Corrosion Damage Atlas for Aircraft Corrosion Management ...

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Corrosion Damage Atlas for Aircraft Corrosion Management and Structural Integrity Assessment

Min Liao Aerospace, National Research Council Canada

1200 Montreal Road, Ottawa, Ontario, K1A 0R6

CANADA

[email protected]

ABSTRACT The National Research Council Canada (NRC) was tasked by the Department of National Defence (DND) to review the Corrosion Prevention and Control Program (CPCP) for all the Royal Canadian Air Force (RCAF) fleets. From this task, NRC and DND identified a need to accurately characterize the type of corrosion present in aircraft and then developed a corrosion damage atlas. This paper presents some results of the NRC developed corrosion damage atlas with “real-world” examples of various types of corrosion that have been found in military aircraft fleets. These images have been collected from various sources for a variety of material systems as well as from destructive and non-destructive inspections (NDI). The damage atlas is expected to allow for a more accurate tracking and assessment of specific forms of corrosion, and to improve corrosion level/severity evaluation for a more effective CPCP program. Several case studies are also presented to demonstrate the importance of the damage atlas for aircraft structural integrity assessment and research.

1.0 INTRODUCTION

As military aircraft fleets continue to age, most often beyond their original design life, they become more susceptible to the time-dependent effects of corrosion. Corrosion has significant impacts on maintenance cost and fleet availability. For example [1][2], the corrosion costs for all aviation and missiles of the United States Department of Defense are $8.97B in FY17 and $10.18B in FY18, which includes the Air Force costs of $5.325B (FY17) and $5.669B (FY18, i.e., 23.6% of total maintenance cost). In FY18, corrosion caused 89,653 non-available days (NAD), about 14.1% of total NAD for the Air Force aviation and missiles. In addition, corrosion can affect structural integrity by accelerating the time to fatigue crack development and stress corrosion cracking for those safety-of-flight structural locations [3]. Therefore a corrosion assessment is required under the Task II of the United States Air Force (USAF) Aircraft Structural Integrity Program (MIL-STD-1530).

To ensure the safe operation of aging fleets, a holistic (cradle-to-grave) corrosion management approach should be adapted based on the same fundamental need to characterize, predict, and monitor structural crack damage. This approach requires the development of accurate corrosion prediction models, reliable health monitoring systems and novel inspection methods for aircraft maintenance. However, accurately identifying the type of corrosion present in service might be difficult since this subject is not extensively studied in the educational system of most countries. To help resolve this deficiency within government defence departments, some documents containing general information on the different types of corrosion affecting an aircraft structure are produced; however, the resulting damage caused by a specific type of corrosion is very often represented by a simple sketch. This schematic does not provide a “real-world” perspective of the corrosion damage and thus does not allow for the accurate determination of the type of corrosion present within the aircraft. More physical details of corrosion can help determine the level of corrosion and then repair actions required for corrosion prevention and control plan (CPCP) for aircraft fleet management.


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To be able to cost-effectively manage corrosion within aging aircraft fleets, assure structural integrity, and then improve availability, with the support of the Department of National Defence (DND), NRC developed a corrosion damage atlas including “real-world” examples of various types of corrosion that have been found in military aircraft fleets. This paper presents some results of the damage atlas, as well as some applications of the damage atlas for aircraft structural integrity assessment and research.

2.0 CPCP REVIEW AND ASSESSMENT

From 2009 to 2011, NRC was tasked by DND to carry out CPCP review and assessment for all the Royal Canadian Air Force (RCAF) air fleets, in order to improve RCAF’s CPCP program. The activities included reviewing the CPCP documents for all fleets to identify common issues/concerns and revising the Canadian Forces Technical Order (CFTO) C-12-010-040-TR-021 that is governing fleet wide corrosion prevention plan. The effort was also expected to provide improved CPCP guidance to DND ASIP managers, support DND contractors for reduction of life-cycle cost, and improve availability of the RACF fleets.

