In Situ Thermal Ablation Repair of Delamination in Carbon Fiber-Reinforced Thermosetting Composites

Abstract

Repairing delamination damage is critical to guarantee the structural safety of carbon fiber-reinforced thermosetting composites. The popular repair approaches, scarf repair and injection repair, can significantly restore the in-plane mechanical performance. However, the out-of-plane properties become worse due to the sacrifice of fiber continuity in these repairing processes, leading to the materials being susceptible under service loads. Here, we propose a novel in situ delamination repair approach of controllable thermal ablation in damage removal, achieving a high repair efficiency without impairing the fiber continuity in carbon fiber/epoxy panels. The epoxy resin in the delaminated region was eliminated under the carbonization temperature in a few minutes, allowing the carbon fiber frame to retain its structural integrity. The healing agent, refilled in the damaged region, was cured by the Joule heating of designed electrodes for 30 min at 80 °C, yielding the whole repair process to be accomplished within one hour. For the delaminated carbon fiber/epoxy panels with thicknesses from 2.5 to 6.8 mm, the in-plane compression-after-impact strength after repair could recover to 90.5% of the pristine one, and still retain 74.9% after three successive repair cycles of the 6.8 mm-thick sample. The simplicity and cost-saving advantages of this repair method offer great potential for practical applications of prolonging the service life of carbon fiber-reinforced thermosetting composites.

1. Introduction

Thermosetting composites offer exceptional durability, good processability, and strength-to-weight advantages; hence, they are ideal for a wide range of composite applications, especially those with rigorous demands, such as aerospace materials [1,2,3]. As an important part of aerospace materials, carbon fiber-reinforced thermosetting composites (CFRTPs), composed of high-performance carbon fibers embedded in a thermoset matrix, have dominated in fabricating aerostructures due to their superior specific strength and stiffness along with favorable corrosion and fatigue resistance [4,5,6]. Facing complex service environmental conditions, CFRTPs are susceptible to low-velocity impacts which can induce internal damages barely noticed by normal visual inspections [7,8]. These internal damages manifest in indentation, matrix cracking, interface debonding, and delamination [9,10], among which delamination occurring between layers of inner structure results in a sharp drop of compression strength, and its propagation causes material failure [11,12]. Therefore, repairing delamination is critical for ensuring the in-service safety of CFRTPs, and, for practical applications, a proper repair process should be easy to handle and cost-effective [13,14].

Repair strategies could be categorized into a self-repair process and others requiring human interventions. The typical self-repair approach is achieved by integrating the healing functionalities into the whole CFRTP during or after the fabrication process [15,16]. These healing agents are released in the damaged region as it suffers an external impact, and hence can repair the damages without manual intervention [17,18]. This automatic repair technique introduces additional healing agents, such as micron-sized particles, into the composites, leading to the fabrication process complex. In addition, the additional healing agent may affect the composite mechanical performance and structural integrity [19,20]. In contrast, the strategies that replace the damaged region with materials that are not inhibited in the composite are widely employed, including resin injection repair [21], patch or plug repair [22], scarf joint repair [23], etc. Among them, injection repair [21,24] and scarf repair [25,26] are two major methods for composite panels given their simplicity in procedure and acceptable repair efficiency. The former is recognized as a low-invasive repair method by injecting a low-viscosity resin into the damaged area, followed by curing the resin at a specified condition. The choice of resin is a key factor, which requires low viscosity and high stability [27]. The latter scarf repair method includes the removal of the damaged site and then a matching replacement bonded into this region, yielding a high-strength recovery efficiency [28]. However, the performance of the scarf repair becomes poor in the case of irregular damaged structures such as corners and edges, because of the difficulty in machining the damaged portion of the laminate into a circular shape. Moreover, a common disadvantage of injection and scarf repair is that the implementation process requires drilling holes in the damaged region [27,29] for either the healing agent injection or damage removal, which inevitably imposes a certain degree of damage on the fiber continuity [27,30]. The fiber rupture can lead to out-of-plane stress concentrations, the accumulation of which causes the composite failure [31,32]. Therefore, developing a repair technique that can avoid undermining the fiber continuity and combine efficient, manageable, and low-cost attributes is essential for CFRTPs’ structural integrity and safe operation.

In this work, we proposed a novel delamination repair strategy for removing damage in CFRTPs by an in situ thermal ablation process, and the repair efficiency was evaluated. We manufactured the CFRTPs, woven carbon fiber-reinforced epoxy resin (CF/epoxy) panels with different thicknesses, and produced delamination in the panels via low-velocity impact experiments. The repair process started with the removal of virgin resin in the delaminated region under a high temperature (T = 800 °C), providing a loading of the healing agent. Then, the healing agent was filled into the damaged region and in situ cured at T = 80 °C where the heat was supplied by the electrical resistance of designed electrodes covered on the delamination surfaces. The optimal parameters of the repair process, such as the burning and healing temperature, were explored, and the repair efficiency was evaluated by comparing the in-plane compressive strength of repaired panels with that of original ones.

