METHOD AND DEVICE FOR DETERMINING A STATE OF A ROTORCRAFT ROTOR

Abstract

A device for determining a state of a rotor of a rotorcraft. The rotorcraft comprises a fuselage and a main rotor provided with a hub rotating about a mast of the rotor and with a plurality of blades whose second ends describe a trajectory defining a tip path plane. The device includes a sensor for measuring an angular velocity of a blade about a pitch axis. The device thus makes it possible to determine a state of the rotor, comprising, for example, estimates of a longitudinal cyclic pitch and of a lateral cyclic pitch of the blade with respect to the tip path plane.

Claims

1. A method for determining a state of a rotor of a rotorcraft, the rotorcraft comprising a fuselage and at least one rotor provided with a mast, a hub and a plurality of blades, each blade comprising a first end and a second end, the hub being connected to the mast and rotating about an axis Of the mast, each blade being connected by a first end to the hub, each blade being able to rotate about at least one articulation axis of the blade, the second end of a blade describing, during a rotation of the hub about the axis, a trajectory close to a mean plane referred to as the tip path plane, wherein the method comprises the following steps: directly measuring the change over time of an angular velocity component of a blade about the articulation axis with respect to an inertial reference frame; and determining a state of the rotor by processing the component.

2. The method according to claim 1 wherein, during the step of determining a state of the rotor, the processing operation is a harmonic analysis of the angular velocity component.

3. The method according to claim 1 wherein the method comprises an intermediate step of detecting the instant a first marker attached to the rotor passes in front of a second marker attached to the fuselage.

4. The method for determining a state of a rotor of a rotorcraft, the rotorcraft comprising a fuselage and at least one rotor provided with a mast, a hub and a plurality of blades, each blade comprising a first end and a second end, the hub being connected to the mast and rotating about an axis of the mast, each blade being connected by a first end to the hub, each blade being able to rotate about at least one articulation axis of the blade, the second end of a blade describing, during a rotation of the hub about the axis, a trajectory close to a mean plane referred to as the tip path plane, wherein the method comprises the following steps: directly measuring the change over time of an angular velocity component of a blade about the articulation axis with respect to an inertial reference frame; and determining a state of the rotor by processing the component, the state of the rotor comprising a longitudinal cyclic pitch and a lateral cyclic pitch of the blade with respect to the tip path plane determined by extracting an amplitude and a phase of the fundamental sinusoidal component of the angular velocity of the blade about the articulation axis, wherein the fundamental sinusoidal component has its frequency equal to that of rotation of the hub of the rotor about the axis.

5. The method according to claim 4 wherein the extraction of the amplitude and the phase of the fundamental sinusoidal component of the angular velocity of the blade about the articulation axis is carried out by at least any one of the following methods: Fourier series decomposition; solving a system of two equations with two unknowns utilizing the instantaneous measurements on two different blades; least squares estimation of the two unknowns of a system of equations utilizing the instantaneous measurements on more than two different blades; synchronous demodulation; recursive least squares method; non-recursive least squares method; and Kalman filtering.

6. The method according to claim 4 wherein the state of the rotor comprises a cone angle of the rotor determined by extracting a mean value of the fundamental sinusoidal component of the angular velocity of the blade about the articulation axis by at least any one of the following methods: Fourier series decomposition; solving a system of three equations with three unknowns utilizing the instantaneous measurements on three different blades; least squares estimation of the three unknowns of a system of equations utilizing the instantaneous measurements of more than three different blades; synchronous demodulation; recursive least squares method; non-recursive least squares method; and Kalman filtering.

7. The method according to claim 4 wherein the state of the rotor comprises: an elevation angle of the blade with respect to the tip path plane; longitudinal and lateral inclination angles and of the tip path plane with respect to a plane of the hub perpendicular to the axis.