A team of NRC researchers first reviewed various CFTOs for four fixed wing aircraft (CC130, CP140, CF188, CC177), two fixed wing lighter aircraft (CC115, CT142/Boeing Dash-6/7/8) and one rotary wing aircraft (CH124). Then a focused review was conducted on a general document C-12-010-040-TR-021[4], i.e., Aircraft Cleaning and Corrosion Control Exterior and Interior, 1997-04-30, Ch/Mod 5-2008-01-16 (referred as TR021-97/08 in this paper), which aims at all RCAF fleets especially a few fleets which do not have specific corrosion documents. The C-12-010-040-TR-021 is equivalent to the USAF Technical Manual, TO-1-1-691, i.e., Cleaning and Corrosion Prevention and Control Aerospace and Non-Aerospace Equipment. During the review, NRC organized a CPCP working group meeting including DND ASIP managers, the Quality Engineering Test Establishment (QETE), and the Royal Military College of Canada (RMC), in which each fleet/party presented the status of their corrosion program, and major issues/gaps related to corrosion. From the CPCP working group meeting and follow-on discussions, NRC summarized a number of findings and recommendations including

• revise/add corrosion definitions in TR021-97/08[4]; provide more realistic corrosion examples/photos (corrosion damage atlas) for the CFTOs, especially TR021-97/08;

• update the outdated (1990s) list for corrosion prevention compounds (CPCs)/corrosion inhibiting compounds (CICs) in the CFTOs following the common MIL standards;

• reassess three levels of corrosion (light, moderate, severe) associated with the non-destructive inspection (NDI) capability and their impacts on structural integrity and fleet management;

• develop research projects to study topics such as corrosion growth rate for corrosion prognostics and effect of CPC/CIC on new materials used in the new aircraft structures (ex. aluminium 7249 in the new CP140 wings).

In addition, DND requested to add engine hot corrosion to the scope of the NRC project. Among the above items, it was strongly recommended that a corrosion damage atlas be developed, which would include images of different types of corrosion and their severity levels. In 2015, NRC developed an aircraft corrosion damage atlas [5], including additional inputs from The Technical Cooperation Program (TTCP) nations. In 2016, DND issued an updated C-12-010-040-TR-021 [6] (herein referred as TR021-2016), which incorporated some of results and recommendations from the NRC CPCP assessment project.



3.0 NRC CORROSION DAMAGE ATLAS

A complete corrosion damage atlas is presented in the NRC report [5]. In the report, corrosion is defined as a chemical or electrochemical reaction between a material, usually a metal, and its environment, that produces a deterioration of the material and its properties. Except for the "precious" metals, such as gold, metals in a refined form are inherently unstable. This instability is what drives the process of corrosion as a refined metal is continually trying to revert to its natural state (the mineral), and some metals do this faster than others. Forgotten, ignored and often deferred, corrosion eventually threatens the integrity of aircraft structures.

The NRC corrosion damage atlas includes “real-world” examples from various corrosion forms and from a variety of material systems, which have been found in military aircraft fleets from RCAF as well as the fleets from some TTCP nations. Eight forms of corrosions are documented, i.e. 1) Pitting corrosion; 2) Intergranular corrosion (IGC); 3) Exfoliation corrosion; 4) Crevice corrosion; 5) Filiform corrosion; 6) Galvanic corrosion; 7) Stress corrosion cracking; 8) Fretting corrosion. The damage atlas includes the images from destructive and non-destructive inspections (NDI) of corrosion-damaged areas. Due to the limited space, this paper only presents some examples on pitting corrosion, intergranular corrosion, and exfoliation corrosion. For comparison, this paper also presents some corrosion sketches updated in the TR021 and explains why and how to use the damage atlas together with the TR021 for more accurate damage characterization, and for further analyses of corrosion effects on structural integrity.

3.1 Pitting Corrosion Some sketches of pitting corrosion in the CFTO TR021 are shown in Figure 1. Note that the sketch Figure 1 a) of TR-021-97/08 is now replaced by two sketches (Figure 1 b) and c)) in TR-021-2016. The new sketches provide more information on electrochemical mechanism of pitting corrosion (Figure 1 b)), and, more importantly, the morphology of pitting (Figure 1 c)) shows the possibility of pitting tunnelling underneath the surface, or spreading under the surface, depending on the material microstructures.

