Composites in General Aviation

6.09 Composites in General Aviation RIC ABBOTT Raytheon Aircraft Company, Wichita, KS, USA 6.09.1 INTRODUCTION

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6.09.2 COMPOSITE USAGE IN GENERAL AVIATION PRODUCTS

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6.09.3 SIMPLE DESIGN GUIDELINES

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6.09.4 COST OF WEIGHT SAVED

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6.09.4.1 Cost/Weight Analysis 6.09.4.2 Cost of Weight Saved

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6.09.5 CERTIFICATION REQUIREMENTS

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6.09.6 TYPICAL CERTIFICATION PLAN

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6.09.7 DAMAGE TOLERANCE EVALUATION

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6.09.7.1 Damage Scenarios 6.09.7.2 Damage Tolerance Tests

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6.09.8 DAMAGE TOLERANCE TEST RESULTS

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6.09.9 INSPECTIONS

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6.09.10 STRUCTURAL REPAIR MANUAL

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6.09.11 SERVICE EXPERIENCE

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6.09.12 CRASHWORTHINESS

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6.09.13 CONCLUSIONS

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6.09.14 REFERENCES

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6.09.1

INTRODUCTION

and in the early 1970s from Cessna. Since then the business has expanded hugely, both at the entry level with sporty fiberglass four-seaters selling for less than $200 000 and at the upper level with intercontinental business jets selling for up to $40 million. Given the wide range of products and market segments, it is to be expected that composite usage also varies widely depending on the manufacturer and the market segment. This usage runs the gamut from hand laid, wet resin saturated, room temperature cured fiberglass to machine laid, autoclave cured, carbon fiber prepreg. In the middle are some of the ªbig ironº of today's market: the Cessna Citation X utilizes composites in control surfaces and

The general aviation industry includes a wide spectrum of manufacturers, products, operators, and service centers. The traditional view of US general aviation manufacturers brings to mind the old big three of small airplanes: Cessna, Beech, and Piper. Except, Cessna is now part of Textron, Beech is now Raytheon Aircraft, and Piper is back in business after several changes of ownership. Their products in the 1960s comprised a range of single piston powered airplanes plus an upper echelon of twins and turboprops. Business jets appeared in the 1950s from Lockheed and Sabreliner, in the 1960s from Learjet and Hawker Siddeley, 1

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Composites in General Aviation

Figure 1 Cirrus SR 20 (photo courtesy of Cirrus Design).

pressure bulkheads as does the Gulfstream V. There is also a trend to produce lower cost parts by employing resin injection methods and one shot, in-tool cure cycles. Why use composites in the first place? The benefits of glass fiber reinforced epoxy (GFRE) are not obvious from a strictly scientific point of view. GFRE is about the same density as aluminum alloy and, after lamination, the axial tensile modulus is less than that of aluminum. The benefits of GFRE lie in low cost materials, low cost tooling, and the ability to mold large and beautifully contoured aerodynamic surfaces. Thus, GFRE is usually the material of choice for the aforementioned sporty four-seaters. Oven cured prepreg or wet lay-up methods facilitate the fabrication of airframes with less parts, less assembly cost, and with modern airfoils which have lower drag than traditional riveted metal construction, not to mention a much more sleek and appealing appearance. Bonus features are freedom from corrosion and fatigue cracking. GFRE combined with honeycomb core is used for fairings and trailing edges for the same reasons on many larger airplanes including the Boeing 747. The introduction of carbon fibers in the 1960s gave the more adventurous designer a whole new world of opportunityÐand a whole new world of lessons to learn. These fibers generally are preimpregnated with epoxy resin and typically require autoclave curing. In simple terms, starting with standard modulus carbon fiber which has 227.5 GPa (33 million psi) typical tensile modulus and

3620 MPa (525 000 psi) typical tensile strength, resin is added for a 60% fiber volume and arrive at a unidirectional prepreg with 138 GPa (20 million psi) typical axial modulus and over 2000 MPa (300 000 psi) typical axial tensile strength. It is not usable for most aerostructures in a purely unidirectional fiber form, but by laying plies in multiple directions (usually orthogonal directions) a usable laminate can be created while trying to maintain about 60% fibers in the primary load direction. The resulting material will have about 83 GPa (12 million psi) axial modulus and over 1240 MPa (180 000 psi) tensile strength. So, it is stronger than aluminum and stiffer than aluminum and it is 58% of the density. So where is the catch? Why don't all new airplanes utilize carbon fiber epoxy to the maximum practical extent? There are three main reasons why these advanced materials are slower to find widespread acceptance than one would expect after studying the apparently fabulous properties: (i) much higher material cost than aluminum, (ii) lack of widely accepted standards for design and certification, and (iii) some serious concerns about maintainability. Or, as some would put it, cost, cost, and . . . cost. 6.09.2

COMPOSITE USAGE IN GENERAL AVIATION PRODUCTS

The Cirrus SR20, shown in Figure 1, is typical of the new generation of entry-level private airplanes. This sleek, 200 mph personal

Composite Usage in General Aviation Products

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Figure 2 Cirrus SR 20 fuselage assembly (photo courtesy of Cirrus Design).

