HK1179584B - Method for detecting whether performance of aircraft component is in the deterioration period and method for maintenance of aircraft - Google Patents

Abstract

The present application discloses a method for detecting whether performance of an aircraft component is in a decline period, comprising: obtaining one or more detection parameters reflecting operation status of the aircraft component; comparing data of the one or more detection parameters with respective desired values; and determining whether the performance of the aircraft component is in the decline period based on a comparison result.

Description

Method for detecting whether performance of aircraft component enters degradation period and aircraft maintenance method

Technical Field

The present invention relates to a method relating to aircraft maintenance, and in particular to a method of detecting that the performance of an aircraft component has entered a decline period.

Background

Aircraft are important vehicles in modern society. Many people travel by air every day. The flight safety of the aircraft must be ensured. When a component of an aircraft fails and does not meet a clearance condition, the aircraft must be stopped for maintenance until the component fails. Thus, failure of a component of an aircraft will likely result in a delay or even a loss of flight of the aircraft.

The current maintenance of aircraft is mainly after-repair or hard-time maintenance. As described above, post-repair is difficult to avoid delays and shutdowns of the aircraft because repair of certain components of the aircraft is time consuming. Sometimes, if the airport has no spare parts available for replacement due to high price or spare parts being used up, etc., this will directly result in the aircraft stopping. In a hard time maintenance mode, a component of the aircraft is repaired or replaced after a fixed period of time. This, while to some extent avoiding delays or stops of the aircraft, has the disadvantage of being too costly. Particularly for some expensive components, the performance may still be good for repair and replacement, which is very wasteful. In addition, for some special cases, the performance of certain components on the aircraft may deteriorate quickly. In this case, the delay and the stop of the aircraft cannot be completely avoided even in the hard time-limited maintenance mode.

Disclosure of Invention

In view of one or more technical problems in the prior art, according to an aspect of the present invention, a method for detecting that performance of an aircraft component enters a decline period is provided, including: acquiring one or more detection parameters reflecting the operation state of the aircraft component; comparing data of the one or more detection parameters with respective predetermined values; and evaluating whether the performance of the aircraft component enters a decline period based on the result of the comparison.

According to another aspect of the invention, there is provided a method of servicing an aircraft, comprising: determining whether the performance of an aircraft component of the aircraft enters a decline period according to the method; scheduling a maintenance plan for the aircraft in response to the performance of the aircraft component entering a decline period; and performing maintenance on the aircraft component of the aircraft.

According to another aspect of the invention, a method for obtaining one or more detection parameters reflecting the operating state of the aircraft component is proposed, comprising: obtaining a plurality of parameters related to the operating state of the aircraft component; correlating data for the plurality of parameters with a failure event for the aircraft component; and determining the detection parameter based on the association of the plurality of parameters with the fault event.

Drawings

Preferred embodiments of the present invention will now be described in further detail with reference to the accompanying drawings, in which:

FIG. 1 is a schematic illustration of a performance curve for an aircraft component according to an embodiment of the invention;

FIG. 2 is a flow diagram of a method of detecting that performance of an aircraft component enters a decline period in accordance with one embodiment of the present invention;

FIG. 3 is a flow chart of a method of obtaining sensed parameters reflecting an operational status of the aircraft component according to an embodiment of the present invention;

FIG. 4 is a flow diagram of a method of detecting performance of an aircraft component according to one embodiment of the invention;

FIG. 5 is a flow diagram of a method of detecting performance of an aircraft component according to one embodiment of the invention;

FIG. 6 is a flow diagram of a method of detecting performance of an aircraft component according to one embodiment of the invention;

FIG. 7 is a schematic diagram of an APU performance variation curve according to one embodiment of the present invention;

FIG. 8 is an example of an A13 message for airbus;

FIG. 9 is a flow diagram of a method of detection of APU performance in accordance with one embodiment of the present invention;

FIG. 10 is a flow diagram of a method of detecting APU performance in accordance with another embodiment of the present invention;

FIG. 11 is a flow diagram of a method of detecting APU performance in accordance with another embodiment of the present invention;

FIG. 12 is a schematic representation of a crew oxygen system performance profile;

FIG. 13 is a flow diagram of a method of detecting crew oxygen system performance according to one embodiment of the present disclosure;

FIG. 14 is a graphical illustration of normalized pressure of oxygen for a crew oxygen system oxygen cylinder versus measurement time in accordance with an embodiment of the present invention;

FIG. 15 is a graphical illustration of normalized pressure of oxygen for a crew oxygen system oxygen cylinder versus measurement time in accordance with an embodiment of the present invention;

FIG. 16 is a schematic diagram of a 24 hour 3 day rolling average leak rate of the crew oxygen system versus measurement time according to the embodiment of FIG. 15; and

FIG. 17 is a flow chart of a method of servicing an aircraft crew oxygen system in accordance with an embodiment of the present invention.

Detailed Description

FIG. 1 is a schematic illustration of a performance curve for an aircraft component according to an embodiment of the invention. As the service time increases, the performance of all aircraft components gradually deteriorates, i.e., the degradation index gradually increases. The degradation index represents how fast the performance of the aircraft component deteriorates. When the decline indexes of the performance of the aircraft components are stable, the performance of the aircraft components is in a stable period; when the performance degradation of the aircraft component is gradually accelerated, the performance of the aircraft component enters a degradation period; when a certain threshold is exceeded, the performance of the aircraft component enters a failure period, and a failure may occur at any time. Adverse consequences can be generated on service quality and flight safety after the airplane components enter the fault period; meanwhile, the system is easy to produce unscheduled maintenance, and causes flight delay and flight stop. There is no means in the prior art to detect whether the performance of an aircraft component has entered a decline period.

The following benefits are achieved for the detection of the decline period: first, when an aircraft component is in the decline phase, the probability of failure is still very low. If the airplane is selected to be overhauled at the moment, the flight safety and the service quality can be guaranteed. Second, when it is detected that the aircraft component is in a degradation period, the airline can schedule maintenance of the aircraft in a timely manner, thereby avoiding unscheduled maintenance and reducing delays in the aircraft. And meanwhile, the cost waste caused by maintenance according to a hard time limit is avoided. Of course, embodiments of the present invention may also be applicable to the detection of a period of failure.

FIG. 2 is a flow diagram of a method of detecting that performance of an aircraft component enters a decline period in accordance with one embodiment of the present invention. As shown in the figure, the detection method 200 of the present embodiment includes: in step 200, one or more detection parameters reflecting the operation state of the aircraft component are acquired; at step 220, comparing data of the one or more detection parameters with corresponding predetermined values; and at step 240, assessing whether the performance of the aircraft component enters a decline period based on the result of the comparison.

One problem to be solved first of all in order to implement a test of the performance of an aircraft component is to test which parameters, i.e. to select which test parameters reflect the operating state of the aircraft component. Taking an airbus A320 airplane as an example, the number of the system data collected by the airplane can be up to 13000. Many of these parameters also reflect, directly or indirectly, the performance of the aircraft component. Therefore, it is a difficult problem how to select a suitable detection parameter from the performance parameters of a plurality of aircraft components, especially for aircraft components with relatively complex structures.

