WO2004002825A1

WO2004002825A1 - Helicopter

Info

Publication number: WO2004002825A1
Application number: PCT/NL2003/000481
Authority: WO
Inventors: Theodoor Van Holten
Original assignee: Technische Universiteit Delft
Priority date: 2002-06-28
Filing date: 2003-06-30
Publication date: 2004-01-08
Also published as: EP1523440A1; AU2003280204A1; NL1020963C2

Abstract

The invention relates to a helicopter provided with a rotor with at least one rotor blade (5), wherein drive means are provided for actively moving the or each rotor blade up and down during rotation of the rotor, in particular about a flapping hinge (8) thereof, so that moments, applied by the rotor to the fuselage (1) are prevented.

Description

Title: Helicopter

The invention relates to a helicopter.

Helicopters are known from practice. As a rule, helicopters have a main rotor which is mounted above the fuselage, with which main rotor lifting and controlling forces can be obtained for flying the helicopter. As a result of the driving moment applied to the rotor by an engine built-in in the fuselage, a reaction torque occurs wanting to rotate the fuselage against the direction of rotation of the rotor. With conventional helicopters, this is prevented by a tail rotor provided on the tail which rotates substantially in a vertical plane. In this way, the yaw of the fuselage, i.e. rotation about an axis parallel to the axis of the main rotor, can be controlled. With such helicopters, the tail rotor therefore also enables the control.

These known, conventional helicopters have as a drawback that the tail rotor is necessary for being able to fly. This tail rotor is expensive and vulnerable and, moreover, requires relatively much power which cannot be used for the drive of the helicopter. Generally, tail rotors demand at least 5 to 10% of the available power. Furthermore, tail rotors are dangerous to the surroundings, particularly when the helicopter is located on or just above the ground.

It has been proposed to equip a helicopter with two contra-rotating main rotors, which can be mounted next to or behind each other, on two axes, or co-axially, one above the other. However, such embodiments are particularly complex and expensive.

Moreover, it has already been proposed to replace the tail rotor by a gas outlet in the tail through which gas can be blown out sideways under pressure, thus preventing the yaw, at least modulating it. However, here, it also holds that thus, a large part of the engine power is lost. Further, for this solution too it holds that it is expensive and relatively complex and, hence, vulnerable. Further, it has been proposed to drive the main rotor by tip drive. Here, gas is blown out under pressure from gas outlets provided near the free ends of the rotor blades of the main rotor, for obtaining rotation of the main rotor. Here, it also holds that this construction is complex, expensive and vulnerable and, furthermore, that it costs much energy.

The object of the invention is a helicopter with which the drawbacks mentioned of the known helicopters are at least partly avoided. In particular, the invention contemplates providing a helicopter with which, during use, in a relatively simple manner, torques occurring as a consequence of the engine are compensated. To that end, a helicopter according to the invention is characterized by the features of claim 1.

With a helicopter according to the invention, at least a part of the rotation of the main rotor, during flight of the helicopter, is obtained by an up and down movement of the or each rotor blade. Due to the linear drive of the rotor blades, the reaction torque of the main rotor is substantially removed and partly compensated by an opposing torque induced in the means mentioned. As a result, a main rotor is obtained which applies virtually no reaction torque on the fuselage, so that a tail rotor is no longer required. Preferably, virtually the entire rotation of the main rotor is obtained by the up and down movement of the rotor blades.

Herein, up and down movement of a rotor blade should at least be understood to include a reciprocal pivoting movement of a rotor blade around a pivot means such as a flapping hinge with which the rotor blade is connected to the axis of the main rotor. Preferably, this flapping hinge is a flapping hinge which, with a conventional rotor, offers some freedom to the respective rotor blade.

