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Damage tolerance and survivability of composite aircraft structures
Damage tolerance and survivability of composite aircraft structures
23
B. Rasuo University of Belgrade, Belgrade, Serbia
23.1
Introduction
The analysis of the dynamic behaviour of damaged constructions can provide knowledge about the acceptable level of damages and likely chances for aircraft survival. The aircraft’s ability to survive even after exposure to severe damage to the vital and loadcarrying parts of the aircraft construction is imperative not only for combat aircraft but also for civil aircraft (Ball, 2003; Raymer, 1999; Richards, Hastings, Rhodes, Ross, & Weigel, 2009; Schwarz & Drake, 2001). Vulnerability, as an element of survivability, is one of the most important exploitable characteristics of contemporary aircraft. With respect to survivability, composite laminated materials demonstrate the best behaviour and results compared to other materials that are currently used in aviation (Rasuo, 1995). Requirements and guidelines for the general programme are contained in the standards MIL-STD-2069, FAR/JAR 25.571, and, for bird strike requirements, FAR/JAR 25.631 (Eschenfelder, 2005; Petit, Bouvet, Bergerot, & Barrau, 2007; Rasuo, 2004). Predicting damage in laminated composite aircraft components due to impact events such as runway debris, hail, and birds is an area of ongoing research (see Figures 23.1e23.3). To reduce certification and development costs, computational methods are required by the aircraft industry to be able to predict structural integrity of composite structures under high-velocity impacts from hard objects, such as metal fragments and stone debris, and from soft or deformable bodies such as birds, hailstones, and tyre rubber. Key issues are the development of suitable constitutive laws for modelling composites in ply, determination of composite parameters from dynamic materials tests, materials laws for deformable impactors, and the efficient implementation of the materials models into finite element codes (Soutis, 2005; Soutis & Beaumont, 2005). Impact damages in aircraft structures made from laminated composites are very complex; the most common are matrix cracking, fibre failure, and delamination (Gerlach, Siviour, Wiegand, & Petrinic, 2012). The ability to predict the initiation and growth of damage is crucial for predicting performance and developing reliable and safe designs of composites (Sultan, Worden, Staszewski, & Hodzic, 2012). By use of simulation in modelling the impact damage, the test costs of aircraft structures made from composite laminates will be reduced (Kreculj & Rasuo, 2009). Structural Integrity and Durability of Advanced Composites. http://dx.doi.org/10.1016/B978-0-08-100137-0.00023-7 Copyright © 2015 Elsevier Ltd. All rights reserved.
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Figure 23.1 Impact-induced mechanical damage of inlet cone of a MiG-21bis fighter aircraft. Courtesy Moma Stanojlovic Company.
Figure 23.2 Hail strike-induced mechanical damage on nose cone of a Cessna 650 aeroplane. Courtesy Prince Aviation Company.
Due to the wide-scale use of composite materials in different aircraft structures, it is necessary to introduce new approaches for impact damage modelling. Efficient methodologies are modelling composite structures by applying specialized finite element methods that take into account macromechanical structural properties, as well as by using numerical methods with complicated analysis codes. In order to validate those methodologies for further use in the structures’ strength design, numerical model verification has to be done. A comparison of the obtained experimental and numerical results is further made to determine if the proposed models are accurate and valid (Kreculj & Rasuo, 2013).
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Figure 23.3 Impact-induced mechanical damage of the thrust reverser door on a Boeing 737-700. Courtesy JAT Tehnika Company.
Composite damages exist at the micro-scale level, while impact loads are applied at the structural level. Hence, it is necessary to consider a multi-scale approach for that kind of problem (Soutis & Beaumont, 2005). The development of suitable constitutive laws for modelling composite laminate failures and material models with finite element codes under impact provides significant assistance in the design and exploitation phases of specified structures. The most well-known failure criteria (two- and three-dimensional) for composite materials are Tsai-Wu, ChangeChang, and Hashin. They are used to predict the level or degree of damage, fracture, and failure of composite structures (Soutis & Beaumont, 2005). Development of computational models and simulations is necessary in studying the onset and growth of impact damages. In numerical simulation of impacts on composite structures, two- or three-dimensional models exist. For impact damage analysis of composite laminates, commercial software ABAQUS and LS-DYNA are most frequently used. Pro/ENGINEER, ANSYS, and some noncommercial software is also suitable for the same purpose. By using it in some circumstances (geometry, boundary conditions, mesh, load, etc.), the distribution of damage, stress, strain, and deformations can be analysed and presented (Kreculj & Rasuo, 2013). Essentially, when considering impact damage modelling in laminated composites of aeronautical structures, the dynamic response of composite laminates under impact is investigated and damage initiation/growth in such structures is predicted (Guyett & Cardrick, 1980; Poon, 1990). For such analysis, numerical modelling and simulations provide important and valuable results. Due to the anisotropy of composite laminates and nonuniform distribution of stress under dynamic loading, the failure process of laminates is very complex. The dynamic response of composite structures subjected to transient dynamic loading has been studied for years in terms of analytical, numerical, and experimental works (Bayandor, Thomson, Scott, Nguyen, & Elder, 2003).
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Experimental methodology for evaluation of damage tolerance and survivability
Taking into account the quantum of considered problems of damage tolerance or survivability of composite aircraft structures in this chapter, the focus is on experimental investigation and methodology for evaluation of damage tolerance or survivability and results for a particularly interesting group of aircraft with vertical take-off and landing, that is helicopters (Vignjevic & Meo, 2002). Helicopters are given special attention in this analysis as they are rather specific, highly vulnerable, and greatly exposed to threats due to their vertical take-off and landing, low speed, flight altitudes, etc. (Kelly, 2011). This chapter presents the analysis of the dynamic behaviour of a heavy transport multipurpose helicopter tail rotor blade (Figures 23.4 and 23.5) before and after ballistic damages caused by 7.9-mm-calibre shoulder weapons that may occur as a consequence of combat as well as terrorist actions (see Figure 23.6). First, non-damaged tail rotor blade behaviour was analysed by exposing the blade to static and dynamic loads in extreme flight conditions; after that, the blade was exposed to long-lasting dynamic loads to define fatigue characteristics; and finally, the blade behaviour after suffering penetrating damage was analysed on its most vital load-carrying part (that is, at the root of the spar), following the same testing program (Figures 23.7 and 23.8). For each of these tests, identical special blades made in the same two-part metal die were used (Carlsson, Adams, & Pipes, 2002; Rasuo, 1995). In the tail rotor blade, the conventional composite materials with epoxy resin matrix, fibreglass filament spar, 18-section laminated fabric skin of fibreglass filament, some carbon filament embedded along the trailing edge, a foam core, leading-edge protection strips, polyurethane, etc., were used (Rasuo, 1997, 2001, 2007).
