MIL-HDBK-17-3F Volume 3,Chapter 8 Supportability be standardized to the maximum extent possible,and repair considerations are appropriate for concept development of any aircraft structural component.This section lists recommendations for design ap- proaches that will improve the repairability of composite aircraft structural components. ● Interchangeable (Net shape Replaceable w/holes) (Rough shape,needs trimming hole drilling) FIGURE 8.2.7 Differences between interchangeable and replaceable panels. 8.2.9.1 General design approach The approach for the composite structures design team needs to be based upon input and knowledge gained from a working relationship established between the design team and airline maintenance person- nel.This can be accomplished through repair workshops,or inquiries,involving airline and OEM cus- tomer support personnel,engineering personnel and involvement with the Commercial Aircraft Composite Repair Committee (CACRC).Reference 8.2.9.1 is a product of this committee and provides general guidance.The time spent within these efforts will provide a broader understanding of the overall envi- ronment in which operators operate.OEM involvement in the CACRC has contributed to addressing the problems that operators voice.The CACRC is pioneering standards and recommendations for the design and maintenance of future composite structure based on current and past experience. Figure 8.2.9.1(a)shows the maintenance development philosophy established during Phase B of the Boeing/NASA Advanced Technology Composite Aircraft Structures(ATCAS)composite fuselage program. Maintenance procedures such as inspection and repair,which are applicable to a service environment, must be considered during design selection.Considerations should be made for bolted repairs,for ex- ample,there should be sufficient edge distances on stiffener and frame flanges,and sandwich edge bands to allow for repair bolts.Skin thicknesses should be sufficient to prevent knife-edges when using 8-16
MIL-HDBK-17-3F Volume 3, Chapter 8 Supportability 8-16 be standardized to the maximum extent possible, and repair considerations are appropriate for concept development of any aircraft structural component. This section lists recommendations for design approaches that will improve the repairability of composite aircraft structural components. Interchangeable (Net shape w/holes) Replaceable (Rough shape, needs trimming & hole drilling) FIGURE 8.2.7 Differences between interchangeable and replaceable panels. 8.2.9.1 General design approach The approach for the composite structures design team needs to be based upon input and knowledge gained from a working relationship established between the design team and airline maintenance personnel. This can be accomplished through repair workshops, or inquiries, involving airline and OEM customer support personnel, engineering personnel and involvement with the Commercial Aircraft Composite Repair Committee (CACRC). Reference 8.2.9.1 is a product of this committee and provides general guidance. The time spent within these efforts will provide a broader understanding of the overall environment in which operators operate. OEM involvement in the CACRC has contributed to addressing the problems that operators voice. The CACRC is pioneering standards and recommendations for the design and maintenance of future composite structure based on current and past experience. Figure 8.2.9.1(a) shows the maintenance development philosophy established during Phase B of the Boeing/NASA Advanced Technology Composite Aircraft Structures (ATCAS) composite fuselage program. Maintenance procedures such as inspection and repair, which are applicable to a service environment, must be considered during design selection. Considerations should be made for bolted repairs, for example, there should be sufficient edge distances on stiffener and frame flanges, and sandwich edge bands to allow for repair bolts. Skin thicknesses should be sufficient to prevent knife-edges when using
MIL-HDBK-17-3F Volume 3,Chapter 8 Supportability countersunk repair fasteners.Fabric outer plies should be considered to help reduce breakout when drill- ing holes for repair bolts in laminates or face sheets.Any lightning strike protection systems that are needed on specific components should be designed to be repairable.It should be assumed that bolted and bonded repairs would follow general practice guidelines. Design for Repair Early development of Well-defined maintenance procedures ADL Design Damage tolerant Efficient,low-cost NDI Load Ultimate design,including procedures to locate damage significant CDT (that always find CD刀 Limit Reliable and simple NDE to Maximum load per fleet lifetime quantify effects of damage Continued safe flight Cost-effective repair with minimal down time when 7 damage is found Allowable Critical Damage Damage Limit Threshold (ADL) (CDT) Increasing Damage Size FIGURE 8.2.9.1(a)Rules for maintainable composite structures. Some composite structural details,while weight and cost efficient,are difficult to repair.Closed hat- section stringers,for instance,are compatible with inexpensive manufacturing techniques and minimum weight,but pose difficulties relative to inspection and attachment in repair applications.The use of blind fasteners should also be kept to a minimum because