In principle, any number of rotor blades suffices, but it is preferred that an even number of blades is used, in particular an even number of blades situated at a regular mutual angle relative to each other. Particularly advantageous is the use of two rotor blades, arranged opposite each other, or two pairs of rotor blades, with the blades of a first pair moving opposite to the rotor blades of the other pair: 2 x anti-symmetric. By using an even number of rotor blades, a symmetrical load is obtained so that a relatively simple construction suffices. Moreover, with an even number of blades, particularly also with two blades, the advantage is achieved that rolling and pitching moments on the fuselage are avoided. The use of two pairs of oppositely up and down moving blades offers the advantage that vibrations, in particular vertical vibrations, can be avoided, at least for the larger part. Also, a so-called (2x) teeter rotor can be used, one or two sets of blades, with the rotor blades of each set fixedly connected to each other. They can pivot like a seesaw on a center axis. Here, the flap forcing is generated such that each time, for one set, the backwardly directed blade has achieved the maximum upward movement, while this is just the opposite for the other set. Its forwardly directed blade has the maximum upward speed. Such a 2x teeter rotor can offer the advantage that the "teeter" pivots are virtually not influenced by the occurring centrifugal loads.

The up and down movement, further also to be called flapping movement, can be obtained by a mechanical drive, further also to be called flap forcing. Preferably, the flapping movement for each pair of rotor blades is effected in accordance with the rate of revolution of the rotor blades, in particular one complete flap forcing (one complete up and downward movement) per revolution of the respective pair of rotor blades, such that the flap forcing effects the flapping movement close to the eigenfrequency, so that occurring moments in the rotor blades adjacent the rotor shaft, at least adjacent and in the coupling of the rqtpr blades to the rotor shaft, are limited to a minimum.

With a helicopter according to the invention, the control means can be designed in a conventional manner, it being understood that the operating means for operating the conventional tail rotor have been replaced with comparable operating means for operating the up and down movement. Conventional swash plate control can be used, operating in an asymmetric manner. With this, it can be effected, when the drive of the flapping movements of the pairs of rotor blades is symmetrical, that the amplitude of the flap forcing, required to eliminate the reaction torque on the fuselage, is not influenced by the further control activities. Through adjustment of the flap forcing, a yaw of the fuselage can be obtained. The control can then simply be designed such that no mutual influence occurs between the yaw, roll and pitch movements.

In a particularly advantageous embodiment, the mechanism for the flap forcing is designed such that the rotor blades are coupled therewith by at least resilient means, preferably relatively slack or flexible resilient means, such that the natural flap frequency closely approximates, and preferably corresponds to the rate of revolution of the rotor blades. This is at least understood to mean a movement such that once per revolution, each blade once reaches a highest position and once a lowest position. As a result, undesired forces, tensions, vibrations and such are even further prevented. The relatively slack resilient means then offer the advantage that a normally occurring "flapping movement" resulting from the normal control forms no hindrance. Herein, relatively slack is at least understood to include a spring system meeting, substantially, the following equation:

Ke² < 0,21Ω²/ ,

K = spring constant (N/m) e = distance from the or each spring to the axis of the rotor shaft

Ω = angular speed

I = moment of inertia around the flapping hinge. It has been found that a spring system per blade which meets this formula remains within approximately 10% deviation from the eigenfrequency. A matchin spring constant K can therefore be indicated as slack.

With a helicopter according to the invention, preferably, at least the flap forcing is controlled such that the maximum moment for the flap forcing of the or a pair of rotor blades occurs when the respective rotor blades extend parallel to the longitudinal axis of the helicopter. As a result, the inclined planes of the rotor blades will be equal but in opposite sense, so that virtually no transverse or longitudinal forces will act on the helicopter as a result of the flap forcing.

It is preferred that the mechanism for the flap forcing is designed such that from standstill, the rotor blades can be put into motion by the rotor shaft, while, preferably gradually, continuously variable and without mechanical couplings, the transition to the flap forcing can be made, so that undesired shocks and such can be prevented while no complicated and relatively expensive starter means are necessary. It is then preferred that between the engine and the rotor shaft, torsion means are included, at least a spring/damper system for insulating the engine and swash plate, in particular relatively slack or flexible torsion means, from each other against vibration, at least partially, so that undesired vibrations resulting from torsion forces in the transmission can be reduced or even be prevented.

In clarification of the invention, exemplary embodiments of a helicopter according to the invention will be further elucidated with reference to the drawing. In the drawing: Fig. 1 shows, in top plan view, a conventional helicopter with forces occurring thereon;

Fig. 2 shows, in side view, a helicopter according to Fig. 1; Fig. 3 schematically shows, in perspective view, a conventional control for a helicopter; 0481

Fig. 4 schematically shows, in side view, a mechanism for flap forcing according to the invention;

Fig. 5 schematically shows a top plan view of a shaft plane of a helicopter with one blade; Fig. 6 shows a view of a rotor blade according to the arrow A in

Fig. 5;

Fig. 7 shows the flow angles on a blade section of a rotor blade; Fig. 8 shows the forces on a rotor blade which contribute to a flapping movement; Fig. 9 schematically shows an alternative embodiment of a mechanism for a flap forcing according to the invention; and

Fig. 10 schematically shows a side view of a helicopter according to the invention without tail.

