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Integration and test of piezocomposite sensors for structure health monitoring in aerospace
Measurement 147 (2019) 106861
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Integration and test of piezocomposite sensors for structure health monitoring in aerospace Filip Ksica a,⇑, Zdenek Hadas a, Jiri Hlinka b a b
Institute of Solid Mechanics, Mechatronics and Biomechanics, Faculty of Mechanical Engineering, Brno University of Technology, Brno, Czech Republic Institute of Aerospace Engineering, Faculty of Mechanical Engineering, Brno University of Technology, Brno, Czech Republic
a r t i c l e
i n f o
Article history: Received 30 October 2018 Received in revised form 13 June 2019 Accepted 22 July 2019 Available online 27 July 2019 Keywords: Structure health monitoring Piezocomposite sensors Non-intrusive testing Structural vibrations
a b s t r a c t The paper presents an innovative approach for Structure Health Monitoring utilizing piezocomposite plates as active sensors. The present work is motivated by the increasing interest in reduction of maintenance costs for aircrafts, which can be achieved by implementing self-sufficient sensor networks into the aircraft to predict risk states and prevent structural failures. These networks could supplement and possibly replace the expensive maintenance inspections. Piezoelectric materials have recently been under a rapid development, have been utilized in various fields of industry and due to their inexpensive and robust nature they can be used as sensors in non-intrusive Structure Health Monitoring networks. The presented approach uses piezocomposite sensors embedded into the structure of a small civil aircraft to directly measure structural vibrations. The functionality of the entire measurement chain was tested during an experimental flight, the acquired data will be used for further research of data-processing, impact detection and structure health prediction. Ó 2019 Elsevier Ltd. All rights reserved.
1. Introduction Over the past decades, air transport has become a rapidly growing industry, as indicated by the increasing amount of goods and passengers transported through the air. This leads to a necessity to increase number of aircrafts and the intensity of their usage, leading to more frequent and thorough maintenance procedures to ensure their safety matches the highest standards. This opens a challenging opportunity for research and development of new technologies that would help reducing operating costs and increasing effectiveness of the aircraft usage. Significant share of the operating costs of an aircraft are maintenance costs, and based on ICAO statistics, for example B757 the maintenance can form up to 23% of the flight (direct) operating costs and up to 12% of the total operating costs of the airline [1]. Typical maintenance procedure is based on a visual inspection of the overall state of the structure, with an addition of various non-destructive testing (NDT) methods to check the state of selected key components. This approach has been in use since times where aircraft structure was mainly from metal materials whose failures were in most cases visually determinable. However, for many parts of an aircraft metal materials have been substituted ⇑ Corresponding author. E-mail addresses: [email protected] (F. Ksica), [email protected] (Z. Hadas), [email protected] (J. Hlinka). https://doi.org/10.1016/j.measurement.2019.106861 0263-2241/Ó 2019 Elsevier Ltd. All rights reserved.
by more popular composite materials, that excel in terms of strength-to-weight ratio and can currently form up to 50% of the total airframe of modern large transport aircrafts [2], and even more for smaller aircraft categories. For composite structures, many failures, such as delamination, happen within the structure itself without being detectable visually, therefore pure visual inspections will become insufficient for both standard maintenance checks and post-repair diagnostics. While today’s aviation is still heavily dependent on visual inspections, a lot of work and effort has been dedicated to developing novel technologies used for non-destructive and automated damage detection, widely known under the title Structural Health Monitoring (SHM), that have high potential of supplementing or even replacing the abovementioned methods, and that can be successfully used for damage detection in tight spaces and locations otherwise hard to reach. Various SHM technologies are described in many research papers, there are also numerous publications summarizing the most important ones [3–6]. These new approaches for damage localization include methods based on propagation of Lamb waves [7–9], ultrasonic waves emitted and detected by a piezoelectric element array. If an obstacle, for example a crack, stands in the way of the waves, there is a change in the received signal spectrum. Another method utilizing piezoelectric sensors is based on pure acoustic emission [10]. This method can detect acoustic signal emitted by crack propagation,
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impact or other localized sources of damage. The evaluation is, however, very complicated due to high sensibility to noise generated during operation and is mostly viable in laboratory environment. Both methods are limited for relatively small areas in which we expect any damage to occur, but are being under development to cover larger sections of the aircraft structure. They are also sensitive to surface quality, environmental conditions such as humidity, temperature and ambient vibrations, making them relatively ineffective for long-term SHM during operation. Another SHM of composite structures is implementation of optical sensors with Fiber Bragg Grating (FBG) [11–13]. Network of optical fibres is embedded into the composite structure, deflection of localized damage translates into change in characteristics of the fibres and is detected. While still being under development, the abovementioned technologies form the mainstream SHM methods, and there are still many other technologies that could potentially form a nondestructive automated sensing network, could contribute to safety and supplement the conventional maintenance procedures. A novel approach utilizing piezocomposite sensors in active mode measuring structural vibrations directly, without the necessity to create a generator – detector network, is proposed. This paper aims to outline the procedure of utilizing piezocomposite sensors for measurement of structural vibrations in real-world conditions. For any novel technologies for SHM in aerospace, the process of their application and testing in real-world situations is limited due to very restrictive safety regulations. Each part of the sensoric network can represent a potential risk to the aircraft operation, therefore the new technology has to be certified before it can be implemented into the aircraft. Previously conducted laboratory tests provided a rough estimate of the capabilities of piezocomposite sensors for SHM of aircraft, but their capabilities in real-world conditions remained unknown. Therefore, for the purpose of this research, a small civil aircraft Cessna 172 was kindly offered for experiments with the piezocomposite sensors embedded within the structure, without any requirement for the tested system certification. It served for verification of laboratory test results and simulations, as well as obtaining information of the aircraft behaviour during various flight stages. It is believed that detailed overview of the experiment, the hardware used, and most importantly the results of conducted flight tests, are necessary for further development of SHM systems based on piezocomposite sensors, especially for modern aircraft predominantly constructed from composite materials. This approach can be used for monitoring of sudden, as well as long term changes in the structure, using spectral analysis of data measured over certain period of time, and advanced techniques (e.g. genetic algorithm) for data comparison can be used to detect and evaluate changes in the structure compared to the default response.
