The landing of a civil transport aircraft is one of the most critical phases due to parametric uncertainties and strong crosswind conditions. In this paper, separate controllers are designed for longitudinal and lateral-directional channels for the landing phase, which is divided into the final approach, flare, and decrab. A multivariable model reference adaptive control scheme is implemented with state feedback for output tracking. The safety and flight performance of the autolanding control system are demonstrated through Monte Carlo simulations of a nonlinear civil transport aircraft model.

By detecting the severe meteorological situations on flight route, airborne weather radar (WXR) can ensure the safety of the aircraft and on-board personnel. Among these critical weather conditions, atmospheric turbulence is one of the main factors that affect flight safety. Atmospheric turbulence detection method that the current WXR adopts mainly is the pulse pair processing (PPP) method, which estimates Doppler spectrum width of weather target echo and compares it with a threshold to determine whether this weather target is turbulence or not. PPP method is simple and easy to implement, but the performance of this method under the condition of low signal-to-noise ratio (SNR) is poor. In this paper, we propose a new turbulence detection method based on the principal component analysis (PCA) approach. This new method uses PCA approach to preprocess the weather target echo and divides it into two parts: the principal component part as signal and the rest part as noise, so as to realize the de-noising function of PCA approach, and it is then combined with PPP method to estimate the spectrum width. Due to the good de-noising performance of PCA approach, this new method improves the detection performance of traditional PPP method especially under the condition of low SNR.

When helicopter is climbing from the ground or flying close to a solid surface, the flow around the lifting rotor can be strongly affected such that it will become time varying. To control the helicopter when the ground effect presents, a fast and accurate simulation technique is required. Traditional finite state inflow model with ground effect still need numerical integration since it cannot directly give prediction below the disk, and the computation of wake is necessary due to its interaction with ground. To improve the efficiency of the model, the adjoint theorem is adopted such that the wake can be solved with closed form equations. The result in hover condition is validated with Hayden model which is correlated well with the experimental data.

Increasing the environmental sustainability of aviation is a key design goal for commercial aircraft for the foreseeable future. From the perspective of structural engineering, this is accomplished through reducing the mass of aircraft components and structures. Advanced manufacturing techniques offer new avenues for design, enabling more complex structures which can have highly tailored properties. One advanced manufacturing concept is the use of 3D printed polymer preforms that are coated with nanocrystalline metal through electrodeposition. This enables the use of high-performance materials in virtually any geometry. To exploit this manufacturing approach, it is incumbent to have well-established mechanical models of the behavior of such hybrid structures. In particular, hybrid polymer–nanometal structures tend to fail due to compressive instabilities. This paper describes a model of local shell buckling, a typical compressive instability, as it applies to hybrid polymer–nanometal structures. The analysis depends upon the Southwell stress function model for radially loaded solids of rotation, and couples this with the Timoshenko analysis of local shell buckling. This combination is applicable to a range of practical configurations for truss-like hybrid structures.

The reduction of the aircraft weight is a great challenge for constructors, allowing them to cut down on fuel consumption and to increase the load carrying. This can be achieved by improving the properties of materials of aircraft parts. One of the most massive parts is the wing itself. In fact, according to a recent review of Aircraft Wing Mass Estimation Methods made by Dababneha and Kipouros (Aerosp Sci Technol, https://doi.org/10.1016/j.ast.2017.11.006 , 2017) an aircraft wing could reach 8811 kg (Megson in Aircraft structure for engineering students, 3rd edn. Arnold, London, (1999) for an Airbus A320-200 and even 39,914 kg (Anderson in Fundamentals of aerodynamic, 3rd edn. McGraw-Hill, Boston, 1999) for a Boeing 747-100 since it is a long-range aircraft. Thus, it seems very interesting to act upon the material choice of wing components but without harming structure stiffness. In fact, composite integration has a direct impact on the stress and the weight of the wing which matters a lot in the conception of an aircraft wing. Such a complex problem requires indeed a multidisciplinary approach based on aerodynamics and structural analysis of both composite material and aluminum alloy. Solving this problem could be of great interest for both industrialists and researchers since it would lighten the path towards full composite wing integration. In this work, we propose a double-approach numerical method to study the effect of different composite materials integrated into aircraft wing skin stiffeners. By this way, this investigation aims to show the improvement given by composite materials to the structure and the achieved mass reduction after cautious integration, taking into consideration the stiffness of structure after load application. It also gives a method based on the best yield on weight ratio to compare the different possibilities of integrations to choose the most relevant material, its stacking and fibers orientations. For that, we relied on both analytical calculation and finite element analysis. A Matlab code was developed to determine structure response to loading such as longitudinal and shear stress and to estimate the global structure’s weight for different materials. We started by working on a wing made from the aluminum alloy AA2024 and then we integrated various composite materials like Glass fiber, S-glass/epoxy, and hm graphite/epoxy. Also, an investigation of the variability of stacking sequences and plies orientations was undertaken. As a result, we obtained a clear assessment of the potential of each material. We finally created, inspired by the Airbus technical data, an accurate finite-element model of an aircraft’s wing that served to study the structural resistance with Abaqus. Coherent results for the selected composite material and for its ply’s stacking and orientation were found.