23.2.1
Vibratory testing
The aim of the tail rotor blade vibratory testing was to determine the basic aeroelastic properties of the blade. The vibratory test program included experimental determination
Figure 23.4 Heavy transport multipurpose helicopter e Mi-8. Courtesy Moma Stanojlovic Company.
Damage tolerance and survivability of composite aircraft structures
Figure 23.5 Heavy transport multipurpose helicopter: the tail rotor blades. Courtesy Moma Stanojlovic Company.
Figure 23.6 Ballistic damage to helicopter’s windshield. Courtesy Moma Stanojlovic Company.
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Figure 23.7 Measurement points in vibratory testing.
Figure 23.8 Helicopter tail rotor blade made of composite laminated materials after ballistic damage made by the bullet of 7.9-mm-calibre shoulder weapons, in-going (left) and out-going (right) penetration.
of the natural oscillation modes and the tail blade natural frequency as well as its structural damping (Ferreira & Fasshauer, 2007; Herman, Orifici, & Mouritz, 2013). All the tests on the helicopter tail rotor blade were performed at Belgrade University, Faculty of Mechanical Engineering and Aeronautical Department Laboratory (Rasuo, 2010a,b, 2011a). The main objective of the testing to be done on non-damaged and damaged helicopter tail rotor blades was to verify experimentally the level of the blade survivability, which is defined as the degradation degree of its vital mechanical characteristics after ballistic damage. On this occasion, only one particular case was analysed: damage on the most vital load-carrying part of the blade; that is, at the root of the spar (Figure 23.8). The resulting damage, together with its in-going and out-going penetration, presents as vital structural damage on the load-carrying part in the area commonly exposed to the heaviest loads. The investigations carried out on both tail rotor blades follow a standard practice used by the majority of scientific and research aeronautical institutions (Carlsson et al., 2002).
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A very robust facility frame made of steel U- and L-profiles tied together with screws was used in the course of the tail rotor blade attachment vibratory testing programme. All the elements used in this testing are shown in Rasuo (2010a,b, 2011a). It can be clearly seen that all the elements were divided into two functional sections: The first one was composed of excitation apparatus, while the second one was made of response-detection equipment (see Figure 23.9). Measuring points were placed along the elastic axes of the blade in the same order, as shown in Figure 23.7. The excitation apparatus consisted of a pulse generator, signal amplifier, digital timer and frequency counter, vibration exciter (shaker), and an aerofoil clamp, while the response-detection group was made of piezoelectric accelerometers, oscilloscope with voltmeters, and a multichannel X-t recorder (see Figure 23.9). The link between the vibration exciter and the rotor blade was formed of a rigidly tied aluminium alloy pipe with adjustable length and by use of a panel aerofoil clamp that was shaped so as to fit the rotor blade cross-section at the location of the application of excitation (see Figure 23.10). For displacement measurement in these investigations, piezoelectric accelerometers were used; they measured displacements at selected points on the blade, as shown in Figure 23.7. A special kind of cement was used to provide a close link between the pick-up and the blade in the course of measurement (Figure 23.10).
ξ
ξ sin(ω t+ θ )
ξ
10
ξ
7
2
1
Response 3
9
8
6 F sinω t
A sinω t f
a sinω t
5
ω = ω(t) f,T
4
Excitation
1. Frame fitting 2. Rotor blade 3. Function generator 4. Digital timer/frequency counter 5. Amplifier 6. Vibration exciter (shaker) 7. Airfoil clamp 8. Piezoelectric sensor 9. Oscilloscope and waveform analysing system 10. X-t recorder
Figure 23.9 Equipment used in vibratory testing. Rasuo (2010a).
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Figure 23.10 Rotor blade in vibratory testing.
23.2.2
Fatigue testing
The fatigue test program included interlaminar separation (delamination) testing and geometric deformation of the blade cross-sections following the fatigue test program during which real tail rotor blade loads were simulated e the same loads to which the blade is exposed under extreme flight conditions, before and after ballistic damage to the blade (Rasuo, 2000, 2011b). Fatigue testing of the root part of the blade spar is one of the most important investigations of the helicopter blades made of composite laminated materials, with respect to their load-carrying ability and survivability check-ups. These tests are carried out to define likely delamination of the composite laminated structure, changes of shape of the root part of the blade, and the loss of its load-carrying ability after exposure to a certain cycle of alternating variable loads, which, on their part, are a consequence of the in-flight, combined load influence. The applied test loads include simulated steady centrifugal, vibratory chordwise bending, vibratory flapwise bending, and vibratory torsional pitch motion (Rasuo, 2011b). The simulated forces’ values were as follows: the centrifugal force was 11,350 daN, while resulting excitation alternating variable load that was vertical to the rotation
Damage tolerance and survivability of composite aircraft structures
Fitting frame
Attach fitting
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Rotor blade model
Cable
Hydraulic actuator
Variable speed drive
Eccentric mechanism
Figure 23.11 Helicopter tail rotor blade fatigue test facility.
plane and originated from vibratory chordwise bending, vibratory flapwise bending, and vibratory torsional pitch motion had a value of 500 daN at the 6.5 Hz (390 rpm) frequency and at the angle of attack of 18 . A special facility frame (Figure 23.11) was constructed to simulate these combined and heavy loads. The facility test frame used in the helicopter tail rotor blade fatigue testing was constructed as a very robust three-dimensional frame made of steel U- and L-profiles and was composed of several basic modules: the facility to which the tail rotor blade was attached and fixed, the excitation group, and modules for centrifugal force simulations (Figures 23.11 and 23.12). The excitation group consisted of an electric motor with a rating of 2.2 kW and a rotation speed of 1420 rpm, a belt drive with a transmission ratio of 1:3, a variablespeed drive (variable reduction gear) with a transmission ratio of 1e3.25, an eccentric
Figure 23.12 The excitation group.
