Pietro Ballarin

Structural bondings has become increasingly critical in aerospace applications, where composite materials are widely employed due to their high specific strength and stiffness. Mode II fracture behaviour remains challenging because of frictional effects between the delaminated surfaces and difficulties in crack tip identification; moreover, the ductile behaviour of adhesives produces large Fracture Process Zones (FPZ) which made critical mode II toughness evaluation. Finally, they are highly sensitive to manufacturing parameters, and the effects of curing pressure remain underexplored. The present study examines the mode II behaviour of co-bonded composite joints through experimental and numerical approaches. Two sets of glass fibre reinforced polymer laminates bonded with adhesive were manufactured using different curing cycles: a nominal autoclave pressure cycle and a reduced-pressure cycle. Mode II fracture properties were experimentally characterized by means of End Notched Flexure tests performed in the presence of differently developed FPZ. The different experimental responses were simulated using finite element models incorporating a Cohesive Zone Model with a parametrized traction-separation law. The optimal laws were identified through a response surface-based optimization procedure using Artificial Neural Network surrogate models trained on finite element simulations. In both configurations, a bilinear–exponential traction–separation law provided the most suitable representation of the interface behaviour. Despite the presence of an extended FPZ, particularly in reduced-pressure specimens, the identified cohesive laws accurately reproduced delamination propagation. The identified toughness values indicate a decrease of 19% of the toughness in reduced-pressure conditions.

In this work, a metallic additive manufactured bracket sensorized with optical fibres was developed with the capability of identifying the applied load magnitude and orientation. Additive manufacturing technology was exploited for the integration of sensors, exploring the possibility of producing internal and external channels for the housing of the optical fibres endowed with fibre Bragg gratings for strain and temperature sensing. The strain transfer between the metallic element and the sensors was studied numerically and experimentally, embedding the fibres into composite patches, co-bonded to the metallic element. A flap track support was selected as application scenario, which was printed in titanium alloy by selective laser melting. A two-step load identification method based on strain sensing was devised. The first model-based step adopts a notional finite element model for the extraction of the influence functions linking applied loads and local strains. The second step is a data-based correction procedure, which involves a limited number of calibration tests on the physical element and the identification of correction factors for the influence functions by means of a genetic algorithm. The description of the algorithm and a first validation by means of virtual tests is presented in detail. A demonstrator of the sensorized bracket was manufactured, proving experimentally the possibility of identifying the loads in conditions different from those used for the data-based correction. Overall, the study provides the methods for designing multifunctional structural connections that integrates structural and load monitoring roles and can be assembled into structures without modifying the conventional manufacturing procedures.

(1) Background: Structural Health Monitoring Systems (SHMSs) can reduce maintenance costs and aircraft downtime. However, their economic impact remains underexplored, particularly in cost–benefit terms. (2) Methods: This study conducted a targeted literature review on all the existing studies consisting of seventeen economic analyses of SHMS applications. Key features—such as SHMS type, structural material, vehicle type, integration stage, and cost elements—were classified to identify prevailing trends and gaps. (3) Results: The analysis revealed a predominance of piezoelectric-based SHMS applied to metallic fixed-wing aircraft, with limited attention to composite structures and e-VTOLs. Most studies focused on maintenance phase impacts, overlooking integration costs during manufacturing. Potential benefits like operational life extension, prognostic capabilities, and safety margin reduction were rarely explored, while critical drawbacks such as detection performance, reliability, and power consumption were underrepresented. Maintenance and fuel costs were the most frequently considered economic drivers; downtime costs were often neglected. (4) Conclusions: Although the majority of reviewed studies suggest a positive economic impact from SHMS implementation, significant gaps remain. Future research should address SHMS reliability, integration during early design stages, and applications to emerging aircraft like e-VTOLs to fully realize SHMS economic advantages.

The implementation of fly-by-wire technologies on rotorcraft requires the knowledge of the loads exerted by the pilot on the command chain during the flight; however, it is not easy to retrieve such information. In this work, a novel composite sensorized helicopter inceptor is developed aimed at monitoring the loads applied by the pilot during the flight for the entire helicopter lifecycle. The inceptor is produced using a soluble mandrel, and the connection elements are produced in polymeric material, allowing the connection of the optical fiber carrying the fiber Bragg grating (FBG) sensors to the interrogation system. Two prototypes of inceptor are manufactured, the first is produced with the aim to calibrate the Finite Element (FE) model of the component. Then the model is exploited to perform a sensor placement optimization based on the minimization of the condition number of the system using a Genetic Algorithm. The second prototype is produced positioning the FBG sensors according to the output of the optimization. Results show that a satisfactory load identification capability is obtained using only two sensors positioned according to the optimization technique. A comparison is also performed using the first prototype equipped with four strain gauges positioned using qualitative criteria: results show that a better load identification quality is obtained with only two sensors positioned according to the placement optimization technique. This work demonstrates the feasibility of a composite sensorized helicopter inceptor, characterized by a complex geometry, capable to provide satisfactory load identification quality with a limited number of sensors.