FIG. 3 is a flow chart of a method for obtaining sensed parameters reflecting an operational status of the aircraft component according to one embodiment of the invention. As shown in the figure, the method 300 for acquiring the detection parameters of the embodiment includes: at step 320, obtaining a plurality of parameters related to the operational status of the aircraft component; at step 340, correlating data of the plurality of parameters with a failure event of the aircraft component; and at step 360, determining the detection parameter based on the association of the plurality of parameters with the fault event.

At step 320, the plurality of parameters associated with the operational status of the aircraft component may be all or a portion of the parameters acquired by the aircraft data system. A large number of parameters which are completely independent of the aircraft component to be detected can be excluded, depending on the physical meaning represented by the individual parameters. According to one embodiment of the invention, the detection parameters are determined on the basis of the physical meaning represented by a plurality of parameters relating to the operating state of the aircraft component.

By utilizing the existing data of a plurality of parameters related to the running state of the aircraft component when the fault event of the aircraft component occurs, which parameters are highly related to the fault event of the aircraft component can be judged through data mining, so that the range of the detection parameters can be determined.

According to one embodiment of the invention, a degree of correlation of the data change of the plurality of parameters with the fault event of the aircraft component is calculated. A fault event may directly affect the deterioration of certain parameters. For example, if a crew oxygen system leaks, the crew oxygen pressure parameter may drop rapidly. If the engine fails, the engine speed will drop rapidly. The deterioration of these parameters represents the occurrence of a fault event. Calculating the correlation between the parameter associated with the operating condition of the aircraft component and the parameter representing the fault event may reflect the correlation between the parameter and the fault event. According to one embodiment of the invention, a Partial Correlation method is used to calculate the Correlation between the parameters related to the operating state of the aircraft component and the parameters representing the fault event. Various partial correlation analysis methods in statistics can be applied to the present embodiment. And if the correlation degree of the parameter and the fault event is found to be larger than a threshold value through calculation, the parameter is taken as a detection parameter. By verifying all parameters related to the operating state of the aircraft component in this way, detection parameters reflecting the operating state of the aircraft component can be obtained.

The value of the threshold determines the final detection parameters and the accuracy of the performance detection. The more parameters the more accurate the detection is of course, however the higher the cost of implementing the detection method. If the performance of an aircraft component is related to multiple parameters, each with a lesser degree of correlation, then the threshold needs to be lowered to incorporate more parameters. According to an embodiment of the invention, the threshold value ranges from 0.3 to 0.5. If the performance of the aircraft component is associated with a few parameters and the correlation with certain parameters is high, the threshold may be increased to reduce unnecessary detections. According to an embodiment of the invention, the threshold value ranges from 0.6 to 0.8.

After obtaining a plurality of detection parameters, it is also possible that the detection parameters may be highly correlated. In this case, one sensed parameter may represent another sensed parameter. Therefore, a part of the detection parameters can be removed by checking the correlation. According to one embodiment of the invention, a correlation between a plurality of detection parameters is calculated; and removing one or more of the plurality of detection parameters based on the correlation between the plurality of detection parameters. By this step, a plurality of detection parameters which are relatively independent and reflect the operation state of the aircraft component are obtained.

By comparing the measured value of the detection parameter with the limit value of the detection parameter, the deterioration degree of the detection parameter can be reflected more intuitively. If the measured value of the parameter under test must be controlled within the limit value of the parameter under test, the parameter under test may be considered to be degraded and the performance of the aircraft component may have entered a degradation period when the measured value of the parameter under test approaches the limit value. If the measured value of the parameter under test is allowed to exceed the limit value of the parameter under test, the parameter under test may be considered to be degraded and the performance of the aircraft component may have entered a degradation period when the measured value of the parameter under test approaches or exceeds the limit value. By integrating the detection results of the plurality of detection parameters, the performance of the aircraft component can be more accurately judged to possibly enter the decline period. According to one embodiment of the invention, the measured value of each test parameter is replaced with a value in a reduced specified state to obtain a more accurate result.

FIG. 4 is a flow diagram of a method of detecting aircraft component performance according to one embodiment of the invention. As shown, the detection method 400 of the present embodiment includes: at step 420, a plurality of sensed parameters reflecting the operational status of the aircraft component are obtained. In step 440, a ratio of the measured value of each of the plurality of measured parameters to the corresponding limit value is calculated. At step 460, a weight is assigned to the ratio of the measured value to the limit value for each sensed parameter. In step 480, the ratio of the weighted measured values of the plurality of detection parameters to the limit values is integrated to obtain a performance reference value of the aircraft component. Thereby, it is detected whether the performance of the aircraft component enters a decline period.

The weights of the respective detection parameters can be estimated from the actual data. According to one embodiment of the invention, the weight of each detection parameter is derived from the degree of correlation of the data change of the plurality of detection parameters with the fault event of the aircraft component.

According to one embodiment of the invention, it is determined that the performance of the aircraft component enters a decline period if the performance reference value of the aircraft component is greater than a threshold value. The threshold value generally needs to be estimated from actual data.

FIG. 5 is a flow diagram of a method of detecting aircraft component performance according to one embodiment of the invention. As shown, the detection method 500 of the present embodiment includes: at step 520, a plurality of sensed parameters reflecting the operational status of the aircraft component are obtained. In step 540, a slope term of the trend of the measured values of the detected parameters over a period of time is calculated.

As the service time increases, the performance of the aircraft components also gradually deteriorates. This attribute can be reflected by the following formula:

X=β0+β1t₀ （1）

wherein X is a measured value of the detection parameter, t₀Is the installation time of the aircraft component, β 0 and β 1 are fitting parameters. Wherein, the beta 1 is a slope term and reflects the variation trend of the detection parameter.

In step 560, the slope term of the trend of the measured values of the one or more test parameters over time is compared to a reference slope term to determine if there is a significant change in the two. And determining that the performance of the aircraft component enters a decline period if the slope term of the change trend of the measured value of the detection parameter in a period of time is significantly changed relative to the reference slope term.

According to one embodiment of the invention, the reference slope term is a slope term of a trend of change over a period of time after initial installation of the aircraft component. According to another embodiment of the invention, the reference slope term is a slope term of the trend of the change of said aircraft component over time in good working condition on other aircraft of the same type. By using the method, the change of the same airplane at different times can be compared, and the comparison can be carried out among different airplanes.

According to one embodiment of the invention, the measured value of the sensed parameter may be replaced with a value in a reduced designated state.

According to one embodiment of the invention, the measured values of the detection parameters are smoothed to reduce the influence of data disturbance. The smoothing process uses a rolling average method of multipoint averaging. Learning, the following formula is used:

X_new=C1X_smooth+C2X_old （2）

wherein, X_oldIs a measured value; x_newIs the smoothed value; x_smoothIs the smoothed value of the neighboring points or the average value of several nearby points; c1 and C2 are weight values. C1 is generally much larger than C2 to increase the smoothing effect.

