This paper describes the development and flight testing of a personal air vehicle by team Harmony for the Boeing GoFly X-prize challenge. For the $1mil grand prize, aircraft were scored by compact size, speed, low noise, and endurance. The team chose a coaxial electric helicopter configuration to maximize rotor area and reduce disk loading for efficiency and acoustic benefits. The rotors were designed through an extensive parametric study using an in-house performance code. Air loads were modeled in HPCMP CREATE™-AV Helios for validation, then used in an inhouse acoustics solver to estimate sound pressure levels. A quiet electric power train was developed, as well as a custom 11kWh, 200lb (90.7kg) battery pack. The flight dynamics of the configuration were modeled and the stability analyzed. Structural analysis was utilized in designing key load-bearing parts. Flight control was implemented with dual, independent, electronically coupled swashplates. First, a 1/3rd scale prototype aircraft was developed to validate the design and acoustic predictions. Then a full-scale, 520lb (235.4kg) prototype with an 8.5ft (2.59m) rotor diameter was developed and accumulated 19.5hrs of testing time. During hovering, the sound pressure levels at 50ft (15.24m) were 73dBA, remarkably low for a rotorcraft. The results of this study underscored the endurance limitation of electric flight due to poor battery performance, as well as the need for reliable, lightweight hardware for such applications.

The paper discusses an experimental parametric study to maximize the hover efficiency of a coaxial rotor system for a micro air vehicle (MAV) that could potentially be launched from a 40 mm grenade launcher to achieve improved mission range and endurance. Prior to testing the rotors in the coaxial configuration, isolated rotor tests were performed on acustom-built hover test stand. The rotor diameter was kept constant at 6 inches (0.015 m), and used two untwisted carbon fiber blades with rectangular planform shape. The blades utilized thin circular cambered plate airfoil sections for improved performance at low Reynolds numbers (Re ≈ 30,000). The parameters that were varied include blade thickness-to-chord ratio (t/c), chord length, pitch angle, and Reynolds number. Thickness-to-chord ratio did not have a significant impact on figure of merit below 4%. Increasing the blade chord at a constant Reynolds number and t/c significantly improved hover efficiency until a chord length of 16.6 mm (solidity, σ=0.14). Across all the cases tested, the optimal pitch angle was around 16- 17 degrees. The optimal rotor from the isolated rotor experiments was used as the baseline rotor for coaxial rotor testing where the effect of vertical rotor separation was investigated with the same blade pitch angle for upper and lower rotors. Overall, the rotor separation had negligible effect on the performance of the upper and lower rotors for a separation distance range from 0.5R to 2R. The highest figure of merit obtained for the coaxial rotor was around 0.55 at Re = 30,000. Across all the vertical separations and disk loading the torque-balanced coaxial rotor system produced almost 1.5 times the thrust of an isolated rotor, which was set at the same pitch angle.

This paper describes the development and flight testing of a compact, re-configurable, hover-capable rotary-wing micro air vehicle that could be tube launched for increasing mission range. The vehicle design features a coaxial rotor with foldable blades, thrust-vectoring mechanism for pitch/roll control and differential rpm for yaw control. The vehicle was stabilized using a cascaded feedback controller implemented on a 1.7-gram custom-designed autopilot. Wind tunnel tests conducted using a single-degree-of-freedom stand demonstrated gust-tolerance up to 5  m/s, which was verified via flight testing. Finally, the 366-gram vehicle was launched vertically from a pneumatic cannon followed by a stable projectile phase, passive rotor unfolding, and transition to a stable hover from arbitrarily large attitude angles demonstrating the robustness of the controller.

This paper discusses the flight testing and system identification of a compact, re-configurable, rotary-wing micro air vehicle concept capable of sustained hover and could potentially be launched from a 40mm grenade launcher. By launching these energy-constrained platforms to a target area, the mission range could be significantly improved. The vehicles used in the paper has a mass of 345 grams. The vehicle design features coaxial rotors with foldable blades, and a thrust-vectoring mechanism for pitch and roll control. Yaw control was accomplished with a specialized counterrotating motor system composed of two independently controlled motors. A comprehensive set of flight experiments were performed to excite the longitudinal, lateral, directional, and heave modes of the vehicle. A linearized statespace model was derived from the flight test data. The model showed that lateral and longitudinal dynamic modes were decoupled from each other and from the other modes of the vehicle. Due to the axisymmetric nature the vehicle design, the longitudinal and lateral stability and control coefficients and their eigenvalues were nearly identical. All of the aerodynamic damping terms were negative and stabilizing except for the pitch and roll acceleration modes. These two unstable modes necessitated the need for pitch and roll feedback control. The final flight dynamics model was compared against flight test data for each state, and the model shown good agreement with the experimental data.

This paper presents an experimental parametric study to maximize the hover efficiency of a coaxial rotor system for a micro air vehicle (MAV) that could be launched from a 40 mm grenade launcher. Towards this, isolated rotor experiments were first conducted to optimize the performance of a single rotor at low Reynolds numbers (Re) through varying parameters including blade pitch angle, thickness-to-chord ratio (t/c), chord length, and Reynolds number. Results showed that t/c had minimal impact on the figure of merit FM) below 4%, while increasing blade chord length significantly improved hover efficiency until a chord length of 16.6 mm (solidity = 0.10). The optimal pitch angle for the isolated rotor was around 16 degrees, and the maximum FM was 0.59 at Re = 70,000. Coaxial rotor experiments were performed using the optimal isolated rotor as the baseline. The vertical separation between the upper and lower rotors had negligible impact on performance for a separation distance range from 0.5R to 3R. However, below a separation distance of 0.5R, the thrust and FM of the lower rotor increased and that of the upper rotor decreased. The highest FM obtained for the coaxial rotor was around 0.60 at Re = 30,000, which remained relatively constant across all vertical separations tested. The coaxial rotor system produced almost 1.66 times the thrust of an isolated single rotor. When compared at the same disk loading, the power loading (thrust/power) for the coaxial rotor was similar to that of an isolated single rotor. A comprehensive analysis (RCAS) was used to model the MAV-scale isolated rotor in hover and the analytical predictions agreed well with the experiments. In order to generate the airfoil lookup table for the RCAS model, water tunnel experiments were conducted at Re = 44,000 using a scaled-up wing with the exact same circular-cambered plate airfoil used in the coaxial rotor blades.