In the NRC corrosion damage atlas, some real-world examples are presented (Figures 2-5), including some images taken from progressive sectioning, with scale bar indicating the pit size. These images reveal that corrosion pitting is one of the most insidious forms of corrosion because the pits are often very small and difficult to see with the naked eye, particularly if they are hidden by general corrosion products or coating. The conditions at the base of a pit can be such that other forms of corrosion, such as intergranular attack can occur, leading to wide subsurface damage as shown in Figure 3. The shape of a pit can be influenced by the grain flow direction as demonstrated by the images in Figure 4. These understandings are critical to evaluate the actual level of pitting and prepare a proper corrosion clean/repair procedure to completely remove the corrosion pitting. Otherwise, even a small amount of remaining corrosion could restart corrosion at an even faster rate.



a) Pitting corrosion [1] (TR-021-97/08)

b) Mechanism of pitting Corrosion [6] (TR-021-2016)

c) Morphology of pitting [6] (TR-021-2016)

Figure 1: Pitting corrosion sketch and updates in DND C-12-010-040-TR-021

Figure 2: Optical micrograph of pits along faying surface of 2024-T3 fuselage lap joint corroded 2024-T3 fuselage lap joint

Figure 3: Optical micrographs showing additional damage at base of pits found on the faying surface of a 2024-T3 fuselage lap joint

Clad layer removed



a) Pitting has occurred in the plane parallel to

the rolling direction b) Pitting has occurred in the short transverse

plane and resulted from the corrosion media attaching to the exposed end grains

Figure 4: Optical micrographs showing grain flow direction in aluminum 2519 on the shape of pit

If corrosion pits are presented in highly loaded structures, the stress concentration of a pit can be sufficient to cause fatigue or stress corrosion cracking to occur prematurely. Figure 5 shows an extensive fatigue crack that formed at a corrosion pit in a 7149-T73511 aluminium extrusion of an F-18 centre fuselage, which was used as a transition structure for a CF188 wing full scale test (FT-245) [7]. This fuselage was from a retired United States Navy (USN) F-18 aircraft that had been in-service for ten years (1984-1994) on an aircraft carrier. Extensive corrosion was reported on the longeron by the USN during its service life, which was repaired by blending out. During the full scale wing test, the cracked component was discovered after a fairing and sealant were removed to perform a scheduled life enhancement procedure. Corrosion pits were found in the area near the apparent crack nucleation site. The corrosion pit measured from this replica image (Figure 5 b)) was used for a corrosion fatigue analysis [7][8], to reproduce the crack growth history matching with the full scale test.

a) Image showing fatigue crack b) Image of corrosion pit (dotted line) at crack

nucleation site – side view

Figure 5: Upper outboard longeron of CF188 centre fuselage (7149-T73511 aluminum extrusion)

3.2 Intergranular Corrosion (IGC) and Exfoliation Corrosion The updated TR021-2016 defines intergranular corrosion and exfoliation under the section (“Intergranular Corrosion”). As proposed by NRC, the TR021-2016 refined the IGC definition with more material microstructure information as, “a highly localized form of dissolution which affects the grain boundary regions in a



polycrystalline metal. Depending on the grain structure of the metal or alloy, intergranular corrosion can produce a network of corrosion or cracking on metal surfaces, occasionally dislodging whole grains, or it may penetrate deep into the metal leaving very little visible evidence of damage. In heavily rolled or extruded metals where grains are flattened and elongated in the direction of working, the presence of intergranular corrosion can lead to layering and flaking producing a delamination effect with surface grains being pushed out by the underlying corrosion products. This is known as exfoliation corrosion and is essentially a severe form of intergranular corrosion occurring in the direction of grain flow”. Figure 6 presents the updated sketches of intergranular corrosion/exfoliation corrosion in the TR021-2016, which however do not show material microstructural features associated with the corrosion damage.

Figure 6: Intergranular/Exfoliation Corrosion in C-12-010-040-TR-021 [6]

The NRC corrosion damage atlas describes intergranular corrosion and exfoliation in two separate sections; it provides additional real life examples, along with NDI inspection results as well as microstructure information.

3.2.1 Intergranular Corrosion (IGC)

A typical example of IGC, found on a fuselage lap joint faying surface, is presented in Figure 7. It shows that IGC produced a network of corrosion through microstructures, penetrated into the thickness of the 2024-T3 material, but leaving little visible evidence of damage on the surface.

a) Intergranular corrosion on faying

surface of a lap joint b) Intergranular corrosion that had penetrated 0.13mm

(0.005 inch) into the thickness on a lap joint

Figure 7: Micrographs showing different levels of intergranular corrosion that had penetrated into the thickness of 2024-T3 material



In another example, shown in Figure 8, anomalies were detected at fastener holes in an aluminium 7178-T6 upper wing skin using NDI techniques. These areas were sectioned and progressively polished to characterize the damage. The results clearly demonstrate the presence of IGC, which had penetrated deep into the thickness (up to 3~4 mm). This would most certainly cause problems in repairing this damage since it is unlikely that the deepest IGC would be removed.