FIG. 6 is a flow diagram of a method of detecting aircraft component performance according to one embodiment of the invention. As shown, the detection method 600 of the present embodiment includes: at step 620, a plurality of sensed parameters reflecting the operational status of the aircraft component are obtained. In step 640, the measured value of the detection parameter in a time period is taken as a sample; in step 660, the measured value of the detection parameter in the same time period before the time period is used as a reference sample; and at step 680, determining whether a significant change has occurred between the sample and the reference sample based on an independent sample test.

A variety of independent sample testing methods in statistics can be applied to the present embodiment. Determining that the performance of the aircraft component enters a decline period if a significant change occurs between a sample of the measured values of the one or more measured parameters and the corresponding reference sample.

According to one embodiment of the invention, the measured value of the sensed parameter may be replaced with a value in a reduced designated state.

After determining whether the performance of the aircraft component of the aircraft enters the decline period according to the method, responding to the performance of the aircraft component entering the decline period, a maintenance plan of the aircraft can be scheduled timely, and the aircraft component of the aircraft can be maintained.

Application example of the onboard auxiliary power unit:

an Airborne Auxiliary Power Unit (AIrborne Autoliary Power Unit), called Auxiliary Power Unit APU for short, is a small turbine engine installed at the tail of an airplane. The main function of the APU is to provide power and air, and there are also a small number of APUs that can provide additional thrust to the aircraft. Specifically, the APU provides power to start the main engine before the aircraft takes off from the ground, so that ground electric and air source vehicles are not needed to start the aircraft. On the ground, the APU also provides power and compressed air to ensure lighting and air conditioning in the passenger cabin and the cockpit. When the airplane takes off, the APU can be used as a standby power supply. After the aircraft lands, the APU still supplies power for illumination and air conditioning.

The function of the APU determines the stability of its operation directly related to the flight cost and quality of service of the aircraft. Moreover, in the absence of ground power supply and air supply guarantee, once an APU fails, the aircraft can be directly disabled. At present, the fault removal and maintenance of the APU are almost post-processing. However, among aircraft devices, the APU is a relatively expensive maintenance device. In addition, the price of the whole part of the APU is high, the cost of the storage spare parts is high, and the repair sending period is up to 4-5 months after the fault. The post-processing maintenance mode ensures that the stable operation of the APU cannot be guaranteed. Moreover, since the time consumption after the APU repair is long, the time consumption is also long, which directly causes the delay and even the stop of the airplane.

FIG. 7 is a schematic diagram of an APU performance variation curve, according to one embodiment of the present invention. As the usage time increases, all APU performance becomes progressively worse, i.e., the fade index increases. When the decay index of the APU performance is stable, the APU performance is in a stable period; when the performance degradation of the APU is gradually accelerated, the performance of the APU enters a degradation period; when a certain threshold value is exceeded, the performance of the APU enters a failure period, and the APU can fail at any time. When the APU enters a fault period, the use of the APU is influenced, and adverse consequences are generated on the service quality and the flight safety; and is easy to cause non-planned maintenance, resulting in flight delay and flight stop. There is no means in the prior art to detect whether the performance of the APU enters the decline period. And certain embodiments of the present invention may implement such detection.

Fig. 8 is an example of an a13 message for airbus. As shown in the figure, the a13 message mainly includes 4 pieces of information, which are: the system comprises a header, APU resume information, operating parameters for starting an aircraft engine and APU starting parameters.

The header consists of CC and C1 sections and mainly comprises flight information of the airplane, a message generation section stage, a bleed valve state, total temperature (namely outside temperature) and other information. The APU resume information is composed of section E1 and comprises information such as APU serial number, operation hour, cycle and the like. The operation parameters for starting the aircraft engine consist of sections N1 to S3; wherein, N1 and S1 represent the operation condition when the first aircraft engine is started, N2 and S2 represent the operation condition when the second aircraft engine is started, and N3 and S3 represent the condition when the APU is slowed down after the APU finishes starting the engine.

The a13 message includes a plurality of parameters related to the operating conditions of the APU. The operation parameters of the starting engine comprise EGT temperature, IGV opening angle, compressor inlet pressure, load compressor inlet temperature, bleed air flow, bleed air pressure, lubricating oil temperature and APU generator load. The parameters of the APU during starting comprise starting time, EGT peak value, rotating speed at the EGT peak value and inlet temperature of a load compressor.

In addition to the parameters in the a13 message, the performance of the APU may be related to other parameters. Taking an airbus A320 airplane as an example, the number of the system data collected by the airplane can be up to 13000. Many of these parameters also reflect APU performance, either directly or indirectly.

For the engine, there are two of the most important criteria affecting the heat engine, the first being age and the second being exhaust temperature EGT. From a physical point of view, the time of use should be a very important parameter. When the APU fails, however, the APU exhaust temperature will rise and approach the limit. Therefore, valuable information is extracted starting from these two parameters. In this example, the external environmental influences, such as altitude, total temperature, generator load, bleed air flow, inlet pressure, load compressor inlet temperature, are eliminated by applying a partial correlation method. The actual data of the APU is analyzed, and the following results are obtained:

in this example, the correlation r is divided into three levels: low degree linear correlation, | r | < 0.4; 0.4 ≦ r | <0.7 is significance correlation; 0.7 ≦ r | <1 is highly linear correlation.

From the analysis results, the age TSR, start-up time STA, engine exhaust temperature EGT, and bleed air pressure PT are weakly correlated, but the inlet guide vane angle IGV and Oil Temperature (OTA) are strongly correlated with age TSR, start-up time STA, engine exhaust temperature EGT, and bleed air pressure PT.

Therefore, under the normal operation condition of all parts of the APU, the inlet guide blade angle IGV and the oil temperature OTA can be represented by the using time TSR, the starting time STA, the engine exhaust temperature EGT and the bleed air pressure PT. On the other hand, the parameters of the use time TSR, the start time STA, the engine exhaust temperature EGT and the bleed air pressure PT are relatively independent and each represents the operating characteristics of a certain APU. By means of the four parameter characteristics, the overall performance condition of the APU can be reflected through effective combination.

