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Ageing of composites in the rotorcraft industry
11 Ageing of composites in the rotorcraft industry K. D R AG A N, Polish Air Force Institute of Technology, Poland
11.1
Introduction to composite structures applied in the rotorcraft industry using the example of PZL
This chapter discusses the manufacturing and structure design of main rotor blades (MRBs); it will cover design schemes, materials used and applications. The main helicopter, as well as composite-elements, manufacturer in Poland is PZL Swidnik S.A. Examples of helicopters manufactured by PZL include W-3 and SW-4 (http://www.pzl.swidnik.pl/). According to data received from the manufacturer (PZL Swidnik S.A., unpublished data, July 2007), the following polymer matrix composite structures are used: • •
composite structures with glass and carbon tapes, roving tows, fabrics such as reinforcement materials, mainly based on epoxy resin; composite sandwich structures with core materials such as Nomex®, glass–epoxy honeycomb and rigid foam plastics based on polyurethane or polimethacrylimide, and with epoxy–glass/carbon woven skin and epoxy adhesive films and adhesive pastes.
The composite materials for the structural elements of PZL helicopters – such as MRBs and tail rotor blades (TRBs) – consist of structures such as shells (glass fabrics with epoxy resin) and carrying straps (glass roving reinforced with epoxy resin). The non-critical elements of rotor blades are the trailing parts (sandwich structures using Nomex or epoxy–glass honeycomb or foam core). The construction of the MRB of an SW-4 helicopter is presented in Fig. 11.1. Composite structures used in the construction of non-critical elements can be divided into two groups: • •
elements working under static loads; elements without static and fatigue strength requirements. 311
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1. Spar (glass roving, glass fabrics/epoxy resin) 2. Blade grip 3. Link 4. Spar carrying strap (glass roving/epoxy resin) 5. Strap (glass fabric/epoxy resin) 6. Honeycomb core (Nomex: aramid fibre/phenolic resin) 7. Strap (glass fabric/epoxy resin) 8. Lightning system shielding (Cu–mesh/phenolic adhesive film) 9. Polyurethane tape 10. Leading edge grip (stainless steel) 11. Fore fairing (stainless steel) 12. Aft fairing (aluminium alloy) 13. Anti-flutter weight (lead) 14. Trailing edge strap (glass fabric/epoxy resin) 15. Trimming tab (aluminium alloy) 16. Trailing edge strap (glass roving/epoxy resin) 17. Aft section skin (glass fabric/epoxy resin) 18. Balance weight seat 19. Blade mooring seat
11.1 Construction of composite MRB of the SW-4 helicopter.
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11 10
Epoxy–carbon elements
13
6
Epoxy–glass elements 5
7 12
1 9 14 4 3 2 8 16
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15
11.2 Composite share in the SW-4 helicopter structure.
Examples of the first group include the following composite elements of helicopters: the horizontal stabilizer of W-3, the cabin structure of SW-4, doors and some elements that are not part of the primary structure of the fuselage. The skin of the horizontal stabilizer of W-3 is made from a glass– epoxy solid structure (tube spar) and from glass honeycomb (trailing part). Figure 11.2 presents the structure of the cabin of an SW-4 helicopter. Composite elements in the cabin structure are made from: • tapes, roving tows, woven hybrid composites with the use of glass and carbon reinforcement and epoxy resin; • sandwich structures with the use of carbon and glass fibre skin and Nomex honeycomb. Examples of elements that do not have static and fatigue strength requirements are cowlings, instrument panels, cowlings for instrument panels, linings, elements of the luggage compartment. The materials used for the helicopter elements are carbon fibres and glass fibres (as reinforcement) and epoxy matrix.
11.2
Potential damage that can occur in a composite main rotor blade
Information about damage, based on original equipment manufacturer (OEM) data as well as on maintenance data from users, will be provided in this chapter. Damage can occur during manufacturing as well as during
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the service life of MRBs. During the manufacturing of composite elements, and also during the maintenance of such elements, different defects can occur.1 According to Birt and Smith:1 One of the most serious is voids in the matrix, which can be further classified as: • delaminations – these are large planar voids occurring at the interfaces between the plies; • discrete voids; • porosity – this can be described as a large number of microvoids, each of which is too small to be of structural significance. It is usually produced during the curing cycle from entrapped air, moisture or volatile products.
Porosity has been a recognized problem for composites for a number of years. According to Birt and Smith, porosity influences the mechanical properties of composites and as a consequence affects the durability of such materials. Depending on the manufacturing process, as well as on manufacturing techniques, there are a number of defects that can occur in the composite structure: (PZL Swidnik S.A., unpublished data, July 2007 and reference 2): • • • • • • • • • • • • •
incompletely cured matrix; incorrect fibre volume fraction; ply misalignment; wavy fibres; ply cracking; delaminations; resin cracks; bonding defects; voids and porosity; foreign object damage (inclusions); resin rich/poor areas; bridging; fibre misalignments.
Different types of damage can occur during manufacturing and also during the service life of composites. Figure 11.3 shows the total percentage share of failure effects for metal and composite MRBs on the basis of data received from the users.3 Ageing issues and manufacturer faults, as well as in-service damage, are the most significant failure effects influencing the MRBs during service life. Figure 11.4 shows the percentage share of failure modes for the MRBs of helicopters. Most of the in-service damage is connected with skin–honeycomb disbonds which lead to skin separation and composite ply cracking, as well as moisture (water ingress) in the honeycomb structure.