of difficulties in removing them in order to make a repair or a replacement.Where fasteners are necessary,the removable types are preferable.Quite of- ten,when removing fasteners performing a repair,the drilling out of blind fasteners the surrounding struc- ture is damaged,thus incurring more cost and down-time.Material choices may also be affected.The designer should avoid the use of different material systems with different curing temperatures on one part. For instance,skins and stiffeners are sometimes precured at 350F(177C)and then,for manufacturing ease,secondarily bonded with 250F(121C)adhesive.This can present problems when the skins or stiffeners are repaired at 350F(177C);the integrity of the 250F(121C)adhesive at the bond interface may be compromised with no indication of degradation. Design concept developments should include parallel efforts to establish maintenance procedures. Maintenance procedures established after design features for manufacturing scale-up are set will typically result in unnecessarily complex repair designs and processes. Another important aspect of concept development critical to maintenance is damage tolerant design practices.The allowable damage limits (ADL)and critical damage thresholds (CDT)defined in Figure 8.2.9.1(a)must be established to support the structural repair manual and inspection procedures. The former allows rapid determination of the need for repair during scheduled inspection,while the latter should be sufficiently large to allow safe aircraft operation between inspection intervals.Knowledge of residual strength and inspection capabilities should allow determination of both ADL and CDT as a 8-17
MIL-HDBK-17-3F Volume 3, Chapter 8 Supportability 8-17 countersunk repair fasteners. Fabric outer plies should be considered to help reduce breakout when drilling holes for repair bolts in laminates or face sheets. Any lightning strike protection systems that are needed on specific components should be designed to be repairable. It should be assumed that bolted and bonded repairs would follow general practice guidelines. Allowable Damage Limit (ADL) Increasing Damage Size Ultimate Maximum load per fleet lifetime Design Load Continued safe flight Limit Critical Damage Threshold (CDT) Efficient, low-cost NDI procedures to locate damage (that always find CDT) Damage tolerant design, including significant CDT Well-defined ADL Design for Repair Early development of maintenance procedures Reliable and simple NDE to quantify effects of damage Cost-effective repair with minimal down time when damage is found FIGURE 8.2.9.1(a) Rules for maintainable composite structures. Some composite structural details, while weight and cost efficient, are difficult to repair. Closed hatsection stringers, for instance, are compatible with inexpensive manufacturing techniques and minimum weight, but pose difficulties relative to inspection and attachment in repair applications. The use of blind fasteners should also be kept to a minimum because of difficulties in removing them in order to make a repair or a replacement. Where fasteners are necessary, the removable types are preferable. Quite often, when removing fasteners performing a repair, the drilling out of blind fasteners the surrounding structure is damaged, thus incurring more cost and down-time. Material choices may also be affected. The designer should avoid the use of different material systems with different curing temperatures on one part. For instance, skins and stiffeners are sometimes precured at 350°F (177°C) and then, for manufacturing ease, secondarily bonded with 250°F (121°C) adhesive. This can present problems when the skins or stiffeners are repaired at 350°F (177°C); the integrity of the 250°F (121°C) adhesive at the bond interface may be compromised with no indication of degradation. Design concept developments should include parallel efforts to establish maintenance procedures. Maintenance procedures established after design features for manufacturing scale-up are set will typically result in unnecessarily complex repair designs and processes. Another important aspect of concept development critical to maintenance is damage tolerant design practices. The allowable damage limits (ADL) and critical damage thresholds (CDT) defined in Figure 8.2.9.1(a) must be established to support the structural repair manual and inspection procedures. The former allows rapid determination of the need for repair during scheduled inspection, while the latter should be sufficiently large to allow safe aircraft operation between inspection intervals. Knowledge of residual strength and inspection capabilities should allow determination of both ADL and CDT as a
