Experimental study on the influence of different laying methods on the thermomechanical properties of carbon fiber/epoxy hybrid composite

2.3.1.1. Cone calorimeter

2.3.1.2. High temperature oxygen index method

2.3.1.3. Vertical/horizontal combustion test

CZF-5 vertical/horizontal combustion tester was used to test the vertical/horizontal combustion rate of the test samples [22]. The effects of different laying methods on the propagation characteristics of carbon fiber/epoxy composites in different directions were analyzed. The experiment was carried out with the test samples in table 2. The size and specification of the test samples shall be made according to the standard. It is made into 150 mm × 10 mm. Each group was tested three times.

Figure 1 shows the sample geometry for the high temperature oxygen index experiment and the vertical/horizontal combustion experiment.

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The high temperature mechanical testing machine developed independently was used. Table 3 shows the performance parameters of the high temperature mechanical testing machine. The carbon fiber/epoxy bidirectional woven fabric (B1 and B3 in table 2), the fuselage frame of electric light aircraft (carbon fiber/epoxy bidirectional woven fabric) and the material of instrument panel (carbon fiber/epoxy prepreg) were tested by microcomputer control. The mechanical properties of carbon fiber/epoxy composites with different laying angles were tested at high temperature.

The experiment was carried out at 25 °C, 50 °C, 100 °C and 150 °C. Three experiments were carried out at each temperature point. The experimental results were averaged. Compared with the mechanical properties at room temperature [23]. The effect of temperature on the mechanical properties of carbon fiber/epoxy composite with the same thickness and different laying angle was analyzed. At the same time, the high temperature mechanical properties of fuselage frame materials and instrument panel materials used in electric light aircraft were compared. Figure 2 shows the equipment appearance.

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The fuselage frame material and instrument panel material (carbon fiber/epoxy prepreg) are numbered T1 and D1 for study and analysis. Table 4 shows the specific corresponding numbers. Figure 3 shows the geometry of the sample.

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Table 9 shows the average ignition time (TTI) of the test samples at different thermal radiation intensities.

Table 9 shows that the ignition time of the test samples is shortened with the increase of thermal radiation intensity under different thermal radiation. This is because the epoxy resin matrix in the carbon fiber/epoxy composite is thermally decomposed into a flammable volatile gas. It is ignited when the concentration of flammable gas reaches a certain value. The time for the pyrolysis reaction of the material is advanced as the heat radiation intensity increases. And the generation of the flammable volatile gas is accelerated. The ignition time of the material is shortened.

Figure 10 shows the relationship between the carbon fiber/epoxy composite laying and ignition time. Figure 10(a) shows that the thickness of the layer has a certain influence on the ignition time of the carbon fiber/epoxy composite. Under the same heat radiation intensity, the ignition time of A1 is the longest and A4 has the shortest ignition time under the same heat radiation. The ignition time of the materials increases with the increase of the laying thickness due to the increase of heat capacity. The time for the epoxy resin matrix to reach pyrolysis temperature is prolonged [25]. So the ignition time is delayed.

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Figure 10(b) and table 9 shows the ignition time of the test samples with different laying angles does not show regularity under the same thermal radiation intensity. The effect of different laying angle on the ignition time of carbon fiber/epoxy composite is not obvious.

Figure 11 shows the heat release rate curve of carbon fiber/epoxy composite with different laying methods under the heat radiation intensity of 35 kW·m⁻².

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Figure 11 shows the heat release rate curves of the experimental samples all show an obvious peak under the same heat radiation intensity, This is because the epoxy matrix in the material is heated to decompose combustible gas. It is ignited to release heat when the gas concentration reaches a certain value. The peak of heat release rate is reached after a period of time.

Table 10 shows the peak HRR of carbon fiber/epoxy composite with different laying methods and the time to reach the peak.

The peak HRR of the test samples increases with the increase of the thermal radiation intensity under different thermal radiation intensities. The time required to reach the peak is advanced. This is because higher thermal radiation intensity can provide energy for the pyrolysis of epoxy resin matrix faster. This makes the time of material pyrolysis earlier [25].

