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The purpose of his paper is to study the radar stealth performance of a Y-type quadrotor with coaxial rotors and parallel rotors.

This Y-type quadrotor is designed as an aerodynamic layout with parallel twin rotors at the front and coaxial twin rotors at the rear. The multi-rotor scattering (MRS) method based on multi-rotor dynamic simulation (MRDS) and electromagnetic scattering module (ESM) is presented. MRDS is used to simulate the complex rotation of parallel rotors and coaxial rotors. ESM is used to calculate the instantaneous radar cross-section (RCS) of the quadrotor.

For a single rotor, the minimum period of the RCS curve at a given azimuth is equal to the basic passage time of the blade, where increasing the speed can shorten this minimum period. When the elevation angle increases, the forward RCS fluctuation of the quadrotor increases, while the average RCS decreases. The change of the roll angle will affect both the mean and the maximum difference of the RCS–time curve at the given lateral azimuth. The increase of the pitch angle will enhance the dynamic amplitude of the RCS–time curve under the forward azimuth.

The research in this article can provide reference for the stealth design of the Y-type quadcopter in the future.

The originality is the establishment of the MRS method. This method could provide value for dealing with the electromagnetic scattering problem of coaxial rotors and parallel rotors.

This paper aims to investigate the mechanisms lying behind the cycloidal rotor under hovering status.

Experiments were conducted to validate the numerical simulation results. The simulations were based on unsteady Reynolds-averaged Navier–Stokes (URANS) equations solver and the sliding mesh technique was used to model the blade motion. 2D and 2.5D simulations were made to investigate the 3D effects of turbulence. The effects of pressure and viscosity were compared to study the significance of the blade motion on force generation.

The 2.5D numerical simulation cannot produce more accurate results than the 2D counterpart. The pitching motion of the blade results in dynamic stall. The dynamic stall vortices induce parallel blade vortex interaction (BVI) upon downstream blades. The interactions between the blades delay the stall of the blade which is beneficial to the thrust generation. The blade pitching motion is the dominant contributor to the force generation and the turbulence is the secondary. Strong downwash in the rotor cage varied the inflow velocity as well as the effective angle of attack (AOA) of the blade.

Cycloidal rotor is a propulsion device that can provide omni-directional vectored thrust with high efficiency and low noise. To understand the mechanisms lying behind the cycloidal rotor helps the authors to design efficient cycloidal rotors for aircraft.

The authors discovered that the blade pitching motion plays primary role in force generation. The effects of the dynamic stall and BVI were studied. The reason why cycloidal rotor can be more efficient was discussed.

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.

IN the theory of the rotaplane developed by Glauert (ref. 20) certain terms are neglected which become important at small angles of incidence of the rotor, i.e., at high forward speed. Lock (ref. 16) has removed some of the approximations in Glauert's theory and obtains the same expression for the profile drag as is given in Appendix I of reference 20, which Glauert derived by consideration of the energy account. This expression, viz.,

