Composite Aircraft Structures Research Articles

Dynamic probability modeling-based inspection intervals optimization for civil aircraft composite structures using Bayesian updating

The evaluation of damage tolerance in composite materials is essential for ensuring the safety of aircraft structures. One of the most challenging aspects of applying probability modeling-based methods to evaluate damage tolerance is determining the actual damage size distributions for in-service aircraft structures. Although existing nondeterministic approaches have been used to optimize inspection intervals of composite structures, few studies have investigated the effects of updates on the actual damage size distribution and its impact on both the probability of structural failure and inspection intervals. This paper proposes a dynamic optimization method for inspection intervals of composite structures based on Bayesian updating. The damage size distribution of the composite structure is characterized by a general stochastic distribution. A Bayesian updating methodology is presented to iteratively update the actual damage size distribution whenever new data becomes available. Based on the constructed probability model, the inspection intervals of composite structures are determined under the objectives of optimal safety and economy for civil aircraft using a Monte Carlo approach. Compared to prior distribution models, the proposed method achieves higher safety for structures during a single inspection, reduces the failure probability of structures throughout their entire service life, and incurs lower maintenance costs. It also enables maintenance personnel to flexibly adjust inspection intervals while facilitating quantitative evaluation of both failure probabilities and maintenance costs associated with these intervals. These findings suggest that the proposed method holds great potential in enabling maintenance personnel to make informed decisions regarding inspection intervals for improved safety and economic performance.

Multi-frequency probabilistic imaging fusion for impact localization on aircraft composite structures

Since the internal barely visible damage of aircraft composite structures caused by the impact is a critical problem, impact monitoring is essential for the integrity and reliability of aircraft composite structures. This paper presents a multi-frequency probabilistic imaging fusion method for localizing impacts on aircraft composite structures. To capture the impact signals, a network of distributed sensors is mounted on the structure. The impact signals are then processed using the continuous wavelet transform (CWT) to extract the multi-frequency narrowband Lamb wave signals. The time difference of arrival (TDOA), a key feature of the impact source, is measured using averaging techniques employed in the normalized variance sequence. Subsequently, a probabilistic imaging function is established, and the TDOA of narrowband Lamb wave signals at each frequency is used as the feature input to generate the multi-frequency probabilistic imaging results. To determine the performance of the imaging results at each frequency, an efficiency index is introduced, allowing for the retention or abandonment of the imaging results. By utilizing the retained multi-frequency probabilistic imaging results, the proposed method achieves impact localization through imaging fusion. Experimental verification is conducted on a stiffened aircraft composite panel, and a comparison is made with two existing methods: the hyperbolic locus imaging method and the virtual time reversal imaging method. The results show that the proposed method can significantly improve localization accuracy compared to the existing methods, and is effective even in the presence of measurement noise.

Repair tolerance assessment for aircraft composite structures using Bayesian updating

Detection and repair of composite damage is crucial to ensure the safety and reliability of aircraft structures. A novel approach to quantitatively evaluate the repair tolerance of composite structures in civil aircraft based on Bayesian updating is presented. The method incorporates historical damage inspection data to determine the prior distribution of damage size, which is then updated with newly collected damage size data using Bayesian theory. Monte Carlo simulation is employed to investigate the probability of failure and estimate maintenance costs, considering various factors such as the frequency and timing of damage events, damage detection, structural strength, gust loads, and maintenance expenses throughout the lifecycle of composite structures. Safety and economic factors are considered to establish a lower threshold for repairs and an upper threshold for maintenance based on the occurrence of accidental impact damage. Verification of the effectiveness and feasibility of a quantitative assessment method for repair tolerance is conducted using damage statistics data from civil aircraft routes utilizing the structural skin panels of composite outer wing. The results demonstrate that the method proposed in conjunction with extensive simulations and full utilization of field damage inspection data can effectively simulate unexpected impact damage situations that may occur during civil aircraft service and evaluate the reliability and economic feasibility of the repair of structure. The research findings hold significant theoretical and practical value for the preparation of documents for continued airworthiness of composite structures, including structural repair manuals and maintenance programs.

