Aircraft monitoring method and system

Abstract

A method of monitoring an aircraft comprising a structural component (2); the method comprises: - associating at least one optical sensor with the structural component (2), - detecting a temperature and a deformation of the structural component (2) by means of the optical sensor, - thermally compensating for the detected deformation as a function of the detected temperature, - defining a significant numerical model of the structural component (2), - selecting a sensor data correlation algorithm to determine an aircraft condition as a function of the defined numerical model and the compensated deformation of the structural component (2), - calculating a plurality of information items relating to the aircraft condition by means of the sensor data correlation algorithm.

Description

[0001] DESCRIPTION

[0002] AIRCRAFT MONITORING METHOD AND SYSTEM

[0003] Technical field

[0004] The present invention relates to an aircraft monitoring method and system. In particular, the invention relates to the field of aeronautics, mainly referring to drones and unmanned vehicles, and of aviation in general.

[0005] In such a context, it is necessary to collect large amounts of data, which must be as reliable as possible both to ensure the proper operation of the aircraft and to detect any faults.

[0006] Background art

[0007] In known systems, the collection of these data occurs by means of multiple sensors. However, the increase in the number of sensors and components associated therewith causes an increase in the weight of the aircraft, with resulting disadvantages in terms of performance and efficiency.

[0008] Document US2018100776A1 describes a device for determining flight parameters of an aircraft based on the measurement of the local deformation of the radome. The solution involves the use of a complex pattern of optical fibers provided with Bragg gratings arranged along the fiber so as to be distributed on the radome once the fiber is installed on the aircraft. FBG (Fiber Bragg Grating) fibers allow measuring the local deformation of the radome, the values of which are processed to determine aerodynamic parameters, in particular speed, angle of attack, and leeway angle.

[0009] US9073623B1 describes the use of FBG sensors for the real-time monitoring of structural aspects of an airplane wing.

[0010] Therefore, a need is felt for a functional redundancy which allows maintaining control of the attitude and performance of the aircraft without significantly affecting its weight.Object of the invention

[0011] It is the object of the present invention to meet such a need by providing a method and system for monitoring the aircraft capable of ensuring data redundancy, even in case of fault of the conventional instrumentation, without a significant increase in weight.

[0012] At least said object is achieved by the present invention, which is characterized by the contents of the independent claims.

[0013] The dependent claims correspond to possible different embodiments of the invention.

[0014] According to an aspect of the description, the description relates to a method of monitoring an aircraft comprising a component.

[0015] The method comprises arranging at least one optical sensor at a component of the aircraft.

[0016] Preferably, the component is a structural component of the aircraft.

[0017] In particular, the component is a structural component designed to withstand aerodynamic loads necessary for flight.

[0018] In other words, the structural component is a component of the aircraft designed to withstand aerodynamic loads which are decisive for flight. In an example, the structural component is an aerodynamic attachment of the aircraft.

[0019] In an example, the structural component is a wing or half-wing of the aircraft. In particular, wing means the set of the two half-wings.

[0020] In an example, the structural component is a fuselage or part of a fuselage of the aircraft.

[0021] In an example, the optical sensor is an FBG (Fiber Bragg Grating) sensor. Advantageously, the FBG sensors are resistant to liquids, electromagnetic fields, and wide temperature ranges, not requiring waterproofing, shielding, or insulation.

[0022] Advantageously, the FBG sensors thus allow the distributed provision of sensors on exposed parts.

[0023] Advantageously, the FBG sensors ensure an advantage in terms of weightand overall size.

[0024] The method comprises detecting a temperature and a deformation of the structural component by means of the optical sensor.

[0025] The method comprises thermally compensating for the detected deformation as a function of the detected temperature.

[0026] The method comprises defining a numerical model of the structural component.

[0027] The numerical model is derived from a parametric identification performed on a physical reference system representative of the structural component. The numerical model can be calibrated upstream by comparing the compensated deformation of the component with the calculated deformation of the structural component.

[0028] Thermal compensation is thus performed for each deformation measurement.