FIG. 9 is a flow diagram of a method of detection of APU performance in accordance with one embodiment of the present invention. As shown in the figure, in the detection method 9000 of the APU performance of this embodiment, in step 9100, the following information of the operation of the aircraft APU is obtained: exhaust temperature EGT, compressor inlet temperature LCIT, start-up time STA, service time TSR and bleed air pressure PT. At step 9200, the difference between EGT and LCIT, EGT-LCIT, STA, TSR and PT, are compared to respective thresholds. According to one embodiment of the invention, the threshold value is a limit value of the respective parameter. In step 9300, respective weights are assigned to the results of the comparisons of EGT-LCIT, STA, TSR, and PT to the respective thresholds. In step 9400, the results of comparisons of EGT-LCIT, STA, TSR, and PT with respective thresholds after considering the weights are integrated. At step 9510, a determination is made whether the integrated result exceeds a first predetermined value. If the integrated result does not exceed the first predetermined value, then in step 9520, it is determined that the performance of the APU is good; at step 9610, it is determined whether the integrated result exceeds a second predetermined value. If the second predetermined value is not exceeded, then in step 9620, the APU performance is judged to be normal; in step 9710, it is determined that the integrated result is greater than a third predetermined value. If the third predetermined value is not exceeded, then at step 9720 it is determined that the APU performance has entered the grace period. If the integrated result exceeds the third predetermined value, then in step 9800, it is determined that the APU performance has entered a failure period.

According to one embodiment of the invention, the information required in step 9100 may be obtained from an APU message, such as the A13 message. For example, an a13 message of the operation of the aircraft APU can be remotely acquired in real time from an international aviation telecommunication group SITA network control center and an ADCC network control center of a national aviation data communication company, and the message decoder decodes the a13 message of the operation state of the aircraft APU to obtain the operation information of the aircraft APU.

And if the APU running state message is not automatically generated in the aircraft data system, adding corresponding sensors and triggering conditions to generate the required APU message. And if the existing APU message in the aircraft data system does not completely cover one or more of the exhaust temperature EGT, the compressor inlet temperature LCIT, the starting time STA, the use time TSR and the bleed air pressure PT, modifying the generation condition of the APU message and increasing one or more missing measurement parameters. The APU message can be transmitted to the data server of the airline company in real time through the ACARS or ATN system, so that the real-time monitoring of the performance of the APU can be realized. Of course, the message transmission mode can also avoid the high cost and human error of the manual mode.

According to one embodiment of the invention, the information required in step 9100 may be obtained directly from the aircraft data system without generating APU messages.

In step 9200, the difference between EGT and LCIT, EGT-LCIT, is set as EGT_Readline。EGT_ReadlineIs the EGT red line value of the APU. EGT_ReadlineDepending on the APU model. Different models of APUs have different EGT red line values, which can be obtained by looking up the relevant manual. The threshold of STA is STA_WarningLineIs the STA performance decay value, which also depends on the APU model. The threshold value of TSR is TSR_rtThe term "wing time reliability" means the time corresponding to 70% wing time reliability of an APU of a certain model. The threshold value of PT is PT_MinMeaning the minimum required bleed air pressure for a particular model APU. Alternatively, the threshold value of PT is PT_BaseLineMeaning the lowest inherent bleed air quantity for a certain model APU when it is operating normally. Comparing EGT-LCIT, STA, TSR and PT with respective thresholds can reflect the degree of deviation of the performance of the current APU from the standard performance of the APU, and thus reflect the degree of performance degradation of the APU. EGT_Readline、STA_WarningLineAnd PT_MinOr PT_BaseLineCan find the phase by looking upThe relevant airplane manual or obtained from the manufacturer. Of course, they can also be obtained by actual experiments. However, TSR_rtDue to the influence of other factors such as geography and maintenance environment, the standard value is deviated to a certain extent. The inventor finds that the aging mode of the APU is Poisson distribution through long-term observation and analysis. In order to obtain more accurate TSR_rtData from which the required TSR can be calculated from the actual data by Poisson distribution_rt. For example, the parameters (such as the mean value) of the poisson distribution followed by the actual using time TSR may be first calculated, and then the corresponding using time TSR when the failure rate is 30% (the stability rate is 70%) may be calculated by using the obtained parameters of the poisson distribution followed by actual using time TSR_rt。

The EGT-LCIT, STA, TSR, and PT may be compared with their respective thresholds in a ratio manner or in a difference manner. To facilitate considering the weights of the various parameters, in step 9200, ratios of EGT-LCIT, STA, TSR, and PT to the respective thresholds are calculated, according to one embodiment of the invention.

EGT-LCIT, STA, TSR and PT have different effects on APU performance, so they need to be assigned different weights. In the case of obtaining the ratios of EGT-LCIT, STA, TSR and PT to the respective thresholds, according to an embodiment of the present invention, take R1, R2, R3 and R4 as the respective weights of EGT-LCIT, STA, TSR and PT, and R1+ R2+ R3+ R4= 1. Based on the observation and analysis of the inventors, the effect of TSR was greatest, so R3 was generally greater than 0.25; the influence of EGT-LCIT and STA may be different for different types of APUs; in contrast, the PT effect is small and R4 is minimal. According to one embodiment of the invention, for an APU of APS3200 model, R3=0.35, R2=0.3, R1=0.2, R4= 0.15. For GTCP131-9A model APU, R3=0.35, R1=0.3, R2=0.2, R4= 0.15.

According to one embodiment of the invention, the performance of the APU is evaluated using the following formula:

wherein, the PDI (Performance Detection index) performance Detection index is a parameter reflecting the performance of the APU. According to the inventors' observation and analysis, if the PDI is less than 0.7, the APU performs well; if the PDI is greater than 0.7 and less than 0.85, the performance of the APU is normally available; if the PDI is greater than 0.85, the APU performance is poor and the fade period has been entered. If the PDI is close to 1, e.g., greater than 0.95, this indicates that the APU has entered a failure period and is likely to fail at any time. Thus, one example of the first predetermined value in step 9510 is 0.7 and one example of the second predetermined value in step 9610 is 0.85; one example of the third predetermined value in step 6710 is 0.95.

The method of the above embodiment of the present invention is further illustrated below by 2 examples.

Example 1: the relevant information for an APU of APS3200 model is as follows: EGT_ReadlineIs 682; STA (station)_WarningLineIs 90; PT_MinIs 3; TSR_rtIt was 5000. The weighting parameters R1=0.2, R2=0.3, R3=0.35, R4=0.15 are taken.

The method comprises the steps of remotely acquiring an aircraft APU message from an SITA network control center or an ADCC network control center in real time, decoding the aircraft APU message through an ACARS message decoder to obtain aircraft APU operation information, and comprises the following steps: the exhaust temperature EGT is 629, the compressor inlet temperature LCIT is 33, the start time STA is 59, the on wing time TSR is 4883 and the bleed air pressure PT is 3.66, by the following equation:

the PDI value was calculated to be 0.85. And judging that the performance of the APU enters a decline period, and planning to maintain the APU of the airplane.

Example 2: the relevant information for the APU of GTCP131-9A model is as follows: EGT_ReadlineIs 642; STA (station)_WarningLineIs 60; PT_MinIs 3.5; TSR_rtIt was 5000. The weighting parameters R1=0.3, R2=0.2, R3=0.35, R4=0.15 are taken.