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mechanism with an adjustable eccentricity of 0e25 mm, and an eccentric crank arm with bonded strain gages for excitation for selection (Figure 23.12). The variable transmission ratio enabled the desired excitation force frequency to be adjusted. A stroboscope was used for an accurate detection of excitation force frequency, whereas the changeable eccentricity of the eccentric arm allowed adjustment of the excitation force intensity (Rasuo, 2011b). The module for the centrifugal force simulation included a section for the static load generation, that is the centrifugal force, the section that transmitted the force to the blade root section, and the blade root attachment fitting, at which centrifugal force was applied at one end and, at the same time, excitation force was applied at the other end. A hydraulic servo-controlled actuator composed of a hydraulic cylinder, distribution system with oil lines and a pump with a servomotor, and control manometer was used as the centrifugal force generator. Its maximum force was 40,000 daN. Thanks to this system, the basic functioning of the facility frame became automatic (Rasuo, 2011b). The fatigue testing program of the root part of the tail rotor blade, aimed at the assessment of its load-carrying ability and survivability, included (in accordance with the standards) time fatigue tests together with the simulated centrifugal forcee relaxing loading program for both damaged and non-damaged blades (Guyett & Cardrick, 1980; Rasuo, 2000, 2011b). The blades were tested for time fatigue by applying the excitation force at a frequency of 6.5 Hz with simultaneous application of fullmagnitude centrifugal force in duration corresponding to 1.5 106 cycles. After every 3 105 cycles, the blade root relaxation was performed by gradually increasing and decreasing the intensity of the centrifugal force in a 0e11350e0 daN range (Figure 23.13).
Figure 23.13 Rotor blade in fatigue testing.
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23.3
651
Results: a case study
As a first step within the vibratory investigations, a harmonic analysis for both tail rotor blade types was performed. After the frequencies for the first four harmonics’ natural (resonant) modes of oscillations were determined, displacement vectors for the first four basic oscillation modes were measured. Some measurement results are shown in Figures 23.14 and 23.15. Figure 23.14 shows the harmonic analysis results of the first four oscillation harmonics for the non-damaged tail rotor blade, while Figure 23.15 gives a comparative presentation of the second oscillation mode for both a nondamaged and a damaged tail rotor blade.
Figure 23.14 Natural modes of oscillation for the non-damaged tail rotor blade.
Figure 23.15 The second oscillation mode for both non-damaged (f2 ¼ 40.2 Hz) and damaged (f 20 ¼ 39.4 Hz) tail rotor blade.
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Figure 23.16 Structural damping for both non-damaged and damaged tail rotor blade.
The tail rotor blades’ structural damping was determined from amplitude reduction of free vibrations. Initially, the blades were excited to vibrate with the first basic (resonant) oscillation mode, with gradually decreasing amplitude due to the damping effects of the structure. The original records obtained from these tests for non-damaged and damaged tail rotor blades at the measured cross-section 10 with time base 1 s/cm are shown in Figure 23.16. The logarithmic decrement of the free vibrations was utilized to characterize the structural damping diagram (Figure 23.17). Its value is determined as follows: d ¼
1 xk ln n xkþn
Figure 23.17 Logarithmic decrement and Q-factor of non-damaged tail rotor blade.
(23.1)
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where n ¼ 10 is the number of observed oscillations and xk is observed initial amplitude in the time interval, whereas the corresponding average value of amplitude is xm ¼
1 ðxk þ xkþn Þ 2
(23.2)
The Q-factor is also commonly used to define the structural damping; it gives relative energy (E) reduction in successive oscillations. The Q-factor is defined as Q ¼
E1 1 z E1 E2 2d
(23.3)
where E1 E2 is the relative energy reduction in successive oscillations (the energy dissipation in successive oscillations). The structural damping results for the non-damaged tail rotor blade expressed by the logarithmic decrement and Q-factor are given in Figure 23.17. In the course of fatigue investigations the behaviour of the blade, particularly of the damaged parts, was closely followed. No further damages or delamination of the structures (i.e. no further changes) were observed at the damaged areas in the course of the testing itself (Heida & Platenkamp, 2012; Rasuo, 2011b). When the fatigue testing program was completed on both non-damaged and damaged blades, further detailed check-ups on the deformation and degradation of the geometrical shape of the blades and delaminations were carried out. No changes were observed on either type of blade. Also, no delamination was observed and the damaged areas of the blades were not expanded in any way (Mikulik & Haase, 2012; Rasuo, 2011b).
23.4
Result analysis and discussion
The vibratory testing results were, surprisingly, the same for non-damaged and damaged heavy transport helicopter tail rotor blades. The obtained differences in frequencies and displacement vectors for some basic types of oscillation were within the 2e5% range. The structural damping results coincided to an even higher degree and the differences in the logarithmic decrement and Q-factor were less than 2% (Figure 23.17). These minor differences in basic dynamic characteristics of nondamaged and damaged helicopter tail rotor blades made of composite laminated materials cannot significantly change or endanger the helicopter flight security. The results of the damaged tail rotor blade fatigue testing have proved that such a severely damaged blade is capable of performing all its vital functions on the helicopter, even after 65 working hours in extremely difficult flight conditions (Ball, 2003; Rasuo, 2004). A very interesting and extremely important result is that the damaged blade survived the entire testing program with all and full loads relevant to the non-damaged blade. This result certainly proves the superiority of composite laminated materials for the production of aircraft vital and load-carrying parts.
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The degradation level of the vital mechanical characteristics after penetrating/ ballistic damage (Figure 23.8) is within such limits that the mission of this heavy transport helicopter (Figure 23.4) can safely be continued and extended much longer than the minimum 30-min flight prescribed by standards to reach the emergency landing site (Ball, 2003; Rasuo, 2004). Considering the results achieved in vibratory testing as well as in fatigue investigations on non-damaged and damaged tail rotor blades made of composite laminated materials, the obtained level of blade survivability is of such a nature that the aircraft could and would survive even considerably worse damages in both the root and other parts of the blade (Figures 23.15e23.17).