[0029] The method comprises selecting a sensor data correlation algorithm (e.g., model-based) to determine an aircraft condition as a function of the numerical model and the compensated deformation of the structural component.

[0030] The method comprises calculating, by means of the sensor data correlation algorithm, a plurality of information items relating to the aircraft condition. The aircraft condition is described by kinematic and dynamic magnitudes, such as: position (coordinates in the reference system), linear velocities (in the three body axes of the aircraft, x-longitudinal, y-lateral, z-vertical), angular rates (pitch, roll, and yaw rates), orientation (attitude angles), and linear and angular accelerations.

[0031] Other calculated information relating to the aircraft can be angles of attack, thermal anomalies, and mechanical failures.

[0032] Advantageously, the correlation algorithm permits correlating the information of the sensor with the numerical model thus permitting the characterization of the aircraft by calculating physical parameters which allow calculating angular rates, accelerations, and load factors.According to an aspect, the method comprises defining a load factor of the aircraft as a function of the plurality of calculated information items relating to the aircraft condition and of a weight of the aircraft.

[0033] The load factor is not directly part of the aircraft condition, but is a magnitude derived from the aircraft condition and the aerodynamic forces.

[0034] The load factor can be calculated from the accelerations and attitude of the aircraft.

[0035] In particular, the acceleration along the vertical axis z, measured in g, is the load factor.

[0036] As already mentioned above, the optical sensor permits detecting the deformation of the structural component.

[0037] In other words, the optical sensor detects wavelengths which are converted into deformation values, e.g., by means of calibration algorithms and / or model-based and data-driven correlation.

[0038] Advantageously, performing a thermal compensation ensures that the detected deformation data reflect only mechanical stresses but not temperature variations.

[0039] Advantageously, in other words, performing a compensation improves the measurement accuracy.

[0040] The deformed shape of the wing can be estimated by analyzing the distribution of deformations along the wingspan.

[0041] Similarly, the distribution of lift produced by the wing can be estimated by means of empirical correlations to the macroscopic deformation that it produces.

[0042] Using the lift estimate under different flight conditions and the distribution of the lift itself, it is possible to determine with precision the weight of the aircraft during flight, and also to estimate the weight distribution of the vehicle, in order to monitor weight and balancing.

[0043] According to an aspect, calculating a plurality of information items relating to the aircraft condition by means of the defined algorithm comprises calculating a lift of the aircraft and / or an acceleration of the aircraft and / oran angular rate of the aircraft and / or an angle of attack of the aircraft.

[0044] The angular rate relative to z is the so-called yaw rate.

[0045] The angular rate relative to x is the so-called roll rate.

[0046] The angular rate relative to y is the so-called pitch rate.

[0047] According to an aspect, defining a significant numerical model of the structural component of the aircraft comprises defining a structural model of the structural component of the aircraft.

[0048] In other words, the numerical model represents, for example, the mechanical behavior of the structural component and depends on the geometry thereof.

[0049] According to an aspect, the method comprises providing an aircraft navigation system for detecting a plurality of aircraft navigation information items and comparing said plurality of aircraft navigation information items with the plurality of calculated aircraft condition information items.

[0050] In an example, the navigation system comprises an inertial platform (INS). In an example, the navigation system comprises an inertial measurement unit (IMU).

[0051] In an example, the aircraft comprises onboard equipment which comprise angle-of-attack sensors.

[0052] Generally, the navigation system can vary depending on the physical model of the considered aircraft.

[0053] Therefore, the devices of the navigation system can vary as a function of the physical model of the considered aircraft.

[0054] In an example, the navigation system comprises a GPS.

[0055] According to an aspect, the method comprises filtering the plurality of navigation information items before comparing said plurality of aircraft navigation information items with the plurality of aircraft condition information items.

[0056] Filtering the information using, for example, a slope-limiting algorithm permits smoothing data by reducing sudden variations due to disturbances. For example, for the accelerations along the vertical axis, it is possible toperform a filtering to reduce the effects due to vibrations.