2. Materials and Methods

2.1. CFRTP Manufacture

The examined continuous CFRTPs consisted of T300 plain woven carbon fabric (Weihai Guangwei Composites Co., Ltd., Weihai, Shandong, China) as the reinforcement and bisphenol A epoxy vinyl ester resin matrix (Tianqi FRP composite materials factory, Dongguan, Guangdong, China). The CF/epoxy panels with different thicknesses of 2.5, 3.6, 6, and 6.8 mm were manufactured by vacuum-assisted resin infusion (VARI) processing, which is elaborated in [33]. The four thicknesses are typical for thin and thick commercial CFPRs, respectively. We named the samples according to their thicknesses as t-x, where x denotes the magnitude of thickness, such as t-2.5 for the panel with a thickness of 2.5 mm.

2.2. Impact Event

The delamination in the CF/epoxy panels was formed by an Instron/9250HV impact tester with a hemispherical impactor (16 mm in diameter) according to the ASTMD7136/D7136-15 standard. The in-plane dimensions of samples were 100 mm in width × 100 mm in length, and the thicknesses were 2.5, 3.6, 6, and 6.8 mm, respectively. The incident energies were set to 12, 18, 28, and 32 J.

2.3. Burning Effect on CF Yarns and CFRTPs

To assess whether CF properties were affected at the temperature of fiber combustion, tensile tests were carried out on the pristine and heated CF yarns of a length of 14 cm, following the ASTM D2256 testing standard in Figure S5. A ZwickRoell universal testing machine was used to perform the test, and the loading speed was set as 2 mm/min. The CF yarns, containing approximately 3000 single fibers, were heated for different durations (0.5–4 min) at 770, 820, and 870 °C, respectively, under the air atmosphere in Figure S5. At such high temperatures, the epoxy resin was carbonized according to the thermogravimetric analysis (TGA) in Figure S2. The surface element and morphology of CF yarns before/after heating were characterized by X-ray photoelectron spectroscopy (Thermo Scientific K-Alpha+; Thermo Fisher Scientific Inc., Waltham, MA, USA), and scanning electron microscopy (SEM; Hitachi S-4300, Tokyo, Japan), respectively.

2.4. Dosage of Healing Agent Estimation

TECHSTORM 481 epoxy resin (Dawn Tianhe Materials Technology (Shanghai) Co., Ltd., Shanghai, China) was mixed with its curing agent at a mass ratio of 100:26. To maintain a constant fiber volume fraction V_f of the repair region, a precise control of added healing agent was required. Given the fact that the volume of liquid-like healing agent (V₁) is approximate to that of its solid phase (=V − V_f), the mass of the healing agent $m={\rho }_{solid}{V}_1\approx {\rho }_{liquid}{V}_1$ was obtained, where ${\rho }_{solid}$ and ${\rho }_{liquid}$ are the density of the healing agent in the solid and liquid states, respectively. The fiber volume fraction (V_f) of original and repaired panels was estimated according to the cross-section area ratio of fiber from their optical images in Figure S1 and listed in Table S1. The average V_f of four panels was (58.9 ± 0.4)%, which dropped to (56.1 ± 0.6)%, after repair.

2.5. In-Plane Compression after Impact (CAI) Tests

CAI tests, following the ASTM D7137 standard, were performed using an Instron universal testing machine (Instron 34TM-30, Norwood, MA, USA) to evaluate the repair efficiency [21,34] in Figure S5. The specimens with the in-plane dimension of 30 mm in width × 100 mm in length were cut from the repaired CF/epoxy panels. They were fixed at one end and compressed on the other end along the longitudinal direction. To prevent the occurrence of buckling in the panel, two metal fixtures were designed to assist in fixing two ends. The loading was controlled at a speed of 2 mm/min. The ultimate compressive strength σc was calculated by:

$${\sigma }_{\mathrm{c}}=\frac{F_{max}}{width \cdot thickness}$$

(1)

where F_max was the maximum force before failure.

The repair efficiency was evaluated by the recovery ratio and loss ratio as follows:

$$recovery ratio=\frac{F_{\mathrm{repaired}}}{F_{\mathrm{pristine}}}$$

(2)

$$loss ratio=\frac{F_{\mathrm{pristine}}-F_{\mathrm{delaminated}}}{F_{\mathrm{pristine}}}$$

(3)

where, F_repaired, F_delaminated, and F_pristine were the compressive peak loads of the repaired, delaminated, and pristine panels, respectively. Both the CAI and tensile tests for each sample were repeated three times, yielding the experimental error being within 10%.