Figure 3 Cessna Citation X (photo courtesy of Cessna Aircraft).

transportation package sells for about $170 000 and was certified to FAR 23 rules in 1998. All major structures, except for the control surfaces, are fabricated using hand laid, fiberglass prepreg, oven cured under vacuum bag at 135 8C (275 8F) (see Figure 2). In small airplanes it is very difficult to build a composite control surface to compete in weight and cost with aluminum surfaces which can be built from sheet aluminum with minimum thickness as low as 0.020 in.

The opposite end of the US general aviation spectrum is filled by the Cessna Citation X and the Gulfstream GV. Both of these airplanes make extensive use of carbon fiber reinforced epoxy (CFRE) but not in primary wing, fuselage, or empennage structures. The Gulfstream GV has landing gear, doors, rudder, elevators, aft pressure bulkhead, and fairings manufactured from CFRE. The Citation X shown in Figure 3 employs CFRE in all control surfaces, and fairings. The aft pressure bulkhead in the

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Composites in General Aviation

Figure 4 Raytheon Premier I.

Citation X is a fabricated from Kevlar fabric prepreg. In both the Citation X and the Gulfstream V, the CFRE parts are hand laid and autoclave cured at 177 8C (350 8F). Designed and priced to be at the entry threshold for business jets, the Raytheon Premier I sells for about $4.25 million. This airplane embodies the most aggressive use yet of CFRE composites outside of the military (see Figure 4). Updated manufacturing methods are employed in construction of the fuselage, horizontal tail, vertical tail, flaps, ailerons, and spoilers. The fuselage is the major CFRE component and is of honeycomb shell construction. As part of a push for greater efficiency, the fuselage shells are fabricated on special tooling and utilize the Cincinnati-Machine Viper Automated Fiber Placement (AFP) system. The tooling concept and the AFP system allow fabrication of the fuselage shell in just two piecesÐthe forward fuselage (pressure cabin and nose section) and the aft fuselage (a nonpressurized section supporting the empennage and engines and containing most of the of the airplane electrical, pressurization, and heating/ cooling equipment). The AFP system is controlled by Acraplace software that translates CATIA solid models representing the fuselage composite design into

precise three-dimensional (3-D) instructions. This enables the machine to place narrow tapes of carbon fiber prepreg into exact position on a male tool to achieve maximum strength with minimum weight, see Figure 5. The machine builds up carbon fiber plies for the inner and outer face sheets in between which is honeycomb core. The final ply on the outer surface is a hybrid fabric woven of carbon fiber and metal wire for lightning strike protection. Presently, this outer fabric ply is hand laid. This type of construction exhibits high resistance to bending and buckling and does away with the maze of internal frames, stringers, longerons, rivets, doublers, and subassemblies seen in conventional metal thin skin assembly. It is, in fact, a true monocoque structure with the required strength and stiffness built into the shell itself. When each fuselage section is finished on the AFP system, it is transferred to a female mold for autoclave cure. In the case of the forward fuselage the result is a pressure cabin formed in a single piece with integral window, windshield, and doorframes, with no joints, rivets, or leak paths. This cabin weighs about 20% less than the equivalent structure in a metallic business jet, see Figure 6.

Simple Design Guidelines

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Figure 5 Automated fiber placement.

Figure 6 One-Shot-Cure pressure cabin shell.

The wing movable surfaces, spoilers, flaps, and ailerons are constructed with CFRE and are also being produced with new and more efficient processes. The wing flaps for the Premier I (see Figure 7) are manufactured using resin transfer molding (RTM). In this process, dry fiber preforms, in both braided sleeve and woven fabric styles, are draped over male tools (mandrels) to form the spars and skins of a multicell airfoil section. The closed mold is evacuated of air by vacuum pump, after which hot liquid resin is injected under high pressure. The mold temperature is then raised to the resin cure temperature at 177 8C (350 8F). The process is less labor intensive than hand lay-up as the use of braided and woven preforms saves cutting, placing, and splicing small pieces from rolls of fabric or tape. Also, this is an in-tool cure process that avoids moving the

tool to the autoclave, uses almost nothing in the way of secondary materials such as vacuum bags and breather plies, and leaves the autoclave free for larger structures. The spoilers and ailerons are also fabricated and cured in individual molds. But the process for these surfaces utilizes internal pressure molding (IPM) in which CFRE prepreg is placed in a closed mold and cured under internal pressure applied by inflatable bladders.

