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Explains the origins of rotor aerodynamic limits for helicopters including retreating and advancing blade limits. Examines the compounding of a helicopter for higher forward speed and reports the conclusions of a student project to design a rotorcraft capable of 300 knots and carrying a payload of 30 passengers.

ACCORDING to historical records the earliest known drawings for an aerial machine that can be classified under the heading of helicopter were made in the fifteenth century by the world renowned Italian scientist and artist Leonardo da Vinci (1452–1519). Probably the Chinese had been making their helicopter toy for some considerable time before da Vinci commenced his experiments. This toy consisted of two feathers, joined together by means of a cork or soft wood boss, to form a crude type of propeller which was pushed up a threaded stick so that upon leaving the stick the propeller rotated at high speed and continued to screw itself up in the air. When the speed of rotation decreased the propeller slowly windmilled down to the ground. A similar toy is still being sold today.

The purpose of this paper is to determine dependencies between a rotor-blade shape and a rotor performance as well as to search for optimal shapes of blades dedicated for helicopter main and tail rotors.

The research is conducted based on computational methodology, using the parametric-design approach. The developed parametric model takes into account several typical blade-shape parameters. The rotor aerodynamic characteristics are evaluated using the unsteady Reynolds-averaged Navier–Stokes solver. Flow effects caused by rotating blades are modelled based on both simplified approach and truly 3D simulations.

The computational studies have shown that the helicopter-rotor performance may be significantly improved even through relatively simple aerodynamic redesigning of its blades. The research results confirm high potential of the developed methodology of rotor-blade optimisation. Developed families of helicopter-rotor-blade airfoils are competitive compared to the best airfoils cited in literature. The finally designed rotors, compared to the baselines, for the same driving power, are characterised by 5 and 32% higher thrust, in case of main and tail rotor, respectively.

The developed and implemented methodology of parametric design and optimisation of helicopter-rotor blades may be used in future studies on performance improvement of rotorcraft rotors. Some of presented results concern the redesigning of main and tail rotors of existing helicopters. These results may be used directly in modernisation processes of these helicopters.

The presented study is original in relation to the developed methodology of optimisation of helicopter-rotor blades, families of modern helicopter airfoils and innovative solutions in rotor-blade-design area.

A theory is developed to describe the dynamic behaviour, control angles to trim and stability derivatives of an aerodynamic servo‐controlled rotor. The analysis is restricted to constant chord rotor blades which are torsionally deformed by the servo flap to give changes in rotor pitch angle, as this is the form in which the system is most likely to be used. Also, the aerodynamic centre and elemental C.G. lines are assumed to coincide with the blade torsion axis. Since the stiff hinge assumption is used, the analysis is applicable to blades with offset or ‘stiff’ flapping hinges, or to cantilever rotors, which can be simulated by a rigid blade with a stiff hinge. Comparison with some N.A.C.A. test tower results shows that the theory developed gives excellent agreement with the available experimental results. In Appendix I the stability of the tip path plane is examined, using the equations derived in the report, and three regions of instability are shown to be present. A practical rotor must be designed to operate below the lowest instability region, as is the case for the N.A.C.A. test rotor. Equations for ∂a1s/∂µ and ∂a1s/∂q are developed for the low speeds near hovering in Appendix II. Other derivatives can be easily derived from the general equations of motion given in Table 1.

This paper aims to compare the comprehensive rotorcraft analyses using the two different blade section property data sets for the blade natural frequencies, airloads, elastic deformations, the trimmed rotor pitch control angles and the blade structural loads of a small-scale model rotor in a blade vortex interaction (BVI) phenomenon.

The two different blade section property data sets for the first Higher-harmonic control Aeroacoustic Rotor Test (HART-I) are considered for the present rotor aeromechanics analyses. One is the blade property data set using the predicted values which is one of the estimated data sets used for the previous validation works. The other data set uses the measured values for an uninstrumented blade. A comprehensive rotorcraft analysis code, CAMRAD II (comprehensive analytical model of rotorcraft aerodynamics and dynamics II), is used to predict the rotor aeromechanics such as the blade natural frequencies, airloads, elastic deformations, the trimmed rotor pitch control angles and the blade structural loads for the three test cases with and without higher-harmonic control pitch inputs. In CAMRAD II modelling with the two different blade property data sets, the blade is represented as a geometrically nonlinear elastic beam, and the multiple-trailer wake with consolidation model is used to consider more elaborately the BVI effect in low-speed descending flight. The aeromechanics analysis result sets using the two different blade section property data sets are compared with each other as well as are correlated with the wind-tunnel test data.

