Nondestructive Evaluation of Composite Repairs

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3Department of Electrical and Computer Engineering, Southern Illinois University ... examination of two different composite repair methods commonly used ...
Nondestructive Evaluation of Composite Repairs Keven R. Mitchell1, Anish Poudel2, Shanglei Li3, Tsuchin Philip Chu2, and Daniel Mattingly1 1 Department of Aviation Technologies, Southern Illinois University 545 North Airport Road, Murphysboro, IL 62966 Tel: (618) 453-9203; Fax: (618) 453-4850; Email: [email protected] 2

Department of Mechanical Engineering and Energy Processes, Southern Illinois University 3 Department of Electrical and Computer Engineering, Southern Illinois University 1230 Lincoln Dr., MC 6603, Carbondale, IL 62901 Tel: (618) 453-7049; Fax: (618) 453-7658; Email: [email protected]

ABSTRACT Composite aircraft structures are found in modern aircraft from Air Transport to General Aviation aircraft designs. Maintenance repair technology varies for each Original Equipment Manufacturer (OEM) and aircraft type. Whereas previous aircraft structural repairs used similar construction, aluminum and rivets, composite aircraft manufacturers certify each aircraft with a mixture of different fiber reinforcements and resin matrix systems. This study is an examination of two different composite repair methods commonly used within the composite aviation industry and how they compare in test using different Nondestructive Evaluation (NDE) techniques. The results obtained from NDE methods were used to evaluate the effectiveness of each method for flaw detection and sizing.

Keywords: Composites, Repair, NDE, IRT INTRODUCTION Composites have been in use on aircraft in one form or another for decades now. Wood laminated props, dope and fabric surfaces are some of the earliest forms found in the older generation aircraft [1]. Today, advanced composites such as carbon fiber reinforced plastics (CFRP), honeycomb materials are increasingly being used in many aerospace structural applications. The main reason for this is, unlike traditional metals and their alloys, composites offer outstanding thermal and physical properties which include high strength and stiffness to weight ratios, low coefficient of thermal expansion, high fatigue resistance, and are less prone to deterioration caused by corrosion and cracking. [2, 3]. However, composite structures are prone to damage which can occur either during the processing stages or during inservice operations. During in-service operations, the composite structures are primarily damaged by lightning strikes, hail damage, low/high velocity impacts caused by runway debris and birds, tool drops, and service vehicle collisions. Other sources of damage include erosion, abrasion, manufacturing defects, excessive heat exposure, and fluid infiltration [3]. Sometimes damage to composite components may not be visible to the naked eye. Therefore, various non-destructive evaluation (NDE) methods are often applied today to identify, characterize these defects and evaluate the extent of damages in structural components [3-8]. In most cases, replacing the entire part is not economically feasible; thus, repairing it is the only viable solution. Furthermore, since the time constraint is an issue, repairs must be performed as quickly as possible so that the aircraft can be returned to service as soon as possible [3]. The standards for repair on sheet metal aircraft are well known and documented with published substitutions and variances. Typically, repair technicians have a supply of rivets, fasteners, and various sheets of metal on hand to mend most damaged metal aircraft. But, the adhesively bonded repair of damaged composite sandwich structures is a complex problem because several factors must be considered to ensure the repair’s effectiveness and structural integrity. Some of these factors include stiffness, strength, stability, operating temperature, durability, and aerodynamic smoothness [9]. With the increase of composite aircraft there is a requirement for standard and effective repair procedures. Several aircraft manufacturers have established unique repair procedures specifically for each type design based on their rigorous tests and analyses. Aircraft Maintenance Technicians (AMT) and inspectors are adapting their technology and training requirements to cope with these composite aircraft. The repair techniques used in the old fiberglass shops on nonstructural components are not suited for the advanced composites used today. Composite repair training centers are being established to provide suitable courses. This paper examines and compares the repair requirements for the two leading general aviation composite aircraft manufactures and the impact using step-heating infrared thermography (IRT) over

Coin-Tap test, and pulse-echo UT. For this work, 2 core 2 composite sandwich test coupon, made with carbon fiber laminate, nomex honeycomb, and glass fiber laminate, was used. Impact damages were simulated in these sandwich coupons and repairs were carried out by using repair procedures of the two leading general aviation composite aircraft manufactures. During the repair process, controlled damages were also induced so as to simulate weak bonds, disbonds, and porosity. For this, improper surface preparation was performed on the repair patches to simulate disbonds, a void was created between the filler and plies to simulate disbonds, and improper curing was conducted to simulate porosity. Finally, Coin-Tap test, pulse-echo UT, and the infrared thermography (IRT) NDE methods were applied to evaluate the composite repair.

