Large Civil Aircraft Research Articles

Phased Array Ultrasonic Inspection of Composite Fuselage Panel

Abstract The composite fuselage panel is a crucial component in large civil aircraft, necessitating higher detection rates and efficiency. Phased array ultrasonic testing can enhance the efficiency and defect detection rate; however, it is essential to consider the structural characteristics of the fuselage to ensure reliable identification of design defects. Based on the material properties, structural features, and defect detection requirements of composite fuselage panels, a test block reflecting their characteristics was designed and fabricated with various prefabricated defects. Ultrasonic phased array method was employed for a series of testing parameter optimization experiments and defect detectability analysis. The detected flaws were quantitatively analyzed and compared with those obtained using conventional ultrasonic automatic scanning techniques to validate the effectiveness of the developed ultrasonic phased array method. The feasibility of utilizing phased array ultrasonics for nondestructive evaluation of wide-body airliner composite fuselage panels is confirmed.

Improved Multi-Body Dynamic Simulation of Landing Gear Drop Test Incorporating Structural Flexibility and Bearing Contact

The investigation of multi-body dynamics (MBD) modeling for landing gear drop tests is a hot topic in the realm of landing gear design. The current results were primarily focused on the multi-rigid body simulation or a simple multi-flexible body simulation, with little regard for the correctness of longitudinal loads and their experimental confirmation, particularly wheel–axle loads. Based on a genuine oleo-pneumatic landing gear drop test of a large civil aircraft, enhanced multi-body dynamics simulation research is carried out, considering the structural flexibility and bearing support by adopting flexible multi-bodies modeling and rigid-flex coupling contacts. When compared to the test data, which purposefully measured the longitudinal wheel–axle loads, the simulation results show that the loads, shock absorber compression, and shock absorber inner pressures are all within good agreement. Furthermore, the influence of structural stiffness and bearing contact was investigated by adjusting the model settings to confirm their importance.

Multiscale Progressive Failure Analysis for Composite Stringers Subjected to Compressive Load.

The fiber-reinforced composite stringer is commonly used in large civil aircraft wing structures. Under compression loads, it exhibits complex failure modes, with matrix cracking being one of the most common. The quantitative analysis of matrix failure is important and difficult. To address this issue, a multiscale method combining the generalized method of cells (GMC) and macroscopic FEM models is employed to quantitatively predict matrix damage and failure. The extent of matrix damage in the composite structure is represented by the number of failed matrix subcells within the repeating unit cells. The 3D Tsai-Hill failure criterion is established for the matrix phase, and the maximum stress failure criterion is applied to the fiber subcell. Upon meeting the criterion, the stiffnesses of the failed subcells are immediately reduced to a nominal value. In the current study, the ultimate loads, failure modes and load-displacement curves of composite stringers subjected to compressive load are obtained by the experiment approach and the proposed multiscale model. The experimental and simulation results show good agreement, and the multiscale analysis method successfully predicts the extent of matrix damage in the composite stringer under compressive load. The number of failed matrix subcells quantitatively evaluates the damage extent within a 2 × 2 GMC model. The findings reveal that matrix subcell failures primarily occur in the 45° and -45° plies of the middle part of the stringer composite.

Gust Load Alleviation Control Strategies for Large Civil Aircraft through Wing Camber Technology

Gust Load Alleviation Control Strategies for Large Civil Aircraft through Wing Camber Technology

Design and Research on the Refit Scheme for Natural Icing Flight Test of Large Civil Aircraft

Design and Research on the Refit Scheme for Natural Icing Flight Test of Large Civil Aircraft

Angular acceleration estimation and aerodynamic parameter identification based on angular velocity equivalent model

Conventional angular acceleration estimation methods generate non-negligible errors when the control surface rapidly deflects; hence, the estimated angular acceleration cannot be used for aerodynamic parameter identification directly. This paper proposes an angular acceleration estimation method based on angular velocity equivalent model. The angular acceleration of the angular velocity equivalent model, whose angular velocity response is consistent with the flight test data, is used as the estimated angular acceleration. Firstly, the equivalent angular velocity model is established by combining the generic aerodynamic model with the rotational dynamic equations of aircraft. Secondly, the angular velocity data and the control surface deflection data are synchronized. Finally, the angular velocity response of the equivalent model is used as the predictor, and the estimated angular acceleration is obtained by extended Kalman filter. The aerodynamic parameter identification and validation are carried out using the estimated angular acceleration of a large civil aircraft flight test data. The results show that the angular acceleration obtained through the proposed angular acceleration estimation method meets the accuracy requirement of aerodynamic parameter identification.

