Trend monitoring of shock absorber condition
By measuring gas pressure and temperature in real time during the shock absorber's working cycle, calculating and analyzing trends, and generating notification signals, the shortcomings of existing shock absorber condition monitoring technologies are addressed, enabling early warning and planned maintenance of shock absorbers.
Patent Information
- Authority / Receiving Office
- CN · China
- Patent Type
- Patents(China)
- Current Assignee / Owner
- MESSIER DOWTY
- Filing Date
- 2020-11-10
- Publication Date
- 2026-06-09
Smart Images
Figure CN114728692B_ABST
Abstract
Description
Background of the Invention
[0002] Traditionally, various types of aircraft use landing gear shock absorbers, similar to those described in FR2946720 A1, to help reduce or dampen stresses experienced by the aircraft during or after landing. These shock absorbers are typically characterized by an oil-pneumatic strut containing compressible gas and hydraulic oil within a damping chamber. Over time, gas and / or oil may leak from the shock absorber, which can reduce its ability to absorb stress and lead to performance degradation.
[0003] Assessing whether a shock absorber performs as expected requires measuring various parameters. In the case of an hydrothermal shock absorber that incorporates both gas and hydraulic fluid, these parameters typically include the gas pressure, hydraulic fluid volume, system temperature, and shock absorber travel (SAT) when the landing gear is supporting the weight of the aircraft.
[0004] Currently, these measurements have been simplified to SAT, aircraft weight, and ambient temperature measurements after the aircraft has completed a predefined number of flight cycles. If those values exceed the "acceptable range," the shock absorbers require maintenance, resulting in the aircraft being grounded until such maintenance is completed.
[0005] The inventors have recognized that current methods for monitoring the condition of shock absorbers suffer from various shortcomings, including the need to manually check various parameters and the inability to provide advance warnings for maintenance needs, which could lead to unexpected grounding of the aircraft. Summary of the Invention
[0006] According to a first aspect of the invention, a method is provided for monitoring the condition of an aircraft landing gear damper, the damper comprising at least one spring chamber containing gas, the method comprising performing a plurality of measurements of gas pressure and temperature, each pair of gas pressure and temperature measurements being performed at the same time in relation to the operating cycle of the damper; calculating a first value based on each pair of gas pressure and temperature measurements; storing the first value in a log; determining a value trend based on the log, the value trend value indicating a prediction of the performance of the damping chamber; and generating a first notification signal in response to determining that the value trend is outside a first range of values.
[0007] Therefore, the inventors devised a method that enables advance warning of maintenance needs when the shock absorber does not require immediate maintenance, allowing the task to be completed at the airline's preferred time and location without accidentally grounding the aircraft.
[0008] According to a second aspect of the invention, a method for monitoring the condition of an aircraft landing gear shock absorber is provided, the shock absorber including at least one damping chamber containing gas, the method comprising: measuring the pressure and temperature of the gas at a predetermined time; calculating a first value based on the measured gas pressure and temperature; determining whether the first value is outside the first predetermined range, and if it is outside the first predetermined range, generating a first notification signal.
[0009] The method further includes determining whether a temperature-corrected gas pressure or value trend is outside a second predetermined range, which encompasses a first predetermined range, and if so, generating a second notification signal to optionally replace the first notification signal.
[0010] Determining the first value may include calculating the temperature-corrected gas pressure.
[0011] Determining the first value may include calculating the difference between the measured pressure and the expected pressure.
[0012] The first range and / or the second range may be predetermined. Alternatively, the first range and / or the second range may be determined based on temperature measurements.
[0013] The pressure and temperature of the gas in the spring chamber can be measured when the shock absorber is fully extended.
[0014] Determining the pressure and temperature of the gas in the spring chamber can occur before the aircraft lands, after the aircraft takes off, or once the aircraft reaches a predetermined altitude, at a predetermined time before landing, in response to the extension of the landing gear, or during a predetermined phase of flight.
[0015] The measured values of gas pressure and temperature can be transmitted to computing devices outside the aircraft, which includes aircraft landing gear dampers, where the computing devices calculate temperature-corrected gas pressure, store the initial value in a log, and / or determine gas pressure trends.
