System identification of a thrust-vectoring, coaxial-rotor-based gun-launched micro air vehicle in hovering flight

This paper describes the development and flight testing of a compact, re-configurable, rotary-wing micro air (MAV) vehicle capable of sustained hover and could potentially be launched from a 40mm grenade launcher. Launching the vehicle as a projectile over the target area could significantly improve the mission range for these energy-constrained platforms. The MAV design features coaxial rotors with foldable blades, and a thrust-vectoring mechanism for pitch and roll control. Yaw control was accomplished by using a specialized counter-rotating motor system comprised of two independently controlled motors. Passive unfolding of the coaxial rotor blades in flight utilizing centrifugal force was demonstrated. A cascaded feedback control strategy was implemented on a 1.7 gram custom-designed autopilot. Systematic wind tunnel tests were conducted with the vehicle on a single degree of freedom stand, which proved the ability of the controller to reject wind gusts up to 6 m/s and stabilize the vehicle during the powered axial descent phase. Different phases of the gun-launched flight sequence were independently verified through targeted flight tests. Free f light testing conducted both indoors and outdoors verified that the vehicle could hover and fly forward in moderate winds. In-flight drop tests were conducted by throttling down the vehicle from a high altitude to attain high axial decent speeds followed by recovery using the rotor thrust to aggressively brake the descent and achieve a stable hover. Finally, the vehicle was launched vertically from a pneumatic cannon followed by a stable projectile phase utilizing the f ins, passive rotor unfolding, and final transition to a stable hover from arbitrarily large attitude angles demonstrating the robustness of the controller, as well as all the sub-systems of the vehicle operating in perfect harmony.

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.

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 of a compact and re-configurable rotary-wing micro air vehicle (MAV) that is capable of sustained hover and could potentially be launched from a 40 mm grenade launcher in the future. Launching the vehicle as a projectile up to the point of operation could significantly improve the mission range for these energy-constrained platforms. The MAV design used coaxial rotors with forldable blades, a thrust-vectoring mechanism for pitch and roll control, and a strict constraint on the outer diameter, which was relaxed to 52 mm for this study. Yaw control was accomplished by using a specialized counter-rotating motor that is composed of two independently controlled motors. Passive unfolding of the coaxial rotor blades utilizing centrifugal force was demonstrated. The vehicle attitude was stabilized in hover using a closed-loop proportional-derivative controller implemented on a 1.7 gram custom autopilot. Through systematic trimming and tuning of the feedback gains, the vehicle was able to achieve stable hover. When the vehicle was subjected to large impulsive itch and roll perturbations, the feedback controller was able to successfully reject the disturbance and retgurn the vehicle to a stable hover within a second. In parallel, an analogue of the flying vehicle or a "dummy" was built and launched using a pneumatic canon to understand the dynamics of the vehicle during the projectile phase without risking the actual flying vehicle. The launch demonstrated that with the right center of gravity location, the present vehicle configuration could be stable during the projectile flight even without fins.

