METHOD FOR MONITORING A GEARBOX IN AN AIRCRAFT, MONITORING SYSTEM FOR A GEARBOX IN AN AIRCRAFT AND AIRCRAFT WITH THE MONITORING SYSTEM

Description

[0001] The invention relates to a method for monitoring a gearbox in an aircraft with the features of the preamble of claim 1. The invention further relates to a monitoring system for a gearbox in an aircraft and to an aircraft with the monitoring system.

[0002] In aircraft and helicopter transmissions, high torques can occur, which can cause wear or natural abrasion of the transmission components. It is therefore necessary to monitor the condition of the transmission to ensure smooth and safe flight operations. For this purpose, sensors for detecting ferromagnetic particles are known, in which a magnetic field is generated to attract the ferromagnetic particles and detect them using a suitable detector.

[0003] For example, German patent application DE 102012219242 A1 discloses a measuring device for detecting ferromagnetic particles that are movably arranged in a liquid volume, wherein the measuring device comprises a magnetic means for generating a magnetic field in the liquid volume, a measuring means for the total electrical resistance of the liquid volume with the ferromagnetic particles contained therein, and a means for functional testing of the measuring device with at least one low-resistance resistor between 50 and 500 ohms for carrying out a resistance measurement.

[0004] US 4,598,280 A discloses both a method and a monitoring system for monitoring a gearbox in an aircraft.

[0005] The invention is based on the objective of proposing a method for monitoring a gearbox that provides a qualitative assessment of the gearbox's condition. Furthermore, it is an objective of the invention to propose a corresponding monitoring system and an aircraft equipped with the monitoring system.

[0006] This problem is solved by a method with the features of claim 1, a monitoring system with the features of claim 6, and an aircraft with the features of claim 11. Preferred or advantageous embodiments of the invention will become apparent from the dependent claims, the following description, and the accompanying figures.

[0007] The invention proposes a method for monitoring a gearbox in an aircraft. The gearbox has a gearbox section with at least one measuring area. The gearbox section can be defined by a gearbox interior or a partial compartment of the gearbox interior. Preferably, the measuring area is located in a region of an oil sump and / or a bottom region of the gearbox. Alternatively, the at least one measuring area can also be located at any other point in the gearbox section where increased wear or abrasion of the gearbox components is to be expected. Ultimately, it is provided that the measuring area is always located in a region where oil flow occurs.For example, an oil nozzle can be used to inject a volume of oil into a bearing, so that as the oil flows through the bearing, it picks up particles which are then conveyed to a particle detection sensor by the further flow of oil.

[0008] Within the measuring area, a first and a second electrical contact are arranged, the two electrical contacts being spaced apart from each other. Preferably, this space forms a gap between the two contacts, thereby electrically isolating them. In particular, the two electrical contacts are connected to a current source, with the circuit being interrupted within the measuring area. The two electrical contacts can be spaced less than 1 mm apart, preferably less than 0.5 mm, and specifically less than 0.1 mm apart. Alternatively or optionally, the electrical contacts can also be used to detect individual particles larger than 0.1 mm, preferably more than 0.5 mm, and specifically more than 1 mm.Ultimately, however, it is not the pure particle size that is decisive for successful detection, but rather the volume formed by the particles, which determines an electrical resistance and is what makes current conduction between the contacts possible in the first place.

[0009] The procedure includes the following steps: Generating a magnetic field in the gearbox section, wherein the magnetic field is configured to move at least one ferromagnetic, in particular electrically conductive, particle arranged in the gearbox section towards the measuring area, such that a bridge, in particular electrically conductive, of one or more of the ferromagnetic particles is formed or is formed between the two electrical contacts; generating a measured value based on an electrical parameter between the two electrical contacts; determining a measure of the criticality of a flight operation based on the measured value.

[0010] The magnetic field can be generated permanently or temporarily, particularly for the duration of the measurement. Preferably, the magnetic field is generated within the gap. In particular, an effective area of ​​the magnetic field, i.e., the area in which the ferromagnetic particles are attracted by the magnetic field, is arranged within or overlapping the measuring area.