[0016] According to another aspect of the invention, a system is provided comprising a shock absorber and a sensor, the shock absorber including a spring chamber, the sensor being configured to measure the temperature and pressure of a gas in the spring chamber, wherein the sensor can transmit the measured values of the gas temperature and pressure to a computing device configured to perform the method according to the invention.
[0017] Brief description of the attached figures
[0018] Embodiments of the invention will now be described by way of non-limiting example only, with reference to the following figures, in which:
[0019] Figure 1 This is a schematic diagram of an aircraft landing gear;
[0020] Figure 2This is a schematic diagram of a shock absorber according to an embodiment of the present invention;
[0021] Figure 3 It is a graph of temperature-corrected pressure values and flight cycle numbers;
[0022] Figure 4 The method according to the first aspect of the invention has been explained; and
[0023] Figure 5 The method according to the second aspect of the present invention has been explained.
[0024] Figure 6 The method according to the third aspect of the present invention has been explained.
[0025] Figure 7 The method according to the fourth aspect of the present invention has been explained.
[0026] Detailed Description of Embodiments of the Invention
[0027] Figure 1 A typical aircraft main landing gear 1 is illustrated schematically. The landing gear includes an oleo-pneumatic shock absorber strut having a main housing 3 coupled to a portion of the aircraft 5 and a sliding piston 7 partially housed within the main housing. The end of the sliding piston not housed within the main housing is coupled to the landing gear wheels.
[0028] Figure 2 A system for monitoring the condition of landing gear shock absorbers according to an embodiment of the present invention has been described. The system includes an oil-pneumatic shock absorber 9 (such as...). Figure 1 (As explained in the text) and sensor 11. The shock absorber includes a sliding piston 12, which is slidably coupled in the outer main housing 14 via a bearing 16. The sliding piston 12 and the housing 14 together define an inner cavity including a spring chamber 18, a recoil chamber 28, and a damping chamber 37. During normal operation, the spring chamber 18 contains a compressible gas 22, such as nitrogen, while the recoil chamber 28 and the damping chamber 37 contain hydraulic oil 20. A shock absorber inflation valve 31 can be connected to the spring chamber 18 via a port to allow gas to be added to the spring chamber 18.
[0029] Sensor 11 is configured to measure the pressure and temperature of the gas in spring chamber 18, for example, by means of a pressure transducer (not described) and a temperature probe (not described), or by a combination of pressure and temperature sensors. Measuring the gas pressure and temperature allows the condition of the shock absorber to be assessed by checking whether the measured values are within a predefined range, and thus assessing its potential ability to perform satisfactorily. The sensor is connected to spring chamber 18 via a port and valve 24. Valve 24 allows sensor 11 to be replaced without losing any gas pressure that might require further maintenance. In other embodiments, the sensor may be connected to spring chamber 18 via a port of shock absorber inflation valve 31. The advantage of having the sensor directly measure the gas in spring chamber 18 is that it provides better pressure reading accuracy and allows measurement of the temperature of the gas itself (rather than the ambient temperature). This accuracy allows for a finer definition of maintenance thresholds due to the lower tolerances that must be accounted for. The proposed system also eliminates the risk of gas pressure loss compared to using a portable pressure gauge. In some embodiments, the valve may be configured to remain closed by pressure within the shock absorber and open by the sensor, thereby reducing the possibility of accidental gas leakage from the valve.
