《纺织复合材料》课程参考文献（Composite Materials Handbook，Volume 3）Chapter 8 Supportability

MIL-HDBK-17-3F Volume 3,Chapter 8 Supportability CHAPTER 8 SUPPORTABILITY 8.1 INTRODUCTION Supportability is an integral part of the design process that ensures support requirements are incorpo- rated in the design and logistics resources are defined to support the system during its operating or useful life.Support resource requirements include the skills,tools,equipment,facilities,spares,techniques, documentation,data,materials,and analysis required to ensure that a composite component maintains structural integrity over its intended lifetime.When the load carrying capability of an aircraft,or product is compromised,(i.e..loss of design function),the damaged structure must be restored quickly and at low cost.Customer requirements can dictate maintenance philosophy,materials availability,and repair capa- bilities that a design team must incorporate throughout the design process.As the contributors to this chapter were primarily from the aircraft industry,the text is slanted towards its particular needs.However, the guiding principles can be beneficial in other composite applications. Since the operating and support cost of a vehicle continues to escalate throughout its life,it becomes imperative to select and optimize those designs that maximize supportability.Life cycle cost,being com- prised of research and development,acquisition,operational and support,and disposal costs,is often a crucial customer requirement for any new weapon system or commercial transport.Often,design changes that enhance producibility,improve vehicle availability,and reduce operational and support costs,far outweigh the short-term increases in acquisition costs.Lost airline profits and reduced wartime readiness are a direct result of designs that did not incorporate supportability early in the design process. Telltale indicators of non-supportable designs include expensive spares,excessive repair times,and un- needed inspections. Aircraft users are often constrained to perform maintenance during aircraft turnaround,after each day's usage,and during scheduled maintenance.Repair time limitations can range from several minutes to several days.In each case users of aircraft containing composite components require durable struc- tures that,when damaged,can be repaired within the available support infrastructure including skills,ma- terials,equipment,and technical data. Composite designs are usually tailored to maximize performance by defining application dependent materials,ply orientation,stiffening concepts,and attachment mechanisms.High performance designs are often less supportable due to increased strain levels,fewer redundant load paths,and a mix of highly tailored materials and geometries.Product design teams should focus on a variety of features that im- prove supportability including compatibility of available repair materials with those used on the parent structure,available equipment and skill,improving subsystem accessibility,and extended shelf-life com- posite repair materials.Structural elements and materials should be selected that are impervious to in- herent and induced damage especially delaminations,low velocity impacts,and hail damage.Each sup- portability enhancement feature results from the designer having an explicit knowledge of the aircraft's operational and maintenance environment and associated requirements and characteristics.Other de- sign considerations also have an impact on supportability including durability,reliability,damage toler- ance,and survivability.A supportable design integrates all the requirements,criteria,and features neces- sary to provide highly valued products in terms of performance,affordability and availability. This section is designed to assist integrated product teams in the development of supportable prod- ucts through five basic sections:1)Introduction-which provides an overview of the Supportability chap- ter;2)Design for Supportability-which provides the designer with design criteria,guidelines and check- lists to ensure a supportable design;3)Support Implementation-which defines and demonstrates those key elements of supportability that must be performed to insure mission success;4)Composite Repairs of Metal Structure-which provides an alternative means to standard metal repair options,and(5)Logistics Requirements-which establishes the support resources needed to maintain the backbone of the support structure.Each section provides the designer and aircraft user with the supportability data and lessons learned that will reduce cost of ownership and improve aircraft availability.Other sections throughout MIL-HDBK-17 discuss the details needed to design supportable components.Sections contained in Vol- 8-1

