FIELD OF THE INVENTION

The invention relates to the field of aircraft guidance, and more particularly the estimation of the alignment of an aircraft relative to a determined trajectory that does not require the use of an absolute satellite positioning system.

TECHNOLOGICAL BACKGROUND

The guidance systems of existing aircraft, and in particular of drones, make it possible to carry out autonomous guidance of an aircraft along a predefined trajectory, corresponding for example to the path of an observation mission. To perform such guidance, the position of the aircraft is determined at regular intervals and compared with a trajectory to be followed. This position is generally determined using a receiver of an absolute satellite positioning system, such as the GPS or Galileo systems.

It may nevertheless happen that the computer of the aircraft is unable to determine the current position of the aircraft, either due to a failure of a component of the aircraft, such as a GPS receiver, or due to unavailability of the signal from the positioning system, for example in the event of jamming of the latter. In this case, the computer cannot guide the aircraft so as to cause it to follow the predetermined trajectory. The aircraft therefore risks crashing at an unknown position and being lost.

To avoid this, the current position of the aircraft can be determined using another system on board the latter, such as an inertial unit continuously measuring the linear and angular accelerations of the aircraft. An integration of the signals provided by this inertial unit then makes it possible to determine the movements of the aircraft and therefore its relative position with respect to the last position provided by the satellite positioning system. Nevertheless, the uncertainty of the position thus determined can be high. The accumulation over time of the differences between the movement determined by integration and the real movement of the aircraft

indeed generates a drift of the position of the aircraft relative to its real position. Such a drift can reach several kilometers per hour of flight since the last position provided by the satellite positioning system.

It has therefore been proposed to additionally use the data measured by a distance meter on the ground in order to correct the position data supplied by the inertial unit and to deduce therefrom corrected position data compensating for the drift of the inertial unit. For this, the deviation meter is connected to a directional antenna of the ground station and is configured to continuously measure the direction in which the aircraft is located relative to a reference direction, for example north. Reference may in particular be made to document FR 3 033 924 in the name of the Applicant for more details on this system.

Such a deviation meter is however not systematically available.

There is therefore a need for a guidance method making it possible to safely guide an aircraft, autonomously, from a remote return point to an airport and to land the aircraft on a runway thereof, despite the unavailability of satellite positioning and despite a pronounced drift in the current position of the aircraft determined from the signals from its inertial unit.

SUMMARY OF THE INVENTION

An objective of the invention is to propose an alternative solution to the use of a ground deviation meter in order to allow the estimation of the position of an aircraft and its landing in a simple and efficient manner, despite the unavailability of positioning by satellite and despite a possible drift of the current position of the aircraft.

For this, the invention proposes an aircraft guidance system comprising:

- a first antenna having a first opening angle at - 3 dB,

- a second antenna having a second opening angle at -

the system being characterized in that:

- the first opening angle is at least twice as large as the second opening angle, and

- within the second opening angle of the second antenna, an absolute value of a difference between the power of a signal received from the first antenna and the power of a signal received from the second antenna is at least equal to 10dB.

Some preferred but non-limiting features of the guidance system described above are as follows, taken individually or in combination:

- the first antenna is omnidirectional.

- the first antenna and the second antenna are coaxial.

- the first opening angle is between 3° and 5° and the second opening angle is between 0.5° and 1.5°.

- the first antenna has a gain of between 25 dB and 35 dB, and the second antenna has a gain of between 35 dB and 50 dB.

- the first antenna and the second antenna are integral in movement, and the system further comprises means for moving the first antenna and the second antenna.

According to a second aspect, the invention proposes a method for autonomous guidance of an aircraft using a guidance system according to one of claims 1 to 6, said method comprising the following steps:

51: positioning of the first antenna and/or the second antenna so that a radioelectric axis of the first antenna and/or of the second antenna points to a supposed position of the aircraft,

52: measurement of a power of a signal received by the first antenna,

53: simultaneously, measurement of a power of a signal received by the second antenna,

54: determination of a difference between the power of the signal received by the first antenna and the power of the signal received by the second antenna, 55: deduction, from the difference thus determined, of a possible misalignment error between, on the one hand, the radioelectric axis of the first antenna and/or the second antenna and, on the other hand, the aircraft.

