Structural health monitoring (SHM) is being widely evaluated by the aerospace industry as a method to improve the safety and reliability of aircraft structures and also reduce operational cost. Built-in sensor networks on an aircraft structure can provide crucial information regarding the condition, damage state and/or service environment of the structure. Among the various types of transducers used for SHM, piezoelectric materials are widely used because they can be employed as either actuators or sensors due to their piezoelectric effect and vice versa. This paper provides a brief overview of piezoelectric transducer-based SHM system technology developed for aircraft applications in the past two decades. The requirements for practical implementation and use of structural health monitoring systems in aircraft application are then introduced. State-of-the-art techniques for solving some practical issues, such as sensor network integration, scalability to large structures, reliability and effect of environmental conditions, robust damage detection and quantification are discussed. Development trend of SHM technology is also discussed.

No abstract

Fatigue crack growth in bolted metallic joints is one of the most serious and concerning types of damage in aircraft structures. The problem of how to quantitatively track the size of a crack in order to determine structural integrity has attracted much attention. In this paper, a novel eddy current array sensing film, with one exciting coil covering the entire thickness and several sensing coils distributed along the axial length of the hole, bonded onto the bolt, is proposed to quantitatively monitor a bolt-hole crack in the radial and the axial directions. Finite element simulation is utilized to investigate the eddy current disturbance caused by the crack growth and to optimize the coil configuration of both exciting and sensing coils in order to improve the crack monitoring capability of the sensing film. The simulation results show that the proposed sensing film has very good capability for tracking cracks when the traces of the exciting coil are designed to be similar to the sensing coils’ and have opposite current direction at the boundary of two adjacent sensing coils. An experiment is conducted to verify the feasibility and effectiveness of the proposed eddy current array sensing film to quantitatively track hole-edge crack growth in bolted joints.

Bolted joints are the key components for enduring primary load and play a vital role in ensuring structural safety for aircrafts. But it is very difficult to analyze the strength and failure modes of bolted joints due to their complexity and nonlinear coupling factors. Therefore, it is greatly essential to estimate potential failure mode at the early age and quantitatively track the damage parameters for calculating residual life and determining structural integrity. In this article, the concept of two-dimensional eddy current array–based sensing film is proposed to estimate the failure modes by identifying the circumferential location of the damage around the bolt hole and quantitatively track the damage parameters including the damage size in the radial direction and the damage depth in the axial direction of bolt hole. Finite element simulation is utilized to study the interaction between eddy current signals and the damage, and optimize the configuration of sensing film. Simulation result shows that a noticeable difference of eddy current signals of sensing film can be clearly seen that damage is located at different circumferential location around the hole. In addition, the exciting coil with opposite current directions at the boundary of axially adjacent coils has a better capacity of differentiating the damage sizes in the radial and axial direction than that with same current directions. The experiment is conducted to demonstrate the effectiveness of the proposed sensing film for estimating fatigue modes and quantitatively tracking damage growth.

Lamb wave-based damage detection for large-scale composites is one of the most prosperous structural health monitoring technologies for aircraft structures. However, the temperature has a significant effect on the amplitude and phase of the Lamb wave signal so that temperature compensation is always the focus problem. Especially, it is difficult to identify the damage in the aircraft structures when the temperature is not uniform. In this paper, a compensation method for Lamb wave-based damage detection within a non-uniform temperature field is proposed. Hilbert transform and Levenberg-Marquardt optimization algorithm are developed to extract the amplitude and phase variation caused by the change of temperature, which is used to establish a data-driven model for reconstructing the reference signal at a certain temperature. In the temperature compensation process, the current Lamb wave signal of each exciting-sensing path under the estimated structural condition is substituted into the data-driven model to identify an interpolated initial temperature field, which is further processed by an outlier removing algorithm to eliminate the effect of damage and get the actual non-uniform temperature field. Temperature compensation can be achieved by reconstructing the reference signals within the identified non-uniform temperature field, which are used to compare with the current acquired signals for damage imaging. Both simulation and experiment were conducted to verify the feasibility and effectiveness of the proposed non-uniform temperature field identification and compensation technique for Lamb wave-based structural health monitoring.