This paper investigates the flocking and obstacle avoidance problem of multi-quadrotor system with a lattice-type structure. To tackle nonlinear dynamics and cooperative consensus of a swarm of quadrotors, we proposed two algorithms (one for free flocking and one for constrained flocking) based on artificial potential field with two cooperative strategies using position tracking and attitude tracking, respectively. Firstly, the inner-loop control model of single quadrotor is provided. Then, we build an artificial potential function to act on the coordinate variable. Third, obstacle avoidance term is added to the flocking algorithm proposed previously. Finally, the position and attitude are taken as the coordinate variables, respectively, to both two algorithms, one for optimizing the distance of each quadrotor until the lattice-type structure is formed and keeping continuously and another for optimizing the distance between quadrotor and obstacles to avoid collision. Simulation results demonstrate the validity of both algorithms.

The application area and the scope of tasks of unmanned aerial system (UAS) are becoming increasingly wider, while the complexity and danger of the working environment are also increasing. When implemented in modern battlefields or complicated circumstances, various dangers which UAS is confronted with would prevent UAS from continuing its mission smoothly. An increase in failure rate is usually inevitable alongside with the perfection of UAS’s performance. In this paper, a modified modeling method for UAS is proposed to solve route planning problems in battlefields or complicated circumstances, and an improved modeling method of performance constraints is established based on analysis of UAS’s physical constraints. And based on the aforesaid research, new design of UAS’s each flight segment is proposed to meet the requirements for different missions. The simulation results show that such design makes the UAS capable to continue its mission.

The “fifth” and “fourth” generation fighters’ supersonic cruise flight comparative analysis approach is proposed and demonstrated. The IVth and Vth generation supersonic cruise flight parameters difference is shown. “Effective supersonic cruise flight” as characteristic feature of the Vth generation is proposed together with it’s criteria. “Supercruise” flight regime is investigated and it is shown that “supercruise” is not optimal from supersonic cruise range yet. “Supercruise” ability does not guarantee the supersonic performance advantage.

Current trends in increasing fuel efficiency and improving aircrafts flight characteristics place engineers before task of reducing the weight of onboard systems and decreasing power consumption with preserving required degree of reliability. In part of aircraft control system, the main stream is the actuators usage with power supply. In this article, three architecture options of the power part of the aircraft control system with different electrification levels are considered. The considered options of the power part of the control system are estimated by three criteria: reliability, mass, power consumption. The selected criteria allow for the preliminary conclusion about the prospects one or another aircraft control system architecture.

Flexible wings with a high aspect ratio own prominent nonlinear effects on the aeroelastic behaviors due to the geometrically large deformation and the nonlinear unsteady aerodynamic force referring to a strongly dynamic stall. This study proposes an efficient method to calculate the nonlinear flutter of the flexible wing in a fast way. First, the geometrically nonlinear deformation of the flexible wing under aerodynamic load is calculated based on a new beam model only having two generalized rotational freedom, which reduces the computing time drastically for engineering practice. Then the deformed flexible wing is regarded as a curved beam to considering the geometrical nonlinearity via the shape of applied curved beam. The curvature of the curved beam is fitted based on its static deformation. Thereby, the flutter of the flexible wing with large deformation can be simply solved by the linear vibration theory of the curved beam. Further, the ONERA dynamic stall model is used to describe the nonlinear unsteady aerodynamic force. In the end, the flutter behavior of the flexible wing is gained using the transfer function method for small computational cost. Illustrative examples demonstrate the validity of the present method through existing analytic solutions and published results.