Cyclic Pitch Input Research Articles

Modeling and experimental study on orientation dynamics of a Mars rotorcraft with swashplate mechanism

The Mars rotorcraft is an auxiliary aerial exploration platform to expand the rovers' patrol range. Attitude control is a fundamental aspect of autonomous flight control. So far, there has been no detailed investigation of the orientation dynamics of the rotor in the Martian atmosphere. A swashplate mechanism for manipulating isolated rotors is proposed to control the blade pitch angle, whose kinematics equation is established based on the closed condition of space vectors. Considering the blade as a rigid beam flapping around a central hinge, the flapping equation of the hingeless rotor is established to analyze the aerodynamic load for orientation adjusting. The Mars rotorcraft orientation dynamics model is constructed to develop an attitude control algorithm by combining the swashplate kinematics equation and the rotor flapping equation. The orientation experiment is conducted to verify the theoretical model and investigate the response law of the body's heading angle to the cyclic pitch input. The experimental results indicate that the attitude response is more sensitive to the blade pitch angle input under a smaller collective pitch angle or a higher rotor speed. The rigid rotor hypothesis may be unsuitable for the case where the collective pitch angle is smaller than 22°. The hovering flight tests prove that the orientation dynamics model proposed in this paper can effectively support the development of the Mars rotorcraft attitude control system.

Aeroelastic responses of airfoil under dynamic stall forced to oscillate by cyclic pitch input

The helicopter blades are involved in a nonlinear aeroelastic system in which unsteady dynamic stall flow and flexible structure are coupled. This problem can be reduced to a 2D aeroelastic problem in which airfoil is forced to oscillate through cyclic pitch input. In this paper, a modified dynamic stall model which focuses on revising the modeling of aerodynamic loads in reattachment stage is adopted in the numerical simulation to predict the aeroelastic responses of the above-mentioned system. The results are validated against experiments, after which parametric studies are performed to investigate the complex aeroelastic responses. In single Degree-Of-Freedom (DOF) system, secondary resonances happen in specific ratios of natural frequency to forcing frequency, which are found in both experiments and simulation. Moreover, as free-stream velocity varies, the aerodynamic stiffness has an apparent influence on the pitching natural frequency of system with low structural frequency. The varying pitching natural frequency brings about different aeroelastic responses including 1/2 subharmonic resonances and quenching phenomenon in single-DOF system as well as internal resonances and combination resonances in 2-DOF system. The aeroelastic system with low structural frequency is further investigated through multiple scale method, which provides an illustration for the mechanism of those phenomena found in the aeroelastic simulation adopting a semi-empirical aerodynamic model.

The adverse cyclic and collective pitch effect in a rotor

AbstractNumerical simulations have been carried out for a 32.16-ft-diameter rotor in autorotational forward flight. Coupled flapping and rotational equations were solved using the transient simulation method (TSM) to ascertain the quasistatic torque equilibrium conditions. The Pitt/Peters inflow theory was adopted in the simulations, and an airfoil look-up table made by a compressible Navier-Stokes solver was used. The adverse cyclic and collective pitch inputs were introduced in a similar fashion to helicopter control in that the cyclic lever is pulled back and the collective lever is pushed down for increasing airspeeds. The simulation results showed that the longitudinal cyclic pitch input combined with a lowered collective pitch increases the rotating torque for a low shaft angle and an advance ratio greater than one, producing both high lift and a high lift-to-drag ratio. Upon introducing the adverse cyclic and collective pitch inputs, the control range broadened, and a torque equilibrium condition was detected at 414.7kt (700ft/s) of airspeed in the simulation.

Analysis of Co-Flow Jet Effect on Dynamic Stall Characteristics Applying to Rotor Airfoils

The performance of helicopter is limited by the phenomenon of dynamic stall. Dynamic stall is attributed to the time-varying angle of attack (AOA) of blade element resulting from blade flapping cyclic pitch inputs and wake flow. Co-flow jet(CFJ), as a novel flow control technique proposed by the joint research team from University Miami (UM) and AFRL has shown great improvement on lift enhancement, drag reduction due to its strong circulation improvement and turbulent mixing effect. Focusing on dynamic stall characteristics, this paper studies the effect of co-flow jet technique applying to rotor airfoil OA309. Numerical simulations show that CFJ can remove dynamic stall phenomenon occurring on rotor airfoils but the momentum coefficient should be carefully chosen so as to obtain high-efficient control effect.

