Abstract:Morphing wings have become one of the important development directions of designing aircraft, and the field has received extensive attention. In this paper, the current status of research at home and abroad was explained, the deformation mechanism was summarised and the development trend was analysed from four aspects: variable span, variable sweep angle, variable chord length and combined deformation. According to the research status in this field and the application requirements of deformable wings, this paper proposed the key technologies to be solved urgently for morphing wings, including deformable skin technology, deformation mechanism, intelligent driver, sensor and control network. It was expected to provide a partial reference for the design and application implementation of deformable wings.

Abstract:The external-combustion wave rotor utilizes the unsteady pressure wave formed by the opening and closing of the ports to perform high-efficiency energy conversion, which has a wide range of application prospects. The typical structures of internal flow field of the external combustion wave rotor are analyzed, the global important results of performance tests are reviewed, and the advantages and challenges of topped gas turbine are summarized.It is pointed out that under the influence of port gradual opening and closing, circumferential clearance leakage, centrifugal force and Coriolis force acceleration, and solid wall heat transfer, the internal flow field structure presents multi-dimensional characteristics, and the leakage loss accounts for the largest proportion.Under model test condition, a pressure gain of 15～30% and a compression efficiency of 70～85% can be obtainedWhen used in conjunction with a gas turbine, it has the advantages of through-flow self-cooling, outstanding steady-state performance, excellent dynamic response and insignificant size effect. However, it is necessary to overcome the bottlenecks such as compact overall stricture design, high performance wave rotor design, wide operating condition wave system structure control, low-loss transition duct design and compound loss suppression.

Abstract:For the public and stakeholders, it is necessary to demonstrate airworthiness safety of AI-based avionics system in acceptable means of compliance, so as to solve problems of trustworthiness, uncertainty and explainability of civil aircraft avionics system caused by introducing artificial intelligence/machine learning technology. First of all, the classification and application status of AI-based avionics system are clarified, and the applicability of existing guidelines and standards are expounded. Then, based on the review of the current research progress, the implementation process and technical details of the certification framework for AI-based avionics system are summarized, including trustworthiness analysis, safety assessment, safety risk mitigation and certification/approval activities. Finally, it puts forward the requirements and implementation methods of the compliance verification of AI-based avionics system in each stage of the whole life cycle, and evaluates the impact of compliance verification on the existing airworthiness system. The study is significant to the design and certification of civil aircraft AI-based avionics system.

Abstract:Aeroelastic modelling and stability analysis of wing with morphing trailing edge have attracted more and more attention. In order to explore the methods and characteristics of active flutter suppression of wings with morphing trailing edge compared with conventional control surfaces, this paper provides a comparative study of active flutter suppression when the aircraft wing is equipped with morphing trailing edges and conventional control surfaces. The complete aeroelastic system of a small aspect ratio wing is modeled using finite element analysis, coupled with unsteady aerodynamics based on Doublet–Lattice Method. Morphing trailing edges and conventional control surfaces deflection modes are considered when generating the unsteady aerodynamic forces. The state-space system is generated by applying rational function approximation technique based on Minimum State method, and the active flutter suppression control law is searched by using the Linear–Quadratic–Gaussian method. The numerical study verified the feasibility of the proposed modeling approach. It shows that the closed-loop system with morphing trailing edges increases the critical flutter speed up to 22%, which is higher than with conventional control surfaces, and requires smaller deflection amplitude.

Abstract:MOEA/D-DE algorithm is easy to implement and widely used to deal with multi-objective optimization problems, while its hyper-parameters CR and F have a great impact on the performance of the algorithm. In this paper, based on the MOEA/D-DE algorithm framework, the Sobol global sensitivity analysis method is used to improve the cross control parameter CR in the differential evolution operator, and the Levy flight strategy is used to control the scale factor F. Then the hyper-parameters in the MOEA/D-DE algorithm obtain adaptive ability. And MOEA/D-DE algorithm with adaptive hyper-parameters (MOEA/D-DEAH) is proposed. The MOEA/D-DEAH MOEA/D-DE with different hyper-parameters settings and NSGAII are tested and compared with function tests and aerodynamic stealth optimization of airfoil. The function test results show that the new algorithm has good performance, and the algorithm can obtain strong robustness by adapting the super parameters. The optimization results of airfoil aerodynamic stealth show that MOEA/D-DEAH algorithm has better optimization results than other algorithms.

