A summary of recent NASA/Army contributions to rotorcraft vibrations and structural dynamics technology

The requirement for low vibrations has achieved the status of a critical design consideration in modern helicopters. There is now a recognized need to account for vibrations during both the analytical and experimental phases of design. Research activities in this area were both broad and varied and notable advances were made in recent years in the critical elements of the technology base needed to achieve the goal of a jet smooth ride. The purpose is to present an overview of accomplishments and current activities of govern and government-sponsored research in the area of rotorcraft vibrations and structural dynamics, focusing on NASA and Army contributions over the last decade or so. Specific topics addressed include: airframe finite-element modeling for static and dynamic analyses, analysis of coupled rotor-airframe vibrations, optimization of airframes subject to vibration constraints, active and passive control of vibrations in both the rotating and fixed systems, and integration of testing and analysis in such guises as modal analysis, system identification, structural modification, and vibratory loads measurement

Concepts for a theoretical and experimental study of lifting rotor random loads and vibrations. Phase 6-B: Experiments with progressing/regressing forced rotor flapping modes

A two bladed 16-inch hingeless rotor model was built and tested outside and inside a 24 by 24 inch wind tunnel test section at collective pitch settings up to 5 deg and rotor advance ratios up to .4. The rotor model has a simple eccentric mechanism to provide progressing or regressing cyclic pitch excitation. The flapping responses were compared to analytically determined responses which included flap-bending elasticity but excluded rotor wake effects. Substantial systematic deviations of the measured responses from the computed responses were found, which were interpreted as the effects of interaction of the blades with a rotating asymmetrical wake

Rotorcraft aviation icing research requirements: Research review and recommendations

The status of rotorcraft icing evaluation techniques and ice protection technology was assessed. Recommendations are made for near and long term icing programs that describe the needs of industry. These recommended programs are based on a consensus of the major U.S. helicopter companies. Specific activities currently planned or underway by NASA, FAA and DOD are reviewed to determine relevance to the overall research requirements. New programs, taking advantage of current activities, are recommended to meet the long term needs for rotorcraft icing certification

Periodic controllers for vibration reduction using actively twisted blades

This paper compares two periodic control methods, the optimal H2 and the periodic static output feedback (POF), to reduce the helicopter rotor vibrations. Actively twisted blades with Macro-Fibre Composite (MFC) piezoelectric actuators are used. The design model is based on a simplified aerodynamic model and on a multi-body model of the Bo 105 isolated rotor with the original blades replaced by actively twisted ones. The performance of the two controllers in alleviating hub loads is verified with improved simulations based on a free-wake model

NASA/Army Rotorcraft Technology. Volume 1: Aerodynamics, and Dynamics and Aeroelasticity

The Conference Proceedings is a compilation of over 30 technical papers presented at this milestone event which reported on the advances in rotorcraft technical knowledge resulting from NASA, Army, and industry rotorcraft research programs over the last 5 to 10 years. The Conference brought together over 230 government, industry, and allied nation conferees to exchange technical information and hear invited technical papers by prominent NASA, Army, and industry researchers covering technology topics which included: aerodynamics, dynamics and elasticity, propulsion and drive systems, flight dynamics and control, acoustics, systems integration, and research aircraft

Development of a High-Performance Low-Weight Hydraulic Damper for Active Vibration Control of the Main Rotor on Helicopters—Part 1: Design and Mathematical Model

The helicopter vibrations generated by the main rotor/gearbox assembly are the principal cause of damage to cockpit instruments and discomfort of the crew in terms of cabin noise. The principal path of vibration transmission to the fuselage is through the gearbox rigid support struts. With the aim of reducing these vibrations, this paper presents the design of a low-weight high-performance active damper for vibration control developed by Elettronica Aster S.p.A. The system is intended to replace the conventional struts and is composed of an electro-hydraulic actuator hosted within a compliant structure. This parallel nested structure allows the system to reach a high-power density. A physics-based mathematical model was used as a design digital twin to optimize the performance to meet the strict requirements. The active damper was designed for a reference application of a 15-seat medium-sized twin-engine helicopter. The model was used to perform the tests specified in the acceptance and testing procedure document, showing the compliance with the requirements of the current design. The damper physical realization, test bench design, experimental campaign, and model validation will be presented in Part 2

Active Vibration Control of Helicopter Rotor Blade by Using a Linear Quadratic Regulator

Active vibration control is a widely implemented method for the helicopter vibration control. Due to the significant progress in microelectronics, this technique outperforms the traditional passive control technique due to weight penalty and lack of adaptability for the changing flight conditions. In this thesis, an optimal controller is designed to attenuate the rotor blade vibration. The mathematical model of the triply coupled vibration of the rotating cantilever beam is used to develop the state-space model of an isolated rotor blade. The required natural frequencies are determined by the modified Galerkin method and only the principal aerodynamic forces acting on the structure are considered to obtain the elements of the input matrix. A linear quadratic regulator is designed to achieve the vibration reduction at the optimum level and the controller is tuned for the hovering and forward flight with different advance ratios

Contributions to the dynamics of helicopters with active rotor controls

This dissertation presents an aeromechanical closed loop stability and response analysis of a hingeless rotor helicopter with a Higher Harmonic Control (HHC) system for vibration reduction. The analysis includes the rigid body dynamics of the helicopter and blade flexibility. The gain matrix is assumed to be fixed and computed off-line. The discrete elements of the HHC control loop are rigorously modeled, including the presence of two different time scales in the loop. By also formulating the coupled rotor-fuselage dynamics in discrete form, the entire coupled helicopter-HHC system could be rigorously modeled as a discrete system. The effect of the periodicity of the equations of motion is rigorously taken into account by converting the system into an equivalent system with constant coefficients and identical stability properties using a time lifting technique. The most important conclusion of the present study is that the discrete elements in the HHC loop must be modeled in any HHC analysis. Not doing so is unconservative. For the helicopter configuration and HHC structure used in this study, an approximate continuous modeling of the HHC system indicates that the closed loop, coupled helicopter-HHC system remains stable for optimal feedback control configurations which the more rigorous discrete analysis shows can result in closed loop instabilities. The HHC gains must be reduced to account for the loss of gain margin brought about by the discrete elements. Other conclusions of the study are: (i) the HHC is effective in quickly reducing vibrations, at least at its design condition, although the time constants associated with the closed loop transient response indicate closed loop bandwidth to be 1~rad/sec on average, thus overlapping with FCS or pilot bandwidths, and raising the issue of potential interactions; (ii) a linearized model of helicopter dynamics is adequate for HHC design, as long as the periodicity of the system is correctly taken into account, i.e., periodicity is more important than nonlinearity, at least for the mathematical model used in this study; and (iii) when discrete and continuous systems are both stable, and quasisteady conditions can be guaranteed, the predicted HHC control harmonics are in good agreement

Rotorcraft Smoothing Via Linear Time Periodic Methods

This research presents the development of an on line linear time periodic rotor vibration controller. The Cramer-Rao bound is developed for a linear time periodic system in order to identify the quality of identified system parameters, which are used in system models for controller development. The methods developed in this work allow model parameters can be verified for accuracy and likewise adjusted to improve controller accuracy