The method comprises the following steps of remotely acquiring an aircraft APU message from an SITA network control center or an ADCC network control center in real time, and decoding the aircraft APU message through an ACARS message decoder to obtain the aircraft APU operation information, wherein the method comprises the following steps: an exhaust temperature EGT of 544, a compressor inlet temperature LCIT of 31, a start time STA of 48, an on wing time TSR of 2642 and a bleed air pressure PT of 3.76, by means of the formula

The PDI value was calculated to be 0.72. And judging that the performance of the APU is normal and the APU can still be normally used.

Compared with the prior art, the embodiment of the invention obtains the PDI value by real-time acquisition of the exhaust temperature EGT, the inlet temperature LCIT of the compressor, the starting time STA, the wing time TSR and the bleed air pressure PT of the APU, and calculates according to the formula (1), and then the accurate detection of the performance of the APU is limited according to the comparison of the PDI value and the preset value. In addition, the ACARS message of the operation state of the aircraft APU is remotely acquired in real time, so that the workload of manual acquisition is reduced, and the working efficiency is improved.

The measurement of EGT and PT is affected by differences in altitude and temperature. According to one embodiment of the invention, the measured EGT and PT are compared to a standard state to remove the effect of altitude and ambient temperature for more accurate detection of APU performance. For example, an altitude of 0 m and a temperature of 50 ℃ may be selected as the standard state, and other altitudes and temperatures may be selected as the standard state.

According to one embodiment of the present invention, the atmosphere correction formula for PT under the standard condition of 0 m altitude and 50 ℃ temperature is

Wherein PT_stdIs the pressure at an altitude of 0 meters, ALT is the altitude or standard altitude, TAT is the ambient or total temperature, m is the air mass, which can take the value of 29. g takes 10 m/s²And R is an adjusting parameter and can take a value of 8.51.

The altitude pressure correction coefficient can thus be derived:

considering the influence of temperature, the final PT correction formula is

Wherein PT_corIs the corrected bleed air pressure, Δ PT is a temperature-dependent function, which can be calculated using the following formula:

ΔPT＝a1TAT²+b1TAT+c1 （6）

wherein TAT is ambient temperature; a1, b1, and c1 are adjustment coefficients. a1, b1 and c1 can be obtained by experimental measurement. According to one embodiment of the invention, a1 has a range of 10^-5Of order, b1 is 10^-2And c1 is between 0 and-1.

After a1, b1 and c1 are experimentally measured, the measured PT can be converted into a corrected PT in a standard state according to equation (6)_cor。

The correction formula for EGT is as follows:

wherein EGT_corIs EGT in the normal state, Δ EGT is a function dependent on temperature, PT_ReqI.e. PT_MinIs the lowest bleed pressure required at engine start, and p1 and p2 are adjustment factors. According to one embodiment of the invention, the value range of p1 is 20-60, and the value range of p2 is 70-100. Specific values for p1 and p2 can be obtained experimentally. For example, at different sea level barometric altitudes, different EGTs were measured while maintaining a temperature of 50 degrees, maintaining a certain power output. Then, the change of the EGT is compared with the EGT of 50-degree sea level air pressure, and the change of the EGT and the temperature are regressed, so that the adjusting coefficient in the correction formula can be obtained.

Δ EGT can be calculated using the following equation:

ΔEGTA＝a2TAT²+b2TAT+c2 （8）

wherein TAT is ambient temperature; a2, b2, and c2 are tuning parameters. a2, b2 and c2 can be obtained by experimental measurement. According to one embodiment of the invention, a2 ranges from 0.005 to 0.02, b2 ranges from 0.5 to 2.5, and c2 ranges from 60 to 100.

After the modified EGT and PT are used, equation (3) can be rewritten as:

according to one embodiment of the invention, for the corrected PDI, if the PDI is less than 0.7, the APU performs well; if the PDI is more than 0.7 and less than 0.8, the performance of the APU is normally available; if the PDI is greater than 0.8, the APU performance is poor and the fade period has been entered. If the PDI is greater than 0.85, it indicates that the APU has entered a failure period. Thus, one example of the first predetermined value in step 6510 is 0.7, and one example of the second predetermined value in step 6610 is 0.8; one example of the third predetermined value in step 6710 is 0.85.

FIG. 10 is a flow diagram of a method of detecting APU performance in accordance with another embodiment of the present invention. As shown, in the APU performance detection method 1000, in step 1010, one or more of an aircraft APU operating exhaust temperature EGT, a start time STA, a bleed air pressure PT, and an IGV angle are obtained. The method for acquiring the performance parameters of the APU described in the embodiment of fig. 9 may be applied to this embodiment.

According to the principles of APU operation, one important parameter reflecting the performance of the APU is EGT, i.e., the APU exhaust temperature. Because the EGT directly reflects the heat energy conversion efficiency of the whole APU when the APU runs at a constant rotating speed. The lower the APU thermal energy conversion efficiency, the higher the value of EGT. Because the control system of the APU can operate the fuel flow valve and the IGV inlet angle to ensure no overtemperature, when the APU is in a state close to the overtemperature and needs to prevent the overtemperature, the PT and IGV angles in the parameters of the APU can reflect the change. STA is a parameter that reflects the overall performance of the APU, including the performance of the starter motor, the performance of the gearbox, and the efficiency of the compressor unit and the power unit (i.e., one compressor and two stages of turbines). By monitoring the four key parameters EGT, IGV, STA and PT, the current performance of the APU and the change trend thereof can be reflected. Moreover, the respective detection of the parameters also facilitates the fault source judgment of the APU and the discovery of the hidden fault.

At step 1020, it is determined whether a significant change in one or more of the exhaust temperature EGT, the start time STA, the bleed air pressure PT, and the IGV angle has occurred. And if one parameter of the exhaust temperature EGT, the starting time STA, the bleed air pressure PT and the IGV angle is changed obviously, judging that the parameter is bad.

For EGT and PT, EGT in the above-described embodiments can be applied_corAnd PT_corInstead of directly obtained EGT and PT, to eliminate the influence of altitude and temperature and obtain more accurate results.

As usage time increases, APU performance also gradually deteriorates. This attribute of the APU performance parameter may be reflected by the following formula:

X=β0+β1t₀ （10）

wherein X is any one of the parameters of exhaust temperature EGT, start time STA, bleed air pressure PT and IGV angle, t₀Is the installation time of the APU, and β 0 and β 1 are fitting parameters. Wherein, the beta 1 is a slope term and reflects the variation trend of the parameter.

According to one embodiment of the invention, a plurality of values of one of EGT, STA, PT and IGV taken over a period of time are fitted to yield a slope term β 1. Comparing β 1 with the slope term as a reference, if the slope terms are significantly different, it is judged that the one of EGT, STA, PT, and IGV has significantly changed. The slope term used as a reference is calculated by using the data of the APU with good working state, and the slope term can be the data after the initial installation of the same APU, and can also be the data of other APUs with good working state of the same type.

According to one embodiment of the invention, after the parameters of the APU installation and the APU are initialized, the initial recorded parameters are averaged to obtain the initial value of each parameter as the respective reference value. The number of the plurality of records is generally greater than or equal to 10 records.

