MIL-HDBK-17-3F Volume 3,Chapter 7-Damage Resistance,Durability,and Damage Tolerance Proof of structure,full-scale static test.The test program starts with an article provided with simu- lated low velocity impact damages,limited by the selected energy cut-off levels and deliberately inflicted at the most stressed areas of the structure.The Ultimate Load capability is demonstrated after fatigue allowing for environmental adverse conditions.This is in line with the means of compliance provided by the AC 20 107A 6 Proof of structure-static,sub s(a):The effects of repeated loading and environ- mental exposure which may result in material property degradation should be addressed in the static evaluation. Proof of structure,full-scale fatigue/damage tolerance test.When considering the effects of ma- terial variability on the repeated load behavior of composite structures,a factor on loads is preferred to a factor on life.The rationale of such approach and the recommended load enhancement factors can be found in References 7.2.1(a)to 7.2.1(d).The demonstration has two parts. First,an enhanced safe life (flaw tolerant)demonstration,to show that no damage will initiate and grow in a structure representative of the minimum quality allowed by the quality control specification(con- sidering not only impact damage but also various manufacturing flaws).This phase is in line with AC 20 107A S 7 Proof of structure-Fatigue/Damage tolerance,(b)fatigue (safe life)evaluation:Fatigue sub- stantiation should be accomplished by component fatigue tests or by analysis supported by test evidence, accounting for the effects of the appropriate environment.The test articles should be fabricated and as- sembled in accordance with production specifications and processes so that the test articles are repre- sentative of production structure.etc. Second,a no-growth demonstration for more severe impact damages,some of which may become detectable at the scheduled inspection intervals.This phase is in line with AC 20 107A S 7 Proof of struc- ture-Fatigue/Damage tolerance,sub $(a):Structural details,elements,and subcomponents of critical structural areas should be tested under repeated loads to define the sensitivity of the structure to damage growth.This testing can form the basis for validating a no growth approach to damage tolerance require- ments A demonstration of the regulatory static load capability is needed to complete this second phase.A"k" value higher than 1.0 can be required depending on the result of a probabilistic approach,if used for certi- fication.It is the second phase of the full-scale test that brings most to the demonstration of the structural safety.At this stage,a precise definition of damage growth is required.For instance,there may be a possibility where an impact damage will grow under the first service loads following the occurrence and, then reach a definite size after a certain time.This is still to be assumed as a"no-growth"situation,since the "growth"is not detrimental to the structural capability.On another hand,a damage can be definitely arrested by a design precaution (a bolt row for instance).Provided regulatory load capability exists after this size extension,the result is comparable to a no-growth situation. 7.3 TYPES,CHARACTERISTICS,AND SOURCES OF DAMAGE Damages are generally discussed in two frames of reference-by stage of occurrence and by physi- cal anomaly.Stage of occurrence is separated into manufacturing and in-service categories.Damages occurring during manufacturing are more accurately classified as"flaws"rather than"damages".They are not distinguished as such in this write-up. Composite aircraft parts can be damaged during manufacturing,shipping.and service.A primary focus in composites is low velocity impacts that can cause significant damage that may not be clearly visible.Sources of such impact damage include falling tools and equipment,runway debris,hail,birds, and collision with other airplanes or ground vehicles.Airplanes can also be damaged by high velocity impacts from discrete source events (e.g.,parts of rotating machinery that fail in turbofan engines and penetrate the engine containment system,the aircraft skin,and supporting structure).All of the above damages can occur to either military or commercial aircraft.Military aircraft may also suffer ballistic dam- age,as may occur in battle. 7-26