[0030] Sensor 11 is arranged communicatively coupled to a computing unit (not described). The computing unit may be housed in a single unit with sensor 11, or it may be located separately at other locations on the aircraft, or together with it remotely from the aircraft. Sensor 11 is configured to measure the pressure and temperature of the gas in spring chamber 18 and transmit these paired temperature and pressure values to the computing unit in response to the satisfaction of predefined conditions. Sensor 11 is configured to measure the gas pressure and temperature in spring chamber 18 at the same point in each operational cycle of the aircraft landing gear, ensuring high repeatability of the measurements and achieving even tighter tolerances within the defined performance range. The predefined conditions can be satisfied when the shock absorber is fully extended or in a state commonly referred to as "unloaded". For example, pressure and temperature measurements can be performed when the landing gear is in the retracted position after takeoff or preferably before landing, as the gas dissolution and temperature will be in their most stable state. Measuring the pressure and temperature of the gas in spring chamber 18 when the shock absorber is fully extended avoids the need to monitor the shock absorber stroke or compression. Sensor 11 can be configured to measure the pressure and temperature of the gas in spring chamber 18 when the landing gear is in the retracted position and a "downshift" command is activated by the flight crew. This ensures that the oil / gas mixture is stable compared to other phases of flight and provides equivalent measurement conditions, such as ambient temperature within the same range, no movement of the landing gear or shock absorbers, and prevention of the influence of gases dissolved in the oil on the measurement. Alternatively, sensor 11 can be configured to measure the pressure and temperature of the gas in spring chamber 18 after a predetermined period of time has elapsed since takeoff. For example, sensor 11 can be configured to measure the pressure and temperature of the gas in spring chamber 18 after the aircraft has reached a predetermined altitude. In some embodiments, the measurement of gas pressure and temperature can occur immediately after takeoff or once the aircraft has reached a predetermined altitude. In some embodiments, pressure and temperature can be measured at a predetermined time before landing, wherein the landing time is estimated based on the aircraft's rate, speed, and altitude, or other systems typically used to facilitate landing procedures. In some embodiments, the measurement of pressure and temperature may occur in response to the extension or deployment of the landing gear, or more generally during a predetermined flight phase. Some embodiments of the invention may use a combination of the above conditions as a trigger for sensor 11 to measure gas temperature and pressure.
[0031] In some embodiments, sensor 11 or computing unit may include internal memory that stores the measured values and transmits them after the aircraft has landed. In some embodiments, the measured values may be transmitted to the computing unit when the aircraft leaves the ground.
[0032] exist Figure 2In the illustrated embodiment, the shock absorber is a single-stage shock absorber that is not separated. However, in other embodiments, the shock absorber may be a two-stage or multi-stage shock absorber. In embodiments with multiple spring chambers, each spring chamber may have a corresponding sensor 11 configured to measure the temperature and pressure of the gas in the respective spring chamber.
[0033] The calculation unit may include a memory on which a log can be stored. The log may store measured values of the temperature and pressure of the gas 22 in the spring chamber 18. The calculation unit processes the paired measured values received by the sensor 11 to generate a first value. In this embodiment, the first value is the temperature-corrected pressure of the gas. The temperature-corrected pressure of the gas can be calculated using Gay-Lussac's law and a predefined reference temperature. In other embodiments of the invention, the first value may be the measured gas pressure. The first value is stored in the log. At the end of each maintenance of the shock absorber, the values in the log may be erased or cleared so that it only contains values measured after the last full maintenance of the shock absorber.
[0034] The calculation unit uses multiple first values or subsets thereof stored in the log to generate a set of corresponding trend values, which form a value trend line based on a predefined algorithm or regression analysis. For example, trend values can be generated by calculating a rolling average of a predetermined number of previously stored first values, or by performing a curve fitting routine on the stored first values to generate the value trend line. The calculation unit compares the value trend values to a first acceptable or compliant stress value range. In this embodiment, the first range is predetermined. If the calculation unit determines that the trend value is outside the first range, and the value in the first range is a stress value within the "tolerance acceptable" range, a first warning or notification signal is generated. In other embodiments, if the calculation unit determines that the trend value is outside a second range (the second range encompasses the first range), a second notification signal is generated. In some embodiments, generating the second notification signal stops the generation of the first notification signal, thus generating only one signal. The first notification signal may indicate that the damper is performing in a suboptimal but acceptable manner, but will soon require maintenance. This allows the aircraft operator to be warned in advance of maintenance needs so that the task can be completed at a preferred time and place. The second notification signal may indicate that the damper is performing unsatisfactorily and requires immediate maintenance.
[0035] In this embodiment, the predetermined values for the first and second ranges are derived from the theoretical isothermal spring curves of the shock absorber across the entire temperature range. In an embodiment where the first value equals the measured gas pressure, the first and / or second ranges can be temperature-corrected pressure ranges based on a predefined reference range and the measured temperature of gas 22. The temperature-corrected ranges can be calculated based on Gay-Lussac's law.
