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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 this paper results arc discussed of wind‐tunnel smoke tests with a small helicopter rotor. The test apparatus includes a specially developed hot‐wire smoke‐generator. An attempt is made to describe the flow pattern in the neighbourhood of a helicopter rotor in the vortex ring state by introducing the spread of the slipstream, causing an airbody around the rotor. By considering the proportions of this airbody the unpleasant behaviour of helicopters in the partial power‐off descent at low forward speeds can be more clearly understood.

THAT the rotaplane rotor is essentially a windmill has hitherto been regarded as self‐evident. It will be shown in the following analysis that, when the profile drag is sufficiently small, only part of the rotor is in the “windmill brake” state. This part receives more torque from the air than can be expended in profile drag and therefore the rotaplane rotor at zero torque must be partly an airscrew in the “vortex ring” state.

Under this heading are published regularly abstracts of all Reports and Memoranda of the Aeronautical Research Council, Reports and Technical Memoranda of the United States National Advisory Committee for Aeronautics and publications of other similar Research Bodies as issued.

THE simple actuator disk theory, first postulated by Froude over sixty years ago, is the basis of most helicopter induced flow theory. This disk is an idealization of a rotor which uniformly accelerates the air with no loss of thrust at the blade tips. It can therefore be regarded as the limit case of a rotor with an infinite number of blades. It is also assumed to be infinitely thin so that no discontinuities in velocity occur on the two sides of the disk.

Aeroplanes PERFORMANCE OF—GENERAL 1112. On the influence of supercharging on the performance of aeroplanes. (9d).

The purpose of this paper is to examine the cross‐coupled responses of a coupled rotor‐fuselage flight dynamic simulation model, including a finite‐state inflow aerodynamics and a coupled flap‐lag and torsion flexible blade structure.

The methodology is laid out based on model development for an articulated main rotor, using the theories of aeroelastisity, finite element and finite‐state inflow formulation. The finite‐state inflow formulation is based on a 3D unsteady Euler‐based concepts presented in the time domain. The most advantages of the model are the capability of modeling dynamic wake effects, tip losses and skewed wake aerodynamics. This is, in fact, a special type of the inflow model relating inflow states, to circulatory blade loadings through a set of first‐order differential equations. A non‐iterative solution of the differential equations has practically altered the model into a simple and direct formulation appending properly to the rest of the helicopter mathematical model. A non‐linear distribution of the induced velocity over the rotor disc is finally obtained by the use of both Legendre polynomials and higher‐harmonic functions. Ultimately, validations of the theoretical results show that the on‐axis response, direct reaction to the pilot input, has a good accuracy both quantitatively and qualitatively against flight test data, and the off‐axis response, cross‐coupled or indirect reaction to the pilot input are improved by this approach of modeling.

Improvements in dynamic prediction of both trim control settings and dynamic cross‐coupled responses of helicopter to pilot inputs are observed.

Further work is required for investigation of the augmented finite state inflow model, including the wake rotation correction factors to describe helicopter maneuvering flight characteristics.

The results of this work support the future researches on design and development of advanced flight control system, incorporating a high bandwidth with low‐phase delay to control inputs and also high levels of dynamic stability within minimal controls cross coupling.

This paper provides detailed characteristics on the mathematical integration problems associated with the advanced helicopter flight dynamics research.

“Introduction to Helicopter Aerodynamics” is written in a simple and easily understandable way. The subject of helicopter performance, sometimes obscured by the complexity of the mathematical analysis, is here explained on the physical grounds and the necessary formulae are evolved using mathematics on the level of an average practical engineer.

Under this heading are published regularly abstracts of all Reports and Memoranda of the Aeronautical Research Council, Reports and Technical Memoranda of the United States National Advisory Committee for Aeronautics and publications of other similar Research Bodies as issued