Abstract:The accurate reconstruction of civil aircraft gust load is of great significance to flight safety.A reconstruction method of dynamic gust load is studied based on the stochastic optimization algorithm. Firstly, based on the aero-elastic coupling equation in the frequency-domain, a time-domain analysis model of the dynamic gust load is established. Secondly, the profile of the gust field is parameterized by Hicks-Henne bump functions. In order to match the measured target values, the parameters of each bump function are optimized based on the stochastic optimization algorithm. Then, the reconstructed profile of the gust field is obtained, as well as the time history of the corresponding dynamic responses and loads. Finally, numerical simulations of the Generic Transport Aircraft (GTA) model experienced the 1-cos shape gust is carried out. The results show that the method established in this paper can be used to recognize the profile of the gust field, and reconstruct the gust dynamic loads. In the numerical examples, the maximum load factor error is 3.5%, and the maximum load error is smaller than 1%.

Abstract:The trajectory prediction model of traditional weapon control algorithm can not effectively predict the complex trajectory of drones.Besides, the existing research on complex trajectory prediction usually consider only one UAV, with the huge amount of calculation. To predict the trajectory of drone swarm quickly and accurately, a method of trajectory prediction for UAVs is proposed. After obtaining the track of drone swarm, it is firstly clustered based on DBSCAN method; Then the trajectory is determined to be simple or complex. Finally,Kalman Filter prediction was used for simple trajectory prediction, while LSTM（Long Short-Term Memory）network was used for complex trajectory prediction. The proposed model is more precise compared with the model using Kalman Filter, as well as consume less computing time compared with the model completely using LSTM network,which can meet the real-time requirement, as well as provide scientific basis for anti-UAV swarm calculation.

Abstract:In the process of robust controller design for aircraft engines, it is difficult to divide the flight envelope region systematically. In this paper, a flight envelope division method for aircraft engines based on thrust fuel consumption characteristics and dynamic pressure fuel consumption characteristics is proposed. According to the thrust, fuel consumption and dynamic pressure characteristics of a turbofan engine during steady-state operation within the full envelope, combined with the objective law of atmospheric conditions, the flight envelope is divided into 65 regions by two division methods, and the parameters of the surrounding small deviation regions and boundary points are replaced by the parameters of the corresponding nominal points of each region. By comparing the parameters of nominal points and boundary points within the full envelope of the engine, the results show that both methods are effective for the full flight envelope division, which provides a theoretical basis for the subsequent aircraft engine controller design.

Abstract:The lunar scientific research station is one of the key aerospace undertakings in my country at present, and it is very important for my country to win the future space race. At present, there is a blank in the field of lunar scientific research stations at home and abroad. The biggest challenge of its performance is that there are no cases and experiences for reference. The key point is that it needs to use the forward design concept in line with my country"s spacecraft development and a complete system engineering task analysis process. . In view of this, this paper adopts the idea of Model Based System Engineering (MBSE, Model Based System Engineering), and proposes a suitable model-based forward process for the system analysis of the lunar scientific research station, and uses the system model as a carrier to analyze the overall analysis of the mission and the mission requirements in turn. Analysis and task application scenario analysis. By carrying out the model-based mission analysis of the lunar scientific research station, the process of the mission analysis of the lunar scientific research station has been preliminarily realized. Development requires activation.

Abstract:It is significant to efficiently and accurately determine the flutter boundary of various aircraft configurations in the aircraft design. In this paper, adopt the linear unsteady aerodynamic method and nonlinear unsteady aerodynamic method for the subsonic and transonic areas, respectively. The nonlinear method uses the transonic small disturbance equation based on the accuracy steady CFD results as inputs. All the flutter solutions use g method for different Mach areas. The flutter analysis is used for an beam and reduced stiffness FEM model of the whole aircraft including empty aircraft and three special fuel cases for choosing the typical altitudes and Mach number covering the flight envelope and determine the flutter envelopes for the four aircraft configuration. The flutter boundary of the aircraft is consistent with flight flutter test. Compared with other analysis methods, the efficiency is obviously improved, especially for a variety of aircraft configurations. Therefore, the method adopted in this paper is an efficient and accurate method for predicting flutter boundary of large aircraft, which is suitable for engineering at present.