[0036] In another embodiment of the invention, the calculation unit compares the determined first value with a first range and / or a second range, i.e., without calculating a trend. If the first value is outside the first range, a first notification signal can be generated. If the first value is outside the second range, a second notification signal can be generated instead of the first notification signal. Similar to the embodiments described above, the first notification signal may indicate that the damper may be performing satisfactorily but will soon require maintenance, while the second notification signal may indicate that the damper may be operating unsatisfactorily and requires immediate maintenance.
[0037] In another embodiment of the invention, the computer unit generates a value trend as described above, but (before or during) also evaluates whether the determined first value is outside the second range. In such an embodiment, determining that the first value is outside the second range will cause the computing unit to generate a second notification signal, regardless of whether the trend value is within the second range.
[0038] In some embodiments, the first and / or second ranges compared with the trend value may be different from the corresponding first and / or second ranges compared with the first value to assess whether a first notification signal or a second notification signal must be generated.
[0039] In another embodiment of the invention, the calculation unit compares the measured gas pressure with the expected theoretical value of the isothermal spring curve at the measured gas temperature to determine the difference or error. The difference between the theoretical optimal value and the measured pressure is then used as a first value, and, depending on the embodiment, may be stored in a log and used to generate an error trend, or directly evaluated against a first range and a second range to assess the health of the shock absorber.
[0040] In some embodiments of the present invention, the first notification signal is communicated only to ground staff serving the flight, while the second notification signal is communicated to both the flight crew and ground staff.
[0041] Figure 3 The text explains the temperature-corrected pressure of the gas based on temperature and pressure measurements during each working cycle of the shock absorber, as well as the trends calculated by the computing unit. Figure 3In the diagram, the z-axis corresponds to the number of flight cycles, while the y-axis represents the temperature-corrected gas pressure. X marks the temperature-corrected gas pressure value calculated by the calculation unit based on the measured temperature and pressure of the gas in the spring chamber 18 of the shock absorber 9 for each operational cycle of the landing gear. The first range, denoted A, bounded by the maximum value y2 and the minimum value y1, represents the fully compliant gas pressure value, i.e., the expected gas pressure value for a shock absorber with the correct gas pressure and oil quantity. The second range, denoted B, bounded by the maximum value y4 and the minimum value y3, represents the gas pressure value that, while not fully compliant, is within the "outside acceptable tolerance" range but indicates that the shock absorber requires comprehensive maintenance in the near future, and continued operation of the landing gear and shock absorber is safe and acceptable within this range. As previously mentioned, the trend line 30 is calculated based on multiple temperature-controlled pressure values.
[0042] In various embodiments of the invention, where the trend is calculated to assess whether a first or second warning signal must be generated, point 32 on trend line 30 indicates that the shock absorber's performance has transitioned from an "acceptable" range to an "outside acceptable tolerance" range. In response to this transition, a first warning signal is generated notifying the shock absorber that it will soon require maintenance. As more measurements are taken, the trend line extends further to point 34, where the shock absorber's performance transitions from an "outside acceptable tolerance" range to an "outside unacceptable tolerance" range, after which the shock absorber's performance may perform unsatisfactorily and require immediate maintenance. Therefore, the shock absorber must be maintained before the next takeoff.
[0043] In those embodiments of the invention that do not calculate trends but use actual temperature-corrected gas pressure values to determine whether they are outside a compliant or acceptable first or second range to assess whether a first or second warning signal must be generated, a first warning signal is generated at point 36 of the work cycle, which is the first temperature-corrected gas pressure value outside the fully compliant range (since the shock absorber was serviced). Note that at this point, trend line 30 is still within the compliant range. Similarly, the first generation of a second warning signal is at point 38 of the work cycle, where the temperature-corrected pressure is outside the “outside acceptable tolerance” range for the first time since the shock absorber was serviced.
[0044] Figure 4 A flowchart illustrating a method for monitoring the condition of an aircraft landing gear shock absorber according to an embodiment of the present invention is provided. References to the landing gear and shock absorbers should be considered as references to the arrangement of the landing gear and shock absorbers, such as regarding... Figures 1 to 3 The subject of discussion.