[0057] In an example, the method comprises combining the plurality of navigation information items with the load factor by means of a data fusion algorithm. According to an aspect, the sensor data fusion algorithm is a complementary filter or a Kalman filter.

[0058] The Kalman filter permits selecting a minimum sensor configuration to obtain a completely optical estimate, independent of the inertial platform. The Kalman filter is an algorithm based on numerical models which continuously updates a condition estimate by combining the measurements with a predictive model.

[0059] The Kalman filter permits modulating the weight assigned to each data source, optimizing the final estimate of the load factor and reducing the effects of disturbance sources (e.g., vibrations, noise, disturbances not included in the measurement).

[0060] The complementary filter is a simpler algorithm which combines two signals with different response characteristics (e.g., Inertial Navigation System + GPS), using frequency fusion (e.g., weighted average).

[0061] Both serve to improve the measurement accuracy.

[0062] Advantageously, using a Kalman filter algorithm permits combining different types of data and improving accuracy.

[0063] Both the Kalman filter and the complementary filters are algorithms used for data fusion from different sensors, which allow improving the result estimates by reducing noise and errors.

[0064] Advantageously, the suggested method according to the present description permits obtaining a functional redundancy, having available the comparison of sensor data and data calculated by means of the algorithms. This large amount of obtained information and the comparison thereof is of crucial importance for maintaining control of the attitude and overall performance of the craft.

[0065] Comparing, by means of the data fusion algorithm, the information relating to the aircraft condition obtained by virtue of the correlation algorithm, withthat obtained from the navigation system improves the system reliability and accuracy, permitting faults in the avionics, equipment, or sensors to be detected, and possibly the information of the faulty devices to be replaced by the information calculated by the FBG-numerical model combination. According to an aspect of the description, the description relates to a system for monitoring an aircraft comprising a component.

[0066] Preferably, the component is a structural component of the aircraft.

[0067] In particular, the component is a structural component designed to withstand aerodynamic loads necessary for flight.

[0068] In other words, the structural components are those subject to the forces which enable flight, e.g., wings, tail surfaces, tail boom. The landing gear can also be considered as a structural component, because it supports loads, e.g., during landing.

[0069] The monitoring system comprises at least one optical sensor associated with the structural component and configured to detect a deformation of the structural component.

[0070] Preferably, the optical sensor is configured to detect a temperature of the structural component.

[0071] The monitoring system comprises a control unit in communication with the optical sensor.

[0072] The control unit is configured to calculate a thermal compensation as a function of the temperature and deformation detected to obtain a compensated deformation.

[0073] The control unit is configured to calculate a plurality of information items relating to the aircraft condition by means of the sensor data correlation algorithm.

[0074] The sensor data correlation algorithm is defined as a function of a significant numerical model of the structural component and of the compensated deformation of the structural component.

[0075] The structural component is, for example, a wing of the aircraft.

[0076] According to an aspect, the structural component comprises an outercasing.

[0077] In an example, the optical sensor is arranged inside the outer casing.

[0078] In an example, the outer casing comprises laminated layers and the optical sensor is arranged between the layers of the outer casing.

[0079] According to an aspect, the system comprises an adhesive layer between the optical sensor and the structural component.

[0080] In a preferred example, the optical sensor is of the FBG type.

[0081] According to an aspect, the system comprises a plurality of optical sensors. For example, the FBG sensors can be inserted into different optical fiber lines.

[0082] According to an aspect, the monitoring system comprises an aircraft navigation system configured to detect at least one aircraft navigation information item.

[0083] The navigation system is in communication with the control unit.

[0084] The control unit is configured to receive and process the aircraft navigation information.

[0085] The control unit is configured to receive the aircraft navigation information and combine the aircraft navigation information with said plurality of pluralities of information items relating to the aircraft condition by means of a sensor data fusion algorithm.

[0086] Advantageously, the system permits extracting additional information which increases data reliability and quality with respect to the use of conventional sensing components alone.

[0087] The monitoring system and method allow measuring an angular rate (speed) relative to axis x.

[0088] The monitoring system and method allow measuring an acceleration along axis x.