6.09.3

SIMPLE DESIGN GUIDELINES

A designer newly assigned to work on composite structures will be required to absorb a large amount of new data and to learn new techniques. They may also be confused by

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Composites in General Aviation

Figure 7 Resin transfer molded wing flap.

new restrictions from the technical team. Some of these will be simply obfuscation and scare stories amounting to nothing more than ªdesign by voodoo.º A sensible designer will retain everything previously understood about metal load-path design and add some simple guidelines that take into account fibrous load paths, anisotropic characteristics, electrical properties, and in-service durability requirements. The certification regulations must be met, of course, and these are discussed later, but one should always remember that the regulations deal with safety and don't include rules for durability (how long a part will last) or economics (how much a part will cost to own and operate). Just as in aluminum design, designing to absolutely minimum standards may result in a part that will not hold up in service. Recognition of fibrous load path design is the major philosophical change for successful composite design. This involves designing laminates that maintain a large percentage of the fibers in the axial direction and not loading laminates through the thickness. A simple rule of mixtures given will provide initial estimates of strength and stiffness, but there must be an awareness of the notch sensitivity of fiber dominated structures. Some extra axial fibers and some extra +458 fibers must be present to transfer loads around stress concentrations created by cutouts, holes, and corners. Application of a first-ply failure criterion is the best way to calculate the maximum allowable load on a particular laminate (i.e., because strength design, stability, and buckling issues in thinner laminates will often restrict the allowable load to less than that permitted by design based purely on strength). Extensive testing is another acceptable way to establish laminate allowables, but this would be very expensive to apply to every laminate in an all-composite design. First-ply failure analysis is admittedly a contentious subject (Sun et al., 1996). In fact, some well-known experts in the business propose that

analysis indicating ply failure in the transverse axis (in other words resin failure) should be ignored as not a real failure mode. But in the experience of the author, these proposals ignore the realities of complex loading in some primary structural parts. In the absence of test data from multidirectional load testing, a simple failure strain cut-off will ensure that loading in one direction will not degrade the strength of plies needed to carry loads in another direction. For example, a special case of multidirectional loading is the top of an aircraft pressure cabin which must carry tension from fuselage bending, shear from fuselage torsion, and hoop tension from internal pressure. Use of first-ply failure as a design criterion will not necessarily create a significant weight penalty as modern resins with improved strain to failure properties are available. A toughened epoxy unidirectional prepreg is in use at Boeing and Raytheon Aircraft that has a typical strain to failure in the transverse axis of over 1.2%. Also, it is recommended that the first-ply failure load in the transverse axis be equated only to a limit load condition. In other words, the calculated failure load of a transversely loaded ply can be exceeded between limit load and ultimate load but not below limit load. Limit load is the maximum design load (or the maximum expected load in a vehicle lifetime), and ultimate load is the limit load increased by a factor of safety, typically 1.5 in civil airplanes. This creates a criterion analogous to that applied to the elastic limit in metal aircraft design. The FAA regulations forbid detrimental permanent deformation below limit load. In the composite analogy, application of first-ply failure criterion practically eliminates the chance of resin cracking below limit load. And after all, as the author often reminds young designers, it is important to keep the rain out! The in-service durability of CFRE structures can be ensured by designing for ultimate load capability with barely detectable impact damage in the laminate, or as it is called in the

Certification Requirements FAR 23 regulations, impact damage at the Threshold of Detectability (TOD). This works due to the fact that TOD impact damage in moderate gauge laminates will induce mainly resin cracking and this damage can be used to represent minor flaws intrinsic to the manufacturing process. When this slightly damaged strength value is used as the ultimate strength property for a laminate it will have the effect of driving the working stresses down below the flaw growth threshold. This is because of a guideline that has been developed after extensive testing both on GFRE and CFRE specimens: there is very little flaw growth below about 50% of the static failure load of a damaged laminate. So, if a laminate damaged with TOD impact damage or containing intrinsic manufacturing defects will carry ultimate load, it follows that in-service load cycles will not cause flaw growth because in civil airplanes these cycles are never more than of two-thirds of limit load. DO NOT APPLY THIS GUIDELINE TO FIGHTER AIRCRAFT OR HELICOPTERS WITHOUT FURTHER VALIDATION. In the first case, the loads are too high; and in the second case, the cycles are too many. In compression critical laminates, designing for ultimate load in the presence of TOD impact damage will probably dominate all other constraints. Analyses of compression strength after impact (CAI) are complicated and will need validation before being accepted in a certification program (Kassapoglou, 1988, 1996; Kassapoglou et al., 1988; Kassapoglou and Abbott, 1988). Therefore, CAI strength values are usually established by well-accepted test methods (SACMA, 1994). Selected test results indicating typical values for damaged specimens are shown in Section 6.09.8.