Robotic compliant assembly for complex-shaped composite aircraft frame based on Gaussian process considering uncertainties

Robots are finding increasing application in aircraft composite structure assembly due to their flexibility and the growing demand of aircraft manufacturers for high production rates. The contact force of the composite frame in a robotic assembly of the aircraft composite fuselage panel can hardly be controlled due to the multi-surface variable contact stiffness caused by compliance and complex shape with multiple mating surfaces. The paper proposes a robotic assembly system for the aircraft composite fuselage frame with a compliant contact force control strategy using the Gaussian process surrogate model. First, a robotic assembly system is introduced, and the global coordinate system transformation model is built. Then, a compliant force control architecture is designed to generate the desired output force. Subsequently, a Gaussian process surrogate model with uncertainties is utilized to model the complicated relationship between the robot’s output force and the normal contact force acting on the mating surface of the composite frame. Furthermore, an optimal contact force control strategy is implemented to improve the contact quality. Finally, an experiment demonstrates that the proposed methodology can ensure that the contact force on each surface is within the limit of the engineering specification and uniformly distributed, improving the quality compared to the traditional assembly process.

A simple image correlation technique for imaging subsurface damage from low-velocity impacts in composite structures

A robust computer vision system is proposed to visualize subsurface barely visible impact damage (BVID) in composite structures through a simple image correlation technique together with a damage imaging condition. This system uses a digital camera to record a video of the surface motion, capturing micron-scale dynamic movement from guided waves propagating on the surface of the structure generated via a sweeping frequency excitation (chirp) up to the ultrasonic frequency range. As the excitation frequency changes during the chirp, waves become trapped within this damaged region, forming standing waves to generate local resonance at specific frequencies. This localized resonance accumulates high wave energy in the region, generating a higher transverse displacement at the damage site compared to the remaining part of the structure. In this work, a simple image correlation technique is proposed to correlate each filtered video frame with the temporal mean of the filtered wavefield video to highlight standing waves at localized damage. The proposed image correlation technique is distinct from digital image correlation (DIC) since it does not use correlation for subset matching to track the movement of surface patterns between two images (video frames). Instead, it uses correlation to directly quantify similarities between corresponding image pixels (windows). Utilizing a single camera greatly simplifies system complexity and hence enhances the practicality and potential for real-time performance over a recently developed technique that utilized 3D DIC with a stereo camera for vision-based BVID detection. To realize the aim, work was conducted in two sequential steps: (1) off-axis 2D DIC was employed rather than 3D DIC to examine the potential of employing a single camera to capture images and extract not only in-plane displacement but also a fraction of transverse displacement in which local resonance is dominant and (2) image correlation was then employed, supplanting the off-axis 2D DIC image processing involved in the first step, to highlight the damage with significantly less processing complexity. This proposed technique using a zero-mean normalized cross-correlation imaging condition, rather than the total wave energy used for DIC-based approaches, is efficient and effective for identifying regions of minute surface movement from local resonance within the damage region without the use of the computationally intensive interpolation to get the sub-pixel gray level information followed by subset matching employed in DIC-based image processing. Two geometrically identical CFRP composite honeycomb panels that had been subjected to low-velocity impacts were used for verification and validation, and two excitation location configurations were tested for each panel. Damaged images produced with correlation for a 100 mm × 100 mm field of view using a single 3-s video, or a total of 19,200 video frames, show accurate damage imaging capabilities regardless of excitation location that exceed that of DIC techniques and is comparable to benchmark damage images obtained from laser Doppler vibrometry, ultrasonic C-scans, and X-ray CT scans. This success of correlation-based imaging of subsurface BVID demonstrates substantial improvements in efficiency and practicality and shows high potential for in situ and real-time computer vision-based nondestructive inspection or structural health monitoring of subsurface BVID in composite aircraft and other critical structures.

Global-local optimization layout method for active piezoelectric sensor driven by structural critical areas

The optimal sensor placement is critical to ensure the integrity of the aircraft composite structures. The active piezoelectric sensors are crucial due to their ability to detect damage through guided wave (GW). However, the current optimization methods for piezoelectric sensor placement do not consider the simultaneous monitoring of the global and local critical areas of damage. Therefore, this study aims at performing the optimal placement of piezoelectric sensors by considering objective constraints and a combination of features that quantify the criticality of different parts of the structure and the global scanning ability of the GW. A novel global-local optimization model incorporating multiple constraints is proposed. It enhances the monitoring ability of the regional and global damage, and optimizes the placement of a limited number of piezoelectric sensors. The feasibility, accuracy, and practicality of the proposed approach are then verified using numerical and experimental cases. Furthermore, the performance of the proposed model is compared with that of public traditional models. The obtained results show the outperformance of the proposed approach.