Certain preferred but non-limiting characteristics of the guidance method described above are the following, taken individually or in combination:

- the first antenna and the second antenna are coaxial during steps S1 to S3.

- the method further includes the following steps, prior to step S5:

56: angular displacement of the first antenna and of the second antenna according to a plurality of depointing angles then repetition of steps S2 to S4 for each depointing angle so as to determine, for each depointing angle, a corresponding deviation, and

57: evaluation of a maximum of the deviations thus obtained.

- During step S6 of movement, the first antenna and the second antenna perform an angular scan in azimuth and/or elevation according to a periodic pattern.

- Steps S6 and S7 are implemented only when the difference determined in step S5, from the assumed position of the aircraft, is less than a determined threshold.

the method further comprises, following step S7, a step of positioning the first antenna and the second antenna so as to substantially align their radioelectric axis with a direction corresponding to the maximum of the deviations thus determined.

- During step S6, the depointing angles are greater than or equal to the second opening angle and less than or equal to twice said second opening angle.

- during step S7, the maximum of the deviations is evaluated by a time convolution method or from a polynomial approximation of degree 2 of the measurements obtained in steps S2 and S3 and associating a given deviation with each depointing angle .

BRIEF DESCRIPTION OF THE DRAWINGS

Other characteristics, objects and advantages of the present invention will appear better on reading the detailed description which follows, and with regard to the appended drawings given by way of non-limiting examples and in which:

FIG. 1 schematically illustrates the transmission diagrams of an example of a first coaxial antenna and of an example of a second coaxial antenna that can be used in a guidance system according to the invention.

FIG. 2 schematically illustrates the transmission diagrams of another example of a first coaxial antenna and of an example of a second coaxial antenna that can be used in a guidance system in accordance with the invention.

FIG. 3 schematically illustrates the transmission diagrams of another example of a first coaxial antenna and of an example of a second coaxial antenna that can be used in a guidance system in accordance with the invention.

Figure 4 illustrates an example of measurement of a difference in dB between the power of the signal received by a first antenna and the power of the signal received by a second antenna, coaxial, pointing to the same radio transmitter and which can be used in a system guide according to the invention, during a sinusoidal sweep (in degrees) around a pointing angle in designation, as well as a polynomial approximation of degree 2 of said measurements.

FIG. 5 represents an example of difference between the power received by the first antenna and the power received by the second antenna used for FIG. 4, with the onset of drift of the aircraft at t=2000 s.

Figures 6 to 8 represent the result of the simulation of the example given when the goniometry algorithm is engaged, Figure 6 illustrating the power of the signals received from the first antenna and from the second antenna, Figure 7 representing the power difference between said

antennas in raw value (dRSSI) and in filtered value (dRSSMIt) and FIG. of filtering without calculation of correction and 2 being the phase of scanning and correction).

FIG. 9 very schematically illustrates an example of an aircraft guidance system according to one embodiment of the invention.

FIG. 10 is a flowchart illustrating the steps of an exemplary embodiment of a method for guiding an aircraft in accordance with the invention.

DETAILED DESCRIPTION OF AN EMBODIMENT

One embodiment of the invention relates to an autonomous guidance system 1 for an aircraft A comprising two antennas 10, 20 whose opening angle is chosen so as to allow, by a simple comparison of the power of their signal respective, to determine if the aircraft A is located in the expected direction or if on the contrary it has deviated from this direction and to determine, if necessary, the direction in which it is actually located by iterations.

More precisely, the guidance system 1 comprises a first antenna 10 having a first opening angle θ1 at −3 dB and a second antenna 20 having a second opening angle θ2 at −3 dB. The first opening angle θ1 is at least twice as large as the second opening angle θ2 and, within the second opening angle θ2 of the second antenna 20, the absolute value of the deviation (difference) between the power of a signal received from the first antenna 10 and the power of a signal received from the second antenna 20 is at least equal to 10 dB.