6.09.4

COST OF WEIGHT SAVED

High-performance composite structures must earn their way into today's cost conscious airplanes by saving weight in addition to the other benefits previously described. The weight saved by substitution of a carefully designed CRFE part for an aluminum part will be about 20% of the aluminum part weight. The cost of 33 million modulus fiber in a unidirectional prepreg certified to aerospace standards is about $65 per lb, about $91 after taking into account cutting losses. The best hand lay-up rates yield a cured part cost of about $219 per lb. The latest machine laid rates are better with cured part costs in the $155 per lb range, very good by aerospace standards, but still

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higher than riveted aluminum parts, which typically cost about $115 per lb. The extra cost of a hand-laid CFRE part over that of a riveted aluminum part may be justified by comparing the extra cost to what the customer is being asked to pay to lift a pound of payload, as shown in the following example:

6.09.4.1

Cost/Weight Analysis

Capital cost of transport capacity, P = (G±n)/d where vehicle gross weight = G, vehicle empty weight = n, and vehicle capital cost = d Cost of riveted aluminum structure, A = 2 labor hours + material Typically, A = $115 per lb Cost of large hand-laid aerospace CFRE structures, C = 2.33 labor hours + material Typically, C = $219 per lb Typical weight reduction when replacing aluminum with CFRE: 20% Cost of replacing a 100-lb aluminum part: = (80 6 C)±(100 6 A) = (80 6 219)±(100 6 115) = $6020

6.09.4.2

Cost of Weight Saved

S = 6020/20 S = $301 per lb S5P, extra cost is justified S>P, extra cost is not justified

Machine lay-up or resin injection processes can sway the balance more toward CFRE parts. However, smaller CFRE parts may be uneconomical when measured by the same analysis due to certain costs that are fixed per part.

6.09.5

CERTIFICATION REQUIREMENTS

The first all-carbon fiber airplane to be certified by the FAA was the Beech Starship. At the time the Starship was developed (early 1980s) the FAA certification regulations did not include any specific mention of composite structures or adhesive bond assembly. After some study, Raytheon engineers proposed to the FAA a certification plan based on published advisory material (FAA, 1984) and certain Boeing specifications to which rights had been

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Composites in General Aviation

purchased. This plan included damage tolerance evaluation for composite structure. Damage tolerance certification had been a requirement since 1978 for Transport Airplanes (airplanes over 6000 kg take-off weight) (CFR, 1996a), but in 1984 the regulations for small airplanes did not require, or even give the option of, damage tolerance certification. Meetings took place with FAA engineers of the Small Airplane Directorate in Kansas City and with the National Resource Specialist for Composites, and in August 1988, the plan was published in the Federal Register in the form of Special Conditions (new regulations applicable to a specific project). Since then the regulations originally published for the Starship have been incorporated into the FAA Small Airplane certification regulations (CFR, 1996b). A damage tolerance evaluation is now required for small airplane composite structures and may be chosen as an option for metallic structures. The new regulations, although requiring some additional expertise in the behavior of composites, actually provide a benefit to manufacturers in that damage tolerance is a more practical method of evaluation for composites than the fatigue life or fail-safe methods that were traditionally applied to metallic structures. The older regulations gave the option of either establishing a replacement fatigue life or of demonstrating fail safety, wherein a load path can be severed and the internal loads shown to redistribute throughout the remaining structure. Modern, affordable composite structures tend to be monocoque in nature, i.e., the skins are load bearing and separate frames and stringers are eliminated. Such structures, without damage, have an infinite fatigue life under normal operating loads and yet cannot be shown to be fail-safe by traditional methods of totally severing one load path. A damage tolerance evaluation concentrates on defects that may be built-in at the time of fabrication or inflicted on the structure during the service life and establishes inspection criteria that will maintain safety throughout the service life. Thus, a damage tolerance evaluation is a more practical method of dealing with the issue of safety during the service life of composite structures. The static strength and stiffness, freedom from flight flutter, lightning strike protection, and emergency landing regulations are the same regardless of the structural materials employed. Regulations that stipulate that materials and processes must be controlled by specifications and that material properties must be established to meet statistical criteria are also unaffected by choice of materials.

6.09.6

TYPICAL CERTIFICATION PLAN

There is quite a lot of flexibility possible under the FAA regulations. However, the plan described is what might be called the classical approach, the type of certification plan that has been used by Boeing, Raytheon, and others and frequently has been described in the literature (Abbott and Kolarik, 1989). A typical FAA certification program for composite structures will include the following tasks: material property coupon tests, laminate analysis, element tests, internal loads analysis, subcomponent tests (optional), and full-scale component tests, airplane ground vibration tests, flight flutter tests, certification reports, and published inspection criteria. A little terminology clarification will help at this stage. Small two-dimensional (2-D) specimens are called coupons and are usually used for material tests; larger but still 2-D specimens are called elements and often are used to represent particular laminates from a larger structure and relatively simple joints; a subcomponent is the next step up the size and complexity scale and these are used for risk reduction or investigative tests of more complex loading situations. Subcomponents are three-dimensional, may involve multidirectional loading such as internal pressure plus shear, and require larger test machines or custom designed test rigs. The final stage is full-scale component testing, for example, testing of a complete wing or fuselage. All of these test types may be employed to develop static strength data or fatigue and flaw growth data. Also, these tests may be conducted on specimens at ambient conditions or on specimens conditioned under temperature/ moisture conditions. Some full-scale components have been tested to demonstrate the load carrying capacity of heat and moisture exposed structures. However, the classical certification method involves full scale testing under ambient conditions only and then using analysis supported by element test results to certify critical structures under the worst environments expected in-service. When finite element analyses are available, the fullscale static test becomes simply a vehicle to verify the calculated internal loads by comparison of predicted strains and deflections with the measured values. The internal load applied on any particular critical laminate can then be compared to the maximum allowable load for that part derived from the laminate analysis. Given the applied loads on all critical laminates and the allowable loads from laminate analyses, a margin of safety can be calculated for each critical laminate. A flow chart for this type of certification program is shown in Figure 8.

