Simulation of Flight Dynamics for Helicopter Icing

Due to the distinct atmospheric conditions, there are notable differences in the rotor aerodynamic characteristics between the Mars helicopters and conventional Earth helicopters. Determining the flight dynamics characteristics of the Mars helicopter presents considerable challenges. This study focuses on a coaxial helicopter as the research subject, specifically analyzing the flapping characteristics of rotor blades in the unique Martian atmospheric environment. The unsteady aerodynamic loads and wake effects on the rotor are calculated utilizing the viscous vortex particle method. Furthermore, the aerodynamic interference of the coaxial dual-rotor system is analyzed across a range of forward flight speeds. Subsequently, a novel and efficient trimming method based on a combination of the viscous vortex particle method and dynamic inflow model for Mars helicopters is proposed to enhance the calculation accuracy and computational efficiency. Moreover, a relaxation factor is introduced to modify the induced velocities at various positions on the rotor disk. Finally, a comprehensive flight dynamics analysis is carried out for the Mars coaxial helicopter. And the results indicate that during the forward flight of the Mars helicopter, the rotors generate noticeable aerodynamic disturbances and there is evident coupling among control channels, with Dutch roll mode and velocity-attitude stability exhibiting dynamic instability as flight speed increases, posing challenges for the design of the flight control system.

In this paper, an optimal controller is investigate d based on the Linear Quadratic Control method. The controller simulation is execut ed based on a small-scale autonomous model helicopter, Yamaha R-50. Genetic algorithms are employed to generate weighting matrices by optimizing the performance index of the control system. As a result, this method produces optimal weighting matrices which further improve the Linear Quadratic Controller. A full state feedback approach is used in the control design process. A Kalman observer is integrated into the controller to predi ct full state variables, since only a limited number of state variables are measured. Finally, th e performance of the controller is evaluated in the time domain with and without disturbances when the model helicopter is in hovering flight and forward flight. I. Introduction utonomous helicopters offer special advantages for major operations in both civilian and military sect ors, particularly in cases that are considered either in accessible or dangerous for human beings, due to th eir vertical take-off and landing (VTOL) capabilities. The contr ol system of autonomous helicopters plays a key rol e in its ability to carry out the assigned task effectively. This has lead to a growing interest in the design and analysis of the control system of autonomous helicopters. In 1990s, classic control theory was widely used for autonom ous helicopter control systems, which is mainly a succe ssive loop closure approach. The control gains are selected separately using tools such as Root locus, Bode and Nyquist plots. The controller parameters are usual ly tuned based on the controller performance measured either in ti me domain or in frequency domain. This design procedure becomes increasingly difficult as more and more loo ps are added, specifically, when the dynamics invol ve multiple inputs, multiple outputs, or multiple feedback loop s with complex performance criteria. To control an autonomous helicopter modeled as a complex Multiple Input Multiple Output (MIMO) system, the performance criteria can not be easily met by the approach based on classis cont rol theory. Several researchers have developed a number of diff erent ways for the control system design of autonom ous helicopters in the literature. Shim et al 1 presented three different control methods using li near robust, fuzzy logic, and nonlinear tracking controls, for an autonomous helicopter in hover and low speed flight. Johnson a nd Kannan 2

The T-50 advanced trainer aircraft combines advanced aerodynamic features and a fly-by-wire flight control system in order to produce a stability and highly maneuverability. The flight control system both longitudinal and lateral-directional axes to achieve performance enhancements and improve stability. The T-50 employs the RSS concept in order to improve the aerodynamic performance in longitudinal axis and the longitudinal control laws employ the dynamic inversion with proportional-plus-integral control method. And, lateral-directional control laws employ the blended roll system both beta-betadot feedback and simple roll rate feedback with proportional control method in order to guarantee aircraft stability. This paper details the design process of developing lateral-directional control laws, utilizing the requirement of MIL-F-8785C and MIL-F-9490D. And, this paper propose the analysis of aircraft characteristics such as dutch-roll mode, roll mode, spiral mode, gain and phase margin about gains for lateral-directional inner loop feedback.

