Experimental and Numerical Study of Tool Effect on Curing Deformation for the Composite Missile Structure

Transcription

1 Experimental and Numerical Study of Tool Effect on Curing Deformation for the Composite Missile Structure Experimental and Numerical Study of Tool Effect on Curing Deformation for the Composite Missile Structure Ye Jinrui 1*, Yue Guangquan 2, Zhang Boming 1, Lu Shan 3, and Cheng Kai 2 1 School of Materials Science and Engineering, Beihang University, China 2 School of Astronautics, Harbin Institute of Technology, China 3 Research Institute of Special Aerospace Materials & Technology, China SUMMARY Composite missile structure is complicated in shape and fabricated by soft-film-assisted Resin Transfer Molding (RTM). The cylindrical tool of composite missile structure was studied in the paper. A completed method of analyzing and estimating the tool deformation was presented. The analysis models and three-dimensional finite element method (FEM) for cure process of composite integrated structure were developed. Thermal deformation and cure shrink were considered in analysis models and the evolution of residual stresses in the entire cure process was studied. The strain and temperature fields in the cylinder-shaped mold of composite missile structure were measured respectively using optical fiber grating and thermocouple. The experimental results show a good agreement with the finite element analysis. The deformation of the tool is added to the composite missile simulation model as a displacement loading to analyze the effects of the tool s curing deformation on the part. The results are valuable for the tool design, tool deformation control, and elimination for the composite missile structure. 1. INTRODUCTION Composite material is characterized by high performance such as specific strength and stiffness, and has become one of the major aeronautic materials. Composite missile structure is applied widely in the aerospace field. The structure is complex, with cylindrical shape, inner stiffeners, and a hatch on one side of the centerline. Soft-filmassisted RTM is adopted to manufacture the complex structure. In the molding process, the curing process of the temperature distribution is essentially heat conduction with nonlinear heat source. Exothermic reaction can lead to higher local temperature of composite materials, which form a complex temperature gradient. The result will lead to non-uniform curing, which can lead to the development of residual stress and then structural deformation. The residual stresses are affected by many factors, such as structural Smithers Rapra Technology, 2011 non-symmetry, layer and process control, etc. One of the main factors is that mold and composite thermal expansion coefficient does not match. Residual stresses that develop in fiberreinforced composite parts that undergo RTM processing while confined to the process tool often lead to dimensional changes such as warpage and springin. There are several phenomena that are responsible for these dimensional changes including mechanical toolpart interaction, through-thickness gradients in the degree of cure, and anisotropic thermo-chemical strains. There are many researches on these factors 1-5, such Fernlund 6,7 studies resin cure shrinkage and anisotropic thermal expansion cause process induced residual stresses in polymer composites. Nuri E. and Kevin P. 8 consider that composite parts manufactured by autoclave processes * Corresponding author: Tel.:(86-010) ; address: yejinrui_1982@126.com(ye Jinrui) have manufacturing distortions due to unavoidable physical mechanisms that take place during the process. KIM and Daniel 9 applied a more direct method to measure the strain in the cure process. They used fiber Bragg grating and resistance strain chip buried inside the composite materials to measure the changing of the strain. It is found that the reciprocity between mold and part leads to evident strain in the cure process. Satish 10 applied shear layer to replace the influence of mold on the composite material by simulation. The method can accurately predict the curing deformation generated by the reciprocity between mold and part. Thus, the cylindrical tool of composite missile structure is studied in the paper. A completed method of analyzing and estimating the tool deformation is presented. The linear finite element thermal-structural analysis model is established. The strain and temperature fields in the cylinder-shaped mold of composite missile structure are measured respectively using optical fiber grating and thermocouple. The experimental results show a good Polymers & Polymer Composites, Vol. 19, Nos. 4 & 5,

