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DOI: 10.1177/1475921719831370

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Real world application of angular scanpulse-echo ultrasonic propagationimager for damage toleranceevaluation of full-scale compositefuselage

WJ Lee1, BH Seo1, SC Hong2, MS Won1 and JR Lee2

AbstractComposite structures are assertively used for new airframe designs and manufacturing in military aircrafts because ofsuperior strength-to-weight ratios and fatigue resistance. Because the composites have different fatigue failure character-istics compared with metals, it is necessary to develop different approaches for the composite fatigue design and testing.In this study, we propose an in situ damage evaluation technology with high spatial resolution during full-scale fatiguetesting of composite aircraft structures. For real composite structure development considering composite fatigue char-acteristics, full-scale fatigue and damage tolerance tests of the composite fuselage structure were conducted to evaluatethe structural characteristics. In the meantime, the laser ultrasonic nondestructive inspection method, called an angularscan pulse-echo ultrasonic propagation imager, which is fully noncontact, real-time, and portable to position it inbetween the complex test rigs, is used to observe in situ damage growth of the composite. Finally, the verification proce-dure assisted by the angular scan pulse-echo ultrasonic propagation imager assures no growth of the initial impact dam-ages after lifetime operation and proves the damage tolerance capability of the developed composite fuselage structure.

KeywordsFull-scale fatigue test, pulse-echo ultrasonic propagation imager, composite fuselage structure, impact damage, damagetolerance

Introduction

Fiber-reinforced composite materials are considered thecore technology for weight reduction of aircraft struc-tures. There are many kinds of aircraft structures madeof composite materials with the help of tailoring andoptimizing the composite structural characteristics. Oneof the main advantages of the composite material com-pared with the metal is its high-specific strength andstiffness. In addition, the direction of reinforced fiberscan be designed considering the structural purpose forthe optimization. The material density of carbon–epoxycomposite (1–2 g/cm3) is much lower than that of tita-nium (4–5 g/cm3), aluminum (2–3 g/cm3), and steel (7–8 g/cm3), which makes it possible to reduce the aircraftstructural weight approximately 30%. These compo-sites have large elastic range with less plastic deforma-tion and show a low strain level.1,2

A disadvantage of a composite is its inhomogeneity.The fiber has good tensile properties but poor compres-sive properties. These composite structures have manyfabrication variables as well as the design variables,which means the exact material properties can be

1The 7th R&D Institute, Agency for Defense Development, Daejeon,

Republic of Korea2Department of Aerospace Engineering, Korea Advanced Institute of

Science and Technology, Daejeon, Republic of Korea

Corresponding authors:

WJ Lee, The 7th R&D Institute, Agency for Defense Development,

Yuseong-gu, PO Box 35, Daejeon 34188, Republic of Korea.

Email: lwj@add.re.kr

JR Lee, Department of Aerospace Engineering, Korea Advanced Institute

of Science and Technology, 291 Daehak-ro, Yuseong-gu, Daejeon 34141,

Republic of Korea.

Email: leejrr@kaist.ac.kr

evaluated when the composite structure is fabricated asinitially designed. It takes a long time to define thematerial data of composites widely, especially forthe fatigue and damage tolerance data. In addition, thecomposites have high stiffness with low ductility, andthe impact energy generates atypical damages andcracks in the composite. Studies on the fracturemechanism made by impact damage and cracking havebeen conducted based on the micro/macro crack initia-tion and crack growth in the fiber and matrix of thecomposite.3–8

However, because composites are inhomogeneous,and ply-stacked structures in general, it is common tohave in-plane delamination and internal damage, whichare difficult to detect and define as the standardizeddamage. As a result, statistical damage tolerance andprobability analysis have been required. The electricalconductivity of the composite is lower than that ofmetal, it is susceptible to the lightning, and the addi-tional copper mesh on the skin is required in specificregions as a result. These complicated fabrication pro-cesses raise the manufacturing cost and quality devia-tion, which require more-precise process control andauthentication.