The overview of the planned experiment was presented on 2019 IEEE international Workshop on Metrology for Aerospace in Italy under the title Application of Piezoelectric Sensors for Structural Health Monitoring in Aerospace and was included in the workshop proceedings [14] without providing any further details or results of the conducted experiment. 2. Piezocomposite sensors: State of art Piezocomposite sensors are new technology invented and patented back in 1999 by NASA (US patent no. 6629341 – Method of fabricating a piezoelectric composite apparatus) [15]. These sensors are composed of thin piezoceramic rods with rectangular cross-section embedded within layers of epoxy matrix. Integrated electrodes are attached to both sides of the plate and sandwiched between polyamide films. They form highly flexible structures, being able to be used as both sensors and actuators. Compared to conventionally used piezoceramic sensors (piezoceramic plates), they are much more robust and flexible, have generally better electromechanical constants and do not need to be attached to conductive surface, because both electrodes are brought out through the polyamide film, as it can be seen in Fig. 1. If voltage is applied to the electrodes, it will bend and flex, distorting any material it is attached to and suppress ambient vibrations. If no voltage is applied, they form highly sensitive strain gauges or energy harvesting devices. Compared to the piezoceramic plates, they can be manufactured in wider spectrum of sizes and thicknesses. Being quite a new technology, piezocomposite sensors have not yet been used quite up to their potential. Up to date, they have seen some use in energy harvesting [17], being viable sources of electrical energy harvested from ambient vibrations of various structures. Due to their flexibility, they can be used as high deformation actuators in low and high frequency applications, or for active vibration control of flexible structures. In case of SHM, a few applications are worth mentioning, for example the abovementioned guided wave propagation technologies [18] and structure monitoring in civil engineering [19]. It is certain, that in the upcoming years they will spread into wider spectrum of sensing technologies, possibly substituting conventional piezoceramic sensors. As the piezocomposite materials are composed of piezoceramic fibres, also available separately in already polarised state, successful application of the commercially available piezocomposite sensors for direct measurement of structural vibrations can lead to further development of smart composite materials with embedded piezosensitive layers that can substitute individual sensors. Also, as most of the damage processes of composite materials happen within the material itself, either as failures of the individual layers or their interface, it would be very beneficial to embed the sensing
Fig. 1. Piezocomposite sensor, mode D33 manufactured my Smart Material [16]
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device/layer into the structure, not on its surface. The current state of their application in aerospace is relatively unexplored, as any novel system needs to go through extensive approval and certification processes before it can be tested on a real aircraft. Therefore, there have been attempts to develop accurate computational models of such materials in order to be used in virtual models of the target structures [20], but fully virtual approach is insufficient for accurate prediction of a complex structure of an aircraft. The physical principle of piezoelectric materials, as used in FEM, can be described with a coupled set of linear equations, one related to the mechanical part (Hooke’s Law) and the other to the electrical part (Maxwell’s equations),
h i fSg ¼ SE fT g þ ½dfEg T fDg ¼ ½d fT g þ ½e fEg T
where fSg is mechanical strain, fT g is mechanical stress, fDg is electrical displacement, fEg is electric field. Piezoelectric materials are characterized by several coefficients, the piezoelectric deformation constants ½d relating the applied mechanical stress to the generated charge on the electrodes, and the dielectric permitivity ½e, and the h i elastic constants of the material SE . These constants have to be either measured in laboratory environment or provided by the manufacturer. In the modelling process, it is important to keep in mind these equations are linear, but the piezocomposite material can be highly nonlinear, mostly due to hysteresis. Furthermore, piezocomposite material, compared to piezoceramic material, contains fibres and matrix, making the analytical and FEM modelling process even more difficult [20]. Therefore, it is beneficial to determine the electrical response experimentally for individual types of piezocomposite sensors and with respect to their application.
from undesirable environmental conditions (e.g. spars inside a wing, stringers inside a fuselage). Determination of these components and suitable location is key for proper function of the sensors, also for mounting into an existing aircraft, accessibility by the personnel has to be taken into account. 3.1. Experimental aircraft It is generally quite difficult to test any novel technologies on an existing aircraft because of very strict safety protocols. Previously untested system can be a potential risk for proper function of the aircraft, and to conduct a test in real-life operating conditions, either private aircraft or partnership is necessary. For the purposes of this work, a small civil aircraft Cessna 172 (Fig.2) was generously provided by one of our partners. It is a four-seat single-engine aircraft with high wing, almost entirely manufactured from sheet aluminium. 3.2. Determination of sensor locations In case of the Cessna 172, accessibility of load bearing components was the main issue. It had to be kept in mind that all the sensors have to be physically attached by wires to a data collection unit, and these wires have to go inside of the structure. Due to that, the only option for sensor placement was the fuselage, that is hollow for the most part and could be accessed via small service door (Fig. 3) on the right hand side of the aircraft or through the storage compartment. The load bearing components inside the fuselage are 7 stringers formed by bent overlaps of the aluminium skin, either L shaped or
3. Measurement of structural vibrations Piezocomposite sensors are an advantageous alternative to other conventional sensing devices due to their small and robust nature, no moving parts and very low manufacturing costs. They can be used for wide spectrum of frequencies and with high electromechanical constants, the output signals have high voltage (1 1000 V) and low current (1 10 lA) nature resulting in high SNR. For SHM purposes, they can be either attached to a surface or embedded within layers of composite material. In this work, they are used as sensors for direct measurement of structural vibrations, thus utilized in active mode (no power source required). As the piezocomposite sensors are sensitive to the change in strain, they have to be mounted on load-bearing structural components, preferably inside the aircraft to provide cover
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Fig. 3. Service door on the right-hand side of the aircraft.
Fig. 2. Photography of a Cessna 172, Source [21]
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C shaped. Two C shaped stringers situated on the bottom, one L shaped on the top and a pair of L shaped on each side (Fig. 4). The C shaped stringers were not suitable for piezocomposite sensor attachment due to their very high curvature and the top one was inaccessible due to the internal structure of the tail spine. For that reason, only the side stringers were a viable option for sensor placement. In the next step, longitudinal locations of the sensors were determined. To ensure sensitivity to all modes of load, i.e. tension, compression, bending and torsion, a simulation these loading modes was performed. A CAD model (Fig. 5) of the fuselage section was created and imported to ANSYS (Fig. 6) where it was meshed and subjected to unitary loads in separate analyses. The von-Mises stress values for the respective loads were measured along each of the side stringers. The loading modes, with respect to the global coordinate system (Z – longitudinal, Y – vertical, X – transverse), included tension in Z direction, bending in X and Y directions and torsion along Z direction. These values were normalized to fit into the interval of 0 1. The absolute values of the stress values were not important at that stage, because the Fig. 6. FEM model of the fuselage section.
aim was to determine locations with maximal values. Also, the real values of load during operation were not known at that point. In order to ensure sensitivity to all loading modes, each position along the longitudinal axis is described with a suitability index, that is normalized geometric mean of the individual normalized von-Mises stress values given by the following equation
hQ suitability index ¼
Fig. 4. Location of stringers inside Cessna 172 fuselage; left: side stringer (green line), right: bottom stringer (red line). (For interpretation of the references to colour in this figure legend, the reader is referred to the web version of this article.)
