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To investigate the use of centre of gravity location on reducing cyclic pitch control for helicopter UAV's (unmanned air vehicles) and MAV's (micro air vehicles). Low cyclic pitch is a necessity to implement the swashplateless rotor concept using trailing edge flaps or active twist using current generation low authority piezoceramic actuators.

An aeroelastic analysis of the helicopter rotor with elastic blades is used to perform parametric and sensitivity studies of the effects of longitudinal and lateral center of gravity (cg) movements on the main rotor cyclic pitch. An optimization approach is then used to find cg locations which reduce the cyclic pitch at a given forward speed.

It is found that the longitudinal cyclic pitch and lateral cyclic pitch can be driven to zero at a given forward speed by shifting the cg forward and to the port side, respectively. There also exist pairs of numbers for the longitudinal and lateral cg locations which drive both the cyclic pitch components to zero at a given forward speed. Based on these results, a compromise optimal cg location is obtained such that the cyclic pitch is bounded within ±5° for a BO105 helicopter rotor.

The reduction in the cyclic pitch due to helicopter cg location is found to significantly reduce the maximum magnitudes of the control angles in flight, facilitating the swashplateless rotor concept. In addition, the existence of cg locations which drive the cyclic pitches to zero allows for the use of active cg movement as a way to replace the cyclic pitch control for helicopter MAV's.

As a short take-off and landing aircraft, FanWing has the capability of being driven under power a short distance from a parking space to the take-off area. The purpose of this paper is to design the take-off control system of FanWing and study the factors that influence the short take-off performance under control.

The force analysis of FanWing is studied in the take-off phase. Two take-off control methods are researched, and several factors that influence the short take-off performance are studied under control.

The elevator and fan wing control systems are designed. Although the vehicle load increases under the fan wing control, the fan wing control is not a recommended practice in the take-off phase for its sensitivity to the pitch angle command. The additional pitch-down moment has a significant influence on the control system and the short take-off performance that the barycenter variation of FanWing should be considered carefully.

The presented efforts provide a reference for the location of the center of gravity in designing FanWing. The traditional elevator control is more recommended than the fan wing control in the take-off phase.

This paper offers a valuable reference on the control system design of FanWing. It also proves that there is an additional pith-down moment that needs to be paid close attention to. Four factors that influence the short take-off performance are compared under control.

FOR the past twenty‐five years inventors and engineers have laboured to design and perfect an airscrew in which pitch change is accomplished automatically by the action of natural forces to which any operating airscrew is subject. Millions of dollars and extensive efforts in this country and abroad have gone into this quest which produced some unusual designs in the past, but has provided aviation today with the practical realization of feasible automatic airscrews. Controllable airscrew designs featuring simple construction and operation have undergone a similar development period. Many factors have influenced this development; such as considerations of cost, mechanical refinement and the state of small aeroplane and engine performance, which in the past would not always have benefited greatly from variable pitch. Today, the advantages automatic and controllable airscrews hold for performance and desirability of the small and medium planes, which are expected to be used widely, warrant thoughtful consideration.

Aircraft pitch control is fundamental for the performance of micro aerial vehicles (MAVs). The purpose of this paper is to establish a simple experimental procedure to calibrate pitch instrumentation and classical control algorithms. This includes developing an efficient pitch angle observer with optimal estimation and evaluating controllers under uncertainty and external disturbances.

A wind tunnel test bench is designed to simulate fixed-wing aircraft dynamics. Key elements of the instrumentation commonly found in MAVs are characterized in a gyroscopic test bench. A data fusion algorithm is calibrated to match the gyroscopic test bench measurements and is then integrated into the autopilot platform. The elevator-angle to pitch-angle dynamic model is obtained experimentally. Two different control algorithms, based on model-free and model-based approaches, are designed. These controllers are analyzed in terms of parametric uncertainties due to wind speed variations and external perturbation because of sudden weight distribution changes. A series of experimental tests is performed in wind-tunnel facilities to highlight the main features of each control approach.

With regard to the instrumentation algorithms, a simple experimental methodology for the design of optimal pitch angle observer is presented and validated experimentally. In the context of the platform design and identification, the similitude among the theoretical and experimental responses shows that the platform is suitable for typical pitch control assessment. The wind tunnel experiments show that a fixed linear controller, designed using classical frequency domain concepts, is able to provide adequate responses in scenarios that approximate the operation of MAVs.

The aircraft orientation observer can be used for both pitch and roll angles. However, for simultaneousyaw angle estimation the proposed design method requires further research. The model analysis considers a wind speed range of 6-18 m/s, with a nominal operation of 12 m/s. The maximum experimentally tested reference for the pitch angle controller was 20°. Further operating conditions may require more complex control approaches (e.g. scheduling, non-linear, etc.). However, this operating range is enough for typical MAV missions.

