Aircraft Structures Research Articles

Autonomous Electrical Current Monitoring System for Aircraft

Aircraft monitoring systems offer enhanced safety, reliability, reduced maintenance cost and improved overall flight efficiency. Advancements in wireless sensor networks (WSN) are enabling unprecedented data acquisition functionalities, but their applicability is restricted by power limitations, as batteries require replacement or recharging and wired power adds weight and detracts from the benefits of wireless technology. In this paper, an energy autonomous WSN is presented for monitoring the structural current in aircraft structures. A hybrid inductive/hall sensing concept is introduced demonstrating 0.5 A resolution, < 2% accuracy and frequency independence, for a 5 A – 100 A RMS, DC-800 Hz current and frequency range, with 35 mW active power consumption. An inductive energy harvesting power supply with magnetic flux funnelling, reactance compensation and supercapacitor storage is demonstrated to provide 0.16 mW of continuous power from the 65 μT RMS field of a 20 A RMS, 360 Hz structural current. A low-power sensor node platform with a custom multi-mode duty cycling network protocol is developed, offering cold starting network association and data acquisition/transmission functionality at 50 μW and 70 μW average power respectively. WSN level operation for 1 minute for every 8 minutes of energy harvesting is demonstrated. The proposed system offers a unique energy autonomous WSN platform for aircraft monitoring.

Structural and flow field analysis of an airfoil subjected to blast-impacted

The shock wave from the explosion of the munition can cause severe damage to the aircraft structure, thereby affecting the aerodynamic performance of the aircraft. This paper aims to propose a Computational Fluid Dynamics (CFD) analysis for the aerodynamic behaviour of the wing after the damage and deformation caused by the impact load. To obtain the damaged deformed wing model for CFD analysis, the explosion damage process was calculated using LS-DYNA software, and the obtained structural deformed wing model was imported into the CFD solver for flow field calculation to obtain lift, drag, and moment coefficients and lift-to-drag ratio. The lift of the damage-deformed wing model decreases due to the destruction of the pressure distribution of the damage-deformed wing model. The lift-to-drag ratio is also affected by shock loading, with smaller values at all considered angles of attack. Similarly, the damage deformed wing model has larger moment coefficients at angles of attack less than 19.5°.

Literature Review on Air Conditioning Heat load Analysis of a Cabin

Abstract: Air conditioning system is necessary in automobile to maintain the comfort condition. Therefore, it is important to study transient response of vehicle under real driving conditions. The Heat Balance Method (HBM) is used for estimating the heating and cooling loads which develops inside the vehicle cabin. The Load Calculation of Automobile Air Conditioning System is calculated and presented. An hourly cooling load is calculated in Tirunelveli region with respect to its latitude. Tons of refrigeration required is also found out from the cooling load calculation and a review on cooling load calculation is also presented. This study gives overall cooling load and AC power consumption which can be used by HVAC engineers to design more efficient car AC systems. Only by knowing the cooling load and source of thermal load, we can develop intelligent system to reduce AC consumption. Environmental Control System (ECS) is a generic term used in aircraft industry for maintaining the required air temperature and pressure inside passenger and crew compartment. The real challenge for an ECS is to operate and supply adequate cooling over a wide range of ground and flight conditions in a most reliable and efficient manner. The most important aspect of ECS is to know the cabin heat loads during air cooling or air heating phases. A precise calculation of aircraft heat load would be a long process requiring detailed knowledge of the aircraft structure and of the quantities involved in the mechanism of heat pick up and interchange in the cabin. This paper describes a method that may be applied in calculating the heat load for any aircraft and calculations are shown for a typical light transport aircraft under the conditions that, the cabin altitude is maintained within 0 ft to 8000 ft (2438.4m) for the aircraft altitude of 0 ft to 25000ft (7620m). The system shall be capable of maintaining temperatures within 59°F to 75°F (15°C to 30°C).

