2008-ahs-cyclopter-with Moble

Experimental Investigation of the Cycloidal-Rotor Concept for a Hovering Micro Air Vehicle Moble Benedict*

Inderjit Chopra†

Manikandan Ramasamy‡

J. Gordon Leishman§

Alfred Gessow Rotorcraft Center Department of Aerospace Engineering University of Maryland College Park, MD 20742

Abstract The viability of cycloidal rotor (cyclocopter) concept for developing hover-capable micro-air vehicles was analyzed through performance and flow field measurements. Detailed parametric studies were conducted to identify the dependence of the rotor performance on the amplitude of the blade pitch, blade airfoil profile, blade flexibility, and rotational speed. All experiments were conducted using a three-bladed, model scale cyclocopter that was tested up to the design rotational speed of 2,000 RPM. While higher pitch angles were found to increase the power loading (thrust/power) of the cyclocopter, the bending and torsional flexibility of the blades deteriorated the performance. Similarly, fixed camber proved to be detrimental to the rotor performance. Thrust measurements suggested the presence of a sidewise force (along with the vertical lift), similar to those found with lifting cylinders. Digital particle image velocimetry (DPIV) measurements made in the wake of the cyclocopter provided evidence of wake skewness, resulting in sideward force. The thrust produced by the cyclocopter was found to increase with a geometric pitch of 45° without showing any signs of stall. This behavior was explained through DPIV measurements that showed the presence of high induced velocities, resulting from the trailing tip vortices. These velocities are comparable to the rotor blade velocities, reducing substantially the aerodynamic angles of attack experienced by the rotor blades, thereby, preventing the occurrence of stall. DPIV measurements also identified several interesting flow features that include the presence of a leading edge vortex, similar to a dynamic stall vortex. The slipstream boundary obtained by following the path of the tip vortices was found to contract, as expected for a lifting rotor. .

Introduction In recent1 years, interest2 has been growing3 in a4 new class of very small fight vehicles called micro air vehicles (MAVs). This interest has developed, in part, by the changing needs of the military as the battlegrounds of the future move to restricted, highly populated urban environments where conventional fixed-wing-based aircraft lose much of their utility. MAVs can also be used for civilian applications such as biochemical sensing, traffic monitoring, border surveillance, fire and rescue operations, forestry, wildlife

surveys, power-line inspection and real-estate aerial photography, to name a few. Many fixed-wing MAVs have been successfully tested [1 – 3]. A good example of a fixed-wing MAV is the Aerovironment Black Widow [4, 5] with a weight of 80 grams and an endurance of about 30 minutes. Even though fixed-wing MAVs are the best performers today within the imposed size and weight constraints, they lack the ability to hover or operate in highly constrained environments which can be important requirements for surveillance. Therefore, studying efficient hovering rotary wing-based concepts would prove useful in developing more versatile MAVs with a larger flight envelope.

1

Graduate Research Assistant. [email protected] Alfred Gessow Professor and Director. [email protected] 3 Assistant Research Scientist. [email protected] 4 Minta Martin Professor. [email protected] Presented at the American Helicopter Society 64th Annual Forum, Montréal, Canada, April 29 – May 1, 2008. Copyright © 2008 by M. Benedict et al. Published by the American Helicopter Society International, Inc. with permission. 2

Several hovering MAVs based on scaled-down single main rotor and coaxial helicopter configurations have been successfully built and flight tested [6, 7]. These MAV scale rotors typically operate in the blade chord Reynolds number range from 10,000 to 60,000. Consequently, they experience much higher profile drag than conventional helicopter rotors. As a result, MAV scale rotors suffer from low aerodynamic efficiency, which translates into poor endurance. The maximum

Figure of Merit achieved to date for an MAV scale rotor is about 0.65 [6]. An MAV based on a cycloidal rotor has been proposed as an alternative to helicopter-based MAVs. A cycloidal rotor is like a horizontal rotary-wing system (Fig. 1), where the span of the blades is parallel to the axis of rotation of the rotor and perpendicular to the direction of flight. The pitch angle of each of the blades is varied periodically about the airfoil quarter-chord as the blade moves around the azimuth of the rotor such that the blade is at positive angle of attack both at the top and bottom positions (Fig. 2). The lift and drag forces produced by each blade can be resolved into vertical and horizontal directions, as shown in Fig. 2. The magnitude and direction of the net thrust vector of the rotor can be changed by varying the amplitude and phase of the cyclic blade pitch. In a cycloidal rotor, each blade element operates at the same conditions (velocity, Reynolds number, angle of attack, centrifugal force), and thus can be set at its best aerodynamic efficiency. Moreover, since the blades are periodically pitched at 1/rev, unsteady mechanisms may augment the lift produced by the blades. A few recent experiments have claimed that cycloidal rotors can be more efficient than a conventional rotor [8]. As the thrust vector of a cycloidal rotor can be instantaneously set to any direction perpendicular to the rotational axis, it may have more maneuverability compared to a helicopter-based MAV. This is an important attribute for highly constrained indoor operations. In the past, most of the tests performed on cycloidal rotors have been at larger scales (Table 1) [8 – 19]. Recent tests on a 6-inch diameter cycloidal rotor at the University of Maryland indicated that this concept is promising even at the lower Reynolds numbers at which MAVs operate [20, 21]. However, these tests were performed using a rigid bench-top test model. In the present work, a preliminary conceptual design study was performed to obtain an estimate of the weight and installed power required for a flight capable cycloidal MAV. A test rotor was designed and built light weight enough to be used on an actual flying MAV. A test rig was designed and built to measure the thrust, torque and RPM of the rotor. The next step was to perform an experimental parametric study to investigate the effect of the rotating speed (RPM), amplitude of blade pitch, blade airfoil profile and blade flexibility on the rotor performance. One important observation made during the tests was the presence of a sidewise force of comparable magnitude to the vertical force. It was also interesting to find that the blades were not stalling even at extremely high geometric angles of attack (45°). To help explain these phenomena, two-dimensional DPIV measurements were made on the rotor.

Fig 1. Cycloidal rotor.