In this description, identical or corresponding parts have identical or corresponding reference numerals. In this description, flapping movement is at least understood to include an up and down movement of at least one rotor blade by pivoting in a flapping hinge. Flap forcing is at least understood to include inducing the flapping movement in a forced manner, i.e. a controlled manner. Flap mechanism is at least understood to include a device for the flap forcing. Herein, shaft plane AN is at least understood to include a normal plane of the rotor shaft RA of a helicopter.

Figs. 1 and 2 schematically show, in top plan view and side view, a helicopter according to the state of the art, provided with a fuselage 1 with a tail 2, extending backwards. In the fuselage 1, an engine 19 is provided, driving a rotor shaft 4, to which rotor blades 5 are attached, one of which is shown in Fig. 1. By driving the engine 19, the rotor blades 5 are driven in the direction B, so that the rotor blades 5 are given an angular velocity Ω. The engine 19 applies a driving moment Q to the rotor blades 5. In reaction thereto, a reaction torque occurs on the fuselage 1, in particular on the engine attachment, represented in Fig. 1 by the arrow RT. As a result of this reaction torque RT, the fuselage 1 will tend to yaw around the rotor shaft 4, at least around a vertical axis, which, in the conventional helicopter, is counteracted by a tail rotor 6 provided at the end of the tail 2. With the aid of pedals 7, the helicopter pilot can regulate the blade angle of the tail rotor 6 such that a greater or smaller transverse force on the tail is obtained, so that the reaction torque RT is compensated. Furthermore, with this, in a simple manner, the desired yaw of the fuselage 1 can be obtained, by slightly increasing or, conversely, decreasing the driving moment of the tail rotor.

As schematically represented in Fig. 3, the rotor blades 5 are connected to the rotor shaft 4 via an assembly of hinges, with which complicated movements of the rotor blades can be effected. This assembly of hinges comprises a flapping hinge 8, a swing hinge 9 and a torsion hinge 10. Around the rotor shaft 4, a swash plate 11 is placed which can be moved for controlling the movements of the rotor blades 5 with the aid of connecting rods 12 connected to the control lever of the pilot. The swash plate is provided through cardanic suspension, so that it can be tilted in all directions, while coupling rods 13 are provided connecting the swash plate to a pivot arm 14 coupled to the torsion hinge 10. Such a control is sufficiently known per se and will not be further described here.

As described in the introduction, the use of tail rotors has as a drawback that they are relatively vulnerable, cause noise pollution and are energetically disadvantageous. Moreover, the maneuverability of a helicopter is not always sufficient. In order to meet these drawbacks, according to the present invention, a helicopter is proposed wherein the reaction torque RT on the fuselage can be completely compensated, at least can be regulated such that in that way, control of the helicopter, in particular yaw, remains possible, while the tail rotor 6 can be omitted. In Fig. 4, schematically, an exemplary embodiment of a flap mechanism 15 according to the invention is shown, which flap mechanism 15 will be elucidated hereinbelow as an example. The flap mechanism 15 comprises a hollow rotor shaft 4, which is provided at the bottom side with an outwardly extending, radial flange 16 with external toothing 17, engaged by a driving sprocket 18 of the schematically represented engine 19. With it, the rotor shaft 4 can be rotated about the longitudinal axis L by the engine 19. Near the top end of the hollow shaft 4, on both sides, brackets 20 are provided in which the flapping hinges 8 for the rotor blades 5 have been provided. For the sake of simplification, the swing hinges 9 and torsion hinges 10 have been omitted in this drawing. These can be provided in a conventional manner, and be driven as shown, for instance, in Fig. 3.

