## 3 Results and Discussion
From Figure 8, it can be seen that the signal generated during the swing movement achieved a frequency ranging from 0.8 to 1.3 Hz. For the verification and confirmation of the results obtained from the spectral analysis, another experiment using the ELMA DT178A 3-axis acceleration and vibration datalogger was conducted. In this case, the experiment was performed so that the end of the aircraft wing was pushed down and released. The initial impulse was created by the 20 cm deviation at the end of the wing. The dynamic response of the system until the steady state was achieved was recorded and analyzed. The main purpose of this experiment was to simulate the step change of the input signal in the mechanical system represented by the aircraft. From the theory of cybernetics, it is known that to obtain the frequency response of the system, it is necessary to apply the step change as the input signal of the system. In this way, it is possible to obtain the natural frequency, which is the frequency at which the system tends to oscillate in the absence of any driving or damping force. It is the frequency at which the structure will tend to vibrate when subjected to certain external forces. These frequencies are dependent on the way in which the mass and stiffness are distributed within the material structure. The flight envelope of any aircraft is determined by the resonant frequencies of the wings and airframe structure; this was the main reason that the tests specified above were used for the verification of the tensile stress measurement. If the aircraft moves too fast, the aerodynamic forces can excite the natural frequencies, causing a resonance, which, in an aircraft, is called flutter. Resonance is a phenomenon in which a dynamic force drives a structure to vibrate at its natural frequency. When a structure is in resonance, a small force can produce a large vibration response with the high amplitude. These oscillations can cause damage to the aircraft structure. This analysis helps to explain the reasons for vibrations, which can cause structural damage during the aircraft construction, and to propose solutions regarding how to reduce them--for example, by the mass or stiffness modification or damping changes. In Figure 9, the FFT analysis of the accelerometer signal can be seen. For comparison, in this case, the same detail of the FFT analysis in the frequency range from 0 to 5 Hz is shown. It is obvious that the bandwidth of the aircraft response was not changed. In this case, it was from 1 to 1.5 Hz. The frequency shift of the response will be the subject of further research. Since these measurements were performed in October with the ambient temperature of 5 \({}^{\circ}\)C, we supposed that the shift was mainly caused by the temperature dependence of the Young modulus. Figure 3:0 FFT analysis of the accelerometer signal obtained during the aircraft wing step movement. This assumption was also analytically confirmed by the performed simulations based on the idealized simplified CAD/CAE (Computer Aided Design/Computer Aided Engineering) model of the aircraft wing connected to the fuselage. The CAD model of the whole aircraft was created in the Creo Parametric software and the consequent simulations, calculations and analyses were performed in the Creo Simulate software. In Figure 10, the simulation model and the results of the modal analysis can be seen. For the first mode of deformation, which is visualized in Figure 10, the calculated natural frequency of the aircraft construction was 1.289 Hz. After the initial modal simulation, the static simulation involving the temperature change from 5 to 30 \({}^{\circ}\)C was performed, which created a basis for the following prestress modal analysis. The resulting natural frequency simulating the increased temperature was 1.057 Hz. The measurement point for the analyses was defined according to the placement of the sensing element, thus simulating the real position of the sensor.