Smart Patch repair solution for cure monitoring of bonded repairs in composite aircraft structures

Composite repair solutions are applied to recover the strength and functionality of composite structures in service without having to replace the whole structural part which is impractical for large sections. However, such bonded repairs are still not implemented in primary aircraft structures given the inability to provide repeatable results, ensure uniform bondline and the absence of nondestructive testing methods for detecting poor bond. In this paper, a dielectric-based technique is proposed to monitor the curing of composite bonded repairs based on Impedance Spectroscopy using smart patches embedded in the laminate by relying on drop-on-demand inkjet printed circuits which provided indicators of the polymer's curing stage whilst introducing minimal thickness in the bondline. The inkjet printed circuit is bonded to the composite plates using a newly developed procedure for high-level applications. The experimental campaign followed a building block approach to iteratively upscale the information from single printed IDTs in polymeric films to large circuits bonded to repaired composite parts with multiple tracks and IDTs. Numerous parameters were explored with the respective influence on the cure monitoring procedure of the thermoset adhesive layer, like the IDT design, the presence of imperfections at the bondline and instrumentation noise. Finally, the embedded smart patch was successfully employed for the detection of multiple impact damages and respective severity on the repair patch.

Impact force localization and reconstruction via gated temporal convolutional network

Techniques for identifying impact forces are crucial for improving the operational reliability and safety of aerospace composite structures. To address the challenges of theoretical modeling that arise from the complexity of composite structures, we propose a Gated Temporal Convolutional Network (GTCN) method. This method establishes inverse mapping relationships between structural vibration responses and impact forces. The GTCN increases the receptive field through dilated convolution, concurrently employing gating mechanisms to extract pivotal features from input signals in the vertical direction. Simulations and experiments are conducted to robustly validate the efficacy and applicability of the proposed impact force identification method. The simulation results demonstrate that the GTCN method is capable of obtaining robust impact force localization and reconstruction results, even amidst conditions of elevated noise levels. The average localization error of the GTCN is kept within 20 mm even at a noise level of 5 dB. Meanwhile, the GTCN method achieves satisfactory identification results for different types of impact forces (Gaussian, semi-cosine, and sawtooth), with an average localization accuracy above 95%. The applicability of the GTCN method to real-world structure is verified in the experiments. The average relative error of the reconstructed forces by GTCN is less than 15%. GTCN achieves stable identification of impact forces when the number of sensors is larger than 2. In both simulations and experiments, the GTCN performs better than Temporal Convolutional Network (TCN) and Convolutional Neural Network (CNN) with higher identification accuracy. The proposed method demonstrates considerable potential for application in the health monitoring of aircraft composite structures.

Effect of tightening process parameters on the T-F curve of bolts in composite structures

Aircraft composite structures are widely joined using bolts. Differences in the tightening process parameters for tightening bolts can significantly affect the tightening torque (T)-preload (F) relationship. Some of the more significant process parameters include lubrication conditions, tightening speed, tightening area, sealant and washer. The experimental study compared the differences in T-F curves, friction coefficients, torque coefficients and yield preload during the tightening process for different values of the process parameters. The focus was on the state of the thread surface, the state of the bearing surface, the diffusion of the lubricating medium and sealant with preload, etc. The distribution of the T-F curve was obtained for all possible values of the process parameters. As the tightening speed increases, both the inter-thread friction coefficient and the macroscopic increase in the coefficient of friction on the bearing surface are observed when tightening the bolt head. The presence of sealant reduces the inter-thread friction coefficient. The size of the T-F curve area reflects the sensitivity of the bolt tightening process to the various process parameters. The size of the T-F curve area reflects the sensitivity of the bolt tightening process to the various process parameters.

Delamination Behavior of CFRP Laminated Plates under the Combination of Tensile Preloading and Impact Loading.

When subjected to impact loading, aircraft composite structures are usually in a specific preloading condition (such as tension and compression). In this study, ballistic tests were conducted using a high-speed gas gun system to investigate the effect of biaxial in-plane tensile preload on the delamination of CFRP laminates during high-speed impact. These tests covered central and near-edge locations for both unloaded and preloaded targets, with the test speeds including 50 m/s, 70 m/s, and 90 m/s. The delamination areas, when impacting the center location under 1000 με, show a 14.2~36.7% decrease. However, the cases when impacting the near-edge location show no more than a 19.3% decrease, and even more delamination areas were observed. In addition, in order to enhance the understanding of experimental phenomena, numerical simulations were conducted using the ABAQUS/Explicit solver, combined with the user subroutine VUMAT with modified Hou criteria. The experimental and simulation results were in good agreement, and the maximum error was approximately 12.9%. The results showed that not only the preloading value but also the impact velocity have significant influences on the delamination behavior of preloaded CFRP laminated plates. Combining detailed discussions, the biaxial tensile preload enhanced the resistance to out-of-plane displacement and caused laminate interface stiffness degradation. By analyzing the influence of the preloading value and impact velocity on competing mechanisms between the stress-stiffening effect and interface stiffness degradation effect, the complex delamination behaviors of laminates under various preloading degrees and impact velocities at different impact locations were reasonably explained.

