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. 8-5MIL-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