Through the hollow rotor shaft 4, a thrust shaft 21 extends, forming a driving element. On the top side of the thrust shaft 21, a yoke 22 is provided, equipped with two arms 23, which reach above the rotor blades 5. Between a free end 24 of each arm 23 and the underlying rotor blade 5, a spring 25 is provided which is relatively slack, such that with an upward movement of the thrust shaft 21, i.e. upwards in Fig. 4, the rotor blades 5 are moved along upwards at a slight delay, pivoted about the flapping hinges 8, while first, the springs 25 are slightly stretched. The purpose of this will be explained further hereinafter. Within the hollow shaft 4, on the thrust shaft 21, a spline 26 is attached, such that it is secured against rotation around the thrust shaft 21 or secured against shifting relative thereto. At the inside of the hollow shaft 4, a toothing 27 is provided, which matches the spline 26, such that the spline 26 and hence the thrust shaft 21 cannot rotate about the longitudinal axis L relative to the hollow rotor shaft 4. but ccim inove up and down in longitudinal direction 28. At the bottom side, the thrust shaft reaches beyond the bottom side of the hollow shaft 4. At the end 29 of the thrust shaft 21, a cross arm 30 is provided, resting by a free end 31, in the form of, for instance a ball, a bearing or the like, on the top side of a plate 32, such that the free end can be actively moved upwards as well as be actively moved downwards. Such a coupling is known from the normal swash plate, which plate is, for instance, round or oval and is suspended by a pivot 33 extending approximately at right angles to the longitudinal axis L. Adjacent a longitudinal edge, at a distance from the pivot 33, in a bearing 34, a pedal connection 35 is provided, such that with the aid of this pedal connection 35, the disc 32 can be pivoted about the pivot 33 between the first position in which the top surface 36 of the disc 32 extends at right angles to the longitudinal axis L, i.e. as a normal plane, while in a second position the top surface 36 is at an inclination relative to the longitudinal axis. The disc 32 is also indicated as non-rotating disc, since this disc can only pivot about the pivot 33, but cannot rotate about the longitudinal axis L.

Preferably, the pivot 33 extends transversely to the normal flying direction (at right angles to the longitudinal direction of the fuselage), while the arm 30 extends under the rotor blades 5, such that the yoke 22 is in the upper or lower position when these blades extend in the longitudinal direction of the fuselage.