8. The method according to claim 7 wherein the method comprises the following additional steps: measuring a longitudinal cyclic pitch and a lateral cyclic pitch of the blade with respect to the hub plane; and calculating longitudinal and lateral inclination angles's) of the tip path plane with respect to the hub plane, the inclination angles of the tip path plane being a function of the longitudinal and lateral cyclic pitches of the blade measured with respect to the hub plane and the estimates of the longitudinal and lateral cyclic pitches of the blade with respect to the tip path plane.

9. The method according to claim 8 wherein the inclination angles of the tip path plane are equal respectively to differences between the longitudinal and lateral cyclic pitches of the blade measured with respect to the hub plane and the estimates of the longitudinal and lateral cyclic pitches of the blade with respect to the tip path plane or indeed equal respectively to a deviation between a corrective term between the longitudinal and lateral cyclic pitches of the blade measured with respect to the hub plane and the estimates of the longitudinal and lateral cyclic pitches of the blade with respect to the tip path plane taking into account a pitch-flap coupling coefficient of the rotor.

10. The method according to claim 8 wherein, during the step of measuring a longitudinal cyclic pitch and a lateral cyclic pitch of the blade, the longitudinal cyclic pitch and the lateral cyclic pitch of the blade are measured at a non-rotating part of a swashplate controlling the collective and cyclic pitch variations of the blade, at servomechanisms actuating the non-rotating part of the swashplate or indeed at a device for controlling the collective and cyclic pitches of the blade.

11. The method according to claim 7 wherein the method also includes a step of calculating a bending moment of the mast of the rotor.

12. The method according to claim 7 wherein the method includes a step of performing an assisted or automatic take-off of the rotorcraft by slaving the tip path plane according to a predetermined inclination setpoint with respect to a horizontal plane.

13. The method according to claim 12 wherein, during the step of performing an assisted or automatic take-off of the rotorcraft, a pilot of the rotorcraft acts solely on a lever for controlling the collective pitch of the blades of the rotor, the cyclic pitches of the blades of the rotor being controlled automatically in order to slave the longitudinal and lateral inclination angles of the tip path plane with respect to a horizontal plane, regardless of the slope of the ground and regardless of the wind acting on the rotorcraft.

14. The method according to claim 11 wherein the method comprises: a step of estimating an aerodynamic force generated by the rotor; a step of estimating the apparent mass of the rotorcraft; and a step of estimating a current mass of the rotorcraft.

15. A device for determining a state of a rotor of a rotorcraft, the rotorcraft comprising a fuselage and at least one rotor provided with a mast, a hub and a plurality of blades, each blade comprising a first end and a second end, the hub being connected to the mast and rotating about an axis of the mast, each blade being connected by a first end to the hub, each blade being able to rotate about an articulation axis of the blade, the second end of a blade describing, during a rotation about the axis, a trajectory close to a mean plane referred to as the tip path plane; wherein the device is configured to implement the method according to claim 1, and comprises: at least one memory; a sensor measuring a change over time of an angular velocity component of a blade about the articulation axis with respect to an inertial reference frame; and at least one calculator determining the state of the rotor.

16. The device according to claim 15 wherein the sensor is a rate gyro arranged in a blade, a sensitive axis of the rate gyro being oriented parallel to the articulation axis of the blade in order to measure the angular velocity component of the blade about the articulation axis.

17. The device according to claim 15 wherein the rotor comprising spherical thrust-bearings, one spherical thrust-bearing connecting a blade to the hub, the sensor comprises an accelerometer whose sensitive axis is oriented parallel to a flapping axis of the blade and one or more rate gyros arranged in a blade, whose sensitive axes are oriented respectively parallel to a pitch axis, the flapping axis and a drag axis of the blade in order to estimate a deflection angle of each spherical thrust-bearing so as to determine the mechanical stresses experienced by the spherical thrust-bearing, predict damage to the spherical thrust-bearing, and reduce the maintenance costs of the spherical thrust-bearing.