And comparing the subsequent parameters with the reference value to obtain the change value of the self. These variation values also conform to equation (10). Their slope terms may also reflect the trend of change of the APU parameters. Therefore, in the present embodiment, the slope term of one of the EGT, STA, PT, and IGV with respect to the change value of the reference value is compared with the slope term of the change value as a reference, and if the slope terms are significantly different, it is determined that the one of the EGT, STA, PT, and IGV has significantly changed. This parameter becomes worse.

According to one embodiment of the invention, the parameter values of one of the EGT, STA, PT and IGV in the consecutive equal-length time periods are compared with independent samples, and if the parameter values of the EGT, STA, PT and IGV are obviously changed, the one of the EGT, STA, PT and IGV is judged to be obviously changed. This parameter becomes worse.

And smoothing the parameter values of the EGT, the STA, the PT and the IGV which are actually measured in order to reduce the interference of fluctuation. According to one embodiment of the invention, the parameter values are smoothed by means of a multipoint smooth average rolling mean. The number of the dots is 3 or more. According to another embodiment of the invention, the parameters are smoothed using the following formula:

X_new=C1X_smooth+C2X_old （11）

wherein, X_oldIs a value before smoothing, i.e. a value actually measured; x_newIs the smoothed value; x_smoothIs a smoothed value, which may be a smoothed value of a neighboring point (e.g., a previous point) or an average value of several nearby points (regardless of the current point); c1 and C2 are weight values, C1 is greater than C2, e.g., C1=0.8, C2= 0.2.

In step 1030, a determination is made as to whether the performance of the APU is degraded, taking into account, in combination, whether one or more of the exhaust temperature EGT, the start time STA, the bleed air pressure PT and the IGV angle has changed significantly.

According to one embodiment of the invention, if any one of EGT, PT, STA and IGV is deteriorated, the performance of the APU is judged to be deteriorated, and the fading period is entered. According to another embodiment of the invention, if the STA is bad, the performance of the APU is judged to be bad, and the decay period is entered. According to another embodiment of the invention, if any two of EGT, PT, STA and IGV go bad, the performance of APU is determined to be degraded, and the degradation period is entered. According to another embodiment of the invention, if both EGT and PT are deteriorated, the performance of APU is judged to be deteriorated, and the decay period is entered.

The embodiments of fig. 9 and 10 may be used simultaneously to more accurately detect the performance of the APU.

FIG. 11 is a flow diagram of a method of detecting APU performance in accordance with another embodiment of the present invention. As shown, in the APU performance detection method 1100, at step 1110, one or both of an aircraft APU operating exhaust temperature EGT and a bleed air pressure PT are obtained. The method for acquiring the performance parameters of the APU described in the above embodiment may be applied to the present embodiment.

In step 1120, the exhaust gas temperature EGT and the bleed air pressure PT are compared with their respective limit values. Specifically, EGT may be related to EGT red line value EGT_RedLineComparing; the bleed air pressure PT may be equal to the lowest bleed air pressure PT required at engine start-up_ReqAnd (6) comparing.

In step 1130, if one of the exhaust gas temperature EGT and the bleed air pressure PT is close to its limit value, it is judged that the parameter is bad. According to one embodiment of the invention, the performance of the APU is judged to enter a decline period if one of the exhaust gas temperature EGT and the bleed air pressure PT deteriorates. According to another embodiment of the invention, the performance of the APU is judged to enter a decline period if both the exhaust gas temperature EGT and the bleed air pressure PT are deteriorating.

According to one embodiment of the invention, the following formula may be used for EGT:

EGT_Tolerance=EGT_RedLine-EGT_cor （12）

wherein, EGT_ToleranceIndicating margin of EGT, i.e. EGT from red line value EGT_RedLineThe distance of (c). Since the APU control system will prevent the EGT from over-warming, when the control mechanism is active, it is an indication that the APU is no longer able to obtain more power by increasing the fueling. The power of the APU is gradually decreased with the increase of the usage time, which indicates that the APU enters the decline phase. Therefore, when EGT is used_ToleranceApproaching 0, indicates that the APU enters the decline phase.

PT is an important observed parameter when the APU enters the decline phase.

According to one embodiment of the invention, the following formula may be used for PT:

PT_Tolerance=PT_cor-PT_Req （13）

wherein PT_ToleranceThe margin for PT, i.e. the distance of PT from the lowest bleed pressure required at engine start-up, is indicated. PT_ToleranceReflects the operation condition of the APU in the decline stage. When PT_ToleranceNear 0, the APU should be replaced.

Example 3: according to the exhaust gas temperature EGT, the external temperature TAT, the altitude ALT and the PT data obtained by the message, the EGT can be obtained by calculation_cor4.49，PT_cor= 3.27. According to the inquiry, the lowest bleed air pressure PT of the air passenger A319 for starting the aircraft_Req= 3.2. Long-term experiment verifies that the red line value EGT of APS3200 model APU_RedLine= 645. From the above performance evaluation formula, one can derive: EGT_Tolerance= 9.49, a value of 9.49/645 close to 0, about 1.4%; PT_Tolerance0.07, the approximation to the 0 value was 0.07/3.2, about 2.2%. Therefore, the EGT and PT parameters are both deteriorated, the APU enters the decline period, and the appropriate time should be selected for replacement, so that the availability of the flights is improved.

The methods of fig. 9-11 may be used simultaneously to more accurately detect the performance of the APU.

Compared with the prior art, the method provided by the embodiment of the invention can realize the performance detection of the APU by acquiring parameters such as the exhaust temperature EGT of the APU, the inlet temperature LCIT of the compressor, the starting time STA, the wing time TSR, the bleed air pressure PT, the angle of the inlet guide vane IGV and the like in real time, and can judge whether the performance of the APU enters the decline period or not, thereby providing good support for an engineer to maintain the APU, ensuring the use of the APU and avoiding the delay and the stop of the airplane caused by the use of the APU. Meanwhile, maintenance and operation control can be carried out in a targeted manner by evaluating the performance of the APU, so that the maintenance cost is greatly reduced.

Application example of unit oxygen system：

FIG. 12 is a graphical representation of a crew oxygen system performance profile. All oxygen systems have a small amount of leakage, so that at a given temperature, a pressure differential of Δ P occurs at different times. And the air leakage rate can be P_LAnd = Δ P/t. When air leakage rate P_LWhen the stability is achieved, the performance of the oxygen system of the unit is in a stable period; when air leakage rate P_LWhen the temperature is gradually increased, the performance of the oxygen system of the unit enters a decay period; when air leakage rate P_LGreater than a threshold value P_LgIn time, the performance of the crew oxygen system enters a failure period, and a failure may occur. The influence is beneficial to flight safety, and the non-planned maintenance is easy to generate, so that the flight delay and the flight stop are caused. There is no means in the prior art to detect whether a crew oxygen system enters a decay period. Such detection may be accomplished according to one embodiment of the present invention.