3. Results and Discussions

3.1. In Situ Thermal Ablation Repair Approach

The typical delamination damage in the CF/epoxy panels was produced by impact events (Materials and Methods; Figure S3) and confirmed by the computed tomography (CT, YXLON FF35) images in Figure S4. The proposed repair method, the thermal ablation repair, is schematically shown in Figure 1. This repair process includes three steps: (a) damage removal by a combustion process; (b) filling, and (c) solidification of the healing agent. Firstly, the virgin resin in the delamination area was eliminated through a combustion process using a butane flamethrower, as shown in Figure 1a. The delaminated region was heated up to 800 °C, the periphery of which was covered by a perforated steel mold to prevent the fire from spreading to undamaged parts. To ensure the completed resin removal, the hole of perforated steel was slightly larger than that of the delaminated area and the combustion residuals were washed out with alcohol. After that, a selected healing agent (TECHSTORM 481) with a low viscosity [35] was filled in the delaminated region as shown in Figure 1b. The electric resistance heating method was employed to increase the temperature in the delaminated region to 80 °C for the purpose of the healing agent being solidified within 30 min. Experimentally, the semicircular notch-shaped copper foils were glued around the repair region at one or both surfaces with the conductive silver, as shown in Figure 1c. The copper foils were connected to a direct-current power, and a series of electric powers (10–24 W) was supplied to generate the resistance heating and hence control the curing temperature of the resin. To ensure that the repair region retained the same thickness and hence identical fiber volume fraction V_f as the pristine sample, a constant pressure of 0.4 MPa was applied at the top surface to expel the overfilled healing agent and to prevent volume expansion.

3.2. Burning Effect on the Mechanical Properties of CF Yarns and CFRTPs

Prior to the investigation of repair efficiency, we examined the impact of the combustion process on the mechanical performance of CF constituents. The CF yarns underwent the burning process with different temperatures and durations, and then were characterized by tensile testing as shown in Figure 2. The burned CF yarns exhibit the expected linear load–displacement relation in Figure 2b–d and broke at a certain strain. The maximum tensile load decreased with increasing burning duration, t, and burning temperature, T, as shown in Figure 2e. Specifically, the maximum tensile load decreased from 310 N, for the pristine sample, to 87.4 N, for the CF yarn that suffered a burning process at 870 °C for 4 min. Given this fact, we selected 800 °C as the burning temperature to remove resin in CFPRs. The sharp drop in the tensile load of the burned CF yarn may have resulted from a loss of adhesion between the CFs because they are loosely dispersed, as shown in Figure S6a. Besides the structural factor, the thermo-oxidative decomposition of CFs may occur at a high temperature, which can diminish the fiber thickness and hence weaken the mechanical properties. In our case, the diameter of burned CF was the same as the pristine one, confirmed by the SEM images in Figure 2a, while the grooves and ridges on the CF surface appeared due to the reduction in the pre-coated sizing agent in the burning process [36]. Accordingly, the surface element analysis in Figure 2f and Table S2 reveals that the carbon element increased by 10.41%, and oxygen and nitrogen elements decreased by 8.57 and 2.05%, respectively, which confirms the decomposition of the sizing agent. These observations suggest that the drop in the adhesion between CFs plays a dominant role in the tensile performance of CF yarn. Interestingly, as carbon fiber yarns are woven as a panel, the CF yarns are unexpectedly structurally robust after burning and without the presence of dispersity as seen for a single yarn. The woven CFs are tightly held together (Figure S6b) under the strain from probably CFs outside the burning part due to their structural integrity [37].

3.3. In Situ Thermal Ablation Repair of Delamination in CF/Epoxy Panel

The thermal ablation repair process started with the damage removal followed by matrix reconstruction. The former was achieved by a controlled burning process (Figure 1a), where only resin was carbonized, enabling the continuity of CFs remaining intact. The surface of burned CFs was barely covered with resin, which is confirmed by the SEM images in Figure 3a, which is in contrast to a large amount of epoxy resin distributed on the CFs extracted from the pristine sample (Figure 3a). Experimentally, the resin removal process takes several minutes depending on the damage size. The complete removal can be identified when the panel is burned through on the other side. In our case, the cylinder-like delamination regions had an identical diameter, therefore, the burn-through duration monotonically increased with the panel thickness as shown in Figure 3b. In detail, the burn-through duration increased from 1 to 9 min as the panel thickness increased from 2.5 to 12 mm, the relationship between which is well captured by the linear function with a slope of 0.74.