18. The device according to claim 15 wherein the device includes a pulse detector measuring instants a first marker attached to the rotor passes in front of a second marker attached to the fuselage

Description

BRIEF DESCRIPTION OF THE DRAWINGS

[0102] The invention and its advantages appear in greater detail from the following description of examples given by way of illustration with reference to the accompanying figures, in which:

[0103] FIGS. 1 to 3 show rotorcraft equipped with a device for determining a state of a rotorcraft rotor;

[0104] FIGS. 4 to 7 show views of a blade;

[0105] FIG. 8 shows a schematic representation of the rotor disk; and

[0106] FIG. 9 shows a graph representing the change over one complete revolution of a rotor of an angular velocity component ω.sub.θ(t) of each blade of this rotor about its pitch axis.

DETAILED DESCRIPTION OF THE INVENTION

[0107] Elements present in more than one of the figures are given the same references in each of them.

[0108] As shown in FIGS. 1 to 3, a rotorcraft 10 comprises a main rotor 11 positioned above a fuselage 16 of the rotorcraft 10 and an anti-torque rear rotor 18 positioned at the rear end of a tail boom 19.

[0109] A body frame (X.sub.F, Y.sub.F, Z.sub.F) is attached to the fuselage 16 of this rotorcraft 10. A longitudinal axis X.sub.F of the fuselage 16 of the rotorcraft 10 is oriented from the rear of the rotorcraft 10 towards the front of the rotorcraft 10. A normal axis Z.sub.F is oriented from top to bottom perpendicular to the longitudinal axis X.sub.F, and a lateral axis Y.sub.F is oriented from left to right perpendicular to the longitudinal X.sub.F and normal Z.sub.F axes.

[0110] The main rotor 11 comprises a mast 12, a hub 13 and blades 14. Each blade 14 has a first end 141 connected to the hub 13 and a second free end 142 as well as a leading edge 143 and a trailing edge 144. The mast 12 is secured to the hub 13 and rotates the hub 13 and the blades 14 about an axis A1 of the mast 12.

[0111] Each blade 14 is also able to rotate about at least one articulation axis, for example about its pitch axis A2, as well as about its flapping axis A3 and its drag axis A4, as shown in FIG. 4.

[0112] FIG. 5 shows the blade 14 and its pitch angle θ.

[0113] FIG. 6 shows the blade 14 and its flap angle β.

[0114] FIG. 7 shows the blade 14 and its drag angle δ.

[0115] Consequently, during a rotation of the hub 13 about the axis A1, each point of the hub 13 moves in a hub plane HP perpendicular to the axis A1 of the mast 12 while the second free end 142 of each blade 14 describes a substantially planar trajectory, close to a mean plane referred to as the “tip path plane TPP”. The trajectory of the second free end 142 of each blade 14 actually experiences slight fluctuations on either side of the tip path plane TPP, which can generally be disregarded. The projection of this trajectory in the tip path plane TPP is substantially circular.

[0116] For example, the pitch axis A2 may be constituted by a straight line contained in the blade 14, integral with the blade 14 and forming a radius of the rotor disk. Likewise, the flapping axis A3 may be constituted by a straight line integral with the blade 14, perpendicular to the pitch axis A2, and whose direction is a tangent to the periphery of the rotor disk at the tip of the free end 142 of said blade 14. The drag axis A4 may be constituted by a straight line integral with the blade 14 and normal to the plane of the rotor disk.

[0117] Furthermore, each blade 14 may be connected to the hub 13, for example by means of three mechanical articulations, or indeed by means of a flexible connection, or indeed by means of a spherical thrust-bearing 17 acting alone as the three articulations. In addition, each blade 14 forms a rotor disk during rotation of the hub 13 about the axis A1 of the mast 12.

[0118] Another body frame (X.sub.R, Y.sub.R, Z.sub.R) is also connected to the fuselage 16 of the rotorcraft 10 and is more precisely integral with the hub plane HP. This other body frame (X.sub.R, Y.sub.R, Z.sub.R) is non-rotating with respect to the fuselage 16 of the rotorcraft 10, a longitudinal axis X.sub.R being formed by a projection on the hub plane HP of the longitudinal axis X.sub.F of the body frame (X.sub.F, Y.sub.F, Z.sub.F), a lateral axis Y.sub.R being formed by a projection on the hub plane HP of the lateral axis Y.sub.F and a normal axis Z.sub.R being oriented perpendicular to the hub plane HP, from top to bottom.

