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  • 1
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    The influence of dynamic inflow and torsional flexibility on rotor damping in forward flight from symbolically generated equations (1985)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: The combined effects of blade torsion and dynamic inflow on the aeroelastic stability of an elastic rotor blade in forward flight are studied. The governing sets of equations of motion (fully nonlinear, linearized, and multiblade equations) used in this study are derived symbolically using a program written in FORTRAN. Stability results are presented for different structural models with and without dynamic inflow. A combination of symbolic and numerical programs at the proper stage in the derivation process makes the obtainment of final stability results an efficient and straightforward procedure.
    Keywords: AIRCRAFT STABILITY AND CONTROL
    Type: Rotorcraft Dynamics 1984; p 221-240
    Format: application/pdf
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  • 2
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    The use of active controls to augment rotor/fuselage stability (1985)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: The use of active blade pitch control to increase helicopter rotor/body damping is studied. Control is introduced through a conventional nonrotating swashplate. State variable feedback of rotor and body states is used. Feedback parameters include cyclic rotor flap and lead-lag states, and body pitch and roll rotations. The use of position, rate, and acceleration feedback is studied for the various state variables. In particular, the influence of the closed loop feedback gain and phase on system stability is investigated. For the rotor/body configuration analyzed, rotor cyclic inplane motion and body roll-rate and roll-acceleration feedback can considerably augment system damping levels and eliminate ground resonance instabilities. Scheduling of the feedback state, phase, and gain with rotor rotation speed can be used to maximize the damping augmentation. This increase in lead-lag damping can be accomplished without altering any of the system modal frequencies. Investigating various rotor design parameters (effective hinge offset, blade precone, blade flap stiffness) indicates that active control for augmenting rotor/body damping will be particularly powerful for hingeless and bearingless rotor hubs.
    Keywords: AIRCRAFT STABILITY AND CONTROL
    Type: Rotorcraft Dynamics 1984; p 1-15
    Format: application/pdf
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  • 3
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    The influence of dynamic inflow and torsional flexibility on rotor damping in forward flight from symbolically generated equations (1984)
    In:  Other Sources
    Details
    Publication Date: 2011-08-19
    Description: The combined effects of blade torsion and dynamic inflow on the aeroelastic stability of an elastic rotor blade in forward flight are studied. The Helicopter Equations for Stability and Loads (HESL) program is extended to derive the governing equations of motion for the blade, and a Lagrangian formulation is used to obtain the equations in generalized coordinates. The program generates the steady-state and linearized perturbation equations in symbolic form and then codes them into FORTRAN subroutines. The coefficients for each equation and for each mode are identified through a numerical program; the latter can also be used to obtain the harmonic balance equations. The governing multiblade equations are derived explicitly using HESL. These equations can accommodate any number of elastic blade modes. Stability results are presented for several hingeless rotor blade structural models, and the influence of dynamic inflow in forward flight with an elastic hingeless rotor is investigated.
    Keywords: AIRCRAFT DESIGN, TESTING AND PERFORMANCE
    Format: text
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  • 4
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    The use of active controls to augment rotor/fuselage stability (1985)
    In:  Other Sources
    Details
    Publication Date: 2011-08-19
    Description: The use of active blade pitch control to increase helicopter rotor/body damping is studied. Control is introduced through a conventional nonrotating swashplate. State variable feedback of rotor and body states is used. Feedback parameters include cyclic rotor flap and lead-lag states, and body pitch and roll rotations. The use of position, rate, and acceleration feedback is studied for the various state variables. In particular, the influence of the closed loop feedback gain and phase on system stability is investigated. For the rotor/body configuration analyzed, rotor cyclic inplane motion and body roll-rate and roll-acceleration feedback can considerably augment system damping levels and eliminate ground resonance instabilities. Scheduling of the feedback state, phase, and gain with rotor rotation speed can be used to maximize the damping augmentation. This increase in lead-lag damping can be accomplished without altering any of the system modal frequencies. Investigating various rotor design parameters (effective hinge offset, blade precone, blade flap stiffness) indicates that active control for augmenting rotor/body damping will be particularly powerful for hingeless and bearingless rotor hubs.
    Keywords: AIRCRAFT STABILITY AND CONTROL
    Type: American Helicopter Society, Journal (ISSN 0002-8711); 30; 13-22
    Format: text
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  • 5
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    High-speed, automatic controller design considerations for integrating array processor, multi-microprocessor, and host computer system architectures (1985)
    In:  Other Sources
    Details
    Publication Date: 2011-08-19
    Description: Modern control systems must typically perform real-time identification and control, as well as coordinate a host of other activities related to user interaction, online graphics, and file management. This paper discusses five global design considerations which are useful to integrate array processor, multimicroprocessor, and host computer system architectures into versatile, high-speed controllers. Such controllers are capable of very high control throughput, and can maintain constant interaction with the nonreal-time or user environment. As an application example, the architecture of a high-speed, closed-loop controller used to actively control helicopter vibration is briefly discussed. Although this system has been designed for use as the controller for real-time rotorcraft dynamics and control studies in a wind tunnel environment, the controller architecture can generally be applied to a wide range of automatic control applications.
