Cycloidal Rotor

5.5 AIRFRAMES 125

flapping amplitude at a given collective pitch setting. Note that rotor thrust was insensitive to flapping amplitude at a constant collective pitch.

Future work should focus on increasing the thrust produced by the rotor by operating at higher rotational speed.

Fitchett and Chopra [65] developed a micro-scale rotor, called the Flotor, that was powered by blade flapping (Figure 5.17). A prototype was constructed and tested in three modes: pure flap-ping, pure rotation, and combined flapping/rota-tion. The geometry of the blades of the Flotor, as well as the rotational speed, was determined based on the wings of bats of similar size and their reduced frequency in cruise flight. The pro-totype rotor had two blades that were flapped in phase, a rotor radius of 80 mm, and a blade aspect ratio of 6.5. To ensure that pure flapping initiated rotor rotation in the correct direction, the blades were constructed with a main spar at the leading edge. This resulted in sufficient elastic twist in the blades to generate the appropriate propulsive forces at low rotational speed. The blades were constructed out of 0.25 mm thick mylar sheet and a carbon fiber framework.

In the pure flapping mode (passive rotation), a maximum disk loading of 10 N/m² was meas-ured, which is low compared to conventional

shaft-driven rotors. For the pure rotation and combined flapping/rotation tests, rigid blades with a circular arc profile were tested. Com-pared to the pure rotation cases, an increase in thrust of up to 20% and a decrease in torque of up to 30% were measured during combined flapping/rotation. Several recommendations were made for future research leading to a flight-capable prototype.

Although the concept of active blade flapping has been shown to enhance the performance of a conventional rotor, the main challenges to this approach are the mechanical complexity of the rotor hub, the inertial forces due to active blade-flapping, and the additional power required by the blade-flapping mechanism. These challenges must be addressed appropriately to enable flight testing of such a configuration.

that is perpendicular to the direction of flight (Figure 5.18). The blade span is parallel to the axis of rotation. As the blades rotate around the azimuth, their pitch angle is varied periodically, typically using a passive mechanism such as a four-bar linkage. Each spanwise blade element operates at about the same conditions—velocity, Reynolds number, angle of attack, centrifugal force—and thus can be designed to operate at its optimum efficiency.

Figure 5.19 shows a cross-section of a six-bladed cycloidal rotor rotating with an angular velocity Ω. Each of the blades produces a lift and a drag force. Blades at the top and bottom posi-tions produce an almost vertical net force, while those at the sides produce small lateral forces because of their reduced angle of attack. The horizontal components of the forces cancel, resulting in a net vertical thrust. In addition, the amplitude and phase of the maximum blade pitch angle may be changed by modifying the configuration of the mechanical linkage. In this way, the magnitude and direction of the net

thrust vector of the rotor can be changed almost instantaneously.

The concept of cycloidal propulsion was first investigated in the 1920s by Kirsten [66] and in the 1930s by Wheatley [67, 68]. These early cycloidal rotors were intended for use in full-scale aircraft. Wind-tunnel tests were performed on 8 ft diameter cycloidal rotors and significant forces were obtained; however, due to incom-plete theoretical knowledge of unsteady aerody-namic effects, it was not possible to accurately predict the performance of these devices. The cycloidal rotor can change the direction of its thrust vector almost instantaneously over a com-plete circle, i.e., over an angular range of 360°.

Because of this unique ability, cycloidal rotors eventually made their way to marine systems, where they are used in tugboats to provide them with low-speed maneuverability. More recently, cycloidal rotors have made a reappearance in aircraft applications. They have been proposed for use on airships [69, 70] and on an UAV of gross weight 600 lb, where the wings are replaced by cycloidal rotors [71]. On a smaller scale, cycloidal rotors of span around 0.8 m have been investigated for VTOL UAVs of take-off mass around 50 kg [72, 73]. These rotors were able to demonstrate a power loading around 12 kg/HP at low thrust that asymptoted to 5 kg/HP at high thrust.

Due to the potential performance benefits of unsteady aerodynamic effects as well as the increased maneuverability afforded by the instantaneous change in thrust vector, cycloidal rotors have been explored for microflyers.

Hwang et al. [74] designed a microscale cyclo-copter with two cycloidal rotors of radius 0.2 m.

Sirohi and Parsons [75] developed a six-bladed, six-inch-diameter cycloidal rotor for a micro-aerial vehicle. Experiments were performed on a prototype to measure the flowfield in the downwash of the rotor as well as the thrust and torque produced at rotational speeds up to 1,200 rpm. The rotor blades had a NACA0010 profile, and the amplitude of the oscillatory blade pitch Ω

FIGURE 5.19 Thrust vectors at each blade cross-section.

5.5 AIRFRAMES 127 angle could be set from 0° to 40°. This translated

into a reduced frequency of around 0.167, which is considered highly unsteady. A time-domain formulation based on the Wagner’s function, in conjunction with downwash predicted based on momentum theory, was used to predict the thrust and torque of the rotor. Good agreement with measured thrust was observed, but there was some discrepancy with measured torque.

These discrepancies were attributed to an over-simplification of the flowfield, especially through the central part of the rotor. The power loading at low thrust settings was observed to be com-parable to that of a conventional helicopter rotor of the same diameter and asymptoted to a lower value at high thrust settings.

Based on experimental results, a micro-aerial vehicle powered by two six-inch-diameter cycloidal rotors was designed (Figure 5.20). The total mass of this vehicle was around 250 g and the rotor speed was around 1,650 rpm.

Benedict et al. [76, 77] performed further exper-imental studies on a cycloidal rotor of the same size, with the goal of optimizing the performance of the cycloidal rotor MAV. The effects of number of blades (ranging from two to six), maximum pitching amplitude, and airfoil camber were investigated. Improved performance was achieved with a larger number of blades, higher pitching amplitude, and uncambered airfoils.

Particle image velocimetry (PIV) measurements indicated a high degree of wake skewness as well as significant rotational flows inside the cycloidal rotor. Aeroelastic modeling of a cycloidal rotor using nonlinear finite elements and multibody simulations with different inflow models [78]

indicated that the wake skewness and resulting side force arises from the mechanical linkage as well as a phase lag due to unsteady aerodynamic effects. Torsional deformations were shown to decrease the thrust produced. Further experi-mental studies on the blade airfoil profile and

FIGURE 5.20 Conceptual twin cycloidal rotor MAV.

location of pitching axis were performed, culmi-nating in the design and successful hover flight of a micro-aerial vehicle with four cycloidal rotors [79], with a total mass of around 750 g and a power loading of 5.6 kg/HP (Figure 5.21).

Future work in this area is expected to focus on the forward flight capability of the cycloidal rotor as well as maneuverability and improve-ment of performance by further harnessing unsteady aerodynamic effects.