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Manufacturer faults 32% In-service damage 11%
Repair faults 6% Maladjustment 6% Pilotage mistakes 2% Ageing 42%
Maintenance personnel faults 1%
11.3 Percentage share of failure effects for the MRBs of helicopters.
Skin separation 55% Disbond 26%
Crack 4% Leakage 3% Dents 3% Fracture 1% Deformation 1%
Moisture 1% Tear 1% Truncation 2% Break 3%
11.4 Percentage share of failure modes for the MRBs of helicopters.
Increased composite use in aerospace constructions has increased the importance of non-destructive testing (NDT) methods capable of identifying flaws in composites.4 There is a necessity for composite inspection to detect manufacturing as well as in-service faults from the point of view of durability of composites.5 The main reasons for the necessity of composite inspection are: • • •
human factors; improper manufacturing parameters (technology); improper materials, etc.
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Quality control must be applied in order to ensure the accuracy of manufactured elements. NDT is a key element for gathering information about structure quality and it helps to achieve higher accuracy and quality of manufactured elements. From the point of view of durability of composite materials, several factors affecting composite structure integrity (e.g. load cycles, temperature changes, low-energy damage, lightning strikes) should be considered. These factors influence the structural performance of composites and may lead to the occurrence of invisible damage in the composite structure. Based on the above, NDT techniques are necessary for the in-service maintenance of composite MRBs and for detection of damage such as: • • • •
disbonds; delaminations; water ingress; low-energy impact damage.
In order to describe the likelihood of detection of the above-mentioned damage types, tests that use reference standards should be applied. In order to perform such tests, it is necessary to develop composite reference standards to be used in NDT equipment calibration for the assessment of damage.6 Testing of such specimens enables, among other things: • the detection of different damages types by the use of a range of technique; • sensitivity description (signal to noise ratio). The results of a sample specimen inspection are presented below. Figure 11.5 shows water ingress detection using laser shearography (for a glass fibre honeycomb specimen with water cells of different sizes and shapes). This technique, with the use of thermal excitation, enables an assessment to be made of the presence water in the blade structure. Figure 11.6 is an example of a carbon fibre plate with teflon inserts of different sizes between plies which create foreign object inclusions. A step sample made from glass fibre with teflon inserts and removable metal shims to create disbonds and delaminations was also considered.7 A range of techniques can be applied for specimen testing: • • •
ultrasonic testing (disbonds, delaminations, foreign object inclusion); mechanical impedance analysis (disbonds); shearography (disbonds, water ingress).
Results obtained during specimen testing gave information on which technique should be used in relation to the following factors: inspected material, thickness range, object geometry, required damage detection, the best sensitivity. A wide range of different techniques provides many possibilities for
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0.79 0.58 0.36
27.8 56.4
0.14
86.3
–0.08
115.0
–0.29
144.9
–0.51
173.9
–0.73
203.8
–0.96 –1.16 0.0
23.8 48.7 72.9 97.8 121.6 146.5 170.6 195.5 219.3 244.
11.5 Water ingress indication with the use of NDT – shearography.
11.6 Carbon fibre specimen with artificial defects.
applying various damage and subsequent detection of damage scenarios. On the basis of this experiment the best techniques were selected for inspection of composite MRBs.
11.3
Low-energy impact damage and durability in a W-3 main rotor blade
Depending on the in-service conditions, the following events can occur: hail strikes and impact damage from bodies such as stones affecting the structural integrity of the MRB. On the basis of these events, low-energy impact tests were performed to assess the durability of composite MRBs to impact damage using advanced non-destructive inspection (NDI) techniques.
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Impact damage affects the structural integrity of composite elements.8 According to Smith, ‘Composite structures can suffer quite severe impact damage without a noticeable surface indentation – known as barelyvisible impact damage (BVID) – and this makes large-area defect detection a necessity for critical structural components’.2 Such accidents, especially those that create non-visible damage, are called low-energy impact damages.9 These damages are not visible to the human eye but can create damages affecting the structural integrity of elements. Very often, the results of low-energy impact damage are: delaminations, disbonds and multiple cracks. Regarding more complex structures such as composite rotor blades, defects that occur can be divided into the following groups: • • •
skin damage (delaminations, cracks); skin–honeycomb damage (disbonds); honeycomb damage (core crush).
For test purposes, the MRB of a W-3 helicopter was used. A special stand for the test was designed and the so-called ‘blunt impactors’ for the drop test were prepared. The term ‘blunt impactors’ indicates that these impactors do not affect the composite surface visually. The main aims of the test were: •
to deliver information about the influence of dropping a mass onto the blade structure (drop energy from 5 to 20 J); • to determine whether any failure modes will occur during the blade test; • to find appropriate, fast and reliable NDT techniques for the evaluation of drop test results; • to describe the influence of energy levels on the damage size. The reason for performing this test was to evaluate the behaviour of composite helicopter blades used in W-3 helicopters in situations that may occur during the service life. The next important issue was to deliver information on whether the maintenance techniques used are capable of finding damages that occur as a result of low-energy impact. The last stage was to evaluate which NDT technique was capable of finding such damages. At the beginning of the test, calibration standards were prepared to determine the sensitivity and accuracy of the techniques selected. One of the specimens was a step sample, and also blade structures with mechanically induced disbonds and delaminations. Tests were applied to these specimens to detect disbonds and delaminations. For the tests, the following techniques were used: mechanical impedance analysis (MIA), pitch–catch, shearography and D-Sight.10
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11.3.1 Test run The first stage of the test involved NDT of prepared W-3 MRB samples in order to detect any manufacturing damages. The next step was to describe the region of interest where impactors could be dropped. For the selected areas of MRBs and referred to the skin and honeycomb interface, the drop test was applied.9 After that stage, NDT was carried out and data were evaluated. The next step was again a drop test followed by NDT. The important issue was the fact that impactors were dropped on the same area as previously. The aim of this test was to analyse structure behaviour under repeated loading conditions and to evaluate the extent of the damage.