[0045] At initial step 40, the temperature and pressure of the gas in the shock absorber are measured. Using the measured temperature and pressure, a first value is then calculated in step 42. The first value is a temperature-corrected gas pressure value. In other methods according to the invention, the first value may be the measured gas pressure. The temperature-corrected gas pressure value is then stored at step 44, and a gas pressure trend is subsequently determined in step 46 based on multiple temperature-corrected gas pressure values stored in a log. Next, in step 48, it is determined whether the gas pressure trend is outside a first range of values, which represents values indicating that the performance of the shock absorber is acceptable or satisfactory. If the gas pressure trend is determined to be outside the first range of values, a first notification signal is generated in step 50. Next, in step 54, it is determined whether the gas pressure trend is outside a second range of values, which represents values indicating that the performance of the shock absorber is safe but will soon require maintenance. If the gas pressure trend is determined to be outside the second range of values, a second notification signal is generated in step 58. If it is determined in step 48 that the gas pressure trend is not outside the first range of values, no further action is taken (step 52). If it is determined in step 54 that the gas pressure trend is not outside the second range of values, no further action is taken (step 56).
[0046] Figure 5 A flowchart illustrating another method for monitoring the condition of an aircraft landing gear shock absorber according to an embodiment of the present invention is provided. Further reference to the landing gear and shock absorber should be considered as a reference to the arrangement of the landing gear and shock absorber, as per [the relevant information]. Figures 1 to 3 The subject of discussion.
[0047] At initial step 60, the temperature and pressure of the gas in the shock absorber are measured. Using the measured temperature and pressure, a first value corresponding to the temperature-corrected gas pressure value is then calculated in step 62. Next, in step 64, it is determined whether the temperature-corrected gas pressure value is outside a first range of values, which indicates that the performance of the shock absorber is acceptable or satisfactory. If the corrected gas pressure value is determined to be outside the first range of values, a first warning signal is generated in step 68. Next, in step 70, it is determined whether the gas pressure value is outside a second range of values, which indicates that the performance of the shock absorber is safe but will soon require maintenance. If the gas pressure trend is determined to be outside the second range of values, a second notification signal is generated in step 74. If it is determined in step 48 that the corrected gas pressure value is not outside the first range of values, no further action is taken (step 66). If it is determined in step 70 that the gas pressure trend is not outside the second range of values, no further action is taken (step 72).
[0048] Figure 6A flowchart illustrating another method for monitoring the condition of an aircraft landing gear shock absorber according to an embodiment of the present invention is provided. Further reference to the landing gear and shock absorber should be considered as a reference to the arrangement of the landing gear and shock absorber, as per [the relevant information]. Figures 1 to 3 The subject of discussion.
[0049] At initial step 76, the temperature and pressure of the gas in the shock absorber are measured. Using the measured temperature and pressure, a first value ΔP is then calculated in step 78. ΔP can be the difference between the measured pressure and the expected ideal pressure for a specific temperature, or the difference between the temperature-corrected pressure and the temperature-corrected ideal pressure. The first value ΔP is then stored in step 80, and a ΔP trend is subsequently determined in step 82 based on multiple ΔP values stored in the log. Next, in step 84, it is determined whether the ΔP trend is outside a first range of values, which indicates values indicating acceptable or satisfactory performance of the shock absorber. If the gas pressure trend is determined to be outside the first range of values, a first warning signal is generated in step 88. Next, in step 90, it is determined whether the gas pressure trend is outside a second range of values, which indicates values indicating that the shock absorber's performance is safe but will soon require maintenance. If the gas pressure trend is determined to be outside the second range of values, a second notification signal is generated in step 58 in step 94. If the gas pressure trend is determined to be outside the first range of values in step 84, no further action is taken (step 86). If it is determined in step 90 that the gas pressure trend is not outside the second range of values, no further action is taken (step 92).
[0050] Figure 7 A flowchart illustrating another method for monitoring the condition of an aircraft landing gear shock absorber according to an embodiment of the present invention is provided. Further reference to the landing gear and shock absorber should be considered as a reference to the arrangement of the landing gear and shock absorber, as per [the relevant information]. Figures 1 to 3 The subject of discussion.