The peak HRR of material A1 is the largest and the time to reach the peak HHR is the longest at the same radiation intensity. The peak HRR of material A4 is the smallest and the time to reach the peak HHR is the shortest. The results show that the heat release rate of carbon fiber/epoxy composite is affected by the thickness of different layers. The peak HRR of carbon fiber/epoxy composite increases, and the time to reach the peak value increases with the increase of layer thickness.

The heat release rate of the test samples with same thickness and different laying angles did not show regularity. The results show that the effect of different laying angle on the heat release rate of carbon fiber/epoxy composite is not obvious.

Mass loss is an important parameter of fire behavior. The content and decomposition rate of the organic components of the composite can be determined in the process of fire by measuring the weight change of the samples [34].

Figure 12 shows the mass loss rate (MLR) of carbon fiber/epoxy composite with different laying methods at the thermal radiation intensity of 35 kW·m⁻². It can be seen that the mass loss of the experimental samples has gone through three obvious stages. Less volatile is released due to the radiation time of material is shorter and the temperature is low in stage I. The concentration of pyrolysis gas can not reach the combustion condition. The quality has not changed significantly. In stage II, the test sample enters the combustion stage. The epoxy matrix is pyrolyzed. The material will burn with flame due to the combustible volatile gas is ignited. The mass loss rate is accelerated obviously. In stage III, combustible materials are exhausted. The mass loss rate reaches the minimum value and remains stable all the time. Most of the combustible materials are consumed after combustion and pyrolysis at this time. The mass loss of the samples is mainly due to the pyrolysis of the epoxy matrix. This can be reflected in the peak value of mass loss rate.

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The combustible gases and volatiles of the test samples are more in unit time. The mass loss rate increases with the increase of radiation intensity. The greater the fire risk is greater in the actual fire field [35]. The time of Peak-MLR appears earlier with the increase of heat radiation flux. It corresponds to the peak time of heat release rate shown in table 10.

Figure 12 shows the effect of different laying angle on the mass loss rate of carbon fiber/epoxy composite is not obvious. In the experiment, the heat radiation flux is set to 35 kW·m⁻², corresponding to the exposure temperature of 716 °C. Figure 13 shows the photo image of fire damage of four test samples (B1–B4) with the same thickness and different laying angle after the cone calorimeter test. The four samples have the same residue after combustion. It can be seen that the four testl samples have obvious stratification. It shows that the epoxy matrix has been exhausted. Only reinforced carbon fiber is left. The oxidation reaction of carbon fiber will be limited when the temperature is lower than about 1000 °C [12]. The oxidation reaction of carbon fiber is limited and the mass loss of carbon fiber is small under this radiation intensity.

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There is a certain regularity between the laying thickness and the mass loss rate of carbon fiber/epoxy composite. Figures 12 and 14 show that the peak MLR and average mass loss rate of the samples with different laying thickness increase with the increase of laying thickness under the same thermal radiation intensity. This is mainly due to the greater mass and thermal capacity of the thick test samples

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The thermal radiation intensity is linearly related to Peak-MLR and Av-MLR in steady state (or quasi steady state) combustion. As shown in Formula (1) [36].

$$\begin{eqnarray}&&m^{\prime} =\displaystyle \frac{1}{L}q+C\end{eqnarray} \tag{ 1 }$$

Where m' represents the mass loss rate including Peak-MLR and average mass loss rate Av-MLR. L is the heat of vaporization. q is the thermal radiation intensity. C is a constant.

The L of the material is the reciprocal of the slope of the fitting line as shown in formula (1). Figure 15 shows the fitting function of mass loss rate and thermal radiation intensity of the test samples with different laying thickness. The influence of heat radiation flux on the Peak-MLR and Av-MLR of carbon fiber/epoxy composite with different laying thickness was analyzed.