Finite Element Method Analysis in Simulation of Induction Heater Using Helical Coil Method

The problem of corrosion is very common in aircraft structure. This phenomenon is happened because of the two different types of material are in contact under moist or water. Along the operational hours of the aircraft structure, this corrosion is identified during the maintenance activity. To study about the corrosion removal by using the helical coil method, in this research the simulation of the removal corrosion in bolt and nut joint are conducted. The aim of the analysis i.e., to get the optimum electric current and frequency for induction heaters to heat stainless steel nuts that cannot be open, or the threads are stuck. The study was carried out so that the stainless-steel nuts reached a temperature of 400°C in one minute. Several variations of the electric current are carried out with magnitudes, namely 30 A, 40 A, 50 A and frequency variations of 300 kHz, 400 kHz, 500 kHz. Based on the variation of the electric current and frequency, the optimal electric current and frequency is generated to heat a stainless-steel nut that cannot be opened, or the screw is stuck at 50 A and 500 kHz with a temperature of 394°C in one minute.

Application of neural network technologies to load monitoring in aircraft structure bearing elements

Economic efficiency in the operation of aircraft local air lines (LA) can be improved by increasing the awareness of the loading and aircraft structure integrity, with the subsequent adjustment of the maintenance program in accordance with the actual operating conditions. The allowance for specific features of loading load-bearing elements at transport aircraft during particular operations provides a significant extending in the safe life (in some cases, by several times). However, at the moment, load monitoring systems in the aviation industry have not yet been implemented universally. The implementation of the well-known approaches developed for monitoring and analyzing loads requires significant changes in the programs and procedures for maintaining airworthiness, including the need to install additional sophisticated measuring equipment. We propose an alternative approach to existing methods of load monitoring without using additional measuring equipment. The main stage is formation of the relationship between the flight parameters recorded by the standard on-board recorder and the loading parameters, which are determined by computational and experimental methods as a result of processing strain gauge data. A sufficient bulk of the strain data is usually obtained during the flight tests at the certification stage. Testing of this technique is considered with reference to the example of the loads in the elements of high lift devices of aircraft flaps. The average error in the estimates of the forces in the connecting rods of the inner flaps does not exceed 6%. The results obtained make it possible to determine with the acceptable accuracy the individual accumulated damageability of an aircraft structure and to assess the residual safe life. The presented approach is a part of the methodological framework necessary for the development and implementation of modern means of analyzing the integrity of the structure, implemented on the basis of on-board monitoring systems of local airlines aircraft.

Optimum alternate material selection methodology for an aircraft skin

Weight penalty has been a challenge for design engineers of aerospace vehicles. Today’s high-efficiency combat aircraft undergoes intense stress and strain during flying missions, which require stronger and stiffer materials to retain structural integrity. Though metallic materials have been successfully used for the construction of aircraft structures and components, metals still have a low strength-to-weight ratio. This paper aims to develop an alternate optimised material selection methodology to replace the metallic skin of a medium-sized military aircraft. The search for the optimum material will result in reduced aircraft weight which will be benefitted by extra payload on the aircraft. The selection methodology is comprised of finding design pressure limits on the aircraft skin, and comparison of properties (strength, elastic modulus, shear modulus, etc.) and performance (safety factor, deflection, and stress) of the existing metallic skin with alternate optimised material. The comparison was made under aerodynamic pressure, bending force, and twisting moment. Carbon Fibre Reinforced Polymer/Epoxy (CFRP) Uni-Directional (UD) prepreg (elastic modolus is 209 GPa) was selected as an alternate optimum material to replace the aluminium alloy skin of the aircraft studied. The selected alternate optimum material resulted in the reduction of aircraft skin weight by 30%.

Learning defects from aircraft NDT data

Non-destructive evaluation of aircraft production is optimised and digitalised with Industry 4.0. The aircraft structures produced using fibre metal laminate are traditionally inspected using water-coupled ultrasound scans and manually evaluated. This article proposes Machine Learning models to examine the defects in ultrasonic scans of A380 aircraft components. The proposed approach includes embedded image feature extraction methods and classifiers to learn defects in the scan images. The proposed algorithm is evaluated by benchmarking embedded classifiers and further promoted to research with an industry-based certification process. The HoG-Linear SVM classifier has outperformed SURF-Decision Fine Tree in detecting potential defects. The certification process uses the Probability of Detection function, substantiating that the HoG-Linear SVM classifier detects minor defects. The experimental trials prove that the proposed method will be helpful to examiners in the quality control and assurance of aircraft production, thus leading to significant contributions to non-destructive evaluation 4.0.

Simulation of cosmic rays inside an aircraft: spectral perturbation and dose reduction due to aircraft structures and contents.