4 i¼1
rnorm;i
i
hQ i max 4i ¼ 1 rnorm;i
where rnorm;i represents the normalized von-Mises stress for each respective loading mode. The geometric mean assures that stresses (strains) for all modes are measurable at each location and as high as possible, excluding the situation that there are some modes that are not detectable. Furthermore, to reflect the fact the piezocomposite sensors measure an averaged value of strain along their entire length, an averaging filter with kernel length adjusted to reflect the final length of the sensor, in this case 1 sample per 1 cm of length. The resulting values of suitability index are illustrated in Fig. 7. The symmetry between left and right stringers was expected, as the model was fully symmetric along the longitudinal – vertical plane. While there are visible peaks on both ends of the stringers, mounting at these locations, the position of 0.5 m for all four stringers was determined as the most suitable and in Fig. 7 is marked by a dashed line. Not only the sensors formed a symmetric array that way, also it was close to the service door and accessibility and mounting was therefore much easier. 3.3. Experimental sensor network
Fig. 5. CAD model of the fuselage section; green: suitable stringers, red: unsuitable stringers. (For interpretation of the references to colour in this figure legend, the reader is referred to the web version of this article.)
After the suitable locations were determined, a preparation of the flight experiment took place. Two main requirements were kept in mind. Firstly, the sensing network should not interfere with the operating systems of the aircraft and must not form any potential risk if any part came loose. Secondly, as the sensors are mounted permanently to the structure, the system must be reduced to an absolute minimum after an experiment is conducted, also it should be reusable for any future experiments. The sensors used in the experiment were commercially available piezocomposite sensors polarized in mode D31, with embedded fibres with cross-section of 180 350 lm and total active area
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Fig. 7. Suitability index for sensor locations on individual stringers.
of 28 7 mm. Based on previous laboratory tests, to achieve the most consistent sensing capabilities of the piezocomposite sensors, the adhesive should have very high elastic modulus and should form as thin layer as possible. For that reason, a hardened highstrength structural epoxy adhesive was used. Also, it is important to mention that if conventional piezoceramic plates were used, the bottom electrode would have to be attached conductively either to the surface of the stringer or to an external wire placed between the plate and the underlying surface. However, it is best
to keep the sensing system galvanically separated from the structure of the aircraft, not only to prevent static electricity accumulated in the metal structure to affect the measurement, but also to ensure that no high voltage feedback from the sensing device gets back to the structure and any attached equipment (GPS module etc.), despite the expected voltage values being relatively small. The attached sensors shown in Fig. 8 form a bottom pair of the sensing array. To enable disconnection of the sensors after an experiment is conducted, a 3D printed enclosure for the connectors
Fig. 8. Mounted sensors (bottom pair).
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is attached next to the sensor, allowing it to leave only the sensor and a short wire to be left in the fuselage, posing marginally no threat to the operation of the aircraft. The enclosure also served another purpose. To prevent the wires breaking of the electrodes due to the mechanical vibrations, a short loop is formed to absorb any shocks along the wires. The wiring itself was looped around the fuselage using zip ties and clip brackets attached with highstrength double sided tape. The entire array of four sensors is illustrated in Fig. 9. The nominal maximum operating voltage for the used sensors is 350 V, but based on our laboratory tests and assumed values of strain, the voltage should not exceed 30 V even in extreme situations and during sharp manoeuvres, which were not expected nor approved to be performed during the test flight. Because it was not known beforehand what actual strain values and respective voltage values would be measured during an operation of the aircraft, an overvoltage protection and voltage divider (Fig. 10) had to be inserted between the sensors and the data collection unit that had only ±5 V analog inputs. The adjustable voltage divider was used to ensure the voltage values fit into an acceptable range and also served as load for energy harvesting, that was part of a different experiment (Fig. 11). From the voltage divider, BNC cables were used to connect the system to a data collection unit (DCU). The DCU was used to
Fig. 11. Data collection unit.
measure raw uncompressed voltage data with sampling frequency of 12,800 Hz and store them to a flash drive in a suitable file format. The A/D converter had resolution of 24 bits, giving us potential 16.7 million levels of resolution within the interval ±5 V. The experiment was therefore set to be on the conservative side, preventing any potential damage to the expensive hardware equipment. It was also necessary to capture any unexpected peaks. The DCU was specifically designed for this experiment to be triggered manually by an operator during flight, measuring specific parts of the test flight and enabling prompt troubleshooting in case that an unexpected event happens. Due to safety precautions, the data was stored in short sequences to achieve redundancy in measured data if any error occurred. The DCU was powered internally from a battery rather than being connected to the internal power supply of the aircraft, mostly to prevent any interference and increase mobility of the device and to fit all the components into one suitcase for easier manipulation and operation during flight (Fig. 11). It is important to keep in mind the sensor sensitivity is slightly different for each individual sensor, mostly due to inconsistencies in manual mounting, adhesive layer thickness, slight error in orientation, manufacturing tolerances etc. However, in long-term these inconsistencies can be monitored and compensated for. 4. Test flight
Fig. 9. Sensing network within the fuselage.
Fig. 10. Overvoltage protection and adjustable voltage divider.
After the entire network was constructed and tested on ground, a flight test was conducted. It took place in local airfield and the entire flight took around 20 min. During that time, over 15 individual measurements were taken, covering engine start-up, take-off, various parts of flight and landing. All four channels were measured and stored for upcoming analyses that took place back on the ground. Four clearly distinguishable segments of the test flight were chosen to illustrate the capabilities of the sensing network, namely engine ignition and movement on the airfield, take-off and ascension, steady flight and landing also with movement on the airfield, as illustrated in Fig. 12 (without engine start-up and field movement being shown). Length of the coloured arrows corresponds with the length of the measured segments. Raw measured voltage signals for the selected flight segments are shown in Fig. 13. As can be seen, the voltage values after the voltage divider, that was set to 1:20 ratio, peaked at around 0.2 V, voltage on the sensors therefore did not exceed 4 V, proving that our setting was conservative enough. The timestamp in the title of the plots is set to 00:00 at the take-off, that took place at around 00:20. In the first graph, the excitation around 20 s time mark was a failed attempt of engine ignition, a second and success-
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Fig. 12. Trajectory of the test flight with highlighted phases.