For the application of composites in aircraft struc-tures, fatigue characteristics should be considered,along with uncertainty. The S–N curve of composites isflat compared with that of metals, which is good for thefatigue performance. However, the composite shows alarge deviation and scatter factor on the fatigue charac-teristics, and it is also susceptible to humidity and tem-perature. As a result, the design of composite structuresrequires enough margin of the safety factor to meet thefatigue life requirement. However, the operating air-craft structures are exposed to impact damage, so theyshould also maintain the structural integrity with thedamage caused by unpredictable impacts. As a result,the damage tolerance design concept should be appliedon the composite aircraft structure and the damageshould not grow during the designed lifetime. Finally, afull-scale aircraft structural fatigue and damage toler-ance test should be conducted to ensure that the dam-ages do not grow to meet the structural integrity afterwhole lifetime of usage.9,10

The impact damage in the composite is hard todetect and define, and it is difficult to predict whether itwill grow or not. In addition, it is hard to define theeffect of the damage on the total fatigue life of the air-craft. To ensure the structural integrity and fatigue lifewith the assumption of damage no-growth, the designlimit load test (DLLT) and design ultimate load test(DULT) are conducted on the full-scale specimens ingeneral. However, this evaluation logic does not includethe specific analysis of the real damage behavior with adamaged area. As a result, the tested structure can be

certified to have structural integrity, even though therelations between the design safety margin and the localdamages are not defined in detail.

In this study, a fatigue and damage tolerance testwas conducted on the full-scale composite fuselagestructure, which includes barely visible impact damages(BVIDs), visible impact damages (VIDs), and largeimpact damages (LIDs). The initial damages aredefined and evaluated by angular scan pulse-echo ultra-sonic propagation imager (PE UPI) to check thegrowth before and after the fatigue life cycle test. Theangular scan PE UPI enables in situ nondestructiveevaluation (NDE) during the full-scale fatigue test ofwhich complicated test rigs does not allow access ofconventional contact ultrasonic testing devices. Thisprocedure can define the damage to grow or not indetail before the limit load and ultimate load test. Thisdamage evaluation process shows a new possibility topredict the structural integrity of the fatigue life anddamage tolerance, with the real-time and visualizeddamage evaluation method with high spatial resolution.

Damage tolerance design process forcomposite structures

Various material property tests evaluate the reliabilityof composite structures after fabrication, even thoughthe composites have various advantages. The fabricatedcomposite structure may include microscale materialdefects with macroscale manufacturing process errorsuch as pore, microcrack, delamination, scratch, impactdamage. These defects can lead to the initial crack andcrack growth in the specific structural load conditionas well as the degradation of composite properties. Thecrack generated by the defects can grow under the spe-cific fatigue load to cause sudden fatigue failure of thestructure. These defects should be evaluated not togrow under the operational load condition by nondes-tructive inspection (NDI) process, which is necessary touse the material in aircraft structures.

A fatigue test on the composite specimens, whichhad representative initial damages, was conducted toidentify the severe damage case of composite fatiguelife. The damages were an open-hole, delamination,scratch, and compression after impact (CAI), and thefatigue tests were conducted under the cyclic load in ahigh operating temperature of 55�C and extremehumidity of 95% conditions. The open-hole specimenshowed the smallest life cycle among the damaged com-posite specimens. After the sample fatigue test, thedamage tolerance design margin could be evaluatedbased on the static failure tests of the open-hole dam-aged specimens. As a result, the aircraft structure wasinitially designed and sized based on the open-hole

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damage tolerance design criteria, and the full-scale air-craft fatigue test should be performed to confirm thestructural integrity and the fatigue life at the end of thedevelopment process.

The aircraft structure consisted of solid laminatesand sandwich composites made by an autoclave curingprocess, and artificial BVIDs were made on the low sta-tic safety margin area. A specially designed impactorwas used to make the impact damages on the surface ofthe composite structures by considering the impactenergy, structural shape of impact locations, and struc-tural boundary conditions to meet the ‘‘barely visible’’requirement. The impact damages on the surface of thecomposite structures are shown in Figure 1 with testconditions in Table 1. The depth of the dent and visibi-lity was determined on each damage, as shown inFigure 1.