Life test result on 4 channel VCELs chip used in 28Gb/s data transfer in space application

For aeronautic applications, in-service Structural Health Monitoring (SHM) is a now established as a key technique to assess the performance and the integrity of composite aircraft structures for real-time in situ monitoring. For space industry, the Assembly Integration and Test including Electrical Ground Support Equipment is an essential tool enabling a satellite or space launcher developer to integrate and validate the electrical functions of its spacecraft before the launching process. In addition, satellite operators and worldwide SatCom manufacturers are willing to improve the on-board processing capacity, functionality and flexibility. In this context, opto-transceiver technology becomes a major building block regarding the highlighted aeronautic, space and ground telecommunication needs.In this context, the paper deals with the robustness evaluation of 850 nm GaAs VCSEL chip on Submount (CoS) used to 4 channel transceivers at 28 Gbit/s. The reliability investigation has been evaluated for two different manufacturers of VCELs and we perform the robustness of the technology regarding the final environment to define the limit usable for space applications. In this paper, we present the first technology of VCSELs. The main failure mechanisms have been defined to establish degradation laws associated. This last action is useful to define the ad equation between technology chosen and mission profile. These results are using for RYTHMS (ReliabilitY of opto-Transceivers for Health Monitoring Systems) Clean sky 2 project.

Dynamic instability analysis of timoshenko FG sandwich nanobeams subjected to parametric excitation

According to the nonlocal theory, the dynamic instability and vibration of functionally graded (FG) porous sandwich nanobeams resting on a visco-elastic foundation under axial harmonic load are studied. In this paper the Timoshenko and nonlocal continuum theory are employed to take shear deformation, rotational bending and small-scale effects into account. The governing equations of motion are extracted using the nonlinear Von-Karman, and Hamilton approaches. Then, to turn the partial differential equations (PDEs) into ordinary differential equations (ODEs) in the case of equations of motion, the method of Galerkin is employed, followed by the use of the multiple time scale method to solve the resulting equations. According to the results of this study, it can be claimed that the porosity ratio is more effective than porosity distribution and nonlocal parameters on the amplitude response and dynamic instability regions. Moreover, to examine thermal increments, foundation characteristics, slenderness, and thickness ratios, the bifurcation diagrams are drawn and discussed. It is found that foundation and geometric characteristics are of the highest significance in the nonlinear response of the Timoshenko nanobeams. The main intention of this study is to present a holistic idea about the porous materials used in sandwich structures which can be used in many composite structures in aircraft and automobiles.

Comparison of Requirements for Composite Structures for Aircraft and Space Applications

In this paper, the aircraft and space vehicle requirements for composite structures are compared. It is a valuable exercise to study composite structural design approaches used in the airframe industry and to adopt methodology that is applicable for space vehicles. The missions, environments, analysis methods, analysis validation approaches, testing programs, build quantities, inspection, and maintenance procedures used by the airframe industry, in general, are not transferable to spaceflight hardware. Therefore, while the application of composite design approaches from aircraft and other industries is appealing, many aspects cannot be directly utilized. Nevertheless, experiences and research for composite aircraft structures may be of use in unexpected arenas as space exploration technology develops, and so continued technology exchanges are encouraged.

Advanced Eddy Current Testing of Carbon Composites

Carbon composites due to their specific properties find applications in various industries, especially in aerospace industry. Widely used carbon fibre reinforced polymers (CFRP) have already been applied even for the aircraft primary structures. The development of advanced diagnostic techniques that are able to easily detect and identify the degradation of the carbon fibre material is still a challenge for various NDT methods. This paper describes the possibility to apply eddy current (EC) for testing of carbon composite structures. Two types of eddy current probes were developed and tested with excellent results. The new conventional eddy current probes are able to reliably and easily detect surface and subsurface discontinuities such as delamination and thickness variation. The probe setting parameters are described for different types of carbon composites (type of matrix and reinforcement, layup). Precise settings are necessary for the successful eddy current testing. It was determined that reliable detection of a minimum surface defect size is Ø1.5 mm for specimens and that eddy currents are able to penetrate into a thickness up to approximately 4 mm depending on the type of carbon composite. Additionally, this paper describes the comparison of eddy current testing with ultrasonic phased array method (PAUT). Composite aircraft structures are very susceptible to impact damage usually detected using PAUT. Therefore, sensitivity and resolution analysis of impact damage detection was performed by means of these two methods. It was found out that the results are very similar to each other when standard pulse-echo technique of PAUT was used. This carbon composite eddy current testing could fully replace or complement other NDT methods such as visual or ultrasonic inspections. Also it could guarantee the product quality in production, aircraft maintenance and in the development of new aircraft structures.