23.5
Conclusions
The stochastic nature of the impact that ballistic damages produce on the helicopter tail rotor blade prevents us from going into more precise and detailed quantitative analysis of the survivability level, and thus we are left only with the possibility to estimate it. The very low levels of differences in the results obtained through the investigation of vibratory characteristics, structural damping, and fatigue characteristics (less than 5%) all point to a low-level vulnerability of composite laminated materials to damages. Well-designed composite laminated structures can provide a high degree of damage tolerance, and in practice, it is still very difficult to utilize the full fibre strength potential of composite structures. That is why those materials yield constructions with exceptionally high levels of survivability e levels that are so important in both military and civil aviation.
23.6 • • • • • • • • • • •
Sources of further information and advice
ASNT e The American Society for Nondestructive Testing: https://www.asnt.org ASTM e American Society for Testing and Materials: http://www.astm.org CEN e European Committee for Standardization: https://www.cen.eu CODAMEIN e Composite Damage Metrics and Inspection, EASA: http://easa.europa.eu/ Composite Structures Damage Tolerance Analysis Methodologies: http://ntrs.nasa.gov/ archive/nasa/casi.ntrs.nasa.gov Crashworthiness of Composite Aircraft Structures: http://www.dtic.mil/dtic/ Damage tolerance of composite structures in aircraft industry: http://www.carboncomposites.eu/ ISO e International Organization for Standardization: www.iso.org Journal of Testing and Evaluation (JOTE): www.astm.org/DIGITAL_LIBRARY/ JOURNALS/TESTEVAL Probabilistic Design of Damage Tolerant Composite Aircraft Structures: http://www.faa. gov/about/office_org/headquarters_offices/ang/offices/tc/http://arc.aiaa.org Rotorcraft Structures and Survivability: http://vtol.org/events/rotorcraft-structures-andsurvivability
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Structural Composites Armour Works e Expert In Survivability: http://www.armourworks. co.uk/products/exterior-platform-protection/composites/structural-composites/ Survivability/Vulnerability Information Analysis Center: http://www.bahdayton.com/SURVIAC/index.htm The Aircraft Combat Survivability Education Website: http://www.aircraft-survivability.com
Acknowledgements The author is grateful for the funding provided by the Ministry of Education, Science and Technological Development of Republic of Serbia through Grant no. TR35006. Also, the author is grateful to the following companies: Prince Aviation, JAT Tehnika, and Moma Stanojlovic, for granted pictures.
References Ball, R. E. (2003). The fundamentals of aircraft combat survivability analysis and design (2nd ed.). New York: AIAA Inc. AIAA Education Series. Bayandor, J., Thomson, R. S., Scott, M. L., Nguyen, M. Q., & Elder, D. J. (2003). Investigation of impact and damage tolerance in advanced aerospace composite structures. International Journal of Crashworthiness, 8(3), 297e306. Carlsson, L. A., Adams, D. F., & Pipes, R. B. (2002). Experimental characterization of advanced composite materials. Cambridge, England: Woodhead Publishing Limited. Eschenfelder, P. (2005). High speed flight at low altitude: hazard to commercial aviation? In 2005 bird strike Committee-USA/Canada seventh annual meeting, Vancouver, BC, bird strike committee proceedings, 7-15-2005. University of Nebraska e Lincoln. Ferreira, A. J. M., & Fasshauer, G. E. (July 2007). Analysis of natural frequencies of composite plates by an RBF-pseudospectral method. Composite Structures, 79(2), 202e210. Gerlach, R., Siviour, C. R., Wiegand, J., & Petrinic, N. (07 February 2012). In-plane and through-thickness properties, failure modes, damage and delamination in 3D woven carbon fibre composites subjected to impact loading. Composites Science and Technology, 72(3), 397e411. Guyett, P. R., & Cardrick, A. W. (July 1980). The certification of composite airframe structures’, symposium on large-scale composite structures. Aeronautical Journal, 84(830/3), 188e203. Heida, J. H., & Platenkamp, D. J. (2012). In-service inspection guidelines for composite aerospace structures. In 18th world conference on nondestructive testing, 16e20 April 2012, Durban, South Africa (pp. 1e14). Herman, A. P., Orifici, A. C., & Mouritz, A. P. (October 2013). Vibration modal analysis of defects in composite T-stiffened panels. Composite Structures, 104, 34e42. Kelly, A. (2011). A materials scientist on: Composites, energy technology, climate changesustainable development. In Celebrating 20 years of DFC, a joint Sheffield-Cambridge conference, deformation and fracture of composites (DFC-11) & structural integrity and multi-scale modelling (SI-5), conference chairs: Costas Soutis, Alma Hodzic and Peter Beaumont, Plenary Lecture, 12e15 April 2011. Cambridge, UK: The Queens College, The University of Cambridge. Kreculj, D., & Rasuo, B. (2009). The problem of impact on aircraft constructions from composite laminates, Tehnika (Vol. LXIV). No. 6, Belgrade (in Serbian), M1eM8.