11.3.2 Test results The techniques used for the inspection were based on a PC interface which made it possible to collect data in the form of pictures and C-scans. This method of data presentation makes it possible to compare data between consecutive test stages. Figure 11.7 presents the results of NDT on the composite skin of a W-3 MRB. Areas marked with polygons show disturbance in the signal and indicate skin–honeycomb disbonds. As shown, differences in the size of the disbond area are related to the different drop energies used. Figure 11.8 shows the impact damage area and its relation to drop energy. Furthermore, results of two stages of the test are shown. The first is connected with the first drop of the impactors and denoted by MIA• •
MIA-1, 2 – first drop, second drop results; P–C-1, 2 (pitch–catch) – first drop, second drop results.
Figure 11.8 shows that an increase in drop energy is associated with an increase in damage area. Moreover, the second stage of the test proved that the impact damage area almost doubles in size. The damage size described differs depending on the NDT techniques used; this is a result of the different sensitivities of the techniques used. In
5J
10J
15J
20J
11.7 NDT results for the composite skin of the W-3 helicopter MRB.
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Damage size (cm2)
70.00 60.00 50.00 36.71 25.38 MIA-2 MIA-3 P–C-2 P–C-3
14.85 8.37 0.00
5
10 15 Drop energy (J)
20
11.8 Damage size in relation to the drop energy.
Table 11.1 Detection capabilities of the selected techniques Drop energy
MIA
Pitch–catch
Shearography
D-Sight
5J 10 J 15 J 20 J
+ + + +
+ + + +
+ + + +
− + + +
+, damage detected; −, damage not detected.
Table 11.2 Damage size (cm2) Drop energy
MIA
Pitch–catch
5J 10 J 15 J 20 J
24.54 41.82 48.14 74.10
13.68 25.12 39.89 54.44
addition, qualitative techniques were also applied, such as shearography and D-Sight. These techniques do not provide information about damage size but only about the existence of damage. Table 11.1 gives information about the detection capabilities of the techniques used. Table 11.2 presents the damage sizes determined from the NDT.
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11.3.3 Conclusions All the techniques applied detected damage (disbonds), with the exception of D-Sight for the smallest damage size (13–24 cm2). Generally, D-Sight is a good option for fast disbond detection. In the work of Heida,10 it was concluded that D-Sight can reliably detect significant impact damage, with a damage area equal to or larger than 5 cm2, within a field of view of about 0.25. However, one very important consideration is that surface curvature in the direction of observation leads to variations in the intensity of reflected light. On the other hand, the signature of surface defects is strongly reduced as a result of environmental light. On the inspected blade, the smallest damage may not be visible because of the geometry and also because of the presence of reflections from the structural elements of the blade. It should be mentioned that the manufacturer does not permit further blade maintenance if the skin–honeycomb disbond area is greater than 80 cm2. Data presented in Fig. 11.8 as well as in Table 11.2 show that the damage size for a drop energy of 20 J is very close to the acceptable limit for further blade maintenance. An ultrasonic pulse–echo technique was also applied to determine the possibility of detection of skin–honeycomb disbonds, as well as the possibility of delamination detection. Figure 11.8 shows the results of an ultrasonic pulse–echo C-scan. For this technique, disbond detection was not as good as that of other techniques due to adhesive in the honeycomb structure which affects the ultrasonic signal. Using this technique it proved possible to detect small delaminations occurring in the drop area. These damages were not detected using any of the other techniques. Further work is now being conducted with the aim of describing the effects of low-energy impacts on the composite blade structure in the critical areas, i.e. (a) the stress concentration area of the blade and (b) the spar structure. Problems to overcome include geometry changes and construction complexity, which may influence signal interpretation. For these purposes, special samples with artificial defects are prepared. At this stage, calibration work has been carried out and trials for the detection of defects are in preparation. Further results will be delivered in the near future.
11.4
Influence of moisture and temperature
The influence of environmental hazards will be discussed in this section, based on OEM information and tests performed in the manufacturer’s facility. Length of service life (airworthiness limitation), expressed in years or months, is mainly affected by moisture (which the composite material may absorb). The influence of other factors such as ultraviolet radiation is
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negligible over the service life as a result of the protection provided by coatings. In addition, manufacturers’ data showed that the environmental resistance of composites exposed to the ultraviolet light did not decrease over a 50-year period. After that period, fatigue strength degradation is approximately 20%, i.e. total fatigue strength equals 80% of the initial value. Laboratory tests carried out by PZL (ageing according to the ASTM D 5229 standard) proved that the composite material reaches total moisture saturation in a stable manner. This process continues over the service life in a similar way, the only difference being the rate of moisture absorption. On the basis of moisture diffusion in the composite material, it is possible to determine such parameters as: diffusion constant and effective equivalent moisture content (Mm). On the basis of the Mm determined, it is possible to calculate the time required for the ageing material to reach a moisture content equal to 99.9% Mm. Calculations performed by the manufacturer proved that for the main composite rotor blades these periods were much greater than in the real service life of these elements. It should be mentioned that during the calculation, the following factors were not taken into consideration: • •
the influence of the existence of protective coatings; the material behaviour in real conditions – i.e. the material absorbing moisture and then drying according to external (weather) conditions.