For a crew oxygen system, the detection parameters are easier to obtain. The oxygen pressure of the oxygen cylinder in the crew oxygen system is the best detection parameter for reflecting the performance of the crew oxygen system. Since the oxygen pressure of the oxygen cylinder in the crew oxygen system is temperature dependent, the oxygen pressure must be obtained simultaneously with the temperature of the oxygen in the oxygen cylinder. However, a temperature sensor is not generally installed in the oxygen system. Therefore, it is necessary to calculate the temperature of the oxygen in the oxygen cylinder from other temperatures that can be measured.

In view of the location of the oxygen cylinder in the crew oxygen system, according to one embodiment of the present invention, the following formula may be used to derive the temperature of the oxygen in the oxygen cylinder:

wherein Tat represents the atmospheric temperature or the outside temperature, Tc represents the cabin temperature, k₁And k₂Is a tuning parameter and satisfies k₁+k₂And (2). According to one embodiment of the present invention, k₁>k₂. That is, the oxygen temperature T and the atmospheric temperature Tat are correlated with the cabin temperature Tc, and the influence of the atmospheric temperature is greater. Of course, other averaging equations may be used to calculate the oxygen temperature.

According to one embodiment of the present invention, k₁=k₂. That is, equation (14) may be rewritten as:

where k is an adjustment parameter. According to one example of the invention, k is a number that is relatively close to the value 1. k. k is a radical of₁And k₂All can pass throughThe actual measurements are obtained, and may also be obtained by statistical analysis.

According to one embodiment of the present invention, k =1 may be taken. Equation (15) can be rewritten as:

while the oxygen temperature thus derived may not be as accurate as equations (14) and (15), it is sufficient for embodiments of the present invention to detect crew oxygen system performance.

After the oxygen temperature is obtained, the measured pressures of the oxygen of the unit at different temperatures can be converted into standard state pressures at standard temperatures for comparison and calculation of the leakage rate. The standard state pressure can be calculated by the following formula:

wherein P is_sIs the pressure of the standard state, T_sIs the standard temperature, P is the measured oxygen pressure, and T is the temperature of the oxygen at the time of measurement. The standard temperature may be 25 ℃. Of course, other temperatures may be used.

FIG. 13 is a flow chart of a method of detecting crew oxygen system performance according to one embodiment of the present invention. In a method 1300 of detecting performance of a crew oxygen system as shown in fig. 13, oxygen pressure data, atmospheric temperature, and cabin temperature of oxygen cylinders in the crew oxygen system are obtained 1310. At step 1320, a crew oxygen message is generated based on the obtained oxygen pressure data of the oxygen cylinders in the crew oxygen system, the atmospheric temperature, and the cabin temperature. In step 1330, the generated crew oxygen message is transmitted to a server for processing the crew oxygen message. In step 1340, the server converts the oxygen pressure of the oxygen cylinders in the crew oxygen system to a standard state pressure at a standard temperature based on the atmospheric temperature and the cabin temperature. The standard temperature may be 25 ℃. Of course, other temperatures may be used.

As shown in FIG. 13, in step 1350, multiple sets of normalized pressure data of the unit oxygen system at different times are obtained in the manner of step 1310-. After a plurality of sets of standard state pressures of oxygen in oxygen bottles in the crew oxygen system at different times and at standard temperatures are obtained, the performance of the crew oxygen system can be determined by processing and evaluating the data.

At step 1360, the sets of normalized pressure data at different times are analyzed to determine if the crew oxygen system performance is degraded. Alternatively, in step 1370, multiple sets of standard state pressure data at different times are compared as one sample to another sample of another set of standard state pressure data for the same type of aircraft to determine if the crew oxygen system performance is degraded.

According to one embodiment of the invention, the segment leakage rate is used to determine if the performance of the crew oxygen system is deteriorating. The section leakage rate of the oxygen system of the unit can be calculated by adopting the following formula:

wherein, t₁Time of flight, t₂Time of flight descent, P_s1Is the unit oxygen standard state pressure P when the airplane takes off_s2The oxygen standard state pressure of the aircraft after landing is obtained. Therefore, the oxygen standard state pressure change delta P of the unit can be changed according to the oxygen standard state pressure before taking off and after landing_sTo determine the performance of the crew oxygen system. For example, if Δ P_s=P_s1-P_s2Above 100 PSI, the performance of the onboard oxygen system deteriorates.

The performance of the crew oxygen system can also be determined according to the flight leakage rate. For example, if the leg leakage rateAbove 48 PSI/day, the performance of the onboard oxygen system deteriorates.

And according to the calculated section leakage rate, estimating the pressure reading of the oxygen system of the unit at a certain temperature. The situation that the oxygen cylinder is not changed in an planned way before flying due to large temperature change of the airplane and the refrigerator after flying in winter can be greatly reduced.

According to one embodiment of the invention, the oxygen pressure P is normalized by the oxygen pressure of the crew oxygen system_sAnd the installation time t of the oxygen cylinder of the oxygen system of the unit_oThe performance of the crew oxygen system is determined by detecting the slope of the fitted curve.

P_sAnd t_oThe relationship conforms to the following formula:

P_s=β1+β2*t_o+μ （19）

wherein, P_sIs the pressure of the standard state, t_oIs the installation time of the oxygen cylinder of the oxygen system of the unit, beta 1 is an intercept term which is related to the flight time; β 2 is a slope term that reflects the hermeticity of the oxygen system; and μ is a random interference term that reflects P_sAnd t_oThe uncertainty in between.

t_oThe mean value of (d) can be expressed as follows:

where n represents the number of sampled data points involved in the calculation.

P_sThe mean value of (d) can be expressed as follows:

where n represents the number of sampled data points involved in the calculation.

According to equations (6) to (8), β 2 can be calculated using the following equation

Beta 2 is a negative value. A smaller value of β 2 indicates a poorer tightness of the crew oxygen system. The performance of the crew oxygen system can be derived by detecting the change in β 2, i.e., the slope term. By comparing the slope term β 2 between different aircraft, the performance of the crew oxygen systems of these aircraft can also be understood.

When the slope detection method is adopted to detect the performance of the oxygen system of the unit, events such as oxygen bottle replacement or oxygenation are preferably not carried out in the time represented by the data points participating in calculation.

According to one embodiment of the invention, the condition that the performance of the crew oxygen system is poor is determined by a method of Independent Sample T Test (Independent Sample Test) of the leakage rate.

Because the time interval of the flight period is short, the possible change of the system pressure is small, the system pressure is easily influenced by the fitting precision of the external temperature and the detection precision of the pressure sensor, and the standard state pressure fluctuation obtained by calculation is large sometimes. Is composed ofAccording to one embodiment of the invention, the air leakage rate is not adopted, and two points with the interval of more than 24 hours are adopted for pressure comparison, namely the leakage rate P with the interval of 24 hours is adopted_L24. Of course, other time intervals, such as time intervals greater than 12 or 36 hours, may also be used. Meanwhile, to eliminate the data dead pixel effect caused by the sampling problem, P is added_L24A3 day rolling average may be used, meaning that all P's within 3 days are calculated_L24Average value of (a). The 3 days are only given as examples, but other days, such as 2-4 days, may of course also be used. Depending on the situation of the data.