In this study, an intricate combinatorial geometry model of a Boeing 777-300ER aircraft was constructed for Monte Carlo transport simulations. The aircraft-induced perturbations of the energy spectra and effective doses of secondary cosmic rays at a typical civil aviation altitude (10km) were investigated on a component-by-component basis, which included neutrons, protons, photons, electrons, positrons, muons and charged pions. Two geomagnetic cutoff rigidities (1.35 and 15.53 GV) and two solar modulation parameters (430 and 1360 MV) were considered in the aforementioned simulations. The characteristics of various cosmic-ray components at six locations along the fuselage were assessed and compared with those of an unperturbed radiation field in the atmosphere. Aircraft structures and contents reduced the effective doses of personnel inside the aircraft to varying degrees, up to an ~32% reduction in the middle section of the passenger cabin. On average, the dose reduction was ~12-16% depending on geomagnetic and solar conditions. Quantifying the aircraft's self-shielding effects can further improve the estimation accuracy of aircrew and passengers' exposure to cosmic radiation. Information regarding the perturbed energy spectra of cosmic rays may be useful for designing onboard experiments or analyzing onboard measurement data.

Principle and Method for Determining the Calendar Safety Life of Aircraft Structural Protection Systems

The calendar safety life of the surface protection system in aircraft structures is the time limit for it to be used without functional failure at a certain level of reliability and confidence. The reliability of such protection systems and the operational safety and economy of the structure are closely related. This paper firstly establishes two methods for determining the calendar safety life of aircraft structural protection systems under a single service environment and in multiple service environments. A method for determining the reliability of the calendar safety life of the aircraft structural protection system was proposed, and an expression of the relationship between the maintenance costs for the aircraft fleet and the reliability of the calendar safety life of the aircraft structural protection system based on the relationship between the amount of corrosion damage to the structural substrate and the corrosion time and the expression of the calendar safety life of the protection system was established. Finally, taking a hypothetical aircraft fuselage wall plate connection structure as an example, an alternating corrosion fatigue test with protection system specimens was carried out. The process for determining the calendar safety life of the structural protection system and its reliability are given. This method is important to ensure the safety of aircraft structures, improve the efficiency of use, and reduce maintenance costs. Generally speaking, the reliability of the calendar safety life of the structure is 99.9%, and after the analysis in this paper, the reliability of the structural protection system is about 70%.

Subsurface impact damage imaging for composite structures using 3D digital image correlation

An integrated system is proposed to visualize subsurface barely visible impact damage (BVID) in composite structures using three-dimensional (3D) digital image correlation (3D DIC). This system uses a pair of digital cameras to record video frames in the field-of-view (FOV) of the structure’s surface, capturing the wavefield generated via chirp excitation in the near-ultrasonic frequency range. Significant pitfalls of previous efforts of damage imaging using two-dimensional DIC have been largely mitigated. First, 3D DIC enables capturing out-of-plane displacements, which are much larger in amplitude versus in-plane displacements that a single camera would be limited to sensing, thus increasing the signal-to-noise ratio. This enhancement in turn increases the sensitivity of the stereo-camera system. Second, a total wave energy (TWE) damage imaging condition is proposed to visualize the local damage region. The monogenic signal obtained via Reisz transform (RT) is employed to compute the instantaneous amplitude, with which the local wave energy can be calculated spatially over time. Since a high displacement amplitude and thus high wave energy will occur in the damage region due to the local resonance, the proposed TWE imaging condition can relax the Nyquist sampling requirement, unlike guided-wave-based structural health monitoring techniques which require fully reconstructing the wavefield and wave modes through sampling that satisfies the Nyquist criterion. As such, a much lower camera frame rate is adequate for the proposed system. Consequently, the maximum spatial resolution of the camera for a given FOV can be achieved at the expense of a reduced frame rate. With the maximized pixel resolution and reduced frame rate for employing the TWE imaging condition, composite structures can be inspected or monitored with a larger FOV. As a result, there is no longer any need to apply signal enhancement techniques, such as sample interleaving, image stitching, or averaging, to increase the effective performance of the camera. Rather than needing thousands of repeated videos for minimizing the incoherent noise, only a single stereo-video with a few seconds of sampling duration is necessary for damage imaging. The use of a powerful piezo-shaker also increases the wave signal amplitude and further enhances sensitivity without permanent adhesion. To demonstrate this stereo-camera concept with the TWE imaging condition, the system was used to image damage in two carbon fiber reinforced polymer composite honeycomb panels, which had been subjected to low-velocity impacts (2 J). For each panel, two excitation configurations were used to verify the robustness of the system. Initial damage maps produced for a 100 × 100-mm FOV using a three-second stereo-video show accurate damage imaging ability that is independent of excitation location and comparable to benchmark damage images computed from laser Doppler vibrometer data and those gathered from ultrasonic and X-ray computerized tomography scans. This efficient and reliable integrated system demonstrated high potential for in-time damage inspection on composite aircraft and other critical structures.