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 presents the design and development of a pitching and plunging (flapping) mechanism for small-scale flight. In order to harness the unsteady lift mechanisms, used by most insects, a biologically inspired flapping/pitching device in conjunction with a rotary wing concept was developed and built. This mechanism attempts to replicate some of the aerodynamic phenomena that enhance the performance of small fliers, replacing the periodic translational motion with a unidirectional circular motion while actively flapping and pitching the rotor blades. In order to find the appropriate combination of phase, amplitude, frequency and rotational speed that leads to enhancement in lift, the device requires uncoupled independent pitch and flap actuation systems to permit the complete mapping of the parameter space. In the device under consideration the phase shift between the flapping and the pitching oscillations can be adjusted from 0 to 360 degrees over a wide range of rotational speeds. Maximum flapping and pitching amplitudes of +/- 23 degree(s) and +/- 20 degree(s) respectively can be attained. Linear displacements of two coaxial shafts are translated into the flapping and pitching motion of the rotor blades. The mechanism was designed to minimize the actuation stroke so that smart materials and conventional actuators such as motors and cams could be used. Kinematic analysis as well as experimental tests were performed. Using a customized test stand thrust and torque produced by the rotor were measured at different angles of attack, in steady-state and under periodical pitching actuation. The results showed that hover efficiency was considerably increased for a range of thrust coefficients. The device was developed based on the University of Maryland's rotary wing Micro Air vehicle (MAV) the MICOR (MIcro COaxial Rotorcraft), an electrically driven 100 g coaxial helicopter. It is anticipated that active flapping and/or pitching could be implemented in the prototype to improve its aerodynamic performance. The present paper will discuss the design and development process of a rotating/pitching/flapping mechanism for MAVs. Test results indicate that unsteady pitching motion can be used to include the aerodynamic effect of delayed stall. Performance measurements confirm that unsteady pitching motion improves efficiency in hover.

In this paper, a new design methodology is introduced to automate the configuration layout design and geometric sizing of rotary-wing micro air vehicles (MAV). The objective of this design-optimization problem is to organize a given set of components and payloads such that the resulting flight vehicle has the most compact overall size and still fulfils the given physical and control constraints. Genetic algorithm (GA) is chosen as the optimization engine because of its proven robust performance. A detailed discussion is presented to explain how the rotary-wing MAV design problem can be formulated as a GA optimization problem. From the case study performed, it is demonstrated that the proposed methodology is able to achieve the design goal.

Micro air vehicles have emerged as a popular option for diverse robotic and teleoperated applications in both open terrain and urban environments because of their inherent stealth and portability. To perform many of the tasks envisioned for micro air vehicles, agility is essential. To date, research efforts to improve agility have focused primarily on constructing complex controllers to enable existing vertical-take-off- and-landing vehicles, such as remote-controlled helicopters and quadrotors, to perform aerobatic maneuvers autonomously. In this work, we adopt a system-level perspective and analyze a new design for a rotary-wing micro air vehicle that utilizes gyroscopic dynamics for attitude control. Unlike traditional vehicles where attitude control moments are generated by aerodynamic control surfaces, the proposed vehicle will leverage the existing angular momentum of its counter rotating components. This paradigm has the potential to yield significant increases in agility when compared to state-of-the-art micro vertical take-off and landing vehicles. The proposed design reduces mechanical complexity by precluding the use of complex mechanisms, such as the swashplate. The capacity to rapidly generate large gyroscopic control moments, coupled with the precision gained from eliminating the need for complex and restrictive aerodynamic models, improves both agility and adaptability. We present the development of a gyroscopically controlled micro air vehicle including comprehensive models of the dynamics and the aerodynamics with an emphasis on the design and analysis of such systems. A dynamics simulator that incorporates these models and mechanical hardware solutions to challenges that arose during prototyping will also be presented.

Coaxial rotor helicopters at the Micro Air Vehicle scale are becoming widely accepted as an easy to fly and stable configuration. This study models the heave and yaw modes of the vehicles dynamics, based on the aerodynamic interaction of the two rotors. The linearized model is represented in state space with a closed form solution. It is then evaluated for its stability and control response. The coupling between the height and heading transient response is evaluated. Based on these characteristics, a technique for control mixing is proposed for the ease of pilot control and also closed loop design. The steady state values of the model are validated using typical parameters of a commercially available vehicle. The predicted thrust sharing between the rotors is found to match with that reported in experiments. The effect of the interaction between the two coaxial rotors on the system performance and stability is analyzed.

