MIL-HDBK-17-3F Volume 3.Chapter 12-Lessons Learned 12.2.3 Damage tolerance Damage tolerance is the measure of the structure's ability to sustain a level of damage or presence of a defect and be able to perform its operating functions.The concern is with the damaged structure hav- ing adequate residual strength and stiffness to continue in service safely:1)until the damage can be de- tected by scheduled maintenance inspection and repaired,or 2)if the damage is undetected,for the re- mainder of the aircraft's life.Thus,safety is the primary goal of damage tolerance.Both static load and durability related damage tolerance must be interrogated experimentally because there are few,if any. accurate analytical methods. There are basically two types of damage that are categorized by their occurrence during the fabrica- tion and use of the part,i.e.,damage occurring during manufacturing or damage occurring in service.It is hoped that the occurrence of the majority of manufacturing associated damage,if beyond specification limits,will be detected by routine quality inspection.Nevertheless,some "rogue"defects or damage be- yond specification limits may go undetected.Consequently,their occurrence must be assumed in the design procedure and subsequent testing(static and fatigue)performed to verify the structural integrity. Service damage concerns are similar to those for manufacturing.Types of service damage include edge and surface gouges and cuts or foreign object collision and blunt object impact damage caused by dropped tools or contact with service equipment.A level of non-detectable damage should be established and verified by test that will not endanger the normal operation of the aircraft structure for two lifetimes.A certain level(maximum allowed)damage that can be found by inspection should be defined such that the vehicle can operate for a specified number of hours before repair or replacement at loads not exceeding design limit.This damage should also be tested(statically and in fatigue)to verify the structural integrity. Delaminations can also be critical defects.However,unless they are very large,historically more than 2 inches(50 mm)in diameter,the problem is mostly with thin laminates.Effects of manufacturing defects such as porosity and flawed fastener holes that are slightly in excess of the maximum allowable are usually less severe.They are generally accounted for by the use of design allowable properties that have been obtained by testing specimens with stress concentrations,e.g.,notches.Most commonly these are specimens with a centered hole.Open holes are typically used for compression specimens while either open or filled holes (holes with an installed fastener)are used for tension testing.(Open holes are more critical than filled holes for compression.Filled holes may be more critical in tension,es- pecially for laminates with ply orientations with a predominate number of plies in the load direction.)Con- sequently,the design allowables thus produced may be used to account for a nominal design stress con- centration caused by an installed or missing fastener,at least to a 0.25 inch(6.4 mm)diameter,as well as accounting for many other manufacturing defects.This is sometimes called the "rogue flaw"approach to laminate design,see Reference 12.2.3. 12.2.4 Durability Durability of a structure is its ability to maintain strength and stiffness throughout the service life of the structure.A structure must have adequate durability when subjected to the expected service loads and environment spectra to prevent excessive maintenance,repair,or modification costs over the service life. Thus,durability is primarily an economic consideration. Metallic structure can be very sensitive to durability issues;major factors limiting life are corrosion and fatigue.Metal fatigue is dictated by the number of load cycles required to start a crack(crack initia- tion)and the number of load cycles for the crack to grow to its critical length,reaching catastrophic failure (crack growth).Crack/damage growth rate is very dependent on the concentration of stress around the crack. In composites,it has been demonstrated that one of the most common damage growth mechanisms is intercracking(delamination).This makes composites most sensitive to compression-dominated fatigue loading.A second common fatigue failure mode is fastener hole wear caused by high bearing stresses. 12-4MIL-HDBK-17-3F Volume 3, Chapter 12 - Lessons Learned 12-4 12.2.3 Damage tolerance Damage tolerance is the measure of the structure's ability to sustain a level of damage or presence of a defect and be able to perform its operating functions. The concern is with the damaged structure hav￾ing adequate residual strength and stiffness to continue in service safely: 1) until the damage can be de￾tected by scheduled maintenance inspection and repaired, or 2) if the damage is undetected, for the re￾mainder of the aircraft's life. Thus, safety is the primary goal of damage tolerance. Both static load and durability related damage tolerance must be interrogated experimentally because there are few, if any, accurate analytical methods. There are basically two types of damage that are categorized by their occurrence during the fabrica￾tion and use of the part, i.e., damage occurring during manufacturing or damage occurring in service. It is hoped that the occurrence of the majority of manufacturing associated damage, if beyond specification limits, will be detected by routine quality inspection. Nevertheless, some "rogue" defects or damage be￾yond specification limits may go undetected. Consequently, their occurrence must be assumed in the design procedure and subsequent testing (static and fatigue) performed to verify the structural integrity. Service damage concerns are similar to those for manufacturing. Types of service damage include edge and surface gouges and cuts or foreign object collision and blunt object impact damage caused by dropped tools or contact with service equipment. A level of non-detectable damage should be established and verified by test that will not endanger the normal operation of the aircraft structure for two lifetimes. A certain level (maximum allowed) damage that can be found by inspection should be defined such that the vehicle can operate for a specified number of hours before repair or replacement at loads not exceeding design limit. This damage should also be tested (statically and in fatigue) to verify the structural integrity. Delaminations can also be critical defects. However, unless they are very large, historically more than 2 inches (50 mm) in diameter, the problem is mostly with thin laminates. Effects of manufacturing defects such as porosity and flawed fastener holes that are slightly in excess of the maximum allowable are usually less severe. They are generally accounted for by the use of design allowable properties that have been obtained by testing specimens with stress concentrations, e.g., notches. Most commonly these are specimens with a centered hole. Open holes are typically used for compression specimens while either open or filled holes (holes with an installed fastener) are used for tension testing. (Open holes are more critical than filled holes for compression. Filled holes may be more critical in tension, es￾pecially for laminates with ply orientations with a predominate number of plies in the load direction.) Con￾sequently, the design allowables thus produced may be used to account for a nominal design stress con￾centration caused by an installed or missing fastener, at least to a 0.25 inch (6.4 mm) diameter, as well as accounting for many other manufacturing defects. This is sometimes called the "rogue flaw" approach to laminate design, see Reference 12.2.3. 12.2.4 Durability Durability of a structure is its ability to maintain strength and stiffness throughout the service life of the structure. A structure must have adequate durability when subjected to the expected service loads and environment spectra to prevent excessive maintenance, repair, or modification costs over the service life. Thus, durability is primarily an economic consideration. Metallic structure can be very sensitive to durability issues; major factors limiting life are corrosion and fatigue. Metal fatigue is dictated by the number of load cycles required to start a crack (crack initia￾tion) and the number of load cycles for the crack to grow to its critical length, reaching catastrophic failure (crack growth). Crack/damage growth rate is very dependent on the concentration of stress around the crack. In composites, it has been demonstrated that one of the most common damage growth mechanisms is intercracking (delamination). This makes composites most sensitive to compression-dominated fatigue loading. A second common fatigue failure mode is fastener hole wear caused by high bearing stresses