MIL-HDBK-17-3F Volume 3, Chapter 8 Supportability 8-1 CHAPTER 8 SUPPORTABILITY 8.1 INTRODUCTION Supportability is an integral part of the design process that ensures support requirements are incorpo￾rated in the design and logistics resources are defined to support the system during its operating or useful life. Support resource requirements include the skills, tools, equipment, facilities, spares, techniques, documentation, data, materials, and analysis required to ensure that a composite component maintains structural integrity over its intended lifetime. When the load carrying capability of an aircraft, or product is compromised, (i.e., loss of design function), the damaged structure must be restored quickly and at low cost. Customer requirements can dictate maintenance philosophy, materials availability, and repair capa￾bilities that a design team must incorporate throughout the design process. As the contributors to this chapter were primarily from the aircraft industry, the text is slanted towards its particular needs. However, the guiding principles can be beneficial in other composite applications. Since the operating and support cost of a vehicle continues to escalate throughout its life, it becomes imperative to select and optimize those designs that maximize supportability. Life cycle cost, being com￾prised of research and development, acquisition, operational and support, and disposal costs, is often a crucial customer requirement for any new weapon system or commercial transport. Often, design changes that enhance producibility, improve vehicle availability, and reduce operational and support costs, far outweigh the short-term increases in acquisition costs. Lost airline profits and reduced wartime readiness are a direct result of designs that did not incorporate supportability early in the design process. Telltale indicators of non-supportable designs include expensive spares, excessive repair times, and un￾needed inspections. Aircraft users are often constrained to perform maintenance during aircraft turnaround, after each day’s usage, and during scheduled maintenance. Repair time limitations can range from several minutes to several days. In each case users of aircraft containing composite components require durable struc￾tures that, when damaged, can be repaired within the available support infrastructure including skills, ma￾terials, equipment, and technical data. Composite designs are usually tailored to maximize performance by defining application dependent materials, ply orientation, stiffening concepts, and attachment mechanisms. High performance designs are often less supportable due to increased strain levels, fewer redundant load paths, and a mix of highly tailored materials and geometries. Product design teams should focus on a variety of features that im￾prove supportability including compatibility of available repair materials with those used on the parent structure, available equipment and skill, improving subsystem accessibility, and extended shelf-life com￾posite repair materials. Structural elements and materials should be selected that are impervious to in￾herent and induced damage especially delaminations, low velocity impacts, and hail damage. Each sup￾portability enhancement feature results from the designer having an explicit knowledge of the aircraft's operational and maintenance environment and associated requirements and characteristics. Other de￾sign considerations also have an impact on supportability including durability, reliability, damage toler￾ance, and survivability. A supportable design integrates all the requirements, criteria, and features neces￾sary to provide highly valued products in terms of performance, affordability and availability. This section is designed to assist integrated product teams in the development of supportable prod￾ucts through five basic sections: 1) Introduction - which provides an overview of the Supportability chap￾ter; 2) Design for Supportability - which provides the designer with design criteria, guidelines and check￾lists to ensure a supportable design; 3) Support Implementation - which defines and demonstrates those key elements of supportability that must be performed to insure mission success; 4) Composite Repairs of Metal Structure – which provides an alternative means to standard metal repair options, and (5) Logistics Requirements - which establishes the support resources needed to maintain the backbone of the support structure. Each section provides the designer and aircraft user with the supportability data and lessons learned that will reduce cost of ownership and improve aircraft availability. Other sections throughout MIL-HDBK-17 discuss the details needed to design supportable components. Sections contained in Vol-