An initial design concept for a micro-coaxial rotorcraft using custom manufacturing techniques and commercial off-the-shelf components is discussed in this paper. Issues associated with the feasibility of achieving hover and fully functional flight control at small scale for a coaxial rotor configuration are addressed. Results from this initial feasibility study suggest that it is possible to develop a small scale coaxial micro rotorcraft weighing approximately 100 grams, and that available moments are appropriate for roll, yaw and lateral control. A prototype vehicle was built and its rotors were tested in a custom hover stand used to measure Thrust and power. The radio controlled vehicle was flown untethered with its own power source and exhibited good flight stability and control dynamics. The best achievable rotor performance was measured to be 42%.

Simplified analytical models for hypersonic lateral-directional stability

The desire for a vertical takeoff and landing (VTOL) aircraft capable of high forward flight speeds is very strong. Compound lift-offset coaxial helicopter designs have been proposed and have demonstrated the ability to fulfill this desire but, with high forward speeds, noise is an important concern that needs to be addressed. The study in this paper utilizes a Rotorcraft Comprehensive Analysis System (RCAS) model of the XH-59 aircraft, in conjunction with a noise prediction code, PSU-WOPWOP, to computationally explore the acoustics of a lift-offset coaxial helicopter. Specifically, unique characteristics of the XH-59 coaxial helicopter noise are identified; and design features and trim settings specific to a compound coaxial helicopter are considered for noise reduction. At some observer locations, there is constructive interference of the coaxial acoustic pressure pulses, such that the two signals add completely. The locations of these constructive interferences can be altered though, by modifying the upper-lower rotor blade phasing, providing an overall acoustic benefit. Significant noise reduction (and power reduction) is possible by reducing rotor tip speeds - an option available because the coaxial rotor in a compound configuration does not need to provide all the propulsive force. A dual-swept tip blade also enables noise reduction at an in-plane, forward, target observer. Compete CFD analyses, coupled with PSU-WOPWOP, should be explored in the future for a comprehensive acoustic evaluation of lift-offset coaxial helicopter noise.

In blue GaN‐based laser diodes with a 40 μm broad‐ridge waveguide, lateral modes of different order are selectively observed by high‐resolution spectroscopy combined with lateral near‐field scanning. Longitudinal mode clusters contribute to different lateral modes, close above the threshold current. Longitudinal mode competition is observed in each longitudinal mode cluster as intensity modulation and wavelength shift, repeating itself with a frequency of around 60–80 MHz. Despite the spectral separation of the different clusters and different lateral mode order, the mode competition processes share the same frequency and show a stable phase relation, especially at low currents. This indicates coupling of separate mode clusters via the charge carrier density. Due to the lateral modes occurring at separate wavelengths, lateral–longitudinal mode competition is found, which not only causes spectral effects but also time‐dependent periodic variations in the lateral near field.

The study deals with the prediction of the aerodynamic performance of the Gun Launched Micro Air Vehicle (GLMAV) platform by using the ANSYS CFX code. A first step consisted of the validation of the numerical simulation on a two-bladed rotor, which has a diameter of 0.20 m. This micro-rotor model is the perfect reproduction at a scale of 1:10 of the rotor bench that ISL had studied in the nineties. The validation was made by the comparison of the computed thrust and torque with the measured ones. These results proved the ability of the ANSYS CFX code to predict the flow field around MAV rotors and gave us confidence for further computations. The second step consisted of an analysis of the grid convergence for the single rotor of the GLMAV platform. Finally, by using the experience gained from all previous computational investigations, the computation of aerodynamic performance of the two coaxial 0.25-m diameter counter-rotating rotors of the GLMAV was carried out in hover flight.

The growing interest of rotary wing UAVs, for military and civilian applications, has encouraged designers to consider miniaturized configurations, more efficient in terms of endurance, payload capability and maneuverability. The purpose of this paper is to study a new configuration of coaxial rotor as applied to a micro aerial vehicle (MAV) with the intention to guarantee the vehicle maneuverability while removing unnecessary control surfaces which would increase wind gust sensitivity. Coaxial rotor configurations maximize the available rotor disk surface and allow for torque cancelation. Tilting rotors may allow for the vehicle control.