Fig 2. Cycloidal rotor thrust vectors.

Previous Research The cycloidal-rotor concept was tested for the first time in 1926 by Nagler [19], when he constructed and tested a “paddlewheel” airplane. Although high lift was measured from the design, the project had to be discontinued because an efficient design solution to sustain the high structural loads on the blades from centrifugal sources could not be found. At the University of Washington, Kirsten [14] investigated the cycloidal-propulsion concept in the 1920s. A large cycloidal propeller was constructed, and its use in air vehicles was investigated. However, later on the study was focused on exploring the use of cycloidal propulsion in marine systems. Today these are used in

tug boats, providing them with the maneuverability to operate in highly constrained locations.

The rotor produced 20.5 lbs of thrust at 825 RPM giving a power loading of 10.25 lbs/hp.

In the 1930s, Wheatley [12, 13], developed a simplified aerodynamic model for a cyclogiro rotating wing. Wind tunnel tests were conducted on a 4-bladed rotor with a span and diameter of 8 ft. NACA 0012 airfoils with a chord of 0.312 ft were used in the blades. The power loading obtained was about 12.5 lbs/hp. Even though significant levels of thrust were measured, there was poor agreement between the theoretical analysis and the test results.

Kim et al. [8 – 11] constructed and tested cycloidal rotors of diameters 5.6, 4.6 and 1.3 ft at the Seoul National University (SNU). All the configurations used NACA 0012 airfoil sections. Some of the specifications of these rotors are given in Table 1. An analysis was conducted using CFD software to predict the aerodynamic characteristics of the cycloidal rotor.

After these early tests, no significant research on cycloidal rotors was reported for decades, until Bosch Aerospace realized their potential in the propulsion of “Lighter Than Air” (LTA) vehicles in 1998 [15 – 17]. Bosch Aerospace, working in collaboration with the Mississippi State University, developed and tested a 6-bladed cycloidal rotor with a diameter and span of 4 ft. The blades had a chord of 1 ft and used NACA 0012 airfoil sections. The maximum pitching angle was fixed at 25 degrees. Rotational speeds up to 650 RPM were tested. The rotor generated 138 lbs of thrust and consumed 28 hp at 600 RPM, i.e. the power loading was around 10.9 lbs/hp. McNabb [18] at the Mississippi State University developed a computer simulation of the cycloidal rotor and the predictions were found to be within five percent of the test results. In 2003, Labiche Aerospace constructed and tested a 3bladed cycloidal rotor with a diameter and span of 2 ft. Table 1. Previous cycloidal rotors. Diameter Span (ft) (ft) Wheatley 1935 Bosch 1998

Number Maximum Reynolds of RPM Number blades

8

8

4

700

5.3X105

4

4

6

550

6.7X105

Labiche‘03

2

2

3

825

1.7X105

SNU ‘06

5.6

3.3

4

450

5.6X105

SNU ‘04

4.6

3.3

4

500

3.4X105

SNU ‘06

1.3

1.1

4

1500

9.5X104

Technion

0.36

0.36

2

5000

4X104

*UMD

0.5

0.5

3,6

1200

1.5X104

Isosilevskii and Levy of the Technion-Israel Institute of Technology conducted a combined experimental and numerical study of a cyclogiro rotor operating at chord Reynolds numbers of about 40,000 [22]. A CFD study revealed the complex flow field experienced by the cycloidal rotor with complex interactions between the blades. The experimental tests were conducted on a model with a span and diameter of 4.3 inches. The blades used NACA 0015 airfoils and had a chord of 0.9 inches. The model could be tested either in a 2-bladed or 4bladed configuration. Numerical predictions showed good agreement with experimentally measured timeaveraged forces. The specifications of all these cycloidal rotors are summarized in Table 1. All the tests mentioned above were hover tests. Many of these studies reported higher power loading (thrust/power) for cycloidal rotors compared to a conventional rotor. One of the most challenging tasks in MAV design is to increase the endurance of the vehicle. High power loading is the key factor for endurance and is the primary impetus for the current research on cycloidal rotors as an efficient means of propulsion for a micro air vehicle. One of the main goals of this study is to carefully investigate whether these claims for higher power loading can be substantiated especially at micro-scales. For this purpose, a parametric study was performed to optimize a cycloidal rotor and compare its power loading to that obtained from a conventional micro rotor operating at the same disk loading.

Cycloidal Rotor Design A preliminary design study was performed to obtain an estimate of the weight and the installed power required for a flight capable cycloidal MAV. The conceptual design used two contra-rotating 3-bladed cycloidal rotors as shown in Fig. 3. A weight breakdown of the vehicle components is provided in Table 2. As the first step, a single cycloidal rotor was built and tested. The main challenge was to design the rotor with the least weight and mechanical complexity because it

causes an offset between the axis of the shaft and the offset ring denoted as L2 in Fig. 4. The three linkages for pitching the blade are connected to the offset ring 120° apart. The other end of the linkages is connected to each of the blades at a point (B) aft of the blade pitching axis (A) (Fig. 5). Together, the system comprises a crankrocker type four-bar linkage, which is used to accomplish the required change in blade pitch angle as shown in Fig. 5. The blades can be set at different pitching amplitude by changing the offset length, L2.

Fig 3. Conceptual design of a cycloidal rotor MAV.