MIL-HDBK-17-3F Volume 3,Chapter 8 Supportability ume 1 include material and structural testing,material types and properties,and joint types;in Volume 3 include materials and processes,quality,design,joints,reliability,and lessons learned needed to supple- ment those decisions that influence supportability. 8.2 DESIGN FOR SUPPORTABILITY 8.2.1 In-service experience The first step toward designing reliable and cost-effective design details is to understand the history of composite structure.Composite materials,as we know them today,were introduced into the commercial aircraft industry during the early 1960's and used mostly glass fiber.Development of more advanced fi- bers such as boron,aramid,and carbon offered the possibility of increased strength,reduced weight,im- proved corrosion resistance,and greater fatigue resistance than aluminum.These new material systems, commonly referred to as advanced composites,were introduced to the industry very gradually and cau- tiously to ensure their capabilities. The early success of the first simple components,such as wing spoilers and fairings,led to the use of advanced composites in more complex components such as ailerons,flaps,nacelles,and rudders.The increased specific stiffness and strengths of composites over aluminum,coupled with weight-driven re- quirements caused by fuel shortages,led to the application of thin-skin sandwich structures.Long-term durability requirements of the original aluminum parts were not fully accounted for when these composite parts were originally designed.To compound the problem further,damage phenomena such as delamination and microcracking were new and complex in comparison to traditional aluminum structure. The original composite parts,particularly thin-gage sandwich panels,experienced durability problems that could be grouped into three categories:low resistance to impact,liquid ingression,and erosion. These parts were either control panels or secondary structure,such as fixed trailing edge panels,and given the emphasis placed on weight and performance,the face sheets of honeycomb sandwich parts were often only three plies or less with a TedlarTM film.This approach was adequate for stiffness and strength,but never considered the service environment where parts are crawled over,tools dropped,and where service personnel are often unaware of the fragility of thin-skinned sandwich parts.Damages to these components,such as core crush,impact damages and disbonds,are quite often easily detected with a visual inspection due to their thin face sheets.However,sometimes they are overlooked,or dam- aged by service personnel,who do not want to delay aircraft departure or bring attention to their acci- dents,which might reflect poorly on their performance record.Therefore,damages are sometimes al- lowed to go unchecked,often resulting in growth of the damage due to liquid ingression into the core. Non-durable design details (e.g.,improper core edge close-outs)also led to liquid ingression. The repair of parts due to liquid ingression can vary depending upon the liquid,of which water and Skydrol(hydraulic fluid)are the two most common.Water tends to create additional damage in repaired parts when cured unless all moisture is removed from the part.Most repair material systems cure at temperatures above the boiling point of water,which can cause a disbond at the skin-to-core interface wherever trapped water resides.For this reason,core drying cycles are typically included prior to per- forming any repair.Some operators will take the extra step of placing a damaged but unrepaired part in the autoclave to dry so as to preclude any additional damage from occurring during the cure of the repair. This is done to assure they will only need to repair the part once.Skydrol presents a different problem. Once the core of a sandwich part is saturated,complete removal of Skydrol is almost impossible.The part continues to weep the liquid even in cure such that bondlines can become contaminated and full bonding does not occur.Removal of contaminated core and adhesive as part of the repair is highly rec- ommended. Erosion capabilities of composite materials have been known to be less than that of aluminum and, as a result,their application in leading edge surfaces has been generally avoided.However,composites have been used in areas of highly complex geometry,but generally with an erosion coating.The durabil- ity and maintainability of some erosion coatings are less than ideal.Another problem,not as obvious as 8-2

MIL-HDBK-17-3F Volume 3, Chapter 8 Supportability 8-2 ume 1 include material and structural testing, material types and properties, and joint types; in Volume 3 include materials and processes, quality, design, joints, reliability, and lessons learned needed to supple￾ment those decisions that influence supportability. 