MIL-HDBK-17-3F Volume 3,Chapter 8 Supportability function of structural location.Damages smaller than the ADL limit may never be discovered,whereas the CDT damage level must always be found by visual inspection. The design of some areas of the structure can be controlled by manufacturing and durability consid- erations. Specific examples of these considerations are minimum gage(to provide a minimum of impact damage resistance and avoid knife-edge at countersink fasteners),stiffener,rib and frame flange width bolt spacing and edge distance requirements,and avoiding rapid ply drops and buildups.Areas of the structure designed to these considerations will,therefore,have higher margins for damage tolerance. Figure 8.2.8.1(b)shows the minimum margins of safety for a composite fuselage side panel,illustrating the "over-designed"regions.These regions have ADLs and CDTs larger than the rest of the fuselage section.Zoned ADL and CDT information should prove useful to operators desiring minimum mainte- nance costs.Structural Repair Manuals quite often point to critical zones on components for special di- rected inspections,so zoned ADL and CDT information could be included in these manuals. Margins of Safety 1.00 Allowable Load Applied Load 0.25 0.15 0.05 F1GURE8.2.9.1(b) Strength margin of safety distribution for a fuselage side panel subjected to ultimate loads. Returning to Figure 8.2.9.1(a),another requirement for maintainable composite structure is the estab- lishment of non-destructive inspection(NDI)and evaluation(NDE)procedures for practical damage loca- tion and quantitative assessment,respectively,during scheduled maintenance.The latter,which may require ultrasonic methods,should only be required to assess the effects of damage found by more easily performed procedures (e.g.,visual). Damage Levels.When damage is found,efficient repair procedures are needed that the operators can accomplish with available resources (tooling,equipment,etc.)and with a minimum amount of airplane down time.In order to develop repair concepts for a broad range of damage scenarios,the repair design philosophy is focusing on more generic repairs that are not damage-specific.This approach will be bene- ficial because generic designs and corresponding repair procedures can be developed for various levels of damage which are,within certain limits,independent of specific damages.This is intended to greatly reduce the need to develop repairs for each damage event as it occurs,providing a higher level of main- tainability.Initially,three damage levels have been defined and are shown in Table 8.2.9.1(a)as they ap- ply to a skin/stringer configuration. 8-18
MIL-HDBK-17-3F Volume 3, Chapter 8 Supportability 8-18 function of structural location. Damages smaller than the ADL limit may never be discovered, whereas the CDT damage level must always be found by visual inspection. The design of some areas of the structure can be controlled by manufacturing and durability considerations. Specific examples of these considerations are minimum gage (to provide a minimum of impact damage resistance and avoid knife-edge at countersink fasteners), stiffener, rib and frame flange width, bolt spacing and edge distance requirements, and avoiding rapid ply drops and buildups. Areas of the structure designed to these considerations will, therefore, have higher margins for damage tolerance. Figure 8.2.8.1(b) shows the minimum margins of safety for a composite fuselage side panel, illustrating the "over-designed" regions. These regions have ADLs and CDTs larger than the rest of the fuselage section. Zoned ADL and CDT information should prove useful to operators desiring minimum maintenance costs. Structural Repair Manuals quite often point to critical zones on components for special directed inspections, so zoned ADL and CDT information could be included in these manuals. FIGURE 8.2.9.1(b) Strength margin of safety distribution for a fuselage side panel subjected to ultimate loads. Returning to Figure 8.2.9.1(a), another requirement for maintainable composite structure is the establishment of non-destructive inspection (NDI) and evaluation (NDE) procedures for practical damage location and quantitative assessment, respectively, during scheduled maintenance. The latter, which may require ultrasonic methods, should only be required to assess the effects of damage found by more easily performed procedures (e.g., visual). Damage Levels. When damage is found, efficient repair procedures are needed that the operators can accomplish with available resources (tooling, equipment, etc.) and with a minimum amount of airplane down time. In order to develop repair concepts for a broad range of damage scenarios, the repair design philosophy is focusing on more generic repairs that are not damage-specific. This approach will be beneficial because generic designs and corresponding repair procedures can be developed for various levels of damage which are, within certain limits, independent of specific damages. This is intended to greatly reduce the need to develop repairs for each damage event as it occurs, providing a higher level of maintainability. Initially, three damage levels have been defined and are shown in Table 8.2.9.1(a) as they apply to a skin/stringer configuration