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Kreculj, D., & Rasuo, B. (June 2013). Review of impact damages modelling in laminated composite aircraft structures. Technical Gazette, 20(3), 485e495. Mikulik, Z., & Haase, P. (2012). Composite damage metrics and inspection, EASA.2010.C13 final report. Hamburg: Bishop GmbH e Aeronautical Engineers. Petit, S., Bouvet, C., Bergerot, A., & Barrau, J.-J. (2007). Impact and compression after impact experimental study of a composite laminate with a cork thermal shield. Composites Science and Technology, 67, 3286e3299. Poon, C. (1990). A review of crashworthiness of composite aircraft structures. Ottawa: National Aeronautical Establishment. Aeronautical Note, NAE-AN-63, NRC NO. 31276. Rasuo, B. (1995). Aircraft production technology. Faculty of Mechanical Engineering, University of Belgrade (in Serbian). Rasuo, B. (1997). Testing of the helicopter tail rotor blade made of composite laminated materials after ballistic damages. In M. L. Scott (Ed.), 11th international conference on composite materials, ICCM-11, Gold Coast, Queensland, Australia, proceedings volume. 2 (Impact), 485e495, Cambridge, England, Woodhead Publishing Limited, July 1997. Rasuo, B. (2000). Full-scale fatigue testing of helicopter blades from composite laminated materials. In The European conference on composite materials, ECCM-9, 4e7 June 2000, Brighton, UK, IOM Communication Ltd. (CD-ROM). Rasuo, B. (2001). Design, fabrication and testing of the helicopter tail rotor blade from composite laminated materials. In 13th international conference on composite materials, ICCM-13, Beijing, China (CD-Rom), June 2001. Rasuo, B. (2004). Aeronautical safeguarding. Belgrade: Military Academy (in Serbian). Rasuo, B. (2007). An experimental methodology for evaluating survivability of an aeronautical constructions from composite materials: an overview. International Journal of Crashworthiness, 12(1), 9e15. Woodhead Publishing Limited, Abington Hall, Abington, Cambridge (Taylor & Francis, London). Rasuo, B. (2010a). Experimental study of the structural damping of composite helicopter blades with different cores. Plastics Rubber and Composites, 39(1), 1e5. Maney Publishing, Institute of Materials, Minerals and Mining, London. Rasuo, B. (2010b). Design, fabrication and testing of the helicopter tail rotor blade from composite materials: an experience, the Royal aeronautical society. In 2010 RAeS structures & materials conference e 2nd aircraft structural design conference, 26e28 October 2010, London, UK (CD-Rom). Rasuo, B. (2011a). An experimental methodology for evaluating survivability or damage tolerance of an aeronautical construction from composite materials. In Celebrating 20 years of DFC, a joint Sheffield-Cambridge conference, deformation and fracture of composites (DFC-11) & structural integrity and multi-scale modelling (SI-5), conference chairs: Costas Soutis, Alma Hodzic and Peter Beaumont, 12e15 April 2011. Cambridge, UK: The Queens College, The University of Cambridge. Rasuo, B. (March 2011b). Experimental techniques for evaluation of fatigue characteristics of laminated constructions from composite materials: full-scale testing of the helicopter rotor blades. Journal of Testing and Evaluation (JTE), 39(2), 237e242. ASTM International, USA. Raymer, D. P. (1999). Aircraft design: A conceptual approach (3rd ed.). Washington: AIAA Inc. AIAA Education Series. Richards, M. G., Hastings, D. E., Rhodes, D. H., Ross, A. M., & Weigel, A. L. (2009). Design for survivability: concept generation and evaluation in dynamic trade space exploration. In Second international symposium on engineering systems, MIT, Cambridge, MA, 15e17 June 2009.
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Schwarz, C. R., & Drake, H. (May 2001). Survivability engineering. In Aerospace systems survivability handbook series, joint technical coordinating group on aircraft survivability (JTCG/AS) (Vol. 4). JTCG/AS-01-D-005, Arlington, VA. Soutis, C. (2005). Fibre reinforced composites in aircraft construction. Progress in Aerospace Sciences, 41(1), 43e51. Soutis, C., & Beaumont, P. W. R. (Eds.). (2005). Multi-scale modelling of composite material systems. Cambridge, England: Woodhead Publishing Limited. Sultan, M. T. H., Worden, K., Staszewski, W. J., & Hodzic, A. (08 June 2012). Impact damage characterisation of composite laminates using a statistical approach. Composites Science and Technology, 72(10), 1108e1120. Vignjevic, R., & Meo, M. (2002). A new concept of a helicopter sub-floor structure crashworthy in impacts on water and rigid surfaces. International Journal of Crashworthiness, 7(3), 321e331.
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B. Rasuo University of Belgrade, Belgrade, Serbia
23.1
Introduction
The analysis of the dynamic behaviour of damaged constructions can provide knowledge about the acceptable level of damages and likely chances for aircraft survival. The aircraft’s ability to survive even after exposure to severe damage to the vital and loadcarrying parts of the aircraft construction is imperative not only for combat aircraft but also for civil aircraft (Ball, 2003; Raymer, 1999; Richards, Hastings, Rhodes, Ross, & Weigel, 2009; Schwarz & Drake, 2001). Vulnerability, as an element of survivability, is one of the most important exploitable characteristics of contemporary aircraft. With respect to survivability, composite laminated materials demonstrate the best behaviour and results compared to other materials that are currently used in aviation (Rasuo, 1995). Requirements and guidelines for the general programme are contained in the standards MIL-STD-2069, FAR/JAR 25.571, and, for bird strike requirements, FAR/JAR 25.631 (Eschenfelder, 2005; Petit, Bouvet, Bergerot, & Barrau, 2007; Rasuo, 2004). Predicting damage in laminated composite aircraft components due to impact events such as runway debris, hail, and birds is an area of ongoing research (see Figures 23.1e23.3). To reduce certification and development costs, computational methods are required by the aircraft industry to be able to predict structural integrity of composite structures under high-velocity impacts from hard objects, such as metal fragments and stone debris, and from soft or deformable bodies such as birds, hailstones, and tyre rubber. Key issues are the development of suitable constitutive laws for modelling composites in ply, determination of composite parameters from dynamic materials tests, materials laws for deformable impactors, and the efficient implementation of the materials models into finite element codes (Soutis, 2005; Soutis & Beaumont, 2005). Impact damages in aircraft structures made from laminated composites are very complex; the most common are matrix cracking, fibre failure, and delamination (Gerlach, Siviour, Wiegand, & Petrinic, 2012). The ability to predict the initiation and growth of damage is crucial for predicting performance and developing reliable and safe designs of composites (Sultan, Worden, Staszewski, & Hodzic, 2012). By use of simulation in modelling the impact damage, the test costs of aircraft structures made from composite laminates will be reduced (Kreculj & Rasuo, 2009). Structural Integrity and Durability of Advanced Composites. http://dx.doi.org/10.1016/B978-0-08-100137-0.00023-7 Copyright © 2015 Elsevier Ltd. All rights reserved.
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Figure 23.1 Impact-induced mechanical damage of inlet cone of a MiG-21bis fighter aircraft. Courtesy Moma Stanojlovic Company.
Figure 23.2 Hail strike-induced mechanical damage on nose cone of a Cessna 650 aeroplane. Courtesy Prince Aviation Company.
Due to the wide-scale use of composite materials in different aircraft structures, it is necessary to introduce new approaches for impact damage modelling. Efficient methodologies are modelling composite structures by applying specialized finite element methods that take into account macromechanical structural properties, as well as by using numerical methods with complicated analysis codes. In order to validate those methodologies for further use in the structures’ strength design, numerical model verification has to be done. A comparison of the obtained experimental and numerical results is further made to determine if the proposed models are accurate and valid (Kreculj & Rasuo, 2013).