Research on icing scaling technology for large civil aircraft icing wind tunnel test

This article explores the icing scaling technology for large civil aircraft icing wind tunnel tests. Based on the typical critical icing conditions of large civil aircraft and the calibration of icing wind tunnel, and through the similarity analysis of water droplet trajectories, water collection, and energy balance, four similar parameters, namely modified inertia parameter, drop energy transfer parameter, stagnation freezing coefficient, and accumulation parameter, were selected as similarity criteria for icing scaling. Based on this criteria, icing scaling was conducted on the critical icing conditions of large civil aircraft, and the scaling results were evaluated by icing code. The results show that the water droplet collection and the ice shape before and after the scaling are in good agreement, and the similarity criteria used are appropriate. Finally, icing wind tunnel test was conducted in 3m × 2m test section using scaled icing conditions. The test results show that the main characteristics of the experimental ice shape are in good agreement with the calculated ice shape, and the calculated ice shape is more critical. The calculated ice shape can be used as a critical ice for aerodynamic wind tunnel tests and flight tests with simulated ice.

Analysis of landing gear walk vibration characteristics excited by aircraft anti-skid brakes

Currently, there is not enough research on the mechanism that causes the vibration of landing gear walk when an aircraft anti-skid brake is activated. This paper aims to provide a comprehensive analysis of the characteristics of the landing gear walk and the associated research. It shows that the walk vibration is caused by a cantilever beam’s tip mass and that the study should focus on the landing gear walk’s characteristics. Firstly, according to the transverse vibration theory of cantilever beams, the transcendental equation for calculating the natural frequency of the landing gear can be obtained. After that, the influence of the structural parameters of the landing gear on each order of the natural frequency is analyzed. To ensure that the conclusions derived are applicable to all landing gear structures, the structural parameters are selected to cover both small unmanned aerial vehicles and large civil aircraft. Finally, the influence of the axial pressure of the landing gear on its inherent frequency is analyzed.

Fluid-structure interaction dynamic analysis of large civil aircraft tank sloshing

To meet the requirements of passengers, the tank volume of large civil aircraft is gradually increasing, which brings certain risks. Under different 0 pertaining conditions during the flight, the liquid in the tank will cause structural failure and affect the aircraft’s safety. Therefore, based on the assumption of an ideal fluid, the smooth particle method (SPH) combined with the finite element method (FEM) is used to establish the coupling dynamic equation between the tank and the liquid. The effects of sloshing impact on the water tank under different filling rates were studied, and the influences of the water wave shape, the center of gravity shift, stress distribution, and deformation of the tank were obtained. The results show that the mass accumulation of water particles occurs in the acceleration stage, the relative center of gravity changes greatly in less water, and the accumulated liquid greatly influences the tank under the action of inertia. The more liquid in the tank is, the more the impact effect on the tank increases. The analysis results provide some reference for the analysis and design of tank anti-sway plates and flight attitude adjustment.

Research on ice shedding experiment for large civil aircraft to exit ice protection mode

While aircraft flying, once the icing detector finds the icing environment, the stall protection system will turn on the icing stall protection mode. After aircraft leaves the icing environment, due to the lack of effective means to determine whether there is still ice on the unprotected surface of the aircraft, it is impossible to exit the icing stall protection mode, which can lead to issues such as high approach speed of the aircraft, affecting landing safety and passenger comfort. To solve this problem, this paper conducts experimental research on ice shedding time under different temperature conditions, simulating the phenomenon of ice shedding during aircraft fight. This work helps formulate the logic of aircraft exiting icing protection mode, and lays a basis for future work of aircraft performance optimization.

