AEROELASTIC STABILITY OF SMART TRAILING-EDGE FLAP HELICOPTER ROTORS

Introduction This paper investigates the aeroelastic stability of a smart helicopter rotor blade with trailing edge flaps. The coupled blade/flap/actuator equations were linearized by using small perturbation motions about a steady trimmed solution. Stability is then determined from an eigenanalysis of these homogeneous equations using a constant coefficient approximation approach and the Floquet method. The baseline correlation of stability calculations without trailing-edge flaps is carried out with wind tunnel test data and DART simulation for a typical 5-bladed bearingless rotor system. The correlation is good for both hover and forward flight conditions. Stability calculations for a rotor with trailing-edge flaps were conducted to examine the effect of flap aerodynamic balance and flap mass-balance. It is shown that both flap aerodynamic over-balancing and flap massunbalancing will have a destabilizing effect on the blade and flap system. The effect of rotor control system stiffness, blade torsion stiffness, actuator stiffness as well as flap mass on blade stability was studied. It is concluded that decreasing of either blade torsion stiffness or actuator stiffness will cause blade torsion-flap flutter problems. Increasing either rotor control system stiffness or flap mass would destabilize the blade/flap system. * Graduate Research Assistant, Student Member AHS t Professor and Director, Fellow AIAA, Fellow AHS Presented at the 42nd Structures, Structural Dynamics and Materials Conference and Adaptive Structures Forum, Seattle, WA, April 16-19, 2001. Copyright © 2001 by J. Shen, and I. Chopra. Published by the American Institute of Aeronautics and Astronautics, Inc. with permission. Helicopters are susceptible to excessive vibration because of the unsteady aerodynamic environment at the rotor disk, nonlinear inertial couplings of slender rotating blades, and complex rotorfuselage interaction effects. A high level of vibration causes fatigue failure of components and human discomfort, seriously affects ride quality and system reliability, increases maintenance costs and degrades equipment performance. Design of helicopters with inherently low vibration is therefore an important goal. Traditional, passive vibration absorbers and islators are routinely used to tackle the vibration problem. However, they incur a large weight penalty and exhibit poor offdesign performance, particularly because of the wide flight envelope of the helicopter. In order to improve vibration-reduction performance across the flight envelope, active control strategies need to be considered. Recent emergence of smart material actuators have opened the domain of smart vibration reduction techniques. Chopra [1] reviews the status of smart structure technology development for application to rotor craft systems. These actuation systems are expected to be compact, light weight, low actuation power, and high band-width devices that can be used for multi-functional roles such as to suppress vibration and noise, increase aeromechanical stability, and improve rotor performance. A large variety of Proude scale [2-5] and Mach scale [6,7] smart rotor models with trailing edge flaps have been tested for some of these qualities. The flap actuators range from piezo-bimorhs, piezo/electrostrictive stacks and piezo/magnetostrictive-induced compositecoupled actuation. Although the research in the application of smart flap rotor systems is American Institute of Aeronautics and Astronautics (c)2001 American Institute of Aeronautics & Astronautics or Published with Permission of Author(s) and/or Author(s)' Sponsoring Organization. in its early phase, it has been demonstrated its capability by various experiment to minimize helicopter vibration by dynamically modifying the aerodynamic loading along the span of blade. To realize such an active control system efficiently on a helicopter rotor, it is important to develop a comprehensive rotor analysis and carry out systematic parametric studies that include aeroelastic stability investigation. The modeling issues of rotor systems with trailing edge flaps to reduce vibration were examined by Millot and Priedmann [8], and Milgram, Chopra and Straub [9]. These analyses, however, assumed prescribed flap deflections and hence neglect actuator dynamics. Recently, the authors [10] presented a comprehensive analysis for blade and flap response and vibration reduction of a rotor with trailing-edge flaps that included actuator dynamics. Coupled actuator, trailing edge flap, and elastic rotor blade equations of motion were formulated using Hamilton's variational principle. These nonlinear, periodic equations were solved using finite elements in space and time. The correlation study for a typical 5-bladed bearingless rotor system (baseline, without flaps) analysis was carried out with wind tunnel test data, and good agreement for blade vibratory flap and lag bending, and torsional moments was observed. Trailing-edge flap modeling was also correlated with CAMRAD II and CAMRAD/JA simulation without including actuator dynamics; the correlation was generally good except for hinge moments perhaps because of different aerodynamic tables used. The effect of actuator dynamics was examined on blade and trailing-edge flap responses, and it was shown that actuator dynamics is important for the calculation of blade loads, especially for a torsionally soft smart actuator system. Inclusion of actuator dynamics results in larger flap control angles as well as more actuation power for different spanwise flap placements. One concern with a smart flap rotor system can be aeroelastic stability. Satisfactory stability characteristics that include blade aeroelastic stability and ground or air resonance can be quite critical to the design and development of a rotor craft with smart flaps. At this stage, there is no focused research on this topic. However, flutter phenomena of control surfaces in fixed-wing aircraft, such as wing-aileron, tail-elevator and rudder, is well studied. Broadbent [11] presented a discussion on flutter of control surfaces and tabs. As an example, the nature of aeroelastic stability of wing-aileron systems is explained by considering the aerodynamic forces that arise from the aileron motion and solving the binary flutter equations of wing bending-aileron and wing torsion-aileron. It was concluded that torsionaileron flutter is predicted with much less reliability than bending-aileron flutter, as it is more sensitive to small changes in the aerodynamic forces; and torsion-aileron flutter usually occurs at higher speeds than bending-aileron flutter for fixed-wing aircraft. The distortion of the control surface itself is usually not important and can be treated as rigid. The control circuit can be considered as a light spring. It is explained that the avoidance of control surface flutter can be achieved by using mass-balance, irreversible controls, and adding more damping. Pung [13] explained the flutter phenomenon by considering energy transfer between wing distortion and aileron deflection, and gave historical remarks on flutter analysis development. Flutter of a two dimensional airfoil with an aileron and a control tab was formulated, and methods to calculate the solution of the flutter determinant were presented. Theodorsen [12] presented the aerodynamic model for an oscillating airfoil or airfoil-aileron combination with three independent degrees of freedom: wing bending, wing torsion, and aileron deflection. The mechanism of flutter was analyzed and analytical solution of flutter speed was determined. The differential equations of motion involving three degrees of freedom were derived by substituting in the aerodynamic forces of the airfoil-aileron combination. The stability solution of the equations is calculated by using a determinant expansion method. The flutter speed solutions of the cases of torsion-aileron, bending-aileron, and bending-torsion were compared with experimental data. The comparison shows good agreement between experimental results and analysis for bending-torsion and bending-aileron cases, and fair agreement for torsion-aileron. Flutter phenomena of control surface is well understood in the fixed wing community, while flutter analysis of trailing-edge flaps in an elastic rotor has not been investigated systematically. Introduction of trailing-edge flap onto rotor blades may introduce many complex aerodynamic and inertial forces that may induce a unique aeroelastic instability. Such an instability can be of a great concern for the rotorcraft designer [14]. The objective of this paper is to examine systematically the aeroelastic stability of a rotor system with American Institute of Aeronautics and Astronautics (c)2001 American Institute of Aeronautics & Astronautics or Published with Permission of Author(s) and/or Author(s)' Sponsoring Organization.