The first antenna 10 therefore has a large opening angle in comparison with the second antenna 20.

For example, the first opening angle θ1 can be between 3° and 5°, typically around 4°, while the second opening angle θ2 can be between 0.5° and 1.5°, typically the order of 1°.

The first antenna 10 can also have a gain of between 25 dB and 35 dB, for example of the order of 30 dB, and the second antenna 20 has a gain of between 35 dB and 50 dB, for example of the order of 40dB.

The guidance system 1 is based on the principle according to which an aircraft A is a radio transmitter so that, when a radio transmitter moves away from a receiving antenna, the power of the signal measured by this antenna decreases. However, it appears that, when a single antenna is used, the attenuation of the power of the signal measured by this antenna can also be due to a plurality of factors including an increase in the distance between the aircraft A and the antenna , a radio transmission problem, an antenna power defect, weather conditions, masking (presence of another radio transmitter between the aircraft A which is being guided and the antennas 10, 20), etc.

On the other hand, the difference between the power of the signals of two antennas 10, 20 pointing at the same radio transmitter remains constant, and this regardless of the distance between the radio transmitter and the two antennas 10, 20. Consequently, if the difference between the power of the signal measured by two given antennas 10, 20 is less than a given threshold or weakens, this necessarily implies that the aircraft A is not aligned with the antennas 10, 20.

The choice of a large-aperture antenna 10 and a small-aperture antenna 20 makes it possible to obtain a power difference sufficient to detect a misalignment of the aircraft A, the difference between the gain of the two antennas 10, 20 being marked so that the measurement accuracy is sufficient to guide the aircraft A. In addition, it makes it possible to perform sufficient angular scanning in the event that a weakening of the power difference is detected.

Reference may in particular be made to Figure 1, which schematically illustrates the emission diagrams of an example of a first antenna

10 and an example of a second antenna 20, which are coaxial. When the radio transmitter is aligned with the axis X1, X2 of the antennas 10, 20 (aircraft A1), the power difference E1 is maximum. On the other hand, when the radio transmitter is misaligned (aircraft A2), the power difference E2 between the two antennas 10, 20 is lower.

It will be understood that the use of the power difference for the guidance of the aircraft A makes it possible to make the determination of the power difference independent:

- the distance between the aircraft A and the antennas 10, 20, provided that the distance remains below a detection limit threshold.

- radio transmission.

- a power fault in the transmitting antenna.

- weather conditions.

- a possible masking.

- etc

Indeed, whatever the situation, the power of the signal received by each of the antennas 10, 20 weakens in a similar manner, so that the difference in power remains constant for the same position and the same pointing of the transmitter. radio in each of these situations.

Preferably, the first and the second antenna 10, 20 are coaxial in order to maximize the overlap of their angular aperture ranges. However, in one embodiment, the first and the second antenna 10, 20 may not be coaxial. In this case, the antennas 10, 20 are positioned so that the opening angle of the first antenna 10, which is large, overlaps the opening angle of the second antenna 20 (see Figure 2).

If necessary, the first antenna 10, which has a large opening angle θ1, can be omnidirectional. The second antenna 20 on the other hand is directional and adjustable (FIG. 3).

The first antenna 10 and the second antenna 20 are integral in movement. By integral in movement, it will be understood here that the first and the second antenna 10, 20 perform the same movements, simultaneously. For this, the first and the second antenna 10, 20 can be fixed solidly together, using a recessed connection, or be separate from each other but moved in a synchronized manner and following the same movements.

Furthermore, the guidance system 1 further comprises means 2 for moving the first and the second antenna 20.

Preferably, the first antenna 10 and the second antenna 20 are moved simultaneously, either by the same moving means 2, or by two separate but synchronized moving means 2.