MIL-HDBK-17-3F Volume 3, Chapter 7 - Damage Resistance, Durability, and Damage Tolerance 7-26 Proof of structure, full-scale static test. The test program starts with an article provided with simulated low velocity impact damages, limited by the selected energy cut-off levels and deliberately inflicted at the most stressed areas of the structure. The Ultimate Load capability is demonstrated after fatigue, allowing for environmental adverse conditions. This is in line with the means of compliance provided by the AC 20 107A § 6 Proof of structure-static, sub § (a) : The effects of repeated loading and environmental exposure which may result in material property degradation should be addressed in the static evaluation. Proof of structure, full-scale fatigue/damage tolerance test. When considering the effects of material variability on the repeated load behavior of composite structures, a factor on loads is preferred to a factor on life. The rationale of such approach and the recommended load enhancement factors can be found in References 7.2.1(a) to 7.2.1(d). The demonstration has two parts. First, an enhanced safe life (flaw tolerant) demonstration, to show that no damage will initiate and grow in a structure representative of the minimum quality allowed by the quality control specification (considering not only impact damage but also various manufacturing flaws). This phase is in line with AC 20 107A § 7 Proof of structure - Fatigue/Damage tolerance, (b) fatigue (safe life) evaluation: Fatigue substantiation should be accomplished by component fatigue tests or by analysis supported by test evidence, accounting for the effects of the appropriate environment. The test articles should be fabricated and assembled in accordance with production specifications and processes so that the test articles are representative of production structure, etc. Second, a no-growth demonstration for more severe impact damages, some of which may become detectable at the scheduled inspection intervals. This phase is in line with AC 20 107A § 7 Proof of structure - Fatigue/Damage tolerance, sub § (a): Structural details, elements, and subcomponents of critical structural areas should be tested under repeated loads to define the sensitivity of the structure to damage growth. This testing can form the basis for validating a no growth approach to damage tolerance requirements. A demonstration of the regulatory static load capability is needed to complete this second phase. A “k” value higher than 1.0 can be required depending on the result of a probabilistic approach, if used for certification. It is the second phase of the full-scale test that brings most to the demonstration of the structural safety. At this stage, a precise definition of damage growth is required. For instance, there may be a possibility where an impact damage will grow under the first service loads following the occurrence and, then reach a definite size after a certain time. This is still to be assumed as a “no-growth” situation, since the "growth" is not detrimental to the structural capability. On another hand, a damage can be definitely arrested by a design precaution (a bolt row for instance). Provided regulatory load capability exists after this size extension, the result is comparable to a no-growth situation. 7.3 TYPES, CHARACTERISTICS, AND SOURCES OF DAMAGE Damages are generally discussed in two frames of reference - by stage of occurrence and by physical anomaly. Stage of occurrence is separated into manufacturing and in-service categories. Damages occurring during manufacturing are more accurately classified as “flaws” rather than “damages”. They are not distinguished as such in this write-up. Composite aircraft parts can be damaged during manufacturing, shipping, and service. A primary focus in composites is low velocity impacts that can cause significant damage that may not be clearly visible. Sources of such impact damage include falling tools and equipment, runway debris, hail, birds, and collision with other airplanes or ground vehicles. Airplanes can also be damaged by high velocity impacts from discrete source events (e.g., parts of rotating machinery that fail in turbofan engines and penetrate the engine containment system, the aircraft skin, and supporting structure). All of the above damages can occur to either military or commercial aircraft. Military aircraft may also suffer ballistic damage, as may occur in battle
MIL-HDBK-17-3F Volume 3,Chapter 7-Damage Resistance,Durability,and Damage Tolerance Concerns about the effects of impact damage can be quite different,depending on the specific design and application.Compressive residual strength of laminated composite material forms is known to de- pend on the extent of delaminations and fiber failure caused by transverse impacts.Tensile residual strength is affected by fiber failure.Impact damage can also affect the environmental resistance of a composite structural component or the integrity of associated aircraft systems.For example,impact dam- age may allow moisture to penetrate into the sandwich core in light-gauge fairing panels or provide a path for fuel leaks in stiffened wing panels.These effects must be understood for safe and economic composite applications. 