Fig. 13. Raw measured data for selected parts of test flight.
ful attempt took place at 30 s time mark. The fourth graph represents landing, touch took place at around 25 s time mark. The measured data was subjected to a spectral analysis in Matlab environment to assess the data further, distinguishing any dominant frequencies, such as engine frequency, to confirm the proper functionality of the hardware. The procedure of analysing the data using spectrogram is depicted in Fig. 14. In the first step, the measured data is separated into segments of fixed length, that is determined with respect to the maximum observable frequency required, such that there are at least 10 samples per period for each frequency. The segment length directly affects the sensitivity of the spectrogram, shorter segments means higher sensitivity to sudden changes in the signal at the cost of lower frequency resolution. With sampling frequency of 12,800 Hz and maximum observable frequency at around 600 Hz, using the equation below a segment length of 0.5 s is calculated.
window length ¼
ðmaximum observable frequencyÞ ðsampling frequencyÞ
ðsamples per periodÞ In the next step, the overall resolution of the Fast Fourier Transform (FFT) results is increased using technique of zero-padding, in which a zero array is added to the data segment. The length of the array should be chosen carefully, as it slows the process of FFT and
only up to a certain point the resolution enhancement is relevant. For our case, the suitable padding was with an array of 10 times length of the original vector. Certain signals may require using a window (e.g. Gaussian window) for the padded segment in order to ensure that the segment starts and ends with the same value to prevent artificial steps in the periodic signal constructed during FFT. In our case, this step was not necessary, as it had no significant impact of the final spectrogram. In the following step, padded data segments were subjected to a FFT and smoothed using averaging filter over 5 samples to slightly reduce the noise in the signal. In the last step, the resulting frequency-domain signals are stacked in time-domain to form a spectrogram. The results of the spectral analysis for characteristic stages of the flight, depicted only in the time-frequency plane, are shown in Fig. 15. Starting from the first figure, the failed engine start-up is evident at around 20 s. At 30 s, the engine start-up was successful and a dominant frequency of approximately 37 Hz is clearly visible. That frequency corresponds to the revolution of the rotor shaft, as the nominal engine speed was 2400 rev.min1 (40 Hz). The multiples of the frequency in the spectrum correspond to the rotor blade frequency (2 blades) and cylinder movement (6 cylinders, 2 pairs of 3). The rest of the lines most likely correspond with natural frequencies of the structure and other moving parts of the aircraft. It was expected that the engine frequency would be the most
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Fig. 14. Procedure of creating high-resolution spectrogram.
Fig. 15. Spectral analysis of the test flight data, first sensor, variable flight stages.
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Fig. 16. Spectral analysis of the test flight data, variable sensors, take-off stage.
dominant one, as the engine is the most significant source of vibrations within the aircraft. However, the measured data can be assessed from a different point of view, and that is comparison between individual sensors (Fig. 16). Based on the comparison between measured data from individual sensors during one stage of the flight, load of individual stringers could be assessed. For example, it is clearly visible that 3rd and 4th stringer (left hand side of the fuselage) were subjected to slightly higher amounts of strain than the other two, providing us an estimate of current loads on the dominant frequency. It is assumed that any localized disturbances (impact, sudden events etc.) or long-term damage development will somehow reflect in the measured response, either by shift in natural frequencies, amplitudes, or by visible wide-spectrum excitation. However, the samples obtained during one test flight are not enough to assess that and methods for thorough data analysis and SHM. When a long-term monitoring experiment is conducted, these data will also serve as a base for fatigue prediction of key components of the structure, meaning if an amplitude of the vibrations exceed certain value (voltage is proportional to a change of strain), any cycles that exceed this value are detected and counted, giving an estimate on how many fatigue cycles happen during a given period of operation (e.g. one flight). Various methods to analyse the measurement data and provide long-term assessment of the structure are currently being developed, but more test flights will have to be conducted before any results in this area are achieved.
5. Future work Due to a limited timeframe and tight schedule of the aircraft operation, there was only a single flight test performed during the initial phase of the research, whose purpose was to test the capabilities and functionality of the proposed sensing method. This method aimed to be as independent of any other existing system of the aircraft and technique currently in use in order to set up a starting point for the future research. Short-term monitoring, as it was performed, is suitable only for detection of sudden changes in the systems, such as impacts and critical failures. It is not possible to evaluate the entire structure based on one set of relatively short measurements. However, the goal of the current SHM techniques is not only to detect sudden changes, but more importantly to monitor slowly developed degradation processes and assess the overall state of the structure. Therefore, this research will be continued with development of long-term monitoring systems, allowing us to use the techniques described in this paper to detect and observe sudden events (e.g. impacts, failures), as well as slow
changes in the system (e.g. crack development, deterioration of material and joints), implement techniques based on comparison of the default response of the structure and the response in its current state, assess them from the perspective of SHM, and provide supplementary feedback to the currently used maintenance procedures. If the application of piezocomposite sensors for SHM proves to be successful, it can contribute towards development of composite materials with embedded autonomous piezocomposite sensing layers used to detect internal structural failures, as well as global state of the structure. 6. Conclusion A novel method for SHM in aerospace utilizing piezocomposite elements as active sensors was proposed and tested on a small civil aircraft. The main goal was providing a functional approach for global structure monitoring of an aircraft without the restrictions of conventionally used methods. In the first step, accessible loadbearing structural elements were determined using a virtual model of the fuselage section and suitable locations for the sensors were allocated. An array of piezocomposite sensors was mounted within the fuselage and connected to a data collection unit. Using these tools, a test flight was conducted and sample data during various stages of the flight were taken. Preliminary analysis of the measured data was performed to confirm functionality of the sensing network and viability of this approach for SHM in aerospace applications. The results of this research will serve as a base for future development of this approach and analysis of the measured data for SHM purposes. As could be seen from Ref. [9], the research of new structural health monitoring approaches is an ongoing activity also for rotorcraft (with an application for structural parts of EC135) and large aircrafts (with application on structural parts of A320). Physical principle of the proposed method is not limited to the small aircraft (it was used for cost effective verification) and can therefore contribute to methods used for large aircrafts as well, for example as large sensor arrays or smart layers within the composite materials. Acknowledgements Authors gratefully acknowledge financial support provided by the ESIF, EU Operational Programme Research, Development and Education within the research project Center of Advanced Aerospace Technology (Reg. No.: CZ.02.1.01/0.0/0.0/16_019/0000826) at the Faculty of Mechanical Engineering, Brno University of Technology.