Damage Tolerance Evaluation

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Figure 8 Certification flow chart for composite structure.

6.09.7

DAMAGE TOLERANCE EVALUATION

Even though regulations may contain different wording, the intent of damage tolerance evaluation is the same regardless of the type of airplane. In general terms, the intent is to ensure long term safety, considering manufacturing quality intrinsic to the processes used, and recognizing that certain damage may occur during service. A certifying agency may express this, in essence, as ªDon't tell me how good the good parts are, tell me how bad the bad ones are, and then tell me how to keep them safe throughout the service life.º

6.09.7.1

Damage Scenarios

Three different damage scenarios normally will be considered. Scenario 1 concerns initial quality, which includes small defects intrinsic to the manufacturing process and acceptable according to the factory inspection standards. Scenario 1 represents the as-manufactured state and therefore the structure must be capable of meeting all requirements in terms of strength, stiffness, safety, and longevity (Kassapoglou and Hammer, 1990). Scenario 2 covers damage that may be inflicted during assembly or in-service. Damage from scenario 2 must exhibit predictable growth, or no growth, during a period of in-service loading (usually expressed in number of inspection intervals) and must be detectable by the specified in-service inspection methods. In addition, the residual strength of the structure with such damage must always be at least equal to the applicable residual strength requirements (typically equal to limit load). Scenario 3 includes damage from discrete sources. Damage resulting under scenario 3 will be obvious to the crew during a flight or will be detected during a

preflight inspection. Therefore, specific residual strength criteria are applied which are concerned with safely completing a single flight; these are sometimes referred to as ªget home loads.º It simplifies the evaluation to first recognize the generic scenarios and potential damage sources, and from those identify the possible damage modes and the desired structural response. From the above definitions it is not difficult to build a matrix such as that shown in Table 1. The damage modes from scenario 1 are typically not a significant problem from the load-capability point of view. However, the potential modes from intrinsic manufacturing quality must be identified and controlled by the manufacturing specifications and factory inspection criteria. Given this, it is usually easy to demonstrate that these small imperfections will not grow under cyclic loads typical of commercial airplane service. Scenario 3 damage is at the opposite end of the scale: these modes of damage are easily detectable and will need attention before further flight (except maybe for an authorized ferry flight to a repair facility). Therefore, inspection and longevity under cyclic loads are nonissues. The scenario that creates the most need for investigation is scenario 2; and a typical test program is described in the following section.

6.09.7.2

Damage Tolerance Tests

Damage tolerance tests either under static loads or cyclic loads for flaw growth data may be conducted using element tests, subcomponent tests, and/or full-scale component tests. Element testing is recommended to evaluate composite structural performance for the various damage modes. It is possible to conduct these evaluations on the full scale test articles, but this is a risky approach and the results will

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Composites in General Aviation Table 1 Damage source and modes. Scenario 1

Source Manufacturing process

Scenario 2 Modes

Small imperfections allowed by the inspection acceptance criteria: porosity, voids, disbonds

Source

Modes

Scenario 3 Source

Modes

Tools baggage hail gravel

Resin cracking, delamination, puncture, crushed core

Severe lightning

Plies burned, puncture

Lightning

Resin burn, delamination, loose rivets Core damage

Bird strike

Delamination growth, disbond growth Resin burn

Engine fire

Delamination, core crush, puncture Puncture, severed load paths Resin burn, delamination

Water intrusion Cyclic loading Engine hot air leakage

be too late to guide the design to a minimum weight and cost configuration. Static testing of elements to validate tolerance to the damage modes in scenario 2 will include impact tests without puncture, puncture tests of detectable size and larger, water ingression tests with freeze/thaw cycles, and lightning strike tests. Strength testing will be performed for the failure modes shown to be critical, based on the finite element analysis. A significant number of undamaged specimens may be tested in order to generate statistical data. This may be desirable in order to establish that unacceptable variability is not introduced by a particular manufacturing process. Composites in the undamaged state are insensitive to fatigue loading at load levels representative of civil airplane in-service loads. Therefore, it is best to concentrate on flaw growth testing, i.e., cyclic testing of deliberately damaged specimens. Even then, specimens must be cycled at high percentages of their static failure load (higher than 50%) in order to generate measurable flaw growth. Cyclic loading at multiple stress levels is desirable to establish the sensitivity of flaw growth to cyclic stress level. An increased number of specimens may be tested at one selected stress level to identify variability in flaw growth. Some specimens may be tested under spectrum loading representing the varying amplitude loads occurring in a service lifetime. There is today no industry-wide acceptance of analytical methods predicting flaw growth rates in composites. In