SECTION (B). GEOMETRIC ASPECTS OF TRIMMED FLIGHT The Relation Between the Various Planes of Reference The preceding analytical investigation of the trimmed conditions has been based upon the use of the particular set of axes advocated in Part I. There we took as the dominant axis the axis of the rotor shaft, and regarded as the dominant plane, plane RS, the plane perpendicular to the rotor shaft. We recall that the choice of this plane as a plane of reference was dictated by the importance attached at the outset to the study of the dynamics of the blade motion. We also pointed out that, from the mechani‐cal point of view, plane SP, the plane of the swash‐plate for in the absence of a swash‐plate, the equivalent ‘plane of no‐feathering’), could alterna‐tively be regarded as the dominant reference plane. But now we sec that if we concentrate on the aerodynamic aspect, the prevalence of the ex‐pressions (?i—?0+?0?1) and (?1—?) ascribe prime importance to the axis of the flapping cone. Now just as in the case of the rotor shaft, we chose to specify its direction by a plane perpendicular to it, so here we shall specify the axis of the flapping cone by any plane perpendicular to it. One such plane is the tip‐path plane: another is the plane through the rotor hub parallel to the tip‐path plane. We shall refer to cither plane as the plane TP, and so long as we are only concerned with directions, no ambiguity will be caused. We now have three planes competing for attention, planes RS, SP and TP, and it is our object to link these planes together and see what roles they play in the description of the motion of the helicopter as a whole, having special regard in all this to the particular position of the helicopter's C.G. It may be added that the relationship between these planes is a prominent feature of certain papers [Refs. (2), (10)], and the phrase ‘the equivalence of flapping and feathering’ is there used to describe the resultant findings. It will be shown that there are two different interpretations of this phrase, and that in one sense, what is implied is no more than the equivalence of alternative geometric descriptions of one and the same motion of a helicopter rotor. On the other hand, the phrase may be taken to mean that a rotor having no flapping hinge at all [such as the one analysed by Squire in Ref. (3)] performs effectively in the same way as one free to flap. Indeed, the results of Ref. (3) are frequently used as the basis of many trim and stability calculations of helicopters with freely flapping rotors, despite the lack of a flapping hinge in the calculations of Ref. (3). We shall see that, in general, this procedure is questionable.

Discusses the 27 papers in ISEF 1999 Proceedings on the subject of electromagnetisms. States the groups of papers cover such subjects within the discipline as: induction machines; reluctance motors; PM motors; transformers and reactors; and special problems and applications. Debates all of these in great detail and itemizes each with greater in‐depth discussion of the various technical applications and areas. Concludes that the recommendations made should be adhered to.

This paper aims to present the results of aerodynamic calculation of the aircraft in tandem wing configuration called VTOL. A presented vehicle combines the capabilities of the classic aircraft and helicopters. The aircraft is equipped with two pairs of tilt-rotors mounted on the tips of the front and the rear wing. The main goal of the presented research was to find the aerodynamic impact of both pairs of tilt-rotors on aerodynamic coefficients of the aircraft. Moreover, the rotors impact on the static stability of the aircraft was investigated too.

The CFD analysis was made for the complete aircraft in the tandem wing configuration. The computation was performed for the model of aircraft which was equipped with the four sub-models of the front and rear rotors. They were modeled as the actuator discs. This method allows for computing the aerodynamic impact of rotating components on the aircraft body. All aerodynamic analysis was made by the MGAERO software. The numerical code of the software was based on the Euler flow model. The used numerical method allows for the quick computation of very complex model of aircraft with a satisfied accuracy.

The result obtained by computation includes the aerodynamic coefficients which described the impact of the tilt rotors on the aircraft aerodynamic. The influence of the angle of attack, sideslip angle and the change of rotor tilt angle was investigated. Evaluation of the influence was made by the stability margin analysis and the selected stability derivatives computation.

Presented results could be very useful in the computation of dynamic stability of unconventional aircraft. Moreover, results could be helpful during designing the aircraft in the tandem wing configuration.

This paper presents the aerodynamic analysis of the unconventional configuration of the aircraft which combines the tandem wing feature with the tilt-rotor advantages. The impact of disturbance generated by the front and rear rotors on the flow around the aircraft was investigated. Moreover, the impact of rotors configuration on the aircraft static stability was found too.

The basis and the extent of theoretical induced velocity calculations are reviewed. The theory is then applied to interference problems involving additional rotors, wings, and tails in the flow. For such cases, the interference effects can be calculated with acceptable accuracy. For hovering, superposition is used to introduce ground effect into the calculations. The resulting flow field offers a qualitative explanation of several previously observed phenomena. It is shown that, because of assumptions inherent in the analysis, the present induced‐flow theory cannot be used to predict the detailed aerodynamic loading on the rotor blades.