[0011] In particular, the bridge is only formed if one or more ferromagnetic particles are located within the measuring area. Preferably, the electrically conductive bridge is formed by bridging the gap between the two electrical contacts, thus closing the circuit and generating the measured value. Depending on the particle size, the bridge can be formed by exactly one ferromagnetic particle or by an agglomeration of several ferromagnetic particles within the measuring area or the gap. The electrical parameter can be, in particular, an impedance. Specifically, when at least one ferromagnetic particle is positioned between the electrical contacts, the impedance changes, and the quantity and / or size of the ferromagnetic particles can be determined based on this change in impedance.Preferably, the real and / or imaginary part of the impedance is evaluated. It should be noted that electrical parameters such as capacitance, frequency, conductivity, or the like can be detected directly or indirectly.

[0012] According to the invention, the measured value can be evaluated using an evaluation criterion to determine the criticality level. For example, the evaluation criterion can be defined by a stored target value or value range against which the recorded measured value is compared. For instance, the measured value can be converted into a comparable value during the evaluation. Alternatively, the measured value can also be evaluated directly. In particular, based on the criticality level, a decision can be made as to whether the gearbox is damaged and thus whether flight operations are at risk.

[0013] The advantage of the invention lies particularly in the fact that, based on the criticality measure, the pilot is given the opportunity to independently decide whether to abort a mission during its course. Furthermore, a qualitative assessment of the gearbox's condition can be made based on the criticality measure. The method also enables the simple and cost-effective detection of ferromagnetic particles in the gearbox, while simultaneously allowing the number and size of the detected particles to be determined by evaluating the electrical characteristics.

[0014] According to the invention, a warning signal is generated based on the criticality measure, and this warning signal then displays a measure of the criticality of flight operations. Preferably, the warning signal is output in real time when a current measurement meets the evaluation criterion. The warning signal can, for example, be output as an electrical signal, which is then made available for further processing by a computing and / or display unit. In particular, the criticality measure can be displayed visually, audibly, and / or haptically. This allows the pilot to be easily informed that there is a risk to flight operations, enabling the pilot to make a decision regarding the mission based on the display.

[0015] According to the invention, the criticality level is divided into at least or exactly two criticality levels, with a corresponding warning signal being generated for each level. Preferably, information regarding the criticality can be transmitted via the warning signal. In particular, at least one associated evaluation criterion can be defined for each criticality level, which must be met for a corresponding criticality indication to occur. Preferably, the criticality levels are classified based on particle size and / or particle number, which can be determined based on the electrical characteristic value or a change in the electrical characteristic value. The stepwise display of the criticality allows the pilot to be easily informed whether the mission must be aborted or can be completed, possibly with limitations.

[0016] According to the invention, the system displays a criticality level of no criticality in a first stage, low criticality in a second stage, and high criticality in a third stage. The first stage is displayed when no particles are detected. The second stage is displayed when a particle with a size smaller than or equal to a predefined maximum permissible particle size is detected, e.g., larger than 10 micrometers and / or smaller than 50 micrometers. Furthermore, the third stage is displayed when at least one particle with a large particle size, e.g., larger than 50 micrometers, and / or a high number of particles are detected. These different criticality levels allow for a reliable classification of gearbox damage, enabling the pilot to better assess the risks associated with flight operations.

[0017] In a further implementation, it is envisaged that the criticality of flight operations is assessed based on at least one additional measurement, which is recorded independently of the primary measurement. Specifically, this additional measurement is recorded simultaneously with the primary measurement and correlated with it, with the criticality then being assessed based on the correlated measurement result. Preferably, the measurement is evaluated using at least one evaluation criterion. By considering at least one additional measurement, measurement reliability can be increased and the potential for misinterpretation of a single measurement can be reduced.

[0018] In a more detailed specification, a low criticality level is indicated if the measured value and / or the subsequent measured value are classified into the second criticality level. Specifically, a separate evaluation criterion is defined for each measured value, according to which the measured value is classified into the respective criticality level. Preferably, the measured values ​​are evaluated and / or assessed based on a decision logic. Optionally, the measured values ​​can be weighted differently. For example, the weighting can be selected depending on the current operating state and / or the measuring point and / or the detected particles. Thus, for example, if particles are detected at typically critical points, a high criticality level can be indicated, even if the subsequent measured value does not meet the corresponding evaluation criterion or is classified into the first or second criticality level.Furthermore, a high criticality level is indicated if the measured value and / or the subsequent measured value are classified in the third criticality level. In particular, a high criticality level is indicated if at least two or all measured values ​​are classified in the third criticality level. Optionally, different weightings can also be considered. Thus, a method is proposed that compares and evaluates different measured values, thereby significantly increasing measurement reliability.