Full-scale fatigue test

The fuselage was constructed by uni-directional andfabric carbon/epoxy laminate and honeycomb sand-wich. All the skins and main structures used the compo-site materials, and a small amount of metals was usedon the engine mounting structure (EMS), hinge fittings,lugs, landing gear supporting structures, and some partof the main bulkhead. The load enhancement factor(LEF) for this composite–metal hybrid structure was1.127 for the full-scale fatigue test. The initial BVIDs

were generated during the component fabrication andstructural assembly process of the full-scale fatigue testspecimen.

A full-scale fatigue test for the composite structurewas conducted according to the following process.First, the fatigue test spectrum was derived based onthe initial mission profile. The composite structure wasfabricated including the initial BVIDs, and the two-lifefatigue load was applied onto the structure.

As shown in Figure 2, after the two-life fatigue test,the VIDs were made by hammer striking on the low-marginal areas of the laminate and sandwich skins, andthen the one-life fatigue test was continued. These dam-ages were generated after a preliminary impact test onthe equivalent specimens in which the proper impactenergy level up to 50 J could be obtained. Because thecomposites have high stiffness and elastic properties,the scale of the dent on the skin was not linear to theapplied energy. It was hard to make a proper dent onthe surface just before the critical energy level, but therewas large damage inside and outside when the energylevel was just over the critical level. In addition, the dif-ferent structural boundary conditions at each pointmade different damage levels, even though the skin hadthe same thickness. As a result, different variables wereset up to make proper VIDs for the full-scale damage-tolerant test.

For a honeycomb sandwich skin, the impact energygenerally makes outer surface damage severe. In

Figure 1. BVIDs on the composite components before aircraft assembly: (a) wing spar flange: impact energy (20.0 J)/dent depth(0.15 mm) and (b) tail wing upper skin: impact energy (3.0 J)/dent depth (0.07 mm).

Table 1. Conditions for impact-induced damage generation.

Impact type Falling weight impact includes single cylinder tubes through which a cylindrical impactor travelsImpactor mass 2 kg, 3 kg, 5.5 kg 6 0.25 kgImpactor tip (ASTM D7136) Hemispherical striker tipTip diameter = 16.0 6 0.1 mm (0.625 in)Hardness: 60–62 HRC

Lee et al. 3

addition, a honeycomb core dent and facesheet delami-nation were generated at the same time. The shape ofthe outer surface dent is clear enough to define as VID.These impact damages should not have an effect on thestructural integrity, and the damage should not growafter the additional one-life fatigue test. Figure 3 showsthe VID on the sandwich skin of the structure.

However, it is also difficult to define the relationsbetween the impact energy and the visual damage levelon the solid laminate composite. In some cases, highimpact above the designed energy level could not gener-ate even BVID on the outer surface of the solid lami-nate, but there was large and severe damage with fiberbreakage on the inner surface. Figure 4(a) shows theouter surface of the impact point, and the dent was

under the BVID level. In addition, Figure 4(b) repre-sents the inner surface with a large damage area, includ-ing fiber breakage. Even though this impact energy wasenough to cause severe damage to the laminated skinstructure, it was hard to detect from the outer side witha visual inspection method. As a result, this type ofhigh impact energy damage with barely visible charac-teristics should be contained in the damage tolerancetest as the most critical undetectable damage type. Thecomposite structure should be designed to withstandthis type of damage, and the structural integrity with adamage no-growth requirement should also be con-firmed. After the implant of the initial damages on thecomposite structure, the fatigue life with damage toler-ance test was conducted.

After the one-life test, LIDs were applied on the min-imum structural margin area. The fatigue cycle wasconducted for three sorties as the minimum period fordamage detection, and the VIDs and LIDs were totallyrepaired. Finally, an ultimate load test was conductedto confirm the structural integrity of the aircraft struc-ture, which had finished the fatigue life cycles.

Damage evaluation

There are not many in situ NDE technologies that canquantitatively assess the growth of composite impactdamage during the full-scale fatigue test. The typicalmanual A-scan was not able to be applied becausecomplicated fixtures, loading and sensing devices didnot allow access of contact ultrasonic testing devices.