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Measurement journal homepage: www.elsevier.com/locate/measurement
Integration and test of piezocomposite sensors for structure health monitoring in aerospace Filip Ksica a,⇑, Zdenek Hadas a, Jiri Hlinka b a b
Institute of Solid Mechanics, Mechatronics and Biomechanics, Faculty of Mechanical Engineering, Brno University of Technology, Brno, Czech Republic Institute of Aerospace Engineering, Faculty of Mechanical Engineering, Brno University of Technology, Brno, Czech Republic
a r t i c l e
i n f o
Article history: Received 30 October 2018 Received in revised form 13 June 2019 Accepted 22 July 2019 Available online 27 July 2019 Keywords: Structure health monitoring Piezocomposite sensors Non-intrusive testing Structural vibrations
a b s t r a c t The paper presents an innovative approach for Structure Health Monitoring utilizing piezocomposite plates as active sensors. The present work is motivated by the increasing interest in reduction of maintenance costs for aircrafts, which can be achieved by implementing self-sufficient sensor networks into the aircraft to predict risk states and prevent structural failures. These networks could supplement and possibly replace the expensive maintenance inspections. Piezoelectric materials have recently been under a rapid development, have been utilized in various fields of industry and due to their inexpensive and robust nature they can be used as sensors in non-intrusive Structure Health Monitoring networks. The presented approach uses piezocomposite sensors embedded into the structure of a small civil aircraft to directly measure structural vibrations. The functionality of the entire measurement chain was tested during an experimental flight, the acquired data will be used for further research of data-processing, impact detection and structure health prediction. Ó 2019 Elsevier Ltd. All rights reserved.
1. Introduction Over the past decades, air transport has become a rapidly growing industry, as indicated by the increasing amount of goods and passengers transported through the air. This leads to a necessity to increase number of aircrafts and the intensity of their usage, leading to more frequent and thorough maintenance procedures to ensure their safety matches the highest standards. This opens a challenging opportunity for research and development of new technologies that would help reducing operating costs and increasing effectiveness of the aircraft usage. Significant share of the operating costs of an aircraft are maintenance costs, and based on ICAO statistics, for example B757 the maintenance can form up to 23% of the flight (direct) operating costs and up to 12% of the total operating costs of the airline [1]. Typical maintenance procedure is based on a visual inspection of the overall state of the structure, with an addition of various non-destructive testing (NDT) methods to check the state of selected key components. This approach has been in use since times where aircraft structure was mainly from metal materials whose failures were in most cases visually determinable. However, for many parts of an aircraft metal materials have been substituted ⇑ Corresponding author. E-mail addresses: [email protected] (F. Ksica), [email protected] (Z. Hadas), [email protected] (J. Hlinka). https://doi.org/10.1016/j.measurement.2019.106861 0263-2241/Ó 2019 Elsevier Ltd. All rights reserved.
by more popular composite materials, that excel in terms of strength-to-weight ratio and can currently form up to 50% of the total airframe of modern large transport aircrafts [2], and even more for smaller aircraft categories. For composite structures, many failures, such as delamination, happen within the structure itself without being detectable visually, therefore pure visual inspections will become insufficient for both standard maintenance checks and post-repair diagnostics. While today’s aviation is still heavily dependent on visual inspections, a lot of work and effort has been dedicated to developing novel technologies used for non-destructive and automated damage detection, widely known under the title Structural Health Monitoring (SHM), that have high potential of supplementing or even replacing the abovementioned methods, and that can be successfully used for damage detection in tight spaces and locations otherwise hard to reach. Various SHM technologies are described in many research papers, there are also numerous publications summarizing the most important ones [3–6]. These new approaches for damage localization include methods based on propagation of Lamb waves [7–9], ultrasonic waves emitted and detected by a piezoelectric element array. If an obstacle, for example a crack, stands in the way of the waves, there is a change in the received signal spectrum. Another method utilizing piezoelectric sensors is based on pure acoustic emission [10]. This method can detect acoustic signal emitted by crack propagation,
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impact or other localized sources of damage. The evaluation is, however, very complicated due to high sensibility to noise generated during operation and is mostly viable in laboratory environment. Both methods are limited for relatively small areas in which we expect any damage to occur, but are being under development to cover larger sections of the aircraft structure. They are also sensitive to surface quality, environmental conditions such as humidity, temperature and ambient vibrations, making them relatively ineffective for long-term SHM during operation. Another SHM of composite structures is implementation of optical sensors with Fiber Bragg Grating (FBG) [11–13]. Network of optical fibres is embedded into the composite structure, deflection of localized damage translates into change in characteristics of the fibres and is detected. While still being under development, the abovementioned technologies form the mainstream SHM methods, and there are still many other technologies that could potentially form a nondestructive automated sensing network, could contribute to safety and supplement the conventional maintenance procedures. A novel approach utilizing piezocomposite sensors in active mode measuring structural vibrations directly, without the necessity to create a generator – detector network, is proposed. This paper aims to outline the procedure of utilizing piezocomposite sensors for measurement of structural vibrations in real-world conditions. For any novel technologies for SHM in aerospace, the process of their application and testing in real-world situations is limited due to very restrictive safety regulations. Each part of the sensoric network can represent a potential risk to the aircraft operation, therefore the new technology has to be certified before it can be implemented into the aircraft. Previously conducted laboratory tests provided a rough estimate of the capabilities of piezocomposite sensors for SHM of aircraft, but their capabilities in real-world conditions remained unknown. Therefore, for the purpose of this research, a small civil aircraft Cessna 172 was kindly offered for experiments with the piezocomposite sensors embedded within the structure, without any requirement for the tested system certification. It served for verification of laboratory test results and simulations, as well as obtaining information of the aircraft behaviour during various flight stages. It is believed that detailed overview of the experiment, the hardware used, and most importantly the results of conducted flight tests, are necessary for further development of SHM systems based on piezocomposite sensors, especially for modern aircraft predominantly constructed from composite materials. This approach can be used for monitoring of sudden, as well as long term changes in the structure, using spectral analysis of data measured over certain period of time, and advanced techniques (e.g. genetic algorithm) for data comparison can be used to detect and evaluate changes in the structure compared to the default response.