The calculations assumed that moisture and temperature affect material in a constant (unchanged) way. The most significant factor is that the material reaches a maximum moisture level in a stable way and longer ageing process do not change moisture content or affect material strength. Because composite material durability parameters decrease in higher temperatures, static strength tests were carried out at temperatures equal to the maximum service life temperature, taking the effect of the sun into consideration. During the tests it was concluded that the degree of strength degradation determined makes it possible to determine an additional static strength safety coefficient. It has been proven that the strength degradation coefficient under high temperature conditions decreases to as low as 0.75. The next issue was to determine strength degradation coefficients during the ageing process in climate chambers. Results show that ageing of materials in conditioning chambers (temperature 80 °C; moisture equal to 85%), with the use of high and low temperatures, affects the static strength of composite materials. These coefficients express the ratio of the strength of the aged specimens (tested in the temperature range −45 °C to 80 °C) to the strength of specimens tested in the environment temperature; the coefficients have a value of about 0.5 depending on the type of structure and the kind of load. A comparison of painted and non-painted elements subjected
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to natural ageing (in the air) shows much lower strength decreases than those subjected to an accelerated ageing process. During natural ageing, the decrease in strength of elements was 40%, in reality, these decreases are lower than 40%. Erosion occurring on the external surface of unpainted specimens was the direct reason for this. During service life, composite materials are protected by the painted coatings that protect the material from erosion as well as from moisture (continued a moisture barrier primer). Moreover, it was proven that composite degradation during accelerated ageing was greater than that in environmental ageing. Tests made on the laboratory samples were confirmed by testing blades that had been aged in natural conditions. During tests the following measurements were made: weight, mass balance and stiffness, additional static tests and assessment of appearance. No significant changes in the static strength or in the mass and stiffness were found. One of the effects of moisture absorption was a decrease in the matrix glass transition temperature; this temperature is very important in relation to composite materials. Above this temperature, rapid change in physical properties can occur that influence mechanical properties, in particular properties that are strongly dependent on the matrix – for example, its interlaminar shear strength. In this respect, it is essential to determine the influence of the ageing process on the glass transition temperature. Glass transition temperatures were determined for the composite materials that were submitted to ageing processes. The glass transition temperature was verified by results of material strength tests at elevated temperatures. These analyses proved that glass transition temperatures exceed maximum helicopter operating temperatures. In this respect, the materials described were recognized as safe for use in these helicopters.
11.5
New techniques for testing composite structures
Techniques such as structural health monitoring (SHM) will be discussed in this section, using information available in the published literature on modern structural integrity assessment. Applying NDT to composites is necessary not only because of their applications in the rotorcraft industry but also their applications in other aircraft structures. The continuous development of the extensive use of composites in aerospace applications creates great demand for advanced NDT techniques as well as for health monitoring techniques. The main advantages of the use of advanced NDT techniques is the possibility of obtaining information about the structural integrity of composites without affecting structure properties and also the collection of data for structure monitoring. The main disadvantages are the necessity to have the aircraft on the ground, the human staff requirements
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and the time required to inspect all relevant locations.11 The use of network sensors, distributed in the structure, for periodical or even continuous monitoring could affect the time required for inspections. SHM enables NDT to be deployed with greater access to more complex structures. According to Roach, ‘The core of SHM is the development of self-sufficient systems, that use built-in distributed sensors/actuators not only to detect structural failures but to monitor the effects of structural usage’.11 The use of sensor networks could be achieved using three approaches, as listed below:11 • • •
in situ sensors; sensor networks with in situ data acquisition; sensors networks with real-time data transmission to a remote site.
In the first group, sensors are permanently installed in the structure. For diagnostic purposes aircraft must be on the ground. All necessary items such as power supply and data acquisition electronics must be delivered to the aircraft. All sensors are connected to the data acquisition system, and data analysis and detection procedures are performed on-site. In this system, manual or automated data collection are substituted by distributed networks. In the second group, distributed networks are equipped with the electronics and memory to record and to store data. ‘Those items are equipped with programmable circuits having the power for automated data logging in flight or on the ground’.11 Data must be gathered by technical staff while the aircraft is on the ground. The third group is quite similar to the second group; the only difference is the use of communication system applications. This system enables wireless transmission of the data collected. For this particular application, software could be developed to send data to maintenance personnel. There is much work dedicated to the possibility of applying SHM to aircraft, as well as to composites. It appears that the use of in situ sensors for the SHM of aircrafts to will be a aviable option in the near future.11
11.6
References
1 E A BIRT, R A SMITH, A review of NDE methods for porosity measurement in fibre-reinforced polymer composites, Insight, 46(11), November 2004, 681–686. 2 R A SMITH, An introduction to the ultrasonic inspection of composites, in Non Destructive Evaluation of Composite Materials, Course, Farnborough, November 2004. 3 K DRAGAN, S KLIMASZEWSKI, In-service NDI of aging helicopters main rotor blades used in Polish Armed Forces, in 9th Joint FAA/DoD/NASA Aging Aircraft Conference, Atlanta, Georgia March 2006.
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4 D ROACH, Enhanced inspection methods to characterize bonded joints: moving beyond flaw detection to quantify adhesive strength, in Air Transport Association Nondestructive Testing Forum, Fort Worth, Texas, October 2006. 5 COMPOSITE QUALIFICATION CRITERIA, in Proceedings of the 51st Annual Forum of the American Helicopter Society, Fort Worth, Texas, May 1995. 6 D ROACH, Development of composite honeycomb and solid laminate reference standards to aid aircraft inspections, Sandia Report, SAND99-05405, Albuquerque, New Mexico. 7 J GIESKE, D ROACH, P WALKINGTON, Ultrasonic inspection technique for composite doubler/aluminium skin bond integrity for aircraft, Sandia Report, SAND980311C, Albuquerque, New Mexico. 8 K DRAGAN, S KLIMASZEWSKI, Low energy impact damage detection in the composite sandwich elements, in 35th Polish National Conference on Non Destructive Testing, Szczyrk, Poland, October 2006. 9 K DRAGAN, S KLYSZ, J LISIECKI, Detection of damages from the low energy impact damage in the composite structures, in Congress of Polish Mechanics, Warszawa, Poland, August 2007. 10 JH HEIDA, D-Sight technique for rapid impact damage detection on composite aircraft structures, NDT.net, 4(6), June 1999. 11 DH ROACH, ‘Smart’ aircraft structures: a future necessity. Health monitoring of aircraft structures using distributed in situ sensor system, High-Performance Composites, January 2007.