The displacement means 2 can comprise, for example, positioners carrying antennas 10, 20, configured to receive pointing orders from a computer 6 (see below) and execute said orders.

The guidance system 1 can also comprise positioning means 3 configured to determine a supposed position of the aircraft A. These means 3 can for example comprise an inertial unit on board the aircraft A and configured to integrate the movements of aircraft A (acceleration and angular speed) to estimate its orientation (roll, pitch and heading angles), its linear speed and its position. To this end, conventionally, the inertial unit 3 comprises accelerometers for measuring the linear acceleration of the aircraft A in three orthogonal directions and gyrometers for measuring the three components of the angular velocity vector (roll, pitch and yaw). The inertial unit 3 also supplies the attitude of the aircraft A (roll angles,

As a variant, the positioning means 3 can comprise an absolute satellite positioning system, such as the GPS or Galileo systems.

Finally, the guidance system 1 comprises a system 4 for receiving signals from the first antenna 10 and from the second antenna 20 and a data processing device 5, 6.

The data processing device 5, 6 can be on board the aircraft A and/or in a ground unit and can comprise one or more communication interfaces 4 and one or more computers 5, 6. For example, the The ground unit and the aircraft A can communicate by radio and each comprise a communication interface 4 of the antenna type. In one embodiment, the data processing device 5, 6 comprises an on-board computer 5, connected to the means for determining the assumed position of the aircraft A, and a ground computer 6.

Each computer 5, 6 can comprise a processor or microprocessor, of the x-86 or RISC type for example, a controller or microcontroller, a DSP, an integrated circuit such as an ASIC or a programmable circuit such as an FPGA, a combination of such elements or any other combination of components making it possible to implement the calculation steps of the guidance method. As we will see later, the computer on the ground 6 can be configured to transmit pointing orders to the displacement means 2, such as positioners, from the positioning information communicated by the positioning means 3, such as an inertial unit, but also an angular pointing error corresponding to the position drift of the aircraft A, a scanning angle in order to create a depointing of the antennas 10,

The communication interfaces 4 for their part can be any interface, analog or digital, allowing the computer(s) to exchange information with the other elements of the guidance system 1 such as the antennas 10, 20 , the displacement means 2 or even the positioning means 3. The communication interfaces can for example comprise an RS232 serial interface, a USB, Firewire, HDMI interface or an Ethernet type network interface.

The determination of the power difference between the first antenna 10 and the second antenna 20 thus makes it possible to correct the drift, even pronounced, of the current position of the aircraft A determined from the signals from its inertial unit 3 (or from any other means of determining the assumed position of the aircraft A) by determining whether the position of the aircraft A actually corresponds to the assumed position, or if it is misaligned with respect to this assumed position

The aircraft A can then be guided according to the following steps, using the guidance system 1 previously described.

During a preliminary step S0 of the guidance method S, an assumed position of the aircraft A is determined.

For example, the assumed position of aircraft A can be determined conventionally by the inertial unit 3 on board aircraft A.

This is however not limiting, the inertial unit 3 possibly being optional. The assumed position of aircraft A can be determined by any means 3. For example, the assumed position of aircraft A can be determined from the last known position of aircraft A, measured by an absolute positioning system via satellite, such as a GPS or Galileo system.

During a first step S1, the first antenna 10 and/or the second antenna 20 are pointed at the assumed position thus determined of the aircraft A (pointing in designation).

For this, the first antenna 10 and the second antenna 20 are moved so that their respective radioelectric axes X1, X2, which are preferably coaxial, intersect the supposed position of the aircraft A.

During a second and a third step S2, S3, the power of the signals received by the first antenna 10 and by the second antenna 20 is measured simultaneously. The power of the signals can in particular be measured in dBm.

During a fourth step S4, the difference between the power of the signal received by the first antenna 10 and the power of the signal received by the second antenna 20 is determined by the data processing device 5, 6, and in particular the ground computer 6.