MIL-HDBK-17-3F Volume 3,Chapter 8 Supportability TABLE 8.2.9.1(a)Example of skin/stringer damage level definitions. Designation Damage Description Repair Level 0 Edge skin delamination or disbond Fastener restraint or injection from stiffening elements resin repair Level 1 Critical damage to a single structural Mechanically fastened or element(skin or stiffener) bonded patch and/or splice Level 2 Multiple occurrences of Level 1 Same as Level 1 (and higher) damage Designs should address repair in such a way that each bay is looked upon as a unit,or building block. Restoration of that unit(rib,frame,stringer,and/or skin)should be designed so that larger multiple-bay damages can be handled with less effort.Structural units are less easily defined for sandwich structure; however,the same general philosophy applies.The strategy behind this approach is to address the re- pair scenarios for a large range of damage at the beginning of the design process to ease the mainte- nance burden. Multiple Options.Another aspect of the approach is to provide operators with multiple options for a given repair situation.Options might include,as examples,temporary vs.permanent repair,bonded com- posite patches versus bolted composite or metal patches.or wet lay-up or prepreg patches versus pre- cured bonded patches.An operator's choice might depend on the severity of the damage,the time avail- able to perform the repair,the operator's facilities and capabilities,inspection/overhaul schedules,and/or current field environmental conditions. Durability versus weight trades.The understanding derived from residual strength analyses and tests will also ultimately lead to cost and weight trades that affect all of the total direct operating costs(DOC). Small increases in manufacturing cost and structural weight may be traded against increased damage tolerance and durability to reduce maintenance costs.Decisions may be required to balance the ADL and CDT.For example,test results for laminate tensile notch sensitivity may show an inverse relationship between small and large notch strength.Under such circumstances it may be desirable to have some ADL capability to avoid having to repair small damages but not at the expense of CDTs that allow suffi- ciently long inspection intervals and satisfactory fail-safe behavior. 8.2.9.2 Repair design issues Skin/stringer structure repair issues.Solid laminate skin/stringer designs are quite often repaired using mechanically fastened external skin patches and nested substructure splice angles.Mechanically fastened repairs require care and accuracy in the drilling of holes and the alignment of parts during as- sembly.Fastener hole breakout is a characteristic problem,commonly solved by using a layer of fabric as the outermost ply for all laminates.Typically,even though there may be other methods to avoid fas- tener hole breakout,there are numerous situations in the real world that challenge a good mechanic's ability to consistently drill high quality holes.Provisions to locate the position of the drilled holes in the structure include alignment marks and templates.Each skin/stringer component design should have laminates lay-ups that have sufficient thickness and numbers of plies in each of the0°，90°，and45°direc- tions so that they are repairable with mechanically fastened patches. Sandwich structure repair issues.Sandwich structure is generally repaired with insitu processed bonded scarf or stepped patches.The typical scarf/step taper ratios employed when repairing thin face sheets of control panels and fixed secondary structure are quite shallow(e.g.,20:1).When repairing sandwich structures with thicker face sheets in more highly loaded areas,however,scarf repairs with these traditional shallow taper ratios result in the removal of a large amount of undamaged material,and 8-19
MIL-HDBK-17-3F Volume 3, Chapter 8 Supportability 8-19 TABLE 8.2.9.1(a) Example of skin/stringer damage level definitions. Designation Damage Description Repair Level 0 Edge skin delamination or disbond from stiffening elements Fastener restraint or injection resin repair Level 1 Critical damage to a single structural element (skin or stiffener) Mechanically fastened or bonded patch and/or splice Level 2 (and higher) Multiple occurrences of Level 1 damage Same as Level 1 Designs should address repair in such a way that each bay is looked upon as a unit, or building block. Restoration of that unit (rib, frame, stringer, and/or skin) should be designed so that larger multiple-bay damages can be handled with less effort. Structural units are less easily defined for sandwich structure; however, the same general philosophy applies. The strategy behind this approach is to address the repair scenarios for a large range of damage at the beginning of the design process to ease the maintenance burden. Multiple Options. Another aspect of the approach is to provide operators with multiple options for a given repair situation. Options might include, as examples, temporary vs. permanent repair, bonded composite patches versus bolted composite or metal patches, or wet lay-up or prepreg patches versus precured bonded patches. An operator's choice might depend on the severity of the damage, the time available to perform the repair, the operator's facilities and capabilities, inspection/overhaul schedules, and/or current field environmental conditions. Durability versus weight trades. The understanding derived from residual strength analyses and tests will also ultimately lead to cost and weight trades that affect all of the total direct operating costs (DOC). Small increases in manufacturing cost and structural weight may be traded against increased damage tolerance and durability to reduce maintenance costs. Decisions may be required to balance the ADL and CDT. For example, test results for laminate tensile notch sensitivity may show an inverse relationship between small and large notch strength. Under such circumstances it may be desirable to have some ADL capability to avoid having to repair small damages but not at the expense of CDTs that allow sufficiently long inspection intervals and satisfactory fail-safe behavior. 