11.1
Introduction to composite structures applied in the rotorcraft industry using the example of PZL
This chapter discusses the manufacturing and structure design of main rotor blades (MRBs); it will cover design schemes, materials used and applications. The main helicopter, as well as composite-elements, manufacturer in Poland is PZL Swidnik S.A. Examples of helicopters manufactured by PZL include W-3 and SW-4 (http://www.pzl.swidnik.pl/). According to data received from the manufacturer (PZL Swidnik S.A., unpublished data, July 2007), the following polymer matrix composite structures are used: • •
composite structures with glass and carbon tapes, roving tows, fabrics such as reinforcement materials, mainly based on epoxy resin; composite sandwich structures with core materials such as Nomex®, glass–epoxy honeycomb and rigid foam plastics based on polyurethane or polimethacrylimide, and with epoxy–glass/carbon woven skin and epoxy adhesive films and adhesive pastes.
The composite materials for the structural elements of PZL helicopters – such as MRBs and tail rotor blades (TRBs) – consist of structures such as shells (glass fabrics with epoxy resin) and carrying straps (glass roving reinforced with epoxy resin). The non-critical elements of rotor blades are the trailing parts (sandwich structures using Nomex or epoxy–glass honeycomb or foam core). The construction of the MRB of an SW-4 helicopter is presented in Fig. 11.1. Composite structures used in the construction of non-critical elements can be divided into two groups: • •
elements working under static loads; elements without static and fatigue strength requirements. 311
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1. Spar (glass roving, glass fabrics/epoxy resin) 2. Blade grip 3. Link 4. Spar carrying strap (glass roving/epoxy resin) 5. Strap (glass fabric/epoxy resin) 6. Honeycomb core (Nomex: aramid fibre/phenolic resin) 7. Strap (glass fabric/epoxy resin) 8. Lightning system shielding (Cu–mesh/phenolic adhesive film) 9. Polyurethane tape 10. Leading edge grip (stainless steel) 11. Fore fairing (stainless steel) 12. Aft fairing (aluminium alloy) 13. Anti-flutter weight (lead) 14. Trailing edge strap (glass fabric/epoxy resin) 15. Trimming tab (aluminium alloy) 16. Trailing edge strap (glass roving/epoxy resin) 17. Aft section skin (glass fabric/epoxy resin) 18. Balance weight seat 19. Blade mooring seat
11.1 Construction of composite MRB of the SW-4 helicopter.
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11 10
Epoxy–carbon elements
13
6
Epoxy–glass elements 5
7 12
1 9 14 4 3 2 8 16
17
15
11.2 Composite share in the SW-4 helicopter structure.
Examples of the first group include the following composite elements of helicopters: the horizontal stabilizer of W-3, the cabin structure of SW-4, doors and some elements that are not part of the primary structure of the fuselage. The skin of the horizontal stabilizer of W-3 is made from a glass– epoxy solid structure (tube spar) and from glass honeycomb (trailing part). Figure 11.2 presents the structure of the cabin of an SW-4 helicopter. Composite elements in the cabin structure are made from: • tapes, roving tows, woven hybrid composites with the use of glass and carbon reinforcement and epoxy resin; • sandwich structures with the use of carbon and glass fibre skin and Nomex honeycomb. Examples of elements that do not have static and fatigue strength requirements are cowlings, instrument panels, cowlings for instrument panels, linings, elements of the luggage compartment. The materials used for the helicopter elements are carbon fibres and glass fibres (as reinforcement) and epoxy matrix.
11.2
Potential damage that can occur in a composite main rotor blade
Information about damage, based on original equipment manufacturer (OEM) data as well as on maintenance data from users, will be provided in this chapter. Damage can occur during manufacturing as well as during
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the service life of MRBs. During the manufacturing of composite elements, and also during the maintenance of such elements, different defects can occur.1 According to Birt and Smith:1 One of the most serious is voids in the matrix, which can be further classified as: • delaminations – these are large planar voids occurring at the interfaces between the plies; • discrete voids; • porosity – this can be described as a large number of microvoids, each of which is too small to be of structural significance. It is usually produced during the curing cycle from entrapped air, moisture or volatile products.
Porosity has been a recognized problem for composites for a number of years. According to Birt and Smith, porosity influences the mechanical properties of composites and as a consequence affects the durability of such materials. Depending on the manufacturing process, as well as on manufacturing techniques, there are a number of defects that can occur in the composite structure: (PZL Swidnik S.A., unpublished data, July 2007 and reference 2): • • • • • • • • • • • • •
incompletely cured matrix; incorrect fibre volume fraction; ply misalignment; wavy fibres; ply cracking; delaminations; resin cracks; bonding defects; voids and porosity; foreign object damage (inclusions); resin rich/poor areas; bridging; fibre misalignments.
Different types of damage can occur during manufacturing and also during the service life of composites. Figure 11.3 shows the total percentage share of failure effects for metal and composite MRBs on the basis of data received from the users.3 Ageing issues and manufacturer faults, as well as in-service damage, are the most significant failure effects influencing the MRBs during service life. Figure 11.4 shows the percentage share of failure modes for the MRBs of helicopters. Most of the in-service damage is connected with skin–honeycomb disbonds which lead to skin separation and composite ply cracking, as well as moisture (water ingress) in the honeycomb structure.