7.3.1 Damages characterized by stage of occurrence 7.3.1.1 Manufacturing Manufacturing damage includes anomalies such as porosity,microcracking,and delaminations result- ing from processing discrepancies and also such items as inadvertent edge cuts,surface gouges and scratches,damaged fastener holes,and impact damage.The inadvertent (non-process)damage can occur in detail parts or components during assembly or transport or during operation.A list of sources of manufacturing defects is given below: Improper cure or processing Improper machining Mishandling Improper drilling Tool drops Contamination Improper sanding Substandard material Inadequate tooling Mislocation of holes or details Most manufacturing damage,if beyond acceptance limits,will be detected by routine quality inspec- tion.For every composite part,there should be acceptance/rejection criteria to be used during inspection of the part.Damage that is acceptable will be incorporated in the substantiation analysis and test pro- gram to demonstrate ultimate strength in the presence of this damage.Some "rogue"defects or damage beyond specification limits may go undetected and consequently,their existence must be assumed as part of damage tolerant design.Establishing the size of the "rogue"or missed flaw is part of the design criteria development process. Examples of rogue flaws occurring in manufacturing include a contaminated bondline surface,or in- clusions such as prepreg backing paper or separation film that is inadvertently left between plies during lay-up.Current inspection methods may not detect these types of defects.As a result,current design practices include the effect of large debonds in damage tolerance criteria which may impose severe weight penalties.In the future,advanced inspection techniques and in-process quality control may lead to less severe criteria.Without adequate inspection techniques,in-process quality controls must be suffi- ciently rigid to preclude this type of defect. 7.3.1.2 Service The main characteristic of in-service damage is that it occurs during service in a random manner. Damage characteristics,location,size,and frequency of occurrence can only be predicted statistically, which involves a large amount of data accumulation.In-service damage is typically classified as non- detectable and detectable(often referred to as non-visible and visible).A part has to be designed in such a way that likely,non-detectable damage (per the selected inspection method)can be tolerated under Ultimate Loads and for the life of the structure.The most common in-service damage is due to an impact event.A list of sources of in-service damage threats is given below: 7-27
MIL-HDBK-17-3F Volume 3, Chapter 7 - Damage Resistance, Durability, and Damage Tolerance 7-27 Concerns about the effects of impact damage can be quite different, depending on the specific design and application. Compressive residual strength of laminated composite material forms is known to depend on the extent of delaminations and fiber failure caused by transverse impacts. Tensile residual strength is affected by fiber failure. Impact damage can also affect the environmental resistance of a composite structural component or the integrity of associated aircraft systems. For example, impact damage may allow moisture to penetrate into the sandwich core in light-gauge fairing panels or provide a path for fuel leaks in stiffened wing panels. These effects must be understood for safe and economic composite applications. 7.3.1 Damages characterized by stage of occurrence 7.3.1.1 Manufacturing Manufacturing damage includes anomalies such as porosity, microcracking, and delaminations resulting from processing discrepancies and also such items as inadvertent edge cuts, surface gouges and scratches, damaged fastener holes, and impact damage. The inadvertent (non-process) damage can occur in detail parts or components during assembly or transport or during operation. A list of sources of manufacturing defects is given below: Improper cure or processing Improper machining Mishandling Improper drilling Tool drops Contamination Improper sanding Substandard material Inadequate tooling Mislocation of holes or details Most manufacturing damage, if beyond acceptance limits, will be detected by routine quality inspection. For every composite part, there should be acceptance/rejection criteria to be used during inspection of the part. Damage that is acceptable will be incorporated in the substantiation analysis and test program to demonstrate ultimate strength in the presence of this damage. Some “rogue” defects or damage beyond specification limits may go undetected and consequently, their existence must be assumed as part of damage tolerant design. Establishing the size of the “rogue” or missed flaw is part of the design criteria development process. Examples of rogue flaws occurring in manufacturing include a contaminated bondline surface, or inclusions such as prepreg backing paper or separation film that is inadvertently left between plies during lay-up. Current inspection methods may not detect these types of defects. As a result, current design practices include the effect of large debonds in damage tolerance criteria which may impose severe weight penalties. In the future, advanced inspection techniques and in-process quality control may lead to less severe criteria. Without adequate inspection techniques, in-process quality controls must be sufficiently rigid to preclude this type of defect. 7.3.1.2 Service The main characteristic of in-service damage is that it occurs during service in a random manner. Damage characteristics, location, size, and frequency of occurrence can only be predicted statistically, which involves a large amount of data accumulation. In-service damage is typically classified as nondetectable and detectable (often referred to as non-visible and visible). A part has to be designed in such a way that likely, non-detectable damage (per the selected inspection method) can be tolerated under Ultimate Loads and for the life of the structure. The most common in-service damage is due to an impact event. A list of sources of in-service damage threats is given below:
MIL-HDBK-17-3F Volume 3,Chapter 7-Damage Resistance,Durability,and Damage Tolerance Hailstones Runway debris Ground vehicles,equipment,and structures Lightning strike Tool drops Birdstrike Turbine blade separation Fire Wear Ballistic damage(Military) Rain erosion Ultraviolet exposure Hygrothermal cycling Oxidative degradation Repeated loads Chemical exposure 7.3.2 Damages characterized by physical imperfection Damage can occur at several scales within the composite material and structural configuration.This ranges from damage in the matrix and fiber to broken elements and failure of bonded or bolted attach- ments.The extent of damage controls repeated load life and residual strength,and is,therefore,critical to damage tolerance Fiber Breakage.This defect can be critical because structures are typically designed to be fiber dominant(i.e.,fibers carry most of the loads).Fortunately,fiber failure is typically limited to a zone near the point of impact,and is constrained by the impact object size and energy.Only a few of the service related events listed in the previous section could lead to large areas of fiber damage. Matrix Imperfections.(Cracks,porosity,blisters,etc.)These usually occur on the matrix-fiber inter- face,or in the matrix parallel to the fibers.These imperfections can slightly reduce some of the material properties but will seldom be critical to the structure,unless the matrix degradation is widespread.Accu- mulation of matrix cracks can cause the degradation of matrix-dominated properties.For laminates de- signed to transmit loads with their fibers(fiber dominant),only a slight reduction of properties is observed when the matrix is severely damaged.Matrix cracks,a.k.a.micro-cracks,can significantly reduce proper- ties dependent on the resin or the fiber/resin interface,such as interlaminar shear and compression strength.For high temperature resins,micro-cracking can have a very negative effect on properties.A discussion of the effects of matrix damage on the tensile strength can be found in Reference 7.3.2(a). Matrix imperfections may develop into delaminations,which are a more critical type of damage. Delamination and debonds.Delaminations form on the interface between the layers in the laminate. Delaminations may form from matrix cracks that grow into the interlaminar layer or from low energy im- pact.Debonds can also form from production non-adhesion along the bondline between two elements and initiate delamination in adjacent laminate layers.Under certain conditions,delaminations or debonds can grow when subjected to repeated loading and can cause catastrophic failure when the laminate is loaded in compression.The criticality of delaminations or debonds depend on: ●Dimensions Number of delaminations at a given location. Location-in the thickness of laminate,in the structure,proximity to free edges,stress concentra- tion region,geometrical discontinuities,etc. Loads-behavior of delaminations and debonds depend on loading type.They have little affect on the response of laminates loaded in tension.Under compression or shear loading,however, the sublaminates adjacent to the delaminations or debonded elements may buckle and cause a load redistribution mechanism,which leads to structural failure.Methods to estimate the criticality of delamination and debonds are presented in Section 7.8.4.2. 7-28
MIL-HDBK-17-3F Volume 3, Chapter 7 - Damage Resistance, Durability, and Damage Tolerance 7-28 Hailstones Runway debris Ground vehicles, equipment, and structures Lightning strike Tool drops Birdstrike Turbine blade separation Fire Wear Ballistic damage (Military) Rain erosion Ultraviolet exposure Hygrothermal cycling Oxidative degradation Repeated loads Chemical exposure 7.3.2 Damages characterized by physical imperfection Damage can occur at several scales within the composite material and structural configuration. This ranges from damage in the matrix and fiber to broken elements and failure of bonded or bolted attachments. The extent of damage controls repeated load life and residual strength, and is, therefore, critical to damage tolerance. Fiber Breakage. This defect can be critical because structures are typically designed to be fiber dominant (i.e., fibers carry most of the loads). Fortunately, fiber failure is typically limited to a zone near the