[0089] The monitoring system and method allow measuring an angular rate (speed) relative to axis y.

[0090] The monitoring system and method allow measuring an acceleration along axis y.The monitoring system and method allow measuring an angular rate (speed) relative to axis z.

[0091] Advantageously, the monitoring system and method allows, in practice, measuring an angular rate (speed) and an acceleration for each of the 3 body axes x, y, z of the aircraft in the 3 main directions where, as mentioned, the acceleration along the vertical axis z, measured in g, is the load factor, the acceleration along axis x is the longitudinal acceleration, the acceleration along axis y is the lateral acceleration.

[0092] The angular rate relative to z is the yaw rate, the angular rate relative to x is the roll rate, the angular rate relative to y is the pitch rate.

[0093] These six outputs (load factor, longitudinal acceleration, lateral acceleration, yaw rate, roll rate, and pitch rate) are the typical outputs of a modem IMU.

[0094] Preferably, a minimum sensor map for measuring rate and acceleration comprises two optical sensors per axis.

[0095] Advantageously, it is possible to obtain a measurement and / or estimation of the three rates and the three accelerations with six sensors.

[0096] Indeed, accelerations and angular displacements (rotations) result in a mechanical effect on the aircraft structure causing deformations of the structural components forming it.

[0097] In practice, aircraft structure means the structure designed to withstand the aerodynamic loads which are decisive for flight.

[0098] In particular, accelerations and angular displacements determine deformations in the structural components designed to withstand aerodynamic loads necessary for flight.

[0099] A combined use of two sensors permits discriminating between two effects (by virtue of the targeted positioning thereof), permitting an estimate of the magnitudes of interest to be obtained (which can be all the more refined and reliable as the number of sensors increases and as the positioning thereof varies).Advantageously, such a system thus allows a redundancy of data with respect to conventional systems by relying on indirect measurements, which are thus not subject to the same failure modes.

[0100] Therefore, the redundancy is not only due to the presence of multiple devices and to multiple data but also to the functional differences of the devices used and to the different types of data.

[0101] Advantageously, the system is minimally invasive by virtue of the reduced complexity of the sensor-provision process.

[0102] Advantageously, such a system permits obtaining data redundancy in a more economical manner as compared to more complex and sophisticated sensor systems.

[0103] Advantageously, such a system allows carrying out a real-time monitoring by providing immediate feedback during tests and all the steps of flight. Advantageously, the system according to the aspects described above permits an increase in data accuracy and amount of information, limiting the addition of hardware, being compatible with the geometries of the aircraft and the requirements to limit its weight.

[0104] Advantageously, the method and system described above permit detecting failures of onboard electronic apparatuses (e.g., such as sensors, equipment), and are suggested as a data source under emergency condition in which said apparatuses are unavailable or not correctly functioning.

[0105] Advantageously, the method and system described permit increasing the performance of the inertial platform, providing greater precision, robustness, and reliability of the navigation data.

[0106] Structural deformations, in particular those of the half-wings, are considered as a data source for processing a more complex system information content, comprising information on navigation, aircraft attitude, and contingent load factor.

[0107] Brief description of the drawingsThe main features of the invention will become more apparent from the following detailed description, given with reference to the accompanying drawings, which depict a mere exemplary and not-limiting embodiment thereof, in which:

[0108] - figure 1 shows a perspective view of part of a monitoring system in accordance with the present description according to a first configuration; - figure 2 shows a perspective view of part of a monitoring system in accordance with the present description according to a second configuration;

[0109] - figure 3 shows a perspective view of part of a monitoring system in accordance with the present description according to a third configuration; - figure 4 shows a diagrammatic top plan view of an example of application of a monitoring system according to the present description;

[0110] - figure 5 shows a diagrammatic bottom plan view of the example in figure 4.

[0111] Detailed description of preferred embodiments of the invention With reference to the accompanying drawings and with particular reference to figure 1 , reference numeral 1 indicates a system for monitoring an aircraft 100 comprising a structural component 2.

[0112] The structural component 2, as shown in figures 1-3, is a wing or half-wing of the aircraft, in a preferred embodiment.