It will be clear that when the sprocket 18 is driven, the hollow shaft 4 will be rotated about the longitudinal axis L, thereby carrying the thrust shaft 21 and the yoke 22, and hence the rotor blades 5, along as a result of the teeth of the spline 26 and the internal toothing 27 intermeshing. When the disc 32 is brought into the first position with the top surface 36 as a normal plane of the longitudinal axis , then the first end 31 of the cross arm 30 will rotate over this top surface 36, while the position of the thrust shaft 21 relative to the hollow shaft 4 will not change in longitudinal direction. As a result, the rotor blades 5 will be rotated, but not be caused to perform flapping movements, at least, not actively. When with the rotor, at least the rotor blades 5 at a standstill, the disc 32 is pivoted from the first position about the pivot 33 to, for instance, the position shown in Fig. 4, then the first end 31 of the cross arm 30 and hence the thrust shaft 21 with the yoke 22 will be moved upwards to the position shown in Fig. 4. As a result, each rotor blade 5 will be pulled, along upwards by the springs 25, pivoted about the flapping hinges 8. When thereupon with the disc 32 in the position as shown in Fig. 4, the sprocket 18 is driven by the engine 19, the rotor shaft 4 will be rotated about the longitudinal axis, taking along the thrust shaft 21 and yoke 22, as well as the cross arm.30. The end 31 will then perform a circular movement along the top surface 36 of the disc 32, while, as a result, the thrust shaft 21 with the yoke 22 will make an up and down movement, the amplitude of which is determined by the length of the cross arm 30 and the angle of inclination of the top surface 36. As a result of this up and down movement of the thrust shaft 21 with the yoke 22, the rotor blades 5 will be forced to make a flapping movement, i.e. an up and down pivoting movement about the flapping hinges 8, induced by the coupling by the springs 25 with the arms 23 of the yoke 22. Preferably, the springs 25 are chosen such that the natural flap eigenfrequency approximates the rate of revolution of the rotor. The first end 31, at least the flap mechanism 15 will, relative to the disc 32, apply a moment relative to the longitudinal axis L, which moment substantially depends on the angle of inclination of the disc 32. The engine 19 too, at least the sprocket 18, will also apply a moment on the rotor shaft 4, relative to the longitudinal axis. By properly selecting the position of the disc 32, at least of the angle of inclination of the top surface 36, these moments can be geared to each other, such that per revolution, on average, no resulting moment is applied to the fuselage. This means that no compensation for such a moment is necessary, so that, for instance, a tail rotor or other means as mentioned in the introduction for compensating such moments can be dispensed with. Thus, a particularly simple, robust, energy-saving and well- controllable helicopter is obtained. Further, it will be clear that through small changes of the angle of inclination of the disc 32 relative to the above-described condition of equilibrium, a resulting moment can be induced in both directions about the longitudinal axis L, so that yaw moments can be generated for the fuselage, with which control of the helicopter becomes possible in a simple manner. The angle of inclination of the disc 32 can be effected by the pilot in a simple manner with the aid the pedal connection 35. Surprisingly, it has appeared that these adjustments have virtually no influence on the normal control of the helicopter as given in Fig. 3. In Fig. 9, an alternative embodiment of a flap mechanism 15 according to the invention is shown, where only the components necessary for the understanding of the invention are schematically represented. In this Figure, the longitudinal axis L is represented by the Z-axis Z. Here, the rotor shaft 4 is provided at a top end with a sprocket 16, to be engaged by an engine (not shown) for rotation of the rotor shaft 4 about the longitudinal axis L. On the rotor shaft 4, an arm 30, extending radially, is provided with a bush 37 in which a thrust shaft 21 is included, which extends approximately parallel to the longitudinal axis L and which can move freely in axial direction 28 in the bush 37. By a ball hinge 31, the bottom end 29 of the thrust axis 21 rests on a disc 32, which is pivotable about an axis 33, indicated as Y-axis. Via a spring 25 and an intermediate rod 25A, the top end of the thrust shaft 21 is connected to the rotor blade 5, at a short distance from the flapping hinge 8, schematically represented by a pin. The operation of this flap mechanism 15 is identical to, at least corresponds to the flap mechanism of Fig. 4, it being noted that here, the flap mechanism 15 is shown for only one rotor blade 5. Upon rotation of the rotor shaft 4 and pivoting of the disc 32 to an inclined position, comparable to that shown in Fig. 4, the ball hinge 31 will trace a path over the top surface 36 of the disc 32, thereby moving the thrust shaft 21 up and down in axial direction 28, as a result of which, again with some delay, the rotor blade 5 will be forced to make a flapping movement, pivoting about the flapping hinge 8. It will be clear that opposite the arm 30 a similar mechanism can be provided for an opposite second rotor blade 5 of a pair of rotor blades, in which case, however, the end 29 resting on the blade 32 can be dispensed with. The thrust bars 21 for the two rotor blades can simply be interconnected by a rigid cross bar. In Fig. 10, schematically, a helicopter 40 according to the invention is shown, where it is clearly visible that no tail is provided. As flap mechanism, a flap mechanism 15 according to Fig. 4 has been schematically drawn, whose yoke 22 with the springs 25 and the rotor shaft 4 are visible. Furthermore, the control rods 13 of the conventional control system are schematically drawn. It will be clear, for that matter, that a helicopter 10 according to the invention can have many shapes, also, for instance, including a tail.

By way of illustration, a theoretic model of a helicopter 1 according to the invention will be described in more detail hereinbelow, with reference to hovering. From the prior art, comparable theories and formulae for forward flight are sufficiently known. During hovering, a downward flow will be induced by the rotor disc 3, the induced velocity Vi as a reaction to the upward thrust of the rotor 5.

From the Figures, it can be derived that the relative flow direction relative to a blade element is determined by Vi and the revolution velocity Ωr, as well as by the vertical movement of the element as a result of an up and down movement dβ/dt r, further to be called "flapping", wherein Ω is the angular speed of the rotor shaft 4 and β is the flapping angle of the blade 5 (Figs. 6 and 8). The resulting flow, relative to the blade element, includes an angle to the shaft plane 3 which equals (Fig. 7):

v_; Λ- r - dβldt

Φ = - - (1)

Ωr

The shape of the cross section of the blade element determines the relative flow direction, which corresponds to zero-lift, represented in Fig. 7 as the Ci = 0 line. The angle between the zero-lift direction and the shaft plane 3 is indicated as pitch angle θ of the blade element, which angle θ is increased or decreased when the pilot tilts the swash plate 11 or moves it up and down (see Fig. 3). The angle α = θ - φ is the angle of attack. The angle between the relative flow velocity and the zero-Hft direction determines the degree of lift on the blade element. The lift force is always perpendicular to the relative flow direction and not to the chord of the blade element. This means that the lift is tilted backwards over an angle φ, at least during the upstroke represented in Fig. 7. During a flap down stroke of the blade, the reverse effect occurs and the lift vector dL can be tilted forward, thus propelling the blade around the rotor shaft.

