A Substructure Synthesis Method with Nonlinear ROM Including Geometric Nonlinearities

Large flexible aircraft are often accompanied by large deformations during flight leading to obvious geometric nonlinearities in response. Geometric nonlinear dynamic response simulations based on full-order models often carry unbearable computing burden. Meanwhile, geometric nonlinearities are caused by large flexible wings in most cases and the deformation of fuselages is small. Analyzing the whole aircraft as a nonlinear structure will greatly increase the analysis complexity and cost. The analysis of complicated aircraft structures can be more efficient and simplified if subcomponents can be divided and treated. This paper aims to develop a hybrid interface substructure synthesis method by expanding the nonlinear reduced-order model (ROM) with the implicit condensation and expansion (ICE) approach, to estimate the dynamic transient response for aircraft structures including geometric nonlinearities. A small number of linear modes are used to construct a nonlinear ROM for substructures with large deformation, and linear substructures with small deformation can also be assembled comprehensively. The method proposed is compatible with finite element method (FEM), allowing for realistic engineering model analysis. Numerical examples with large flexible aircraft models are calculated to validate the accuracy and efficiency of this method contrasted with nonlinear FEM.

Design of sliding mode flight control system for a flexible aircraft

The evolution of large transport aircraft is characterized by longer fuselages and larger wingspans, while efforts to decrease the structural weight reduce the structural stiffness. Both effects lead to more flexible aircraft structures with significant aeroelastic coupling between flight mechanics and structural dynamics, especially at high speed, high altitude cruise. The lesser frequency separation between rigid body and flexible modes of flexible aircraft results in a stronger interaction between the flight control system and its structural modes, with higher flexibility effects on aircraft dynamics. Therefore, the design of a flight control law based on the assumption that the aircraft dynamics are rigid is no longer valid for the flexible aircraft. This paper focuses on the design of a flight control system for flexible aircraft described in terms of a rigid body mode and four flexible body modes and whose parameters are assumed to be varying. In this paper, a conditional integral based sliding mode control (SMC) is used for robust tracking control of the pitch angle of the flexible aircraft. The performance of the proposed nonlinear flight control system has been shown through the numerical simulations of the flexible aircraft. Good transient and steady-state performance of a control system are also ensured without suffering from the drawback of control chattering in SMC.