[0119] A reference frame (X, Y, Z) may also be integral with the tip path plane TPP. This reference frame (X, Y, Z) is non-rotating with respect to the fuselage 16 of the rotorcraft 10, a longitudinal axis X being formed by a projection on the tip path plane TPP of the longitudinal axis X.sub.F of the body frame (X.sub.F, Y.sub.F, Z.sub.F), a lateral axis Y being formed by a projection on the tip path plane TPP of the lateral axis Y.sub.F and a normal axis Z being oriented perpendicular to the tip path plane TPP, from top to bottom.

[0120] The rotorcraft 10 comprises a device 1 for determining a state of the main rotor 11 of the rotorcraft 10. This device 1 is based principally on utilizing the sinusoidal oscillations of at least one blade 14 of the main rotor 11 of the rotorcraft 10 about at least one articulation axis of this blade 14 in order to determine this state of the main rotor 11 of the rotorcraft 10.

[0121] However, this device 1 can also be adapted in order to determine the state of the rear rotor 18 of the rotorcraft 10 by utilizing the sinusoidal oscillations of at least one blade of the rear rotor 18 of the rotorcraft 10 about at least one articulation axis of this blade.

[0122] The device 1 for determining a state of the main rotor 11 of the rotorcraft 10 may be based on utilizing the sinusoidal oscillations of the pitch of at least one blade 14 of the main rotor 11 about its pitch axis A2. The device 1 may also be based on utilizing the sinusoidal oscillations of the movements of a blade 14 about its flapping axis A3 and/or its drag axis A4.

[0123] The state of the rotor comprises, in particular, a longitudinal cyclic pitch θ.sub.1S/TPP and a lateral cyclic pitch θ.sub.1C/TPP of a blade 14 of the main rotor 11 with respect to said tip path plane TPP. The state of the rotor may also comprise the cone angle β.sub.0 of the blade 14 as well as longitudinal and lateral inclination angles β.sub.1C and β.sub.1S of the tip path plane TPP with respect to the hub plane HP and/or an aerodynamic force F generated by the main rotor 11.

[0124] The device 1 for determining a state of the main rotor 11 of the rotorcraft 10 comprises a calculator 2 provided with a memory 3. In addition, the device 1 includes a sensor 4 on at least one blade 14 of the main rotor 11. This device 1 is configured to implement a method for determining a state of the main rotor 11 of the rotorcraft 10 in order to determine the state of the main rotor 11 of the rotorcraft 10.

[0125] The memory 3 of the calculator 2 stores at least one executable code of an algorithm for carrying out this method. The calculator 2 executes this code. The calculator 2 may, for example, comprise at least one processor and at least one memory, or at least one programmable logic array, at least one integrated circuit, at least one programmable system, at least one logic circuit, at least one analog circuit, these examples not limiting the scope given to the expression “calculator”. The calculator 2 may be a calculator dedicated to carrying out this method or indeed a calculator of the rotorcraft 10 having a plurality of functions.

[0126] The sensor 4 comprises, for example, at least one rate gyro arranged in a blade 14. A sensitive axis of the rate gyro may be oriented parallel to the pitch axis A2 of the blade 14. The sensor 4 measures an angular velocity component ω.sub.θ(t) of a blade 14 about the pitch axis A2 of this blade 14 with respect to an inertial reference frame. In this way, the sensor 4 makes it possible to measure the change over time of this angular velocity component ω.sub.θ(t) of a blade 14 about its pitch axis A2 with respect to an inertial reference frame. The measurement ω.sub.θ of a rate gyro whose sensitive axis is oriented parallel to the pitch axis A2 may be written as follows:

ω.sub.θ={dot over (θ)}+Ω.Math.β≅Ω.Math.(β.sub.0+θ.sub.1S.Math.cos ψ−θ.sub.1C.Math.sin ψ),

{dot over (θ)} being the derivative with respect to time of the pitch angle of the blade 14;
Ω being the speed of rotation of the hub 13 of the rotor 11;
β being the flap angle of the blade 14;
β.sub.0 being the mean flap angle, or cone angle or elevation angle of the blade 14 with respect to the tip path plane TPP (by definition of the tip path plane TPP, the sine and cosine components of the flap angle, β.sub.1c and β.sub.1s, are zero when they are referenced to this plane, hence the equality β=β.sub.0) and;
θ.sub.1S and θ.sub.1C being the longitudinal and lateral cyclic pitches of the blade 14.

[0127] A sensitive axis of the rate gyro may also be oriented parallel to the flapping axis A3 of the blade 14. The sensor 4 measures an angular velocity component ω.sub.β(t) of a blade 14 about the flapping axis A3 of this blade 14 with respect to an inertial reference frame. In this way, the sensor 4 makes it possible to measure the change over time of this angular velocity component ω.sub.β(t) of a blade 14 about its flapping axis A3 with respect to an inertial reference frame. The measurement ω.sub.β of a rate gyro whose sensitive axis is oriented parallel to the flapping axis A3 may be written as follows:

ω.sub.β={dot over (β)}+Ω.Math.θ≅Ω.Math.(θ.sub.0+θ.sub.1S.Math.sin ψ+θ.sub.1C.Math.cos ψ),

{dot over (β)} being the derivative with respect to time of the flap angle (zero when the reference plane is the tip path plane TPP);
Ω being the speed of rotation of the hub 13 of the rotor 11;
θ being the pitch angle of the blade 14;
θ.sub.0 being the collective pitch of the blade 14; and
θ.sub.1S and θ.sub.1c being the longitudinal and lateral cyclic pitches of the blade 14.

[0128] A sensitive axis of the rate gyro may also be oriented parallel to the drag axis A4 of the blade 14. The sensor 4 measures an angular velocity component ω.sub.δ(t) of a blade 14 about the drag axis A4 of this blade 14 with respect to an inertial reference frame. In this way, the sensor 4 makes it possible to measure the change over time of this angular velocity component ω.sub.δ(t) of a blade about its drag axis A4 with respect to an inertial reference frame. The measurement ω.sub.δ of a rate gyro whose sensitive axis is oriented parallel to the drag axis A4 may be written as follows:

ω.sub.δ={dot over (δ)}+Ω,

{dot over (δ)} being the derivative with respect to time of the drag angle; and
Ω, the speed of rotation of the hub 13 of the rotor 11.

[0129] The sensor 4 may also comprise a multi-axis rate gyro or indeed two rate gyros whose sensitive axes may be oriented respectively parallel to the pitch axis and the flapping axis A3 of the blade 14. In this way, the sensor 4 makes it possible to measure the change over time of the angular velocity components ω.sub.θ(t), ω.sub.β(t) and ω.sub.δ(t) of a blade 14 respectively about its pitch axis A2, its flapping axis A3 and its drag axis A4 with respect to an inertial reference frame.

[0130] The sensor 4 may also comprise a triaxial rate gyro or indeed three rate gyros whose sensitive axes may be oriented respectively parallel to the pitch axis A2, the flapping axis A3 and the drag axis A4 of the blade 14.

[0131] The sensor 4 may also comprise an accelerometer whose sensitive axis is oriented parallel to the flapping axis A3 of the blade 14 in order to estimate a static value of the angle of the blade 14 about its drag axis A4. This accelerometer may be positioned, for example, vertically in line with a spherical thrust-bearing 17, as shown in FIG. 7. The measurement γ.sub.β of this accelerometer positioned in this way can thus be written as follows:

γ.sub.β=δ.Math.r.Math.Ω.sup.2,

δ being the drag angle;
r being the distance between the accelerometer and the axis of rotation A1; and
Ω being the speed of rotation of the hub 13 of the rotor 11.

