzy genetic algorithm can provide the ability to adjust weights to improve localization accuracy adaptively. As for the impact force prediction, this paper adopts a model based on a Convolutional Neural Network-Gated Recurrent Unit combined with a Squeeze-Excitation attention mechanism (CNN-GRU-SE), which is capable of predicting the impact force occurring in the flat plate and reinforced structure region of the aircraft under different energy conditions. Finally, the impact of incorporating a transfer learning approach on model performance and training cost is investigated for fuselage regions with different structures.

One-shot drilling process for thin CFRP/Aluminium alloys stacks

ABSTRACT The aerospace industry pursues lighter, stronger, and more fuel-efficient aircraft, relying on lightweight materials. This study investigates one-shot drilling of thin hybrid stacks, including carbon fiber-reinforced polymer (CFRP, 2.2 mm thick), and aluminum alloys (Al2024-T3, 1.6 mm thick, and Al7075-T6, 1.0 mm thick), crucial in modern aircraft structures. The challenge of thin materials lies in their reduced flexural stiffness, which significantly impacts the quality of the final hole and the effects of the process parameters on torque and thrust force. The experimental approach involves drilling each material separately to analyze thrust force, torque, and hole quality, then comparing outcomes with one-shot drilling in thin hybrid stacks. The results highlight that thin hybrid stacks are challenging to drill because their reduced flexural stiffness significantly impacts the final hole quality and the effects of the process parameters on torque and thrust force differently from what typically occurs for the drilling of conventional hybrid stacks.

Machining Temperature, Surface Integrity and Burr Size Investigation during Coolant-Free Hole Milling in Ti6Al4V Titanium Alloy

Modern Aircraft structures use titanium alloys where the processing of holes becomes essential to assemble aerospace parts. Considering the limitations of drilling, the study evaluates the helical milling for hole processing in Ti6Al4V. The experimental evaluation was conducted by considering burr size, surface roughness, machining temperature, and microhardness under coolant-free conditions. The axial feed and cutting speed were varied at three levels, and nine experiments were conducted. The results exhibit a lower machining temperature during helical milling than during drilling. In addition, the helical milling helped to lower the surface roughness and size of the exit burrs. However, helical-milled holes showed higher subsurface microhardness than conventionally drilled holes. The process variables were influential on machining temperature magnitude. The highest recorded temperature of 234.7 °C was observed at 60 m/min of cutting speed and 0.6 mm/rev feed. However, the temperature rise did not affect the microhardness. Strain hardening associated with mechanical deformation was the primary mechanism driving the increase in microhardness. Helical-milled holes exhibited an excellent surface finish at lower axial feeds, while chatter due to tool deformation at higher feeds (0.6 mm/rev) diminished the surface finish. The surface roughness increased by 98% when the cutting speed increased to 60 m/min from 20 m/min, while a moderate increment of 28% was observed when the axial feed increased to 0.6 mm/rev from 0.2 mm/rev. Furthermore, the formation of relatively smaller burrs was noted due to significantly lower thrust load and temperature produced during helical milling.

Efficient structural optimisation of composite materials aircraft wings

Contemporary computational analysis of modern aircraft structures demands rapid and efficient, yet accurate enough tools for the conceptual and preliminary design stages. On that front, a novel optimisation framework, utilising low-cost numerical tools for sizing composite materials aircraft wings subject to stiffness, strength and dynamic aeroelastic constraints is proposed herein. The structural representation of the numerical model is based on the well-established equivalent-plate methodology, capable of reducing the size and computational cost of the associated problem. An equivalent-plate model of a modern transport aircraft wing is developed and compared to its equivalent 3D Finite Element Method model. Results by means of natural frequencies and modes, indicate excellent accordance between the numerical models. An optimisation framework is then presented, with the ply thicknesses of a baseline lay-up being assigned as design variables. The developed framework succeeds to guide the mass of the wing to a minimum while satisfying the constraints under a critical loading scenario. Optimal lay-ups for the skins, spar webs and ribs as well as spar and rib caps dimensions are obtained. The presented optimisation framework, exhibiting high accuracy and efficiency, constitutes a robust numerical tool for the early design stages of composite aircraft wings.