With reference to Fig. 7, the total aerodynamic force resisting the angular motion of the blade element about the rotor shaft 4 is (under small angle assumption):

dF = dL - φ + dD (2)

where dD is the profile drag caused by viscous effects. The contribution of the blade element to the shaft power for the shaft 4 required for driving the rotor 5 at a steady angular speed is given by:

dP_Λ = dF Ωr = dL - v, + dL - dβ I dt - r + dD - Ωr (3)

and the total shaft power required for the shaft 4 is:

where the integral relative to the blade azimuth angle ψ determines the average value during one revolution and the second integral is the total of all elements of the blade between root ^'and. tip. -.:(r=R). The contribution of the term dL.Vi in (3) to the total pQweϊ^'(4 ,i called induced power Pi. This represents the power corresponding to the kinetic energy in the flow by inducement of downward velocity Vi. Therefore, the power is intrinsically associated with the lifting of the helicopter weight.

The contribution of the term dD.Ωr to the total power is indicated with profile power P . This is associated with the losses due to viscous effects. Both the induced power Pi and the profile power P_p are known from manuals relating to helicopter flight and are not essential to the present invention. Here, of particular interest are the contribution of dL.dβ/dt.r to the total power indicated as Pβ. From \dL.r = M_a , where M_a is the aerodynamic flap moment of the blade about the flapping hinge, follows

For the sake of simplification, it is assumed that the flapping hinge 8 of a blade is placed on the rotor shaft 4, as represented in Figs. 6 and 8. If the weight of the blade is not taken into consideration, the moments around the flapping hinge are due to aerodynamic forces on the blade, the centrifugal force and the mechanical moment applied to the blade root. The aerodynamic forces contribute (M_a (Ψ), as discussed earlier. The centrifugal effects tend to stretch the blade in the shaft plane and the corresponding moment around the flapping hinge (small angle assumption) is

, which is the moment of inertia of the blade around the flapping hinge 8. The moment resulting from the flap mechanism 15 is represented as Mfi(ψ). The equation of motion for the flapping of the blade is then:

Id²β ldt² + Ω²Iβ = M_a{t)+ M {t) (6)

which represents harmonic variations of the flapping angle β. The undamped eigenfrequency is Ω, equal to the rotor velocity. This means that once-per-rev-excitations such as M_a- variations as a result of the swash plate and once-per-rev flap forcing leads to resonance of the blade motion. The flap reaction does not become very great, as the moment M_a also comprises a strong aerodynamic damping factor (in the order of 50% of critical damping for a typical helicopter blade), so that the resulting blade flap motion will follow from: β(ψ) = a₀ - a_x cos ψ -b_x sin ψ where ψ = Ωt (7)

Introducing the result (7) into the equation of motion (6) shows:

M_a{ψ) = Ω²Ia_Q -M_β(ψ) (8)

and inserting this result into equation (5) leads to:

It is clear that the integral represents the average power supplied to the rotor by the flap mechanism 15. When this power is indicated as Pfiap, equation (9) gives that Pβ= -Pfiap. The total shaft power required for maintaining a constant rotation of the rotor follows from:

P_Λ ^ P, + P_p - P_Jbp (10)

From which it appears that the required shaft power (P_sh) can be eliminated by making Pfia sufficiently large. The required driving power is then entirely ensured by the flap mechanism 15. In this case, no reaction moment will occur. The power required for the flap mechanism 15 for obtaining rotor velocity is equal to the shaft power which it replaces. For a rotor with only one blade it holds that it is particularly hard to control. For two or more blades, in particular an even number of blades on a rotor, it holds, by contrast, that this is well controllable.