[0132] Since the distance r and the speed of rotation Ω are known, it is thus possible to deduce the drag angle δ(t) and, in particular, its mean or static component δ.sub.0.

[0133] In addition, the rotorcraft 10 may comprise an AHRS device 6 supplying, in particular, specific forces applied to the fuselage 16 of the rotorcraft 10, for example along the axes X.sub.F, Y.sub.F, Z.sub.F.

[0134] The rotorcraft 10 shown in FIG. 1 comprises a single sensor 4 arranged on a blade 14 of the main rotor 11.

[0135] According to FIG. 2 and FIG. 3, the rotorcraft 10 comprises several sensors 4, one sensor 4 being arranged on each blade 14 of the main rotor 11 of the rotorcraft 10.

[0136] For example, according to FIG. 3, the rotorcraft 10 comprises five blades 14 and five sensors 4. For example, one sensor 4 is positioned at the sleeve 20 of each blade 14.

[0137] The rotorcraft 10 shown in FIGS. 1 and 2 has the axis A1 of the mast 12 of its rotor 11 inclined by a few degrees. This inclination is typically between 3° and 4° with respect to the normal axis Z.sub.F, towards the front. Consequently, the hub plane HP is inclined with respect to a plane formed by the longitudinal X.sub.F and lateral Y.sub.F axes.

[0138] In the rotorcraft 10 shown in FIG. 2, the device 1 may also comprise a pulse detector 5 detecting instants when a first marker 115 attached to the main rotor 11 passes in front of a second marker 165 attached to the fuselage 16. The pulses generated by this pulse detector 5 thus enable the calculator 2 to determine the azimuth ψ of each blade 14 about the axis A1 of the mast 12 of the rotor 11 as well as the speed of rotation Ω of the hub 13 about its axis A1. This pulse detector 5 comprises, for example, a Hall-effect cell secured to the second marker 165 of the fuselage 16, and a magnet positioned secured to the first marker 115 of the main rotor 11.

[0139] The rotorcraft 10 may also comprise a complementary sensor 7 arranged, for example, on a non-rotating part of a swashplate 17 controlling the pitch variations of the blades 4 as shown in FIG. 2 or indeed at each servomechanism actuating the non-rotating part of this swashplate 17 or indeed at a device for controlling the pitch variation of the blades 4. This complementary sensor 7 can thus directly or indirectly measure a pitch angle θ.sub.HP with respect to the hub plane HP.

[0140] FIG. 8 schematically shows the rotor 11 as well as the tip path plane TPP and the hub plane HP. Since the tip path plane TPP and the hub plane HP are generally not parallel, the longitudinal θ.sub.1S and lateral θ.sub.1C cyclic pitches of each blade 14 of the rotor 11 are different depending on whether they are measured with respect to the tip path plane TPP or indeed with respect to the hub plane HP. The relations for the transition between these two reference planes are written, for example for a rotor 11 having zero pitch-flap coupling, as follows:

θ.sub.1C/TPP=θ.sub.1C/HP+β.sub.1S,  (E3)

θ.sub.1S/TPP=θ.sub.1S/HP−β.sub.1C,  (E4)

where:
θ.sub.1S/TPP and θ.sub.1C/TPP: longitudinal and lateral cyclic pitches of a blade 14 of the rotor 11 with respect to the tip path plane TPP;
θ.sub.1S/HP and θ.sub.1C/HP: longitudinal and lateral cyclic pitches of a blade 14 of the rotor 11 with respect to the hub plane HP; and
β.sub.1C and β.sub.1S, the inclination of the tip path plane TPP with respect to the hub plane HP, respectively longitudinally (i.e., about the lateral axis Y.sub.F) and laterally (i.e., about the longitudinal axis X.sub.F).

[0141] For a rotor 11 having non-zero pitch-flap coupling, the relations for the transition between reference planes may comprise a corrective term, namely the pitch-flap coupling coefficient.