8.2 DESIGN FOR SUPPORTABILITY 8.2.1 In-service experience The first step toward designing reliable and cost-effective design details is to understand the history of composite structure. Composite materials, as we know them today, were introduced into the commercial aircraft industry during the early 1960's and used mostly glass fiber. Development of more advanced fi￾bers such as boron, aramid, and carbon offered the possibility of increased strength, reduced weight, im￾proved corrosion resistance, and greater fatigue resistance than aluminum. These new material systems, commonly referred to as advanced composites, were introduced to the industry very gradually and cau￾tiously to ensure their capabilities. The early success of the first simple components, such as wing spoilers and fairings, led to the use of advanced composites in more complex components such as ailerons, flaps, nacelles, and rudders. The increased specific stiffness and strengths of composites over aluminum, coupled with weight-driven re￾quirements caused by fuel shortages, led to the application of thin-skin sandwich structures. Long-term durability requirements of the original aluminum parts were not fully accounted for when these composite parts were originally designed. To compound the problem further, damage phenomena such as delamination and microcracking were new and complex in comparison to traditional aluminum structure. The original composite parts, particularly thin-gage sandwich panels, experienced durability problems that could be grouped into three categories: low resistance to impact, liquid ingression, and erosion. These parts were either control panels or secondary structure, such as fixed trailing edge panels, and given the emphasis placed on weight and performance, the face sheets of honeycomb sandwich parts were often only three plies or less with a Tedlar™ film. This approach was adequate for stiffness and strength, but never considered the service environment where parts are crawled over, tools dropped, and where service personnel are often unaware of the fragility of thin-skinned sandwich parts. Damages to these components, such as core crush, impact damages and disbonds, are quite often easily detected with a visual inspection due to their thin face sheets. However, sometimes they are overlooked, or dam￾aged by service personnel, who do not want to delay aircraft departure or bring attention to their acci￾dents, which might reflect poorly on their performance record. Therefore, damages are sometimes al￾lowed to go unchecked, often resulting in growth of the damage due to liquid ingression into the core. Non-durable design details (e.g., improper core edge close-outs) also led to liquid ingression. The repair of parts due to liquid ingression can vary depending upon the liquid, of which water and Skydrol (hydraulic fluid) are the two most common. Water tends to create additional damage in repaired parts when cured unless all moisture is removed from the part. Most repair material systems cure at temperatures above the boiling point of water, which can cause a disbond at the skin-to-core interface wherever trapped water resides. For this reason, core drying cycles are typically included prior to per￾forming any repair. Some operators will take the extra step of placing a damaged but unrepaired part in the autoclave to dry so as to preclude any additional damage from occurring during the cure of the repair. This is done to assure they will only need to repair the part once. Skydrol presents a different problem. Once the core of a sandwich part is saturated, complete removal of Skydrol is almost impossible. The part continues to weep the liquid even in cure such that bondlines can become contaminated and full bonding does not occur. Removal of contaminated core and adhesive as part of the repair is highly rec￾ommended. Erosion capabilities of composite materials have been known to be less than that of aluminum and, as a result, their application in leading edge surfaces has been generally avoided. However, composites have been used in areas of highly complex geometry, but generally with an erosion coating. The durabil￾ity and maintainability of some erosion coatings are less than ideal. Another problem, not as obvious as

MIL-HDBK-17-3F Volume 3,Chapter 8 Supportability the first,is that edges of doors or panels can erode if they are exposed to the air stream.This erosion can be attributed to improper design or installation/fit-up.On the other hand,metal structures in contact or in the vicinity of these composite parts may show corrosion damage due to: Inappropriate choice of aluminum alloy Damaged corrosion sealant of metal parts during assembly or at splices Insufficient sealant and/or lack of glass fabric isolation plies at the interfaces of spars,ribs and fit- tings Assessing operator experience with composite structure is,taken as a whole,an extremely difficult task.A survey of operators provides responses depending on the composite application ranging from horror stories for thin skinned sandwich structures,to outstanding success for thick skinned sandwich or solid laminate primary structures.Some of the facts and data that are available are the detailed reports that were received from the operators on parts involved in the NASA-sponsored Advanced Composites Energy Efficiency(ACEE)program,which supported the design and fabrication of composite parts such as the B727-200 elevators and the B737 spoilers and horizontal stabilizers.Five