In the case of a weakening of the power difference with respect to an expected power difference, the computer 5 can send displacement commands to the displacement means 2, for example to positioners carrying the first antenna and the second antenna 20, so as to move them angularly (step S6) according to a plurality of depointing angles and to point their radioelectric axis X1, X2 to a position different from the assumed position established during the preliminary step S0.

During step S6, the first and the second antenna 10, 20 are moved angularly in azimuth and/or in elevation.

Optionally, during step S6, at each movement, the depointing angle at which the first and second antennas 10, 20 are moved is greater than or equal to the second aperture angle θ2 and less than or equal to twice said second opening angle 02.

Steps S2 to S6 are then repeated until the power difference is maximum, or at least reaches a predefined threshold value corresponding to an admissible alignment between the radioelectric axis X1, X2 of the antennas 10, 20 and the aircraft A. The pointing associated with the maximum power deviation then roughly indicates the direction of aircraft A.

As a variant, during step S6, the first antenna 10 and the second antenna 20 can carry out a sweep according to a predefined pattern, the maximum power deviation then being determined from the different power deviations determined for each angle of depointing of the angular scan in order to deduce the direction of the aircraft A.

Preferably, the angular scan is performed in a periodic pattern.

For example, FIG. 4 illustrates an example of measurement of a difference between the power of the signal received by a first antenna 10 and the power of the signal received by a second antenna 20, coaxial and pointing at the same radio transmitter during of a sinusoidal sweep around a pointing angle in designation.

In another example, the scanning can follow a Lissajous curve in order to allow a good population of the ends of the scanned angular zone while ensuring a cross passage on the main lobe of the signals from the first and from the second antenna 10, 20.

In order to determine the pointing correction angle and to align the radio axis X1, X2 of the antennas 10, 20 with the direction of the aircraft A, in a first embodiment, the ground computer 6 (or any another processing device) can for example establish a polynomial approximation of degree 2 (parabolic regression) of the measurements in the sense of the least squares which connects the scanning angle (on the abscissa, corresponding to the angle between the measurement direction during the scanning and the supposed position of the aircraft A determined during the preliminary step S0) at the deviation of the powers obtained at the step S3. If necessary, the least squares can optionally be weighted in order to take into account the degree of confidence associated with each measurement.

The pointing correction angle is then obtained by determining the abscissa of the maximum of the polynomial of degree 2 thus established (step S7).

If necessary, to ensure the robustness of the algorithm:

- the correction can be limited to the maximum amplitude of the sweep and/or - the correction can be filtered over several sweep periods using a low-pass filter (for example a Kalman filter) whose constant time can for example be set to a quarter of the scanning period and/or

- the sweep can be triggered on a filtered power deviation criterion and/or

- the scan can be terminated on a convergence criterion of the correction.

Once the scan has been performed and the pointing angular error, representing the position drift of the aircraft A and/or a possible initial pointing error due to incorrect initial setting of the orientation of the carrier positioners 2 on the ground, determined , the antennas 10, 20 are moved by the carrier positioners 2 (or any other suitable moving means) so as to point to the real position of the aircraft A thus identified.

This first embodiment makes it possible to determine the pointing correction angle. However, the presence of secondary lobes in the signals from the antennas 10, 20 (see for example FIG. 1) is likely to raise the level of the signals measured at the level of the edges of the main lobes and can produce correction calculations in the opposite direction in because of the regression calculation that generates convex and non-concave solutions. The maximum then becomes a minimum.

In the case where the solution is convex, the computer 6 can apply a linear regression then select the maximum of this linear regression over the scanning interval. Furthermore, the scanning amplitude can be chosen as a function of the filtered value of the difference in powers, so that the amplitude is all the greater as the filtered value is low.

In a second embodiment, which can be combined with the first embodiment, the maximum of the power difference is determined by a time convolution method (step S7). If necessary, this embodiment makes it possible to take account of the delay in the control of the displacement means 2 by introducing a delay time in azimuth and in elevation, which makes it possible to associate the difference in powers with the misalignment actually applied to the antennas 10, 20.