METHODS This section describes the fabrication of a sandwich test coupon, the repair process, and the NDE method applied to evaluate the composite repair. The raw materials used to fabricate a 2 core 2 composite sandwich test coupon include Carbon Fiber Fabric 3K 2x2 Twill Weave 5.8oz/yd, AHN 7800 commercial grade Nomex Honeycomb 1/8 inch thick, Fiberglass Cloth 7781-05, and West 2 part Epoxy system. To balance the need for a representative number of layers for test parts and availability of material, a symmetrical 2-ply layup of [0/+45]s was used. Sandwich test coupons were prepared by using a wet-layup with vacuum bag curing. The test coupons were 6 inch x 11.25 inch x .16 inch and was impacted by hammer to simulate the outer laminate skin and core damage. The first step of any composite repair is to determine the extent of any suspected damage. Both general aviation aircraft manufactures recommend using visual inspection of both sides of the composite surface whenever accessible. Visual inspections should be followed by a coin tap inspection of any suspected damage areas. Coin tap inspections are specific for each aircraft OEM. Any suspected damage found during coin tap inspections is referenced to OEM Aircraft Maintenance Manuals (AMM) for placement of subassemblies structures, and core termination edges. Damage assessment is one of the few times a technician can make improvements as long as the inspections are visual in nature. The techniques used include the use of high-intensity lighting, endoscopes, eye looms, lighted magnifying glasses, borescope, and mirrors. Visual inspections should not be limited by eyesight, but accompanied with a hand surface inspection. By running one’s hands over the surface of a suspected damaged area, the technician can often detect imperfections and anomalies [10, 11]. Other aircraft manufacturers requires it’s AMM to “Push the middle of the area to be tested with his thumb if no cracks are found in the skin. If he can feel the skin of the aircraft hitting the core of a sandwich (or other layer/component), the skin is considered disbonded and must be repaired” [12]. Other NDE methods such as infrared thermography, ultrasonics, radiography, and laser shearography are also applied in direct coordination with the individual aircraft OEM under repair. For this work, sandwiched core damage repair is compared to laminated sandwiched core damages on the upper left wing skin of one of the manufacturer at Wing Station 120 as shown in Figure 1.

Figure 1: A schematic portraying upper left wing skin (top view) of one of the composite aircraft manufacturer. The damages induced in the test coupons were a single dent with the maximum depth of 0.05 inch (1.27 mm). The outer face sheet and core were delaminated with crushed core material. Since each aircraft OEM's wing is constructed and attached differently to the fuselage, the wing station location will be related to the center line of the wing; clear of any spars and ribs. The inner laminate sheet i.e. fiberglass suffered no damage or disbonding from the core.

Both OEMs consider this repair to be major structural damage and both have published repair procedures for upper wing skin sandwiched core damage. Table 1 shows the wing laminated sandwiched core repair comparisons for both OEMs. The repair fabric material and structural resin used for this work is different from the OEM’s because of the availability of material in the laboratory. The substitution of materials affects the strength of the coupons but not the methods used. This study is an examination of two different composite repair methods used and not the strength of each repair. Table 1: Comparisons of the wing laminated sandwich core repair for different Aircraft manufacturer

Original ply count Replacement plies Repair fabric material Scarf distance per ply Structural resin Cloth to mixed resin ratio Core replacement

OEM “A”

OEM “B”

2 2 Carbon Fiber Fabric 1.0 in (25.4 mm) MGS L418 100 : 42 grams (+/-) 4

2 4 Carbon Fiber Fabric 0.5 in (12.7 mm) MGS L418 No published ratio

Replaced damaged core (1inch dia.) with bonding paste

Replaced damaged core core (1inch dia.) with nomex honeycomb and core splice (honeycomb wrapped on the edges with the foaming tape).