Figure 8: Results from a progressive polishing carried out on the section B (left) that did not contain any visible exfoliation in 7178-T6. Arrows indicate the location of the deepest intergranular corrosion attack

Figure 9 shows a group of images from a bathtub bay panel (inside of fuel bays) of a helicopter. This panel was an aluminium honeycomb structure with an inner layer of 0.012” thick 2024-T3 sheet, and an outer layer of 0.008” thick 2024-T3 sheet. After the bays were opened, corrosion damage was found on the inner sheet especially near the drainage areas where water condensation was evident. As the aircraft fleet mainly operated in a maritime environment, the corrosion was estimated to have formed due to a mixture of fresh water and salty air. From the fractography (Figure 9, c)-d)), the corrosion was determined to be intergranular corrosion [9]. Some of the corrosion damage found was significant enough to penetrate through the thickness of the aluminium sheets, which could form sharp edges and cut the fuel cells in the fuel bays.

a) Photograph of intergranular corrosion present in 2024-T4 aluminium honeycomb structure b) Location of sample taken along the dotted lines



c) Transverse and Longitudinal views of the sample d) Zoom-in transverse view of the sample

Figure 9: Photograph of IGC present in 2024-T3 aluminium panel (a)), and detailed examination of corroded area to determine this corrosion type (from b) to d)) [9]

3.2.2 Exfoliation Corrosion

In the past decades, NRC has carried out extensive research on aircraft exfoliation corrosion, using naturally corroded samples collected from manufacturers and retired aircraft. A “typical” exfoliation found on aircraft upper wing skin made from 7186 aluminium alloys and containing steel fasteners is presented in Figure 10. The exfoliation typically forms at the edge of a hole containing a steel fastener and can grow outwards, sometimes towards the adjacent fasteners, as shown in Figures 10 a) and b). Investigations have indicated that the exfoliation at fastener holes is caused by a galvanic action when electrolytes penetrate the space between the steel fastener and fastener hole [10]. Depending on the severity of the exfoliation, the damage surface could appear to have a bulge as shown in Figures 10 c) and d).

a) Upper wing skin (7178-T6) showing locations of exfoliation (circled) at fastener holes. The exfoliation

present was ground flat

b) Upper wing skin (7178-T6) showing locations of exfoliation (arrows)

Transverse

Longitudinal



c) Upper wing skin (7178-T6) containing exfoliation

d) Upper wing skin (7178-T6) containing severe exfoliation with bulging of surface

Figure 10: Images of various levels of exfoliation present at fastener holes in upper wing skins

Once exfoliation progresses beyond the fastener head it can manifest itself as a bulge on the exterior surface as shown in Figure 10 c) and d). This is also clearly shown in Figure 11 where a section of a corroded upper wing skin was cut and progressively polished. However it should be noted that this bulging does not always occur away from the hole as demonstrated in the cross-sectional image shown in Figure 11 c).

a) Optical image of specimen showing locations of polishing sections

b) Optical micrograph of polished wing section from section i shown in the figure a) above

c) Optical micrograph of polished wing section from section k shown in the figure a) above. Note

that the bulge is located away from the fastener head but this is not always the case

Figure 11: Optical images of an exfoliated 7178-T6 upper wing skin

For exfoliation underneath the surface or beneath a fastener head, visual or enhanced visual techniques would have difficulty to detect the corrosion. Other methods such as eddy current (EC) can be used to detect corrosion when it is still beneath the fastener heads, but removal of the fasteners is necessary to enable insertion of the probe inside

i k


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the hole. To prevent this time consuming operation, ultrasonic (UT) methods based on normal and oblique incidence angles can be used since exfoliation corrosion propagates parallel to the surface and the corrosion products have different acoustic impedance as compared to the base alloy. However, these methods can only detect the corrosion in the first layer unless there is adequate coupling such as sealant or adhesive between the layers of the structures.