This time convolution method, which is used for calculating the position of the signal maximum, is based on the following assumptions:

- for low depointing angles, the gain of the antenna with greater gain behaves like a paraboloid of revolution.

- for low depointing angles, the gain of the lower gain antenna is considered constant.

- the scanning on each of the axes (vertical and horizontal) is of sinusoidal form of the type s c = a c sin (h c wί), where x represents either the azimuth axis or the elevation axis. This scan is applied to each axis over a period of T=2pIw.

The ARSSI power difference can then be modeled as a paraboloid of revolution by the following formula:

where: ARSSI is the deviation of the powers measured in steps S2 and S3

e Az and e EI are the angular errors in azimuth and elevation, respectively

s Az and s Eί are the azimuth and elevation scan angles, respectively

Q is the half-angle of opening at -3 dB of a fictitious paraboloid antenna whose gain profile corresponds to the difference between the gain profiles of the antennas 10, 20.

The time convolution method then consists in calculating, over a scanning period of T= 2 p/w, the following quantities on each of the scanning axes (i.e. in azimuth and elevation):

e Az = J o ARSSl X a Az dt and C El = J Q ARSSI X a El dt

The angular error e Az (in azimuth) and the angular error e (in elevation) are therefore proportional to the convolution product of the deviation of the ARSSI powers and the scanning pattern on the corresponding axis :

e 2 c Az

£AZ 3CL 2 Az T

and

Note that this result is valid for {n Az , n El } e N and n Az ¹ n El and that a 2 Az and a 2 m correspond to the amplitudes of the sweeps defined above (s c = a c sin (h c wί)).

If necessary, step S6 can only be implemented when the power difference is less than a predetermined threshold.

Example :

A simulation was carried out with real data recorded during the flight of an aircraft A. During this flight, an elongation of 140 km was reached, the wearer was moving at 36 m/s following a straight trajectory (phase 1 ) then orbital or spiral approaching the ground station (phase 2).

The first antenna 10 had a gain of 30 dB and a first opening angle 01 at - 3 dB of 0.9°, while the second antenna 20 5 had a gain of 44 dB and a second opening angle 02 at - 3 dB of 4.0.

The powers received on the two antennas 10, 20 (44dB dish and 30dB patch) have been artificially degraded by simulating a depointing angle. The power signal (RSSI) of each antenna has been corrected io as follows:

where: Oi is the opening angle at -3 dB of antenna i (first antenna 10 or second antenna 20)

θi is the offset angle of the antenna i (first antenna 10 or second antenna 20) with respect to the radioelectric axis Xi (X1 or X2).

The drift was simulated by a drift velocity of 2 m/s normal to the pointing axis, which is a pessimistic case. The antenna drift angle was simulated by:

where: DTLS VA is the distance between the antennas 10, 20 on the ground and the aerial vehicle. To accentuate the drift effect in value and speed, the actual distance of the test serving as the basis for the simulation was artificially reduced by 25 km in the simulation.

25 tstart is the simulation drift start time (2000 s here)

Vdrift is the drift speed in m/s (2 m/s here)

The noise on the gain measurement was set at 3 dB and the model describing the gain variation around the maximum was a parabola. The mismatch was 0.8°.

30 The angular displacement of the aircraft was measured every 14 s.

Figure 5 shows the shape of the powers received in the absence of a direction-finding algorithm for a start of drift at t=2000 s. It can be seen in this figure that the signal difference (on the ordinate) between the two antennas 10, 20 is gradually reduced when the misalignment exceeds the first opening angle θ1.

The curves in Figures 6 to 8 show the result of the simulation with a direction finding algorithm in accordance with the first embodiment engaged where:

0 corresponds to the preparation phase after resetting the power difference filtering (follows the output of a scan sequence)

1 corresponds to the filtering phase without correction calculation

2 corresponds to the scanning and correction phase

For this simulation, the scan is performed with a period of seven seconds in order to accumulate enough measurement points on a scan to make an accurate correction calculation and to be sure not to exceed the speed capabilities of the positioner.