Monitoring the integrity of aeronautical structures is fundamental for safety. Structural Health Monitoring Systems (SHMSs) perform real-time monitoring functions, but their performance must be carefully assessed. This is typically done by introducing artificial damages to the components; however, such a procedure requires the production and testing of a large number of structural elements. In this work, the damage detection performance of a strain-based SHMS was evaluated on a composite helicopter rotor blade root, exploiting a Finite Element (FE) model of the component. The SHMS monitored the bonding between the central core and the surrounding antitorsional layer. A damage detection algorithm was trained through FE analyses. The effects of the load’s variability and of the damage were decoupled by including a load recognition step in the algorithm, which was accomplished either with an Artificial Neural Network (ANN) or a calibration matrix. Anomaly detection, damage assessment, and localization were performed by using an ANN. The results showed a higher load identification and anomaly detection accuracy using an ANN for the load recognition, and the load set was recognized with a satisfactory accuracy, even in damaged blades. This case study was focused on a real-world subcomponent with complex geometrical features and realistic load conditions, which was not investigated in the literature and provided a promising approach to estimate the performance of a strain-based SHMS.

Monitoring the integrity of aeronautical structures is fundamental for safety. Structural Health Monitoring Systems (SHMS) perform real-time monitoring functions, but their performances must be carefully assessed. In this work the damage detection performances of a strain-based SHMS, were evaluated on a composite helicopter rotor blade root. The SHMS monitored the bonding between the central core and the surrounding antitorsional layer. A damage detection algorithm was trained through Finite Element analyses performed on the component. The effects of the loads variability and of damages were decoupled by including a load recognition step in the algorithm. Load recognition was accomplished in different ways: version#1 adopted an Artificial Neural Network (ANN) trained on blades in pristine condition; version#2 was trained also on damaged blades; version#3 adopted a calibration matrix method. Anomaly detection, damage assessment, and localization were performed by using an ANN. Results showed a higher load identification and anomaly detection accuracy for version#1 and version#2 compared to version#3. Gaussian noise reduced anomaly detection and damage assessment performances, and reduced damage localization accuracy for bigger damages. The damage detection performance were obtained for a fibre optic based-SHMS, including the trade-off between probability of detection and false alarm rate for different damage sizes.

The economic impact of Structural Health Monitoring Systems based on optical fibre sensors is assessed in the development of composite helicopter rotor blades. Hence, the focus of this analysis is on the helicopter’s Beginning Of Life stage. Two applications of the Structural Health Monitoring System are considered in the development of composite blades: curing cycle development and accomplishment of laboratory and flight certification tests. Optical fibre sensors measure the temperature field during the curing cycle and strain field during the laboratory tests and allow load identification during the load survey activity. It was found that Structural Health Monitoring Systems can potentially lead to economic benefits during the development of the blade provide that a reduction in the number of curing cycles and number of blades tested is achieved as a consequence of the improvement of the temperature and strain field quality. Moreover, an economic benefit could be achieved during the load survey activity, needed to complete the certification of the composite blade, avoiding the periodical maintenance of the applied strain gauges acquiring the strains during the flight.

The paper presents technological, experimental and numerical studies aimed at the development of a morphing nozzle in composite material with a deformable divergent section, which is designed to control the vector of thrust in rockets without using mechanisms or localized flexible joints. The nozzle design is characterized by corrugated composite walls and longitudinal elements that make possible a separation of load paths involved in bending response, bearing of internal pressures and related force resultants. Mechanical requirements and conditions of pressure, temperature and gas velocity in the morphing part of the nozzle are defined by selecting and investigating two potential application scenarios. Thereafter, both technological aspects and the assessment of mechanical performances are addressed by designing, manufacturing and testing a composite demonstrator with morphing capabilities. A lightweight and flexible thermal protection system is proposed, designed and numerically analyzed. Finally, a virtual demonstrator of the composite morphing nozzle is developed. Results point out the possibility of manufacturing composite elements with axial load bearing capability that may undergo significant bending deformation without failures. The integration of a flexible thermal protection systems leads to a promising concept for the development of innovative morphing nozzle for rocket engine applications.