shipsets of B727 eleva- tors have accumulated more than 331,000 hrs.and 189,000 cycles;108 B737 spoilers have accumulated more than 2,888,000 hrs.and 3,781,000 cycles.Five shipsets of B737 horizontal stabilizers,which incor- porated laminate torque boxes and sandwich ribs,had amassed over 133,500 flight hours and 130,000 landings as of May.1995.The service exposure data collected for these parts have not indicated any durability or corrosion problems.One B737-200 aircraft with the ACEE stabilizers was removed from ser- vice after 19,295 flight cycles and 17,302 flight hours,and one stabilizer was acquired by Boeing for a detailed tear-down inspection.The stabilizer was found to be in excellent condition with no fatigue dam- age,and the only corrosion discovered was some minor pitting found in some fastener holes of the alu- minum trailing edge fittings.This was determined to be due to a fastener sealing practice which has since been obsoleted.Several repairs have been satisfactorily performed on the 727 elevators and remaining 737 horizontal stabilizers which are still in service. The in-service success of these ACEE components is in part due to the integrated teams which de- veloped them.The teams for both the B727 sandwich elevators and the B737 stiffened-skin configured horizontal stabilizers considered maintainability during the developmental programs.They devised repair and inspection schemes,and for each component,Maintenance Planning Manuals were compiled and released as part of the NASA contractual obligation.The airlines,United for the ten B727 elevators,and Delta and Mark Air for the five shipsets of B737 stabilizers,were in essence part of the teams who planned these documents.As mentioned above,both of these components have been damaged and repaired using the repair schemes designed for them.In all of the instances,the repairs were satisfacto- rily performed in-place on the aircraft. An in-service evaluation,launched in 1980,with twenty-two airbrakes/spoilers(14 fabricated with car- bon-epoxy tape,and 8 fabricated from carbon-epoxy fabric)installed on Air France A300 aircraft,is still going on.Non-destructive inspections(visual and ultrasonic)are performed on aircraft and in the labora- tory during the service life.Thirteen airbrakes are still on aircraft,and seven have been withdrawn from service for testing to assess stiffness and residual strength.As of November,1995,these components had accumulated 405,698 flight hours and 236,588 flight cycles.The component with the most time in service had accumulated 32,069 flight hours and 16,802 flight cycles.Bolted repairs(metal patches for temporary,and composite precured patches for permanent repairs)were designed.Two components have been repaired with blind fasteners to arrest manufacturing produced disbonds between the skins and ribs.Some minor corrosion pitting was found on the aluminum(7075)spar at the central fitting splice due to the protect scheme having been damaged and not restored during assembly.A modification of the trailing edge was implemented early in the program;the rubber one being replaced by a solid carbon one. As an example of successful thicker solid laminate structure,the ATR 72 outer wing box has accumu- lated 1,429,539 flight cycles and 1,163,333 flight hours since entering into service in 1989.The aircraft with the most time service has accumulated 23,343 flight cycles and 14,988 flight hours.Service experi- ence has been very good with only one accidental damage being reported:an aircraft crashed into a han- gar door at a speed of 15 miles/hr(25 km/hr).The composite outer wing box was repaired using bolted 8-3

MIL-HDBK-17-3F Volume 3, Chapter 8 Supportability 8-3 the first, is that edges of doors or panels can erode if they are exposed to the air stream. This erosion can be attributed to improper design or installation/fit-up. On the other hand, metal structures in contact or in the vicinity of these composite parts may show corrosion damage due to: • Inappropriate choice of aluminum alloy • Damaged corrosion sealant of metal parts during assembly or at splices • Insufficient sealant and/or lack of glass fabric isolation plies at the interfaces of spars, ribs and fit￾tings Assessing operator experience with composite structure is, taken as a whole, an extremely difficult task. A survey of operators provides responses depending on the composite application ranging from horror stories for thin skinned sandwich structures, to outstanding success for thick skinned sandwich or solid laminate primary structures. Some of the facts and data that are available are the detailed reports that were received from the operators on parts involved in the NASA-sponsored Advanced Composites Energy Efficiency (ACEE) program, which supported the design and fabrication of composite parts such as the B727-200 elevators and the B737 spoilers and horizontal stabilizers. Five shipsets of B727 eleva￾tors have accumulated more than 331,000 hrs. and 