2 Ye Jinrui, Yue Guangquan, Zhang Boming, Lu Shan, and Cheng Kai agreement with the finite element analysis. The deformation of the tool is added to the composite missile simulation model as a displacement loading to analyze the effects of the tool s curing deformation on the part. 2. SOFT-FILM-ASSISTED RTM Soft-film-assisted RTM,a shaping method by expansion of the mold, refers to injecting resin through RTM system to mold cavity and curing, in which the fabric preform is compressed by thermal expression male die. Its pressurized method is different from general composite molding technology. Curing pressure of general composites forming process is from external pressure source which needs pressing equipment or vacuum equipment,such as vacuum suction, hot pot and mold, etc. However, pressure of the soft-film-assisted RTM is put on fiber perform by thermal expansion of the core model which has a great linear expansion coefficient. This method does not need external pressure source and is suitable for integral shaping of complex products, especially for the shaping of composite structure which has a complex cavity. It can overcome such disadvantages as the difficulty of delivering external pressure uniformly and demoulding hard or unable to. According to products structure, soft-film-assisted RTM needs two molds: rigid female die and flexible male mold. Dimensions of female mold s cavity are overall dimensions of composite structure; the form of core mold is the same as that of composite structure. All of dimensions are smaller than true sizes of components. The decrease is called technology gap. It is need to use this gap to control pressure points and the size of pressure. In the composite molding, firstly we capped core mold with dry reinforcement to form composite preform, and then closed cavity block, using resin transfer molding or other methods to press resin into the mold and cured. 3. INFLUENCE OF MOLD ON CURING DEFORMATION The composite missile structure can be simplified as a cylindrical structure, containing a number of horizontal and vertical white stiffeners. There are 7 windows at the side near the central axis, as shown in Figure 1. In making the airplane fuselage with composite materials, the progress of compound and molding are finished at the same time with a designed mold. In the progress of molding, the different thermal expansion coefficient between metal mold and the structure, the usage of soft mold, the deformation of the frame mold and the temperature distribution are the important factors that lead the deformation of structure curing. Between of those factors, the effect of soft mold is complex, so we regard it as part of the structure. Therefore, only metal mold is considered, which have a higher coefficient of linear expansion than the structure. Before the curing of the resin, there will be a stress gradient distribution in the through-thickness direction, which is because in the direction the stress transfer is poor; the fibers nearby the mold are under the tension stress; the shear modulus of the Figure 1. The simplified composite missile structure composite structure is low at that time. As the progress advancing, the residual stress will be cured in the structure. After the progress, the relieving of the residual stress caused the deformation of structure. Furthermore, the thermal conductivity and the figure of the mold directly effect the temperature distribution of the structure, which will affect the size and distribution of the residual stress at last. In the curing of resin-based composites, the different thermal expansion coefficient caused the interaction between metal mold and the structure. The structure adheres on the mold surface under the curing press, then as the transform of the temperature, there will be shearing strength between them, that causes the inside residual stress. On the gel point and the glass-transition temperature T g, the progress will be separate into three states: viscous liquid I; viscoelastic solid II; viscoelastic solid III (as shown in Figure 2, the curing cure of bismaleimide-triazine resin 5428). In the state of viscous liquid, there is no or only little interaction between them; however, in the state of viscoelastic solid II, there will be tensile stress inside the structure because of the shear modulus and the shear stress. But the stress distributes equably and there not be a stress gradient 260 Polymers & Polymer Composites, Vol. 19, Nos. 4 & 5, 2011