[0019] In a further specification, it is provided that an additional measured value is output upon detection of at least one further ferromagnetic particle in a further measuring area within the gearbox. In particular, this further measuring area can be located in the gearbox section or in a further gearbox section. Preferably, ferromagnetic particles are detected in this further measuring area, as previously described. Specifically, ferromagnetic particles can be detected at at least two, preferably more than three, and in particular more than five different positions or measuring areas within the gearbox. Alternatively or optionally, the further measured value is output upon a change in a gearbox parameter. For example, one or more known gearbox parameters such as vibrations, noise, speed, torque, gearbox temperature, or the like can be evaluated for this purpose.A method is therefore proposed which is characterized by high measurement accuracy, whereby a failure of the gearbox can be detected early and reliably by taking several measured values ​​into account.

[0020] It is preferred that the electrical parameter be an ohmic resistance. In particular, particles with a small cross-section exhibit a higher resistance than particles with a larger cross-section. For example, the ohmic resistance can be determined based on Ohm's law and provided as a measured value. Thus, the particle size can be determined based on the ohmic resistance. Alternatively, or optionally, the electrical parameter is an electrical voltage. In particular, a voltage drop occurs across the electrically conductive bridge, so that the voltage difference can be used to calculate the measured value. For example, the electrical voltage can be determined based on Ohm's law. Thus, the particle size can be determined based on the voltage or a change in the voltage.

[0021] A further aspect of the invention is a monitoring system designed and / or suitable for monitoring a gearbox in an aircraft. The monitoring system is designed and / or suitable for carrying out the method as previously described.

[0022] The monitoring system comprises at least one particle detection sensor designed to detect ferromagnetic particles within a measuring area of ​​the gearbox. Specifically, the particle detection sensor is located within the gearbox section of the gearbox, and the particle detection sensor defines the measuring area. The measuring area is understood to be the region around the sensor within which the ferromagnetic particles are detected. For this purpose, the particle detection sensor has two mutually oriented electrical contacts located within the measuring area and connected to an electrical circuit. Preferably, the particle detection sensor is designed as a chip detection sensor, which is configured and / or suitable for detecting ferromagnetic and, in particular, metallic chips. Such sensors are also known, among other things, as magnetic chip detectors.

[0023] The particle detection sensor includes a magnet designed and / or suitable for generating a magnetic field. The magnet can be a permanent magnet or an electromagnet. Its function is to move one or more ferromagnetic particles arranged in the gearbox housing into the measuring range of the particle detection sensor, so that an electrically conductive bridge of one or more of the ferromagnetic particles forms between the two electrical contacts. The magnet is preferably arranged such that the area of ​​effect of the magnet, in which the ferromagnetic particles are magnetically attracted, is located between the two contacts and / or within the measuring range. For example, the magnet can be positioned in the gap between the two contacts.If one or more ferromagnetic particles are present between the electrical contacts, they can be detected based on the electrical characteristics between the two contacts. The particle detection sensor is designed to output a measured value based on this electrical characteristic. In particular, the measured value can be output as an electrical measurement signal.

[0024] The monitoring system also includes an evaluation unit designed and / or suitable for evaluating the measured value in order to determine a measure of the criticality of flight operations. Preferably, the particle detection sensor and the evaluation unit are interconnected via signal transmission. In particular, the evaluation unit contains at least one evaluation criterion based on which the measured value is mapped to the measure of criticality. For example, the evaluation unit can be implemented as a hardware or software module in a flight computer. The evaluation unit is configured.

[0025] In a further embodiment, the monitoring system includes at least one additional particle detection sensor, which is designed to detect ferromagnetic particles in a further measuring area of ​​the gearbox. Specifically, the measuring area and the further measuring area are arranged at different positions within the gearbox, particularly in different and / or separate gearbox sections. Alternatively, the two measuring areas can also be arranged adjacent to each other, particularly in a common gearbox section. For this purpose, the further particle detection sensor is optionally arranged in the gearbox section or the further gearbox section. In particular, the particle detection sensor and the further particle detection sensor are of identical design.For example, the monitoring system can have more than two, in particular more than four, and specifically more than six particle detection sensors. The additional particle detection sensor is configured to output a further measurement value upon detecting at least one further ferromagnetic particle. Preferably, all particle detection sensors are connected to the evaluation unit via signal transmission. The evaluation unit is configured to assess a measure of the criticality of flight operations based on the initial measurement value and the subsequent measurement value. In particular, the evaluation unit is configured to assess the criticality using a decision logic.