Ad-hoc solutions for ultrasonic inspection of highly complex aircraft composite structures

Nowadays, the ultrasonic inspection (UT) is the most common non destructive technique (NDT) for composite components in the aerospace industry, due to its high accuracy, reliability and the degree of industrialization. Although UT is widely developed, new manufacturing processes and part concepts are continuously pushing the technology for new improvements and applications. This is the case of the geometries resulting from novel out of autoclave composites manufacturing technologies, Liquid Resin Infusion (LRI) or Resin Transfer Moulding (RTM), allow the fabrication of parts in one shoot, with focus in bringing answers to the challenge of manufacturing fully assembled parts and avoiding secondary assembly phases but also limiting the accessibility for NDT inspections. In this sense, one of the critical issues in the aeronautic sector is the improvement of the inspection of geometries with difficult access and small dimensions. The requirement of the aeronautic industry of guaranteeing the quality of the primary composite structures implies the inspection of the whole component through certified technologies. This constraint, together with the lack of technological solutions for the inspection of regions with limited accessibility is blocking the industrial implementation of optimized manufacturing processes. The present work describes the design and development of adaptive hardware solutions for the single and phased array UT inspection of internal structures on a highly integrated composite aeronautic component. The final solution has been obtained as result of an iterative design process and manufactured through additive manufacturing technology (Powder Bed Laser Fusion, formerly Selective Laser Sintering) using polyamide. This device was validated in a for inspection of flat surfaces of stiffener internal structure, belonging to the composite flap demonstrators of lengths: 3 meters in a first approach, and 8 m for final validation. This novel inspection solution has been performed in the framework of the FLAP project to develop and innovate the manufacturing process for complex mobile surfaces in a collaboration between Aernnova, through its Composites Division, and CATEC teams.

Investigation on Failure Modes and Joining Of Aircraft Composite Structures

Composite materials are widely used in diverse applications, and extensive research have been performed to enhancethe understanding of the mechanical behavior of joined composite materials. It is mainly to develop design procedures for taking maximum advantage of their properties. Here in these joints the load is transferred between the disjointed components across the bonding surfaces of the joint. Composite joints have special characteristics which complicates its failure investigation. Through systematic scientific investigation, it is possible to recognize that the strength of the adhesively bonded joint is higher than that of the mechanically fastened joint. This paper presents investigations on joints and bonded materials in failure modes of joining of aircraft composite structures.

Infrared defect recognition technology for composite materials

This study mainly involves the methods and experiments of using infrared thermal wave imaging detection technology to detect internal defects in aircraft composite materials. The results were discussed and analyzed. In this paper, the feasibility of the experiment was verified by simulation. In simulation, the minimum accuracy of detectable defects is 4 mm radius under the mesh division accuracy with a correlation coefficient of 5. On this basis, an accurate detection method and prototype nondestructive testing system for defects of aircraft composite materials based on infrared imaging detection technology were designed, which can realize the identification and positioning of defects in aircraft composite material structures, including type, size and accurate depth of defects. Finally, the data collected by the infrared detection system was recognized through YOLO neural network. The test result shows the confidence level for water point defect is more than 0.9, while the confidence level for crack defect is about 0.5. This research result will expand the use case of infrared nondestructive testing technology around the world, and the transformation of the results will help to solve the maintenance problems of aircraft in general aviation.