Table 2. Component breakdown of cycloidal MAV. Component

Weight (g)

% Total

Electronics

22

7.7

Li-Po Battery (1500 mAh) Motors

40

13.9

39

13.6

Servos

12.5

4.4

Rotors (Combined) Structure

155

54.0

18.5

6.4

Total

287

100

had to be used on a flying vehicle. Fig. 3 shows the cycloidal MAV rotor, which has a three-bladed design with a diameter and span of 6 inches. The blades used NACA 0010 airfoil sections with a one inch chord. The main structural elements of the design consisted of two carbon fiber end plates, to which each of the blades was attached. The blade was pitched about two pitch bearings, one on the end plate at the root end and the other at the tip. The end plates were also connected to each other by the hollow carbon fiber rotor shaft, which rotates about two bearings at the root end. The mechanism devised for achieving the required blade pitch motion around the rotor is a passive system and the only power penalty incurred for its operation is the friction associated with its moving components. The blade pitching mechanism consists of mainly two bearings, bearing 1 and bearing 2 as shown in Fig. 4. The inner ring of bearing 1 is fixed on the rotating shaft and the outer ring is fixed inside the offset disk such that the axis of the offset disk is offset by a length L2 from the shaft axis, as shown in Fig. 4. The offset disk is fixed inside the inner ring of bearing 2. An offset ring is fixed outside the outer ring of bearing 2. This arrangement

The blades were fabricated using a +45/-45 carbon composite prepreg wrapped around a foam core. The symmetric NACA 0010 profile was used for the airfoils because of the constraint that they must operate effectively at both positive and negative angles of attack. The rotor was driven by a brushless Hacker Motor, Model B20 36S. Previous tests on the University of Maryland prototype rotor have shown that mechanical power losses constituted almost 75 percent of the total power consumption [21]. Therefore, extreme care was taken while building the rotor to reduce friction and mechanical interference. The fabricated rotor (excluding the motor) weighed 78 grams. The rotor was carefully balanced to minimize vibration levels. The rotor structural design went through a large number of iterations to reach the stage where the rotor could operate at a speed of 2,000 RPM without any mechanical problems. At 2,000 RPM and 40° pitching amplitude, the rotor produced 150 grams of thrust, which is sufficient for the cycloidal MAV to hover (using two such rotors).

Bearing 1 Rotor Shaft

Bearing 2 Offset Disk Offset=L2

Linkage Offset Ring

Fig. 4 Blade pitch changing mechanism.

Parametric studies Experimental Setup Several experiments were conducted on the prototype cycloidal rotor (Fig. 6) to investigate the effect of the rotating speed (RPM), amplitude of blade pitch, blade airfoil profile, and the blade flexibility on rotor performance. A test setup was designed and built to measure the thrust, torque and RPM of the rotor. The schematic of the experimental setup is shown in Fig. 7. The thrust was measured by a thrust load cell and the torque was measured using a torque load cell. A Hall sensor was used to generate a 1/rev signal to measure the RPM. The total power was determined from the torque and RPM measurements. Test data was acquired on a laptop by using LabView data acquisition software. A range of tests were performed to characterize the performance of the rotor. Measurements were taken at blade pitching amplitudes of 25°, 30°, 35°, 40° and 45°, and for rotational speeds ranging from 400 to 2,000 RPM.

camber. Figure 8 shows the different kinds of blade profiles tested. All the blades tested had a span of six inches and a uniform chord of one inch.

DPIV Setup Schematic of the setup used for the digital image velocimetry measurements are shown in Figs. 9 and 10. As shown in the schematics, the measurements were made by placing the camera and laser source at two different orientations. In both cases, the camera was placed orthogonal to the laser light sheet.

A L4 B L1

L3 Fig. 6 Cycloidal MAV rotor.

L2

Fig. 5 4-bar blade pitch changing mechanism.

The blade profiles tested were the NACA 0010, reverse NACA 0010 (where the trailing edge faces the flow instead of leading edge), 6% thick flat plate blades with different symmetric leading edge wedge angles (12°, 8°, 5° and 3°), 6% thick flat plate blades with symmetric 5deg leading edge and trailing edge wedge angles (5deg LE&TE), flat plate blades with symmetrically sharpened leading edges (having 3%, 2% and 1% thickness-to-chord ratios) and 3% thick sharpened leading edge flat plate blade with 7% circular

Fig. 7 Experimental setup. The first setup (setup A) was prepared to study the shed wake behind the trailing edge of the rotor blades. The laser sheet was oriented along the span of the rotor, as shown in Fig. 9. The spatial location of the tip vortices, and their strength can be measured. This is essential for

understanding the effects of the induced flow field produced by the tip vortices. In the second setup (setup B), the positions of the laser and camera were swapped, as shown in Fig. 10. Therefore, in setup B, the laser sheet is placed at the midspan of the rotor, and the camera placement provides chordwise flow velocities. With setup B, two sets of measurements were taken. In the first set of measurements, the camera was focused on a larger area below the rotor to study the rotor wake. In the second setup, the camera was moved closer to the rotor to capture the flow structures around the blades.

A fully-articulated optical arm was used to locate the light sheet in the required region of focus. The two lasers were fired with a pulse separation time of 10 µs. The interrogation region was focused on a particular region of interest within the image using 200-by-150 nodes, with a spatial resolution of 1 mm on either side.

The flow at the rotor was seeded with a mineral oil fog. The entire test area was uniformly seeded before each sequence of measurements. It should be remembered here that DPIV technique actually measures the flow velocity of the seed particles. Therefore, it is essential to have as small a particle as possible to prevent particle tracking errors. The average size of the seed particles were 0.22 microns in diameter, which was small enough to minimize the particle tracking errors [23]. The DPIV system included dual Nd:YAG lasers that were operated in phase synchronization with the rotor, the optical arm to transmit the laser light into the region of interrogation, a digital CCD camera with 2 mega-pixel resolution placed orthogonally to the laser light sheet, a high-speed digital frame grabber, and DPIV cross-correlation analysis software. The laser could be fired at any blade phase angle around the shaft, enabling DPIV measurements to be made at any required wake age.

Fig 9. DPIV setup A

Fig 10. DPIV setup B

PIV Image Processing

Fig 8. Cross sections of all the blades tested The laser light sheet was capable of being pulsed at frequencies up to 15 Hz, and was used in synchronization with the rotor frequency to illuminate planes in the flow field.

For image processing, a recursive technique called deformation grid correlation was used [24]. This procedure is based on window shifting technique; however, the second window is both sheared and translated instead of just using simple translation [25]. This technique, especially the introduction of shear, was found to be appropriate for measuring the high velocity gradients found inside rotor wake flows. Detailed analysis on the uncertainty associated with the process has also been made [25].