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Figure 23.3 Impact-induced mechanical damage of the thrust reverser door on a Boeing 737-700. Courtesy JAT Tehnika Company.
Composite damages exist at the micro-scale level, while impact loads are applied at the structural level. Hence, it is necessary to consider a multi-scale approach for that kind of problem (Soutis & Beaumont, 2005). The development of suitable constitutive laws for modelling composite laminate failures and material models with finite element codes under impact provides significant assistance in the design and exploitation phases of specified structures. The most well-known failure criteria (two- and three-dimensional) for composite materials are Tsai-Wu, ChangeChang, and Hashin. They are used to predict the level or degree of damage, fracture, and failure of composite structures (Soutis & Beaumont, 2005). Development of computational models and simulations is necessary in studying the onset and growth of impact damages. In numerical simulation of impacts on composite structures, two- or three-dimensional models exist. For impact damage analysis of composite laminates, commercial software ABAQUS and LS-DYNA are most frequently used. Pro/ENGINEER, ANSYS, and some noncommercial software is also suitable for the same purpose. By using it in some circumstances (geometry, boundary conditions, mesh, load, etc.), the distribution of damage, stress, strain, and deformations can be analysed and presented (Kreculj & Rasuo, 2013). Essentially, when considering impact damage modelling in laminated composites of aeronautical structures, the dynamic response of composite laminates under impact is investigated and damage initiation/growth in such structures is predicted (Guyett & Cardrick, 1980; Poon, 1990). For such analysis, numerical modelling and simulations provide important and valuable results. Due to the anisotropy of composite laminates and nonuniform distribution of stress under dynamic loading, the failure process of laminates is very complex. The dynamic response of composite structures subjected to transient dynamic loading has been studied for years in terms of analytical, numerical, and experimental works (Bayandor, Thomson, Scott, Nguyen, & Elder, 2003).
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Experimental methodology for evaluation of damage tolerance and survivability
Taking into account the quantum of considered problems of damage tolerance or survivability of composite aircraft structures in this chapter, the focus is on experimental investigation and methodology for evaluation of damage tolerance or survivability and results for a particularly interesting group of aircraft with vertical take-off and landing, that is helicopters (Vignjevic & Meo, 2002). Helicopters are given special attention in this analysis as they are rather specific, highly vulnerable, and greatly exposed to threats due to their vertical take-off and landing, low speed, flight altitudes, etc. (Kelly, 2011). This chapter presents the analysis of the dynamic behaviour of a heavy transport multipurpose helicopter tail rotor blade (Figures 23.4 and 23.5) before and after ballistic damages caused by 7.9-mm-calibre shoulder weapons that may occur as a consequence of combat as well as terrorist actions (see Figure 23.6). First, non-damaged tail rotor blade behaviour was analysed by exposing the blade to static and dynamic loads in extreme flight conditions; after that, the blade was exposed to long-lasting dynamic loads to define fatigue characteristics; and finally, the blade behaviour after suffering penetrating damage was analysed on its most vital load-carrying part (that is, at the root of the spar), following the same testing program (Figures 23.7 and 23.8). For each of these tests, identical special blades made in the same two-part metal die were used (Carlsson, Adams, & Pipes, 2002; Rasuo, 1995). In the tail rotor blade, the conventional composite materials with epoxy resin matrix, fibreglass filament spar, 18-section laminated fabric skin of fibreglass filament, some carbon filament embedded along the trailing edge, a foam core, leading-edge protection strips, polyurethane, etc., were used (Rasuo, 1997, 2001, 2007).
23.2.1
Vibratory testing
The aim of the tail rotor blade vibratory testing was to determine the basic aeroelastic properties of the blade. The vibratory test program included experimental determination
Figure 23.4 Heavy transport multipurpose helicopter e Mi-8. Courtesy Moma Stanojlovic Company.
Damage tolerance and survivability of composite aircraft structures
Figure 23.5 Heavy transport multipurpose helicopter: the tail rotor blades. Courtesy Moma Stanojlovic Company.
Figure 23.6 Ballistic damage to helicopter’s windshield. Courtesy Moma Stanojlovic Company.
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Figure 23.7 Measurement points in vibratory testing.
Figure 23.8 Helicopter tail rotor blade made of composite laminated materials after ballistic damage made by the bullet of 7.9-mm-calibre shoulder weapons, in-going (left) and out-going (right) penetration.
of the natural oscillation modes and the tail blade natural frequency as well as its structural damping (Ferreira & Fasshauer, 2007; Herman, Orifici, & Mouritz, 2013). All the tests on the helicopter tail rotor blade were performed at Belgrade University, Faculty of Mechanical Engineering and Aeronautical Department Laboratory (Rasuo, 2010a,b, 2011a). The main objective of the testing to be done on non-damaged and damaged helicopter tail rotor blades was to verify experimentally the level of the blade survivability, which is defined as the degradation degree of its vital mechanical characteristics after ballistic damage. On this occasion, only one particular case was analysed: damage on the most vital load-carrying part of the blade; that is, at the root of the spar (Figure 23.8). The resulting damage, together with its in-going and out-going penetration, presents as vital structural damage on the load-carrying part in the area commonly exposed to the heaviest loads. The investigations carried out on both tail rotor blades follow a standard practice used by the majority of scientific and research aeronautical institutions (Carlsson et al., 2002).
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A very robust facility frame made of steel U- and L-profiles tied together with screws was used in the course of the tail rotor blade attachment vibratory testing programme. All the elements used in this testing are shown in Rasuo (2010a,b, 2011a). It can be clearly seen that all the elements were divided into two functional sections: The first one was composed of excitation apparatus, while the second one was made of response-detection equipment (see Figure 23.9). Measuring points were placed along the elastic axes of the blade in the same order, as shown in Figure 23.7. The excitation apparatus consisted of a pulse generator, signal amplifier, digital timer and frequency counter, vibration exciter (shaker), and an aerofoil clamp, while the response-detection group was made of piezoelectric accelerometers, oscilloscope with voltmeters, and a multichannel X-t recorder (see Figure 23.9). The link between the vibration exciter and the rotor blade was formed of a rigidly tied aluminium alloy pipe with adjustable length and by use of a panel aerofoil clamp that was shaped so as to fit the rotor blade cross-section at the location of the application of excitation (see Figure 23.10). For displacement measurement in these investigations, piezoelectric accelerometers were used; they measured displacements at selected points on the blade, as shown in Figure 23.7. A special kind of cement was used to provide a close link between the pick-up and the blade in the course of measurement (Figure 23.10).