Rotor burst

Ground equipment collision

Puncture

other words, there is no universally accepted equivalent in composites to the fracture mechanics methods used in metal structures analyses; thus the need for flaw growth testing. Honeycomb construction has a particular advantage in maintaining the residual strength of a pressure vessel wall after incurring large size damage from sources such as those described in scenario 3. This is due to the honeycomb shell stiffness imparting great resistance to crack bulging that in thin skin structures is a source of high crack extension forces. Tests to validate residual strength in the presence of large puncture damage may be conducted on test cylinders or on element specimens loaded to simulate internal pressure loading. These days, certification of major load carrying structure requires that full-scale components such as wings, fuselages, and tail structures are tested through a sequence of loads representing at least two lifetimes of expected mission loads (FAA, 1998). Each lifetime will consist of thousands of load cycles, including wing lift, fuselage reactions, tail loads, pressure cycles, and landing loads. The extent of manufacturing defects acceptable to the factory inspection criteria may be built deliberately into the test articles or may be simulated by mechanically inflicting damage such as delaminations, disbonds, and disrupted fiber load paths. Damage will also be inflicted mechanically in the structure to simulate inservice damage. These situations include lighting strike, hail damage, runway damage, and

Damage Tolerance Test Results

Figure 9

Virgin and comparative residual strengthsÐhoop tension.

tool impacts. These damage modes will be tested through as much as two lifetimes of cyclic loading to prove that the full-scale structures are, in fact, damage tolerant and that damage will not grow in an unpredictable manner and will always be detected by the inspection procedures to be used in-service. Larger damage may be inflicted later in the full-scale cyclic testing program to simulate impacts with ground service equipment, impacts with hangar doors and other aircraft (hangar rash) and poor maintenance practices; all bad things that occasionally happen in the loading, handling, and maintenance of civil airplanes. The larger damage modes will normally be detected before the next flight and so the demonstration for these modes may consist of a relatively few flights of cyclic loading and inclusion in the residual strength tests. After completion of cyclic and flaw growth testing, the major components, wing, fuselage, and tail, will be subjected to load tests to verify that, in spite of all the load cycles and inflicted damage, the remaining structure will still carry the required residual strength loads. These are typically limit condition flight loads and or pressure loads expected to be encountered during the service life of the aircraft, except for the larger damage modes that are associated with the ªget homeº reduced loads criteria. 6.09.8

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DAMAGE TOLERANCE TEST RESULTS

Selected examples of element test results are discussed in this section and illustrated in the accompanying figures. The results shown were

obtained from samples representing fuselage shell construction on a corporate type airplane. However, scale matters and different results could be obtained from tests on samples representing larger airplanes because of greater laminate thicknesses required to carry the basic pressure and bending loads. In the case of hoop tension loading from internal pressure, designing for damage tolerance will not impose a serious weight penalty. In the case of cabin pressure an additional factor of safety required in the regulations means that ultimate pressure is twice the normal operating pressure. The ultimate pressure must be carried with the undamaged panel; so, just a little additional material will enable the panel to meet the required residual strength load with large puncture damage. The residual strength required for the pressure case is about 60% of the ultimate pressure (see Figure 9). In the case of longitudinal tension loading from fuselage bending, the residual strength requirement is limit load, i.e., about 67% of the ultimate load. To carry that load with a large puncture, a small amount of material must be added. In both cases a more robust structure would result if ultimate loads were to be carried with impact damaged panels. This may be required by the regulations unless impact damage is readily detectable with the inspection methods proposed. A similar situation exists in the case of compression loading from fuselage bending. Designing to the residual strength requirement with large puncture damage would imply using approximately 85% of the allowable virgin strength (see Figure 10). But if an impactdamaged panel is to be good for ultimate load

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Composites in General Aviation

Figure 10 Virgin and comparative residual strengthsÐcompression.

