SELF-HEALING OF COMPOSITE STRUCTURES IN A SPACE ENVIRONMENT R S Trask1, I P Bond1 and C O A Semprimoschnig2 1
Department of Aerospace Engineering, University of Bristol, Queen's Building, University Walk, Bristol. BS8 1TR. U.K,
[email protected] 2 European Space Agency, ESTEC, Materials Physics & Chemistry Section PO BOX 299, Keplerlaan 1, NL - 2200 AG Noordwijk, The Netherlands.
[email protected] ABSTRACT The use of functional repair components stored inside hollow reinforcing fibres is being considered as a selfrepair system for future composite structures. This paper considers the problem of introducing a liquid healing resin, contained within hollow glass storage vessels, within a space environment. The problem of resin outgassing of a commercially available 2-part epoxy resin system and the thermal cycling of the glass storage vessels, and their sealing caps, are discussed. The mechanical property assessment of the baseline hollow fibre laminate, the damaged hollow fibre laminate and the healed hollow fibre laminate is discussed revealed that a self-healed laminate had a residual strength of 87% compared to an undamaged baseline laminate and 100% compared to an undamaged self-healing laminate. This study provides clear evidence that a FRP laminate containing hollow fibre layers can successfully self-heal. 1. INTRODUCTION Terrestrial composite structures can be damaged in many ways, e.g. unexpected mechanical overload, fatigue loading at stress concentrations, heat/chemical attack or impact damage arising from collisions between moving objects. Foreign object impact to single skin laminates can result in a drastic reduction in composite strength, elastic moduli, and structural durability and damage tolerance characteristics. Composite structures operating in the space environment are vulnerable to impact damage resulting from collisions with micrometeoroids and orbital debris. The relative velocity of man-made debris ranges from zero, for objects in the same orbit, to approximately 11 km/s for objects in retrograde orbits [1]. In comparison, collisions with meteoroid particles take place with an average velocity of 19km/s [2]. The damage to the brittle composite structures consists of penetration holes with adjacent surface damage and some internal ply delamination. The internal damage may be anisotropic, following the structure of the fibres. For complete penetrations the rear surface damage area (surface spallation) is frequently larger than the entry hole (ratios of 5:1 for spallation damage-
to-hole size are often observed for carbon/epoxy laminates) [3]. Space composite laminates can also suffer from internal matrix microcracking. This type of damage can occur within the lifetime of the composite structure due to its exposure to the repeated thermal cycling within the space environment. The magnitude of the microcracks will be dependent upon certain laminate and environmental properties (i.e. mission orbit, thermal properties, fibre and resin properties, number of cycles, stacking sequence, temperature differential, bond strength, stress/strain at micro/macro levels and matrix morphology). Any form of self-healing system contained within a space composite laminate must be able to repair damage resulting from collisions with micrometeoroids and orbital debris or damage arising from thermal cycling.
2. SELF-HEALING REPAIR OF SPACE COMPOSITE STRUCTURES The use of functional healing components stored inside composite materials to restore physical properties after damage has been advocated since the 1990’s [4]. Incorporation of hollow glass fibres within a structural composite laminate offers the advantage of being able to store additional functional components for composite self-repair systems [5]. To extend the existing terrestrial self-healing concept for composite structures operating in space, certain aspects of the environment, the nature of the damage threat, the location of the hollow filaments, and the choice of repair resin have to be considered before a technical solution can be derived. In this research programme the self-healing is defined as autonomous, i.e. all the healing materials are contained within the composite and no external repair is required to initiate the repair. Alternative repair methods could be active autonomous healing which requires the replenishment of the healing medium or non-autonomous self-healing which requires some external impetus to initiate the repair.
In the self-healing approach adopted by Bristol University, hollow glass fibres are used because they offer the advantage of being able to store additional functional components for composite self-repair systems as well as acting as reinforcement. This approach has been investigated by a number of workers [6-10]. A typical hollow fibre self-healing approach used within composite laminates could take the form of fibres containing a one-part resin system, a two-part resin and hardener system or a resin system with a catalyst or hardener contained within the matrix material. The approach at Bristol has been illustrated in Figure 1. This figure shows (top left) the 30µm hollow fibres, UV dye emitting from a fracture plane in a hollow fibre composite (bottom left) and the healing sequence of 2-part epoxy resin with UV dye using time lapse photography (right).