3 Experimental and Numerical Study of Tool Effect on Curing Deformation for the Composite Missile Structure distribution in the through-thickness direction. After demoulding, the structure reinstates without causing warpage. The interaction that causes the deformation of structure curing happens in the state of viscoelastic solid III. The shear stress caused by the interaction transfers in the layers, but in that state the shear modulus and the stress transfer of the composite structure is low and the farther from the mold surface the smaller the shear stress is, so there will be a stress gradient distribution in the through-thickness direction. As the status change of resin (the resin transfer from viscoelastic solid to elastic solid when the glass- transition temperature T g is higher than curing temperature), the composite structure is formed and the gradient of residual stress is left in it. After the progress of demoulding, the residual stress will release and cause the deformation of structure. So this paper mainly analyses the interaction caused by different thermal expansion coefficient between metal mold and the structure in the state of viscoelastic solid. 4. STRUCTURAL MODEL OF FINITE ELEMENT ANALYSIS The strain and temperature field of points on the approximate cylindrical mold which is used to produce composite missile structure is measured by optical fiber grating and thermocouple technology. Composite missile structure is loaded with the deformation of mold as displacement loads. The Influence of mold deformation on curing deformation of the missile structure is simulated. The model was established by finite element analysis software ANSYS. The unit has capabilities such as plastic, super elastic, stress stiffening, creep, large deformation and large strain, etc. The unit can also simulate deformation of the little or no Compressible super-elastic material by using mixed formula. Figure 2. The standard curing temperature, curing degree and T g of bismaleimide 5428 resin Both ends are fixed constraints, and the temperature is raised from 20 C to 200 C. The maximum displacement of mold deformation is mm. Obtain the displacements of mold and part, which are loaded in to the part, and get the final deformation. The final deformation is shown in Figure 3, and the maximum displacement is mm. The measured strain of unidirectional laminate is considered as thermoshrinking strain. The model contains strains generated by cure shrinkage and thermo- deformation. In finite element analysis, all strains are analyzed for thermo-stress problem. And the other mechanical properties of composite materials are set as that the resin is completed cured. The strains in 0 and 90 directions of unidirectional plates e x and e y are measured by experiments. They are expressed by: ε x = ε x 0 ν yx ε y 0 ε y = ε y 0 ν xy ε x 0 (1) (2) where n is the poison ratio of composite unidirectional plates. The curing shrinkage is equivalent to the thermal shrink, thenε x 0 andε y 0 can be expressed by: ε x 0 = a x ΔT ε y 0 = a y ΔT (3) (4) where, DT is the changing of temperature, a x and a y are the equivalent coefficients of thermal expansion. Assign DT = 1, from equations (1)-(4) the equivalent coefficients of thermal expansion a x and a y can be obtained. Then set the a x and a y of materials in the finite element analysis, and the load is DT = EXPERIMENTS Experiment A: select cylindrical mold and heat it inside an oven. Arrange optical fiber gratings at the generatrix and radial direction of mold. Thermocouple is arranged at each optical fiber grating to measure the distribution of temperature. Experiment B: overlay unidirectional T700/9368 on the mold, form the Polymers & Polymer Composites, Vol. 19, Nos. 4 & 5,

4 Ye Jinrui, Yue Guangquan, Zhang Boming, Lu Shan, and Cheng Kai composite part, and measure the deformation. It is assumed that the temperature inside oven distributes even. After measuring the strain, the deformation of mold is obtained. If the measurements agree with the simulation results under the same conditions, the deformation results are considered as the displacement loads added on the missile structure to simulate the influence of mold-deformation on the curing deformation of missile structure. At last, contrast the experimental results of A and B. Figure 3. The deformation generated by mold When the temperature inside oven distributes uneven, apply each measured temperature by thermocouple as load into numerical simulation. If the strain by simulation agrees with the measured result by optical fiber grating, the simulation method is valid and the deformation obtained by simulation will be loaded on the composite structure to simulate the influence of mold deformation on the curing deformation of the missile structure. Contrast the experimental results of A and B. The purpose of these experiments is to analyze the strain changing at different locations of mold surface. Measure the surface temperature of mold and analyze the distribution of temperature field. In the experiments, we select cylindrical curing mold, 8 optical fiber gratings, and 32 thermocouples. The optical fiber gratings and thermocouples are even arranged at the axial and radial directions of the cylindrical mold. Select bismaleimide/ carbon fiber with size of 300 mm 300 mm 2 mm, and overlay is [0] 16. Arrange three monitoring points at 0 and 90 directions, as shown in Figure 4, in order to observe whether the strain distributes even at the surface. Also arrange three monitoring points along 0 and 90 thickness direction. There are total 8 optical fiber gratings and 12 monitoring points. The changing strain at different thickness positions along 0 and 90 directions of unidirectional laminate is shown in Figures 5 and 6, which illustrate the curing process. It is illustrated that the strain of X1 which nears to the curing mold is maximum, and the strains of X2 (1667 με), X3 (1627 με) which are away mold are similar and smaller than that of X1. The result is generated by the reciprocity between mold and composite part during curing process, and the stress gradient is formed along the thickness direction. The strains at different Figure 4. Positions of measurement points for FBG 262 Polymers & Polymer Composites, Vol. 19, Nos. 4 & 5, 2011

