Fuzzy-Logic-Based Health Monitoring and Residual-Life Prediction for Composite Helicopter Rotor (original) (raw)
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On the effect of progressive damage on composite helicopter rotor system behavior
Composite structures, 2007
Composite rotor blades are routinely used in modern helicopters. Damage in composite helicopter rotor blades is investigated in this work. Effects of the key damage modes in composite materials such as matrix cracking, debonding/delamination and fiber breakage on various properties of the composite rotor blade such as stiffnesses, frequencies, deflection, root forces, root moments and strains in forward flight are studied using an aeroelastic analysis. The composite rotor blade is modeled as a thin walled composite beam and includes the effect of transverse shear, elastic couplings and restrained warping. Matrix cracking is modeled at the laminate level and debonding/ delamination and fiber breakage at the lamina level and included in the formulation by adjusting the A, B and D matrices for composite laminates. A stiff in-plane rotor blade with a two-cell airfoil section with [0/±45/90] s family of laminates is considered. An aeroelastic analysis of the helicopter rotor based on finite elements in space and time is used to study the effects of key damage modes in a composite rotor in forward flight.
On the effect of matrix cracks in composite helicopter rotor blade
Composites science and technology, 2005
This paper investigates the feasibility of an on-line damage detection capability for helicopter main rotor blades made of composite material. Damage modeled in the composite is matrix cracking. A box-beam with stiffness properties similar to a hingeless rotor blade is designed using genetic algorithm for the typical [±h m /90 n ] s family of composites. The effect of matrix cracks is included in an analytical model of composite box-beam. An aeroelastic analysis of the helicopter rotor based on finite elements in space and time is used to study the effects of matrix cracking in the rotor blade in forward flight. For global fault detection, rotating frequencies, tip bending and torsion response, and blade root loads are studied. It is observed that the effect of matrix cracking on lag bending and elastic twist deflection at the blade tip and blade root yawing moment is significant and these parameters can be monitored for online health monitoring. For implementation of local fault detection technique, the effect on axial and shear strain, for matrix cracks in the whole blade as well as matrix cracks occurring locally is studied. It is observed that using strain measurement along the blade it is possible to locate the matrix cracks as well as to predict density of matrix cracks.
Fatigue of Aircraft Structures, 2019
The fatigue test was carried out on an element of a rotor blade removed from the Mi-2 helicopter. The purpose of the test was to check the fatigue strength of the repaired rotor blade. Metal composite rotor blades have a metal spar in the form of a box and the trailing sections in the form of metallic honeycomb sandwich panels. The trailing sections are bonded to the spar. The repair had been carried out at the point where the trailing section became debonded from the spar at the Air Force Institute of Technology in Warsaw using a methodology developed for carrying out repairs of rotor blades’ damage. All types of the Mi family helicopters are equipped with metal composite rotors blades. Depending on MTOW (Maximum Take-Off Weight) and destination of helicopters, blades differ in dimensions, but their design solutions are practically the same. For this reason, the developed repair methodology can be used for all characteristic rotor blades structures for Mi helicopters. The fatigue t...
Progressive Failure Analysis of Helicopter Rotor Blade Under Aeroelastic Loading
Aviation, 2020
Unlike metal structure, composite structures don’t give any clue till the fatal final collapse. The problem is more complicated when applied load on the structure is aeroelastic in nature. Under such loading, composite laminate experiences stresses. The first layer failure happens when stresses in the weakest ply exceed the allowable strength of the laminate. This initial layer-based failure changes overall material characteristics. It is important now to degrade the composite laminate characteristics for the subsequent failure prediction. The constitutive relations are required to be updated by the reduction in stiffness. The rest of the undamaged laminates continue to take the load till the updated strength is reached. In the present work, layer wise progressive failure analysis under aeroelastic loading has been performed by the inclusion of different failure criteria which allow for the identification of the location of the failure. ANSYS APDL environment has been used to model ...