An overview of icing scaling parameters for supercooled large drop icing conditions in icing wind tunnels

Aircraft icing is a serious threat to flight safety, especially the abnormal ice accumulation on the wing surface under the icing conditions of supercooled large droplets (SLD). Icing wind tunnel test is one of the main means of conducting research on aircraft icing, and the icing scaling law is the theoretical basis for selecting the similarity of characteristic parameters. In order to deeply understand the problems related to icing similarity in icing wind tunnel test under SLD icing conditions, the test capabilities on SLD icing conditions for the existing ice wind tunnel and the common similar parameter scaling methods were reviewed. The icing similarity transformation of SLD icing test was emphasized and the advantages, disadvantages, and applicability of each icing scaling law were analyzed. It can provide some references for the SLD icing wind tunnel test for large civil aircraft.

A Follow-up Loading System during the Retraction and Extension Process of a Large Civil Aircraft

This article introduces a design scheme for a distributed load follow-up loading system for large passenger aircraft flaps. The follow-up loading system includes tape lever components, follow-up pulley and drive components, force control actuators, fixed pulley and cable components, and support structures. At the same time, a loading control system was also designed. The dynamic loading system can simulate the deformation of flaps under distributed aerodynamic loads during the entire retraction and extension process, greatly improving the fidelity of the test environment and comprehensively assessing the interference between flaps and surrounding structures under all operating conditions.

Design of a large-scale model for wind tunnel test of a multiadaptive flap concept

The design and application of morphing systems are ongoing issues compelling the aviation industry. The Clean Sky - program represents the most significant aeronautical research ever launched in Europe on advanced technologies for greening next-generation aircraft. The primary purpose of the program is to develop new concepts aimed at decreasing the effects of aviation on the environment, increasing reliability, and promoting eco-friendly mobility. These ambitions are pursued through research on enabling technologies fostering noise and gas emissions reduction, mainly by improving aircraft aerodynamic performances. Within the Clean Sky framework, a multimodal morphing flap device was designed based on tight industrial requirements and tailored for large civil aircraft applications. The flap is deployed in one unique setting, and its cross section is morphed differently in take-off and landing to get the necessary extra lift for the specific flight phase. Moreover, during the cruise, the tip of the flap is deflected for load control and induced drag reduction. Before manufacturing the first flap prototype, a high-speed (Ma = 0.3), large-scale test campaign (geometric scale factor 1:3) was deemed necessary to validate the performance improvements brought by this novel system at the aircraft level. On the other hand, the geometrical scaling of the flap prototype was considered impracticable due to the unscalability of the embedded mechanisms and actuators for shape transition. Therefore, a new architecture was conceived for the flap model to comply with the scaled dimensions requirements, withstand the relevant loads expected during the wind tunnel tests and emulate the shape transition capabilities of the true-scale flap. Simplified strategies were developed to effectively morph the model during wind tunnel tests while ensuring the robustness of each morphed configuration and maintaining adequate stiffness levels to prevent undesirable deviations from the intended aerodynamic shapes. Additionally, a simplified design was conceived for the flap-wing interface, allowing for quick adjustments of the flap setting and enabling load transmission paths like those arising between the full-scale flap and the wing. The design process followed for the definition of this challenging wind tunnel model has been addressed in this work, covering the definition of the conceptual layout, the numerical evaluation of the most severe loads expected during the test, and the verification of the structural layout by means of advanced finite element analyses.

Correction to: From Bananas to Large Civil Aircraft: An On-Going Quest for Value-Added in the Computation of Countermeasures at the WTO

Correction to: From Bananas to Large Civil Aircraft: An On-Going Quest for Value-Added in the Computation of Countermeasures at the WTO

Numerical optimization of transonic natural laminar flow nacelles

Natural laminar flow nacelle is a promising technology for drag reduction. In this paper, an optimization platform is established for the design of transonic axisymmetric and three-dimensional natural laminar flow nacelles for large civil aircraft. The platform adopts the class/shape transformation method for geometric parameterization, a four-equation transition model for transition prediction, and the differential evolution algorithm combined with the radial basis function surrogate model as the optimization algorithm. The optimized axisymmetric nacelle demonstrates approximately 31% chord length of laminar flow, with the drag reduction of 13.3%. The influence of the Reynolds number and inlet mass flow rate on the optimization result is also investigated. The axis-symmetric nacelle optimization method is further used for the section profile design of a non-axisymmetric nacelle. An equivalent method is used to simulate the different local flow angles at different sections in the circumferential direction of the non-axisymmetric nacelle by using different inlet mass flow rates of the axisymmetric nacelle. The optimized natural laminar flow nacelle maintains over 30% chord length of laminar flow with robustness to the change of the freestream angle of attack. The total drag of the non-axisymmetric nacelle is reduced by 5.4% under cruise conditions.