[0113] In an embodiment, the structural component can be a fuselage 101 or a part of the fuselage 101 of the aircraft 100.

[0114] The structural component 2 comprises an outer casing 4.

[0115] The system 1 comprises an optical sensor 3 associated with the structural component 2.

[0116] In an embodiment, shown for example in figure 1 , the sensor 3 is arranged inside the outer casing 4.

[0117] In an embodiment, the outer casing 4 comprises laminated layers 4’, 4”. In an embodiment, as shown in figure 2, the sensor 3 is arranged betweenthe laminated layers 4’, 4” of the outer casing 4.

[0118] In an embodiment, as shown in figure 3, the sensor 3 and the outer casing 4 are attached by an adhesive layer 5.

[0119] In an embodiment, the adhesive layer is identified by a support matrix. The optical sensor 3 is configured to detect a deformation of the structural component 2.

[0120] The optical sensor 3 is configured to detect a temperature of the structural component 2.

[0121] In a preferred embodiment, the optical sensor 3 is an FBG sensor.

[0122] For example, the optical sensors 3 can be inserted in different optical fiber lines L1, L2.

[0123] The sensor 3 detects wavelengths which are converted into deformation values by means of appropriate algorithms.

[0124] The system 1 comprises a control unit (not shown) in communication with the sensor 3.

[0125] The control unit is configured to calculate a thermal compensation as a function of the temperature and the detected deformation to obtain a compensated deformation.

[0126] The control unit is configured to calculate a plurality of information items relating to the aircraft condition by means of a sensor data correlation algorithm.

[0127] Information relating to the aircraft condition means, for example, information relating to the position, attitude, and altitude.

[0128] The sensor data correlation algorithm is defined as a function of a significant numerical model of the structural component and of the detected deformation of the structural component.

[0129] The numerical model represents, for example, the mechanical behavior of the structural component 2.

[0130] In a preferred embodiment, the monitoring system 1 comprises an aircraft navigation system (not shown) configured to detect at least one aircraft navigation information item.In an embodiment, the navigation system comprises a Global Positioning System GPS.

[0131] In an embodiment, the navigation system comprises an Inertial Measurement Unit I MU.

[0132] In an embodiment, the navigation system comprises an Inertial Navigation System INS.

[0133] In an embodiment, the navigation system comprises angle-of-attack sensors.

[0134] The navigation system is in communication with the control unit.

[0135] The control unit is configured to receive and process the aircraft navigation information.

[0136] The control unit is configured to combine at least the aircraft navigation information with the plurality of information items relating to the aircraft condition by means of a sensor data fusion algorithm.

[0137] In an embodiment, the sensor data fusion algorithm is a complementary filter.

[0138] In an embodiment, the sensor data fusion algorithm is a Kalman filter. The description is directed to a method of monitoring an aircraft comprising a structural component 2.

[0139] The method comprises associating at least one optical sensor 3 with the structural component 2.

[0140] The method comprises detecting a temperature and a deformation of the structural component 2 by means of the optical sensor 3.

[0141] The method comprises thermally compensating for the detected deformation as a function of the detected temperature.

[0142] The method comprises defining a significant numerical model of the structural component 2.

[0143] The model is representative of the behavior of the structural component 2. The method comprises selecting a sensor data correlation algorithm to determine an aircraft condition as a function of the numerical model and the compensated deformation of the structural component 2.The method comprises calculating, by means of the sensor data correlation algorithm, a plurality of information items relating to the aircraft condition. In an embodiment, the method comprises defining a load factor of the aircraft as a function of the plurality of calculated information items relating to the aircraft condition and of a weight of the aircraft.

[0144] In an embodiment, calculating a plurality of information items relating to the aircraft condition by means of the defined algorithm comprises calculating a lift of the aircraft and / or an acceleration of the aircraft and / or an angular rate of the aircraft and / or an angle of attack of the aircraft.

[0145] In an embodiment, defining a significant numerical model of the structural component 2 of the aircraft comprises defining a structural model of the structural component 2 of the aircraft.