In this research project an epoxy resin system (Cytec’s Cycom 823) was selected as the healing resin. This system was chosen for a number of reasons. These include the need to match the composite baseline laminate (epoxy 913 with E-glass reinforcement), the two part nature of the system permitting its inclusion in separate storage vessels, the viscosity drops to a minimum within the operating temperature profile, and the mixed system reaches gelation within 30 minutes. Furthermore, it was observed experimentally that the individual elements of the 2-part system were sufficiently robust to survive the composite laminate curing processing route (120°C for 1 hour) whilst housed at their respective interfaces within the fibre stack of the composite laminate. Although a carbon fibre or glass fibre composite laminate could have been selected as the baseline material it was decided to select a glass fibre laminate since it permits better visualisation of the damage when viewed with transmission light. To enhance the visualisation UV dye can also be used. Finally a comment concerning the damage introduction method should be made. Although the damage event within the space environment is likely to be initiated by either a high velocity impact or thermal cracking it was decided that these two approaches were beyond the scope of this project. Instead, a static three point bend test fixture was used to initiate the damage (shear cracks and delaminations) within the composite laminates. Although the damage would not be representative of the damage observed in the space environment it would permit a repeatable evaluation of the self-healing approach. 3. VERIFICATION OF THE SELF-HEALING APPROACH IN A SPACE ENVIRONMENT
Fig. 1: Healing process using hollow fibres; Top left 30µm hollow fibres, Bottom Left - UV dye emitting from a fracture plane in a hollow fibre composite; and, Right - the healing sequence of 2-part epoxy resin with UV dye using time lapse photography (right). The exact nature of the self-healing method will depend upon (1) the nature and location of the damage, (2) the potential self-healing repair resins, and (3) the influence of the operational environment. The selfhealing fibres could be introduced within the composite laminate as additional plies at each interface, at critical interfaces or as individual filaments spaced at predetermined distances within each ply. In this research project, the self-healing filaments were introduced at four critical ply interfaces were identified within the same project and reported previously in [11].
The space environment significantly influences the selection and performance of the different candidate resin systems. In the high vacuum of space, composite materials desorb moisture, which can lead to large dimensional changes in the structure. Alternatively, the high vacuum environment of space can initiate the mass loss of polymeric materials due to outgassing of low molecular weight components. All composite materials are assessed to determine the level of outgassing with the aim of minimising this risk to avoid the threat of contamination to mission critical spacecraft components (e.g. optics). Clearly, a liquid resin system exposed to high vacuum is likely to have a high outgassing rate. However, in the case of the current self-healing approach, the only time the liquid resin and hardener will come into contact with the high vacuum is when the structure has been damaged. Furthermore, when the resin comes into contact with the hardener the curing process will be initiated and therefore the rate of outgassing should decrease as the
liquid transforms to a solid. To evaluate what takes place under these circumstances an outgassing experiment investigating the resin, hardener and the mixed resin and hardener was devised.
(a)
4. OUTGASSING TEST EVALUATION The standard screening test method employed by ESA to determine the outgassing performance of candidate space materials is given in document ECSS-Q-70-02A [12]. This document describes a thermal vacuum test to determine the outgassing properties of solid materials proposed for use in the fabrication of spacecraft and associated equipment. The test defines the outgassing properties of materials at 125°C under 10-3 Pa for 24 hours. This is determined by measuring the sample before and after the test to ascertain the recovered mass loss (RML) (i.e. total mass loss of the specimen without the absorbed water), and the collected volatile condensable material (CVCM) collected on a condenser plate in close proximity to the sample. The collector plate was kept at a constant 25°C during the experiment. The outgassing test method (as specified in ECSS-Q70-02A) tends to be conducted on solid materials with very low outgassing rates. However, since a two-part liquid system (with potentially high outgassing rate) was going to be assessed it was decided to undertake some measures to avoid excessive contamination of the highly sensitive test equipment. A sample of material was placed in a bespoke, single Knudsen cell, which was placed in a small vacuum system equipped with a cold trap and evacuated by a turbo-molecular pump. In addition to capturing the mass loss, a mass spectrometer was used to investigate the nature of the outgassing products as the temperature of the Knudsen cell was slowly increased to 125°C. The experimental arrangement is indicated in Fig. 2.