Figure 11 Virgin and comparative residual strengthsÐshear.

then only 65% of the virgin strength can be used. This may not be a serious penalty as the maximum compression loads occur in the fuselage bottom from down-bending load cases and the lower fuselage is usually reinforced by cargo or passenger floor structure. Maximum shear loading occurs along the side of the fuselage. Again, designing to carry limit load (67% of ultimate load) with large puncture damage is a slight weight penalty, approximately 82% of the maximum virgin strength can be used. In this case, 82% of virgin strength will also accommodate impact damage at ultimate load (see Figure 11). Figure 12 shows an assemblage of test results that illustrate the trend from no damage to massive damage in a pressure test cylinder

with CFRE honeycomb walls. Massive damage is the type of wall puncture that could only occur from a serious collision with ground support equipment such as steps, generator carts, refueling equipment, baggage handling equipment, and so on. As mentioned previously, this type of damage will be obvious and should be detected before flight. However, just in case, tests have been conducted to determine cylinder wall residual strength with massive damage. The trend revealed has proven typical of test results from big and small airplane honeycomb pressure walls: with larger and larger damage, a residual strength threshold becomes apparent. This confirms the benefit of honeycomb structure in resisting skin crack bulging and containing damage growth.

Damage Tolerance Test Results

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Figure 12 Honeycomb pressure shell residual strength.

Figure 13 Constant amplitude compression flow-growth test results.

Flaw growth test results for large puncture damage under constant amplitude cyclic compression loading are shown in Figure 13. These data conveniently are interpreted using a bestfit straight line in a log±log format. This type of plot can then be used to assess important damage tolerance characteristics of the laminates. Composites typically show a higher

flaw growth threshold than aluminum. This makes the flaw growth life vs. stress line much shallower than that for aluminum. This implies a greater scatter in the life axis, i.e., small changes in cyclic stress level will cause relatively large changes in flaw growth life. The significance of scatter in flaw-growth life can be assessed by estimating and plotting a B-Basis

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Composites in General Aviation

(90% probability of survival with 95% confidence) stress life-line assumed to be parallel to the mean-life plot. A prime use of these data is to structure the full-scale cyclic test to run in a rational, yet economical manner. The scatter in flaw growth life may be accounted for by application of a load enhancement factor which increases the applied loads and therefore reduces the number of test lifetimes needed to achieve a B-Basis relationship between test lives and service lives. The flaw growth threshold may be determined by extrapolation of the B-Basis life-line out to, say, 10 million cycles. All load cycles that induce stresses below the flaw growth threshold may be eliminated from the fullscale test load spectrum.

6.09.9

INSPECTIONS

Based on interpretation of the available test and analysis results, inspection criteria will be developed to maintain safety and economy throughout the planned service life of the airplane. The inspection requirements will be published in the airplane Maintenance Manual or the Structural Inspection and Repair manual (SRIM). Airplanes are subject to routine structural inspections. These start with a ªwalk-aroundº inspection performed before every flight, and progress through 200 h and annual inspections. Special structural inspections are introduced at a starting threshold and continue at a subsequent repeating interval throughout the service life. These inspections will be integrated with the routine inspections whenever possible, but in some cases, the special structural inspections may require some structural disassembly in order to provide adequate access. The published inspection procedures will include inspection methods, threshold time at which to start composite damage detection, and subsequent frequency of inspections. A factor is usually applied so that allowance is made for the damage to exist over several inspection intervals, depending on the criticality of the structure. So far, most airplane projects using composites in primary structures have been able to rely on visual inspection as the primary damage detection method. This is partly a benefit of extensive testing which has been performed to confirm the damage tolerance capability of composites. Another factor is probably the relatively short time that primary structural composites have been in service. As service experience builds up, more sophisticated inspection methods may become economically available and be may

become necessary, as shown by the experience with aluminum structures.

6.09.10

STRUCTURAL REPAIR MANUAL

The introduction of composite parts into civil aviation service without planned support in the way of allowable damage criteria, repair procedures, repair materials, repair kits, and availability of support personnel is an invitation to have severely disgruntled customers who will surely besmirch the reputation of composites. At the time of introduction of structural parts into airplane customer service, decisions were made by different manufacturers to either provide the fullest possible information and training on composite repair or to restrict the information on materials and construction and in essence say ªCall your local dealer.º The best choice is for the structural repair manual (SRM) to show areas in which standard field repairs are permissible, specify materials, procedures, and available repair kits, and allow cosmetic repair of minor damage. Certain structures will be restricted to factory supervised repair; these tend to be areas which are heavily reinforced, such as spars, bulkheads, windshield frames, and doorframes. Supervised repairs also may be required if the magnitude of damage exceeds the limits for standard field repair specified in the SRM. A typical decision flow diagram for field repairs is shown in Figure 14.

6.09.11

SERVICE EXPERIENCE

The service experience with composite structures in general aviation has been excellent. Starships have been flying since the late 1980s and no problems with major structure have been encountered. Control surfaces and pressure bulkheads on the Gulfstream GV have also encountered no major service problems. Composite stabilizer structures have been in use on Beech 1900 commuters flying typically 2500 h per year since the mid-1980s; and, again, no problems related to the composites have been reported.