AGUSTA Experience on Damage Tolerance Evaluation of Helicopter Components
2000
Within the fatigue evaluation of the EHIOI, Agusta has MAIN AND TAIL ROTORS carried out a specific program of flaw tolerance evaluation of the primary loading path. The program is close to completion The Main Rotor Head is made by the Main Rotor Hub, with a and this paper provides a summary of the most relevant titanium core and carbon fibre-epoxy loop windings enclosed results. by glass-epoxy box. Two different configurations of blade For composite components, damage size was increased grips can be used: the foldable configuration for naval considering both manufacturing discrepancies greater than the variants, made by Inboard and Outboard Tension Links, with minimum quality standard and impact damages clearly hybrid titanium frames and composite plates, and the 'onedetectable during visual inspections. The favourable data piece' design Civil Tension Link made by a full carbon and achieved are based on the 'no growth' concept. glass fibre-epoxy laminate. Most of the Main Rotor Blade The metal parts of the main rotor head were evaluated by shear loads are reacted by the titanium Support Cone, via the enhanced safe life method and fail safe capability. The slow Elastomeric Bearings. crack growth approach was instead applied for the Rear The Tail Rotor is a 'see-saw' made by carbon fibre-epoxy Fuselage End Fittings, which connect the Tail Unit. Blades and the composite Hub, made by carbon fibre loops All these evidences can be used in addition to the windings with glass fibre-epoxy wrap. comprehensive safe life evaluation of the aircraft to improve Both Main and Tail Rotor have Elastomeric Bearings, the maintenance and the repair actions, providing relevant damage tolerance capabilities. Based on this experience, application of flaw tolerance criteria will be carried out on the new helicopters in development All the composite parts were evaluated according to AC 20phase. 107A, proving a flaw tolerance demonstration considering:
Study of impact on helicopter blade
Engineering Failure Analysis, 2012
This article presents a study of damage in structures that are similar to helicopter blade sections, subjected to an impact. These complex composite structures were impacted by a steel ball of 125 g at impact speed ranging from 30 to 130 m/s. This led to properly highlight the kinematics of the impact and to define the sequence of the damage's mechanisms. An explicit FE model is also presented. The damage modelling of the roving is performed through a scale change. It allows a good representation of observed experimental behaviour. As the mesh density is low, it can be used for the modelling of a real structure.
Application of Damage Tolerance principles to the design of helicopters
International Journal of Fatigue, 2009
Helicopters are very specific aircrafts, for which the design against fatigue phenomena is a particularly important and complex problem. The peculiarity lays in the load spectrum that is composed by a high number of low-amplitude cycles, which result from the mechanical rotation of the rotor blades. The fatigue design methodology most commonly applied by the helicopter community was based on the Safe-Life philosophy, applied anyhow with a particular approach. Since 1989, the Airworthiness Regulations evolved towards the application of Damage Tolerance principles also to rotorcraft. This change has forced the helicopter industries to review their design methodologies, and to face new problems, linked with fracture mechanics applications to their typical structures. Flaws, accidental damages and manufacturing discrepancies must be accounted for, in addition to the retirement life based on Safe Life. Two approaches are mainly used to improve the fatigue assessment: establish the retirement life of parts considering the possibility of initial flaws or cracks of realistic size, or supplement retirement lives by inspection plans defined on the basis of test and analysis. If safety is assured through inspections, with intervals defined by means of crack growth analysis, it is obvious that more refined crack growth models are necessary, because most of those currently used in the fixed wing industry are based on semi-empirical basis and contain hidden tuning factors that are related to the typical aeroplanes spectra. Moreover, unnecessary conservatism gives rise to very short inspection intervals, that cannot be practically implemented by operators in an inspection program. The paper reviews the current requirements and presents and discusses the methodologies that the helicopter industry adopts for demonstrating compliance with such regulatory requirements. In addition, recommendations are given on research and development activities required for refining the defined methodologies and for their better implementation.
Modeling progressive damage accumulation in thin walled composite beams for rotor blade applications
Composites science and technology, 2006
In the present study, models of the key damage modes in composite materials such as matrix cracking, debonding/delamination and fiber breakage are included in a thin walled composite beam analysis for helicopter rotor blade applications. The effects of transverse shear, elastic couplings and restrained warping are also included while modeling the thin walled composite beams. Matrix cracking is modeled at the laminate level and debonding/delamination and fiber breakage at the lamina level and included in the formulation by adjusting the A, B and D matrices for composite laminates. The beam analysis is used to investigate the effect of damage on the various properties of thin walled composite beams such as out-of-plane stiffness, in-plane stiffness and torsion stiffness and deflections such as bending slopes and twist under an applied load. A two-cell airfoil section with [0/±45/90] s family of laminates having stiffness properties equivalent to a stiff in-plane hingeless rotor blade is considered. It is found that the bending and torsion stiffness loss due to matrix cracking is about 6-12% and 25-30%, respectively, and due to debonding/delamination is further 6-8% and 40-45%, respectively. Most of the bending stiffness loss is observed in the fiber breakage damage mode. It is also observed that the combination of the static tip bending and torsion response can be used to predict length of the damaged part of the thin walled composite beams, damage location and damage level.