The overview of the planned experiment was presented on 2019 IEEE international Workshop on Metrology for Aerospace in Italy under the title Application of Piezoelectric Sensors for Structural Health Monitoring in Aerospace and was included in the workshop proceedings [14] without providing any further details or results of the conducted experiment. 2. Piezocomposite sensors: State of art Piezocomposite sensors are new technology invented and patented back in 1999 by NASA (US patent no. 6629341 – Method of fabricating a piezoelectric composite apparatus) [15]. These sensors are composed of thin piezoceramic rods with rectangular cross-section embedded within layers of epoxy matrix. Integrated electrodes are attached to both sides of the plate and sandwiched between polyamide films. They form highly flexible structures, being able to be used as both sensors and actuators. Compared to conventionally used piezoceramic sensors (piezoceramic plates), they are much more robust and flexible, have generally better electromechanical constants and do not need to be attached to conductive surface, because both electrodes are brought out through the polyamide film, as it can be seen in Fig. 1. If voltage is applied to the electrodes, it will bend and flex, distorting any material it is attached to and suppress ambient vibrations. If no voltage is applied, they form highly sensitive strain gauges or energy harvesting devices. Compared to the piezoceramic plates, they can be manufactured in wider spectrum of sizes and thicknesses. Being quite a new technology, piezocomposite sensors have not yet been used quite up to their potential. Up to date, they have seen some use in energy harvesting [17], being viable sources of electrical energy harvested from ambient vibrations of various structures. Due to their flexibility, they can be used as high deformation actuators in low and high frequency applications, or for active vibration control of flexible structures. In case of SHM, a few applications are worth mentioning, for example the abovementioned guided wave propagation technologies [18] and structure monitoring in civil engineering [19]. It is certain, that in the upcoming years they will spread into wider spectrum of sensing technologies, possibly substituting conventional piezoceramic sensors. As the piezocomposite materials are composed of piezoceramic fibres, also available separately in already polarised state, successful application of the commercially available piezocomposite sensors for direct measurement of structural vibrations can lead to further development of smart composite materials with embedded piezosensitive layers that can substitute individual sensors. Also, as most of the damage processes of composite materials happen within the material itself, either as failures of the individual layers or their interface, it would be very beneficial to embed the sensing
Fig. 1. Piezocomposite sensor, mode D33 manufactured my Smart Material [16]
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device/layer into the structure, not on its surface. The current state of their application in aerospace is relatively unexplored, as any novel system needs to go through extensive approval and certification processes before it can be tested on a real aircraft. Therefore, there have been attempts to develop accurate computational models of such materials in order to be used in virtual models of the target structures [20], but fully virtual approach is insufficient for accurate prediction of a complex structure of an aircraft. The physical principle of piezoelectric materials, as used in FEM, can be described with a coupled set of linear equations, one related to the mechanical part (Hooke’s Law) and the other to the electrical part (Maxwell’s equations),
h i fSg ¼ SE fT g þ ½dfEg T fDg ¼ ½d fT g þ ½e fEg T
where fSg is mechanical strain, fT g is mechanical stress, fDg is electrical displacement, fEg is electric field. Piezoelectric materials are characterized by several coefficients, the piezoelectric deformation constants ½d relating the applied mechanical stress to the generated charge on the electrodes, and the dielectric permitivity ½e, and the h i elastic constants of the material SE . These constants have to be either measured in laboratory environment or provided by the manufacturer. In the modelling process, it is important to keep in mind these equations are linear, but the piezocomposite material can be highly nonlinear, mostly due to hysteresis. Furthermore, piezocomposite material, compared to piezoceramic material, contains fibres and matrix, making the analytical and FEM modelling process even more difficult [20]. Therefore, it is beneficial to determine the electrical response experimentally for individual types of piezocomposite sensors and with respect to their application.
from undesirable environmental conditions (e.g. spars inside a wing, stringers inside a fuselage). Determination of these components and suitable location is key for proper function of the sensors, also for mounting into an existing aircraft, accessibility by the personnel has to be taken into account. 3.1. Experimental aircraft It is generally quite difficult to test any novel technologies on an existing aircraft because of very strict safety protocols. Previously untested system can be a potential risk for proper function of the aircraft, and to conduct a test in real-life operating conditions, either private aircraft or partnership is necessary. For the purposes of this work, a small civil aircraft Cessna 172 (Fig.2) was generously provided by one of our partners. It is a four-seat single-engine aircraft with high wing, almost entirely manufactured from sheet aluminium. 3.2. Determination of sensor locations In case of the Cessna 172, accessibility of load bearing components was the main issue. It had to be kept in mind that all the sensors have to be physically attached by wires to a data collection unit, and these wires have to go inside of the structure. Due to that, the only option for sensor placement was the fuselage, that is hollow for the most part and could be accessed via small service door (Fig. 3) on the right hand side of the aircraft or through the storage compartment. The load bearing components inside the fuselage are 7 stringers formed by bent overlaps of the aluminium skin, either L shaped or
3. Measurement of structural vibrations Piezocomposite sensors are an advantageous alternative to other conventional sensing devices due to their small and robust nature, no moving parts and very low manufacturing costs. They can be used for wide spectrum of frequencies and with high electromechanical constants, the output signals have high voltage (1 1000 V) and low current (1 10 lA) nature resulting in high SNR. For SHM purposes, they can be either attached to a surface or embedded within layers of composite material. In this work, they are used as sensors for direct measurement of structural vibrations, thus utilized in active mode (no power source required). As the piezocomposite sensors are sensitive to the change in strain, they have to be mounted on load-bearing structural components, preferably inside the aircraft to provide cover
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Fig. 3. Service door on the right-hand side of the aircraft.
Fig. 2. Photography of a Cessna 172, Source [21]
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C shaped. Two C shaped stringers situated on the bottom, one L shaped on the top and a pair of L shaped on each side (Fig. 4). The C shaped stringers were not suitable for piezocomposite sensor attachment due to their very high curvature and the top one was inaccessible due to the internal structure of the tail spine. For that reason, only the side stringers were a viable option for sensor placement. In the next step, longitudinal locations of the sensors were determined. To ensure sensitivity to all modes of load, i.e. tension, compression, bending and torsion, a simulation these loading modes was performed. A CAD model (Fig. 5) of the fuselage section was created and imported to ANSYS (Fig. 6) where it was meshed and subjected to unitary loads in separate analyses. The von-Mises stress values for the respective loads were measured along each of the side stringers. The loading modes, with respect to the global coordinate system (Z – longitudinal, Y – vertical, X – transverse), included tension in Z direction, bending in X and Y directions and torsion along Z direction. These values were normalized to fit into the interval of 0 1. The absolute values of the stress values were not important at that stage, because the Fig. 6. FEM model of the fuselage section.
aim was to determine locations with maximal values. Also, the real values of load during operation were not known at that point. In order to ensure sensitivity to all loading modes, each position along the longitudinal axis is described with a suitability index, that is normalized geometric mean of the individual normalized von-Mises stress values given by the following equation
hQ suitability index ¼
Fig. 4. Location of stringers inside Cessna 172 fuselage; left: side stringer (green line), right: bottom stringer (red line). (For interpretation of the references to colour in this figure legend, the reader is referred to the web version of this article.)