The predicted blade natural frequencies using the two different blade section property data sets at non-rotating condition are quite similar to each other except for the natural frequency in the fourth flap mode. However, the natural frequencies using the predicted blade properties at nominal rotating condition are lower than those with the measured blade properties except for the second lead-lag frequency. The trimmed collective pitch control angle with the predicted blade properties is higher than both the wind-tunnel test data and the result using the measured blade properties in all the three test cases. The two different blade property data sets both give reasonable predictions on the blade section normal forces with BVI in the three test cases, and the two analysis results are reasonably similar to each other. The blade elastic deformations at the tip using the measured blade properties are correlated more closely with the wind-tunnel test data than those using the predicted blade properties in most correlation examples. In addition, the predictions of blade structural loads can be slightly or moderately improved by using the measured blade properties particularly for the oscillatory flap bending moments. Finally, the movement of the sectional centre of gravity location of the uninstrumented blade has a moderate influence on the blade elastic twist at the tip in the baseline case and the oscillatory flap bending moment in the minimum noise case.

The present comparison study on rotor aeromechanics analyses using the two different blade property data sets will show the influence of blade section properties on rotor aeromechanics analysis.

This paper is the first attempt to compare the aeromechanics analysis results using the two different blade section property data sets for all three test cases (baseline, minimum noise and minimum vibration) of HART-I in low-speed descending flight.

The purpose of this paper is to present a simulation method applied for investigation of helicopter ground resonance phenomenon.

The considered physical model of helicopter standing on ground with rotating rotor consists of fuselage and main transmission gear treated as stiff bodies connected by elastic elements. The fuselage is supported on landing gear modeled by spring-damper units. The main rotor blades are treated as set of elastic axes with lumped masses distributed along blade radius. Due to Galerkin method, parameters of blades motion are assumed as a combination of bending and torsion eigen modes. A Runge–Kutta method is applied to solve equations of motions of rotor blades and helicopter fuselage.

The presented simulation method may be applied in preliminary stage of helicopter design to avoid ground resonance by proper selection of landing gear units and blade damper characteristics.

Ground resonance may occur in form of violently increasing mutual oscillations of helicopter fuselage and lead-lag motion of rotor blades. According to changes of stiffness and damping characteristics, simulations show stable behavior or arising oscillations of helicopter. The effects of different blade balance or defect of blade damper are predicted.

The simulation method may help to determine the envelope of safe operation of helicopter in phase of take-off or landing. The effects of additional disturbances as results of blades pitch control as swashplate deflection are introduced.

In‐plane vibration of a balanced helicopter rotor is caused by variations with azimuth of the in‐plane forces acting on individual blades. These forces may be summarized under three headings: ‘Induced forces’ caused by the inclination of elemental lift vectors relative to the axis of rotation. ‘Profile drag forces’: variations are caused by changes with azimuth angle of the angle and airspeed of the individual blade elements. ‘Coriolis forces’, which are caused by blade flapping, which brings about a variation of blade moment of inertia about the axis of rotation. Equations are developed in this paper for the resultant hub force due to each of these forces, on the assumptions of small flapping hinge offset. It is assumed that blades are linearly twisted and tapered, an assumption which in practice can be applied to any normal rotor. It is shown that by suitably inclining the mechanical axis it is possible to balance out the worst induced and profile drag vibrations by the coriolis one, which can be made to have opposite sign. If the mechanical axis is fixed in the fuselage, this suppression is fully effective for one flight condition only. In multi‐rotor helicopters, vibration suppression can be extended over a much wider range by varying the fuselage attitude. The logical result of this analysis is, for single rotor helicopters, a floating mechanical axis which can be adjusted or trimmed by the pilot. This would be quite simple to do on a tip‐driven rotor, and has already been achieved with a mechanical drive on the Doman helicopter. The more important causes of vibration from an unbalanced rotor are next con‐sidered, attention here being confined principally to fully articulated rotors, which are the most difficult to balance because the drag hinges tend to magnify all in‐accuracies in finish and balance. From a brief discussion of the vertical vibration of an imperfect rotor it is shown that some contemporary methods of ‘tracking’ are fundamentally wrong. Finally the vibration due to tip‐mounted power units is described. In discussing the effect of a vibratory force on a helicopter a simple response chart is developed, and it is thought that its use could well be accepted as a simple standard for general assessment purposes. In the development of equations for vibration the following points of general technical interest are put forward: An equation for induced torque is developed which includes a number of hitherto neglected parameters. A new form of equation for mean lift coefficient of a blade is suggested. The simple Hafner criterion for flight envelopes is shown to give rise to considerable error, and the use of Eq. (28) is suggested in its place. The variation of profile torque with forward speed is given, and the increase due to ? varying round the disk is expressed as an explicit equation, thus allowing considerable improvement in the present methods of allowing for this effect.

THE continual development of helicopter rotor systems has so far resulted in the use of about six main types, and it will be of value briefly to recapitulate their advantages and disadvantages in order to obtain a balanced picture against which the stiff‐hinged rotor can be judged.

THIS paper presents a summary of the method and results of a general investigation into the performance characteristics of ‘single’ autogyro and helicopter rotors, which was a preliminary to the establishment, by the firm the author serves, of a helicopter division.

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.