The study shows the design of an effective pitch angle observer, based on a simple experimental approach, which achieved locally optimum estimates at the test conditions. Additionally, the instrumentation and design of a test bench for typical pitch control assessment in wind tunnel facilities is presented. Finally, the study presents the development of a simple controller that provides adequate responses in scenarios that approximate the operation of MAVs, including perturbations that resemble package delivery and parametric uncertainty due to wind speed variations.

Vertical take-off is commonly adopted in most insect-mimicking flapping-wing micro air vehicles (FMAV) while insects also adopt horizontal take-off from the ground. The purpose of this paper is to study how insects adjust their attitude in such a short time during horizontal take-off by means of designing and testing an FMAV based on stroke plane modulation.

An FMAV prototype based on stroke plane rotating modulation is built to test the flight performance during horizontal take-off. Dynamic gain and decoupling mixer is added to compensate for the nonlinearity during the rotation angle of the stroke plane getting too large at the beginning of take-off. Force/torque test based on a six-axis sensor validates the change of aerodynamic force and torque at different rotation angles. High-speed camera and motion capture system test the flight performance of horizontal take-off.

Stroke plane modulation can provide a great initial pitch toque for FMAV to realize horizontal take-off. But the large range of rotation angles causes nonlinearity and coupling of roll and yaw. A dynamic gain and a mixer are added in the controller, and the FMAV successfully achieves horizontally taking off in less than 1 s.

The research in this paper shows stroke plane modulation is suitable for insect’s horizontal take-off

Optimal switching angles are investigated for minimizing accumulated numerical errors when the dual‐Euler method is used in the simulation of angular rotation.

First, round‐off errors are theoretically modeled with a simplified mathematical representation of rotation. Round‐off errors take critical roles in the vicinity of indefinite points because they cause major numerical inaccuracy in very large numerical values represented with limited binary numbers. Optimal switching angles of (±π/4, ±3π/4) are derived and numerically examined. With a more practical and severe rotational model, the switching angles are numerically tested.

In conclusion, switching pitch angles of (±π/4, ±3π/4) yield near minimum numerical errors in angular parameters of pitch, yaw, and roll if truncation errors are not dominant by using high‐order integration algorithms and small step sizes. It is also noticed that accumulated numerical errors increase dramatically if pitch and roll angles are switched beyond the optimal angles with a little margin.

Optimal switching angles in the dual‐Euler method are identified based on the truncation error analysis. The mechanism of accumulated numerical errors in the dual‐Euler method, which depends on switching angles, is also revealed.

The purpose of this paper is to define an optimal pitching profile for the blades of a cycloidal rotor which minimizes the mean power consumption for a given mean thrust of the rotor.

A simple analytical model of the kinematics and aerodynamics of a cycloidal rotor is defined first to obtain expressions for thrust and power depending on the pitching profile and geometrical parameters of the rotor. Then, Lagrange optimization is applied to obtain the optimal pitching schedule under hovering conditions. Finally, results of the theoretical analysis are compared with those of a two-dimensional computational fluid dynamics (CFD) model.

Results of the optimization suggest that the optimal profile is a combination of sinusoidal functions. It is shown that the adoption of the optimal pitching schedule could improve the power efficiency of the rotor by approximately 25 per cent.

The possibility to increase the efficiency of a cycloidal rotor by acting on its pitching schedule could be a significant factor of success for this alternative propulsion concept.

The present work represents the first attempt at a definition of an optimal pitching profile for a cycloidal rotor. Moreover, although being carried out on the basis of simplified analytical considerations, the present investigation sets a methodological framework which could be successfully applied to the design of similar kinds of systems.

Provision must be made on all aeroplanes for the rapid egress of passengers and crew in emergency. Fulfilment of the requirements detailed below will normally be regarded as sufficient, but departures may be permitted or required in particular cases.

After briefly outlining the main features of the variable‐pitch propeller, this paper proceeds to describe the development of the piston‐engined hydraulically operated propeller as a brake, both in the air and on the ground. Examples are given of the magnitude of the braking effort of a propeller when windmilling under controlled conditions and when in reverse pitch under power. The advent of the gas turbine, originally intended as a means of jet propulsion, opened up a new field of application for the variable‐pitch propeller and this application with its attendant problems and their solution is discussed. Three types of gas‐turbine power plant, together with the appropriate propeller arrangements are reviewed. These arc: (I) the direct‐connected turbine; (2) the compound‐compressor turbine; and (3) the free‐propeller turbine.

THE present paper gives, in abbreviated form, the theory of blade motion and of static and dynamic stability of single‐rotor helicopters. Limitations of space do not permit of full discussion and the article should be taken as only an introduction to the somewhat complex problems of helicopter stability and control.