This paper studies the design of a simulator to simulate the general behaviour of an aircraft in its lateral motion. The study will be in two parts: the aircraft lateral natural system and the design of the control laws. In the aircraft lateral natural system, all modes will be discussed including the Pure Roll mode, the Dutch Roll mode, and the Spiral mode. In the design of the control laws, two types of control laws were discussed which are the yaw damper and the poles placement method. The result is a complete simulator for aircraft lateral motion which allows the user to simulate and study the lateral characteristic of an aircraft given the desired input. The simulator resulted from this study is suitable to be used for teaching and learning as the changes in parameters yield instantaneous results that can be displayed.

In this paper, detailed development of a nonlinear aeroelastic coupled trim model of a twin-cyclocopter in forward flight is presented. Twin-cyclocopter consists of two cycloidal rotors as main thrusters and a conventional nose rotor for pitch-torque balance. It is shown that five control inputs (mean and differential rpm, mean and differential phase offset of cyclorotors, rpm of nose rotor) are needed to balance three moments and two forces on cyclocopter in forward flight while forces along lateral direction remain balanced at all stages. In this coupled trim procedure, blade aeroelastic response equations and vehicle trim equations are solved together by simultaneously updating control inputs and blade response. To obtain the blade response and forces for a given set of control inputs, an aeroelastic model of cyclorotor and an aerodynamic model of the conventional nose rotor in forward flight is developed. The nonlinear aeroelastic model of the cyclorotor is developed by coupling unsteady aerodynamic model of cyclorotor in forward flight with a geometrically exact beam based structural framework capable of predicting large bending and torsional deflections of rotor blade. Towards this, complex aerodynamics of the cyclorotor is thoroughly investigated and various underlying phenomena, such as dynamic virtual camber, effects of near and shed wake and leading-edge vortices are rigorously modeled. A modified Double Multiple Streamtube (D-MS) model is implemented to capture the complex dynamic inflow characteristics of cyclorotor in forward flight. The present model is validated with previously published in-house experimental data on the performance of a trimmed cyclorotor at different forward speeds.

Aerodynamic characteristics of rigid coaxial rotor by wind tunnel test and numerical calculation

Accurate prediction of complex rotor-airframe integration is essential for optimum design and accurate analysis of rotorcraft. Over the years, a broad array of computational methodologies has been employed to pre dict the rotor-airframe interaction. In the present study three leading computational approaches are developed and employed in the study of rotor -airframe integration problems. These approaches include an overset grid -based, unstructured tetrahedral -based, and unstructured Cartesian grid - based methods. The computational results are compared with experimental data on a generic rotor - airframe configuration tested at Georgia Tech and on the Robin helicopter geometry tested at NASA Langley. The agreemen t with experimentally measured surface pressure is found to depend strongly on the assumed rotor load distribution . These studies are valuable in helping develop and document the capability of these approaches to provide accurate predictions of rotor -airframe interaction. Introduction The flowfield around a single rotor helicopter, whether in forward flight or hover, is difficult to model due to the presence of strong vorticity which lingers in the vicinity of the rotor. In modern helicopter designs, the rotor blades are placed very close to the fuselage for compact configurations and to reduce the drag associated with the mast. In such configurations, the vortex system generated by the rotor blades interacts with the body surface in a complex unsteady manner. Rotors and airframes designed without taking into account the rotor -airframe aerodynamic interactions tend to perform significantly below their predicted capabilities. 1 Over the years, a broad array of computational methods have been employed to predict the rotor -airframe interaction. Several reviews of these methods are available. 2,3,4 These methods include modeling of individual rotors and tip vortices and modeling the rotor as a single actuator disk.

