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The purpose of this paper is to present a novel optimization approach to design a robust H-infinity controller.

To use a modified particle swarm optimization (PSO) algorithm and to search for the optimal parameters of the weighting functions under the circumstance of the given structures of three weighting matrices in the H-infinity mixed sensitivity design.

This constrained multi-objective optimization is a non-convex, non-smooth problem which is solved by a modified PSO algorithm. An adaptive mutation-based PSO (AMBPSO) algorithm is proposed to improve the search accuracy and convergence of the standard PSO algorithm. In the AMBPSO algorithm, the inertia weights are modified as a function with the gradient descent and the velocities and positions of the particles.

The AMBPSO algorithm can efficiently solve such an optimization problem that a satisfactory robust H-infinity control performance can be obtained.

The purpose of this paper is to make performance improvements and timely critical execution enhancements for operational flight program (OFP). The OFP is core software of autonomous control system of small unmanned helicopter.

In order to meet the time constraints and enhance control application performance, two major improvements were done at real‐time operating system (RTOS) kernel. They are thread scheduling algorithm and lock‐free thread message communication mechanism. Both of them have a direct relationship with system efficiency and indirect relationship with helicopter control application execution stability through improved deadline keeping characteristics.

In this paper, the suitability of earliest deadline first (EDF) scheduling algorithm and non‐blocking buffer (NBB) mechanism are illustrated with experimental and practical applications. Results of this work show that EDF contributes around 15 per cent finer‐timely execution and NBB enhances kernel's responsiveness around 35 per cent with respect to the number of thread context switch and CPU utilization. These apply for OFP implemented over embedded configurable operating system (eCos) RTOS on x86 architecture‐based board.

This paper illustrates an applicability of deadline‐based real‐time scheduling algorithm and lock‐free kernel communication mechanism for performance enhancement and timely critical execution of autonomous unmanned aerial vehicle control system.

This paper illustrates a novel approach to extend RTOS kernel modules based on unmanned aerial vehicle control application execution scenario. Lock‐free thread communication mechanism is implemented, and tested for applicability at RTOS. Relationship between UAV physical and computation modules are clearly illustrated via an appropriate unified modelling language (UML) collaboration and state diagrams. As experimental tests are conducted not only for a particular application, but also for various producer/consumer scenarios, these adequately demonstrate the applicability of extended kernel modules for general use.

This paper aims to present the design of a stability augmentation system (SAS) in the longitudinal and lateral axes for an unstable helicopter.

The feedback controller is designed using linear quadratic regulator (LQR) control with full state feedback and LQR with output feedback approaches. SAS is designed to meet the handling qualities specification known as Aeronautical Design Standard (ADS‐33E‐PRF). A helicopter having a soft inplane four‐bladed hingeless main rotor and a four‐bladed tail rotor with conventional mechanical controls is used for the simulation studies. In the simulation studies, the helicopter is trimmed at hover, low speeds and forward speeds flight conditions. The performance of the helicopter SAS schemes are assessed with respect to the requirements of ADS‐33E‐PRF.

The SAS in the longitudinal axis meets the requirement of the Level 1 handling quality specifications in hover and low speed as well as for forward speed flight conditions. The SAS in the lateral axis meets the requirement of the Level 2 handling quality specifications in both hover and low speed as well as for forward speed flight conditions. The requirements of the inter axis coupling is also met and shown for the coupled dynamics case. The SAS in lateral axis may require an additional control augmentation system or adaptive control to meet the Level 1 requirements.

The study shows that the design of a SAS using LQR control algorithm with full state and output feedbacks can be used to meet ADS‐33 handling quality specifications.

The paper aims at developing a novel algorithm to estimate high-order derivatives of rotorcraft angular rates to break the contradiction between bandwidth and filtering performance because high-order derivatives of angular rates are crucial to rotorcraft control. Traditional causal estimation algorithms such as digital differential filtering or various tracking differentiators cannot balance phase-lead angle loss and high-frequency attenuation performance of the estimated differentials under the circumstance of strong vibration from the rotor system and the rather low update rate of angular rates.

The algorithm, capable of estimating angular rate derivatives to maximal second order, fuses multiple attitude signal sources through a first-proposed randomized angular motion maneuvering model independent of platform dynamics with observations generated by cascaded tracking differentiators.

The maneuvering flight test on 5-kg-level helicopter and the ferry flight test on 230-kg-level helicopter prove such algorithm is feasible to generate higher signal to noise ratio derivative estimation of angular rates than traditional differentiators in regular flight states with enough bandwidth for flight control.

The decrease of update rate of input attitude signals will weaken the bandwidth performance of the algorithm and higher sampling rate setting is recommended.

Rotorcraft flight control researchers and engineers would benefit from the estimation method when implementing flight control laws requiring angular rate derivatives.

A purely kinematic randomized angular motion model for flight vehicle is first established, combining rigid-body Euler kinematics. Such fusion algorithm with observations generated by cascaded tracking differentiators to estimate angular rate derivatives is first proposed, realized and flight tested.

This paper seeks to present a feedback error learning neuro‐controller for an unstable research helicopter.

Three neural‐aided flight controllers are designed to satisfy the ADS‐33 handling qualities specifications in pitch, roll and yaw axes. The proposed controller scheme is based on feedback error learning strategy in which the outer loop neural controller enhances the inner loop conventional controller by compensating for unknown non‐linearity and parameter uncertainties. The basic building block of the neuro‐controller is a nonlinear auto regressive exogenous (NARX) input neural network. For each neural controller, the parameter update rule is derived using Lyapunov‐like synthesis. An offline finite time training is used to provide asymptotic stability and on‐line learning strategy is employed to handle parameter uncertainty and nonlinearity.