Results and Discussion Rotor Forces

rotational speed. Therefore, if φ was a result of the blade kinematics, it should not have varied with rotational speed. Therefore, the fact that

The coordinate system used for the cycloidal rotor is shown in Fig. 11. The azimuthal position of the blade, Ψ is measured counter clockwise from the negative Z-axis. The blade pitch angle, θ, is measured with respect to the tangent of the blade’s circular path. Intuitively, it appears that with this kind of blade kinematics, the rotor should only produce a vertical force (Tz). However, the tests showed that the rotor also produces a sidewise force (Ty), whose magnitude is comparable to that of the vertical force (Figure 11). This was further confirmed by the DPIV studies, which showed the presence of a skewed wake. A detailed explanation of this phenomenon is included in the section on the DPIV results. The present experimental rig was made in such a way that the whole rotor setup can be rotated along the rotor shaft axis to any orientation along the rotor shaft axis. Basically, rotating the rotor setup, results in a change in the direction of eccentricity. At 0°, the eccentricity is pointing vertically upwards (along the positive Z axis) causing the maximum blade pitch angle to occur at the extreme top and bottom positions. The thrust load cell is oriented in such a way that it can only measure the forces along the Z-axis. Therefore, to measure Tz, the whole rotor setup is set at 0° and the force in the Z-direction is measured. To measure the sidewise force Ty, the rotor setup is rotated by 90° and locked in that position. Now Ty is pointing in the Z-direction and it can be measured by the load cell. The resultant force, TRes is calculated from the measured Tz, and Ty as,

Fig 11. Schematic of the forces produced by the cycloidal rotor

TRe s = Tz2 + Ty2

The angle made by the resultant force TRes with the vertical, φ is given by ⎛ Ty ⎞ ⎟ ⎝ Tz ⎠ To confirm whether the rotor was producing the resultant thrust at an angle φ with respect to the vertical, the rotor setup was rotated by φ (so that the resultant thrust is acting along the Z-axis) and the thrust was measured. The measured resultant thrust matched well with the resultant thrust calculated from Tz, and Ty.

φ = tan −1 ⎜

Figure 11 shows the schematic of the forces produced by the rotor. Figure 12 and 13, respectively, show the variation of Tz, and Ty with rotational speed for the rotor with NACA 0010 blades. Both the components appear to be vary with the square of rotational speed. Figure 14 shows the variation of the resultant force phase, φ , with rotational speed for the NACA blades. From the figure it can be seen that φ increases with the rotational speed. The blade pitch angle variation along the rotor azimuth is independent of the

Fig 12. Vertical force vs. RPM for the rotor with NACA 0010 blades. Φ varies with rotational speed proves that the reason for the sidewise force may be mainly aerodynamic and partly a result of the bending and torsional deformation of the blades caused by centrifugal loads. The aerodynamic reason behind this phenomenon is explained in detail in the DPIV section. In the remaining part of the paper, the thrust refers to the resultant thrust (TRes).

generation rotor model results for which 75% of the total power consumed was mechanical power. This is mainly due to the fact that the present rotor is almost 1/5th the weight of the previous generation of rotor, even though the rotors have the same blade span and diameter.

Fig 13. Sidewise force vs. RPM for the rotor with NACA 0010 blades.

Fig 15. Power breakup for the rotor with NACA 0010 blades.

Effect of Rotational Speed (RPM)

Fig 14. Resultant force phase vs. RPM for the rotor.

Power Breakdown The total power required for the rotor was obtained from the torque and RPM measurements. The total power includes the aerodynamic power consumed by the blades and the mechanical power required for the whole mechanism. Aerodynamic power includes the induced power, profile power, rotational losses and the aerodynamic power required for pitching the blades. Tare tests were carried out at different RPMs after removing the blades to measure the mechanical power consumed by the rotor during its rotation. The mechanical power measurements were subtracted from the total power measurements to obtain the aerodynamic power required to rotate the blades. Figure 15 shows the power breakdown for the rotor with NACA 0010 blades. Mechanical power constituted only around 10 percent of the total power. This was a major improvement from the previous

Figures 16 and 17 respectively show the variation of resultant thrust and aerodynamic power for the rotor with NACA 0010 blades with RPM, for different blade pitching amplitudes. As expected, the thrust and the aerodynamic power vary as the square and cube of RPM respectively. Figure 18 shows the variation of aerodynamic power loading (thrust/power) with thrust for different pitching amplitudes. The thrust level was changed by varying the rotational speed. As expected, the power loading varies as (RPM)-1 because thrust is a function of the square of RPM and power varies as the cube of RPM. From Fig. 18, it can also be seen that for the rotor with regular NACA 0010 blades the optimum angle (for maximum power loading) is 40°. Power loading decreased as the pitching amplitude was increased to 45°. As the RPM changed from 400 to 2,000, the chord Reynolds number changes by a factor of five. To see any effect of Reynolds number, the thrust and the power has to be non-dimensionlised with RPM using T CT = ρ AΩ 2 R 2 P CP = ρ AΩ3 R3 where T is the resultant thrust, P is the aerodynamic power, Ω is the rotational speed, ρ is the density of air

Fig 16. Thrust vs. RPM for the rotor with NACA 0010 blades.

Fig 19. CT vs. RPM for the rotor with NACA 0010 blades.

Fig 20. Cp vs. RPM for the rotor with NACA 0010 blades. Fig 17. Power vs. RPM for the rotor with NACA 0010 blades.

Fig 18. Power loading vs. thrust with NACA 0010 blades.

(1.25 Kg/m3), R is the radius of the rotor and A=bD is the projected rectangular area of the cycloidal rotor. b is the blade span and D is the diameter of the rotor. The present rotor had a diameter and blade span of 6 inches. In Figs. 19 and 20, respectively, CT and CP are plotted against RPM to see the effect of Reynolds number on blade lift and power. CT remains almost constant with RPM showing that this range of Reynolds number variation is not sufficient to produce significant changes in lift performance of the airfoils. This was true for almost all the blades tested, except for the extremely flexible ones where the flexibility effects showed up at higher RPMs because of the transverse centrifugal loading on the blades. However, there is a moderate increase in CP with increase in RPM which was not expected. Again, this can be attributed to the torsional flexibility of the blades rather than any Reynolds number effects.