Figure 12 shows two-dimensional amplitude C-scan inspection results along with photographs of one test coupon. Figure 12 a) is the top-view photograph and Figure 12 b) is the ultrasound C-scan image of the top surface obtained by gating the top-surface echo. Corrosion and other surface features, seen in the C-scan image, correlate nicely with the visible features on the photograph. Figure 12 c) is the ultrasound C-scan image of the entire coupon, obtained by using the back-wall echo, that shows discontinuities in the entire ultrasonic beam path including the top surface. In order to see the internal corrosion sites better, image subtraction was used and the result is presented in Figure 12 d). In this figure, lighter areas correspond to surface features and darker areas correspond to internal corrosion. The image shows internal exfoliation or intergranular corrosion, pointed out by the arrows, extending beyond the fastener holes. However, the presence of fasteners prevents interrogation of areas beneath the fastener heads. Regardless of the method employed, NDI to quantify the level of exfoliation that is present in an aircraft component is a very time consuming task.

a) Photograph showing visible surface features

b) C-scan image of the top-surface

c) C-scan image of the entire specimen


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(d) Subtraction of (b) and (c) showing internal exfoliation corrosion as pointed by arrows

Figure 12: Exfoliation coupon and NDI results

From the above examples, it is seen that the NRC corrosion damage atlas provides more details on material microstructural information, NDI inspection and fractography results. These results are expected to provide the front line maintenance personnel with more in-depth corrosion knowledge to better determine the corrosion levels and associated maintenance actions, but also would allow the ASIP managers to assess the impact of corrosion on structural integrity and airworthiness.

4.0 CORROSION EFFECT ON STRUTURAL INTEGRITY

The CFTOs like the TR021 categorize three levels of corrosion, i.e., Light corrosion (a depth of 0.001” maximum); Moderate corrosion (as deep as 0.010”); and Severe corrosion (including pitting deeper than 0.010”). This categorization approach does not consider the loading/stress levels at the structural locations, and has no details on the effect of different levels of corrosion on structural integrity. One reason for this is the perception that corrosion does not pose a significant safety risk but is primarily a maintenance issue. Experience, however, has shown that this, in fact, is not the case since failures have been attributed to corrosion damage. When the corrosion is located at even moderately stressed structures, it could increase the risk of stress corrosion cracking or corrosion fatigue cracking. Since the 1990s, NRC has carried out extensive work to study corrosion effects on structural integrity; in order to develop various methodologies/tools and generate the necessary data to introduce the “Anticipate and Manage” philosophy for the USAF as well as extend the Holistic Structural Integrity Process (HOLSIP) to include environmental degradation (www.holsip.com). This section summarizes a few case studies for corrosion examples that are also presented in the damage atlas.

4.1 Effect of Pitting Corrosion For the pitting corrosion example presented in Figure 5, a corrosion fatigue analysis [7] was originally carried out to reproduce the fatigue life of the F-18 upper board longeron. Additional analyses [8] used different methods, i.e. strain-life (SL, using the tool CI89), damage tolerant (DT, using the tool AFGROW), and a holistic lifing method (using the tool ECLIPSE from Analytical Process/Engineered Solutions (APES)). The holistic method/tool accounted for the concurrent interaction between corrosion and fatigue. The corrosion pit growth rate was assumed to follow a cube root power law relation, as shown in Figure 13 a). Figure 13 b) presents all the analysis results in comparison with NRC full scale test results, with and without corrosion effects. It is shown that the typical DT analysis estimated the shortest life; the SL estimated the longest life; and the holistic method/tool (ECLIPSE) estimated the life closest to actual service experience. This NRC research indicated that the holistic method is capable of quantifying the impact of an evolving corrosion pit on aircraft structural integrity.


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Corrosion pit growth rate [7]

Figure 13 Corrosion (pitting) and fatigue modeling, F-18 example [7][8]

4.2 Effect of Intergranular Corrosion (IGC) The IGC example presented in Figure 9 was analysed by NRC [11] to support DND’s airworthiness assessment and maintenance actions. Because an initial maintenance schedule shows that the contractor proposed repair procedure, for the entire fleet, would take several years to complete, DND wanted to know what level of corrosion damage that would be present in the aircraft which would be deferred to be repaired later. Based on the damage information, an airworthiness risk assessment was required to determine the level of risk for these aircraft. NRC proposed an evaluation method to estimate the size of the corrosion damage. This method was based on some laboratory accelerated corrosion tests (Figure 14) [12][13], and a power law equation, d=A(t)1/B (d: corrosion depth; A: material/environment parameter; t: time; B: exponent) assumed for 2024-T3. Although the lab tests were not exactly the same as the fuel bay panel corrosion process, a comparative study was carried out by conservatively assuming that all aircraft panels were corroded and the corrosion grew at a faster rate (ex., B=2 is more aggressive than B=3). NRC results [11] were used by DND to complete a qualitative airworthiness risk assessment, leading to a decision on a proper repair schedule to ensure the fleet availability.

