Engineering topics

Nguyen, Khanh Q.

Publications and source records attributed to Nguyen, Khanh Q..

5 records

NASA NTRS2004

Blade-Pitch Control for Quieting Tilt-Rotor Aircraft

A method of reducing the noise generated by a tilt-rotor aircraft during descent involves active control of the blade pitch of the rotors. This method is related to prior such noise-reduction methods, of a type denoted generally as higher-harmonic control (HHC), in which the blade pitch is made to oscillate at a harmonic of the frequency of rotation of the rotor. A tilt-rotor aircraft is so named because mounted at its wing tips are motors that can be pivoted to enable the aircraft to take off and land like a helicopter or to fly like a propeller airplane. When the aircraft is operating in its helicopter mode, the rotors generate more thrust per unit rotor-disk area than helicopter rotors do, thus producing more blade-vortex interaction (BVI) noise. BVI is a major source of noise produced by helicopters and tilt-rotor aircraft during descent: When a rotor descends into its own wake, the interaction of each blade with the blade-tip vortices generated previously gives rise to large air-pressure fluctuations. These pressure fluctuations radiate as distinct, impulsive noise. In general, the pitch angle of the rotor blades of a tilt-rotor aircraft is controlled by use of a swash plate connected to the rotor blades by pitch links. In both prior HHC methods and the present method, HHC control signals are fed as input to swash-plate control actuators, causing the rotor-blade pitch to oscillate. The amplitude, frequency, and phase of the control signal can be chosen to minimize BVI noise.

Betzina, Mark D.↗

NASA NTRS1997

Numerical Simulation of Unsteady Aerodynamic Models

This report documents the results of the numerical simulations of unsteady aerodynamic models. The results focus on numerical accuracy and efficiency, and the robustness of the numerical methods. The aerodynamic models includes the classical Wagner and Kussner functions and the Leishman-Beddoes dynamic stall model. The simulations includes the numerical approximations of the Duhamel's integrals using both indicial (step) and impulse responses, the numerical integrations of the state-space models, and the exact solutions. The report also presents the conversion among different model representations.

Nguyen, Khanh Q.↗

NASA NTRS1995

Parametric Investigation of The Aeromechanic Stability of Bearingless Rotors

The objective of this investigation is to conduct a parametric investigation on the aeromechanic stability of modern bearingless rotors. To ensure aeromechanic stability, modem bearingless; rotors use elastomeric dampers to augment the blade inplane damping. The augmented dampers are necessary to avoid aeromechanic instabilities such as air and ground resonance on soft-inplane rotors. The prevention of air and ground resonance depends largely on the damping level of the rotor-fuselage system during the critical frequency-crossing between the rotor and the fuselage inplane motions. The blade inplane damping, critical in ensuring aeromechanic stability of a rotor, depends not only on the damper sizes but also on the aeroelastic properties of the rotor blade. The results of this investigation provide insight into the source of inplane damping. The parametric study is carried out analytically using the University of Maryland Advanced Rotorcraft Code, or UMARC.

Nguyen, Khanh Q.↗

NASA NTRS1994

Higher harmonic control analysis for vibration reduction of helicopter rotor systems

An advanced higher harmonic control (HHC) analysis has been developed and applied to investigate its effect on vibration reduction levels, blade and control system fatigue loads, rotor performance, and power requirements of servo-actuators. The analysis is based on a finite element method in space and time. A nonlinear time domain unsteady aerodynamic model, based on the indicial response formulation, is used to calculate the airloads. The rotor induced inflow is computed using a free wake model. The vehicle trim controls and blade steady responses are solved as one coupled solution using a modified Newton method. A linear frequency-domain quasi-steady transfer matrix is used to relate the harmonics of the vibratory hub loads to the harmonics of the HHC inputs. Optimal HHC is calculated from the minimization of the vibratory hub loads expressed in term of a quadratic performance index. Predicted vibratory hub shears are correlated with wind tunnel data. The fixed-gain HHC controller suppresses completely the vibratory hub shears for most of steady or quasi-steady flight conditions. HHC actuator amplitudes and power increase significantly at high forward speeds (above 100 knots). Due to the applied HHC, the blade torsional stresses and control loads are increased substantially. For flight conditions where the blades are stalled considerably, the HHC input-output model is quite nonlinear. For such cases, the adaptive-gain controller is effective in suppressing vibratory hub loads, even though HHC may actually increase stall areas on the rotor disk. The fixed-gain controller performs poorly for such flight conditions. Comparison study of different rotor systems indicates that a soft-inplane hingeless rotor requires less actuator power at high speeds (above 130 knots) than an articulated rotor, and a stiff-inplane hingeless rotor generally requires more actuator power than an articulated or a soft-inplane hingeless rotor. Parametric studies for a hingeless rotor operating in a transition flight regime and for an articulated rotor operating at the level-flight boundary (high speed and high thrust conditions) indicate that blade parameters including flap, lag, torsion stiffness distributions, linear pretwist, chordwise offset of center-of-mass from elastic axis and chordwise offset of elastic axis from aerodynamic center can be selected to minimize the actuator power requirements for HHC.

Nguyen, Khanh Q.↗

NASA NTRS1994

Aeroelastic Stability of Modern Bearingless Rotors: A Parametric Investigation

The University of Maryland Advanced Rotorcraft Code (UMARC) is utilized to study the effects of blade design parameters on the aeroelastic stability of an isolated modern bearingless rotor blade in hover. The McDonnell Douglas Advanced Rotor Technology (MDART) Rotor is the baseline rotor investigated. Results indicate that kinematic pitch-lag coupling introduced through the control system geometry and the damping levels of the shear lag dampers strongly affect the hover inplane damping of the baseline rotor blade. Hub precone, pitchcase chordwise stiffness, and blade fundamental torsion frequency have small to moderate influence on the inplane damping, while blade pre-twist and placements of blade fundamental flapwise and chord-wise frequencies have negligible effects. A damperless configuration with a leading edge pitch-link, 15 deg of pitch-link cant angle, and reduced pitch-link stiffness is shown to be stable with an inplane damping level in excess of 2.7 percent critical at the full hover tip speed.

Nguyen, Khanh Q.↗