Modified Voce-Type Constitutive Model on Solid Solution State 7050 Aluminum Alloy during Warm Compression Process

The 7050 alloy is a kind of Al-Zn-Mg-Cu alloy that is widely used for aircraft structures. Although the deformation behavior of the solid solution state 7050 aluminum alloy is critical for engineering and manufacturing design, it has received little attention. In this study, the room and warm compression behavior of the solid solution-state 7050 alloy was researched, and a modified model with variable parameters was built for the flow stress and load prediction. The isothermal compression tests of the solid solution-state 7050 alloy were performed under the conditions of a deformation temperature of 333–523 K, a strain rate of 10−3–10−1 s−1, and a total reduction of 50%. The strain-stress curves at different temperatures were corrected by considering interface friction. The flow stress model of aluminum was established using the modified Voce model. For evaluating the modified Voce model’s prediction accuracy, the flow stresses calculated by the model were compared with the experimental values. Consequently, for assessing its prediction abilities in finite element applications, the whole compression process was simulated in the finite element analysis platform. The results sufficiently illustrated that the modified Voce-type model can precisely predict the complex flow behaviors during warm compression. This study will guide the prediction of the warm compression load and the optimization of the heat treatment process of the alloy.

An Efficient Procedure for Bonding Piezoelectric Transducers to Thermoplastic Composite Structures for SHM Application and Its Durability in Aeronautical Environmental Conditions

Piezoceramic transducers (PCTs) bonded to carbon fiber-reinforced plastic (CFRP) composite structures must be durable as well as remain properly bonded to the structure in order to provide reliable data for accurate guided-wave-based structural health monitoring (SHM) of aeronautical components. The current method of bonding transducers to composite structures through epoxy adhesives faces some shortcomings, such as difficult reparability, lack of weldability, longer curing cycles, and shorter shelf life. To overcome these shortcomings, a new efficient procedure for bonding the transducers to thermoplastic (TP) composite structures was developed by utilizing TP adhesive films. Application-suitable TP films (TPFs) were identified and characterized through standard differential scanning calorimetry (DSC) and single lap shear (SLS) tests to study their melting behavior and bonding strength, respectively. Special PCTs called acousto-ultrasonic composite transducers (AUCTs) were bonded to high-performance TP composites (carbon fiber Poly-Ether-Ether-Ketone) coupons with a reference adhesive (Loctite EA 9695) and the selected TPFs. The integrity and durability of the bonded AUCTs in aeronautical operational environmental conditions (AOEC) were assessed in accordance to the standard Radio Technical Commission for Aeronautics DO-160. The AOEC tests performed were operating low and high temperatures, thermal cycling, hot-wet, and fluid susceptibility tests. The health and bonding quality of the AUCTs were evaluated by the electro-mechanical impedance (EMI) spectroscopy method and ultrasonic inspections. The AUCT defects were created artificially and their influence on the susceptance spectra (SS) was measured to compare them with the AOEC-tested AUCTs. The results show that a small change occurred in the SS characteristics of the bonded AUCTs in all of the adhesive cases after the AOEC tests. After comparing the changes in SS characteristics of simulated defects with that of the AOEC-tested AUCTs, the change is relatively smaller and therefore it can be concluded that no serious degradation of the AUCT or the adhesive layer has occurred. It was observed that the most critical tests among the AOEC tests are the fluid susceptibility tests, which can cause the biggest change in the SS characteristics. Comparing the performance of the AUCTs bonded with the reference adhesive and the selected TPFs in the AOEC tests, it was seen that some of the TPFs, e.g., Pontacol 22.100 outperforms the reference adhesive, while the other TPFs have similar performance to that of the reference adhesive. Therefore, in conclusion, the AUCTs bonded with the selected TPFs can withstand the operational and environmental conditions of an aircraft structure, and hence, the proposed procedure is easily installed, reparable, and a more reliable method of bonding sensors to aircraft structures.