[0142] The pitch angle θ comprises a collective component θ.sub.0 and a cyclic component that can be resolved into a longitudinal cyclic component θ.sub.1S and a lateral cyclic component θ.sub.1C.

[0143] The method for determining a state of the main rotor 11 of the rotorcraft 10 that can be implemented by the device 1 comprises at least the following two steps:

[0144] directly measuring the change over time of an angular velocity component ω.sub.θ(t) of a blade 14 about its pitch axis A2 with respect to an inertial reference frame; and

[0145] determining a state by processing the angular velocity component ω.sub.θ(t).

[0146] The angular velocity ω.sub.θ(t) of a blade 14 about its pitch axis A2 in the reference frame (X,Y,Z) attached to the tip path plane TPP is first measured by a sensor 4 arranged on this blade 14. This angular velocity ω.sub.θ(t) is equal to the derivative with respect to time of the angle θ.sub.TPP of this blade 14 in the same reference frame (X,Y,Z) and can therefore be expressed in the form of a limited Fourier series development as a function of the angular orientation Ω.Math.t of the rotor 11 about the axis A1 such that:

ω.sub.θ(t)=Ω.Math.sin β.sub.0+Ω.Math.θ.sub.1S/TPP.Math.cos(Ω.Math.t)−Ω.Math.θ.sub.1C/TPP.Math.sin(Ω.Math.t),  (E6) where:

Ω.Math.sin β.sub.0: mean value, or continuous component;
(Ω.Math.θ.sub.1S/TPP) and (−Ω.Math.θ.sub.1C/TPP): harmonic coefficients of order 1, corresponding to the frequency Ω/2π; and
β.sub.0: cone angle or elevation angle of the blade with respect to the tip path plane TPP, defined between the tip path plane TPP and the blade 14.

[0147] FIG. 9 comprises a graph showing the change over one complete revolution of the rotor 11 of the angular velocity components ω.sub.θ(t) of each of the five blades 14 of the rotor 11 about their pitch axis A2. This graph shows the signals from the five sensors 4 equipping the five blades 14 of the main rotor 11 of the rotorcraft 10 shown in FIGS. 2 and 3.

[0148] From the expression (E6) of the angular velocity ω.sub.θ(t) of a blade 14, the cone angle β.sub.0 as well as the longitudinal cyclic pitch θ.sub.1S/TPP and the lateral cyclic pitch θ.sub.1C/TPP of the blade 14 with respect to the tip path plane TPP, constituting a state of the rotor 11, can be estimated following this processing solely of the angular velocity component ω.sub.θ(t), for example by a harmonic analysis, by means of the calculator 2, according to at least one of the following known methods:

[0149] Fourier series decomposition;

[0150] solving a system of three equations with three unknowns utilizing the instantaneous measurements on three different blades 14;

[0151] solving a system of five equations with three unknowns utilizing the instantaneous measurements of the five blades 14, the 5/3 overdetermination being solved, for example, by the least squares method;

[0152] synchronous demodulation;

[0153] recursive least squares method;

[0154] non-recursive least squares method; and

[0155] Kalman filtering.

[0156] These estimates of the cone angle β.sub.0 and of the longitudinal θ.sub.1S/TPP and lateral θ.sub.1C/TPP cyclic pitches of the blade 14 are determined by extracting, respectively, a mean value of the angular velocity ω.sub.θ(t), as well as an amplitude and a phase of the sinusoidal component of order 1 of the angular velocity ω.sub.θ(t) by means of the calculator 2.

[0157] The method may comprise an intermediate step of measuring the instant when the first marker 115 attached to the rotor 11 passes in front of the second marker 165 attached to the fuselage 16 by means of the pulse detector 5 such that the calculator 2 can determine the azimuth ψ and the speed of rotation Ω of each blade 14 about the axis A1 and, in particular, of each blade 14 equipped with a sensor 4. Subsequently, the calculator 2 may determine the phase of the sinusoidal component of the angular velocity ω.sub.θ(t) of the blade 14 about its pitch axis A2.