Effect of Blade Pitching Amplitude In Fig. 21, the variation of CT for the rotor with NACA 0010 blades is plotted with respect to the blade’s pitching amplitude for various rotational speeds. CT increases linearly from a pitching amplitude of 25° to 40°, whereas it shows either a small increase or decrease (for some rotational speeds) from 40° to 45° pitching amplitude. This may be because the NACA blades might have reached their maximum CL condition. Figure 22 shows the variation of CP with pitching amplitude, and there is no dramatic increase in power from 40° to 45° pitching amplitude. This shows that the blade might not be completely stalled at 45° pitching angle. Figure 23 shows the variation of CT with blade pitching amplitude for five different blades, namely, the NACA 0010, reverse NACA 0010, 6% flat plate with 5° leading edge wedge angle (5deg LE), 6% flat plate with 5° leading edge and 5° trailing edge wedge angle (5deg LE&TE) and 3% thick flat plate with sharpened leading edge, at a rotational speed of 2000 RPM. In all these cases, except for the regular NACA 0010, CT increases linearly from 25° to 45° pitching amplitude showing no signs of stall. From Fig. 24 it is clear that even CP increases steadily with blade pitch, clearly indicating that the stall does not occur on the blades. It was surprising to see the blades were not stalling even at a geometric pitch angle of 45°. However, from the DPIV studies, it was seen that the induced velocities seen in the rotor wake were extremely high and were comparable with the blade velocity itself. Therefore, even though the geometrical angle of attack was 45°, high induced velocities probably decreased the effective aerodynamic angles of attack and kept them below the stall values. The DPIV studies also showed the presence of a dynamic stall vortex at the leading edge, and flow reattachment at 40° pitch angle. The delay of the stall to higher pitch angles (dynamic stall) may occur because of the 1/rev pitching oscillation of the blades.

Fig 21. CT vs. pitching amplitude for the rotor with NACA 0010 blades.

Fig 22. CP vs. pitching amplitude with NACA blades.

Fig 23. CT vs. pitching amplitude for five different blades at 2000 RPM.

Fig 24. CP vs. pitching amplitude for five different blades at 2000 RPM.

Figure 25 shows the variation of the power loading (thrust/power) with RPM for the rotor with reverse NACA 0010 blades at different pitching amplitudes. Clearly 30° pitching amplitude produced better power loadings and 45° produced the worse power loading for all thrust conditions.

Fig 25. Power loading vs. thrust for the reverse NACA blades. Figure 26 shows the variation of aerodynamic power loading (thrust/power) with RPM for the 5° leading edge and trailing edge wedge angle (5deg LE&TE) blades. It can be seen that 40° pitching amplitude produced the maximum power loading at the high thrust condition (>70g) and 45° pitching angle produced almost the same power loading as the 40° case at extremely high thrust conditions (>100g).

Fig 26. Power loading vs. thrust for the 5 deg LE & TE blades. Figure 27 shows the variation of power loading for the 6% thick 5° leading edge wedge angle blades (5deg LE). In this case at extreme high thrust cases (>100g), 45° pitching

amplitude produced the maximum power loading for the same value of thrust. For the cases where T<100g, 40° pitching amplitude produced the maximum power loading. Again, the 25° pitching amplitude had the lowest power loading amongst all the cases.

Fig 27. Power loading vs. thrust for the 5 deg LE blades. Figure 28 shows the variation of power loading for the rotor with 3% thick sharpened leading edge flat plate blades. In this case, 45° pitching amplitude is better than all the other cases especially at the high thrust case. This is mainly a result of the increase in blade’s bending stiffness with pitch angle which will be explained in the next sections.

Fig 28. Power loading vs. thrust for the 3% flat plate blades. The general observation from this study, is that at the higher thrust levels (higher rotational speed), the rotor performs better when the blades are set at higher pitching amplitudes of 40° and 45°. Because all these studies

were directed towards building a hover-capable cycloidal rotor MAV, the low thrust regime is not of much interest. The reason for higher power loading at higher pitch angles can be partly because the bending stiffness of the blades increases with pitch angle. This is very evident for some of the thinner blades (3% flat plate blade) tested (Refer to Fig. 28). The other reason can be the superior aerodynamic performance of these blades at higher pitch angles. Note that the power loading varies as (RPM)-1. Therefore, it is more efficient to operate the rotor at a higher pitch angles and a lower RPMs than operating at a lower pitch angle and higher RPMs for producing the same thrust. This is true as long as the blade does not stall.

case, the thrust level was changed by changing the rotational speed. At the highest thrust, the reverse NACA blades produced 25% more power loading than the NACA blades. The 3% flat plate blade was the worst performer from flexibility effects. Figure 35 shows the variation of power loading with thrust for the 30° pitching amplitude. Again, in this case, the reverse NACA blades produce higher power loading than the regular

Effect of Blade Airfoil Profile Figure 29 shows the variation of CT with rotational speed for five blades at 25° pitching amplitude. It can be seen that the NACA 0010 produces the maximum thrust at all the rotational speeds. CT, for the 3% flat plate blade dropped at higher rotational speeds, probably a result of blade flexibility effects. Figure 30 shows the variation of CT for 30° pitching amplitude. Again, the NACA 0010 blades produced more thrust than other blades at all rotational speeds. At 35° pitching amplitude (Fig. 31), 5° leading edge wedge angle blades are very close to NACA in terms of thrust producing capability. Again, for 40° and 45° pitching amplitudes (Figs. 32 and 33), the NACA blades performed better among blades in the terms of thrust at a particular rotational speed. However, from these results, it can be concluded that at these low Reynolds numbers (< 22,000), the blade section does not have much of an effect on the lift production. Even reverse NACA or flat plates perform comparably to the NACA blades.

Fig 30. CT vs. RPM at 30 deg pitching amplitude.

Fig 31. CT vs. RPM at 35 deg pitching amplitude.