    Keywords: COMPUTER OPERATIONS AND HARDWARE
    Format: text
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  • 6
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    Hover test of a full-scale hingeless rotor (1984)
    In:  Other Sources
    Details
    Publication Date: 2018-12-01
    Description: The performance and aeroelastic stability in hover of a 9.8-m diameter, hingeless helicopter rotor system was evaluated. Rotor performance and inplane damping data were obtained for rotor operation beteen 350 and 425 rpm for thrust coefficients (CT/sigma) between 0.0 and 0.12. At constant rotor thrust, a minimum in rotor inplane damping was measured at 400 rpm. Good agreement is shown between experimental performance data and predicted performance. The influence of different aerodynamic inflow models on predicting damping levels is also shown. The best correlation with experimental stability data was obtained when a dynamic inflow model was used instead of static or quasistatic inflow models. Comparison with other full scale, hingeless rotor data in hover is presented. The hingeless rotor data and data from a full scale, bearingless main rotor test performed on the same general purpose test apparatus were compared. Although the bearingless rotor was more highly damped at design tip speed and 1-g thrust operation, greater sensitivity to operating conditions is shown. At low thrust levels the bearingless main rotor is less damped than the hingeless rotor.
    Keywords: AIRCRAFT DESIGN, TESTING AND PERFORMANCE
    Format: text
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  • 7
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    Rotorcraft research testing in the National Full-Scale Aerodynamics Complex at NASA Ames Research Center (1985)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The unique capabilities of the National Full-Scale Aerodynamics Complex (NFAC) for testing rotorcraft systems are described. The test facilities include the 40- by 80-Foot Wind Tunnel, the 80- by 120-Foot Wind Tunnel, and the Outdoor Aerodynamic Research Facility. The Ames 7- by 10-Foot Subsonic Wind Tunnel is also used in support of the rotor research programs conducted in the NFAC. Detailed descriptions of each of the facilities, with an emphasis on helicopter rotor test capability, are presented. The special purpose rotor test equipment used in conducting helicopter research is reviewed. Test rigs to operate full-scale helicopter main rotors, helicopter tail rotors, and tilting prop-rotors are available, as well as full-scale and small-scale rotor systems for use in various research programs. The test procedures used in conducting rotor experiments are discussed together with representative data obtained from previous test programs. Specific examples are given for rotor performance, loads, acoustics, system interactions, dynamic and aeroelastic stability, and advanced technology and prototype demonstration models.
    Keywords: AIRCRAFT DESIGN, TESTING AND PERFORMANCE
    Type: NASA-TM-86687 , REPT-85140 , NAS 1.15:86687
    Format: application/pdf
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  • 8
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    Hover test of a full-scale hingeless helicopter rotor: Aeroelastic stability, performance and loads data (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A hover test of a full-scale, hingeless rotor system was conducted in the NASA Ames 40- by 80-foot wind tunnel. The rotor was tested on the Ames rotor test apparatus. Rotor aeroelastic stability, performance, and loads at various rotational speeds and thrust coefficients were investigated. The primary objective was to determine the inplane stability characteristics of the rotor system. Rotor inplane damping data were obtained for operation between 350 and 425 rpm (design speed), and for thurst coefficients between 0.0 and 0.12. The rotor was stable for all conditions tested. At constant rotor rotational speed, a minimum inplane dampling level was obtained at a thrust coefficient approximately = 0.02. At constant rotor lift, a minimum in rotor inplane damping was measured at 400 rpm.
    Keywords: AIRCRAFT DESIGN, TESTING AND PERFORMANCE
    Type: NASA-TM-85892 , A-9573 , NAS 1.15:85892
    Format: application/pdf
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  • 9
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    A full-scale wind tunnel investigation of a helicopter bearingless main rotor (1981)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A helicopter bearingless main rotor was tested. Areas of investigation included aeroelastic stability, aerodynamic performance, and rotor loads as a function of collective pitch setting, RPM, airspeed and shaft angle. The rotor/support system was tested with the wind tunnel balance dampers installed and, subsequently, removed. Modifications to the rotor hub were tested. These included a reduction in the rotor control system stiffness and increased flexbeam structural damping. The primary objective of the test was to determine aeroelastic stability of the fundamental flexbeam/blade chordwise bending mode. The rotor was stable for all conditions. Damping of the rotor chordwise bending mode increases with increased collective pitch angle at constant operating conditions. No significant decrease in rotor damping occured due to frequency coalescence between the blade chordwise fundamental bending mode and the support system.
    Keywords: AIRCRAFT DESIGN, TESTING AND PERFORMANCE
    Type: NASA-TM-81321 , A-8696
    Format: application/pdf
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  • 10
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    Formulation of coupled rotor/fuselage equations of motion (1979)
    In:  Other Sources
    Details
    Publication Date: 2019-06-27
    Description: The governing equations of motion of a helicopter rotor coupled to a rigid body fuselage are derived. A consistent formulation is used to derive nonlinear periodic coefficient equations of motion which can be used to study coupled rotor/fuselage dynamics in forward flight. The methodology of rotor/fuselage coupling is clearly described and the importance of an ordering scheme in deriving consistent nonlinear equations of motion is reviewed. The final equations which are presented in partial differential form can be used to model coupled rotor/fuselage aeroelastic response or stability problems.
    Keywords: AIRCRAFT DESIGN, TESTING AND PERFORMANCE
    Type: Vertica; 3; 3-4,; 1979
    Format: text
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