Apply Initial Flaws/Damage (BVID, Delamination, Debonding, Open-hole) Apply VID (Visible Impact Damage)

Apply LID (Large Impact Damage)

1 2

(1) (2) (3) (5) (6) (7) (8) (9) (10)

Apply Ini�al Flaw/Damage (BVID*, Delamina�on, Debonding, Open-hole) Apply VID*

1 2

(4)

3

Apply LID*

4

Repair VID&LID

(11) (12) (13)

3

4

Figure 2. Aircraft structure fatigue/damage tolerance test procedure with Category I, II, and III damages.

Figure 3. VID on the honeycomb sandwich skin.

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In addition, the inspection results are highly dependenton material properties and thickness as well as inspec-tors’ skill. Sometimes, probe contact and couplantapplication are not allowed. Recently, scanning laserultrasonic methods and ultrasonic propagation imagingtechnologies have shown high feasibility for an in situand remote NDE tool during the full-scale structuretests.11–13 The laser ultrasonic propagation imager(UPI) is classified into guided wave UPI and bulk-waveultrasonic propagation imager (B-UPI), depending onthe laser ultrasonic modes used. The B-UPI is classifiedagain into pulse-echo and through-transmission sys-tems. In the PE UPI, the ultrasonic generation and sen-sing laser beams are coincident at one point on thetarget. The generated ultrasound will be propagatedover the thickness and then reflected. The reflectedwave is captured by another sensing laser. The sensingmechanism at one point is repeated as the laser beam ismoving during the scanning of the target surface. Thetypical scanning mechanisms are linear and angular

scan modes. Linear scan is realized by an X-Y lineartranslation stage and angular scan is done by a lasermirror scanner. The laser mirror scanner (LMS) makespossible a rapid raster scan based on angular motionsof the two laser mirrors mounted on the galvano-motors in the scanner. Figure 5(a) shows the angularscan PE UPI system, which includes a Q-switched laseras the generation laser and a laser heterodyne interfe-rometer as the sensing laser, and beam combinationoptics and a galvano-motorized mirror scanner. Thesystem body includes controllers for the two lasers andthe LMS, as well as a data acquisition board. They areall synchronized for rapid scanning measurement up toa pulse repetition rate of 5 kHz. The system body alsoincludes a graphical user interface (GUI) platform forautomatic inspection, damage visualization in real time,and post image processing. The basic inspection resultis PE ultrasonic wave propagation imaging (UWPI)video, which is generated right after the scanning inreal time.15 In addition, various time window

Figure 4. VID on the solid laminate skin: (a) outer damage and (b) inner damage.

Figure 5. (a) Configuration of angular scan pulse-echo ultrasonic propagation imager14 and (b) implementation of system for in situNDE of a fuselage section.

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amplitude mapping (VTWAM) can be applied as postprocessing to enhance damage visualization.16 Figure5(b) shows how the system implemented for in situNDE of the full-scale composite fuselage during thedamage tolerance test.

In situ NDE results

Damage evaluation was conducted on the BVID andVID area in the solid laminate and sandwich skin forthe fuselage. Impact damages were visualized using PEUWPI freeze frame and VTWAM, which are damagevisualization techniques in angular scan PE UPI.Figure 6 shows VID on the honeycomb sandwich skinstructure. The circular dent shows that the sandwichouter skin was damaged with fiber breakage. The scan-ning area of the combined sensing and excitation laserbeam was 100 mm 3 100 mm, and the standoff dis-tance between the fuselage skin and the UPI scanninglaser head was 1.86 m, and fully noncontact measure-ment was conducted. Figure 5 shows typical inspectionconfiguration and it should be noted that the line ofsight is very limited until the target to be inspected byfixtures, and loading and strain sensing equipment. Thesampling frequency of the ultrasonic wave was10 MHz, and the scanning interval was 0.25 mm. Thepulse-echo ultrasonic wave propagation video wasobtained in real time after scanning. The total scanningtime for the area of the interest of 100 mm 3 100 mmwas 32 s under a 5-kHz pulse repetition rate.

Figure 6 shows the scan area and the visualizationresult of the impact area before the fatigue test. The cir-cular dark region shows the impact damage area, andthe white lines show fiber breakage on the overlappedimage. Figure 7 shows a comparison of the damage

before the fatigue life test and after the test where straingauges are also visible but not for the NDI but for thefatigue test. The impact damage did not show visiblegrowth detectible by the angular scan PE UPI with aspatial resolution of 0.25 mm during the life cycle test.