Purpose This paper aims to provide the international aeronautical community with details of the development of a new disruptive technology for aircraft propulsion. Design/methodology/approach This paper describes the results achieved by a small Austrian aeronautical innovations company in developing a cyclogyro propulsion system capable of vertical launch and efficient forward flight. The research team progressed from concept definition and simulation (2004-2006), through experimental validation and concept demonstration (2006), component optimization (2006-2012), full system demonstration (2012-2014) and examination of ability to scale (both larger and smaller) (2015 onwards). This paper provides details of the results of each of these stages. Findings The research team proved that cyclogyro propulsion can be used for the vertical launch, and that, in forward flight, it has the potential to achieve efficiency beyond the range of conventional fixed wing and rotorcraft. Research limitations/implications This research indicates that the efficiency increases with forward speed within the range achieved in standard wind tunnels (up to 35 m/s). This efficiency appears to be caused by a unique chamber effect within the cyclogyro rotor assembly. Future research should be conducted to analyse this chamber effect in greater detail and to test the cyclogyro rotor for speeds beyond 35 m/s. Practical implications This work indicates that cyclogyro propulsion could have the potential to provide vertical launch, high speed and highly efficient aircraft that have reduced wing span, no external rotors and exceptional agility. This technology could therefore be feasible for vertical take-off and landing aircraft that can safely form densely packed swarms. Social implications It could be researched as an efficiency increase in forward flight completely different to existing propulsion systems. This could open a way for a more efficient air traffic in future and faster reduction of CO2 and NOX emission an allow an environment-friendlier air travelling. Originality/value This paper provides the details of the first cyclogyro aircraft to have flown and will serve the aeronautical community by stimulating the debate on this new disruptive technology.

Aerodynamic interference between the main and tail rotor can have a strong negative influence on the flight mechanics of a conventional helicopter. Significant unsteadiness in the tail rotor loading is encountered under certain flight conditions, but the character of the unsteadiness can depend on the direction of rotation of the tail rotor. Numerical simulations, using Brown's vorticity transport model, of the aerodynamic interaction between the main and tail rotors of a helicopter are presented for a range of forward and lateral flight trajectories. Distinct differences are predicted in the behavior of the system in left and right sideward flight that are consistent with flight experience that the greatest fluctuations in loading or control input are required in left sideways flight (for a counterclockwise rotating main rotor). These fluctuations are generally more extreme for a system with tail rotor rotating top-forward than top-aft. Differences are also exposed in the character of the lateral excitation of the system as forward flight speed is varied. The observed behavior appears to originate in the disruption of the tail rotor wake that is induced by its entrainment into the wake of the main rotor. The extent of the disruption is dependent on flight condition, and the unsteadiness of the process depends on the direction of rotation of the tail rotor. In intermediate-speed forward flight and right sideward flight, the free stream delays the entrainment of the tail rotor wake far enough downstream for the perturbations to the rotor loading to be slight. Conversely, in left sideward and quartering flight, the free stream confines the entrainment process close to the rotors, where it causes significant unsteadiness in the loads produced by the system.

Systematic design methodology for development and flight testing of a variable pitch quadrotor biplane VTOL UAV for payload delivery

Dynamic stall on an oscillatory pitching NACA 0012 airfoil was simulated via CFD under time-varying freestream conditions, which are representative of fast forward flight for a helicopter and which lead to transonic effects at the minimum of the pitch angle. The selected flow conditions were representative of a helicopter blade section at around 80% of the blade radius at the extreme of the forward flight speed, i.e. at around Mach 0.3. A sinusoidal pitch angle variation of 15° ± 10° and an out-of-phase sinusoidal freestream variation of Mach 0.48 ± 0.3 were applied at a mean reduced frequency of 0.029. The mean Reynolds number was approximately 3.3x10⁶. First, the simulation results for steady freestream dynamic stall were carefully validated with experiment. Then, the steady and unsteady freestream dynamic stall simulations were compared. Key findings include the lift coefficient hysteresis loop changing direction for the variable freestream conditions, as well as the peak values of lift, drag and negative pitching moment coefficients increasing by factors of approximately 2∼4. Results suggest that the mechanism of dynamic stall changes dramatically due to the occurrence of transonic flow at minimum angle of attack and confirms that inclusion of time-varying freestream in dynamic stall simulations is important for the extremes of the forward flight speed.