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Manufacturer faults 32% In-service damage 11%
Repair faults 6% Maladjustment 6% Pilotage mistakes 2% Ageing 42%
Maintenance personnel faults 1%
11.3 Percentage share of failure effects for the MRBs of helicopters.
Skin separation 55% Disbond 26%
Crack 4% Leakage 3% Dents 3% Fracture 1% Deformation 1%
Moisture 1% Tear 1% Truncation 2% Break 3%
11.4 Percentage share of failure modes for the MRBs of helicopters.
Increased composite use in aerospace constructions has increased the importance of non-destructive testing (NDT) methods capable of identifying flaws in composites.4 There is a necessity for composite inspection to detect manufacturing as well as in-service faults from the point of view of durability of composites.5 The main reasons for the necessity of composite inspection are: • • •
human factors; improper manufacturing parameters (technology); improper materials, etc.
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Quality control must be applied in order to ensure the accuracy of manufactured elements. NDT is a key element for gathering information about structure quality and it helps to achieve higher accuracy and quality of manufactured elements. From the point of view of durability of composite materials, several factors affecting composite structure integrity (e.g. load cycles, temperature changes, low-energy damage, lightning strikes) should be considered. These factors influence the structural performance of composites and may lead to the occurrence of invisible damage in the composite structure. Based on the above, NDT techniques are necessary for the in-service maintenance of composite MRBs and for detection of damage such as: • • • •
disbonds; delaminations; water ingress; low-energy impact damage.
In order to describe the likelihood of detection of the above-mentioned damage types, tests that use reference standards should be applied. In order to perform such tests, it is necessary to develop composite reference standards to be used in NDT equipment calibration for the assessment of damage.6 Testing of such specimens enables, among other things: • the detection of different damages types by the use of a range of technique; • sensitivity description (signal to noise ratio). The results of a sample specimen inspection are presented below. Figure 11.5 shows water ingress detection using laser shearography (for a glass fibre honeycomb specimen with water cells of different sizes and shapes). This technique, with the use of thermal excitation, enables an assessment to be made of the presence water in the blade structure. Figure 11.6 is an example of a carbon fibre plate with teflon inserts of different sizes between plies which create foreign object inclusions. A step sample made from glass fibre with teflon inserts and removable metal shims to create disbonds and delaminations was also considered.7 A range of techniques can be applied for specimen testing: • • •
ultrasonic testing (disbonds, delaminations, foreign object inclusion); mechanical impedance analysis (disbonds); shearography (disbonds, water ingress).
Results obtained during specimen testing gave information on which technique should be used in relation to the following factors: inspected material, thickness range, object geometry, required damage detection, the best sensitivity. A wide range of different techniques provides many possibilities for
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0.79 0.58 0.36
27.8 56.4
0.14
86.3
–0.08
115.0
–0.29
144.9
–0.51
173.9
–0.73
203.8
–0.96 –1.16 0.0
23.8 48.7 72.9 97.8 121.6 146.5 170.6 195.5 219.3 244.
11.5 Water ingress indication with the use of NDT – shearography.
11.6 Carbon fibre specimen with artificial defects.
applying various damage and subsequent detection of damage scenarios. On the basis of this experiment the best techniques were selected for inspection of composite MRBs.
11.3
Low-energy impact damage and durability in a W-3 main rotor blade
Depending on the in-service conditions, the following events can occur: hail strikes and impact damage from bodies such as stones affecting the structural integrity of the MRB. On the basis of these events, low-energy impact tests were performed to assess the durability of composite MRBs to impact damage using advanced non-destructive inspection (NDI) techniques.
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Impact damage affects the structural integrity of composite elements.8 According to Smith, ‘Composite structures can suffer quite severe impact damage without a noticeable surface indentation – known as barelyvisible impact damage (BVID) – and this makes large-area defect detection a necessity for critical structural components’.2 Such accidents, especially those that create non-visible damage, are called low-energy impact damages.9 These damages are not visible to the human eye but can create damages affecting the structural integrity of elements. Very often, the results of low-energy impact damage are: delaminations, disbonds and multiple cracks. Regarding more complex structures such as composite rotor blades, defects that occur can be divided into the following groups: • • •
skin damage (delaminations, cracks); skin–honeycomb damage (disbonds); honeycomb damage (core crush).
For test purposes, the MRB of a W-3 helicopter was used. A special stand for the test was designed and the so-called ‘blunt impactors’ for the drop test were prepared. The term ‘blunt impactors’ indicates that these impactors do not affect the composite surface visually. The main aims of the test were: •
to deliver information about the influence of dropping a mass onto the blade structure (drop energy from 5 to 20 J); • to determine whether any failure modes will occur during the blade test; • to find appropriate, fast and reliable NDT techniques for the evaluation of drop test results; • to describe the influence of energy levels on the damage size. The reason for performing this test was to evaluate the behaviour of composite helicopter blades used in W-3 helicopters in situations that may occur during the service life. The next important issue was to deliver information on whether the maintenance techniques used are capable of finding damages that occur as a result of low-energy impact. The last stage was to evaluate which NDT technique was capable of finding such damages. At the beginning of the test, calibration standards were prepared to determine the sensitivity and accuracy of the techniques selected. One of the specimens was a step sample, and also blade structures with mechanically induced disbonds and delaminations. Tests were applied to these specimens to detect disbonds and delaminations. For the tests, the following techniques were used: mechanical impedance analysis (MIA), pitch–catch, shearography and D-Sight.10
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11.3.1 Test run The first stage of the test involved NDT of prepared W-3 MRB samples in order to detect any manufacturing damages. The next step was to describe the region of interest where impactors could be dropped. For the selected areas of MRBs and referred to the skin and honeycomb interface, the drop test was applied.9 After that stage, NDT was carried out and data were evaluated. The next step was again a drop test followed by NDT. The important issue was the fact that impactors were dropped on the same area as previously. The aim of this test was to analyse structure behaviour under repeated loading conditions and to evaluate the extent of the damage.
