A method of monitoring structural integrity of a repair made to a laminated structure made of composite material, wherein the laminated structure is part of an aircraft skin including interwoven wire fabric wires for providing lightning protection, and wherein the laminated structure is subjected to loads during flight of the aircraft, comprising: (a) placing a sensor between layers of composite material of a repair patch, the repair patch including an interwoven wire fabric layer, and the sensor being capable of outputting data representing a current structural characteristic of surrounding composite material after the composite material has been cured; (b) curing the composite material while the repair patch is in contact with a repair site of a parent structure made of composite material such that the wires of the interwoven wire fabric layer of the repair patch overlap the interwoven fabric wires of the aircraft skin to produce a repaired parent structure having an embedded sensor; (c) after the curing step, acquiring and recording baseline data from the embedded sensor which represents a structural characteristic of the repaired parent structure; (d) after the baseline data has been acquired and recorded, subjecting the repaired parent structure to loads; (e) acquiring and recording post-loading data from the embedded sensor at a time subsequent to or during step (d), said post-loading data representing a structural characteristic of the repaired parent structure; (f) processing the baseline data and post-loading data in a manner that identifies differences between the respective baseline and post-loading data indicative of structural change in the repaired parent structure; and (g) determining whether the identified differences indicate a structural integrity change to the repaired parent structure in excess of a specified threshold, wherein steps (e) through (g) are performed by a computer system. (a) placing a sensor between layers of composite material of a repair patch, the repair patch including an interwoven wire fabric layer, and the sensor being capable of outputting data representing a current structural characteristic of surrounding composite material after the composite material has been cured; (b) curing the composite material while the repair patch is in contact with a repair site of a parent structure made of composite material such that the wires of the interwoven wire fabric layer of the repair patch overlap the interwoven fabric wires of the aircraft skin to produce a repaired parent structure having an embedded sensor; (c) after the curing step, acquiring and recording baseline data from the embedded sensor which represents a structural characteristic of the repaired parent structure; (d) after the baseline data has been acquired and recorded, subjecting the repaired parent structure to loads; (e) acquiring and recording post-loading data from the embedded sensor at a time subsequent to or during step (d), said post-loading data representing a structural characteristic of the repaired parent structure; (f) processing the baseline data and post-loading data in a manner that identifies differences between the respective baseline and post-loading data indicative of structural change in the repaired parent structure; and (g) determining whether the identified differences indicate a structural integrity change to the repaired parent structure in excess of a specified threshold,