4.3 Effect of Exfoliation Corrosion As shown in Figures 10-12, exfoliation corrosion can spread/grow not only on surface but also penetrate/grow into the thickness. The depth of exfoliation can affect the structural integrity more than its area on the surface [14]. In 2004, to support a project on corrosion effects on structural integrity (CESI) sponsored by the USAF, NRC developed a statistical analysis tool [15] to estimate the depth of exfoliation using an existing grind-out database, which contained the grind-out width and depth measurements taken after repair actions. However, to be able to use this database an assumption had to be made, i.e., the dimensions of the initial exfoliation would match the grind-out dimensions. Due to the large scatter that is typically present in such a database, plus the potential measurement error on surface corrosion (width), a nonparametric statistical method was developed at NRC to determine statistically meaningful depth estimations from the grind-out database. A schematic showing the analysis tool is shown in Figure 15 a) and the estimated results are shown in Figure 15 b). In addition, the depth estimation can be used to investigate the capability of the NDI techniques for determining the depth of exfoliation corrosion. The NDI results can also be combined with the developed tool to provide better depth estimation with higher confidence levels.


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a) Penetration test on AA2024-T3 1.9 cm plate, in ST, L and LT directions [12][13]

b) Penetration tests for 8.9 cm thick plate of AA2024-T3 in oxygenated 1 M NaCl at -580 mV, in ST, L

and LT directions [12][13]

Figure 14 Penetration test results on AA2024-T3 plates [12][13]

Exfoliation depth calculator using a grind-out

database [15] Estimated depth vs. width with different

probability and errors [18]

Figure 15: Estimating exfoliation depth using grind-out database and surface width measurement

In an effort to move away from the current costly “Find-It-Fix-It” maintenance approach toward the “Anticipate and Manage” approach, NRC conducted further research to evaluate the effects of exfoliation on residual strength and remaining fatigue life. Many static and fatigue tests were completed using pristine, artificially corroded, and naturally corroded specimens (7075-T6511) [16]. Both thermographic and ultrasonic NDI were performed to determine the maximum depth and 3D profiles of the exfoliation damage. Fatigue crack origins were determined with the aid of a scanning electron microscope and the nucleating mechanisms in the exfoliation regions were found and documented. An analytical model [17] was developed based on a ‘soft inclusion’ technique and 3D FEM. This technique automatically generates the 3D geometry of the ‘soft inclusion’ (damage zone) from the ultrasonic NDI input (Figure 16). For engineering applications, a simplified fatigue model [18] was developed for estimating the remaining fatigue life of the exfoliated specimens. The NRC study found that, exfoliation may not have severe detrimental effect on residual strength and stress concentration (Figure 16); some exfoliated coupons

0.000.020.040.060.080.100.120.140.16

0.0 0.5 1.0 1.5 2.0 2.5 3.0 3.5Width (inch)

Dep

th (i

nch)

Steel 360&733Steel FastenersAluminum FastenersDEPTH, LOWER(P=10%), ERR=0.1"DEPTH, MEAN(P=50%), ERR=0.1"DEPTH, UPPER (P=90%), ERR=0.1"DEPTH, B-BASIS(P=90%,CL=95%), ERR=0.1"


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had longer lives than those where corrosion had been ground-out; a calibrated fatigue model is capable of estimating the remaining fatigue life.

Figure 16: Modelling of exfoliation corrosion effect on upper wing skin (7178) structural integrity

In collaboration with RUAG Aerospace, NRC carried out further research on exfoliation modeling and testing to investigate the effects of exfoliation corrosion on the static and fatigue behaviour of F-5E upper (7075-T651) and lower (7075-T7351) wing panels [19][20][21]. An engineering framework was developed to model exfoliated areas by the NRC ‘soft inclusion’ model (Figure 17) and fatigue metrics (for instance pit depth) for pitting corrosion in the lower wing skin. The developed framework provides the structural engineer with a method to assess the impact of corrosion on the structural integrity of a wing, and provides an engineering basis to decide on corrosion repair actions (when and how).

a) F-5E wing exfoliation corrosion and modelling b) Analytical and test life comparison