Individual pitch control for 2‐bladed wind turbines via multiblade multilag transformation

AbstractTwo‐bladed wind turbines have regained the attention of the community thanks to the advantages in manufacturing cost provided by the lower number of blades and the ease of implementation of effective passive systems for load reduction (ie, teetering pin). Considering both teetering and nonteetering architectures, the dynamics of 2‐bladed turbines is different from that of 3‐bladed machines especially in terms of how multiples of the rotor frequencies in blade signals are translated on the fixed system. Such characteristics have hampered the adoption of active control laws for load mitigation based on individual pitch control, extensively studied for 3‐bladed turbines. A basic element for control development allowing to capture the essence of the relationship between signals on the blades and the fixed system on 2‐bladed turbines–represented by the Coleman transformation for 3‐bladed turbines–has not been identified yet. The present paper tries to fill the gap, presenting an extended transformation–called multiblade multilag–applicable to turbines with an arbitrary number of blades, providing a systematic way to link rotor signals to fixed system signals, thus allowing the application of control algorithms for individual pitch control developed for 3‐bladed turbines to the 2‐bladed case. The paper addresses the problem first at a theoretical level, and subsequently providing applicative results from simulations on virtual models of teetering and nonteetering 2‐bladed turbines. The proposed transformation algorithm and control laws allow to effectively reduce some relevant loads and motions respectively in the nonteetering and teetering scenarios, through a cyclic pitch input.

Single-shot temperature- and pressure-sensitive paint measurements on an unsteady helicopter blade

Unsteady pressure-sensitive paint (PSP) measurements were acquired on an articulated model helicopter rotor of 0.26 m diameter in edgewise flow to simulate forward flight conditions. The rotor was operated at advance ratios (free stream velocity normalized by hover tip speed) of 0.15 and 0.30 at a cycle-averaged tip chord Reynolds number of 1.1 × 105, with collective and longitudinal cyclic pitch inputs of 10° and 2.5°, respectively. A single-shot data acquisition technique allowed a camera to record the paint luminescence after a single pulse of high-energy laser excitation, yielding sufficient signal-to-noise ratio to avoid image averaging. Platinum tetra(pentafluorophenyl) porphyrin (PtTFPP) in a porous polymer/ceramic binder served as the PSP. To address errors caused by image blurring and temperature sensitivity, a previously reported motion deblurring algorithm was implemented and the temperature correction was made using temperature-sensitive paint measurements on a second rotor blade. Instantaneous, unsteady surface pressure maps at a rotation rate of 82 Hz captured different aerodynamic responses between the two sides of the rotor disk and were compared to the nominally steady hover case. Cycle-to-cycle variations in tip unsteadiness on the retreating blade were also observed, causing oblique pressure features which may be linked to three-dimensional stall.

小型2重反転回転翼機の前進飛行に関する研究

The toy model coaxial helicopter for entertainment has a characteristic that it goes left when the operation for forward motion is made with the controller. The reason for this characteristic was made clear as follows: The phase difference between cyclic pitch and flapping motion is varied with the forward velocity. For this helicopter, the azimuth angles for the cyclic pitch inputs to the lower rotor are determined for a hovering flight. Then, the tip path plane of the lower rotor is inclined to the left and forward, when the operation for forward motion is kept. And the fuselage moves to the left and forward. On the other hand, during the forward motion, the inclination of the tip path plane of the upper rotor is varied by the stabilizer bar, whose motion is mainly determined by the fuselage attitude. Then, there are both moments when the upper rotor suppresses the fuselage motion generated by the lower rotor and when it promotes the fuselage motion.

Effects of Auxiliary Lift and Propulsion on Helicopter Vibration Reduction and Trim

This paper examines the vibration reductions caused by the introduction of auxiliary lift and propulsion, individually, as well as in combination, on a light [5800-lb (2640 kg)] helicopter with a four-bladed hingeless rotor, at flight speeds close to the maximum cruise velocity of the baseline helicopter. The changes in trim (vehicle orientations and control settings) because of auxiliary lift and propulsion are also examined in detail, and the fundamental mechanisms that produce the changes in trim and associated vibration reductions are identified. Based on results using a comprehensive aeroelastic analysis, it was concluded that auxiliary lift, alone, produces relatively small reductions in vibration. On the other hand, significant vibration reductions were obtained through auxiliary propulsion alone. A combination of lift and propulsion was most effective and reduced the vibration index by over 90%. It was also observed that auxiliary lift significantly reduces the main rotor thrust but increases the nose-down pitch attitude and tip-path-plane forward tilt to provide the required propulsive force. This increases the downwash through the rotor disk and requires a larger rotor longitudinal cyclic pitch input. In contrast, auxiliary propulsion that minimizes vibration produces little reduction in main rotor thrust, but results in a slightly nose-up pitch attitude (the auxiliary propulsion exceeds vehicle drag) along with a backward tilt of the tip-path plane. This decreases the downwash through the rotor disk and requires a smaller rotor longitudinal cyclic pitch input. A combination of auxiliary lift and propulsion minimizes vibration results in an even larger backward tilt of the tip-path plane and a net upwash through the rotor disk. The rotor collective pitch undergoes little change as a result of auxiliary lift, even though the main rotor thrust is decreased. In contrast, for auxiliary propulsion it decreases significantly even though the rotor thrust undergoes only small reductions. This counterintuitive observation is explained. The reduced downwash with auxiliary propulsion, or upwash with combined lift and propulsion, puts the rotor in a partial autorotation state, drastically reducing the induced drag, main rotor torque, and power. Auxiliary lift produces modest reductions in main rotor power, primarily because of a reduced profile drag associated with lower rotor loading. Because the rotor loading is lower with auxiliary lift than with auxiliary propulsion, but larger vibration reductions are produced with the latter, it can be deduced that vibration reductions are less a result of “unloading” of the rotor per se and more because of overall changes in trim, especially the reduction in longitudinal cyclic pitch (seen with auxiliary propulsion).