The component of velocity, μ0ΩR, along the x‐axis is specified, and also X, the angle of inclination of the flight path. The concept of the ‘internal’ and ‘external’ parameters of the motion is introduced, the ‘internal’ parameters consisting of certain combinations of the control and flapping angles θ0, θ1, θ2, β0, β1, β2 and the ‘external’ parameters consisting of μ0, Λ0, Θ, X, ∈ and z*, where Θ is the inclination of the x‐axis to the horizontal, and e the inclination to the vertical of the rotor resultant force‐vector. For a given centre of gravity position of the helicopter the ‘external’ parameters are shown to be determinable independently of the ‘internal’ parameters, subject to certain assumptions regarding the fuselage drag coefficient. The fundamental ‘internal’ parameter then emerges as (β1—Θ), which physically is the fore or aft inclination to the vertical of the rotor cone axis. The value of this parameter is found as the root of a quadratic equation, and not a linear equation as hitherto. It is shown that at high values of μ0 the rotor force‐vector docs not coincide with the rotor cone axis (as it is commonly supposed to do), the difference between (β1—Θ) and ∈ amounting possibly to 5 deg., of which no more than 1½ deg. can be accounted for by the rotor mean drag force component. Particular attention is paid to the case of horizontal flight, in which the true speed is μ0ΩR see Θ, the factor sec Θ being important. It is shown that at a given true speed the following quantities are independent of the centre of gravity position of the helicopter: (i) ∈, (ii) the quantity (λ1—Λ0+λ0β1), i.e. the component of the air velocity along the rotor cone axis, (iii) the horse‐power of the engine. In addition the following ‘internal’ parameters, (iv) (β1—Θ), (v) (θ2—β1), (vi) (θ1—β2), are very nearly independent of C.G. position, their variations with C.G. position only just becoming noticeable at large values of μ0. This feature enables a much simplified solution to the trim problem to be obtained by working in terms of a fictitious C.G. position chosen to make Θ zero. The genuinely C.G.‐sensitive parameters θ1, θ2, β1, β2 and Θ are subsequently converted to their true values corresponding to the actual C.G. position. The effect of variation in blade moment of inertia is examined through variations in the associated parameter γ, and it is shown that both the ‘external’ and ‘internal’ conditions are substantially unaltered, save for (θ1+ β2) and the mean coning angle β0, which do vary considerably with γ. The analysis is accurate at least so far as μ05 and the results are shown to be consistent with an energy equation. The only sources of error in the dynamics of the problem lie in the rejection of higher powers of μ0 and of flapping harmonics of β beyond the first. In order to assess the effects of this rejection, an analysis is made of a rigid rotor system free only to feather without flapping, and an exact dynamical solution for horizontal flight is obtained, subject to the same aerodynamical assumptions as before. At a given value of μ0 see Θ the results differ slightly from the case of the flapping rotor because now the purely feathering rotor applies a pitching couple mechanically to the fuselage via the shaft, which the flapping rotor is not able to do. However, by calculating the value of this couple, and supposing that the helicopter with the flapping rotor experiences either an aerodynamic fuselage pitching moment of the same amount or alternatively an appropriate C.G. shift, the two cases are reduced to essentially equivalent physical systems, particularly if the mean coning angle, β0, is deliberately arranged to be zero by proper choice of γ. A recalculation of the flapping case is then made; if its theory were exact, its solution would coincide with the known exact solution of the purely feathering rotor case. The discrepancies are shown to be very small, and thus the validity of rejecting μ06 is established for the earlier flapping rotor analysis. Part III is concerned with the theoretical solution of the feathering‐cumflapping rotor case, and also with a geometrical account of the trim configuration. Part IV contains the solution for the purely feathering rotor together with a discussion of the large number of numerical results and related tables, obtained on a Pegasus Computer.

The purpose of this paper is to propose an analytical model for the calculation of forces ripples in bearingless permanent magnet motor.

The model is based on a spectral modeling of the forces based on a spectral analysis of the air-gap flux density.

The proposed methodology allows the prediction of the ripples of the levitation according the design of the machine and the topology of the winding. It can be used in the design procedure of bearingless motor in order to avoid mechanical resonance of the rotor shaft.

This model cannot be used for high saturated machines.

The main originality of this paper consists in the evaluation of the field harmonics according to their sources and the contribution of the different harmonics to the force ripples. This can help the designer to focus on reducing or eliminating only the harmonics of flux density which create the force ripples.