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A1:

$$\begin{eqnarray}&&{m}_{peak}^{{\prime} }=0.00979q+0.25827,\,{R}^{2}=0.92826\end{eqnarray} \tag{ 2 }$$

$$\begin{eqnarray}&&{m}_{Av}^{{\prime} }=0.38643q+8.03578,\,{R}^{2}=0.98920\end{eqnarray} \tag{ 3 }$$

A2:

$$\begin{eqnarray}&&{m}_{peak}^{{\prime} }=0.00864q+0.16155,\,{R}^{2}=0.98125\end{eqnarray} \tag{ 4 }$$

$$\begin{eqnarray}&&{m}_{Av}^{{\prime} }=0.32373q+7.23621,\,{R}^{2}=0.98983\end{eqnarray} \tag{ 5 }$$

A3:

$$\begin{eqnarray}&&{m}_{peak}^{{\prime} }=0.00962q+0.04161,\,{R}^{2}=0.96579\end{eqnarray} \tag{ 6 }$$

$$\begin{eqnarray}&&{m}_{Av}^{{\prime} }=0.25868q+5.54088,\,{R}^{2}=0.99750\end{eqnarray} \tag{ 7 }$$

A4

$$\begin{eqnarray}&&{m}_{peak}^{{\prime} }=0.00833q+0.00626,\,{R}^{2}=0.94190\end{eqnarray} \tag{ 8 }$$

$$\begin{eqnarray}&&{m}_{Av}^{{\prime} }=0.2119q+4.94824,\,{R}^{2}=0.96167\end{eqnarray} \tag{ 9 }$$

where ${m}_{peak}^{{\prime} }$ and ${m}_{Av}^{{\prime} }$ represent the peak MLR and the average MLR, respectively.

The fitting function of each test sample adopts the formula with higher R². The formula can be used not only to predict the mass loss rate of experimental samples under different heat radiation fluxes, but also to determine the vaporization heat of materials. The vaporization heat of A1, A2, A3 and A4 were calculated as 2.588 MJ·kg⁻¹, 3.089 MJ·kg⁻¹, 3.866 MJ·kg⁻¹, and 4.713 MJ·kg⁻¹, respectively. Vaporization heat L is a solid property in engineering calculation [37]. It can be used to characterize the combustion rate under a specific heat flow. It can be seen that the higher the thickness of carbon fiber/epoxy composite, the smaller the vaporization heat through calculation. This shows that the smaller the thickness of carbon fiber/epoxy composite with the same laying angle, the faster the burning rate. This is consistent with the results of vertical/horizontal combustion rate experiments.

Figures 16(a) to (d) show the stress-strain curves of four samples at 25 °C, 50 °C, 100 °C and 150 °C. The trend of the stress-strain curves of the four samples is similar under the four temperature environments. In the initial stage, the stress and strain also increase with the increase of tensile displacement and decrease rapidly when it reach the peak value.

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Figures 16(a) to (d) show the maximum stress values of B1, B3, T1 and D1 at 25 °C are basically the same as those at 50 °C. This is because the epoxy resin as the matrix has not started to decompose at 50 °C. The temperature has little effect on its mechanical properties. The maximum stress values of the four samples at 100 °C is slightly lower than that at 25 °C. This is because the epoxy begins to decompose at 100 °C [40]. However, the whole mechanical properties of the experimental samples have not been greatly affected. The maximum stress of the sample at 150 °C is lower than that at 25 °C. This is due to the poor heat resistance of epoxy [41]. The epoxy resin in each layer of the sample has been decomposed at 150 °C. The trend of the stress-strain curves of the four samples is similar under the four temperature environments. At the beginning, the stress and strain increase with the increase of tensile displacement. It drops rapidly when the peak value is reached.

The tensile failure strength is the maximum stress of the test sample from the beginning of tensile test to the time of fracture failure [42]. It is usually used to characterize the strength and toughness of materials. The tensile modulus of elasticity refers to the stress required to produce unit elastic deformation under the action of external force [43]. It is used to measure the difficulty of elastic deformation of materials. Figures 17 and 18 show the relationship between the tensile failure strength and the tensile modulus of the test samples with the change of temperature.