Figure 17: Modelling of exfoliation corrosion effect on F-5E upper wing (7075) structural integrity


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5.0 DISCUSSIONS

Corrosion Growth Rate: One of the major gaps present today is the existence of accurate corrosion growth rate models [22], especially for proactive maintenance actions and structural integrity assessment. The models that do exist were developed using accelerated corrosion tests (such as the exponential rule for maximum pit depth growth). However there are no “transfer functions” to relate accelerated corrosion tests to actual growth rates. Since corrosion growth rates vary dramatically depending on the type of material, corrosion process (e.g. pitting versus exfoliation versus stress corrosion cracking) and atmospheric conditions, studies need to be carried out to either determine realistic corrosion growth rates or develop transfer functions to relate accelerated rates to realistic ones. Presently, a more realistic rate can only be estimated by extrapolating the corrosion rate from previous metal loss or grind-out database (ex. Figure 15). This can be done through analysis by measuring what is believed to be the most severe damage metric and then reverse growing the damage exponentially over some time period. In addition, there are no estimates as to the time required to breakdown a corrosion protective coating, such as cladding on Aluminium 2024-T3. Both these rates are required to accurately predict the future state of a component susceptible to corrosion.

Corrosion of New Material: Some new materials have been developed with more corrosion resistance, for example the new aluminium alloys 7249-T76511 extrusion replaced the legacy 7075-T6 extrusion in the CP140 new wings. Figure 18 presents preliminary results [23] from the laboratory accelerated corrosion testing (ASTM EXCO and ANCIT) for three aluminium alloys, 7075-T7351 forging, 7249-T76511 extrusion, and 7075-T6 extrusion [23]. It is shown that, compared to 7075-T6 extrusion, 7249-T76511 extrusion has much less corrosion damage which means more corrosion resistance, but this new material still gets corroded under the test environment. For the new material, the data on in-service damage experience, corrosion growth rate, and effect of existing or updated CPC/CIC would be useful to ensure an effective CPCP for long term fleet management. The corrosion damage atlas should be updated with more real-world examples including the new materials.

Figure 19: Accelerated corrosion tests for aluminum alloys, 7075-T7351, 7249-T76511, and 7075-T6


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6.0 CONLUDING REMARKS

This paper provides a brief summary of a NRC project to review the CPCP documents for all the RCAF fleets. Based on that project a corrosion damage atlas was developed at NRC for more accurate damage characterization and more effective corrosion management. The NRC corrosion damage atlas includes real-world examples with photos showing corrosion damage size and morphology, detailed material microstructure information, as well as non-destructive and destructive inspection/fractography results. This paper highlighted some examples from the damage atlas, and demonstrated that the NRC corrosion damage atlas can be used in conjunction with the RCAF corrosion control manuals (ex. C-12-010-040/TR-021) to improve corrosion level/severity evaluation. Furthermore, the presented corrosion examples were used for assessing their effects on structural integrity in order to assure aircraft safety and continued airworthiness. In the end this paper discussed a major gap/need on corrosion growth rate model/data. This paper suggests to continue updating the NRC corrosion damage atlas by collecting more real-world examples from service and including new materials. This way, it is expected that operators can continue improving military aircraft corrosion management with reduced cost and improved availability.

ACKNOWLEDGEMENTS

This work was completed with financial support from DRDC (Defence Research and Development Canada).

Specifically the author would like to acknowledge Nick Bellinger who led the effort to develop the NRC corrosion damage atlas. Nick passed away on August 9th, 2015; this paper is dedicated to him in memory of his significant contributions to aging aircraft corrosion research.


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REFERENCES

[1] Estimated Impact of Corrosion on Cost and Availability of DoD Weapon Systems -FY18 Update, LMI, March 2018.

[2] Estimated Impact of Corrosion on Cost and Availability of DoD Weapon Systems -FY17 Update, LMI, March 2017.

[3] MIL-STD-1530D, Department of Defense Standard Practice, Aircraft Structural Integrity Program (ASIP), October 2016.

[4] Department of National Defence Canada, AIRCRAFT CLEANING AND CORROSION CONTROL EXTERIOR AND INTERIOR, C-12-010-040/TR-021, 1997-04-30 (Ch/Mod 5 2008-01-16) (referred as TR021-97/08 in this paper).

[5] Bellinger, N.C., Airframe Corrosion Damage Atlas, LTR-SMM-2015-0128, June 2015.