ξ
ξ sin(ω t+ θ )
ξ
10
ξ
7
2
1
Response 3
9
8
6 F sinω t
A sinω t f
a sinω t
5
ω = ω(t) f,T
4
Excitation
1. Frame fitting 2. Rotor blade 3. Function generator 4. Digital timer/frequency counter 5. Amplifier 6. Vibration exciter (shaker) 7. Airfoil clamp 8. Piezoelectric sensor 9. Oscilloscope and waveform analysing system 10. X-t recorder
Figure 23.9 Equipment used in vibratory testing. Rasuo (2010a).
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Figure 23.10 Rotor blade in vibratory testing.
23.2.2
Fatigue testing
The fatigue test program included interlaminar separation (delamination) testing and geometric deformation of the blade cross-sections following the fatigue test program during which real tail rotor blade loads were simulated e the same loads to which the blade is exposed under extreme flight conditions, before and after ballistic damage to the blade (Rasuo, 2000, 2011b). Fatigue testing of the root part of the blade spar is one of the most important investigations of the helicopter blades made of composite laminated materials, with respect to their load-carrying ability and survivability check-ups. These tests are carried out to define likely delamination of the composite laminated structure, changes of shape of the root part of the blade, and the loss of its load-carrying ability after exposure to a certain cycle of alternating variable loads, which, on their part, are a consequence of the in-flight, combined load influence. The applied test loads include simulated steady centrifugal, vibratory chordwise bending, vibratory flapwise bending, and vibratory torsional pitch motion (Rasuo, 2011b). The simulated forces’ values were as follows: the centrifugal force was 11,350 daN, while resulting excitation alternating variable load that was vertical to the rotation
Damage tolerance and survivability of composite aircraft structures
Fitting frame
Attach fitting
649
Rotor blade model
Cable
Hydraulic actuator
Variable speed drive
Eccentric mechanism
Figure 23.11 Helicopter tail rotor blade fatigue test facility.
plane and originated from vibratory chordwise bending, vibratory flapwise bending, and vibratory torsional pitch motion had a value of 500 daN at the 6.5 Hz (390 rpm) frequency and at the angle of attack of 18 . A special facility frame (Figure 23.11) was constructed to simulate these combined and heavy loads. The facility test frame used in the helicopter tail rotor blade fatigue testing was constructed as a very robust three-dimensional frame made of steel U- and L-profiles and was composed of several basic modules: the facility to which the tail rotor blade was attached and fixed, the excitation group, and modules for centrifugal force simulations (Figures 23.11 and 23.12). The excitation group consisted of an electric motor with a rating of 2.2 kW and a rotation speed of 1420 rpm, a belt drive with a transmission ratio of 1:3, a variablespeed drive (variable reduction gear) with a transmission ratio of 1e3.25, an eccentric
Figure 23.12 The excitation group.
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mechanism with an adjustable eccentricity of 0e25 mm, and an eccentric crank arm with bonded strain gages for excitation for selection (Figure 23.12). The variable transmission ratio enabled the desired excitation force frequency to be adjusted. A stroboscope was used for an accurate detection of excitation force frequency, whereas the changeable eccentricity of the eccentric arm allowed adjustment of the excitation force intensity (Rasuo, 2011b). The module for the centrifugal force simulation included a section for the static load generation, that is the centrifugal force, the section that transmitted the force to the blade root section, and the blade root attachment fitting, at which centrifugal force was applied at one end and, at the same time, excitation force was applied at the other end. A hydraulic servo-controlled actuator composed of a hydraulic cylinder, distribution system with oil lines and a pump with a servomotor, and control manometer was used as the centrifugal force generator. Its maximum force was 40,000 daN. Thanks to this system, the basic functioning of the facility frame became automatic (Rasuo, 2011b). The fatigue testing program of the root part of the tail rotor blade, aimed at the assessment of its load-carrying ability and survivability, included (in accordance with the standards) time fatigue tests together with the simulated centrifugal forcee relaxing loading program for both damaged and non-damaged blades (Guyett & Cardrick, 1980; Rasuo, 2000, 2011b). The blades were tested for time fatigue by applying the excitation force at a frequency of 6.5 Hz with simultaneous application of fullmagnitude centrifugal force in duration corresponding to 1.5 106 cycles. After every 3 105 cycles, the blade root relaxation was performed by gradually increasing and decreasing the intensity of the centrifugal force in a 0e11350e0 daN range (Figure 23.13).
Figure 23.13 Rotor blade in fatigue testing.
Damage tolerance and survivability of composite aircraft structures
23.3
651
Results: a case study
As a first step within the vibratory investigations, a harmonic analysis for both tail rotor blade types was performed. After the frequencies for the first four harmonics’ natural (resonant) modes of oscillations were determined, displacement vectors for the first four basic oscillation modes were measured. Some measurement results are shown in Figures 23.14 and 23.15. Figure 23.14 shows the harmonic analysis results of the first four oscillation harmonics for the non-damaged tail rotor blade, while Figure 23.15 gives a comparative presentation of the second oscillation mode for both a nondamaged and a damaged tail rotor blade.
Figure 23.14 Natural modes of oscillation for the non-damaged tail rotor blade.
Figure 23.15 The second oscillation mode for both non-damaged (f2 ¼ 40.2 Hz) and damaged (f 20 ¼ 39.4 Hz) tail rotor blade.
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Figure 23.16 Structural damping for both non-damaged and damaged tail rotor blade.
The tail rotor blades’ structural damping was determined from amplitude reduction of free vibrations. Initially, the blades were excited to vibrate with the first basic (resonant) oscillation mode, with gradually decreasing amplitude due to the damping effects of the structure. The original records obtained from these tests for non-damaged and damaged tail rotor blades at the measured cross-section 10 with time base 1 s/cm are shown in Figure 23.16. The logarithmic decrement of the free vibrations was utilized to characterize the structural damping diagram (Figure 23.17). Its value is determined as follows: d ¼
1 xk ln n xkþn
Figure 23.17 Logarithmic decrement and Q-factor of non-damaged tail rotor blade.