Examples of flap mechanisms 15 are shown in Figs. 4 and 9. Naturally, many variants are possible, such as scissor mechanisms, hydraulic, pneumatic and/or electric mechanisms and the like. In a simple and independent manner, these mechanisms can be combined with a conventional control mechanism as shown, for instance, in Fig. 3. Important to note is an essential difference between the flap mechanism 15 and the application of cyclic pitch by the swash plate. Both are periodical, with a frequency of once- per-rev, but cyclic pitch is a-symmetrical (the blades are moved to the same extent but in the opposite direction) whereas the movement initiated by the flap mechanism is symmetrical. The working principle of a flap mechanism according to the invention has already been described hereinabove. When the flap forcing is only just sufficient, no shaft power needs to be transmitted to have the rotor rotate at a constant speed, so that the entire power is available for the flap mechanism 15. In this situation, a moment is applied to the inclined surface 36 which is equal in size but opposite in direction to the moment applied by the main engine on its connection, in the helicopter 1. As a result thereof, no reaction torque occurs on the helicopter.

The angle of inclination of the surface 36 is controlled by the pilot. The amplitude of the flap mechanism, in particular of the thrust shaft 21, determines the division of the entire engine power between the flap mechanism 15 and the rotor shaft 4. In this way, therefore, it is controlled whether or not a reaction torque remains and in what direction it is directed. In this way, therefore, a powerful yaw control can be realized for the helicopter, by adjusting the angle of inclination of the plate 32, at least the surface 36.

As a consequence of an asymmetrical, cyclic pitch, the factor dβ₀/dt.Mfi in the integral of equation (9) becomes equal in magnitude but opposite for the two blades and therefore adds nothing to Pfiap for the two blades. Conversely, the term dβfi/dt.Mfi in equation (9), where βfi represents the portion of the total flap angle that is caused by the flap forcing, is equal in magnitude as well as in sign for the two blades and is therefore the only term contributing to the flapping power Pfiap. Therefore, for a two-blade rotor system, equation (9) is as follows: - flapjolal ldt{ψ)-M_β{ψ)- d{ψ) (11)

where N is the number of blades, in the present case N=2. Therefore, for a two-blade system as shown, there is no interference between the swash plate control and the prevention of torque. In other words, there is no interference between the yaw, roll and longitudinal control of the helicopter. Therefore, a helicopter according to the present invention has conventional control characteristics both while hovering and in other flying movements.

As equation (6) describes a resonance situation for once-per-rev excitation, dβfi/dt(ψ) is in phase with Mα(ψ). In case the critical value is damped by approximately 50%, the aerodynamic moment M_a in equation (6) can be written as

where ζ=0.5. The amplitudes of βfi en Mα bear the following relation to each other:

β_flΩ²l = M_β (12)

So that equation (11) gives:

P_βap._lolal = Iβ_β ² (13)

When typical values for a relatively light helicopter are filled in, it is found that the magnitude of the flap angles required for elimination of a remaining torque needs to be in the order of 0.1 radials, a value comparable to the angles required for controlling the helicopter.

A fluctuation will occur in the vertical force on the fuselage of the helicopter 1 as a result of vertical accelerations of the blade masses and as a result of fluctuations in the lifting powers. In view of the relatively small value of the flapping angles as described hereinabove, a conventional vibration insulating system can be used to obviate this problem. A different solution may be the use of a four-bladed rotor where each time two pairs of blades, diametrically opposed, make flapping movements in anti-phase. As a result, the fluctuating vertical forces are exactly averted. The 2 x teeter rotor configuration can also be particularly advantageous.

From the above description, it clearly follows that a helicopter according to the present invention can be controlled in a conventional manner while during flight the drive can be completely effected by the flap mechanism. When starting the helicopter 1, the plate 32 will be brought into a substantially horizontal position, so that virtually the entire engine power is transmitted to the rotor shaft 4 for setting the rotor shaft with blades in motion. As soon as the rotor shaft with the blades has sufficient angular speed, the disc 32 can be tilted, so that the drive can slowly and gradually be taken over by the flap mechanism 15, so that a smooth transition is obtained.

The invention is not limited in any way to the exemplary embodiments represented in the description and the drawings. Many variations thereon are possible within the framework of the invention as outlined by the claims.