[0026] In an alternative or optionally supplementary embodiment, the monitoring system is provided to have at least one or exactly one gearbox monitoring sensor, which is designed and / or suitable for detecting a gearbox parameter. In principle, the gearbox monitoring sensor can be a temperature, speed, or torque sensor. Preferably, however, the gearbox monitoring sensor is an acceleration sensor suitable for detecting vibrations in the gearbox. The gearbox monitoring sensor is designed to output the gearbox parameter as an additional measured value and transmit it to the evaluation unit. For this purpose, the gearbox monitoring sensor and the evaluation unit are interconnected via a signal connection. The evaluation unit is designed to assess a measure of the criticality of flight operations by evaluating the electrical measured value and the gearbox parameter.Specifically, the sensor data from several transmission monitoring sensors can be used and correlated with the measured values ​​of at least one particle detection sensor to assess criticality.

[0027] In a further embodiment, the monitoring system includes a display device configured to indicate a level of criticality for flight operations. Specifically, the display device is located, or can be located, within the pilot's field of vision, preferably inside the cockpit. The display device is preferably designed as a light source or a display that can assume at least two different display states to indicate the criticality. For example, the display states can be distinguished by showing different light signals and / or messages.

[0028] In a specific embodiment, a V-shaped gap is formed between the two contacts, designed to accumulate multiple ferromagnetic particles. In particular, the two contacts are arranged at an angle of more than 20 degrees, preferably more than 50 degrees, and specifically more than 80 degrees, to each other. Ferromagnetic particles of varying sizes can accumulate through the V-shaped gap. Measurements taken over time can then indicate whether the wear is normal or critical.

[0029] Another aspect of the invention relates to an aircraft with the monitoring system as previously described. The aircraft can, in principle, be configured as an airplane, air taxi, or the like. Preferably, however, the aircraft is configured as a helicopter.

[0030] Further features, effects and advantages of the invention will become apparent from the following description of a preferred embodiment of the invention and the accompanying figures. These show: Figure 1 is a highly schematic representation of an aircraft with a monitoring system as an embodiment of the invention; Figure 2 is a schematic representation of a particle detection sensor of the monitoring system according to Figure 1 Figure 3 shows the particle detection sensor in the same representation as in Figure 2 in a first detection state; Figure 4 shows the particle detection sensor in the same representation as in Figure 2 in a second detection state; Figure 5 shows the particle detection sensor in the same representation as in Figure 2 in a third detection state; Figure 6 shows a decision logic for assessing the criticality of flight operations using a schematic block diagram.

[0031] Figure 1Figure 1 shows a highly schematic block diagram of an aircraft 1, represented here only as a functional block, as an embodiment of the invention. For example, aircraft 1 is designed as a helicopter.

[0032] The aircraft 1 has a gearbox 2, which serves to translate a drive torque. For example, the gearbox 2 can translate the drive torque from an engine to a thruster and / or to one or more rotors.

[0033] The gearbox 2 is preferably designed as a mechanical gearbox, e.g. a gear gearbox, which is subjected to high torques and speeds during flight operations.

[0034] The gearbox 2 can be divided into several gearbox sections 3, 4, which are designed, for example, as subsections of a gearbox interior of the gearbox 2. For example, the gearbox sections 3, 4 can be spatially separated from each other, e.g., wet and dry compartments. Alternatively, the gearbox sections 3, 4 can also define a common interior space of the gearbox 2.

[0035] The aircraft 1 has a monitoring system 5, which serves to monitor the gearbox 2 during flight operations. For this purpose, the monitoring system 5 includes at least one particle detection sensor 6, which serves to detect ferromagnetic particles. These ferromagnetic particles can form, for example, as a result of abrasion or damage to the gearbox components due to high loads. The particle detection sensor 6 is located in the first gearbox section 3 in order to detect any ferromagnetic particles that may be present therein. Upon detection of one or more ferromagnetic particles, the particle detection sensor 6 outputs a measured value.

[0036] Optionally, the monitoring system 5 can include an additional particle detection sensor 7, which is arranged in the second gearbox section 4 for detecting ferromagnetic particles. This additional particle detection sensor 7 thus enables the detection of ferromagnetic particles at different positions within the gearbox 2. Upon detecting one or more ferromagnetic particles, the additional particle detection sensor 7 outputs a further measured value.