(b)
(c) Fig. 2: Outgassing test set-up: (a) Al-foil crucible containing the mixed resin system; (b) Knudsen cell with heater and thermocouples housing the Al-foil crucible; (c) Vacuum chamber housing Knudsen cell connected to mass spectrometer. Prior to the commencement of the outgassing experiment the Cycom 823 liquid resin and hardener where mixed at the recommended 4:1 ratio. At room temperature the mixed sample has a viscosity of 250cP and will not reach full cure for 6 days. A measured amount of mixed resin is then poured into a small Aluminium-foil crucible (see Fig. 2a). The crucible is then placed in a small Knudsen cell, provided with a heater and two thermocouples (see Fig. 2b). The Knudsen cell is contained within a small, vacuum chamber, which was evacuated by a two-stage, rotary vane, mechanical pump. This provided a typical pressure throughout the test of between 2x10-2 and 6x10-3 millibar. From the low-vacuum stream, a sample of the gas was removed via a needle valve into a highvacuum system, pumped by a turbo-molecular pump. The high-vacuum system operated at typically not more than 2x10-5 throughout the testing. This system was equipped with a high-vacuum gauge and a mass spectrometer which was used to monitor the volatile components from the samples throughout the test. Four tests were undertaken to understand how the liquid resin and hardener were affected by temperature and vacuum. The first two tests considered the liquid resin and hardener whilst the third and fourth tests considered two different heating rates (at 3.33°C/min and 1.67°C/min) for a 4:1 resin-to-hardener mix ratio. The test results for the fast ramp rate have given below.
Partial pressure 10^-7 to 10^-6 Torr
Temperature [C] 25
26
35
41
48
56
64
70
76
92 100 106 113 119 125
0
3
5
7
9
11
12
14
16
20
10 9 8 7 6 5 4 3 2 1 0 22
23
25
27
31
Time [minutes] Hydrogen
Methane
H-carbons
CO2
45
Fig. 3: Recorded partial pressures for Cycom 823 mixed resin system at fast ramp rate (RT to 125°C in 30 minutes) for specific elements and mass index numbers The experimental data indicated in Fig. 3 is a measure of the rate of ‘outgassing’ of the different chemical elements as a function of temperature within the partial pressure range of 10-7 to 10-6 Torr (a range chosen specifically to capture the outgassing of the high mass index elements). In general, the recorded spectra size of the individual mass index number (MIN) in the fast temperature ramp is higher than the corresponding MIN in the slow ramp test, i.e. the accelerated rate has induced rapid outgassing. The fast ramp rate results indicated in Fig. 3 show a rapid increase in the formation of hydrocarbon products from 76°C to 113°C before the rate of outgassing returns to the baseline level once the material has attained 125°C. The reduction in the measured outgassing products is due to cure of the healing resin (typically 15 minutes after attaining 95°C – according to manufacturer’s data). In addition to the experiments to determine the outgassing behaviour of a two-part epoxy RML measurements were taken according to the test guidelines in ECSS-Q-70-02A. In this standard, the general acceptance limits for outgassing materials (not manufactured for optical devices, or in their vicinity) for RML is < 1% and for CVCM is < 0.1%. The mass loss measurements for the Cycom 823 system are detailed in Table 1.
Table 1: Mass loss measurements for Cycom 2-part epoxy liquid resin ramped to 125°C within 60 minutes and held for 24 hours Outgassing property description Total specimen mass before test Total specimen mass after test Mass of material before test Final mass of collector plates after test Initial mass of collector plates before test
TML % =
Wo − W f
CVCM % =
Wm
Mg 473.98 409.57 448.10
Wg
61.24
Wp
60.53
× 100
Wg − W p Wm
Wo Wf Wm
(1)
(2)
The TML of the liquid specimen over 24 hours was 14.4% as calculated by Equation (1). The CVCM of the two-part epoxy system over 24 hours was 0.16% as calculated by Equation (2). As expected, the outgassing performance of the liquid epoxy resin system falls outside the existing ESA guidelines, i.e. guidelines generated for cured materials and not for liquids. In the context of producing a self-
healing material, the Cycom 823 two-part resin system was selected because it has a very low viscosity and resin gelation occurs within 15 minutes at 90°C. For a standard commercially available epoxy resin this system provides the best healing attributes within the operating environment. It has to be stated that this test is the worst case for a cracked hollow fibre. An undamaged and sealed fibre shall be mostly impermeable to outgassing. As in a full composite the amount and mass of uncured resin in the hollow fibre will be only a fraction of the total mass of a composite the derived values have to be seen in a perspective and therefore the obtained results are overestimating the outgassing of a full composite. Furthermore, if, as the outgassing tests suggest, the rate of cure can happen fast enough then the rate of outgassing then the relative magnitude of the outgassing products will be diminished even further. To confirm the observation this phenomenon will need further investigation.