8.2.9.2 Repair design issues Skin/stringer structure repair issues. Solid laminate skin/stringer designs are quite often repaired using mechanically fastened external skin patches and nested substructure splice angles. Mechanically fastened repairs require care and accuracy in the drilling of holes and the alignment of parts during assembly. Fastener hole breakout is a characteristic problem, commonly solved by using a layer of fabric as the outermost ply for all laminates. Typically, even though there may be other methods to avoid fastener hole breakout, there are numerous situations in the real world that challenge a good mechanic's ability to consistently drill high quality holes. Provisions to locate the position of the drilled holes in the structure include alignment marks and templates. Each skin/stringer component design should have laminates lay-ups that have sufficient thickness and numbers of plies in each of the 0°, 90°, and 45° directions so that they are repairable with mechanically fastened patches. Sandwich structure repair issues. Sandwich structure is generally repaired with insitu processed bonded scarf or stepped patches. The typical scarf/step taper ratios employed when repairing thin face sheets of control panels and fixed secondary structure are quite shallow (e.g., 20:1). When repairing sandwich structures with thicker face sheets in more highly loaded areas, however, scarf repairs with these traditional shallow taper ratios result in the removal of a large amount of undamaged material, and
MIL-HDBK-17-3F Volume 3,Chapter 8 Supportability hence,very large patch sizes.In these situations,repairs may be combinations of scarfed and external patches,so that the repair sizes can be minimized.Flush repairs may be required for some components for aerodynamic reasons or to prevent chafing.Also,thick face sheets require thick patches,which may require special processing to achieve proper consolidation.Patch and bondline porosity are of particular concern with normal field processing,which is accomplished with vacuum pressure and heat blankets. Lower temperature cures are generally preferred due to concerns over causing additional damage via vaporization of water that has infiltrated the core.Also,the surrounding structure may act as a heat sink, making it difficult to achieve and control the higher temperatures with heat blankets,and may contribute to thermal gradients that can result in warpage or degradation of the surrounding structure.For thick sand- wich,heat blankets on both sides of the structure may be required to control the through thickness tem- perature.Still,the shorter processing times generally associated with higher temperature cures are very attractive in terms of minimizing the out-of-service time for a damaged airplane. Sandwich moisture ingression issues.Consideration must be given to moisture ingression when designing maintainable,repairable sandwich structures.Sandwich designs must address the effects of moisture in the core,both by minimizing the degree of moisture ingression,and by determining what its presence does to the performance of the structure.Moisture ingression can occur through face sheet damages,and part edge and end seals,so special care must be taken to design durable sandwich parts. Unfortunately,to make durable face skins,additional thickness is needed,and this may not be desirable from a performance point of view.Durable edge and end seals can be designed,see Ref.8.2.1.1.When repairing damaged sandwich structures,a drying cycle is typically performed prior to the accomplishment of any bonded repair.This is performed so that any retained moisture does not interfere with the curing cycle.There have been numerous cases of face skins blowing off sandwich components during the vac- uum bag heating cure cycle. 8.3 SUPPORT IMPLEMENTATION A repair has the objective of restoring a damaged structure to an acceptable capability in terms of strength,durability,stiffness,functional performance,safety,cosmetic appearance or service life.Ideally, the repair will return the structure to original capability and appearance. The design assessment of a repair for a given loading condition involves the selection of a repair con- cept,the choice of the appropriate repair materials and processes,then specifying the detailed configura- tion and size of the repair.Most repairs are basically designed as a joint to transfer load into and out of a patch.To ensure that the repair configuration will have adequate strength and stiffness,the repair joint must be analyzed to predict its strength. The selection of the type of load-transfer joint to be used for a patch/strap is a tradeoff between sim- plicity,strength and stiffness.The easier configurations are generally not as strong as the more difficult ones.It is critical that the materials and process information is available prior to the system being put into place. 