When the drift speed reaches approximately 4° per minute, the power difference between the two antennas 10, 20 cannot be maintained at the highest level, which translates into dragging causing the aircraft to leave the effective zone of the main lobe of the second antenna 20.

CLAIMS

1. Aircraft guidance system (A) comprising:

- a first antenna (10) having a first opening angle

(01) to -3dB,

- a second antenna (20) having a second opening angle (θ2) at -3 dB,

the system (1) being characterized in that:

- the first opening angle (01) is at least twice as large as the second opening angle (02), and

- within the second opening angle (θ2) of the second antenna (20), an absolute value of a difference between the power of a signal received from the first antenna (10) and the power of a signal received of the second antenna (20) is at least equal to 10 dB.

2. System (1) guidance according to claim 1, wherein the first antenna (20) is omnidirectional.

3. System (1) according to one of claims 1 or 2, wherein the first antenna (10) and the second antenna (20) are coaxial.

4. System (1) according to one of claims 1 to 3, wherein the first opening angle (01) is between 3° and 5° and the second opening angle (02) is between 0.5° and 1.5°.

5. System (1) according to one of claims 1 to 4, wherein the first antenna (10) has a gain of between 25 dB and 35 dB, and the second antenna (20) has a gain of between 35 dB and 50dB.

6. System (1) according to one of claims 1 to 5, wherein the first antenna (10) and the second antenna (20) are integral in

movement, and the system (1) further comprises means for moving (2) the first antenna (10) and the second antenna (20).

7. Method (S) for autonomous guidance of an aircraft (A) using a guidance system (1) according to one of claims 1 to 6, said method

(S) comprising the following steps:

51: positioning of the first antenna (10) and/or the second antenna (20) so that a radioelectric axis (X1, X2) of the first antenna (10) and/or of the second antenna (20) points on an assumed position of the aircraft (A),

52: measurement of a power of a signal received by the first antenna

(10),

53: simultaneously, measurement of a power of a signal received by the second antenna (20),

S4: determination of a difference between the power of the signal received by the first antenna (10) and the power of the signal received by the second antenna (20),

S5: deduction, from the difference thus determined, of a possible misalignment error between, on the one hand, the radioelectric axis (X1, X2) of the first antenna (10) and/or the second antenna and , on the other hand, the aircraft.

8. Method (S) according to claim 7, in which the first antenna (10) and the second antenna (20) are coaxial during steps S1 to S3.

9. Method (S) according to one of claims 7 or 8, further comprising the following steps, prior to step S5:

S6: angular displacement of the first antenna (10) and of the second antenna (20) according to a plurality of depointing angles then repetition of steps S2 to S4 for each depointing angle so as to determine, for each depointing angle, a corresponding gap, and

S7: evaluation of a maximum of the deviations thus obtained.

10. Method (S) according to claim 9, in which, during step S6 of displacement, the first antenna (10) and the second antenna (20) carry out an angular scan in azimuth and/or in elevation according to a pattern periodic.

11. Method (S) according to one of claims 9 or 10, in which steps S6 and S7 are implemented only when the deviation determined in step S5, from the assumed position of the aircraft (A), is below a determined threshold.

12. Method (S) according to claim 10, further comprising, following step S7, a step of positioning the first antenna (10) and the second antenna (20) so as to substantially align their radioelectric axis ( X1, X2) with a direction corresponding to the maximum of the deviations thus determined.

13. Method (S) according to one of claims 9 to 12, in which, during step S6, the depointing angles are greater than or equal to the second opening angle (θ2) and less than or equal to twice said second opening angle (02).

14. Method (S) according to one of claims 9 to 13, in which, during step S7, the maximum of the deviations is evaluated by a time convolution method or from a polynomial approximation of degree 2 measurements obtained in steps S2 and S3 and associating a given deviation with each depointing angle.