The paper presents a six-degree of freedom mathematical model of a light sport aircraft and analytically reveals aircraft dynamic response both longitudinal and lateral-directional stability. The model composes of both the kinematics equations of motion and the kinetics equations of aerodynamic forces and moments, known as the aerodynamic model equations. Simulation results show responses of the perturbed dynamic system at trim condition, which indicates dynamic stability of both short-period pitching oscillation and phugoid in longitudinal axes. The spiral, roll subsidence, and dutch roll modes in lateral-directional axes are also dynamic stable as well. These are essential to understand and evaluate the dynamic behavior, stability, safety, and other aspects of the designed aircraft through a mathematical model before conducting an operational flight test.  

The present authors proposed the “quasi-neutral dihedral-effect and directional-stability” (QNDD) airplane, which is a modified neutral dihedral-effect and directional-stability (NDD) airplane for its realization. The QNDD airplane provides a superior tolerance to crosswinds compared to a conventional one by modifying geometric properties: dihedral angle of a main wing and vertical tail volume to achieve small values of and , which results in a small effect of crosswind gust on an airplane’s attitude. While the QNDD airplane yields small and , the NDD airplane yields and equal to zero. When the changes of and due to the wing deformations are considered, the NDD airplane is ideal, while the QNDD airplane is more practical. The QNDD airplane had four real eigenvalues rather than classically two reals and one pair of complex conjugates. The eigenvalues seen in a conventional airplane as a Dutch roll mode were seen to be reals in a QNDD airplane, and the mode was overdamped, which is different from a set of eigenvalues of a conventional one. From the Bode diagrams and unsteady simulations with a discrete gust applied as a crosswind, it was confirmed that the QNDD airplane had a higher tolerance to the crosswind.

Mission of Unmanned Aerial Vehicle (UAV) “Elang Avionik” is surveillance and aerial photographs. Therefore the flight of UAV must be stable and controlable, and first step activity is dynamic modelling and stability analisys. The problems of UAV system is disturbance, noise of sensor, MIMO and uncertainty dynamic model. For good result using the multivariable robust control, with some step research that is: (1)modeling and stability analysis, (2) design and implementation of PID control system, (3) flight dynamic parameter identification, (4) design and implementation of hardware in the loop simulation, (5) design and implementation of multivariable robust control, (6) test and evaluation of system. Simulation result show that the eigen value in longitudinal is: phugoid mode = –0,061293±0,40526i and non-oscillation mode = –6,1121±4,9253. In lateral directional is:dutch roll mode = –0,91089±5,7994i, spiral mode= –0,036563, and roll subsidence mode = –12,7181. Location of poles system on the left of imaginary axis, the means that the character of system is dynamic stable. But settling time to steady state condition is very long and improved by control system design.
 Key word: State space, Longitudinal, Lateral, Stable static, Stable dynamic

The analysis of the flight of insects, birds and ornithopters is difficult because there is limited aerodynamic theory applicable for the large amplitude unsteady wing motions. Furthermore the complex wing shapes and wing motions encountered in insect and bird flight entails complex aerodynamic interactions. A systematic theoretical and experimental study has been pursued by the senior author for the past several years to gain a better understanding of the aerodynamic, dynamic and mechanical problems of flapping wing flight. The primary intent has been to establish a strong technical base necessary for the development of the ornithopter as an alternative to the helicopter. The helicopter is best suited for hovering flight and suffers greatly reduced efficiency as forward speed is increased. On the other hand, the ornithopter can achieve hovering flight, albeit with much more mechanical complexity, but promises high efficiency in forward flight. The method of attack for this study of the ornithopter has been to depart gradually from the conventional aircraft configuration so that the large body of knowledge accumulated in the development of the airplane could be applied to the study of the ornithopter. In this paper the results of the aerodynamic studies are presented because of their applicability to the modeling of insect and bird flight.