8.3.1 Part Inspection Presence of damage in aircraft composite parts is usually found in the course of a routine on-line in- spection,depot inspection,or,for large damages,noticed by the pilot.The predominant mode of inspec- tion is visual with more sophisticated modes of inspection performed at the depot.Once damage is iden- tified visually in-service,the damage should be characterized quantitatively by measuring dent depth,ex- tent of surface damage,and length of scratches before proceeding to more complex NDI.This will gen- erally consist of tap testing to define the boundary between damaged and undamaged portions of the structure and followed,for major repairs,with instrumented NDI techniques(ultrasonics,radiography,etc.) to locate the through the thickness characteristics of the damage.At a depot other NDI methods,such as shearography or thermography,may be available.A good general reference on inspection methods is SAE ARP 5089 "Composite Repair NDI and NDT Handbook"(Reference 8.3.1).A summary of common nondestructive test methods and their utilization is shown in Table 8.3.1. 8-20
MIL-HDBK-17-3F Volume 3, Chapter 8 Supportability 8-20 hence, very large patch sizes. In these situations, repairs may be combinations of scarfed and external patches, so that the repair sizes can be minimized. Flush repairs may be required for some components for aerodynamic reasons or to prevent chafing. Also, thick face sheets require thick patches, which may require special processing to achieve proper consolidation. Patch and bondline porosity are of particular concern with normal field processing, which is accomplished with vacuum pressure and heat blankets. Lower temperature cures are generally preferred due to concerns over causing additional damage via vaporization of water that has infiltrated the core. Also, the surrounding structure may act as a heat sink, making it difficult to achieve and control the higher temperatures with heat blankets, and may contribute to thermal gradients that can result in warpage or degradation of the surrounding structure. For thick sandwich, heat blankets on both sides of the structure may be required to control the through thickness temperature. Still, the shorter processing times generally associated with higher temperature cures are very attractive in terms of minimizing the out-of-service time for a damaged airplane. Sandwich moisture ingression issues. Consideration must be given to moisture ingression when designing maintainable, repairable sandwich structures. Sandwich designs must address the effects of moisture in the core, both by minimizing the degree of moisture ingression, and by determining what its presence does to the performance of the structure. Moisture ingression can occur through face sheet damages, and part edge and end seals, so special care must be taken to design durable sandwich parts. Unfortunately, to make durable face skins, additional thickness is needed, and this may not be desirable from a performance point of view. Durable edge and end seals can be designed, see Ref. 8.2.1.1. When repairing damaged sandwich structures, a drying cycle is typically performed prior to the accomplishment of any bonded repair. This is performed so that any retained moisture does not interfere with the curing cycle. There have been numerous cases of face skins blowing off sandwich components during the vacuum bag heating cure cycle. 8.3 SUPPORT IMPLEMENTATION A repair has the objective of restoring a damaged structure to an acceptable capability in terms of strength, durability, stiffness, functional performance, safety, cosmetic appearance or service life. Ideally, the repair will return the structure to original capability and appearance. The design assessment of a repair for a given loading condition involves the selection of a repair concept, the choice of the appropriate repair materials and processes, then specifying the detailed configuration and size of the repair. Most repairs are basically designed as a joint to transfer load into and out of a patch. To ensure that the repair configuration will have adequate strength and stiffness, the repair joint must be analyzed to predict its strength. The selection of the type of load-transfer joint to be used for a patch/strap is a tradeoff between simplicity, strength and stiffness. The easier configurations are generally not as strong as the more difficult ones. It is critical that the materials and process information is available prior to the system being put into place. 8.3.1 Part Inspection Presence of damage in aircraft composite parts is usually found in the course of a routine on-line inspection, depot inspection, or, for large damages, noticed by the pilot. The predominant mode of inspection is visual with more sophisticated modes of inspection performed at the depot. Once damage is identified visually in-service, the damage should be characterized quantitatively by measuring dent depth, extent of surface damage, and length of scratches before proceeding to more complex NDI. This will generally consist of tap testing to define the boundary between damaged and undamaged portions of the structure and followed, for major repairs, with instrumented NDI techniques (ultrasonics, radiography, etc.) to locate the through the thickness characteristics of the damage. At a depot other NDI methods, such as shearography or thermography, may be available. A good general reference on inspection methods is SAE ARP 5089 “Composite Repair NDI and NDT Handbook” (Reference 8.3.1). A summary of common nondestructive test methods and their utilization is shown in Table 8.3.1