189,000 cycles; 108 B737 spoilers have accumulated more than 2,888,000 hrs. and 3,781,000 cycles. Five shipsets of B737 horizontal stabilizers, which incor￾porated laminate torque boxes and sandwich ribs, had amassed over 133,500 flight hours and 130,000 landings as of May, 1995. The service exposure data collected for these parts have not indicated any durability or corrosion problems. One B737-200 aircraft with the ACEE stabilizers was removed from ser￾vice after 19,295 flight cycles and 17,302 flight hours, and one stabilizer was acquired by Boeing for a detailed tear-down inspection. The stabilizer was found to be in excellent condition with no fatigue dam￾age, and the only corrosion discovered was some minor pitting found in some fastener holes of the alu￾minum trailing edge fittings. This was determined to be due to a fastener sealing practice which has since been obsoleted. Several repairs have been satisfactorily performed on the 727 elevators and remaining 737 horizontal stabilizers which are still in service. The in-service success of these ACEE components is in part due to the integrated teams which de￾veloped them. The teams for both the B727 sandwich elevators and the B737 stiffened-skin configured horizontal stabilizers considered maintainability during the developmental programs. They devised repair and inspection schemes, and for each component, Maintenance Planning Manuals were compiled and released as part of the NASA contractual obligation. The airlines, United for the ten B727 elevators, and Delta and Mark Air for the five shipsets of B737 stabilizers, were in essence part of the teams who planned these documents. As mentioned above, both of these components have been damaged and repaired using the repair schemes designed for them. In all of the instances, the repairs were satisfacto￾rily performed in-place on the aircraft. An in-service evaluation, launched in 1980, with twenty-two airbrakes/spoilers (14 fabricated with car￾bon-epoxy tape, and 8 fabricated from carbon-epoxy fabric) installed on Air France A300 aircraft, is still going on. Non-destructive inspections (visual and ultrasonic) are performed on aircraft and in the labora￾tory during the service life. Thirteen airbrakes are still on aircraft, and seven have been withdrawn from service for testing to assess stiffness and residual strength. As of November, 1995, these components had accumulated 405,698 flight hours and 236,588 flight cycles. The component with the most time in service had accumulated 32,069 flight hours and 16,802 flight cycles. Bolted repairs (metal patches for temporary, and composite precured patches for permanent repairs) were designed. Two components have been repaired with blind fasteners to arrest manufacturing produced disbonds between the skins and ribs. Some minor corrosion pitting was found on the aluminum (7075) spar at the central fitting splice due to the protect scheme having been damaged and not restored during assembly. A modification of the trailing edge was implemented early in the program; the rubber one being replaced by a solid carbon one. As an example of successful thicker solid laminate structure, the ATR 72 outer wing box has accumu￾lated 1,429,539 flight cycles and 1,163,333 flight hours since entering into service in 1989. The aircraft with the most time service has accumulated 23,343 flight cycles and 14,988 flight hours. Service experi￾ence has been very good with only one accidental damage being reported; an aircraft crashed into a han￾gar door at a speed of 15 miles/hr (25 km/hr). The composite outer wing box was repaired using bolted

MIL-HDBK-17-3F Volume 3,Chapter 8 Supportability carbon-epoxy and metal patches (see Figure 8.2.1(a)),while all the metal parts of the center box were replaced due to permanent deformations.One aircraft exhibited erosion of the outer ply at the leading edge of the upper skin,and a chamfer was introduce in the design,and no other problems have been reported. a)Damaged ATR 72 carbon outer wingbox:carbon front spar and carbon wing skins b)Repaired carbon spar before repair of wing skins FIGURE 8.2.1(a)Repairs to badly damaged ATR 72 wing. Production carbon-epoxy sandwich parts,such as trailing edge panels,cowls,landing gear doors, and fairings have demonstrated weight reduction,delamination resistance,fatigue improvement and cor- rosion prevention.The poor service records of some parts can be attributed to fragility,the inclusion of non-durable design details,poor processing quality,porous face sheets(insufficient thickness),and badly installed or poorly sealed fasteners.Many of the design problems were a result of insufficient technology 8-4