[0051] At initial step 96, the temperature and pressure of the gas in the shock absorber are measured. Using the measured temperature and pressure, a first value ΔP is then calculated in step 98. ΔP can be the difference between the measured pressure and the expected ideal pressure for a specific temperature, or the difference between the temperature-corrected pressure and the temperature-corrected ideal pressure. Next, in step 100, it is determined whether ΔP is outside a first range of values, which indicates values indicating that the performance of the shock absorber is acceptable or satisfactory. If ΔP is determined to be outside the first range of values, a first notification signal is generated in step 104. Next, in step 104, it is determined whether ΔP is outside a second range of values, which indicates values indicating that the performance of the shock absorber is safe but will soon require maintenance. If ΔP is determined to be outside the second range of values, a second notification signal is generated in step 110. If it is determined in step 48 that ΔP is not outside the first range of values, no further action is taken (step 108). If it is determined in step 106 that ΔP is not outside the second range of values, no further action is taken (step 108).
[0052] However, it will be understood that other methods according to embodiments of the present invention may include some of the steps, measurements, and actions described above. Furthermore, the sequence of steps described above does not imply that certain steps cannot be performed simultaneously with the described steps or in a different order.
[0053] Although the invention has been described above with reference to one or more preferred embodiments, it should be understood that various changes or modifications may be made without departing from the scope of the invention as defined in the appended claims. The word "comprising" may mean "including" or "consisting of," and therefore does not exclude the presence of any element or step other than those listed in the claims or specification. The mere fact that certain measures are stated in mutually different dependent claims does not imply that combinations of these measures cannot be advantageously used.
Claims
1. A method for monitoring the condition of a shock absorber (9) of an aircraft landing gear (1), the shock absorber (9) comprising at least one spring chamber (18) containing gas (22), the method comprising: Multiple measurements are performed on the gas pressure and temperature in the spring chamber, with each gas pressure and temperature measurement being performed at the same time in relation to the working cycle of the shock absorber. The first value is determined based on each pair of gas pressure and temperature measurements; Store the first value in the log; Based on the logs, a value trend line for the first value is calculated, and the value trend line indicates performance prediction; as well as A first notification signal is generated in response to determining that the value on the trend line is outside a first range (A) of values, wherein the measurement of the gas pressure and temperature occurs before the aircraft lands.
2. The method as described in claim 1, characterized in that, It also includes determining whether the first value or the value on the trend line is outside a second range (B) that covers the first range (A), and if so, generating a second notification signal as a replacement or supplement to the first notification signal.
3. The method as described in claim 1 or 2, characterized in that, The values of gas pressure and temperature are communicated to a computing device outside the aircraft, which includes the aircraft landing gear shock absorbers, wherein the computing device calculates the temperature-corrected gas pressure, stores the first value in the log, and / or determines a trend line for the value of the first value.
4. The method as described in claim 1 or 2, characterized in that, The gas pressure and temperature are measured at one of the following times: once the aircraft has reached a predetermined altitude; in response to the extension of the landing gear; or when the landing gear is in the retracted position.
5. The method according to any one of claims 1-2, characterized in that, Determining the first value includes calculating the temperature-corrected gas pressure.
6. The method according to any one of claims 1-2, characterized in that, Determining the first value involves calculating the difference between the measured pressure and the expected pressure.
7. The method as described in claim 2, characterized in that, The first range (A) and / or the second range (B) of the value are predetermined.
8. The method as described in claim 2, characterized in that, The first range (A) and / or the second range (B) of the values are determined based on the temperature measurement.
9. The method according to any one of claims 1-2, characterized in that, The gas pressure and temperature are measured when the shock absorber (9) is fully extended.
10. A system for monitoring the condition of a landing gear (1) shock absorber (9) of an aircraft, comprising: Including the shock absorber (9) of the spring chamber (18); and Sensor (11) is configured to measure the temperature and pressure of the gas (22) in the spring chamber (18); The sensor (11) is arranged to transmit measurements of the temperature and pressure of the gas (22) to a computing device configured to perform the method described in any of the preceding claims.
11. An aircraft landing gear (1) comprising the system as claimed in claim 10.