Figure 8 shows impact damage on the solid laminateskin. The outer skin shows a very small dent under theBVID level, but the visualized figure shows a largedamage area in the laminate skin. The shape and direc-tion of the damaged area were related to the fiber direc-tion of the inner ply, which was a break. The scanningimage shows the damage on the inner surface of theskin through the laminate. Figure 9 shows that therewas no distinguishable growth of damage during thefatigue life test.

For the test of combined damage characteristics,large impact damage (LID) was made near the BVIDpoint on the honeycomb sandwich skin. The damagewas evaluated before and after design limit load test(DLLT), and the results are shown in Figure 10.

Figure 10(a) shows initial BVID made on the skinstructure. The impact energy caused outer skin damagewith a core dent, and the skin was delaminated. Thediagonal white band is the overlapped region of skinply, and the dimples are hexagonal core. Figure 10(b)shows damage combination of BVID and LID. Theadditional LID made a deep dent on the sandwich skin,and it also could affect the existing damage. This kindof damage also should be considered the severe case,because the combined damage area can generate unpre-dictable large delamination in the structure. Accordingto comparison of the scan results, as shown in Figure10(b) and (c), there was no distinguishable increment ofthe damage shape and size as a result, which meansthe combined impact damages and cracks meet the

Figure 6. NDI and visualization of the impact damage on the sandwich fuselage skin: (a) scan area and (b) inspection result.

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Figure 7. Comparison of impact damage on the sandwich skin: (a) before and (b) after the fatigue test using the inspection resultsof PE UWPI freeze frames and VTWAMs.

Figure 8. NDI and visualization of the impact damage on the solid laminate skin: (a) scan area and (b) inspection result.

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no-growth requirement after DLLT. This result con-firms that these design damages are acceptable for thefatigue life requirement of the aircraft structure. TheDULT was conducted after repairing VIDs and LIDs,and the structure met the requirement, as expected.

Damage tolerance (DT) test aims at showing nodamage growth after fatigue loading. For that, indeed,we need high spatial resolution image to make sure itbecause the damage images before and after test mustbe identical to pass the DT test. In this point of view,the results presented here showed successful demonstra-tion of the required function.

Conclusion

A full-scale composite aircraft structure was developed,and a damage tolerance test for fatigue life

requirements was conducted. Based on the fatigue lifetest of the damaged composite specimens, the damagetolerance design margin was defined initially. After thefabrication of the aircraft structure, various impactdamages were made on the composite fuselage struc-ture, and a damage tolerance test was conducted.Basically, the damage tolerance test of the compositestructure includes the DLLT and DULT after finishingfatigue cycle loading to show that the damages do notaffect the structural integrity. In this study, a new anddetailed in situ NDE method for the damage was pro-posed using the angular scan PE UPI. With the help ofthe angular scan PE UPI with real-time, noncontact,remote evaluation and small line-of-sight area require-ment, damages and cracks could be monitored duringdamage tolerance tests, and both the inside and outsideof the damaged skin could be identified with

Figure 9. Comparison of impact damage on the solid laminate skin: (a) before and (b) after the fatigue test using the inspectionresults of PE UWPI freeze frames and VTWAMs.

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Figure 10. Comparison of impact damage: (a) initial damage, (b) before, and (c) after the damage limit load test using theinspection results of PE UWPI freeze frames and VTWAMs.

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visualization. Furthermore, it was confirmed that thedamage and cracking did not grow after the fatiguespectrum cyclic loading and DLLT, directly. By tracingthe behavior of various damages in the compositestructure with visualization, more detailed damage tol-erance design, damage control, and certification ofdamage tolerance tests will be possible.

Declaration of conflicting interests

The author(s) declared no potential conflicts of interest withrespect to the research, authorship, and/or publication of thisarticle.

Funding

The author(s) received no financial support for the research,authorship, and/or publication of this article.

ORCID iD

JR Lee https://orcid.org/0000-0003-0742-4722

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