Abstract:It is significant to improve the simulation accuracy by improving the loading method to reduce the error between the numerical simulation and the test results. The ablation damage characteristics of composite laminates under different forms of lightning current were explored, the Braginskii arc radius expansion formula was improved, and a numerical model of arc expansion radius suitable for lightning strike process was established. A finite element model of CFRP laminate was established based on ABAQUS and the ablation damage under different current loading modes was analyzed. The results show that the finite element model can simulate the lightning strike ablation damage of CFRP laminates. Compared with other loading methods, the damage area, damage depth and damage volume caused by extended moving loading are in good agreement with the test results. The effects of current expanding and movement should be taken into account when simulating lightning strike ablation damage of composite materials. Keywords: composite; lightning strike; ablation damage; finite element model; current loading method

Abstract:With the development of multi-electric aircraft, the researches on the structure improvement and fault diagnosis of aviation transformer rectifier units (TRUs) are increasing, but research on the characteristics and effects of typical fault modes of different types of TRUs is lacked. Based on the analysis of TRU structures, the failure mode and effects analysis (FMEA) and fault tree analysis (FTA) of TRUs are studied to determine typical failure modes of TRUs. Then, test prototypes of various TRUs are established by Simulink, and the electrical signal characteristics of TRUs under different fault modes are analyzed and compared; the detailed fault modes of 8 TRUs are defined. The results show that the hard faults of TRUs have great impact on the system, and the high-pulse TRUs have better stability under fault modes and lower reliability, which provides a basis for fault diagnosis and fault location research of TRUs.

Abstract:Civil aircraft is assembled by various of and a large number of parts and components, which requires a clear classification of aircraft products in order to establish an efficient technology control mechanism throughout the whole lifecycle. This paper adopts system engineering methodology to identify the stakeholders of aircraft products. On the basis of analyzing and summarizing the demand of key stakeholders, the paper proposes a source-based civil aircraft product classification solution and criteria which is used to distinguish specified parts from equipment. A unified, standardized and multi-dimensional civil aircraft product classification and definition method has been formed. The case of certain type aircraft product classification indicates the proposed source-based civil aircraft product classification solution is convenient for engineers to quickly determine product categories in practical work, which ensures the accuracy and effectiveness of product classification, provides a basis for civil aircraft product design and technical management, and satisfies the application requirements of civil aircraft projects.

Abstract:Six Sigma DMAIC method was introduced to analyze the engine installation process in order to solve the problem of low first time yield of engine clearance size in aircraft assembly plant C. Multiple regression analysis and other methods were used to determine the key factors affecting the clearance size and their influence degree, and the optimal parameter range was obtained by linear fitting, and a detailed control plan was made. The analysis results of the control chart show that the improved clearance size is in a stable state of the process, and the first time yield of engine installation has increased from 65.2% before the improvement to 95%. Conclusion: Quality control based on Six Sigma DMAIC method has a positive effect on reducing production process variation and improving production efficiency.

Abstract:As the most critical and core part of the aircraft, the aero-engine has been serving in extreme environment such as high temperature and high load for a long time. Fatigue failure is one of the main causes of engine structure failure. With the development of industry, the problem of very high cycle fatigue(VHCF) of aero-engine materials becomes increasingly prominent. This paper summarizes the research status of VHCF of typical engine materials and describes the application situation of VHCF test technology. Including very high cycle axial vibration fatigue, bending vibration fatigue, torsional vibration fatigue, composite vibration fatigue test loading technology, temperature control technology and damage monitoring technology, The development of ultra-high cycle fatigue test technology for typical aero-engine materials in China is also prospected.

