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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.

IN the September, 1956, issue of AIRCRAFT ENGINEERING, Mr McClements and Sq./Ldr. Armitagc surveyed ‘Helicopter Developments during the Post‐War Years’. The purpose of the present article is to discuss the subsequent developments which have taken place during the last six years. This period has been one of continued expansion in helicopter development and of much more widespread utilization of helicopters in both military and commercial operation. Since their initial development, over 150 different types of helicopter have been successfully flown. This figure includes many experimental machines or prototypes built by small groups. The major helicopter constructors have put about 40 types into quantity production and over 10,000 helicopters have been built in the Western countries. These have been predominantly for the military services, the large majority being built in U.S.A., but something approaching 2,000 have been used in commercial operations.

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.

The purpose of this paper is to advance the multiphysics analysis of helicopter rotors under icing conditions by coupling the iced rotor’s aerodynamics, analyzed by CFD, with the rotor’s structural characteristics, analyzed by CSD.

The current work introduces supercomputer-based computational approaches capable of assessing the impact of ice accretion on the aerodynamics, blade dynamics, vibrations and loading of a rotorcraft. The rigid and elastic motions of the blades are accounted for through a loose coupling of the flow solver to a multibody dynamics solver. The coupling framework allows for comprehensive aeroelastic simulations of iced rotors in hover and in forward flight.

The flow and structural modules were validated on a full helicopter configuration in forward flight using the ROBIN experimental model. The tip structural deflections were in very close agreement with the experimental measurements.

The results of the CFD analyses are limited by the available experimental results they can be compared to. In dry air CFD, three-dimensional (3D) experiments occur first and CFD is then compared to them; in icing, the opposite is true: 3D experiments (if they are ever done, as they are very expensive) chase CFD and sometimes never occur.

This paper presents an outline of how CFD and computational stress dynamics (CSD) analyses can be linked and provides a toolbox for deeper investigation of the complex flows over helicopters operating under difficult in-flight icing conditions.

More and more helicopters are designed to be able to operate in hostile environments such as rescuing and saving lives over the oceans or mountains, conditions under which icing encounters cannot be avoided.

A loosely coupled CFD/CSD framework that accounts for the rotor blades structural response to aerodynamic loading and ice accretion in hover and forward flight has been presented. This versatile and cost-effective framework provides a more accurate estimation of the helicopter rotor performance and its degradation due to icing encounters during the early design stages than traditional CFD tools.

The purpose of this paper is to study the conceptual design and optimisation of a compound gyroplane. A study of a compound gyroplane configuration and its characteristics was performed to develop a sizing program.

The vertical takeoff and landing capabilities of a helicopter are particularly important. The need for efficient hover and the effectiveness of forward flight in the helicopter can cause conflicts within the design process. The designers usually wish to increase the helicopter's maximum forward speed. Recently, the compound aircraft is one of the concepts considered for the purpose of expanding the flight envelope of rotorcraft. The study of the compound gyroplane showed its advance capabilities for this purpose. Understanding its characteristics, a number of calculations are conducted to implement a sizing program for compound gyroplanes based on the conventional helicopter sizing process.

The results of the sizing program were validated using existing aircraft data such as the Challis Heliplane, Carter Copter, FB‐1 Gyrodyne, and Jet Gyrodyne. The program is appropriate to size a compound gyroplane at the conceptual design phase. An optimisation study was also performed to enhance sizing results. The compromise between the rotor lift sharing factor and the ratio of the wing span (Bw) to rotor diameter (D) was solved by choosing the total gross weight (TOGW) as the objective function, while the design variables are compromising factors. The optimum results showed that the TOGW of all four kinds of compound gyroplanes was considerably reduced.

A conceptual sizing program for unconventional compound aircraft was developed. The study showed that an optimum design process is necessary to enhance the sizing results.

This paper aims to improve the computational efficiency and to achieve high-order accuracy for the computation of helicopter rotor unsteady flows in forward flight during the industrial preliminary design stage.