A low-power Bluetooth-based wireless sensor network and its global confliction-solving impact localization method

During the whole service lifetime of aircraft structures with composite materials, impacts are inevitable and can usually cause severe but barely visible damages. Since the occurrences of impact are random and unpredictable, it is a hotspot direction to develop an online impact monitoring system that can meet strict limitations of aerospace applications including small size, light weight, and low power consumption. Piezoelectric (PZT) sensor, being able to generate impact response signals with no external power and cover a large-scale structure with only a small amount of them, is a promising choice. Meanwhile, for real systems, networks with multiple nodes are normally required to monitor large-scale structures in a global way to identify any impact localization confliction, yet the existing studies are mostly evaluated with single nodes instead of networks. Therefore, in this paper, based on a new low-power node designed, a Bluetooth-based digital impact monitoring PZT sensor network is proposed for the first time with its global confliction-solving impact localization method. Evaluations of the system as a network are researched and analyzed on a complex real aircraft wing box for a global confliction-solving impact localization, showing a satisfying high accuracy.

Comparison of Fatigue Life of AA2050-T852 and AA7050-T7452 Alloy Forgings at Different Orientations

Abstract This study compares the fatigue performance between the AA2050-T852 alloy and AA7050-T7452 alloy traditionally used in aircraft structures. Rotating beam fatigue tests were performed on AA2050-T852 and AA7050-T7452 specimens cut from forged plates with different orientations and at different thickness levels. The samples were tested at 345 MPa (50 ksi), 276 MPa (40 ksi), 207 MPa (30 ksi), and 172 MPa (25 ksi). Scanning Electron Microscope (SEM) images showed mixed-mode failure in all samples. However, the equiaxed dimples were more abundant and finer in the AA2050-T852 alloy. The results indicated that AA2050-T852 alloy demonstrated better fatigue life performance compared to AA7050-T7452 alloy at almost all stresses. No significant difference in fatigue anisotropy was found in one alloy versus the other.

Quantification of hidden corrosion in aircraft structures using enhanced D-Sight NDT technique

In this paper, the enhancement of the D-Sight non-destructive testing technique was proposed by enriching it with a possibility of quantification of hidden corrosion typical for metallic fuselages of aircrafts. The proposed approach is based on the developed image processing technique, which allows the quantifying of the spatial dimensions of a corroded area from D-Sight images, which is supported with finite element modeling to achieve information about the depth and degree of corrosion of a given structure. The D-Sight inspections were validated using metrological measurements and the digital image correlation technique to introduce correction factors for the evaluations of dimensions of a corroded area using D-Sight due to overestimations typical of this technique. The performed studies allowed determination of a lower bound of the detectability of hidden corrosion using the D-Sight technique. The proposed enhancements will improve inspections of aircraft structures, where hidden corrosion remains a serious and unsolved problem.

The Effect of Rivet Arrangement on the Strengths of Lap Joints and Lap Joint Design Methods

To address the impact of rivet arrangement on the strengths of riveted lap joints, the failure modes and failure mechanisms of riveted lap joints were first studied using finite element analysis software. Next, the effects of the number of rivets, rivet rows, rivet arrangement, and row spacing on the lap joint strength were studied using the peak load as the evaluation index. Then, we proposed the concept of line load density to solve the problem that a varying rivet spacing and rivet edge distance will change the width of the sheet and thus the maximum load capacity, which is used as an index to study the effect of rivet spacing and rivet edge distance on the lap strength. Finally, a spring–mass model was developed to study the forces present in multi-row riveting. The model could accurately calculate the force–displacement curves during tensioning. The results show that when multiple rivet rows are used, higher stress concentrations cause the plate to first fracture at an outer rivet row with more rivets; therefore, the rivets should be arranged such that there are more rivets in the middle and fewer rivets on both sides. When the total rivet strength is greater than the remaining strength of the plate, the numbers of rivets and rivet rows have limited effects on the lap joint strength; however, this primarily affects the damaged form of the lap joint member. When the rivet spacing is less than 5d, the lap strength increases with increases in the rivet spacing, and when the rivet spacing is greater than 5d, the lap strength does not change significantly with increases in the rivet spacing. When the rivet edge distance is less than 3d, the lap joint strength increases with increases in the rivet edge distance, and when the rivet edge distance is greater than 3d, it has a limited effect on the lap joint strength. The rivet row spacing has no significant effect on the lap joint strength. The results of this study are valuable for improving the strengths of riveted structures in aircraft.