[0037] Alternatively or optionally, the monitoring system 5 can include a transmission monitoring sensor 8, which serves to detect a transmission parameter. For example, the transmission monitoring sensor 8 is designed as an acceleration sensor, which is arranged to detect vibrations in or on the transmission 2. The transmission monitoring sensor 8 then outputs the transmission parameter as an additional measured value.

[0038] The monitoring system 5 includes an evaluation unit 9, which is configured to evaluate the measured values. Sensors 6, 7, and 8 are connected to the evaluation unit 9 via signal transmission. For example, the evaluation unit 9 forms an integral part of a flight computer.

[0039] The evaluation unit 9 includes an evaluation module 10, which is configured to evaluate the measured values ​​based on at least one evaluation criterion in order to determine a measure of the criticality of flight operations. Furthermore, the evaluation unit 9 may include an assessment module 11, which is configured to correlate the measured values ​​of sensors 6, 7, and 8 with each other and to assess the criticality based on an assessment criterion.

[0040] The evaluation unit 10 is configured to output a warning signal when at least one evaluation criterion and / or assessment criterion is met, the warning signal being provided to a display unit 12. The display unit 12 can, for example, be configured as a light source, with the display unit 13 converting the warning signal into a light signal to indicate the degree of criticality of the flight operation. For this purpose, the display unit 10 is preferably arranged within the pilot's field of vision, and the pilot decides, based on the light signal, whether the mission must be aborted or whether the mission can be completed.

[0041] The Figure 2 Figure 6 shows a schematic representation of the particle detection sensor 6 as a further embodiment of the invention. It should be noted that the following description of the function is applicable analogously to the further particle detection sensor 7.

[0042] The particle detection sensor 6 has a first and a second electrical contact 13, 14, which are electrically separated from each other by a distance A and are arranged within the respective gear section 3, 4. The two contacts 13, 14 are connected to an electrical circuit which is interrupted in the area of ​​the contacts 13, 14 by the gap formed between the contacts 13, 14.

[0043] The particle detection sensor 6 has a magnet 15 which generates a magnetic field F within the gear section 3. For example, the magnet 15 can be a permanent magnet or an electromagnet. The magnet 15 is preferably arranged between the two contacts 13, 14, so that a ferromagnetic particle located in the gear section 3 is moved into a measuring area M of the particle detection sensor 6 due to the magnetic interaction. The measuring area M is understood to be the area of ​​the particle detection sensor 6 in which the ferromagnetic particles can contact or adhere to the contacts 13, 14.

[0044] The Figures 3 to 5 Each figure shows the detection of different particle sizes or numbers by the particle detection sensor 6, as already demonstrated in Figure 2As described above, in a detection state, at least one particle 17 is arranged between the two contacts 13, 14, wherein the presence of one or more particles 17 is detectable between the two contacts 13, 14 based on an electrical characteristic, in particular an impedance change. Preferably, depending on the particle size and / or number, an electrically conductive bridge 16 is formed from one or more particles 17, which bridges the two contacts 13, 14, thus closing the circuit.

[0045] The measured value output by the particle detection sensor 6 is evaluated using an evaluation criterion stored in the evaluation module 10 to determine a measure of the criticality of flight operations. The electrical parameter can be, for example, an ohmic resistance or a voltage, which are evaluated according to Ohm's law. The measure of criticality is determined based on the detected particle size and / or number. For example, a particle 17 with a small cross-section has a higher electrical resistance than a particle 17 with a larger cross-section, so the particle size can be deduced based on the electrical resistance. Additionally, it is possible to determine, as a function of time, how many particles 17 have accumulated on the particle detection sensor 6 within a defined period.

[0046] For example, the evaluation unit 9 is designed to divide the criticality level into three stages, with the display unit 12 indicating no criticality in the first stage, low criticality in the second stage, and high criticality in the third stage. The classification of the stages depends on the determined particle size or number. For example, the first stage is displayed if no particles 17 are detected by the particle detection sensor 6. The second stage is displayed if a single particle 17 with a maximum permissible particle size, e.g., less than or equal to 50 micrometers, is detected, as is the case in Figure 3 This is shown. For example, the third stage is displayed when a particle 17 is detected that is larger than the maximum permissible particle size, as shown in Figure 4shown, and / or if several particles 17 are detected by the particle detection sensor 6 within the specified time period, as shown in Figure 5 As shown, for example, the display device 12 can be configured to show the individual stages as different colors, e.g., traffic light colors. For example, the first stage can be displayed as a green light signal, the second stage as a yellow light signal, and the third stage as a red light signal. This will give the pilot the opportunity to decide whether it is necessary to abort the flight operation.

