This paper presents a practical approach for health monitoring of an aircraft fuselage using vibration measurements with piezoelectric transducers. For the test specimen, a fuselage element of an Airbus A320, the feasibility of health monitoring of stringers, frames and panels, is studied. The proposed structural health monitoring (SHM) system consists of three major components: vibration measurement, signal processing and damage diagnosis. By using applied piezoelectric patch actuators, flexural waves are excited which propagate along the monitored component. The used signal for the actuator is a broadband sine-sweep signal with an upper frequency of 100 kHz. The structural response is measured with piezoelectric patch transducers, which generate a charge proportional to the induced strain due to the vibration. These sensors are positioned on different locations around the exciting piezoelectric actuator. The sampled time signal is used to compute the complex frequency response functions between actuator and sensor. After measuring the vibration behaviour of the undamaged structure, the structure is artificially damaged by saw cuts to simulate cracks. After that the vibration behaviour is measured again in the same way. The measurements of the damaged and undamaged state are now evaluated with different mathematical algorithms, like root mean square value, norm, energy content, analysis of the phase and correlation coefficient. The aim is to find a method that needs very small computation effort and that even might be implemented by an analogue circuit. Finally the data are analysed to find appropriate damage metrics. It can be shown that some of the tested methods are a sensitive damage metric. With only a few actuators and sensors cracks in the outer panel with a length of 30-60 mm and in the stringer profile with a length of 10 mm can be found.