The theoretical results are validated using simulation studies based on a nonlinear six degree‐of‐freedom helicopter undergoing an agile maneuver. The neural controller performs well in disturbance rejection is the presence of gust and sensor noise.

The neuro‐control approach presented in this paper is well suited to unmanned and small‐scale helicopters.

The study shows that the neuro‐controller meets the requirements of ADS‐33 handling qualities specifications of a helicopter.

This paper aims to present the novel methodology of computational simulation of a helicopter flight, developed especially to investigate the vortex ring state (VRS) – a dangerous phenomenon that may occur in helicopter vertical or steep descent. Therefore, the methodology has to enable modelling of fast manoeuvres of a helicopter such as the entrance in and safe escape from the VRS. The additional purpose of the paper is to discuss the results of conducted simulations of such manoeuvres.

The developed methodology joins several methods of computational fluid dynamics and flight dynamic. The approach consists of calculation of aerodynamic forces acting on rotorcraft, by solution of the unsteady Reynold-averaged Navier–Stokes (URANS) equations using the finite volume method. In parallel, the equations of motion of the helicopter and the fluid–structure-interaction equations are solved. To reduce computational costs, the flow effects caused by rotating blades are modelled using a simplified approach based on the virtual blade model.

The developed methodology of computational simulation of fast manoeuvres of a helicopter may be a valuable and reliable tool, useful when investigating the VRS. The presented results of conducted simulations of helicopter manoeuvres qualitatively comply with both the results of known experimental studies and flight tests.

The continuation of the presented research will primarily include quantitative validation of the developed methodology, with respect to well-documented flight tests of real helicopters.

The VRS is a very dangerous phenomenon that usually causes a sudden decrease of rotor thrust, an increase of the descent rate, deterioration of manoeuvrability and deficit of power. Because of this, it is difficult and risky to test the VRS during the real flight tests. Therefore, the reliable computer simulations performed using the developed methodology can significantly contribute to increase helicopter flight safety.

The paper presents the innovative and original methodology for simulating fast helicopter manoeuvres, distinguished by the original approach to flight control as well as the fact that the aerodynamic forces acting on the rotorcraft are calculated during the simulation based on the solution of URANS equations.

Despite of the numerous characteristics of the multirotor unmanned aircraft systems (UASs), they have been termed as less energy-efficient compared to fixed-wing and helicopter counterparts. The purpose of this paper is to explore a more efficient multirotor configuration and to provide the robust and stable control system for it.

A heterogeneous multirotor configuration is explored in this paper, which employs a large rotor at the centre to provide majority of lift and three small tilted booms rotors to provide the control. Design provides the combined characteristics of both quadcopters and helicopters in a single UAS configuration, providing endurance of helicopters keeping the manoeuvrability, simplicity and control of quadcopters. In this paper, rotational as well as translational dynamics of the multirotor are explored. Cascade control system is designed to provide an effective solution to control the attitude, altitude and position of the rotorcraft.

One of the challenging tasks towards successful flight of such a configuration is to design a stable and robust control system as it is an underactuated system possessing complex non-linearities and coupled dynamics. Cascaded proportional integral (PI) control approach has provided an efficient solution with stable control performance. A novel motor control loop is implemented to ensure enhanced disturbance rejection, which is also validated through Dryden turbulence model and 1-cosine gust model.

Robustness and stability of the proposed control structure for such a dynamically complex UAS configuration is demonstrated with stable attitude and position performance, reference tracking and enhanced disturbance rejection.

Helicopter rotor blades are usually aerodynamically limited by the severe conditions present in every revolution: strong shock wave boundary layer interaction on the advancing side and dynamic stall on the retreating side. Therefore, different flow control strategies might be applied to improve the aerodynamic performance.

The present research is focussed on the application of passive rod vortex generators (RVGs) to control the flow separation induced by strong shock waves on helicopter rotor blades. The formation and development in time of the streamwise vortices are also investigated for a channel flow.

The proposed RVGs are able to generate streamwise vortices as strong as the well-known air-jet vortex generators. It has been demonstrated a faster vortex formation for the rod type. Therefore, this flow control device is preferred for applications in which a quick vortex formation is required. Besides, RVGs were implemented on helicopters rotor blades improving their aerodynamic performance (ratio thrust/power consumption).

A new type of vortex generator (rod) has been investigated in several configurations (channel flow and rotor blades).

AGAINST the background briefly outlined together with the development of increasingly sophisticated powerplants as well as the research activities conducted into the operational aspects of rotating wing aircraft, it was realised that major benefits could be expected from fully exploiting health monitoring techniques. Applications would not only be for future helicopters but also for retrospective action for those currently in service. The Civil Aviation Authority (CAA) has conducted extensive investigations into many aspects of these proposals and published a report by a working group in 1985, the membership of which was drawn from industry, the Ministry of Defence (MOD) had the CAA. This report has been prepared to satisfy HARP recommendation No 11.

THE successful development of the gas turbine powerplant made available large quantities of compressed air to the aircraft designer, who realised that the aerodynamics of the aircraft could be significantly modified by its use. This is currently used in naval aircraft where blown flaps have reduced the take‐off and landing speeds. Experimental aircraft have flown with higher lift coefficients on the wings than has been realised using blown flaps. One such aircraft is the B.A.C.‐Hunting 126 jet flap aircraft which uses an internal form of the jet flap. An alternative form of the jet flap has been suggested in America in which the efflux from jet engines mounted in pods below the wing is deflected on to the lower surface of conventional trailing edge flaps which deflect the gas flow and so form an external jet flap.