Masakazu Shimanuki

Development of damage monitoring system for aircraft structure using a PZT actuator/FBG sensor hybrid system

This paper presents a part of the research results on a damage monitoring system using PZT actuators/FBG sensors for advanced composite material structures of new-generation aircrafts. To achieve weight reduction of the aircraft structure, these advanced composite materials have gradually been employed for the primary structure. It is expected that when these materials are extensively employed, an efficient bonded structure such as a hat-shaped stringer will be utilized for the aircraft structure. However, these bonded structures have critical problems such as debonding and delamination at the interfaces of the laminate. Further, a single-step molding process of the structure elements is necessary in order to ensure low cost and thus affordability. However, this low-cost process results in an increase in the non-destructive inspection (NDI) cost. Therefore, an innovative damage monitoring system is required for structural health management. In the present study, the authors have developed a hybrid sensor system that can detect the elastic waves launched from the piezo transducer (PZT) actuator using a high-speed and high-accuracy fiber Bragg grating (FBG) sensor to resolve the issues mentioned above. In this study, the conceptual design of an aircraft that can employ this damage monitoring system was carried out. Subsequently, the application area was selected based on cases of certain kinds of damage. Further, the validity of the damage monitoring system for the verification of the structural integrity of the aircraft was discussed. Next, in order to verify the elastic wave detectability of the FBG sensor, it was confirmed that an elastic wave of 300 kHz is detectable at a distance of 5 cm between the PZT actuator and FBG sensor using an aluminum sheet and CFRP cross-ply laminate and also by considering the relationship between sensor length and sensitivity. Through the present research results, the possibility of applying the damage monitoring system to the composite material bonding structure in an aircraft is presented.

Feasibility studies on active damage detection for CFRP aircraft bonding structures

This paper presents a part of the feasibility study for employing a damage monitoring system using a PZT actuator/FBG sensor. The goal of this research is to improve the present safety level of aircrafts by ensuring structural integrity and reducing maintenance/operation costs of advanced composite materials that will be employed for the primary structures of the new generation aircraft. Our novel sensor system employing a PZT actuator and an FBG optical fiber sensor can detect several types of damage, such as delamination and debonding, which are expected to occur in the highly efficient bonding structures of an aircraft wing box when the main/tail wing section is designed using composite materials. In this system, elastic waves will be transmitted into the structure by PZT actuators and received by small-diameter FBG optical fiber sensors that are embedded in the critical section of the primary structure of an aircraft. The onset and growth of damage can be detected with a very high accuracy through the change in the received elastic waves. In this study, a conceptual design was implemented in order to employ the novel system. We also selected the optimum application candidate area in order to verify the effects of the system. Further, the investigation and experiment on the novel sensor system that uses a bonded PZT actuator and an FBG sensor was carried out, and the basic damage detection technique was established based on the experimental results of the received elastic wave. Furthermore, compressive tests were carried out using the coupon specimen with an embedded small-diameter or standard-diameter optical fiber sensor. It was verified that the coupon specimen with an embedded small-diameter optical fiber suffered no degradation of its material properties.

Damage growth detection of composite laminate using embedded FBG sensor/PZT actuator hybrid system

This paper presents a part of the study conducted for developing a damage diagnostic system for an advanced composite material that can be utilized in next-generation aircraft structure. The authors have been working on a detection of elastic wave which can be launched from the PZT actuators, using small- and normal-diameter FBG optical fiber sensors that are bonded to the surface of the CFRP laminate under different conditions. Based on the results, it was verified that it is possible to achieve a high-accuracy detection of elastic wave by using FBG sensors bonded to the surface of the CFRP laminate. It was also verified that the damages generated on the inside of the composite material may be detected by the waveform analysis of the received elastic wave. In this study, the authors succeeded in the embedment of small-diameter FBG optical fiber sensors into the bonding surface of the double-lap type coupon specimen, which simulates the bonding structure of the CFRP composite structure. In this study, we also clarified several issues pertaining to the conditions, methods, and techniques involved in fiber embedding. An optical loss was observed during the embedment process, which may result in the loss of both accuracy and reliability. Based on these observations, the authors developed embedding techniques for optical fiber sensors that can reduce this optical loss. Additionally, the possibility of detecting an elastic wave, which was launched from the PZT actuators bonded to the surface of the coupon and directed to the host material, was verified using double-lap type coupon specimen having embedded small-diameter FBG optical fiber sensors at the bonding surface. Therefore, this specimen has provided an artificial defect that simulates the delamination generated at the bonding interface. Based on the measurements of the elastic wave, it was verified that the change in the elastic wave depends on the damage length, which is caused by the artificial defect. Moreover, based on the analysis of the received elastic wave, the possibility of damage detection was confirmed. The successful development of this damage monitoring system would ease the implementation of structural health monitoring system in aircraft structures in the near future.