[0047] The Figure 6 A flowchart illustrates a decision logic for assessing criticality during flight operations.

[0048] In a first process step V1, the measured value of the particle detection sensor 6 is provided to the evaluation unit 9 and, if necessary, converted into an evaluable value.

[0049] In a second process step V2, the measured value is evaluated and classified into one of the three criticality levels S1, S2, S3.

[0050] To avoid misinterpretation of the measured value, the measured value of the additional particle detection sensor 7 and / or the transmission monitoring sensor 8 is added in a third process step V3 and classified into one of the three criticality levels S1, S2, S3, if the measured value of the particle detection sensor 6 is classified into the second or third criticality level S2, S3 in the second process step V2.

[0051] In a fourth process step V4, a warning signal is generated based on the respective criticality level S1, S2, S3 and provided to the display device 12, which is configured to display the respective criticality level S1, S2, S3 based on the warning signal. The display device 12 continues to show the first level S1 until a particle 17 is detected by the particle detection sensor 6.

[0052] If the measured value of the particle detection sensor 6 is classified into the second or third criticality level S2, S3 during the second process step V2, and the measured value of the further particle detection sensor 7 and / or the transmission monitoring sensor 8 is classified into the third criticality level S3 during the third process step V3, a warning signal is generated during the fourth process step V4 to indicate the third criticality level S3.

[0053] If, however, the measured value of the particle detection sensor 6 is classified as the second or third stage S2 or S3 during the second process step V2, and the measured value of the further particle detection sensor 7 and / or the gearbox monitoring sensor 8 is classified as the second criticality stage S3 during the third process step V3, a warning signal indicating the second criticality stage S2 is generated during the fourth process step V4. Thus, the pilot receives a qualitative statement about the condition of the gearbox 2, allowing the pilot to decide, based on this display, whether to end the flight or abort it. Reference sign

[0054] 1 Aircraft 2 Transmission 3 First transmission section 4 Second transmission section 5 Monitoring system 6 Particle detection sensor 7 Additional particle detection sensor 8 Transmission monitoring sensor 9 Evaluation unit 10 Evaluation module 11 Evaluation module 12 Display unit 13 First contact 14 Second contact 15 Magnet 16 Conductive bridge 17 Ferromagnetic particle Distance FMagnetic field MMeasuring range V1-V4Process steps S1-S3Criticality levels

Claims

1. A method of monitoring a gearbox (2) in an aircraft (1), wherein the gearbox (2) comprises at least one gearbox section (3) with at least one measuring area (M), wherein a first and a second electrical contact (13, 14) are arranged in the measuring area (M) at a distance (A) with respect to each other, in which: - a magnetic field (F) is generated in the gearbox section (3), wherein the magnetic field (F) is configured to move at least one ferromagnetic particle (17) disposed in the gearbox section (2) towards the measuring area (M), so that a bridge (16) produced from one or more of the ferromagnetic particles (17) is formed between the two electrical contacts (13, 14); - a measured value is generated on the basis of an electrical parameter between the two electrical contacts (13, 14); - a measure of the criticality of a flight operation is determined by an evaluation module (10) on the basis of the generated measured value, wherein a warning signal is generated on the basis of the measure of the criticality and, based on the warning signal, the measure of the criticality of the flight operation is indicated, wherein the measure of the criticality is classified into at least two criticality levels (S1, S2, S3), and a respective warning signal is generated for each of the criticality levels (S1, S2, S3), and wherein no criticality is indicated in a first criticality level (S1), a low criticality is indicated in a second criticality level (S2) and a high criticality is indicated in a third criticality level (S3), characterized in that a detected particle size and / or number of particles is determined for the generated measured value using an evaluation criterion stored in the evaluation module (10), wherein the measure of the criticality is classified into the first criticality level (S1) if the number of particles detected within a specified time period is equal to zero, wherein the measure of the criticality is classified into the second criticality level (S2) if the detected particle size is smaller than or equal to a specified maximum permissible particle size and the number of particles detected within the specified time period is equal to one, and wherein the measure of the criticality is classified into the third criticality level (S3) if the detected particle size is larger than the specified maximum permissible particle size and / or the number of particles detected within the specified time period is greater than one.