Helicopter Vibration Reduction Using Fixed-System Auxiliary Moments

Reductions in helicopter rotor hub vibratory loads that can be obtained through the introduction of steady auxiliary pitching and rolling moments in the fixed system are examined. The auxiliary moments allow changes in vehicle attitude and rotor pitch control inputs, such that vibrations can be minimized, while vehicle equilibrium is still satisfied. Because auxiliary moments can be generated by variation of the pitch of the right and left halves of the horizontal tail, the weight penalty, complexity, and performance losses associated with auxiliary lift and thrust (conventional compounding) would be alleviated. A comprehensive rotorcraft aeroelastic analysis was modified to include the fixed system auxiliary moments in the trim equations, and vibration reductions were examined for two different cases: 1) a light BO-105 type helicopter with a four-bladed hingeless rotor at high flight speeds (advance ratio of 0.35) and 2) a medium-weight UH-60 type helicopter with a four-bladed articulated rotor at a more moderate flight speed (advance ratio of 0.25). For the BO-105 type aircraft, a nosedown auxiliary pitching moment and a roll-left auxiliary rolling moment could reduce the in-plane vibratory hub forces by 20-30% when used individually and by 90% when used in combination. For the UH-60 type aircraft, a nosedown auxiliary pitching moment and a roll-right auxiliary rolling moment produced reductions in the range of 15-65% in all components of vibratory hub loads. For both aircraft, the nosedown auxiliary pitching moment increased the nosedown pitch attitude of the aircraft and reduced the rotor longitudinal cyclic flapping and the longitudinal cyclic pitch input. For the BO-105, the roll-left auxiliary moment increased the leftward shaft tilt and no longer required the tip path plane to tilt left relative to the hub plane. This decreased the lateral cyclic pitch input. However, for the UH-60, the roll-right auxiliary moment actually reversed the roll attitude of the aircraft (shaft now tilting right). This required a large leftward tilt of the tip path plane relative to the hub plane and increased the lateral cyclic pitch requirement. When implementation of the auxiliary moments through collective and differential pitch inputs to the horizontal tail was examined, it was observed that the expected reductions in vibration were indeed achieved, but that the auxiliary roll moment generated was smaller than expected due to one-half of the horizontal tail being stalled.

Active Control of Gimballed Rotors Using Swashplate Actuation During Shipboard Engagement Operations

An analysis has been developed to investigate the feedback control of gimballed rotor systems using swashplate actuation during transient shipboard engagement operations. The equations of motion for a rigid, three-bladed gimballed rotor system were derived using a Newtonian force summation method. Aerodynamic forces were simulated with a blade-element linear attached flow model. The resulting equations of motion were time integrated along a specified rotor-speed profile. A time domain linear-quadratic-regulator optimal control technique was applied to the equations of motion to minimize the transient rotor response. The physical limits of collective and cyclic pitch inputs in both magnitude and rate were enforced. The maximum transient gimbal tilt angle was reduced by as much as 56% with full knowledge of the ship air-wake environment and 42% with only partial knowledge of the ship airwake environment. Increasing the physical limits of collective and cyclic pitch inputs achieved up to a 70% reduction in the maximum transient gimbal tilt angle.

Dynamic analysis of the rigid rotor system.

A unified dynamic analysis for the rotor motion of the rigid or nonarticulated rotor system in a hovering state is presented. The dynamic behavior of the rigid rotor as well as that of the conventional rotor is expressed in a simple linearized mathematical form by using complex variables. The available control moment and the necessary phase shift for a given cyclic pitch input are obtained in general formulas and charts. Generally, the control and damping moments of the rigid rotor system increase with the increment of the napping stiffness, but there is a critical stiffness at which the aforementioned moments will be maximum. Optimum napping stiffness may be given so as to avoid unfavorable coupling motion between roll and pitch. It is not always necessary to install the control gyro, because similar stabilizing effect may be obtained by using the proper blade configuration without a special gyro system. The necessary control phase advance is also given to avoid cross control of the rigid rotor system.