Fig 29. CT vs. RPM at 25 deg pitching amplitude. Figure 34 compares the power loading of the different blades at different thrust levels for 25° pitching amplitude. In this

NACA blades at higher thrust levels. At 35o pitching amplitude there was little variation between different blades (Fig. 36). The reverse NACA blades performed better than all of the other blades. However, at higher pitching angles of 40o and 45o the regular NACA blades performed marginally better than the reverse NACA blades (Figs. 37 and 38). Although the blade with 5o wedge angle at the leading edge (5deg LE) was better than the blade with the 5o wedge angle on both leading and trailing edge (5deg LE&TE) in terms of thrust

Fig 32. CT vs. RPM at 40 deg pitching amplitude.

Fig 35. Power loading vs. thrust at 30 deg pitching amplitude.

Fig 33. CT vs. RPM at 45 deg pitching amplitude. Fig 36. Power loading vs. thrust at 35 deg pitching amplitude. production, both blades had almost the same power loading for all the pitching amplitudes. With increase in pitching angle for the 3% flat plate, the power loading started to improve. One important conclusion from these results, is that at these low Reynolds numbers, the blade profile does not necessarily have to be a conventional airfoil shape. This is reinforced by the fact that even the reverse NACA airfoil blade perform as well as the regular NACA profile blade.

Effect of Flexibility

Fig 34. Power loading vs. thrust at 25 deg pitching amplitude.

With a cycloidal rotor the centrifugal force acts in the transverse direction, unlike a conventional rotor. In most of the present tests, the centrifugal force from the blade mass itself was about 20 times the aerodynamic force experienced by the blade. In the test rotor, the blades

have a fixed-fixed support at the ends and large bending deformation was observed especially with the less thick blades at higher rotational speeds. As far as the torsion is concerned, the blades have a fixed boundary condition at the root from the rigid pitch link and free boundary condition at the tip. The pitching axis of the blade is along the quarterchord. For most of the flat plate blades, the center of gravity (CG) is along the mid-chord and the centrifugal force is acting along the CG. Therefore the centrifugal loading can also cause torsional deformation of the blade.

quite similar flexibilities. However, the same does not hold true for the 3% flat blade, which had nearly 1/8th of the stiffness of a 6% plate blade. Most of the drop in power loading observed for the 3% thick blades was not a result of aerodynamic reasons but may be attributed to the elastic deformation of the blades. To examine this effect more clearly, similar tests were carried out on the 1%, 2%, 3%, and 6% thick 5deg LE blades with sharpened leading edge. Figure 39 shows the variation of CT with rotational speed for all the 4 blades at 40o pitching amplitude. From a purely aerodynamic perspective, not much variation in CT would be expected for these four different blades. However, the results showed otherwise because of the flexibility of the blades. The 6% thick flat plate produced a maximum of 40% more thrust than for the 1% thick blade. A clear trend could be seen from the test results. At a particular rotational speed, the lower the stiffness of the blade the lower was the thrust produced. The same held true for the power loading as shown in Fig. 40. At the maximum value of thrust for the 1% thick blade, the power loading for the 6% blade was almost 80% more than for the 1% thick blade. The conclusion from these tests, is that blade flexibility in both bending and torsional mode deteriorates the performance of the cycloidal rotor.

Fig 37. Power loading vs. thrust at 40 deg pitching amplitude.

Fig 39. CT vs. RPM for different thickness blades at 40 deg pitching amplitude. Fig 38. Power loading vs. thrust at 45 deg pitching amplitude.

Effect of Leading Edge and Trailing Edge Wedge Angle

The next step was to perform systematic tests to investigate the effects of blade bending and torsional flexibility on rotor performance. All the blades which were tested had different flexibilities. The effects of flexibility may not be very pronounced among the NACA, reverse NACA, 6% thick 5deg LE and 6% thick 5deg LE&TE blades because they had

Some of the test results shown in the previous sections showed that a conventional airfoil profile is less important for delivering performance at these low Reynolds numbers (< 22,000). In some of the cases, the reverse NACA blades produced higher power loading than regular NACA blades. It appeared as if sharpened

leading edge blade may perform better than the rounded leading edge of the conventional airfoil. Therefore, 6% thick flat plate airfoils with symmetric wedge angles of 12°, 8°, 5°, 3° and also the one with 5° wedge angle on both leading and trailing edge (5deg LE&TE) were tested at 25° pitching amplitude (Refer to Fig. 8 for the definition of wedge angle). Only symmetric sharpening of the blades can be used because the blades of the cycloidal rotor undergo periodic pitching motion and have to operate at both positive and negative angle of attack.

Fig 40. Power Loading vs. Thrust for different thickness blades at 40 deg Pitching Amplitude

decreasing the wedge angle decreases the stiffness of the blade. It is not possible to vary the wedge angle of the blade without varying the stiffness of the blade (if the chord, thickness and blade material is not changed). Therefore, some component of the variation in the performance of the blades must be attributed to the change in the flexibility of the blade from the change in wedge angle and not just the aerodynamics at the leading edge. The overall conclusion is that a moderate amount of leading edge sharpening (~5°) produced a slightly better power loading.

Fig 42. Power loading vs. thrust for different LE angle blades at 25 deg pitching amplitude.

Effect of Camber Several tests on conventional rotors at the same Reynolds number range have shown that a small amount of camber can significantly improve the performance of the rotor. However, camber is not an obvious choice for the cycloidal rotor blade because it has to operate at both positive and negative angles of attack. Therefore, if it has positive camber, as the blade traverses the upper half, it will have reverse camber for the lower half as shown in Fig. 43. However, a test was still conducted to see whether the improved performance (from positive camber) during the upper half can compensate for the degraded performance during the lower half.