The impact of periodic axial loads on nonlinear dynamic instability behavior of Inconel 625 pipes

In various engineering fields like aerospace and aircraft structures or marine and offshore platforms, constitutive material of critical components should be made of specific materials that can work properly in the required workspace. Such materials must have excellent properties such as high mechanical strength as well as great resistance to corrosion, oxidation, and creep. Inconel 625 is a superalloy that is chosen as constitutive material of critical components due to its great abilities. On the other hand, since investigating Inconel 625 pipe has not been done yet, different mechanical characteristics of using structures made of Inconel 625 should be assessed. Additionally, doing so would be necessary to gather information for current industrial affairs and also future investigations. Therefore, the nonlinear dynamic instability response of axially loaded Inconel 625 pipes is investigated in the current article. The pipe structure is modeled via the Donnell shell theory and nonlinear von Kármán theory. The motion equations of pipes are established by applying the Hamiltonian approach. Then, in order to alter the nonlinear derived partial differential equations into the Mathieu-Hill equation, both Navier's solution and Airy stress function are implemented. Additionally, the amplitudes of steady-state oscillation of the Inconel 625 pipe are determined by employing Bolotin's method. Eventually, the impacts of various effective parameters on the nonlinear dynamic behaviors of Inconel 625 pipes are evaluated. The results indicate static and dynamic load factors possess a remarkable effect on instability exciting areas and steady-state vibration amplitudes of pipe. Moreover, the dynamic instability response of the pipe is dependent on the radius-to-thickness and length-to-radius ratios, and also how the ratios are affected depends on the wave number.

Assessing the Life Cycle Environmental Performance and Economic Costs of Composite Materials in Certain Aircraft Structure

The aviation sector is looking to replace conventional metals for aircraft doors with composite materials due to the latter’s potentially favorable combination of mechanical properties and low weight. However, little is known about the environmental impacts and economic costs associated with the production of such composite doors. This study conducts a Life Cycle Assessment and Life Cycle Costing on an example composite aircraft door to quantify its environmental and economic impacts. In addition, uncertainty analysis has been performed to enhance the quality of the assessment.

Research on the stress intensity factor of crack on the finite width plate with edge damage based on the finite element method

In aircraft structure maintenance, when we remove edge cracks of some structures, such as aircraft skin, or frame edges, the removal depth and radius affect the stress intensity factor of the crack tip, thereby affecting the service life of the structure. In this paper, we use the finite element method to calculate the crack stress intensity factor of the finite width plate and compare it with the engineering analysis method, which proves the effectiveness of the finite element method. We further research the influence of the crack stress intensity factor with different edge damage depths and radii of the finite width plate, which will provide a reference for aircraft structure edge damage tolerance analysis in engineering.

Integrating the Digital Twin of Decision Support Systems in Aeronautics

A Decision Support System (DSS) is a computerized system that helps users make decisions. In the field of air transport, RAL has developed DSSs to support decision making in various settings, including surface transportation and national security. The purpose of this chapter is to explore the maintenance hypothesis of conditional status checking and propose supporting concepts, such as enhanced care and proactive maintenance. These concepts are further enhanced by robust validation and strategies to improve the effectiveness of care in an extraordinary way. Therefore, a decision support system framework is necessary to build today's flying maintenance framework. Maintenance Decision Support Systems (MDSS) offer a valuable tool for aeronautical support, enhancing maintenance efficiency and cost control. However, further research is required to delve into the associated topics of flying MDSS. Due to the diverse configurations and intricate structures of modern aircraft, a significant number of aircraft parameters need to be evaluated. To facilitate further research, it is essential to establish systematic procedures for compiling a valuable database and define research criteria to assess the sufficiency of existing information. Considering the cost issues associated with continuous flight position information, determining the criteria for data download should guide further research.