Abstract:It is significant to determine the equivalent static load of impact load for the strength design and verification of aircraft structures. The formula to calculate shock load dynamic effect coefficient is established basing on terminal peak saw-tooth shock pulse describing the time history of shock load, single freedom dynamic response model and the equivalent principle of model displacement. Then the function relationship between dynamic effect coefficient and the product of shock load duration and model natural frequency is deduced from trigonometric function inequality. The flow chart to determine equivalent static load of shock load is drawn according to the above theory. Finally this method is used to predict the equivalent static load of arresting hook system and aircraft model ditching process. The comparative results show that the maximal relative error of dynamic effect coefficient of arresting impact load is below 4% between theoretic estimated values and numerical simulated values by FEM, the relative error of equivalent static pressure during model ditching test is below 1.6% between theoretic estimated value and determined design value.

Abstract:To deal with the complex load conditions, diverse constraints and time-consuming problems faced during the design of the modular reconfigurable unmanned aircraft vehicle (UAV), taking the modular reconfigurable UAV in which different wings can be changed by using a common fuselage to meet different flying missions as the research object, different wings’ bending moment constraints application methods are studied in depth. A simultaneous multi-configuration optimization technique considering the wing’s front and back bars’ bending moment constraints is proposed. A typical modular reconfigurable UAV’s wing structure is designed and optimized by using the proposed method. Compared with the design results obtained by using the typical two traditional optimization formulations, the proposed technique is proven validate and convergent. And this method has some value in the practical engineering fields.

Abstract:Abstract：In order to complete the strength analysis and airworthiness verification of aircraft structure, it is necessary to carry out full-size static strength test of aircraft. For different load condition, the reliability and validity of test result are directly affected by whether the balancing design of test load can reflect the real stress condition of the evaluation area. In the full-size static strength test of aircraft, In ordor to simplify the complex loading scheme.Traditional balance loading is used to ensure the loading is accurate in the asseessment area, while the loading is simplify to some loading points in the non assessment area.In this paper, we find out that to the thin-wall structure like civil aircraft, traditional balacncing load in the non asseessment area have an effect on asseessment area,even if the the non asseement area is far away from the asseement area.Taking a front fuselage serious condition of a large civil aircraft as an example,the load design scheme of full-size static strength test of aircraft is studied based on the FEM model,By comparing the result of FEM analysis under theoretical design load and common test load, the optimized test load balancing scheme is put forward, the optimized blancing load is based on the real load in the condition we choose,and the balancing load in the non asseessment area doesn’t change the balace mode of the aircraft. The calculated results show that the internal force under optimized test load is basically identical with it under theoretical design load. According to the research in this paper, we can conclude that the Saint Venmant principl is inapplicable to the thin-wall structure like civil aircraft, in ordor to achieve the vertification purpose, the balancing load we exert should consistent with the real situation.

Abstract:Failure rate and average failure rate play a very important role in the reliability research of mechanical products. They and average residual life determine the life distribution and reliability of structures. Based on the basic index of reliability measurement, the rela-tionship between fatigue reliability coefficient and reliability was deduced when the life distribution is two parameter Weibull distri-bution. A method for calculating the structural failure rate and average failure rate was proposed based on fatigue reliability coeffi-cient, and the corresponding calculation process was derived. Based on the existing fatigue life test data of integral lap panel structure, the proposed method and the traditional integral method are used to calculate the average failure rate and error analysis of integral lap panel structure. The results show that the formula is simple, easy to calculate, and has the characteristics of high accuracy and relia-bility. It is of great significance for FMECA and maintenance decision-making of mechanical structure.

Abstract:The monitoring of hole edge cracks in metal structures with holes is of great significance for ensuring flight safety and enhancing aircraft reliability. In order to monitor the crack propagation length at the hole edge, the fatigue loading test was carried out on the aluminum alloy plate with hole edge pre-cracks, and the relationship curve between the crack length at the hole edge and fatigue loading times (a-N curve) and the data of the center wavelength of the fiber grating strain sensor changing with the crack propagation at the hole edge were obtained. The method of judging the crack propagation and penetration at the preformed corner by using the center wavelength data was put forward. Then, the envelope of the collected central wavelength data is extracted by extremum method, cubic spline interpolation method and wavelet analysis, and the data envelope is obtained. Finally, the relationship between the data envelope and the a-N curve is analyzed by BP neural network, and a monitoring model that can calculate the crack propagation length at the hole edge through the data collected by the fiber grating strain sensor is obtained, which lays a foundation for the health monitoring of the aircraft structure in the future.