The thermosetting resin, TECHSTORM 481, was used as a healing agent given its low viscosity and thermal stability. Its curing temperature, T = 80 °C, was realized by means of the electric resistance heating method as shown in Figure 4a,c. Two copper foil arcs, acting as electrodes for generating heat, were placed surrounding the damaged region at the top surface (Figure 4a), the temperature of which, estimated by the thermal imager, is uniformly distributed in Figure 4c. The temperature at the center (T_c) of the damaged region increases with increasing electrothermal heating duration and supply powers in Figure 4d and Figure S7. For all samples, T_c was largely enhanced by 200–500% in the first 100 s of heating, and then its increase became smooth with T_c reaching a plateau as heating time > 300 s. The plateau temperature T_p was in the range of 60–160 °C, which increases with the increasing supply power and decreasing panel thickness in Figure 4f. For each thickness, T_p =80 °C, the curing temperature, was achieved by adjusting the magnitude of supply power, as listed in Table S3. For instance, for the t-2.5 sample, T = 80 °C was obtained when the supply power was 10.5 W. Along the thickness direction, the temperature distribution was evaluated by the relative temperature difference between the top and bottom surfaces, $\mathsf{\Delta }T=\frac{2 \left(T_{\mathrm{top}}-T_{\mathrm{bottom}}\right)}{T_{\mathrm{top}}+T_{\mathrm{bottom}}}$, as shown in Figure 4e, which exhibits a strong thickness dependence. As the thickness was below 3.6 mm (t-2.5 and t-3.6 samples), ΔT was well below 10%, the acceptable experimental error, while it largely increased to ~13% as the panel thickness reached 6 mm, due to probably the poor thermal conductivity along the thickness direction. To compensate the poor heat transport ability, we add an additional pair of electrodes at the bottom surface for heat generation in Figure 4b. This strategy significantly reduces ΔT from ~13% to ~2% for both t-6 and t-6.8 (Figure 4e), providing an efficient heat generation approach for samples with large thicknesses. Increasing curing temperature can significantly reduce the resin curing [38] and hence the whole repair process duration. Compared with the typical scarf and inject repair methods that take 2–5 days when performed at room temperature [21,38,39], the repair duration within one hour is relatively fast.

3.4. Repair Efficiency Analysis

Repair efficiency was evaluated by the sample’s CAI performances. The representative compressive load–displacement curves of CF/epoxy panels under four conditions (pristine, delamination, burning, and repair) are shown in Figure S8. All of the specimens demonstrate a monotonic increase in compressive load with displacement until the failure occurs. The corresponding compressive strengths of the CF/epoxy panels with different thicknesses are shown in Figure 5a. For the pristine panels, the compressive strength increased from 110.65 to 178.89 MPa as panel thickness increased from 2.5 to 6 mm, and a further increase of thickness to 6.8 mm led to a small decrease of compressive strength to 164.39 MPa. Forming a delamination with a diameter of 2.5 mm in the panels, the compressive strength decreased by 30–50%, while the complete removal of resin led to a further reduction in compressive strength, which was only ~30% of that of pristine panels. This was anticipated given the large porosity (~40%) was introduced when the resin is replaced by air in the damaged region. When the healing agent fully filled the pores and the matrix was reconstructed, the compressive strength of repaired panels recovered to the magnitude in the range of 102.25–162.34 MPa, close to that of the pristine state (110.65–178.89 MPa). The recovery ratio, qualified by F_repaired/F_pristine, reached 87% for all four thicknesses, as shown in Figure 5b, although the loss ratio, qualified by F_pristine−F_delaminated/F_pristine, spanned from 30% to 50%. To explore the stability of the repaired panels, we examined the compressive performance of the t-6.8 CF/epoxy panel under three successive delamination–repair cycles, as shown in Figure 5c. The compressive strength of the repaired panels decreased to 122.73 MPa after three cycles, while the recovery ratio still achieve 74.94% in the third repair process. The slight reduction in compressive strength may account for the weak interfacial adhesion between the reconstructed matrix and carbon fibers due to some char residue left on the surfaces as shown in Figure S9.

4. Conclusions

In summary, we developed a novel thermal ablation repair technique, which can repair the delamination of CFRTPs at a high efficiency, and without undermining the fiber continuity and structural integrity. We prepared the CF/epoxy panels via the VARI method and pre-damaged them in low-velocity impact tests to produce typical delamination. The optimal parameters of the thermal ablation approach were investigated and its repair efficiencies were evaluated by CAI tests. For the resin removal, a suitable temperature and complete resin removal are two important factors determining the repair efficiency. The char residue of resin can weaken the interfacial adhesion between the fiber and reconstructed matrix leading to a low recovery ratio. The compressive strength recovery ratio reached ~87% and slowly decreased to ~75% after three successive delamination–repair processes. The repair efficiency of this method is robust with the panel thickness and delamination, suggesting a universal feature of this thermal ablation repair approach. This approach is expected to apply for repairing large, complex, and irregular aerospace materials.