11.3.2 Test results The techniques used for the inspection were based on a PC interface which made it possible to collect data in the form of pictures and C-scans. This method of data presentation makes it possible to compare data between consecutive test stages. Figure 11.7 presents the results of NDT on the composite skin of a W-3 MRB. Areas marked with polygons show disturbance in the signal and indicate skin–honeycomb disbonds. As shown, differences in the size of the disbond area are related to the different drop energies used. Figure 11.8 shows the impact damage area and its relation to drop energy. Furthermore, results of two stages of the test are shown. The first is connected with the first drop of the impactors and denoted by MIA• •
MIA-1, 2 – first drop, second drop results; P–C-1, 2 (pitch–catch) – first drop, second drop results.
Figure 11.8 shows that an increase in drop energy is associated with an increase in damage area. Moreover, the second stage of the test proved that the impact damage area almost doubles in size. The damage size described differs depending on the NDT techniques used; this is a result of the different sensitivities of the techniques used. In
5J
10J
15J
20J
11.7 NDT results for the composite skin of the W-3 helicopter MRB.
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Damage size (cm2)
70.00 60.00 50.00 36.71 25.38 MIA-2 MIA-3 P–C-2 P–C-3
14.85 8.37 0.00
5
10 15 Drop energy (J)
20
11.8 Damage size in relation to the drop energy.
Table 11.1 Detection capabilities of the selected techniques Drop energy
MIA
Pitch–catch
Shearography
D-Sight
5J 10 J 15 J 20 J
+ + + +
+ + + +
+ + + +
− + + +
+, damage detected; −, damage not detected.
Table 11.2 Damage size (cm2) Drop energy
MIA
Pitch–catch
5J 10 J 15 J 20 J
24.54 41.82 48.14 74.10
13.68 25.12 39.89 54.44
addition, qualitative techniques were also applied, such as shearography and D-Sight. These techniques do not provide information about damage size but only about the existence of damage. Table 11.1 gives information about the detection capabilities of the techniques used. Table 11.2 presents the damage sizes determined from the NDT.
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11.3.3 Conclusions All the techniques applied detected damage (disbonds), with the exception of D-Sight for the smallest damage size (13–24 cm2). Generally, D-Sight is a good option for fast disbond detection. In the work of Heida,10 it was concluded that D-Sight can reliably detect significant impact damage, with a damage area equal to or larger than 5 cm2, within a field of view of about 0.25. However, one very important consideration is that surface curvature in the direction of observation leads to variations in the intensity of reflected light. On the other hand, the signature of surface defects is strongly reduced as a result of environmental light. On the inspected blade, the smallest damage may not be visible because of the geometry and also because of the presence of reflections from the structural elements of the blade. It should be mentioned that the manufacturer does not permit further blade maintenance if the skin–honeycomb disbond area is greater than 80 cm2. Data presented in Fig. 11.8 as well as in Table 11.2 show that the damage size for a drop energy of 20 J is very close to the acceptable limit for further blade maintenance. An ultrasonic pulse–echo technique was also applied to determine the possibility of detection of skin–honeycomb disbonds, as well as the possibility of delamination detection. Figure 11.8 shows the results of an ultrasonic pulse–echo C-scan. For this technique, disbond detection was not as good as that of other techniques due to adhesive in the honeycomb structure which affects the ultrasonic signal. Using this technique it proved possible to detect small delaminations occurring in the drop area. These damages were not detected using any of the other techniques. Further work is now being conducted with the aim of describing the effects of low-energy impacts on the composite blade structure in the critical areas, i.e. (a) the stress concentration area of the blade and (b) the spar structure. Problems to overcome include geometry changes and construction complexity, which may influence signal interpretation. For these purposes, special samples with artificial defects are prepared. At this stage, calibration work has been carried out and trials for the detection of defects are in preparation. Further results will be delivered in the near future.
11.4
Influence of moisture and temperature
The influence of environmental hazards will be discussed in this section, based on OEM information and tests performed in the manufacturer’s facility. Length of service life (airworthiness limitation), expressed in years or months, is mainly affected by moisture (which the composite material may absorb). The influence of other factors such as ultraviolet radiation is
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negligible over the service life as a result of the protection provided by coatings. In addition, manufacturers’ data showed that the environmental resistance of composites exposed to the ultraviolet light did not decrease over a 50-year period. After that period, fatigue strength degradation is approximately 20%, i.e. total fatigue strength equals 80% of the initial value. Laboratory tests carried out by PZL (ageing according to the ASTM D 5229 standard) proved that the composite material reaches total moisture saturation in a stable manner. This process continues over the service life in a similar way, the only difference being the rate of moisture absorption. On the basis of moisture diffusion in the composite material, it is possible to determine such parameters as: diffusion constant and effective equivalent moisture content (Mm). On the basis of the Mm determined, it is possible to calculate the time required for the ageing material to reach a moisture content equal to 99.9% Mm. Calculations performed by the manufacturer proved that for the main composite rotor blades these periods were much greater than in the real service life of these elements. It should be mentioned that during the calculation, the following factors were not taken into consideration: • •
the influence of the existence of protective coatings; the material behaviour in real conditions – i.e. the material absorbing moisture and then drying according to external (weather) conditions.