Subsurface impact damage imaging for composite structures using 3D digital image correlation

An integrated system is proposed to visualize subsurface barely visible impact damage (BVID) in composite structures using three-dimensional (3D) digital image correlation (3D DIC). This system uses a pair of digital cameras to record video frames in the field-of-view (FOV) of the structure’s surface, capturing the wavefield generated via chirp excitation in the near-ultrasonic frequency range. Significant pitfalls of previous efforts of damage imaging using two-dimensional DIC have been largely mitigated. First, 3D DIC enables capturing out-of-plane displacements, which are much larger in amplitude versus in-plane displacements that a single camera would be limited to sensing, thus increasing the signal-to-noise ratio. This enhancement in turn increases the sensitivity of the stereo-camera system. Second, a total wave energy (TWE) damage imaging condition is proposed to visualize the local damage region. The monogenic signal obtained via Reisz transform (RT) is employed to compute the instantaneous amplitude, with which the local wave energy can be calculated spatially over time. Since a high displacement amplitude and thus high wave energy will occur in the damage region due to the local resonance, the proposed TWE imaging condition can relax the Nyquist sampling requirement, unlike guided-wave-based structural health monitoring techniques which require fully reconstructing the wavefield and wave modes through sampling that satisfies the Nyquist criterion. As such, a much lower camera frame rate is adequate for the proposed system. Consequently, the maximum spatial resolution of the camera for a given FOV can be achieved at the expense of a reduced frame rate. With the maximized pixel resolution and reduced frame rate for employing the TWE imaging condition, composite structures can be inspected or monitored with a larger FOV. As a result, there is no longer any need to apply signal enhancement techniques, such as sample interleaving, image stitching, or averaging, to increase the effective performance of the camera. Rather than needing thousands of repeated videos for minimizing the incoherent noise, only a single stereo-video with a few seconds of sampling duration is necessary for damage imaging. The use of a powerful piezo-shaker also increases the wave signal amplitude and further enhances sensitivity without permanent adhesion. To demonstrate this stereo-camera concept with the TWE imaging condition, the system was used to image damage in two carbon fiber reinforced polymer composite honeycomb panels, which had been subjected to low-velocity impacts (2 J). For each panel, two excitation configurations were used to verify the robustness of the system. Initial damage maps produced for a 100 × 100-mm FOV using a three-second stereo-video show accurate damage imaging ability that is independent of excitation location and comparable to benchmark damage images computed from laser Doppler vibrometer data and those gathered from ultrasonic and X-ray computerized tomography scans. This efficient and reliable integrated system demonstrated high potential for in-time damage inspection on composite aircraft and other critical structures.

Riveting damage behavior and mechanical performance assessments of CFRP/CFRP single-lap gasket-riveted joints

Riveting is one of dominant joining methods in carbon fiber reinforced polymer (CFRP) structures, however, riveting is prone to cause CFRP local damage. In this study, a riveting method with a gasket was developed for CFRP/CFRP joints. Riveting damage behaviors were assessed, and the parametric study of the rivet rod length was also carried out. Moreover, both tensile shear and pull-off performances were assessed under quasi-static loading, including the discussion of failure modes. Results demonstrated that the riveting damage mainly occurred in CFRP plies near the rivet deformed head due to the excessive deformation of the rivet tail along both rivet radial direction and rivet axial direction. The over long rivet rob caused a askew deformation of the rivet tail resulting in the most serious CFRP damage. Joining the CFRP/CFRP structures by using an appropriate gasket can obviously reduce the riveting damage and enhance the joining strength of the composite specimen. This paper provided a useful joining method for aircraft composite structures.

In-situ integration and performance verification of large-scale PZT network for composite aerospace structure

Piezoelectric transducer (PZT) based structural health monitoring (SHM) technology has been proved to be effective in increasing safety and reliability of composite aircraft structures. However, the attachment of PZT network to the host structure is considered as a weak link when facing the long term durability requirement of aerospace SHM, which should be overcome for aerospace SHM. In this paper, a surface-mounted co-curing method is creatively proposed to realize in-situ integration of large-scale PZT network and composite structure. By jointly controlling curing temperature and pressure, the proposed method can realize integration of large-scale PZT network and structural surface with high reliability and performance consistency. Compared with the conventional integration methods, the proposed method does not affect the manufacturing process or reduce the mechanical property of structure, and can be easily implemented whenever the host structure is available. To evaluate the effectiveness of the proposed integration method, a large-scale PZT network with an overall dimension of 1100 mm × 600 mm is integrated with a carbon fiber composite panel of wing box. The integration caused influence on functional integrity is first assessed by performing electro-mechanical impedance based theoretical analysis and experimental investigation. Then guided wave signals of the integrated large-scale PZT network are also acquired and analyzed, which proves the good signal repeatability and consistency of the network. At last, the SHM performance of the network is verified by conducting impact damages on the composite panel and performing damage monitoring, experimental results show that accurate monitoring is realized.