6.09.12

CRASHWORTHINESS

Safety in the event of an emergency landing has proven to be outstanding. A nose landing gear collapsed during a landing of one of the Starship test airplanes; the airplane was flown home, repaired, and was returned to service in

Conclusions

15

Figure 14 Decision flow chart for field repair.

10 days. The repairs were made by procuring blank parts from the factory, cleaning out the damaged areas, and splicing replacement sections into place by bonding and riveting. An even more spectacular event in occurred in Denmark in February 1994. Starship number 35 ran off the runway into a snow bank at approximately 130 mph. The right hand main gear collapsed, the other main gear and the nose gear were sheared off (not torn out, but the aluminum forgings severed) from the force of hitting the snow bank at high speed. The right hand wing was dragged along the ground and, as a result, suffered damage to the flaps, vertical stabilizer, and rudder. The nose section was damaged by the nose gear being severed and forced upward into the structure. The cabin underbelly was crushed through skidding along without the landing gear. Crew and passengers were, of course, well shaken, but were otherwise unhurt. No fuel was spilled, no seats came loose, no windshield or window glass was broken, or even cracked, and the cabin was undistorted enabling the cabin door to open normally. The crew and passengers unbuckled their seat belts and walked away. A team was sent to survey the damage and list the replacement parts needed. Later the airplane was repaired on-site by a crew of five technicians plus one engineer, one inspector, and one service manager. Some parts with localized damage were repaired using techniques published in the SRM which allows damage to be repaired on-site by trained service staff. For more extensive damage, blank parts were delivered from the factory and were used as stock from which to cut replacement panels that were then bonded and/or mechanically fastened into

place. Aircraft systems such as landing gear, propellers, hydraulics, antennae, etc., were replaced with factory parts. The repairs were finished and the airplane rolled out for flight test in July 1994, much to the surprise of the insurance company and the Danish aviation authorities who were convinced that a metal airplane would have totaled by such an incident. 6.09.13

CONCLUSIONS

Modern manufacturing methods enable the fabrication of composite primary load carrying structures for commercial aircraft use which are low cost as well as low weight. These structures require a damage tolerance evaluation for certification to Part 23 or Part 25 of the FAA regulations. Composite structures can be designed to tolerate large puncture damage with little weight penalty. It may not be required by regulation to carry ultimate loads after impact damage (it depends on the inspections specified). However, a more robust and more economical to operate product will result when composite structures are designed to carry ultimate load with impact damage. A rational damage scenario and supporting element test program will assist considerably the damage tolerance evaluation. Composite structures are relatively insensitive to fatigue loading, and a high (compared to aluminum) flaw-growth threshold may be defined from flaw-growth test results. Scatter in flaw-growth life should be examined in order to establish a load-enhancement factor for use in the full-scale cyclic tests. The application of a

16

Composites in General Aviation

flaw-growth threshold and a load-enhancement factor will enable full-scale fatigue tests to be conducted more economically than on equivalent metal structures. With a combination of careful design, rational testing, and advanced manufacturing techniques, we can expect to see continuing growth in applications of composites in civil aviation primary structures. 6.09.14

REFERENCES

R. Abbott and A. Kolarik, in `SAMPE 34th International Symposium and Exhibition', May 1989, eds. G. A. Zakrzewski, D. Mazenko, S. T. Peters and C. D. Dean, SAMPE, Covina, CA, 1989, vol. 34, pp. 283± 289. 14 CFR, Federal Aviation Regulation 25.571, Amendment 25-86, February, 1996a.

Copyright # 2000 Elsevier Science Ltd. All rights reserved. No part of this publication may be reproduced, stored in any retrieval system or transmitted in any form or by any means: electronic, electrostatic, magnetic tape, mechanical, photocopying, recording or otherwise, without permission in writing from the publishers.

14 CFR, Federal Aviation Regulation 23.573, Amendment 23-48, February, 1996b. FAA Advisory Circular AC-107A, April, 1984. FAA Advisory Circular AC 25.571-1C, April, 1998. C. Kassapoglou, Comp. Struct. Int. J., 1988, 9, 139±159. C. Kassapoglou, J. Comp. Tech. Res., 1996, 18(4), 274± 284. C. Kassapoglou and R. Abbott, in `The 29th SDM Conference', Williamsburg, VA, 1988. C. Kassapoglou and J. Hammer, J. Am. Helicopter Soc., 1990, 35(4), 46±52. C. Kassapoglou et al., J. Comp. Tech. Res., 1988, 10, 65± 73. C. T. Sun, B. J. Quinn, J. Tao and D. W. Oplinger, `Comparative Evaluation of Failure Analysis Methods for Composite Laminates', DOT/FAA/AR-95/109, May, 1996. Suppliers of Advanced Composite Materials Association, Recommended Test Method, SRM 2R-94, `Compression after Impact Properties of Oriented Fiber-Resin Composites', 1994.

Comprehensive Composite Materials ISBN (set): 0-08 0429939 Volume 6; (ISBN: 0-080437249); pp. 165±180