MIL-HDBK-17-3F Volume 3, Chapter 8 Supportability 8-4 carbon-epoxy and metal patches (see Figure 8.2.1(a)), while all the metal parts of the center box were replaced due to permanent deformations. One aircraft exhibited erosion of the outer ply at the leading edge of the upper skin, and a chamfer was introduce in the design, and no other problems have been reported. a) Damaged ATR 72 carbon outer wingbox: carbon front spar and carbon wing skins b) Repaired carbon spar before repair of wing skins FIGURE 8.2.1(a) Repairs to badly damaged ATR 72 wing. Production carbon-epoxy sandwich parts, such as trailing edge panels, cowls, landing gear doors, and fairings have demonstrated weight reduction, delamination resistance, fatigue improvement and cor￾rosion prevention. The poor service records of some parts can be attributed to fragility, the inclusion of non-durable design details, poor processing quality, porous face sheets (insufficient thickness), and badly installed or poorly sealed fasteners. Many of the design problems were a result of insufficient technology

MIL-HDBK-17-3F Volume 3,Chapter 8 Supportability transfer from development programs such as NASA-ACEE.Many flight control and secondary structural panels were designed using composite materials without consideration of the applicability of composites and the service environment.As a result,many components were designed around weight-efficient sandwich configurations with face sheets of only two or three plies.Not only is the damage resistance of these components poor,but they are difficult to seal from fluids. Fragility,so much an issue in these thin-gage sandwich structures,is much less an issue in thicker- gage primary structures (sandwich or solid laminate)such as main torque boxes of empennages and wings,and fuselages.The thicker skins of the Boeing B777 and the Airbus series of composite empen- nage main torque boxes,the US Air Force B-2 wings and fuselage,the ATR 72.US Air Force F-22,the US Navy/Marine F/A 18,and the RAF/Royal Navy/US Marines AV8B Harrier outboard wing boxes,to list a few examples,are much more damage resistant than the vast number of light gage sandwich flight con- trol and secondary structural components that are currently in service,and are still being introduced to service.In addition to their highly damage resistant primary structural components,the latest Airbus air- craft and the B777 have incorporated improved light gage composite structural designs.However,they will still be more vulnerable to damage in service than the primary structural components,due to their minimum gage,mainly sandwich,construction. 8.2.2 Inspectability During the design of composite structural components consideration should be given to the inspection methods available to both the manufacturer and the customer.Typical composite in-process non- destructive inspection(NDI)methods available to the manufacturer are:visual,through-transmission ul- trasonics(TTU),pulse-echo ultrasonics,x-ray,and other advanced NDI methods such as enhanced opti- cal schemes and thermography.Most airlines and military operators use visual inspections supplemented with both mechanical(i.e.,some form of tap test)and electronic(i.e.,pulse echo and low-frequency bond testing)to locate damage.Because of the predominance of visual inspections,provisions should be made during the design phase for complete external and internal access for visual inspection of all components, regardless of whether they are critical primary structural components or secondary structures such as fairings.If a visual inspection indicates potential damage,then the more sophisticated inspection tech- niques can be used to provide more accurate damage assessments.Additional suggestions can be found in Section 8.3.1. 8.2.2.1 General design guidelines Whether one chooses a laminate stiffened skin or a sandwich configuration for a specific component, there are inspectability issues within each configuration category.For example,the use of closed hat stiffeners to stiffen laminate skins.while extremely efficient from a structural point of view.create three areas in the skin and stiffener that are difficult to inspect by any method(Figure 8.2.2.1(a),section(a)).A blade stiffener,on the other hand,has only the one difficult inspection area(Figure 8.2.2.1(a),section (b)). The adhesive fillets of the closed-hat stiffener.and the rolled noodle of the blade stiffener.are contributors to these inspection difficulties.These areas are difficult to inspect during the manufacturing process,and are even more of a problem for the service operator with limited access to the internal surfaces. With a sandwich configuration there are inspection difficulties associated with potted areas,detection of fluids that have leeched into the sandwich honeycomb core,disbonds of face sheets,foam core,and damages within the core.Also difficult for operators are inspections of bondlines of stiffeners or frames that are bonded to the internal face sheets of sandwich components(Figure 8.2.2.1(b)).When airplane operators are forced to use inspection methods that are subjective,i.e.,the tap test,they are handi- capped by lack of knowledge of damage sizes and criticality.This is a significant problem for operators, and while sandwich structural configurations can be very efficient from a performance point of view,they tend to be fragile,easily damaged,and difficult to inspect.Interestingly some airline operators prefer sandwich over laminate stiffened skins from a repair point of view,but virtually all express frustration with the durability and inspection of sandwich structures. 8-5