Study on the Bearing Capacity of Stringer of Large Civil Aircraft Composite Fuselage after Impact Damage

This paper analyzes the post-impact pressure loss test data of eight test configurations of composite fuselage hat-type stringers and measures each composite stringer’s residual compressive strength. It compares the bearing capacity and failure strain of different configurations of stringers with and without BVID initial impact damage. Based on the test data and analysis, the optimal configuration of the stringer is selected, which plays a guiding role in the structural design of composite panels.

From Bananas to Large Civil Aircraft: An On-Going Quest for Value-Added in the Computation of Countermeasures at the WTO

ABSTRACT This article discusses the methodology used in the quantification of the level of authorized countermeasures in trade disputes settled under the auspices of the World Trade Organization (WTO)’s dispute settlement system. It discusses the legal framework pertaining to the quantification of the level of the remedies provided for under the WTO agreements in case of failure by the respondent to timely comply with the rulings and recommendations of the Dispute Settlement Body and where no agreement on compensation is reached: suspension of concessions and countermeasures. It analyzes whether the current understanding of ‘damage’ in WTO law is consistent with the prevailing economic reality and suggests an alternate interpretation that would account for value actually added on the territory of the complainant as a basis for quantification of the level of countermeasures. Finally, the article responds to ‘pragmatic’ arguments advanced by those who conceptually oppose to a methodological reform that would authorize countermeasures based on the share of value-added attributable to the complainant rather than on the gross amount of exports of the good affected by the impugned measure.

Research on Static Test Verification Technology of the Connection Between APU and Fuselage of Civil Aircraft

This paper studies the static strength verification of the connection between APU and fuselage of large civil aircraft, which provides a verification method for the airworthiness verification of the connection between APU and fuselage. Based on the internal force analysis results of APU finite element, this paper selects the static test load, completes the design from the theoretical load to the test load, designs a set of APU dummy loading device to introduce the load to complete the static test. The finite element analysis model is verified by the test results. The verified finite element analysis, limit load and limit load static strength test are used to show the airworthiness compliance of the static strength of the connection between APU and fuselage.

Diameter-thickness ratio influence on large civil aircraft fuse pins’ strength

The emergency break-away technology of civil aircraft is particularly important to the safety of aircraft, and the fuse pin is the key to the realization of emergency break-away technology. A type of civil aircraft pylon fuse pin was taken as the research object. Firstly, the finite element structure model of the fuse pin was established by ABAQUS, and the explicit dynamic algorithm was used to study the break-away process and load of the fuse pins. Secondly, the shear failure test of the fuse pin was designed to verify the accuracy of the finite element model. Finally, the influence of the diameter-thickness ratio on the failure mode and load of the fuse pins was studied. Comparing the failure load with the theoretical value, it can be concluded that the simulated and theoretical values of the failure load of the fuse pins were in good agreement under a certain range of diameter -thickness ratio, which provided references for the selection of the real fuse pins.

Aircraft Lateral-Directional Aerodynamic Parameter Identification and Solution Method Using Segmented Adaptation of Identification Model and Flight Test Data

Current aerodynamic parameter identification methods take the consistency between the response of the motion model reconstructed with the identified aerodynamic parameters and flight data as an optimization index. Since there is only one index constraint in aerodynamic parameter optimization, and more than ten aerodynamic parameters to be identified simultaneously have different sensitivity to the optimization index, the accuracy of identification results of parameters with low sensitivity is poor, especially for those with small absolute value. To solve these problems, a new idea, having aerodynamic parameters solved by balance equations and identified in stages, is proposed. Based on the motion characteristics of different lateral-directional maneuvers, the scheme of combined maneuvers is designed, and the step-by-step identification procedure is established. An estimation method of angular acceleration with angular velocity equivalent model is proposed to solve the problem that angular acceleration cannot be measured directly. A lateral-directional aerodynamic parameter identification method using segmented adaptation of flight data and identification models is established based on the identifiability of flight data. The lateral-directional aerodynamic parameter identification results of a large civil aircraft show the identification results can be obtained with higher accuracy by adopting the proposed identification method and the flight test scheme.