4 i¼1
rnorm;i
i
hQ i max 4i ¼ 1 rnorm;i
where rnorm;i represents the normalized von-Mises stress for each respective loading mode. The geometric mean assures that stresses (strains) for all modes are measurable at each location and as high as possible, excluding the situation that there are some modes that are not detectable. Furthermore, to reflect the fact the piezocomposite sensors measure an averaged value of strain along their entire length, an averaging filter with kernel length adjusted to reflect the final length of the sensor, in this case 1 sample per 1 cm of length. The resulting values of suitability index are illustrated in Fig. 7. The symmetry between left and right stringers was expected, as the model was fully symmetric along the longitudinal – vertical plane. While there are visible peaks on both ends of the stringers, mounting at these locations, the position of 0.5 m for all four stringers was determined as the most suitable and in Fig. 7 is marked by a dashed line. Not only the sensors formed a symmetric array that way, also it was close to the service door and accessibility and mounting was therefore much easier. 3.3. Experimental sensor network
Fig. 5. CAD model of the fuselage section; green: suitable stringers, red: unsuitable stringers. (For interpretation of the references to colour in this figure legend, the reader is referred to the web version of this article.)
After the suitable locations were determined, a preparation of the flight experiment took place. Two main requirements were kept in mind. Firstly, the sensing network should not interfere with the operating systems of the aircraft and must not form any potential risk if any part came loose. Secondly, as the sensors are mounted permanently to the structure, the system must be reduced to an absolute minimum after an experiment is conducted, also it should be reusable for any future experiments. The sensors used in the experiment were commercially available piezocomposite sensors polarized in mode D31, with embedded fibres with cross-section of 180 350 lm and total active area
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Fig. 7. Suitability index for sensor locations on individual stringers.
of 28 7 mm. Based on previous laboratory tests, to achieve the most consistent sensing capabilities of the piezocomposite sensors, the adhesive should have very high elastic modulus and should form as thin layer as possible. For that reason, a hardened highstrength structural epoxy adhesive was used. Also, it is important to mention that if conventional piezoceramic plates were used, the bottom electrode would have to be attached conductively either to the surface of the stringer or to an external wire placed between the plate and the underlying surface. However, it is best
to keep the sensing system galvanically separated from the structure of the aircraft, not only to prevent static electricity accumulated in the metal structure to affect the measurement, but also to ensure that no high voltage feedback from the sensing device gets back to the structure and any attached equipment (GPS module etc.), despite the expected voltage values being relatively small. The attached sensors shown in Fig. 8 form a bottom pair of the sensing array. To enable disconnection of the sensors after an experiment is conducted, a 3D printed enclosure for the connectors
Fig. 8. Mounted sensors (bottom pair).
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is attached next to the sensor, allowing it to leave only the sensor and a short wire to be left in the fuselage, posing marginally no threat to the operation of the aircraft. The enclosure also served another purpose. To prevent the wires breaking of the electrodes due to the mechanical vibrations, a short loop is formed to absorb any shocks along the wires. The wiring itself was looped around the fuselage using zip ties and clip brackets attached with highstrength double sided tape. The entire array of four sensors is illustrated in Fig. 9. The nominal maximum operating voltage for the used sensors is 350 V, but based on our laboratory tests and assumed values of strain, the voltage should not exceed 30 V even in extreme situations and during sharp manoeuvres, which were not expected nor approved to be performed during the test flight. Because it was not known beforehand what actual strain values and respective voltage values would be measured during an operation of the aircraft, an overvoltage protection and voltage divider (Fig. 10) had to be inserted between the sensors and the data collection unit that had only ±5 V analog inputs. The adjustable voltage divider was used to ensure the voltage values fit into an acceptable range and also served as load for energy harvesting, that was part of a different experiment (Fig. 11). From the voltage divider, BNC cables were used to connect the system to a data collection unit (DCU). The DCU was used to
Fig. 11. Data collection unit.
measure raw uncompressed voltage data with sampling frequency of 12,800 Hz and store them to a flash drive in a suitable file format. The A/D converter had resolution of 24 bits, giving us potential 16.7 million levels of resolution within the interval ±5 V. The experiment was therefore set to be on the conservative side, preventing any potential damage to the expensive hardware equipment. It was also necessary to capture any unexpected peaks. The DCU was specifically designed for this experiment to be triggered manually by an operator during flight, measuring specific parts of the test flight and enabling prompt troubleshooting in case that an unexpected event happens. Due to safety precautions, the data was stored in short sequences to achieve redundancy in measured data if any error occurred. The DCU was powered internally from a battery rather than being connected to the internal power supply of the aircraft, mostly to prevent any interference and increase mobility of the device and to fit all the components into one suitcase for easier manipulation and operation during flight (Fig. 11). It is important to keep in mind the sensor sensitivity is slightly different for each individual sensor, mostly due to inconsistencies in manual mounting, adhesive layer thickness, slight error in orientation, manufacturing tolerances etc. However, in long-term these inconsistencies can be monitored and compensated for. 4. Test flight
Fig. 9. Sensing network within the fuselage.
Fig. 10. Overvoltage protection and adjustable voltage divider.
After the entire network was constructed and tested on ground, a flight test was conducted. It took place in local airfield and the entire flight took around 20 min. During that time, over 15 individual measurements were taken, covering engine start-up, take-off, various parts of flight and landing. All four channels were measured and stored for upcoming analyses that took place back on the ground. Four clearly distinguishable segments of the test flight were chosen to illustrate the capabilities of the sensing network, namely engine ignition and movement on the airfield, take-off and ascension, steady flight and landing also with movement on the airfield, as illustrated in Fig. 12 (without engine start-up and field movement being shown). Length of the coloured arrows corresponds with the length of the measured segments. Raw measured voltage signals for the selected flight segments are shown in Fig. 13. As can be seen, the voltage values after the voltage divider, that was set to 1:20 ratio, peaked at around 0.2 V, voltage on the sensors therefore did not exceed 4 V, proving that our setting was conservative enough. The timestamp in the title of the plots is set to 00:00 at the take-off, that took place at around 00:20. In the first graph, the excitation around 20 s time mark was a failed attempt of engine ignition, a second and success-
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Fig. 12. Trajectory of the test flight with highlighted phases.