point of impact, and is constrained by the impact object size and energy. Only a few of the service related events listed in the previous section could lead to large areas of fiber damage. Matrix Imperfections. (Cracks, porosity, blisters, etc.) These usually occur on the matrix-fiber interface, or in the matrix parallel to the fibers. These imperfections can slightly reduce some of the material properties but will seldom be critical to the structure, unless the matrix degradation is widespread. Accumulation of matrix cracks can cause the degradation of matrix-dominated properties. For laminates designed to transmit loads with their fibers (fiber dominant), only a slight reduction of properties is observed when the matrix is severely damaged. Matrix cracks, a.k.a. micro-cracks, can significantly reduce properties dependent on the resin or the fiber/resin interface, such as interlaminar shear and compression strength. For high temperature resins, micro-cracking can have a very negative effect on properties. A discussion of the effects of matrix damage on the tensile strength can be found in Reference 7.3.2(a). Matrix imperfections may develop into delaminations, which are a more critical type of damage. Delamination and debonds. Delaminations form on the interface between the layers in the laminate. Delaminations may form from matrix cracks that grow into the interlaminar layer or from low energy impact. Debonds can also form from production non-adhesion along the bondline between two elements and initiate delamination in adjacent laminate layers. Under certain conditions, delaminations or debonds can grow when subjected to repeated loading and can cause catastrophic failure when the laminate is loaded in compression. The criticality of delaminations or debonds depend on: • Dimensions • Number of delaminations at a given location. • Location - in the thickness of laminate, in the structure, proximity to free edges, stress concentration region, geometrical discontinuities, etc. • Loads - behavior of delaminations and debonds depend on loading type. They have little affect on the response of laminates loaded in tension. Under compression or shear loading, however, the sublaminates adjacent to the delaminations or debonded elements may buckle and cause a load redistribution mechanism, which leads to structural failure. Methods to estimate the criticality of delamination and debonds are presented in Section 7.8.4.2
MIL-HDBK-17-3F Volume 3,Chapter 7-Damage Resistance,Durability,and Damage Tolerance Combinations of Damages.In general,impact events cause combinations of damages.High-energy impacts by large objects (i.e.,turbine blades)may lead to broken elements and failed attachments.The resulting damage may include significant fiber failure,matrix cracking.delamination,broken fasteners, and debonded elements.Damage caused by low-energy impact is more contained,but may also include a combination of broken fibers,matrix cracks and multiple delaminations.There is some experimental evidence that,for relatively small damage sizes,impact damage is more critical than other defects(see Figures 7.3.2(a)and(b),References 7.3.2(b)and(c)).Note that all of the data shown in these figures are for damage sizes less than 2 inches(50 mm).Some results for damages greater than 2 inches(50 mm) suggest large holes or penetrations are at least as severe as equivalent sizes of impact damage. Damage diameter(in.)or porosity 0.5 1.0 1.5 2.0 1.2 1.0 0.8 54-mm hole 0.6 0.4 △Porosity 0.2 o Delamination △Open hole ●Filled hole and O Impact damage delamination 6.4 mm,laminate 12.7 25.4 38.1 50.8 Damage diameter(mm) FIGURE 7.3.2(a)Relative severity of defect damage on static compression strength. Flawed Fastener Holes.Improper hole drilling,poor fastener installation,and missing fasteners may occur in manufacturing.Hole elongation can occur due to repeated load cycling in service.Such issues can effectively extend the size of the hole and lead to assumptions that the hole is open(or filled,depend- ing on which leads to the greater notch sensitivity).The notch sensitivity of a composite has generally been dealt with by using semi-empirical analyses. 7-29
MIL-HDBK-17-3F Volume 3, Chapter 7 - Damage Resistance, Durability, and Damage Tolerance 7-29 Combinations of Damages. In general, impact events cause combinations of damages. High-energy impacts by large objects (i.e., turbine blades) may lead to broken elements and failed attachments. The resulting damage may include significant fiber failure, matrix cracking, delamination, broken fasteners, and debonded elements. Damage caused by low-energy impact is more contained, but may also include a combination of broken fibers, matrix cracks and multiple delaminations. There is some experimental evidence that, for relatively small damage sizes, impact damage is more critical than other defects (see Figures 7.3.2(a) and (b), References 7.3.2(b) and (c)). Note that all of the data shown in these figures are for damage sizes less than 2 inches (50 mm). Some results