[0146] In an embodiment, the method comprises providing a navigation system of the aircraft to detect a plurality of navigation information items of the aircraft and comparing said plurality of navigation information items of the aircraft with the plurality of calculated information items relating to the aircraft condition.

[0147] In an embodiment, the method comprises filtering the plurality of navigation information items before comparing said plurality of navigation information items of the aircraft with the plurality of information items relating to the aircraft condition.

[0148] In an embodiment, the method comprises combining the plurality of navigation information items with the load factor by means of a data fusion algorithm.

[0149] In an embodiment, the sensor data fusion algorithm is a complementary filter or a Kalman filter.

[0150] The data read by the FBG sensors are integrated by means of the numerical model permitting a plurality of information items related to the aircraft condition to be calculated.

[0151] The information is further compared with the data provided by the navigation apparatuses by means of a suitable fusion or data sensor fusion algorithmto improve the robustness of the information, detect fault conditions, replace missing information due to faults / hostile operating environments.

[0152] Examples of calculation of vertical acceleration, roll rate, and pitch rate, verified by the Applicant, are described below with reference to figures 4 and 5.

[0153] The structural component 2 is a wing of the aircraft 100 and comprises a first half-wing 2s and a second half-wing 2d.

[0154] Example 1 (figure 4).

[0155] Consider a pair of optical sensors 3as and 3ad, placed on the wing 2 at a height yo from the root of the first half-wing 2s and from the root of the second half-wing 2d, respectively.

[0156] The optical sensors are positioned as close as possible to the torsional center / point of minimum sensitivity to the effects of the flexural-torsional coupling of the structure of the aircraft 100.

[0157] The optical sensors 3as, 3ad can be considered sensitive only to deformations due to loads which induce a bending moment on the wing 2, and in particular to vertical acceleration (load factor) and to the roll moment (and therefore roll rate).

[0158] Vertical acceleration leads to a symmetrical deformation contribution on the two half-wings 2s, 2d.

[0159] Vice versa, the roll moment leads to an antisymmetric contribution on the two half-wings 2s, 2d.

[0160] The Applicant has mathematically demonstrated the correlation between the Bragg wavelengths reflected by the optical sensors 3as, 3ad, the mechanical deformations generated at the aforesaid optical sensors, and the kinematic magnitudes of the aircraft 100 considered here.

[0161] In particular, the following applies:

[0162] £3ad=£n + £roll

[0163] £3as — £n - £ o / /

[0164] where:£3ad : mechanical deformation of the point in which sensor 3ad is present; £3as : mechanical deformation of the point in which sensor 3as is present; £n: mechanical deformation at the measurement point induced by the variation of the load factor;

[0165] £roii : mechanical deformation at the measurement point induced by the variation of the roll rate.

[0166] From which it follows that, adding and subtracting the deformations measured by the respective optical sensors 3as, 3ad mounted on the left and right half-wings 2s, 2d, it is possible to isolate the effects of load factor (nz) and roll rate (p).

[0167] Example 2 (figure 5).

[0168] Consider a pair of sensors 3bs, 3bd placed on the wing 2 at a height yo from the root of the first half-wing 2s and from the root of the second half-wing 2d, and at a height xo from the center of torsion, respectively.

[0169] Both sensors 3bs, 3bd are sensitive to both the contribution of vertical acceleration and roll, and the contribution of longitudinal deformation induced by the flexural-torsional coupling of the torque generated symmetrically on the two half-wings 2s, 2d by the pitch rate of the aircraft 100.

[0170] Therefore, the following applies:

[0171] £3bd=£n + Eroll + Epitch

[0172] £3bs=£n—Eroll + Epitch

[0173] where:

[0174] £3bd: mechanical deformation of the point in which sensor 3bd is installed; £3bS: mechanical deformation of the point in which sensor 3bs is installed; Eroi . mechanical deformation induced by the variation of the roll rate;Cpitch: mechanical deformation induced by the variation of the pitch rate.