MIL-HDBK-17-3F Volume 3, Chapter 8 Supportability 8-5 transfer from development programs such as NASA-ACEE. Many flight control and secondary structural panels were designed using composite materials without consideration of the applicability of composites and the service environment. As a result, many components were designed around weight-efficient sandwich configurations with face sheets of only two or three plies. Not only is the damage resistance of these components poor, but they are difficult to seal from fluids. Fragility, so much an issue in these thin-gage sandwich structures, is much less an issue in thicker￾gage primary structures (sandwich or solid laminate) such as main torque boxes of empennages and wings, and fuselages. The thicker skins of the Boeing B777 and the Airbus series of composite empen￾nage main torque boxes, the US Air Force B-2 wings and fuselage, the ATR 72, US Air Force F-22, the US Navy/Marine F/A 18, and the RAF/Royal Navy/US Marines AV8B Harrier outboard wing boxes, to list a few examples, are much more damage resistant than the vast number of light gage sandwich flight con￾trol and secondary structural components that are currently in service, and are still being introduced to service. In addition to their highly damage resistant primary structural components, the latest Airbus air￾craft and the B777 have incorporated improved light gage composite structural designs. However, they will still be more vulnerable to damage in service than the primary structural components, due to their minimum gage, mainly sandwich, construction. 8.2.2 Inspectability During the design of composite structural components consideration should be given to the inspection methods available to both the manufacturer and the customer. Typical composite in-process non￾destructive inspection (NDI) methods available to the manufacturer are: visual, through-transmission ul￾trasonics (TTU), pulse-echo ultrasonics, x-ray, and other advanced NDI methods such as enhanced opti￾cal schemes and thermography. Most airlines and military operators use visual inspections supplemented with both mechanical (i.e., some form of tap test) and electronic (i.e., pulse echo and low-frequency bond testing) to locate damage. Because of the predominance of visual inspections, provisions should be made during the design phase for complete external and internal access for visual inspection of all components, regardless of whether they are critical primary structural components or secondary structures such as fairings. If a visual inspection indicates potential damage, then the more sophisticated inspection tech￾niques can be used to provide more accurate damage assessments. Additional suggestions can be found in Section 8.3.1. 8.2.2.1 General design guidelines Whether one chooses a laminate stiffened skin or a sandwich configuration for a specific component, there are inspectability issues within each configuration category. For example, the use of closed hat stiffeners to stiffen laminate skins, while extremely efficient from a structural point of view, create three areas in the skin and stiffener that are difficult to inspect by any method (Figure 8.2.2.1(a), section (a)). A blade stiffener, on the other hand, has only the one difficult inspection area (Figure 8.2.2.1(a), section (b)). The adhesive fillets of the closed-hat stiffener, and the rolled noodle of the blade stiffener, are contributors to these inspection difficulties. These areas are difficult to inspect during the manufacturing process, and are even more of a problem for the service operator with limited access to the internal surfaces. With a sandwich configuration there are inspection difficulties associated with potted areas, detection of fluids that have leeched into the sandwich honeycomb core, disbonds of face sheets, foam core, and damages within the core. Also difficult for operators are inspections of bondlines of stiffeners or frames that are bonded to the internal face sheets of sandwich components (Figure 8.2.2.1(b)). When airplane operators are forced to use inspection methods that are subjective, i.e., the tap test, they are handi￾capped by lack of knowledge of damage sizes and criticality. This is a significant problem for operators, and while sandwich structural configurations can be very efficient from a performance point of view, they tend to be fragile, easily damaged, and difficult to inspect. Interestingly some airline operators prefer sandwich over laminate stiffened skins from a repair point of view, but virtually all express frustration with the durability and inspection of sandwich structures