for damages greater than 2 inches (50 mm) suggest large holes or penetrations are at least as severe as equivalent sizes of impact damage. Damaged strength/Undamaged strength Damage diameter (in.) or % porosity 1.2 1.0 0.8 0.6 0.4 0.2 0 0.5 1.0 1.5 2.0 54- mm hole Porosity Open hole Impact damage 6.4 mm laminate o Delamination Filled hole and delamination 0 12.7 25.4 38.1 50.8 o o o Damage diameter (mm) FIGURE 7.3.2(a) Relative severity of defect damage on static compression strength. Flawed Fastener Holes. Improper hole drilling, poor fastener installation, and missing fasteners may occur in manufacturing. Hole elongation can occur due to repeated load cycling in service. Such issues can effectively extend the size of the hole and lead to assumptions that the hole is open (or filled, depending on which leads to the greater notch sensitivity). The notch sensitivity of a composite has generally been dealt with by using semi-empirical analyses
MIL-HDBK-17-3F Volume 3,Chapter 7-Damage Resistance,Durability,and Damage Tolerance 1.0 ∠2%porosity 0.8 -50.1-mm diam delamination Visible impact damage 0.6 xew 12.7 mm diam hole 0.4 'ssel]s 0.2 0 101 102 103 104 105 106 107 Cycles to failure FIGURE 7.3.2(b)Relative severity of defect damage on compression fatigue strength,R=10. 7.3.3 Realistic impact energy threats to aircraft As discussed in Section 7.2.2,certification of aircraft composite structure requires the establishment of realistic impact energy level cut-offs for Ultimate Load considerations.A conservative assumption is to set the energy level at a 90%probability,analogous with the concept of a B-basis strength value.This then means that the realistic energy cut-off has been selected in such a way that,at the end of lifetime of the aircraft,no more than 10%of them will have been impacted with an energy value equal to this cut-off level or higher.For these 10%corresponding to a more damaged situation,and then possibly not being able to comply with the Ultimate Load requirements.damage tolerance considerations will demonstrate the regulatory safety level. Letting Eco energy cut-off value,and with Pa the probability,per flight,to encounter one impact with an energy E=Eco,then,(1-Pa)is the probability for an aircraft to have encountered either no impact or impacts of a lower energy on that flight.In fact the risk of low velocity impact damage is not likely to occur during the actual flight,but during the various operations associated with this flight,e.g.,aircraft servicing and a shared part of the risk associated with the scheduled inspections. Then it follows that: (1-Pa)n is the probability to have never encountered any impact with an energy of E>Eco after n flights,and then the probability to have encountered at least one damage created by an impact with an energy of E>Eco after n flights is given by: P=1-(1-Pa) Assuming that: n=50,000 flights for a short/medium range commercial aircraft, P=0.1 Then Pa 2.1 x 105 7-30
MIL-HDBK-17-3F Volume 3, Chapter 7 - Damage Resistance, Durability, and Damage Tolerance 7-30 Cyclic stress, max/Static stress, max. 12.7 mm diam hole Cycles to failure 1.0 0.8 0.6 0.4 0.2 0 101 102 103 104 105 106 107 Visible impact damage 2% porosity 50.1-mm diam delamination FIGURE 7.3.2(b) Relative severity of defect damage on compression fatigue strength, R=10. 7.3.3 Realistic impact energy threats to aircraft As discussed in Section 7.2.2, certification of aircraft composite structure requires the establishment of realistic impact energy level cut-offs for Ultimate Load considerations. A conservative assumption is to set the energy level at a 90% probability, analogous with the concept of a B-basis strength value. This then means that the realistic energy cut-off has been selected in such a way that, at the end of lifetime of the aircraft, no more than 10% of them will have been impacted with an energy value equal to this cut-off level or higher. For these 10% corresponding to a more damaged situation, and then possibly not being able to comply with the Ultimate Load requirements, damage tolerance considerations will demonstrate the regulatory safety level. Letting Eco = energy cut-off value, and with Pa the probability, per flight, to encounter one impact with an energy E ≥ Eco, then, (1-Pa) is the probability for an aircraft to have encountered either no impact or impacts of a lower energy on that flight. In fact the risk of low velocity impact damage is not likely to occur during the actual flight, but during the various operations associated with this flight, e.g., aircraft servicing and a shared part of the risk associated with the scheduled inspections. Then it follows that: (1-Pa)n is the probability to have never encountered any impact with an energy of E ≥ Eco after n flights, and then the probability to have encountered at least one damage created by an impact with an energy of E ≥ Eco after n flights is given by: n P = 1 - (1-Pa) Assuming that: n = 50,000 flights for a short/medium range commercial aircraft, P = 0.1 Then Pa = 2.1 x 10-6