Нелинейное деформирование и устойчивость композитной цилиндрической оболочки при комбинированном нагружении изгибающим моментом и краевой поперечной силой

Currently, polymer composite materials are widely used in the modern aircraft structures. Their application significantly reduces the weight of the structure, while maintaining its strength and stiffness characteristics. A large number of works have been published on the study of the strength of such structures, but the issues of strength and stability during their nonlinear deformation remain unresolved. There is a small number of works on the study of strength and stability of thin-walled shells made of polymer composite materials. The latter is especially necessary for thin-walled aircraft fuselage structures, where the loss of composite skin stability is unacceptable. The problem of determining the influence of the stacking order of monolayers in a skin on the strength and stability of composite material shells under nonlinear deformation remains unsolved. Methods for calculating the strength and stability of thin-walled composite structures, regarding the nonlinearity of the initial stress-strain state, are not well developed. Therefore, the development of reliable and efficient methods for calculating shells made of composite materials is an urgent task. The article describes solving the problem of strength and stability of cylindrical composite shells under arbitrary loading using finite element methods and Newton-Kantorovich linearization. Critical loads have been determined in the course of solving a geometrically nonlinear problem using the Sylvester criterion. The stability of a circular cylindrical shell made of a polymer composite material has been studied under combined loading by a bending moment and a transverse force. The influence of the nonlinearity of deformation, methods of stacking monolayers on the shell critical loads has been determined.

УЗАГАЛЬНЕННЯ МОДЕЛІ ФОЛЬКЕРСЕНА НА ВИПАДОК ОСЬОВОЇ СИМЕТРІЇ

Thin-walled structures may contain defects as cracks and holes that are leftovers of the material the construction, is made of or they occur during the operation as a result of, for example, mechanical damage. The presence of holes in the plate causes a concentration of stresses at the boundary of the holes and ultimately leads to premature failure of the structural element. Repair of local damage of modern aircraft structures can be made by creating overlays that are glued to the main structure. The overlay takes on part of the load, unloading the damaged area. This method of repair provides tightness and aerodynamic efficiency to the structure. The calculation of the stress state of such glued structures is usually performed by using the finite element method. The classic models of the stress state of overlapped joints are one-dimensional. That is, the change of the stress state along only one coordinate is considered. At the same time, the connections of a rectangular form are also considered. The purpose of this work is to create a mathematical model of the stress state of circular axisymmetric adhesive joints and to build an appropriate analytical solution to the problem. It is assumed that the bending of the plates is absent; the deformation of the plates is even by thickness. The adhesive layer works only on the shift. The main plate and the overlay are considered isotropic. The solution is built on polar coordinates. The stress state of the connection depends only on the radial coordinate, i.e. one-dimensional. The solution is obtained in analytical form. This mathematical model is a generalization of the classical model of the adhesive connection of Volkersen to a circular or annular region and is considered for the first time. Boundary conditions are met exactly. The satisfaction of marginal conditions, as well as boundary conditions, leads to a system of linear equations with respect to the unknown coefficients of the obtained solutions. The model problem is solved and the numerical results are compared with the results of calculations performed by using the finite element method. It is shown that the proposed model has sufficient accuracy for engineering problems and can be used to solve problems of the design of aerospace structures.

Comparative Study of Energy Absorption Capability of Flat Plate Coupons Made by CFRP Plain Weave Fabric Composites

Despite years of governmental and academic institutions’ researches, no experimental standards are established for evaluating crush Specific Energy Absorption SEA for plain weave fabric woven carbon-fiber-reinforced composites used in modern aircraft structures as elements of the boxes to mitigate damage during crush events. At the laboratory scale, this paper proposes a comparative study of energy absorption capability of flat plate coupons made by CFRP plain weave fabric composites. A new fixture design and setup were created with hydraulic pressure and drop tower machines to carry out tests of flat plate composite specimens under quasi-static and low velocity on-axis crash loading. For investigating parameters sensibility of triggers and layups, numerical and experimental results of four trigger types and three stacking sequences were compared. A confrontation between experimental and pre-developed UL-Crush numerical material model results confirms that coupons with 0˚ oriented central plies and saw teeth or corrugated triggers dissipates higher energy during crush, compared to coupons with 90˚ or 45˚ oriented central plies and chamfer 45˚ or steeple triggers. An efficient and simplified experimental methodology was developed to measure and investigate different parameters influencing SEA of composites under crush load. Comparison between experimental and UL-Crush material model confirms the performance of such simulation tool.