After removing damaged material (1 inch diameter) and establishing a proper ply count and orientation, repairs were carried out by using the wet layup process as required from both OEMs. For this, first core replacement or fills were applied per the Aircraft Maintenance Manuals (AMM) or Structure Repair Manuals (SRM) and the fiber placement laminate was replaced. The Laminating Transfer Method (LTM) was used during the repair process which is comparable to laminating in place except the repair plies are laid-up on release film first, cut to size, wet-out and stacked in reverse order for placement of plies. Once prepared, the stack was lifted up off the repair table and placed face down on the damaged area. This method simplifies multiple ply repairs while maintaining orientation and replacement schedule.

Figure 2: OEM “A” damage core cell shown with replacement ply order The ply orientation in the repair described by each OEM is different with the order, size, and replacement schedule. OEM “A” removes the damaged surface plies and crushed core. The core was replaced with bonding paste and the original damaged plies with two plies arranged in a “wedding cake” fashion, first ply being the largest and the final ply the smallest as illustrated in Figure 2. Each replacement ply was staggered at 1 in (24 mm) for carbon fibers. The replacement ply orientations for the damaged area were 0 ° and 45 °. However, OEM “B” replaces the plies in reverse order compared to OEM “A”. OEM “B” removes the damaged surface plies and crushed core. The core was replaced with nomex honeycomb (Cut out the same size as the damaged area) and core splice. The original damaged plies are placed in a fashion such that the first ply is the smallest and each consecutive

ply larger and overlapping the preceding ply with the final ply being the largest as shown in Figure 3. The ply orientation applied was -45°, +45°, -45°, and +45° such that +45° was the outer ply. In order to intentionally create a bad repair in both cases, damages were simulated by improper surface preparation, creating a void between the filler and plies, by not applying the core splice, and improper curing.

Table 2 describes the test coupons prepared during the repair process using two different OEMs. Figure 4 shows the repair coupons prepared by following two different OEMs instruction.

Figure 3: OEM “B” 2:1 ply replacement with largest ply placement last

Table 2: Description of the test coupons Panel Naming

Repair Procedure

Repair Type

Simulated Defect Types

A1

OEM “A”

Good

None

A2

OEM “A”

Bad

No core splice applied in the replaced core to simulate disbond and improper curing to simulate the porosity.

B1

OEM “B”

Good

None

B2

OEM “B”

Bad

Void between the filler and plies to simulate disband, and Poly Vinyl Alcohol (PVA) release applied on the left half of the repair area to simulate a weak bond.

Figure 4: Pictures of cured repaired sandwich composite. For this study, the step-heating infrared thermography (IRT) method was applied to inspect the damaged areas in the sandwich test coupons. The experimental setup used for conducting IRT tests is shown in Figure 5. This IRT system is the same as described in [13]. Except, for this study four low power (500 Watts) halogen bulbs were used as a heat source and a longer heat flux application was used. Prior to conducting the experiments, the surfaces of the sample were painted with a water soluble dry graphite paint lubricant for emissivity correction.