5 Experimental and Numerical Study of Tool Effect on Curing Deformation for the Composite Missile Structure thickness positions which are vertical to the fiber direction are similar with each other. Because the composite part contains mainly resin at 90 direction and the reciprocity between mold and composite part influences the strain little. Figure 5. The strain and fitting curve in positions with different thickness in parallel fiber orientation of unidirectional laminate Thus, when we measure the strains at 0 and 90 directions of unidirectional laminate in curing process, the monitoring point should be arranged away the curing mold, where the influence of reciprocity between mold and composite part is little. When the thickness of part is large, we can neglect this influence and obtain the strain of unidirectional laminate at the free state. Select the same material and experimental method, and the size of part with [0/45/90] 2 overlay is changed to 100 mm 100 mm 0.75 mm. The strain at X3 position along fiber direction is 1283 με. The strain at YP2 position vertical to fiber direction is με after lowering the temperature. By Eqs. (1)-(4) the equivalent coefficient of thermal expansion is calculated, a x = 1216 x 10-6 and a x = 2835 x By finite element analysis it is obtained that the maximum warp deformation is 18.8 mm. The experimental measured result is 21.5 mm. Figure 7 illustrates the warp deformation by simulations and experiments, which agree with each other. Figure 6. The evolutions of strains in vertical fiber orientation of unidirectional laminate Figure 7. The simulation and experiment results 6. CONCLUSIONS The paper analyzes the influence of mold on the curing deformation of composite structure. Take the simplified cylindrical missile structure fabricated by soft-film- assisted RTM into consideration, the linear finite element thermal-structural model is produced. A completed method of analyzing and estimating the tool deformation has been presented and verified by experiments. We apply optical fiber grating and thermocouple to measure the strain and temperature of the cylindrical mold of the composite Polymers & Polymer Composites, Vol. 19, Nos. 4 & 5,

6 Ye Jinrui, Yue Guangquan, Zhang Boming, Lu Shan, and Cheng Kai missile structure. The experimental results show a good agreement with simulation results by the finite element analysis. The research results are valuable for design of composite missile structural mold, controlling the deformation degree of mold or removing its influence on the structural curing deformation. REFERENCES 1. Radford D.W. and Rennick T.S., Journal of Reinforced Plastics and Composites, 19 (2000) Potter K.D., Campbell M., and Langer C., Composites: Part A, 36 (2005) Twigg G., Poursartip A., and Fernlund G., Composite Science and Technology, 63 (2003) Twigg G., Poursartip A., and Fernlund G., Composites: Part A, 35 (2004) Twigg G., Poursartip A., and Fernlund G., Composites: Part A, 35 (2004) Fernlund G., Rahman N., and Courdji R., Composites: Part A, 33 (2002) Johnston A., Hubert P., and Fernlund G., Science and Engineering of Composite Materials, 5 (1996) Nuri E. and Kevin P., Composites: Part A, 36 (2005) Kim Y.K. and Daniel I.M., Journal of Composite Materials, 36 (2002) Satish K.B. and Lloyd V.S., Composites: Part A, 36 (2005) Polymers & Polymer Composites, Vol. 19, Nos. 4 & 5, 2011