[0175] Given, by means of example 1, the values of nzand p, it is possible to calculate the pitch rate q by summing the deformations and the use of a correlation coefficient (the roll contribution cancels out automatically, the load factor is already known).

[0176] The load factor generally induces a symmetrical bending moment on the two half-wings 2s, 2d, a roll induces an asymmetrical bending moment on the two half-wings 2s, 2d, and the pitch angle is proportional to the torsional moment on the two half-wings 2s, 2d.

[0177] By applying the optical sensors 3 close to the elastic axis of the wing 2, only bending induced by the bending moment generated by roll or a variation of the load factor is read.

[0178] With a second fiber far from the elastic axis, a deformation induced by the bending moment is always read, but generated by a variation in pitch. With optical sensors 3 positioned symmetrically on the two half-wings 2s, 2d, the roll effect is canceled.

[0179] Advantageously, the monitoring method and system allow obtaining a measurement and / or estimation of the three rates and the three accelerations.

[0180] One measurement point on the left half-wing, one measurement point on the right half-wing placed on the neutral axis of the wing at the same distance from the fuselage, preferably on the upper surface of the wing are sufficient for the load factor (they can also read the roll) .

[0181] To also estimate the rates, it is necessary to double the sensor map, adding two further measurement points far from the elastic axis, preferably on the lower surface of the wing.

[0182] With two pairs of optical sensors on the wing or the tail section, it is possible to measure bending, torsion, load moment, moments induced by the commands on the tail plane.Structural deformations, e.g., those of the half-wings, are considered as a data source for processing an informational content comprising information on navigation, attitude of the aircraft, and instantaneous load factor.

Claims

CLAIMS1. A method of monitoring an aircraft (100) comprising a structural component (2, 2s, 2d), said monitoring method comprising:- associating at least one optical sensor (3, 3as, 3ad, 3bs, 3bd) with the structural component (2, 2s, 2d),- detecting a temperature and a deformation of the structural component (2, 2s, 2d) by means of the optical sensor (3, 3as, 3ad, 3bs, 3bd), - thermally compensating for the detected deformation as a function of the detected temperature,- defining a significant numerical model of the component (2, 2s, 2d), - selecting a sensor data correlation algorithm to determine an aircraft condition as a function of the numerical model and the compensated deformation of the structural component (2, 2s, 2d),- calculating a plurality of information items relating to the condition of the aircraft (100) by means of the sensor data correlation algorithm.

2. The monitoring method according to claim 1, comprising defining a load factor of the aircraft (100) as a function of the plurality of calculated information items relating to the condition of the aircraft (100) and of a weight of the aircraft (100).

3. The monitoring method according to any one of the preceding claims, wherein calculating a plurality of information items relating to the condition of the aircraft (100) by means of the defined algorithm comprises calculating a lift of the aircraft (100) and / or an acceleration of the aircraft (100) and / or an angular rate of the aircraft (100) and / or an angle of attack of the aircraft (100).

4. The monitoring method according to any one of the preceding claims, wherein defining a significant numerical model of the structural component (2, 2s, 2d) of the aircraft comprises defining a structuralmodel of the structural component (2, 2s, 2d) of the aircraft.

5. The monitoring method according to any one of the preceding claims, comprising providing a navigation system of the aircraft (100) for detecting a plurality of navigation information items of the aircraft (100) and comparing said plurality of navigation information items of the aircraft (100) with the plurality of calculated information items relating to the aircraft condition.

6. The method according to claim 5, comprising filtering the plurality of navigation information items before comparing said plurality of navigation information items of the aircraft (100) with the plurality of information items relating to the aircraft condition.

7. The monitoring method according to claim 2 and according to claim 5 or 6, comprising combining the plurality of navigation information items with the load factor by means of a data fusion algorithm.

8. The monitoring method according to claim 7, wherein the sensor data fusion algorithm is either a complementary filter or Kalman filter algorithm.