[6] Department of National Defence Canada, AIRCRAFT CLEANING AND CORROSION CONTROL EXTERIOR AND INTERIOR, C-12-010-040/TR-021, 2016-11-17 (referred as TR021-2016 in this paper).

[7] Liao, M., Bellinger, N., and Komorowski, J.P., Rutledge, B., and Hiscocks, R., Corrosion fatigue prediction using holistic life assessment methodology. The Proceedings of the 8th International Fatigue Congress (FATIGUE 2002), Stockholm, Sweden, June 2002, published by EMAS UK, edited by A.F. Blom, Vol. 1, pp. 691-700.

[8] Liao, M., Bellinger, N.C., Komorowski, J.P., Chapter17-Corrosion fatigue analysis of an F-18 Longeron, published in Corrosion Fatigue and Environmentally Assisted Cracking in Aging Military Vehicles, RTO-AG-AVT-140, Research and Technology Organisation of NATO. 2011 Mar, ISBN 978-92-837-0125-5.

[9] Butler, C., Fuel cell bay corrosion on CH149 Cormorant, QETE Report A017406, 2007.

[10] Hoeppener, D.W. and Chandrasekaran, V., Corrosion and corrosion fatigue predictive modeling state of the art review, submitted as a part of the research program “Modeling Corrosion Growth on Aircraft Structure” subcontract Number NCI USAF 9061-008, April 30, 1998.

[11] Liao, M., Bellinger, N.C. and Butler, C., Evaluation of Corrosion Damage Growth in CH149 (Cormorant) Helicopter, LM-SMPL-2007-0179, 2007.

[12] Zhang, W., and Frankel, G.S., Localized Corrosion Growth Kinetics in AA2024 Alloys, J. Electrochem. Soc. 149, B510 (2002).

[13] Huang, T.S., and Frankel, G.S., Localized corrosion growth kinetics in Al alloys, the 6th FAA/DoD/NASA Aging Aircraft Conference, 2002.

[14] Liao, M., and Komorowski, J. P., Effect of prior exfoliation corrosion on fatigue and fracture behavior of aging aircraft structures and materials, LTR-SMPL-2002-0136, June 2002.

[15] Bellinger, N.C., Liao, M., T. Benak, A. Marincak, F. Taremi, Corrosion effects on structural integrity, LTR-SMPL-2004-0165, August 2004.


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[16] Liao, M., Renaud, G., Bellinger, N.C., Backman, D., and Forsyth, D.S., Effects of exfoliation corrosion on static and fatigue behavior of aircraft materials and structures -- testing and modeling studies. The 8th Joint NASA/FAA/DoD Conference on Aging Aircraft, January 2005, Palm Springs, California, USA.

[17] Liao, M., Bellinger, N.C., Komorowski, J.P., Modeling the effects of prior exfoliation corrosion on fatigue life of aircraft wing skins. International Journal of Fatigue. 2003 Nov 30;25(9):1059-67.

[18] Liao, M., Mills, T., Chapter 15- Exfoliation corrosion and fatigue modelling, published in Corrosion Fatigue and Environmentally Assisted Cracking in Aging Military Vehicles, RTO-AG-AVT-140, Research and Technology Organisation of NATO. 2011 Mar, ISBN 978-92-837-0125-5.

[19] Renaud, G., and Liao, M., Static and fatigue strength analysis of an upper wing skin with exfoliation corrosion, Proceeding of the 26th International Congress of Aeronautic Science, Anchorage, Alaska, USA, 2008.

[20] Uebersax, A., Huber, C., Renaud, G., Liao, M., Structural integrity of a wing upper skin with exfoliation corrosion. The 25th Symposium of the International Committee on Aeronautical Fatigue (ICAF2009), Rotterdam, Netherlands, June 2009.

[21] Uebersax, A., Zehnder R., Renaud G., and Liao M., Evaluation of the structural impact of corrosion damage on a wing, International Congress of Fatigue (Fatigue2014), Melbourne, Australia, March 2014.

[22] Bellinger, N., A proposed roadmap for corrosion structural integrity research, LTR-SMPL-2013-0142, November 2013.

[23] Liao, M., Bombardier, Y., Fatigue crack growth rate testing and modeling for a new aircraft material (7249): comparison and discussion, 2014 Holistic Structural Integrity Process (HOLSIP) Workshop, Salt Lake City, UT.

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