The integral arbitrary Lagrangian–Eulerian form of unsteady compressible Navier–Stokes equations with low Mach number preconditioned pseudo time terms based on non-inertial frame of reference undergoing rotating and translating was derived and discretized in the framework of multi-block structured finite volume grid using three types of spatial reconstruction schemes, i.e. the third-order accurate monotonic upwind scheme for conservation laws, the fifth-order accurate weighted essentially non-oscillatory and the fifth-order accurate weighted compact nonlinear schemes.

The results show that the present non-inertial computational method can obtain comparable results with other methods, such as the dynamic overset method, and make sure that the higher-order spatial schemes can significantly improve the tip vortex resolution.

The computational grid used by the present method remained static during the whole unsteady computation process, with only local deformations induced by blade cyclic pitch and other operating motions, which greatly reduced the complexity of grid motion and enhanced the efficiency and robustness.

A NUMBER of approaches to the calculation of rotor downwash have already been discussed. Broadly spsaking, the methods of Castles and DeLeeuw and Squire and Mangler are the same. In both methods the downwash at the rotor disk is assumed to be perpetrated in a helical downwash sheet which, as the slipstream, extends below the rotor to infinity. The downwash in the disk due to the bound vortices, and the additional downwash in the disk which is induced by the helical sheets in the slipstream (Castles and DeLeeuw substitute downwash rings for helices, in the interest of mathematical simplicity) is calculated, on the assumption of an infinite number of lightly loaded blades. The final results of Castles and DeLeeuw on the one hand, and Squire and Mangier on the other, are in very wide disagresment. This disagreement is principally due to the fact that, whereas the first investigation assumes constant circulation along the blade (ideal twist and taper), Mangier and Squire assume a ‘practical’ variation of the form likely to be encountered on an untwisted untapered blade. We conclude that the radial distribution of lift on a helicopter blade will have a profound effect on the downwash pattern: which in turn will affect the calculated lift.

Some of the fundamental problems associated with the aerodynamic configuration of a single‐rotor remotely piloted helicopter are discussed in this paper. A method for selecting the aerodynamic shape of a fuselage and determining the locations and parameters of a horizontal stabilizer and a tail rotor in the preliminary design analysis is given. The application of this method to a remotely piloted single main rotor and tail rotor helicopter developed at the Nanjing University of Aeronautics and Astronautics is described. A raindrop shaped fuselage is chosen for the remotely piloted helicopter and the low drag characteristics of the fuselage are demonstrated from the wind‐tunnel experimental data. The experimental results on the pitching moment characteristics of the horizontal stabilizer are also presented and analyzed. The design considerations of the horizontal stabilizer and tail rotor of the helicopter are discussed in detail.

HELICOPTER performance methods currently in use fall into two main groups. The smallest is typified by Refs. and, and evaluates performance by considering the forces acting on each element of the blade. To do this it is first necessary to determine the flapping and pitch angles of the rotor, and then to integrate the elemental forces. The equations thus developed are much too complicated for design office use, and simplifications such as Ref. are achieved only at the expense of limiting the work to untwisted and untapered blades. Moreover the basic theory so far published is incomplete since it is tacitly assumed that the rotor thrust is normal to the no‐feathering axis or the tip path plane.

This paper aims to build an accurate mathematical model which is necessary for control design and attitude estimation of a miniature unmanned rotorcraft and its subsequent conversion to an autonomous vehicle.

Frequency-domain system identification of a small-size flybar-less remote controlled helicopter is carried out based on the input–output data collected from flight tests of the instrumented vehicle. A complete six degrees of freedom quasi-steady dynamic model is derived for hover and cruise flight conditions.

The veracity of the developed model is ascertained by comparing the predicted model responses to the actual responses from flight experiments and from statistical measures. Dynamic stability analysis of the vehicle is carried out using eigenvalues and eigenvectors. The identified model represents the vehicle dynamics very well in the frequency range of interest.

The model needs to be augmented with additional terms to represent the high-frequency dynamics of the vehicle.

Control algorithms developed using the first principles model can be easily reconfigured using the identified model, because the model structure is not altered during identification.

This paper gives a practical solution for model identification and stability analysis of a small-scale flybar-less helicopter. The estimated model can be easily used in developing control algorithms.