2. The method of claim 1, characterized in that the measure of the criticality is assessed on the basis of at least one further measured value which is detected independently of the generated measured value.

3. The method of claim 2, characterized in that a low criticality is indicated if the generated measured value and / or the further measured value are classified into the second criticality level (S2), and in that a high criticality is indicated if both the generated measured value and / or the further measured value are classified into the third criticality level (S3).

4. The method of claim 2 or claim 3, characterized in that the further measured value is output upon a detection of at least one further ferromagnetic particle (17) in a further measuring area (M) of the gearbox (2) and / or is generated upon a change in a gearbox parameter of the gearbox (2).

5. The method of one of the preceding claims, characterized in that the electrical parameter is an ohmic resistance and / or a voltage.

6. A monitoring system (5) for a gearbox (2) in an aircraft (1), with at least one particle detection sensor (6) for the detection of ferromagnetic particles (17) in a measuring area (M), - wherein the particle detection sensor (6) is arranged in a gearbox section (3) of the gearbox (2), - wherein the particle detection sensor (6) comprises two electrical contacts (13, 14) which are distanced from each other, - wherein the particle detection sensor (6) comprises a magnet (15) for generating a magnetic field (F), wherein the magnet (15) is configured to move a ferromagnetic particle disposed in the gearbox section (3) into the measuring area (M) of the particle detection sensor (6), so that an electrically conductive bridge (16) produced from one or more of the ferromagnetic particles (17) is formed between the two electrical contacts (13, 14), - wherein the particle detection sensor (6) is configured to output a measured value on the basis of an electrical parameter between the two electrical contacts (13, 14), with an evaluation device (9) for evaluating the measured value, - wherein the evaluation device (9) is configured to determine a measure of the criticality of the flight operation by evaluating the measured value, wherein a warning signal is generated by the evaluation device (9) on the basis of the measure of the criticality and the measure of the criticality of the flight operation is indicated on the basis of the warning signal, wherein the measure of the criticality is classified into at least two criticality levels (S1, S2, S3) and a respective warning signal is generated for each of the criticality levels (S1 , S2, S3), and wherein no criticality is indicated in a first criticality level (S1), a low criticality is indicated in a second criticality level (S2) and a high criticality is indicated in a third criticality level (S3), characterized in that the evaluation device (9) comprises an evaluation module (10) in which an evaluation criterion is stored in order to determine a detected particle size and / or number of particles for the measured value using the evaluation criterion, wherein the measure of the criticality is classified into the first criticality level (S1) if the number of particles detected within a specified period of time is equal to zero, wherein the measure of the criticality is classified into the second criticality level (S2) if the detected particle size is less than or equal to a specified maximum permissible particle size and the number of particles detected within the specified period of time is equal to one, and wherein the measure of the criticality is classified into the third criticality level (S3) if the detected particle size is greater than the specified maximum permissible particle size and / or the number of particles detected within the specified period of time is greater than one.

7. The monitoring system (5) of claim 6, characterized by at least one further particle detection sensor (7) for detecting ferromagnetic particles (17) in a further measuring area (M) of the gearbox (2), wherein the further particle detection sensor (7) is arranged in the gearbox section (3) or a further gearbox section (4), wherein the further particle detection sensor (7) is configured to output a further measured value upon detecting at least one further ferromagnetic particle (17), wherein the evaluation device (9) is configured to assess a measure of the criticality of the flight operation on the basis of the measured value and the further measured value.

8. The monitoring system (5) of claim 6 or claim 7, characterized by at least one gearbox monitoring sensor (8) for detecting a gearbox parameter of the gearbox (2), wherein the gearbox monitoring sensor (8) is configured to output the gearbox parameter as a further measured value, wherein the evaluation device (9) is configured to assess a measure of the criticality of the flight operation on the basis of the measured value and the gearbox parameter.

9. The monitoring system (5) of one of claims 6 to 8, characterized by an indicator device (12), wherein the indicator device (12) is configured to indicate the criticality of the flight operation.

10. The monitoring system (5) of one of claims 6 to 9, characterized in that a V-shaped gap is formed between the two contacts (13, 14), the gap being configured to accumulate several of the ferromagnetic particles.

11. An aircraft (1) characterized by a monitoring system (5) according to one of claims 6 to 10.