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The tensile failure strength and modulus of elasticity of B1 are 220.04 Mpa and 4187.21 Mpa at 25 °C, respectively. At 100 °C, the tensile failure strength and elastic modulus are 216.76 Mpa and 3853.1 Mpa at 100 °C, respectively. Compared with 25 °C, the tensile failure strength and elastic modulus decreased by 1.49% and 7.98% respectively. The tensile failure strength and elastic modulus of B3 decreased by 5.79% and 11.40%, respectively. The tensile failure strength and elastic modulus of T1 decreased by 11.18% and 13.72%, respectively. The tensile failure strength and elastic modulus of D1 decreased by 6.71% and 9.76%, respectively. It can be seen that the decrease is not large. This is because the epoxy resin just began to decompose at 100 °C. The mechanical properties of the four samples are basically stable within the temperature range of 25 °C–100 °C. The mechanical properties are not affected by temperature.

The tensile strength and elastic modulus of the test samples decreased sharply due to the gradual decomposition of the matrix epoxy resin from 100 °C to 150 °C. The tensile failure strength and elastic modulus of B1 at 150 °C are 171.82Mpa and 1936.67Mpa. Compared with 25 °C, it decreased by 21.91% and 53.75%, respectively. The tensile failure strength and elastic modulus of B3 decreased by 47.42% and 67.90%, respectively. The tensile failure strength and elastic modulus of T1 decreased by 27.12% and 50.80%, respectively. The tensile failure strength and elastic modulus of D1 decreased by 16.15% and 24.30%, respectively. The mechanical properties of the samples at 150 °C are greatly affected by the temperature.

Figures 17 and 18 show that the tensile failure strength and tensile modulus of the experimental samples decrease with the increase of temperature. The mechanical properties of carbon fiber/epoxy bi-directional woven fabric (B1, B3, T1) are better than that of carbon fiber unidirectional prepreg (D1) at 25 °C, 50 °C, 100 °C and 150 °C. The tensile mechanical properties of B3 are better than those of B1 and T1 among three kinds of carbon fiber/epoxy bi-directional woven fabrics. This is because B3 is laid in [(0°/90°)₆/(±45°)₂]. There are 8 layers in total. There are 6 layers of fiber along the direction of stress. The laying method of B1 is [(0°/90°)/(±45°)]₄. There are 8 layers in total. There are four layers of fiber along the direction of stress. The laying method of T1 is [(0°/90°)/±45°]₂ [±45 °/(0°/90°)]₂. There are 8 layers in total. There are four layers of fiber along the direction of stress. The laying mode of D1 is [(±45°)/0°/45°/0°/–45°/0°]. There are six floors in total. There are three layers of fiber along the direction of stress. The fiber can bear more load along the direction of stress. The number of layers of B3 fiber along the stress direction is more than that of the other three test samples. B3 can bear large load. The load-bearing performance of B1, T1 and D1 decreased in turn.

The mechanical properties of carbon fiber/epoxy composites were studied at high temperature. Regression analysis of the relationship between tensile strength and temperature was carried out. The formula of the relationship between the tensile failure strength and temperature of four test samples is obtained. As shown in table 11.

Goodness of fit:

For nonlinear equations:

• (1)  
  Calculate the sum of squared residuals Q = ∑(yy*)² and ∑y², where y represents the measured value and y* represents the predicted value;
• (2)  
  Find the goodness of fit R²  =  1 − (Q/∑y²)^1/2;

For the linear equation: R² = ∑(y* − y_average)²/∑(y − y_average)²;

The goodness of fit of four regression equations was calculated. Select the equation with the best goodness of fit as the fitting function. Table 12 shows the goodness of fit and the selected fit function for each sample.

Figures 19(a) to (d) show the curve of fitting function and test data.

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Figure 19 shows the fitting function has a good agreement with the measured data of carbon fiber/epoxy composite in the range of 25 °C to 150 °C. It can reflect the change rule of high temperature mechanical properties of carbon fiber/epoxy composite with temperature in the temperature range.