The calculations assumed that moisture and temperature affect material in a constant (unchanged) way. The most significant factor is that the material reaches a maximum moisture level in a stable way and longer ageing process do not change moisture content or affect material strength. Because composite material durability parameters decrease in higher temperatures, static strength tests were carried out at temperatures equal to the maximum service life temperature, taking the effect of the sun into consideration. During the tests it was concluded that the degree of strength degradation determined makes it possible to determine an additional static strength safety coefficient. It has been proven that the strength degradation coefficient under high temperature conditions decreases to as low as 0.75. The next issue was to determine strength degradation coefficients during the ageing process in climate chambers. Results show that ageing of materials in conditioning chambers (temperature 80 °C; moisture equal to 85%), with the use of high and low temperatures, affects the static strength of composite materials. These coefficients express the ratio of the strength of the aged specimens (tested in the temperature range −45 °C to 80 °C) to the strength of specimens tested in the environment temperature; the coefficients have a value of about 0.5 depending on the type of structure and the kind of load. A comparison of painted and non-painted elements subjected
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to natural ageing (in the air) shows much lower strength decreases than those subjected to an accelerated ageing process. During natural ageing, the decrease in strength of elements was 40%, in reality, these decreases are lower than 40%. Erosion occurring on the external surface of unpainted specimens was the direct reason for this. During service life, composite materials are protected by the painted coatings that protect the material from erosion as well as from moisture (continued a moisture barrier primer). Moreover, it was proven that composite degradation during accelerated ageing was greater than that in environmental ageing. Tests made on the laboratory samples were confirmed by testing blades that had been aged in natural conditions. During tests the following measurements were made: weight, mass balance and stiffness, additional static tests and assessment of appearance. No significant changes in the static strength or in the mass and stiffness were found. One of the effects of moisture absorption was a decrease in the matrix glass transition temperature; this temperature is very important in relation to composite materials. Above this temperature, rapid change in physical properties can occur that influence mechanical properties, in particular properties that are strongly dependent on the matrix – for example, its interlaminar shear strength. In this respect, it is essential to determine the influence of the ageing process on the glass transition temperature. Glass transition temperatures were determined for the composite materials that were submitted to ageing processes. The glass transition temperature was verified by results of material strength tests at elevated temperatures. These analyses proved that glass transition temperatures exceed maximum helicopter operating temperatures. In this respect, the materials described were recognized as safe for use in these helicopters.
11.5
New techniques for testing composite structures
Techniques such as structural health monitoring (SHM) will be discussed in this section, using information available in the published literature on modern structural integrity assessment. Applying NDT to composites is necessary not only because of their applications in the rotorcraft industry but also their applications in other aircraft structures. The continuous development of the extensive use of composites in aerospace applications creates great demand for advanced NDT techniques as well as for health monitoring techniques. The main advantages of the use of advanced NDT techniques is the possibility of obtaining information about the structural integrity of composites without affecting structure properties and also the collection of data for structure monitoring. The main disadvantages are the necessity to have the aircraft on the ground, the human staff requirements
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and the time required to inspect all relevant locations.11 The use of network sensors, distributed in the structure, for periodical or even continuous monitoring could affect the time required for inspections. SHM enables NDT to be deployed with greater access to more complex structures. According to Roach, ‘The core of SHM is the development of self-sufficient systems, that use built-in distributed sensors/actuators not only to detect structural failures but to monitor the effects of structural usage’.11 The use of sensor networks could be achieved using three approaches, as listed below:11 • • •
in situ sensors; sensor networks with in situ data acquisition; sensors networks with real-time data transmission to a remote site.
In the first group, sensors are permanently installed in the structure. For diagnostic purposes aircraft must be on the ground. All necessary items such as power supply and data acquisition electronics must be delivered to the aircraft. All sensors are connected to the data acquisition system, and data analysis and detection procedures are performed on-site. In this system, manual or automated data collection are substituted by distributed networks. In the second group, distributed networks are equipped with the electronics and memory to record and to store data. ‘Those items are equipped with programmable circuits having the power for automated data logging in flight or on the ground’.11 Data must be gathered by technical staff while the aircraft is on the ground. The third group is quite similar to the second group; the only difference is the use of communication system applications. This system enables wireless transmission of the data collected. For this particular application, software could be developed to send data to maintenance personnel. There is much work dedicated to the possibility of applying SHM to aircraft, as well as to composites. It appears that the use of in situ sensors for the SHM of aircrafts to will be a aviable option in the near future.11
11.6
References
1 E A BIRT, R A SMITH, A review of NDE methods for porosity measurement in fibre-reinforced polymer composites, Insight, 46(11), November 2004, 681–686. 2 R A SMITH, An introduction to the ultrasonic inspection of composites, in Non Destructive Evaluation of Composite Materials, Course, Farnborough, November 2004. 3 K DRAGAN, S KLIMASZEWSKI, In-service NDI of aging helicopters main rotor blades used in Polish Armed Forces, in 9th Joint FAA/DoD/NASA Aging Aircraft Conference, Atlanta, Georgia March 2006.
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4 D ROACH, Enhanced inspection methods to characterize bonded joints: moving beyond flaw detection to quantify adhesive strength, in Air Transport Association Nondestructive Testing Forum, Fort Worth, Texas, October 2006. 5 COMPOSITE QUALIFICATION CRITERIA, in Proceedings of the 51st Annual Forum of the American Helicopter Society, Fort Worth, Texas, May 1995. 6 D ROACH, Development of composite honeycomb and solid laminate reference standards to aid aircraft inspections, Sandia Report, SAND99-05405, Albuquerque, New Mexico. 7 J GIESKE, D ROACH, P WALKINGTON, Ultrasonic inspection technique for composite doubler/aluminium skin bond integrity for aircraft, Sandia Report, SAND980311C, Albuquerque, New Mexico. 8 K DRAGAN, S KLIMASZEWSKI, Low energy impact damage detection in the composite sandwich elements, in 35th Polish National Conference on Non Destructive Testing, Szczyrk, Poland, October 2006. 9 K DRAGAN, S KLYSZ, J LISIECKI, Detection of damages from the low energy impact damage in the composite structures, in Congress of Polish Mechanics, Warszawa, Poland, August 2007. 10 JH HEIDA, D-Sight technique for rapid impact damage detection on composite aircraft structures, NDT.net, 4(6), June 1999. 11 DH ROACH, ‘Smart’ aircraft structures: a future necessity. Health monitoring of aircraft structures using distributed in situ sensor system, High-Performance Composites, January 2007.