[0158] In addition, the state of the rotor may include the longitudinal and lateral inclination angles β.sub.1C and β.sub.1S of the tip path plane TPP with respect to the hub plane HP and/or an aerodynamic force F generated by the rotor 11.

[0159] In order to determine these inclination angles β.sub.1C and β.sub.1S of the tip path plane TPP with respect to the hub plane HP, the method may include the following additional steps:

[0160] measuring a longitudinal cyclic pitch θ.sub.1S/HP and a lateral cyclic pitch θ.sub.1C/HP of the blade 14 with respect to the hub plane HP perpendicular to the axis A1, by means of the complementary sensor 7; and

[0161] calculating longitudinal and lateral inclination angles β.sub.1C and β.sub.1S of the tip path plane TPP with respect to the hub plane HP, as functions of the longitudinal θ.sub.1S/HP and lateral θ.sub.1C/HP cyclic pitches of the blade 14 measured with respect to the hub plane HP and estimates of the longitudinal θ.sub.1S/TPP and lateral θ.sub.1C/TPP cyclic pitches of the blade 14 with respect to the tip path plane TPP. Where appropriate, these functions also involve the pitch-flap coupling coefficient.

[0162] The method may also include a step of calculating a bending moment of the mast 12 of the rotor 11. Knowing this bending moment of the mast 12 of the main rotor 11 of the rotorcraft 10 advantageously makes it possible to estimate the mechanical stresses experienced by the mast 12 of the main rotor 11 in particular. This bending moment of the mast 12 is determined as a function of the moment of inertia of the blades 14, the speed of rotation Ω of the hub of the rotor 11 and of the eccentricity of the flapping axis A3 of the blades 14.

[0163] The method may also comprise a step of performing an assisted or automatic take-off of the rotorcraft 10 by slaving the tip path plane TPP to a predetermined inclination set point with respect to a horizontal plane of a terrestrial reference frame defined perpendicularly to the direction of terrestrial gravity. During this stage of performing an assisted or automatic take-off of the rotorcraft 10, a pilot of the rotorcraft 10 acts solely on a lever for controlling the collective pitch of the blades 14 of the rotor 11, the cyclic pitches of the blades 14 of the rotor 11 of the rotorcraft 10 being controlled automatically in order to slave the longitudinal and lateral inclination angles of the tip path plane TPP with respect to a horizontal plane in a terrestrial reference frame, possibly utilizing the roll and pitch angles of the fuselage of the rotorcraft measured, for example, by the AHRS device 6.

[0164] The method may also comprise a step of estimating an aerodynamic force F generated by the rotor 11, a step of estimating the apparent mass of the rotorcraft 10 and a step of estimating a current mass M of the rotorcraft 10.

[0165] The cone angle β.sub.0 makes it possible to determine in a known manner the aerodynamic force F generated by the rotor 11. This aerodynamic force F helps balance the rotorcraft 10 during flight by opposing, in particular, the apparent mass of the rotorcraft 10 during hovering flight. The current mass M of the rotorcraft 10 can then be estimated as a function of this aerodynamic force F generated by the rotor 11, this apparent mass of the rotorcraft 10 and the specific force experienced by the rotorcraft 10 determined, for example, by the AHRS device 6.

[0166] Furthermore, the main rotor 11 may comprise spherical thrust-bearings 17 respectively connecting a blade 14 to the hub 13. Estimates of the three deflection angles of each spherical thrust-bearing 17 can be determined by means of the sensor 4 so as to determine the mechanical stresses experienced by each spherical thrust-bearing 17. Thereafter, damage to each spherical thrust-bearing 17 can be predicted in order to reduce the maintenance costs of the spherical thrust-bearings 17.

[0167] Naturally, the present invention may be subjected to numerous variations as to its implementation. Although several embodiments are described above, it should readily be understood that it is not conceivable to identify exhaustively all possible implementations. It is naturally possible to replace any of the means described with equivalent means without going beyond the ambit of the present invention.