Structural Life Management in a Combat Aircraft

The paper briefly describes some of the underlying concepts, issues and strategies for ensuring efficient management of structural integrity of an airframe throughout the useful life of an aircraft—especially that of a military combat aircraft. Much of the discussion in the paper is based on the experience of the authors during the design and development phase of the Indian light combat aircraft, namely, LCA Mk-I Tejas, and the information available about the LCA in the open literature or public domain. Even though not an ageing aircraft, LCA is taken as an example to facilitate bringing out various issues and strategies to resolve them, particularly with relevance to the Indian aeronautical scenario. The use of both metallic and composite materials in modern combat aircraft structures necessitates different design strategies to be adopted for their structural integrity throughout their lifetime. In this respect too, the LCA serves as a good example. Within the limited scope of the paper, the description and discussions are primarily limited to some of the important aspects, such as the fatigue degradation in metallic materials and impact damage in composites—which do form the bulk of the effort of life management of airframe structures. Other issues such as corrosion in metals and environmental degradation in composites are also briefly mentioned. A brief description of the aircraft and its structure is given initially. Major steps in designing for structural life management, such as, estimation of static loads, derivation of fatigue spectra for the intended usage, degradation through fatigue or impact damage, lifing philosophies, estimation of life and monitoring of health and usage of individual aircraft are described. A brief mention is also made about structural health monitoring (SHM) as a future direction in life management. It is expected that the exposition in the paper will help understand issues and strategies that would eventually be required as a combat aircraft gets aged in service.

Effects of Flexural and Mass Axes Position of Simple Rectangular Wing on Flutter Speed

The flexibility of the modern aircraft structures paves the way for the study of aero elasticity during the verification procedures of aircraft design and stability. The structures of wing will cautiously vibrate, either linearly or non-linearly due to the coupling effect between the aerodynamic, inertial and elastic forces of the wing structure, so the designing of aircraft involves two major aero elastic considerations such as wing torsional divergence and flutter. However, flutter problems are considered from the field of dynamic stability, the modal representation of flutter analysis is done by setting up the lifting surface as a linear ordinary differential equation, then changed to Eigen values and further stability analysis has been done. In order to complete the flutter analysis as an analytical tool, two important aerodynamic theories are taken into account and we identified and solved the classical flutter derivatives by using Eigen value method and also the position of flexural and mass axis of the simple rectangular wing on flutter speed has been analysed.

Prediction of crack growth life 2A97-T3 lithium aluminum alloy under Mini-Twist load spectrum

An important aspect of life analysis and damage tolerance design of modern aircraft structures is a reasonable model and method for prediction of fatigue crack growth behavior and growth life under actual flight spectrum loading. Mini-Twist load spectrum and constant amplitude crack propagation curves test have been conducted towards the 2A97-T3. Coefficients of Paris equation and Walker equation for constant amplitude crack propagation curves are given. The Wheeler retardation model and AFFDL closure model are used to predict the aboved crack growth life. The predicted results of Wheeler retardation model and AFFDL closure model agree well with the experimental results.

Thermal plasticity deformation buckling analysis on the carbon fiber-reinforced aluminum matrix composite material

As light weight and durability have become the important design objectives for the modern aircraft structures, the carbon fiber-reinforced resin matrix composites are widely applied to the aircraft structures for it can reduce the airframe weight effectively at the same time of guaranteeing the load-carrying capacity. In aircraft structures, the mechanical connection remains the important connection pattern for the composite material and the laminated structures; however, existence of the connecting hole has resulted in the stress concentration of structures while the borehole fatigue breakdown of parts is the main failure mode of aircraft structures. According to the low fatigue life of the aircraft connection structures, research on the hole extrusion and reinforcement technology on the carbon fiber-reinforced composite material and the laminated structure is made in this paper.