9. The monitoring method according to any one of the preceding claims, wherein associating at least one optical sensor (3, 3as, 3ad, 3bs, 3bd) with the structural component (2, 2s, 2d) comprises associating a first optical sensor (3as, 3bs) with a first half-wing (2s) of the aircraft (100) and associating a second optical sensor (3ad, 3bd) with a second halfwing (2d) of the aircraft (100), said first and second optical sensors (3as, 3bs, 3ad, 3bd) being arranged on opposite sides of a fuselage (101) of said aircraft (100).

10. The monitoring method according to claim 9, wherein said first optical sensor (3as, 3bs) and said second optical sensor (3ad, 3bd) are arranged symmetrically with respect to said fuselage (101).

11. The monitoring method according to claim 9 or 10, wherein said first optical sensor (3as, 3bs) and said second optical sensor (3ad, 3bd) are arranged substantially at an elastic axis of first half-wing (2s) and second half-wing (2d).

12. The monitoring method according to any one of claims 9 to 11, wherein said first optical sensor (3as, 3bs) and said second optical sensor (3ad, 3bd) are arranged substantially at a neutral axis of first halfwing (2s) and second half-wing (2d).

13. A system for monitoring an aircraft comprising a structural component (2, 2s, 2d), said monitoring system comprising at least one optical sensor (3, 3as, 3ad, 3bs, 3bd) associated with the structural component (2, 2s, 2d) and configured to detect a temperature and a deformation of the structural component (2, 2s, 2d), said monitoring system comprisinga control unit in communication with said optical sensor (3, 3as, 3ad, 3bs, 3bd) and configured to calculate a thermal compensation as a function of the detected temperature and deformation to obtain a compensated deformation,the control unit being configured to calculate, by means of the sensor data correlation algorithm, a plurality of information items relating to the condition of the aircraft (100),said sensor data correlation algorithm being defined as a function of a significant numerical model of the structural component (2, 2s, 2d) and of the compensated deformation of the structure component (2, 2s, 2d).

14. The monitoring system according to claim 13, wherein the structural component (2, 2s, 2d, 101) comprises an outer casing (4, 4’, 4”), the optical sensor (3, 3as, 3ad, 3bs, 3bd) being arranged inside said outer casing (4).

15. The monitoring system according to claim 13 or 14, wherein the outer casing (4, 4’, 4”) comprises laminated layers, the optical sensor (3, 3as, 3ad, 3bs, 3bd) being arranged between said laminated layers of the outer casing (4, 4’, 4”).

16. The monitoring system according to any one of claims 13 to 15, wherein the optical sensor (3, 3as, 3ad, 3bs, 3bd) is attached to the structural component (2, 2s, 2d) by means of an adhesive layer (5).

17. The monitoring system according to any one of claims 13 to 16, wherein the optical sensor (3, 3as, 3ad, 3bs, 3bd) is of the Fiber Bragg Grating (FBG) type.

18. The monitoring system according to any one of claims 13 to 17, comprising a navigation system of the aircraft (100) configured to detect at least one navigation information item of the navigation information of the aircraft (100), said navigation system being in communication with the control unit, the control unit being configured to receive said at least one navigation information item of the aircraft (100) and combine said navigation information item of the aircraft (100) with said plurality of information items relating to the condition of the aircraft (100) by means of a sensor data fusion algorithm.

19. The monitoring system according to any one of claims 13 to 18, comprising a first optical sensor (3as, 3bs) associated with a first halfwing (2s) of the aircraft (100) and a second optical sensor (3ad, 3bd)associated with a second half-wing (3d).

20. The monitoring system according to claim 19, wherein said first optical sensor (3as, 3bs) and said second optical sensor (3ad, 3bd) are arranged symmetrically with respect to a fuselage (101) of said aircraft (100).

21. The monitoring system according to claim 19 or 20, wherein said first optical sensor (3as, 3bs) and said second optical sensor (3ad, 3bd) are arranged substantially at an elastic axis of first half-wing (2s) and second half-wing (2d).

22. The monitoring method according to any one of claims 19 to 21, wherein said first optical sensor (3as, 3bs) and said second optical sensor (3ad, 3bd) are arranged substantially at a neutral axis of first halfwing (2s) and second half-wing (2d).

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