Fig 41. CT vs. RPM for different LE angle blades at 25 deg pitching amplitude. As far as the thrust is concerned, all the blades were very similar, as shown in Fig. 41. As shown in Fig. 42, all the blades produced very similar power loadings except for the 3° wedge angle case. The lower power loading for this case may again be a result of blade flexibility effects because

Two 3% thick flat plate blades, one with zero camber and the other with 7% circular camber were tested at the lowest pitching amplitude of 25o. The results did not turn out as expected. In Fig. 45 it can be seen that the power required by the cambered blades was almost three times that of the uncambered blades. This clearly shows that the reverse camber caused the blades to stall. It is worthwhile to note that even after complete stall during the entire lower half, the cambered blade produced more

thrust than the zero cambered blade at all rotational speeds (Fig. 44). From Fig. 46 it can be seen that the power loading for the cambered blade is only half that of the uncambered blades. However, it may dramatically improve the performance of the cycloidal rotor if the camber of the blade can be actively changed so that it has a positive camber during both the upper and lower half of the cycle.

DPIV Results The DPIV results are discussed in two categories: (1) Spanwise view, where the trailing tip vortices were analyzed, and (2) chordwise view, where the analysis on the sidewise force produced by the rotor was performed.

Fig 45. CP vs. RPM for cambered blades at 25 deg pitching amplitude.

Fig 43. Motion of a cambered blade around azimuth.

Fig 46. Power loading vs. thrust for cambered blades at 25 deg pitching amplitude.

Tip Vortex Measurements

Fig 44. CT vs. RPM for cambered blades at 25 deg pitching amplitude.

The tip vortices trailing from the cycloidal rotor blade will affect the induced flow distribution, and can play a substantial role in determining the overall performance of a cycloidal rotor. This can be expected based on the observations of a conventional rotor wake [26]. These vortices are even more important at low Reynolds numbers because of their relative size and flow structure compared with the blade chord [27]. Therefore,

quantitatively measuring the strength of the tip vortices becomes essential.

Fig 47. Schematic showing the evolution of the tip vortices. The tip vortex measurements were made using setup A (Figure 9). The camera was focused on the lower blade and the wake below it. A finite wing (cycloidal rotor blade) always produces tip vortices from its free tip because of the pressure difference between the upper and lower surfaces of the blade [28]. Because each of the cycloidal rotor blade has two free tips, it is expected to produce two counter rotating tip vortices, as shown in Fig. 47. Unlike conventional rotors, where the root of the blade experiences very low velocity, the blade tips of a cycloidal rotor experience the same velocity. As a result, the tip vortex from both the left and right tips should be similar. This phenomenon is captured by the DPIV measurements as shown in the Fig. 48 (a – c) , which shows the evolution of both tip vortices. In these figures, the distances are non-dimensionlised by the blade span and the velocity is normalized by blade velocity. Each figure shows the phase averaged velocity vectors superimposed on the vorticity contours for a wide range of wake ages. 0° wake age corresponds to the alignment of the laser light sheet with the trailing edge of the blade as the blade reaches the lower most point in its trajectory. The measurements were performed at 6 different wake ages, namely, 30°, 45°, 60°, 75°, 90° and 105°. Because there are three blades, after 120° the flow structures become repeated.

Figure 48 (a) shows the vorticity contours at 30° of wake age. In the figure, the two trailing vortices from the right blade tip and the two vortices from the left tip can be clearly seen. Adjacent tip vortices on the same side are 120° apart, which is the phase angle between the blades. As expected, it can be seen that the left and right tip vortices have opposite vorticity. Because the entire blade operates at the same speed and because all the blade sections are set at the same pitch angles, it could be expected that the left and right tip vortices should have the same strength. However, for all the wakages (Fig. 48 (a – c)), it can be seen that the right vortex is stronger than the left one. This may be attributed to torsional deformation in the blade. The camera is captured the tip vortices trailed by the blades as it reaches the lowest most point of its trajectory. As explained in the previous section, the left end of the blade has a fixed boundary condition for torsion and the right end has a free torsional boundary condition. Because the center of gravity (CG) of the blade is behind the pitching axis, at this blade location the centrifugal force produces a nose-up torsional deformation that increases the geometric angle of attack of the blade at all sections [Fig. 49]. However, because of these particular torsional boundary conditions, the blade will twist causing a maximum increase in the geometric angle of attack at the right end and zero increase at the left end. Higher angle attack at the right end can lead to a stronger tip vortex. Figures 48 (b – c) (45° to 105° wake age) shows the downward convection of the vortices. It can also be seen that as the vortex convects downwards, it gradually diffuses. Figure 50 shows the time-averaged flow field in the rotor wake where the color contour represents absolute velocity. A very distinct slipstream boundary can be seen. Because the cycloidal rotor produces a thrust, a contracting wake structure is expected. This can be clearly seen from the figure. The contracting wake structure of the cycloidal rotor looks similar to that of a conventional rotor [27].

Vortex Properties Figure 51 shows the variation of swirl velocity along a section (section AA as shown in Fig. 47) across the wake. The distance is non-dimensionlised by the blade span, and the measured flow velocity is normalized by the blade velocity. As the wake age increases, the core sizes of the both the left and right tip vortices increases, which are associated with reductions in their peak swirl velocity. This shows the influence of viscous and turbulent diffusion [29]. An important observation is the

(a) ζ = 30deg

(c) ζ = 60deg

(b) ζ = 45deg

(d) ζ = 75deg

(e) ζ = 90deg (f) ζ = 105deg Fig 48. DPIV measurements showing the presence of a pair of tip vortices from either side of the cyclocopter blade, ζ = 30° to 105deg.

change in fluid momentum. Therefore, the velocity of the fluid in the wake should have components in both the vertical and horizontal direction, which may produce a skewed wake, as shown in Fig. 52.

Fig 49. Blade section at the lower most point.

Fig 51. Velocity profiles across the rotor wake by taking the sections across the tip vortices at all the six wake ages. The PIV setup B (Fig. 10) was used to obtain the velocity measurements in the plane of the rotor perpendicular to the blade span. The results are shown in the Fig. 53. The direction of the velocity vectors clearly indicates the presence of a skewed wake.

Fig 50. Time averaged velocity measurements showing the wake contraction of the cycloidal rotor. high induced velocities in the wake; the induced velocities are 60 to 70 % of the blade velocity. This can produce a high induced angle of attack ( φ ) at the blade. The angle attack of the blade, α is given by, α = θ −φ where θ is the blade pitch angle. Therefore, the high induced angle reduces the blade angle to attack to a smaller value and can explain the reason why the blades do not stall even at a geometric pitch angle of 45°.