Simulating of Bird Strike on Aircraft Laminated Glazing

A bird strike is a critical problem in the context of safety in the aviation industry. All modern aircraft structures are designed with account of likely collision with birds. Thus, aviation standards in force require that the aircraft construction would allow the crew to conclude the flight safely after collision with a 1.81-kg bird. A method for analysing the stress-strained state of laminated airplane glazing at different operational factors is presented. The method includes a technique for strength analysis of the laminated airplane glazing at bird impact, and a technique for analysis of excess pressure. The model of laminated glazing is based on the refined first-order theory accounting for transverse shear strains, thickness reduction and normal element rotation inertia in each layer. The mathematical model of the pressure impulse authentically reproducing the bird impact is based on experimental research. Theoretical results are in good agreement with experimental data, thus allowing to recommend the method to develop new airplane glazing elements.

The application of thin-walled integral constructions in aviation as exemplified by the SAT-AM project

The high level of durability and reliability is a strict requirement for today's aircraft constructions. The most important aspect is the mass of the structure which has a decisive influence on both the volatile and technical properties of an aircraft as well as its economical operation. Modern aircraft structures, or more precisely their load-bearing structures, are almost exclusively manufactured as thin-walled ones which meet the requirement to minimize the mass of the structure. While the local loss of coverage stability is permissible under operating load conditions, exceeding critical load levels of the structural skeleton elements (e.g. frames, stringers) is practically synonymous with some destruction of the structure. It results in a necessity for the continual improvement of both design methods and the improvement of structural solutions of aviation structures. By reducing the thickness of the cover and, at the same time, introducing densely spaced stiffening longitudinal elements, a structure with considerably higher critical loads can be obtained and, consequently, a more favorable distribution of gradients and stress levels which directly contributes to an increase in fatigue life. This paper attempts to analyze the above-mentioned problem more closely, using the example of a densely ribbed rectangular plate.

Crack growth analysis and residual life estimation of structural elements under mixed modes

Attention in this work is focused to crack growth analysis and residual life estimation of metal structural components under mixed modes I/II. Primary attention of this investigation is numerical simuation crack growth trajectory. For this purpose FEM is used together with the maximum tangential stress (MTS) criterion. To determine crack growth trajectoties of cracked structural components under mixed modes convencional singular finite elements are used. In this investigation specimen with two holes and crack between its is considered. The crack between two holes is defined to achieve mixed mode I/II crack growth trajectory.The main objective of this work is to develop a computational analysis methodology to simulate realistic crack growth and to predict remaining life and residual strength of complex aircraft and another flight structures. Modern aircraft structures are designed using damage tolerance philosophy. This design philosophy envisions sufficient strength and structural integrity of the aircraft to sustain major damage and to avoid catastrophic failure. An special attention in this work is focused to study the effects of initial angle of crack between two holes on crack growth trajectories. Computation crack growth trajectories are compared with own experimental results. There are good agreements in determining the path of the crack growth and life estimation. Good agreement between computation and experimental crack growth trajectory is obtained.

Synergistic effects of spray-coated hybrid carbon nanoparticles for enhanced electrical and thermal surface conductivity of CFRP laminates

Carbon fibre reinforced plastics (CFRPs) are intensively used in modern aircraft structures because of their superb specific mechanical properties. Unfortunately their electrical and thermal conductivities are not sufficiently high for some applications like electromagnetic interference (EMI) shielding and lighting strike protection (LSP). The addition of external metallic structures, such as aluminium or copper mesh, is generally required with a compromise in terms of increased mass and manufacturing cost as well as reduced corrosion resistance. In the present work spray coating of carbon nanoparticles was utilized as a simple method to locally increase the electrical and thermal suface conductivity of CFRPs. The combined use of carbon nanotubes (CNTs) and graphene nanoplatelets (GNPs) synergistically reduced the CFRPs surface resistivity by four orders of magnitude (from 2–3 Ω/sq to 3 × 10−4 Ω/sq) and increased the thermal conductivity by more than 7 times (from 200 W m−1 K−1 to 1500 W m−1 K−1), opening up possibilities for the replacement of metallic mesh structures for EMI shielding and LSP. An analytical model was introduced based on a one-dimensional heat conduction approach to predict the effective thermal conductivity for the hybrid nanofiller coating layer and its findings showed good agreement with experimental data.