As described, a helicopter 1 according to the invention can have any desired form, provided or not provided with a tail. A helicopter according to the invention can be designed for passengers as well as for freight. Naturally, a rotor with flap mechanism according to the invention could be combined with conventional control mechanisms, such as a relatively small tail rotor, tip drive with the aid of ram -jets or such systems described in the introduction, while the handling of the helicopter can be improved. Naturally, the blade geometry of the rotor blades can be adjusted in a simple manner. Many variations are possible for the flap mechanisms, a-s, long as^'they can force the rotor blades 5, in a positively controlled manner, to make an up and down movement, at least a pivoting movement about a flapping hinge, which movement can be controlled by the pilot, at least a control system of the helicopter. Optionally, electronic or mechanical means can be .provided for controlling the amplitude of the flapping movement of the rotor blades in a semi-automatic manner, for instance for limiting maximum allowable remaining torques. Also, flap mechanisms can be used with which the blades can move up and down perpendicularly.

These and many comparable variations are understood to fall within the framework of the invention as outlined by the claims.

Claims

1. A helicopter, provided with a rotor with at least one rotor blade, characterized in that drive means are provided for actively moving the or each rotor blade up and down during rotation of the rotor, in particular about a flapping hinge thereof.

2. A helicopter according to claim 1, wherein the rotor has an even number of blades.

3. A helicopter according to claim 1 or 2, wherein the drive means are arranged for moving the or each rotor blade up and down at least once per revolution of the rotor.

4. A helicopter according to any one of the preceding claims, wherein the drive means comprise adjustment means for adjusting the amplitude of the up and down movement of the or each rotor blade.

5. A helicopter according to any one of the preceding claims, wherein the drive means comprise an engine which is in driving connection with a shaft of the rotor, while the drive means, during use, apply a torque to said fuselage which is opposite to the reaction torque applied to the fuselage of the helicopter by the engine as a result of the driving moment of the engine.

6. A helicopter according to claim 5, wherein the adjustment means are adjustable, such that the torque applied during use by the drive means on the fuselage is regulable between a value which is smaller than said reaction torque and a value which is greater than said reaction torque, the arrangement being such that yaw of the helicopter can thus be adjusted.

7. A helicopter according to any one of the preceding claims, wherein the rotor comprises a hollow shaft within which a drive element is move able in longitudinal direction, which drive element is connected, via a coupling element, to the or each rotor blade, at a distance from a flapping hinge through which the respective rotor blade is connected to the shaft.

8. A helicopter according to claim 7, wherein the coupling element extends substantially above the rotor.

9. A helicopter according to any one of the preceding claims, wherein the rotor comprises a shaft around which a tiltably arranged guiding element with a running surface is provided, wherein a driving element is provided which is rotatable about the same longitudinal axis as the shaft of the rotor and is moveable in longitudinal direction, which driving element is coupled to the shaft of the rotor in the sense of rotation, the drive element, near a bottom side, being provided with a running element resting on the guiding element, at a distance from said longitudinal axis, such that upon rotation of the driving element about said longitudinal axis and adjustment of the guiding element such that the running surface includes an angle with the longitudinal axis, the driving element is moved in a reciprocating manner in longitudinal direction.

10. A helicopter according to claim 9, wherein angle adjustment means are provided for adjusting the guiding element, which angle adjustment means comprise operating means which are placed in the helicopter instead of means for operating a tail rotor as in a conventional helicopter.

11. A helicopter according to any one of the preceding claims, wherein the helicopter has substantially no tail, at least no tail provided with compensating means such as a tail rotor or air blow-out means.

12. A helicopter according to any one of the preceding claims, wherein the control means for the helicopter are substantially conventional control means, apart from means for controlling the up and down movement of the or each rotor blade.

13. A helicopter according to any one of the preceding claims, in particular according to any one of claims 7 or 8, wherein the drive means for moving the rotor blades up and down are connected to the rotor blades via spring means.

14. A helicopter according to claim 13, wherein the spring means are connected to the respective rotor blades at a distance from the or each pivot by which the respective rotor blade is connected to the rotor shaft.

15. A helicopter according to claim 13 or 14, wherein the spring means are dimensioned and designed such that they are geared to the eigenfrequency of the up and down movement of the rotor blades.

16. A helicopter according to any one of the preceding claims 13 — 15, wherein the spring means are relatively slack, in particular slack to such an extent that the spring system meets the formula Ke² < 0.21Ω²/ .