A Basic Study of CFRP Laminates with Embedded Prestrained SMA Foils for Aircraft Structures

This paper presents some basic research results for the application of the smart materials and structural systems to aircraft structure. The authors conducted quasi-static load-unload tests and fatigue tests on several kinds of quasi-isotropic carbon fiber-reinforced plastic (CFRP) laminates with embedded prestrained shape memory alloy (SMA) foils. The damage behavior and the fatigue behavior were evaluated based on the effects of the recovery stress of SMA, and the relationship between the applied strain and the transverse crack density is discussed. It was found that the developed smart materials achieved a maximum improvement of 34% in the onset strain of the transverse cracks and a maximum improvement of 60% in the onset strain of delamination as compared with the standard CFRP laminates in the static condition. However, the fatigue properties of the smart material were not improved due to the material degradation of the 90°-ply at 80°C. The fatigue test by the multistep strain method using a structural element specimen was used to verify the results of coupon tests. The results indicate the possibility of using coupon data as a design database.

Damage growth monitoring for a bonding layer of the aircraft bonding structure

This study is about some of the results of the test performed for the purpose of developing the damage monitoring system with the advanced composite bonding structure applied to a next generation aircraft. Through the past researches, we succeeded in receiving hundreds of kHz of elastic waves (lamb waves) launched from PZT byusing an optical fiber sensor bonded to or embedded in a specimen. Furthermore, by using an FBG optical fiber sensor embedded in the bonding interface of a CFRP coupon specimen or a structure element specimen with skin/stringer bonded structure or bonded to the surface of the specimen, the test results have been proven and the fact is verified that regarding the structure of composites, variations of elastic waves according to damage growth can be received with high accuracy. The authors also suggest it is possible to detect a damage, which is generated inside composites by calculating the elastic waves. For this study, we manufactured a structural element specimen where a small-diameter optical fiber sensor is embedded in the bonding interface, which is simulated a skin/stringer bonding structure of actual composite structures. We also developed the system, which is detecting elastic (lamb) wave up to 1MHz on our own and optimized it according to the corresponding specimen. Furthermore, an artificial damage is installed to critical area of the structural element specimen as a damage origin point. It is verified that our monitoring system can detect the variations of elastic waves accompanying the damage of 20mm2 occurring and growing from the artificial damage by the applied cyclic load.

Damage suppression system using embedded SMA (shape memory alloy) foils in CFRP laminate structures

This paper presents an overview of the demonstrator program with respect to the damage growth suppression effects using embedded SMA foils in CFRP laminates. The damage growth suppression effects were demonstrated for the technical verification in order to apply to aircraft structure. In our previous studies, the authors already confirmed the damage growth suppression effects of CFRP laminates with embedded pre-strained SMA foils through both coupon and structural element tests. It was founded that these effects were obtained by the suppression of the strain energy release rate based on the suppression of the crack opening displacement due to the recovery stress of SMA foils through the detail observation of the damage behavior. In this study, these results were verified using the demonstrator test article, which was 1/3-scaled model of commercial airliner fuselage structure. For the demonstration of damage growth suppression effects, the evaluation area was located in the lower panel, which was dominated in tension load during demonstration. The evaluation area is the integrated stiffened panel including both “smart area” (CFRP laminate with embedded pre-strained SMA foils) and “conventional area” (standard CFRP laminate) for the direct comparison. The demonstration was conducted at 80 degree Celsius in smart area and room temperature (RT) in conventional area during quasi-static load-unload test method. As the test results, the demonstrator test article presented that the damage onset strain in the smart area was improved by 30% for compared with the conventional area. Therefore, the successful technical verification of the damage onset/growth suppression effect using the demonstrator presented the feasibility of the application of smart material and structural system to aircraft structures.