A strength-model for laser joined hybrid aluminum–titanium transition structures

Abstract Besides aluminum (Al) and titanium (Ti) alloys carbon fiber reinforced plastics (CFRP) are increasingly demanded for modern aircraft structures. Therefore, suitable Al–Ti and Al–CFRP joining techniques are of great interest. A novel concept for Al–CFRP joints uses titanium laminates as transition elements. In this paper, the tensile strengths of double-sided laser beam joined hybrid structures of aluminum sheets and single-layer as well as multi-layer titanium parts are investigated by varying the edge shape and the laminate build-up. The full-double-15°-scarf-edges produce the maximum strength. It is shown that a model can explain and predict the different failure behaviors and strengths.

The Use of Wireless Sensor Technologies for Condition Monitoring of Modern Aircraft Structures

Monitoring of the aircraft structures’ during the pre-flight testing is a critical task of the aerospace industry. One of the most promising solutions, not yet widely applied, is continuous monitoring of aircraft structures using wireless sensor network technology. The brief summary of the proposed system is the following: special sensors (devices which detect potential deformations), send signals to the local motes (autonomous computing device equipped with a wireless transmitter). Information from motes is gathered by routers which then transfer the aggregated information to the datacenter. Motes are used for collection and primary processing of the data from sensors, which then sent to the routers, and, ultimately, gateway of corporate network. Applications of corporate network control and define flexible patterns for processing of the information received from sensors. As an example of hardware components for preprocessing and data transmission the system described in this study uses motes and routers from MEMSIC, for measuring the level of deformation - sensors from Tokyo Sokki Kenkyujo. This network structure allows to centralize data collection modes in the process of testing; implement continuous data collection at a defined frequency; process and display data in real-time.

Engineered-Based Machining Parameters Analysis for Aircraft Structural Parts

Modern aircraft structures need to be further integrated and parts to be enlarged, the structural parts which require high accuracy of size, shape, and good deformed surface finish, so digital machining process has become key technology in aviation industry. Machining parameters selection is an important phase of machining process, it has very important influence on the dimensional quality of an aircraft structural part. This paper develops a methodology to get an optimal machining parameters in an aircraft structural part machining process, which can satisfies the quality specifications and minimizes the expected value of the sum of machining costs at the same time. An engineered-based model is adopted to describe the machining process of aircraft structural part firstly and then a quantitative machining parameter evaluation methods driven by cost constraint is proposed, an optimal machining parameter is obtained through a computer experiment model which can explore a large number of machining process parameter alternatives. At last, we illustrate the validity and the significance of the engineered-based machining parameters analysis method for aircraft structural parts by a case study. The results have shown that the proposed method could significantly increase the efficiency of aircraft structural parts machining process.

Predictive Modeling for Tool Deflection and Part Distortion of Large Machined Components

Machined monolithic components provide the foundation for modern aircraft structures requiring high performance designs in terms of weight, strength, and fatigue properties. Part distortions arising from machining-induced stresses and tool deflection resulting from high dynamic cutting forces frustrate manufacturing and assembly processes, necessitating expensive in-process and post-machining corrective measures aimed at eliminating their effects. Component weight and cost requirements are often compromised by adding thicker component section designs or off-machine part flipping processes as a distortion control mechanism. Similarly, small depths of cut are taken to minimize cutting forces and tool deflection. Additional spring passes are also taken to eliminate any undercut errors introduced by unanticipated deflection and distortion problems. The ability to accurately predict and minimize tool deflections and part distortion via simulations can significantly reduce manufacturing and assembly costs. This paper presents physics-based models for predicting tool deflection and part distortions by considering the appropriate physics for each problem. Dynamic cutting forces predicted by physics-based machining models, and tool compliance properties are incorporated into a detailed linear elastic deflection model in order to predict in-process deflections along a computer numerical control (CNC) machining toolpath. Similarly, bulk stress state and machining-induced stresses for large, monolithic part machining are taken into account for predicting part distortions. Sources of stresses may include heat treatment, quenching, forging, and machining operations. Bulk stress data from heat treatment predictive models can also be imported and mapped onto the workpiece finite element model. CNC part programs, along with stresses arising from corresponding tooling, are processed and analyzed. Results of a validation study for workpiece distortion predictions are presented for a number of monolithic, thin-walled components. Predictions for tool deflection are also compared against experimental measurements for multiple cutting configurations and tool diameters. Good correlation is found between predictions and measurements of deflection.