The samples were inspected before and after the test with the vacuum pump output being monitored using mass spectrometry to give an indication of any filament rupture (i.e. tube fracture will release molecules that will be detected by mass spectrometry, providing an indication of elapsed cycles before rupture). End capped hollow fibre Thermocouple
Liquid N2 Cooling channel
Heater plate
A point to mention though is the question whether a pressure build up within the hollow fibre could lead to cracking or whether micro-cracking of the hollow fibre/encapsulant could trigger any detrimental outgassing reactions. To asses whether this is likely to occur a series of thermal cycling tests were undertaken to evaluate the thermal stability of the storage vessels and two different end capping materials. The results of these tests are discussed in the next section. 5. THERMAL CYCLING OF THE SELFHEALING STORAGE FILAMENTS Determining the adverse effect of operational temperature extremes upon the self-healing filaments is a key requirement for any self-healing concept in a space environment. To assess the self-healing concept in this environment a purpose built test fixture was manufactured to assess the structural continuity of the seal at the end of the hollow fibres (see fig 4.). The specimens (10 individual hollow fibres each with an external diameter of 100µm ± 5µm) were cycled between ±100°C for 100 cycles at a heating rate of 10°C per minute. The ten samples were sandwiched in individual grooves in an Al block (see Fig 4b). Three thermocouples were pushed up inside the Al block such that they were only 5mm away from the internal cavity. The Al block had two channels machined in the bottom face to permit two copper cooling pipes using liquid Nitrogen to be secured to the specimen block. The final assembly was located upon a heater plate. This arrangement permitted the temperature profile to be controlled using a balance between the heater plate and the liquid Nitrogen and eliminated any thermal lag.
Fig 4. Thermal cycling set-up; (a) hollow fibre with end cap seal; (b) Al storage chamber with 10 cavities each housing a 100µm diameter fibre by 100mm long; (c) thermal cycling chamber In this experiment two different end-capping methods (2-part Cycom 823 epoxy and 1-part high modulus silicone sealant [Bostik 100HMA]) applied to 10 different filaments was undertaken. The test was arranged such that a number of filaments contained (1) air, i.e. they were unfilled, (2) filled with Cycom 823 resin or (3) filled with Cycom 823 hardener. This approach investigated any effect on the sealing method by the healing material within the hollow fibre. In the first experiment difficulty with locating the 100µm diameter fibres within their 100mm long cavity was experienced. This resulted with the loss of 2 storage tubes due to fibre crushing. These two failures were recorded by the mass spectrometer. Apart from
the crushing failures no other failure was recorded during the experiment of either the epoxy or silicon end caps. When viewed under the microscope the continuity of the end caps for both sealing methods was confirmed. The only significant difference between the two methods was their ease of application and the consistency of forming the meniscus at the end of the hollow tubes. The high viscosity of the silicone system is maintained throughout the room temperature curing process and hence results in a controlled ‘cap’. Conversely, the 2-part epoxy requires an increase in temperature to initiate the cure. However, before gelation occurs the resin system attains it’s lowest viscosity value (15cP) thus making the control of the end cap difficult. An example of the two different sealing methods is given in Fig 5. In addition to the process control problems the outgassing performance of the silicone is superior to the epoxy and hence was selected for further evaluation.
(a)
(b)
(c)
Fig. 5: End capping methods on hollow glass storage vessels; (a) 2-part epoxy correctly sealing over exposed end, (b) 2-part epoxy not sealing the exposed fibre end, (c) silicone correctly sealing over exposed fibre end. In the second test ten samples were sealed with silicone and tested using the same procedure outlined above. The mass spectrometry was used to monitor the samples throughout the experiment and didn’t record any failures. This was confirmed by visual examination after the experiment was concluded. 6. MECHANICAL PROPERTY ASSESSMENT 6.1 Self-healing laminate manufacture The hollow glass fibre diameter was chosen to have an external diameter of 60 µm ± 3 µm and an internal diameter of ~40 µm yielding a hollowness fraction (the ratio of internal area to external area) of ~55%. Once manufactured the individual fibres are then consolidated within a 913 epoxy resin film, which was selected to match the baseline laminate material. The
healing resin can then be drawn into the individual lamina using a vacuum. Once the ends have been sealed (in this research the fibres were sealed with Bostik Bond Flex 100HMA high modulus silicone sealant) the infused hollow fibre layers (which can be considered as standard prepreg sheets) are then laid-up according to the required stacking sequence and processed according to the manufacturer’s guidelines. Once fully cured the laminate was sectioned into individual samples measuring 25mm wide by 100mm long. After cutting, the edges of the samples are sealed with a two-part rapid curing epoxy system (Araldite Rapid) to prevent any healing resin loss through the exposed ends of the hollow fibres. 6.2 Flexural strength determination Four-point bend flexural strength testing according to ASTM D6272–02 [13] was selected to characterise the strength and stiffness of the baseline and self-healed samples. A detailed investigation of this work has been reported previously [11], but for clarity a brief summary will be given below again. The four-point loading configuration ensures a region of uniform bending stress in the area of the damage between the loading noses, rather than a peak stress under the nose at the point of worst damage. A support span to depth ratio of 32:1 with the loading noses positioned at one third of the span was selected. The tests were conducted using a loading rate of 5mm/min on a Roell Amsler hydraulic test frame fitted with a 25kN load cell. A Linear Potentiometric Displacement Transducer (LPDT) was used to record mid-span deflection, which was logged through a PC running Instron data acquisition software. In the case of the damaged samples a three point bend test was used to initiate matrix shear cracks and delamination damage with the composite laminate. In the case of the laminate containing the self-healing resin and hardener lamina this was permitted to heal by heating to 100°C (from ambient) in an air circulation oven and held at this temperature for 2 hours. This curing cycle was selected to match the typical operational temperature profile of a composite structure in space undertaking two low earth orbits (LEO). The flexural strength results for the five different laminate configurations (baseline, baseline + damage, self-heal baseline, self-heal baseline + damage, and self-heal baseline damaged + healed) are given in Table 1.