(23.1)
Damage tolerance and survivability of composite aircraft structures
653
where n ¼ 10 is the number of observed oscillations and xk is observed initial amplitude in the time interval, whereas the corresponding average value of amplitude is xm ¼
1 ðxk þ xkþn Þ 2
(23.2)
The Q-factor is also commonly used to define the structural damping; it gives relative energy (E) reduction in successive oscillations. The Q-factor is defined as Q ¼
E1 1 z E1 E2 2d
(23.3)
where E1 E2 is the relative energy reduction in successive oscillations (the energy dissipation in successive oscillations). The structural damping results for the non-damaged tail rotor blade expressed by the logarithmic decrement and Q-factor are given in Figure 23.17. In the course of fatigue investigations the behaviour of the blade, particularly of the damaged parts, was closely followed. No further damages or delamination of the structures (i.e. no further changes) were observed at the damaged areas in the course of the testing itself (Heida & Platenkamp, 2012; Rasuo, 2011b). When the fatigue testing program was completed on both non-damaged and damaged blades, further detailed check-ups on the deformation and degradation of the geometrical shape of the blades and delaminations were carried out. No changes were observed on either type of blade. Also, no delamination was observed and the damaged areas of the blades were not expanded in any way (Mikulik & Haase, 2012; Rasuo, 2011b).
23.4
Result analysis and discussion
The vibratory testing results were, surprisingly, the same for non-damaged and damaged heavy transport helicopter tail rotor blades. The obtained differences in frequencies and displacement vectors for some basic types of oscillation were within the 2e5% range. The structural damping results coincided to an even higher degree and the differences in the logarithmic decrement and Q-factor were less than 2% (Figure 23.17). These minor differences in basic dynamic characteristics of nondamaged and damaged helicopter tail rotor blades made of composite laminated materials cannot significantly change or endanger the helicopter flight security. The results of the damaged tail rotor blade fatigue testing have proved that such a severely damaged blade is capable of performing all its vital functions on the helicopter, even after 65 working hours in extremely difficult flight conditions (Ball, 2003; Rasuo, 2004). A very interesting and extremely important result is that the damaged blade survived the entire testing program with all and full loads relevant to the non-damaged blade. This result certainly proves the superiority of composite laminated materials for the production of aircraft vital and load-carrying parts.
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Structural Integrity and Durability of Advanced Composites
The degradation level of the vital mechanical characteristics after penetrating/ ballistic damage (Figure 23.8) is within such limits that the mission of this heavy transport helicopter (Figure 23.4) can safely be continued and extended much longer than the minimum 30-min flight prescribed by standards to reach the emergency landing site (Ball, 2003; Rasuo, 2004). Considering the results achieved in vibratory testing as well as in fatigue investigations on non-damaged and damaged tail rotor blades made of composite laminated materials, the obtained level of blade survivability is of such a nature that the aircraft could and would survive even considerably worse damages in both the root and other parts of the blade (Figures 23.15e23.17).
23.5
Conclusions
The stochastic nature of the impact that ballistic damages produce on the helicopter tail rotor blade prevents us from going into more precise and detailed quantitative analysis of the survivability level, and thus we are left only with the possibility to estimate it. The very low levels of differences in the results obtained through the investigation of vibratory characteristics, structural damping, and fatigue characteristics (less than 5%) all point to a low-level vulnerability of composite laminated materials to damages. Well-designed composite laminated structures can provide a high degree of damage tolerance, and in practice, it is still very difficult to utilize the full fibre strength potential of composite structures. That is why those materials yield constructions with exceptionally high levels of survivability e levels that are so important in both military and civil aviation.
23.6 • • • • • • • • • • •
Sources of further information and advice
ASNT e The American Society for Nondestructive Testing: https://www.asnt.org ASTM e American Society for Testing and Materials: http://www.astm.org CEN e European Committee for Standardization: https://www.cen.eu CODAMEIN e Composite Damage Metrics and Inspection, EASA: http://easa.europa.eu/ Composite Structures Damage Tolerance Analysis Methodologies: http://ntrs.nasa.gov/ archive/nasa/casi.ntrs.nasa.gov Crashworthiness of Composite Aircraft Structures: http://www.dtic.mil/dtic/ Damage tolerance of composite structures in aircraft industry: http://www.carboncomposites.eu/ ISO e International Organization for Standardization: www.iso.org Journal of Testing and Evaluation (JOTE): www.astm.org/DIGITAL_LIBRARY/ JOURNALS/TESTEVAL Probabilistic Design of Damage Tolerant Composite Aircraft Structures: http://www.faa. gov/about/office_org/headquarters_offices/ang/offices/tc/http://arc.aiaa.org Rotorcraft Structures and Survivability: http://vtol.org/events/rotorcraft-structures-andsurvivability
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Structural Composites Armour Works e Expert In Survivability: http://www.armourworks. co.uk/products/exterior-platform-protection/composites/structural-composites/ Survivability/Vulnerability Information Analysis Center: http://www.bahdayton.com/SURVIAC/index.htm The Aircraft Combat Survivability Education Website: http://www.aircraft-survivability.com
Acknowledgements The author is grateful for the funding provided by the Ministry of Education, Science and Technological Development of Republic of Serbia through Grant no. TR35006. Also, the author is grateful to the following companies: Prince Aviation, JAT Tehnika, and Moma Stanojlovic, for granted pictures.
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Schwarz, C. R., & Drake, H. (May 2001). Survivability engineering. In Aerospace systems survivability handbook series, joint technical coordinating group on aircraft survivability (JTCG/AS) (Vol. 4). JTCG/AS-01-D-005, Arlington, VA. Soutis, C. (2005). Fibre reinforced composites in aircraft construction. Progress in Aerospace Sciences, 41(1), 43e51. Soutis, C., & Beaumont, P. W. R. (Eds.). (2005). Multi-scale modelling of composite material systems. Cambridge, England: Woodhead Publishing Limited. Sultan, M. T. H., Worden, K., Staszewski, W. J., & Hodzic, A. (08 June 2012). Impact damage characterisation of composite laminates using a statistical approach. Composites Science and Technology, 72(10), 1108e1120. Vignjevic, R., & Meo, M. (2002). A new concept of a helicopter sub-floor structure crashworthy in impacts on water and rigid surfaces. International Journal of Crashworthiness, 7(3), 321e331.