Fig. 13. Raw measured data for selected parts of test flight.
ful attempt took place at 30 s time mark. The fourth graph represents landing, touch took place at around 25 s time mark. The measured data was subjected to a spectral analysis in Matlab environment to assess the data further, distinguishing any dominant frequencies, such as engine frequency, to confirm the proper functionality of the hardware. The procedure of analysing the data using spectrogram is depicted in Fig. 14. In the first step, the measured data is separated into segments of fixed length, that is determined with respect to the maximum observable frequency required, such that there are at least 10 samples per period for each frequency. The segment length directly affects the sensitivity of the spectrogram, shorter segments means higher sensitivity to sudden changes in the signal at the cost of lower frequency resolution. With sampling frequency of 12,800 Hz and maximum observable frequency at around 600 Hz, using the equation below a segment length of 0.5 s is calculated.
window length ¼
ðmaximum observable frequencyÞ ðsampling frequencyÞ
ðsamples per periodÞ In the next step, the overall resolution of the Fast Fourier Transform (FFT) results is increased using technique of zero-padding, in which a zero array is added to the data segment. The length of the array should be chosen carefully, as it slows the process of FFT and
only up to a certain point the resolution enhancement is relevant. For our case, the suitable padding was with an array of 10 times length of the original vector. Certain signals may require using a window (e.g. Gaussian window) for the padded segment in order to ensure that the segment starts and ends with the same value to prevent artificial steps in the periodic signal constructed during FFT. In our case, this step was not necessary, as it had no significant impact of the final spectrogram. In the following step, padded data segments were subjected to a FFT and smoothed using averaging filter over 5 samples to slightly reduce the noise in the signal. In the last step, the resulting frequency-domain signals are stacked in time-domain to form a spectrogram. The results of the spectral analysis for characteristic stages of the flight, depicted only in the time-frequency plane, are shown in Fig. 15. Starting from the first figure, the failed engine start-up is evident at around 20 s. At 30 s, the engine start-up was successful and a dominant frequency of approximately 37 Hz is clearly visible. That frequency corresponds to the revolution of the rotor shaft, as the nominal engine speed was 2400 rev.min1 (40 Hz). The multiples of the frequency in the spectrum correspond to the rotor blade frequency (2 blades) and cylinder movement (6 cylinders, 2 pairs of 3). The rest of the lines most likely correspond with natural frequencies of the structure and other moving parts of the aircraft. It was expected that the engine frequency would be the most
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Fig. 14. Procedure of creating high-resolution spectrogram.
Fig. 15. Spectral analysis of the test flight data, first sensor, variable flight stages.
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Fig. 16. Spectral analysis of the test flight data, variable sensors, take-off stage.
dominant one, as the engine is the most significant source of vibrations within the aircraft. However, the measured data can be assessed from a different point of view, and that is comparison between individual sensors (Fig. 16). Based on the comparison between measured data from individual sensors during one stage of the flight, load of individual stringers could be assessed. For example, it is clearly visible that 3rd and 4th stringer (left hand side of the fuselage) were subjected to slightly higher amounts of strain than the other two, providing us an estimate of current loads on the dominant frequency. It is assumed that any localized disturbances (impact, sudden events etc.) or long-term damage development will somehow reflect in the measured response, either by shift in natural frequencies, amplitudes, or by visible wide-spectrum excitation. However, the samples obtained during one test flight are not enough to assess that and methods for thorough data analysis and SHM. When a long-term monitoring experiment is conducted, these data will also serve as a base for fatigue prediction of key components of the structure, meaning if an amplitude of the vibrations exceed certain value (voltage is proportional to a change of strain), any cycles that exceed this value are detected and counted, giving an estimate on how many fatigue cycles happen during a given period of operation (e.g. one flight). Various methods to analyse the measurement data and provide long-term assessment of the structure are currently being developed, but more test flights will have to be conducted before any results in this area are achieved.
5. Future work Due to a limited timeframe and tight schedule of the aircraft operation, there was only a single flight test performed during the initial phase of the research, whose purpose was to test the capabilities and functionality of the proposed sensing method. This method aimed to be as independent of any other existing system of the aircraft and technique currently in use in order to set up a starting point for the future research. Short-term monitoring, as it was performed, is suitable only for detection of sudden changes in the systems, such as impacts and critical failures. It is not possible to evaluate the entire structure based on one set of relatively short measurements. However, the goal of the current SHM techniques is not only to detect sudden changes, but more importantly to monitor slowly developed degradation processes and assess the overall state of the structure. Therefore, this research will be continued with development of long-term monitoring systems, allowing us to use the techniques described in this paper to detect and observe sudden events (e.g. impacts, failures), as well as slow
changes in the system (e.g. crack development, deterioration of material and joints), implement techniques based on comparison of the default response of the structure and the response in its current state, assess them from the perspective of SHM, and provide supplementary feedback to the currently used maintenance procedures. If the application of piezocomposite sensors for SHM proves to be successful, it can contribute towards development of composite materials with embedded autonomous piezocomposite sensing layers used to detect internal structural failures, as well as global state of the structure. 6. Conclusion A novel method for SHM in aerospace utilizing piezocomposite elements as active sensors was proposed and tested on a small civil aircraft. The main goal was providing a functional approach for global structure monitoring of an aircraft without the restrictions of conventionally used methods. In the first step, accessible loadbearing structural elements were determined using a virtual model of the fuselage section and suitable locations for the sensors were allocated. An array of piezocomposite sensors was mounted within the fuselage and connected to a data collection unit. Using these tools, a test flight was conducted and sample data during various stages of the flight were taken. Preliminary analysis of the measured data was performed to confirm functionality of the sensing network and viability of this approach for SHM in aerospace applications. The results of this research will serve as a base for future development of this approach and analysis of the measured data for SHM purposes. As could be seen from Ref. [9], the research of new structural health monitoring approaches is an ongoing activity also for rotorcraft (with an application for structural parts of EC135) and large aircrafts (with application on structural parts of A320). Physical principle of the proposed method is not limited to the small aircraft (it was used for cost effective verification) and can therefore contribute to methods used for large aircrafts as well, for example as large sensor arrays or smart layers within the composite materials. Acknowledgements Authors gratefully acknowledge financial support provided by the ESIF, EU Operational Programme Research, Development and Education within the research project Center of Advanced Aerospace Technology (Reg. No.: CZ.02.1.01/0.0/0.0/16_019/0000826) at the Faculty of Mechanical Engineering, Brno University of Technology.
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