Table 1. Summary of flexural strength and percentageretained strength Specimen ID
Specimen description
Flexural strength [MPa]
% retained strength
Group A
Baseline laminate {[0°/+45°/90°/-45°]2s} – no damage
668 ± 13
100
Group B
Baseline laminate with selfhealing layers at critical interfaces {±45°/90°} – no damage
559 ± 12
84
Group C
Damaged (2500N indentation) baseline laminate
479 ± 32
72
Damaged baseline laminate with self-healing layers, 494 ± 7 no repair
74
Damaged baseline laminate with Group E self-healing layers, with 2 hours 578 ± 28 self-healing at 100°C
87
Group D
The experimental results shown in Table 1 illustrate the ultimate flexural strength and the percentage-retained strength for the different sample configurations. The results indicated that the inclusion of hollow fibres gives in an initial strength reduction of 16%. It was found that the baseline laminate (Group A) and the laminate containing the hollow fibre layers (but no healing) had comparable low energy impact damage tolerance both in terms of damage size and residual failure strength (typically 72-74%). After healing of the damage site was undertaken it was found that a selfhealed laminate (Group E) had a residual strength of 87% compared to an undamaged baseline laminate (Group A) and 100% compared to an undamaged selfhealing laminate (Group B).
The thermal outgassing of the 2-part Cycom 823 epoxy resin system was observed to fall outside the existing ESA guidelines, i.e. guidelines generated for cured materials and not for liquids. However, it has to be stated that this test is the worst case for a cracked hollow fibre. An undamaged and sealed fibre shall be mostly impermeable to outgassing since it will be located within the heart of the composite laminate. A final point to note concerns the selection of the healing resin. At present a resin system specifically designed for operation within the space environment does not exist. For this research project, the Cycom 823 twopart resin system was selected because it has a very low viscosity and resin gelation occurs within 15 minutes at 90°C. For a standard commercially available epoxy resin this system provides the best healing attributes of balancing low viscosity with relatively quick gelation thereby permitting infusion of the damage whilst minimising the extent of outgassing. A resin system optimised for this environment would clearly be more beneficial. The thermal cycling tests illustrated the robustness of the hollow fibre and the two different end capping materials for surviving the LEO thermal environment. A series of mechanical tests were undertaken to evaluate the influence on the flexural strength of incorporating hollow fibre plies into a baseline laminate. The results indicated that the inclusion of hollow fibres gives in an initial strength reduction of 16%. It has been shown that the damage site was undertaken it was found that a self-healed laminate had a residual strength of 87% compared to an undamaged baseline laminate and 100% compared to an undamaged self-healing laminate. This result would, therefore, suggest that biomimetic repair is now possible for advanced composite structures.
7. CONCLUSIONS
8. ACKNOWLEDGMENTS
This work has shown that a hollow-fibre self-healing approach can be used for the repair of advanced composite structures. In the course of this study it has been shown that hollow glass fibres containing a twopart healing resin can be manufactured and incorporated within a standard autoclave processing technique, thereby, indicating that this healing approach could be readily applied to existing composites manufacturing techniques. Through the specific placement of four self-healing plies to match the critical damage threat it has been shown that a standard commercial available two-part epoxy resin system (Cycom 823) can be used for the repair of internal matrix cracking and delaminations through the thickness of a 16 ply laminate.
The authors wish to acknowledge the financial support of the European Space Agency, ESTEC Contract No.: 18131/04/NL/PA, throughout the duration of this project.
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[2]
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