Hood with 4 500W Halogen lamps

IR Camera

Computer

Power for Heat Source Controller

Test Coupon

Figure 5: Infrared Thermography setup

RESULTS The first step of any composite repair is to determine the extent of any suspected damage. Prior conducting IRT, we inspected the reapired test coupons by using coin tap-test and pulse-echo UT. The Coin-Tap test was not able to distinguish the differences between the good and bad repairs. The pulse-echo UT was also not able to generate any convincing results. Therefore, step-heating IRT was applied to test all test coupons which had simulated impact damages. A continuous heat flux was applied for approximately 2 sec., test coupons were then allowed to cool, and the temperature responses were recorded which allowed for a temperature variation to be seen within the defect areas. The IR camera recorded the data for 300 frames at 0.03 frames/sec. The best image in the sequence with the highest thermal contrast (IR image showing most of the defects within the sample) was used for the analysis. The damages induced in the test coupons were a single dent with the maximum depth of approx. 0.05 inch (1.27 mm) which was verified by Visual inspection. Prior to conducting IRT, the damaged test coupon’s edges were sealed by using flash tape to run pulse-echo immersion ultrasonic tests (UT) at 1 MHz. The main reason for this was to prevent moisture ingression into the honeycomb sandwich coupons. Figure 6 shows the IR image result and surface temperature-time curve for regions of interest (ROIs) obtained for coupon A1. The temperature difference from the impact damage region to the sound area in all test coupons were approximately 12 ºC to 14 ºC. In the IR image results as shown in Figure 6 (a), the edges of the test coupons were colder as compared to other areas of the test panels. This was mainly due to the possible water intrusion along the edges during the immersion ultrasonic testing. 61 ROI 1 ROI 2

Temperature

54 46 39 31 24 0

(a)

30 60 90 120 150 180 210 240 270 300 Frames

(b)

Figure 6: IRT test results for the test coupon A1 (a) best contrast images at frame 62, (b) surface temperaturetime curve for ROIs.

After the repair process was completed, repaired test coupons were inspected by using step-heating IRT to evaluate the repair for all 4 test coupons. Figure 7 shows the IRT test results for the repaired test coupons. The IR image results shown in Figure 7(a) are the best contrast image obtained at frame 59 for the repaired test coupon A1. This is considered to be a good repair and the IR image results do not show any anomalies in the repair region.

4th ply

3rd ply

2nd ply st Replaced Core 1 ply (a)

Fiber Porosity wrinkling (b)

Voids

2nd ply 1st ply

Bonding paste (c)

Weak bond areas (d)

Figure 7: IRT test results for the repaired test coupons (a) coupon A1, good repair, (b) coupon A2, bad repair, (c) coupon B1, good repair, (d) coupon B2, bad repair. Figure 7(b) is the best contrast image obtained at frame 81 for the repaired test coupon A2. This is a bad repair which contains no core splice in the replaced honeycomb and was not cured properly. The IR image results show porosity distribution and fiber wrinkling in the repair region. The disbond due to no core splice was also revealed in the repair areas. Figure 7(c) is the best contrast image obtained at frame 67 for the repaired test coupon B1. This is a good repair and the IR image results do not show any anomalies in repair region. Figure 7(d) is the best contrast image obtained at frame 63 for the repaired test coupon B2. This is a bad repair which contains improper surface preparation on the right side of the repair area. The application of PVA release foam created a weaker bond between the cured Carbon laminate and repair patches. As a result these areas are characterized as hot areas in the IR image results.

CONCLUSIONS This study evaluates two repair techniques based on two commercial OEM’s by using Infrared Thermography method. The test coupons manufactured for this work were impacted by a hammer to simulate the outer laminate skin and core damage. Step-heating (SH) infrared thermography testing (IRT) was implemented to determine the extent of impact damages in all test coupons. Upon visual inspection, it was determined that all test coupons had a single dent with the maximum depth of approximately 0.05 inch (1.27 mm). The core area replaced was 1 inch in diameter for all test coupons. SH IRT was able to identify the impact damages during the damage assessment process. During the repair process, two commercial repair procedures were utilized. In addition, two bad repairs were fabricated in each of the OEM repairs. In order to intentionally create a bad repair in both cases, damages were simulated by improper surface preparation, creating void between the filler and plies, using two methods, not applying the core splice, and improper curing. SH IRT was able to distinguish between the good repair and bad repair. Whereas, the Coin Tap

inspections found no defects. The good repairs displayed consistent colorization not revealing hot or cold spots in the repaired areas. However, the bad repairs were characterized by the hot spots in the IRT results. We demonstrated that IRT can be an efficient and fast tool for examining the repair areas in addition to the Coin Tap inspections. If applied on a periodic basis during maintenance and repairs the processed data can then be used to confirm serviceable repairs and monitor any potential problem areas.

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2.

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