An essential part of the SHM validation effort is to check the presence and adequacy of the methods required to validate the correct functionality of each SHM task, which can be targeted at detecting structural faults. The ultimate proof of the correct functionality is validation evidence, e.g. crack detection evidence, observed during the operation of the aircraft. However, the occurrences of structural faults such as cracks are infrequent, and hence, years of flight tests might be required to collect validation evidence; small numbers of flights would be only sufficient to prove the system's “fitness for flight” and would be insufficient to prove “fitness for purpose”. Validation evidence can be collected during laboratory tests by inducing faults in structural specimens and examining the SHM detection capability. However, collecting validation evidence using simple structural specimens during laboratory tests would never provide the conclusive proof of correct functionality for the aircraft complex structural assembly. Therefore, a generalisation technique and calibration approach would be required to extrapolate from laboratory specimens to actual aircraft. This paper will present the technical details of such a validation method with the illustration example being the primary task of damage location detection in a composite structure.