Chordwise View As mentioned previously, the force measurements showed the presence of a sideward force even though it was not expected (Fig. 52). The presence of sideward force means a horizontal

Fig 52. Schematic of the skewed wake structure.

To capture the leading edge vortex, the camera and laser source was kept as in setup B and the camera was moved closer to the rotor to focus on the blade when it reaches its maximum pitch angle. It can be seen from Fig. 54 that the DPIV measurements clearly captured the leading edge vortex on the blade. This can play additional role in enhancing lift on the blades.

Conclusions

Fig 53. DPIV measurements using setup B showing the skewed wake structure. The skewness of the wake and the sidewise force produced by the cycloidal rotor can be explained using the analogy of a spinning cylinder in a flow. In the present case the cycloidal rotor acts like the spinning cylinder [30], and the inflow acts like the freestream.

In this paper the viability of a cycloidal rotor for propelling an MAV-scale vehicle has been investigated through performance and flow field measurements. The conceptual design study of a cycloidal-rotor MAV was performed. The cycloidal-MAV as designed weighs 290 grams and uses two contra-rotating three-bladed rotors for propulsion and control. The cyclocopter rotor was built light weight enough to be used on an actual flight capable MAV. An experimental setup was designed and built to measure the thrust, torque and RPM of the rotor. A parametric study was conducted to investigate the effect of the rotating speed (RPM), amplitude of the blade pitch, blade airfoil profile, and blade flexibility on the rotor performance. The rotor with normal NACA 0010 blades, at 2,000 RPM and 40° pitching amplitude produced 150 grams of thrust, which is sufficient for the cycloidal MAV to hover (using two such rotors). The parametric study shows that the best power loading (thrust/power) was obtained at 40° pitching amplitude for most of the stiff blades tested. Because the power loading varies as (RPM)-1, it is more efficient to operate the rotor at a higher pitch angle and a lower RPM than operating at a lower pitch angle and higher RPM to produce the same amount of thrust, provided the blade does not stall.

Fig 54. DPIV measurements showing the leading edge vortex on top of the blade In a cycloidal rotor, there a 1/rev high amplitude (25 to 45°) pitching motion associated with the blade along with the rotation. This is an ideal condition for the development of a leading vortex prolonging the stall to a higher angle of attack (i.e. dynamic stall – see Ref 31).

The blades that are stiffer in bending and torsion produced better power loading than the flexible ones. For the flexible blades, the power loading improved with an increase in pitch angle because the pitch angle increases the stiffness of the blade in the direction of bending caused by the centrifugal force. It is interesting to note that reverse NACA 0010 blades provided better power loading than for normal NACA 0010 blades at low blade pitching amplitudes. However, at a fixed operating speed, the blade that produced the maximum thrust at all pitch angles was regular NACA 0010 airfoil. Fixed camber had a detrimental effect on the rotor performance. However, the performance of the cycloidal rotor can be dramatically improved if the camber of the blade can be actively changed so that it has a positive camber at all the azimuthal locations. The blades tested with different leading edge wedge angles did not differ significantly in their performance.

One of the notable observations from the rotor tests is the presence of a sideward force of magnitude comparable to the vertical force. The fact that the ratio of the sidewise force with the vertical force (phase of the resultant force) varies with rotor speed proves that the reason for the sidewise force is aerodynamic and has limited relationship to the kinematics of the blade. The DPIV wake measurements also showed the presence of skewed flow, which reinforced the sideward force measurements. Another interesting observation was the absence of blade stall even at an extremely high pitch angle of attack of 45°. This could be partly explained using the DPIV measurements, which showed extremely high induced velocities in the rotor wake. This can keep the aerodynamic angle of attack below the stall value. The chordwise DPIV measurements also showed the presence of a leading edge vortex and flow reattachment aft of leading edge. Because the blades are periodically pitching at an amplitude of 45° and at a frequency of 1/rev, the leading edge vortex may be a dynamic stall vortex. This may be another reason for delaying the stall to higher angle of attack, and also produce more lift.

4.

Grasmeyer, J. M. and Keennon, M. T., “Development of the Black Widow Micro Air Vehicle,” 39th Aerospace Sciences Meeting and Exhibit, Reno, Nevada, January 8—11, 2001.

5.

Keenon, M. T. and Grasmeyer, J. M., “Development of the Black Widow and Microbat MAVs and a Vision of the Future of MAV Design,” AIAA/ICAS International Air and Space Symposium and Exposition, The Next 100 Years, Dayton, OH, July 17, 2003.

6.

Bohorquez, F. and Pines, D., “Rotor and Airfoil Design for Efficient Rotary Wing Micro Air Vehicles,” Proceedings of the 61st Annual American Helicopter Society Forum, Grapevine, TX, June 1—3, 2005.

7.

Hein, B. and Chopra, I., “Hover Performance of Micro Air Vehicles: Rotors at Low Re,” Proceedings of the American Helicopter Society International Specialists' Meeting on Unmanned Rotorcraft, Chandler, AZ, January 18—20, 2005.

8.

Kim, S.J., Yun, C.Y., Kim, D., Yoon, Y. and Park, I., “Design and Performance Tests of Cycloidal Propulsion Systems,” Proceedings of the 44th AIAA/ASME/ASCE/AHS Structures, Structural Dynamics, and Materials Conference, Norfolk, VA, April 7—10, 2003.

9.

Hwang, I. S., Hwang, C. P., Min, S. Y., Jeong, I. O., Lee, C. H., Lee., Y. H. and Kim, S. J., “Design and Testing of VTOL UAV Cyclocopter with 4 Rotors,” Proceedings of the 62nd Annual Forum of the American Helicopter Society, Phoenix, AZ, April 29—May 1, 2006.

Acknowledgements This research was carried out under the Multidisciplinary University Research Initiative (MURI) grant W911NF0410176 from the Army Research Office with Dr. Tom Doligalski as Technical Monitor. The authors would like to acknowledge Mr. Bradley Johnson for his contributions in the PIV part of this work.

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