According to one embodiment of the invention, the 24-hour 3-day rolling average leakage rate P reflecting the performance characteristics of the crew oxygen system is calculated by using the following formula_L-avg24，：

Where n represents the number of data points in 3 days.

According to one embodiment of the present invention, if it is desired to determine whether a change in the oxygen performance of a crew member occurs over a certain period of time, the data over the set of time periods may be taken as a set of samples; at the same time, another set of data for the same type of aircraft is taken as a set of samples. P of two groups of data samples_L-avg24And comparing, and determining whether the two groups of data have significant changes according to statistical probability so as to judge the time period and the degree of the performance deterioration of the oxygen system of the unit.

According to one embodiment of the present invention, first, P is calculated for 2 sets of data_L-avg24And calculate P_L-avg24The variance. Assume S1²Is a first group P_L-avg24Variance of (including n items of data), S2²Is a second group P_L-avg24(containing m data) variance. Due to S1²/S2²The F value should be determined by finding the F distribution table by difference, subject to the F (n-1, m-1) distribution. And whether the two groups of data have obvious difference can be judged according to the F value. Two sets of data can be considered to be significantly different if the probability of the two sets of data belonging to the same distribution is examined to be less than 2.5%.

Other independent sample T-test methods may also be used to determine if there is a significant difference between the two sets of data. If this difference is significant, it indicates that there is a significant change in the performance of the crew oxygen system. If the performance of the oxygen system of the unit is obviously changed, which group of data represents the performance deterioration of the oxygen system of the unit can be easily judged according to the average value of the permeability.

The independent template test method for average leakage rate can use data of the same airplane in different time periods, and can also use data of the same type of airplane in different time periods. Therefore, this method is flexible. Moreover, the checking mode is not limited by whether the oxygen bottle is replaced or not and oxygenation is carried out, and the method can be used for comparing whether the performance of the unit oxygen system is obviously changed or not before and after the oxygen bottle is replaced and oxygenation is carried out.

The following examples are provided to illustrate how the method of the present invention can be used to detect whether a significant change in the performance of a crew oxygen system has occurred.

Fig. 14 is a graph illustrating the normalized pressure of oxygen in the oxygen cylinder of the crew oxygen system versus the measurement time, according to one embodiment of the present invention. In fig. 14, broken lines represent the standard state pressures of the actual sampling conversion, and straight lines represent lines regressed according to the standard state pressures of oxygen and the measurement time, respectively. The detection is carried out by adopting a formula (22) of a slope detection method, and the leakage rate of the oxygen system of the unit is overlarge, the slope is-0.024929 and is much smaller than the normal slope which is lower than-0.015. This reflects the degradation of the crew oxygen system and the decay period has been entered.

Fig. 15 is a graph illustrating the normalized pressure of oxygen in the oxygen cylinder of the crew oxygen system versus the measurement time, according to one embodiment of the present invention. The figure shows a process for replacing a crew oxygen system oxygen cylinder at a time. The dots in FIG. 15 represent the normalized pressure for the actual sample transition. FIG. 16 is a graph illustrating the 24 hour 3 day rolling average leakage rate of the crew oxygen system versus the time of measurement according to the embodiment of FIG. 15. Two sets of data before and after replacing the oxygen cylinder are used as two samples, and an independent sample T test method is adopted to test whether the two samples are the same. The calculation shows that the probability of the two sets of data being identical before and after the replacement of the oxygen cylinder is zero. The performance of the oxygen system of the unit is poor, and the average leakage rate is 2 times of the original average leakage rate. The performance of the crew oxygen system has entered the decay period.

As can be seen from the embodiments of fig. 14 to fig. 16, the method for detecting the performance of the crew oxygen system according to the present invention can obtain whether the performance of the crew oxygen system is deteriorated or not by processing and analyzing the oxygen pressure data and the temperature data of the crew oxygen system obtained in the crew oxygen message, and by calculating a slope or performing a T test on an independent sample, and the like, and enter a performance decay period or a failure period of the crew oxygen system.

FIG. 17 is a flow chart of a method of servicing an aircraft crew oxygen system in accordance with an embodiment of the present invention. In a method 1700 of aircraft crew oxygen system maintenance as shown in fig. 17, oxygen pressure data, atmospheric temperature, and cockpit temperature of oxygen cylinders in the crew oxygen system are obtained 1710. At step 1720, a crew oxygen message is generated according to the acquired oxygen pressure data of the oxygen cylinder in the crew oxygen system, the atmospheric temperature and the cockpit temperature. In step 1730, the generated crew oxygen message is transmitted to the server. In step 1740, the server processes the crew oxygen message to obtain the standard pressure of the oxygen bottle in the crew oxygen system at the oxygen standard temperature. In step 1750, it is determined whether the performance of the crew oxygen system is degraded based on the plurality of sets of standard state pressure data at different times. In step 1760, if the performance of the crew oxygen system is degraded, a suitable time is scheduled for maintenance of the crew oxygen system.

The above embodiments are provided only for illustrating the present invention and not for limiting the present invention, and those skilled in the art can make various changes and modifications without departing from the scope of the present invention, and therefore, all equivalent technical solutions should fall within the scope of the present invention.

Claims (22)

1. A method of detecting that performance of an aircraft component enters a decline period, comprising:

acquiring one or more detection parameters reflecting the operation state of the aircraft component;

comparing data of the one or more detection parameters with respective predetermined values; and

evaluating whether the performance of the aircraft component enters a decline period based on the results of the comparison;

wherein the predetermined value is a limit value of the one or more detection parameters;

wherein the step of comparing the data of the one or more detection parameters with respective predetermined values comprises: calculating a difference or ratio of the measured value of the one or more sensed parameters to the corresponding limit value;

the method further comprises: assigning a weight to the ratio of the measured value to the limit value for each of the one or more measured parameters; wherein the weight for each detection parameter is derived from a degree of correlation of a data change of the plurality of detection parameters with a fault event of the aircraft component;

wherein the step of assessing whether the performance of the aircraft component enters a decline period based on the result of the comparison comprises integrating the weighted ratios of the measured values of the plurality of measured parameters to the limit values to derive a performance reference value for the aircraft component; and determining that performance of the aircraft component enters a decline period in response to the performance reference value for the aircraft component being greater than a threshold value.

2. The method of claim 1, wherein the step of obtaining one or more detection parameters reflecting the operational status of the aircraft component comprises:

obtaining a plurality of parameters related to the operating state of the aircraft component;

correlating data for the plurality of parameters with a failure event for the aircraft component; and

determining the detection parameter based on the association of the plurality of parameters with the fault event.

3. The method of claim 2, wherein the data for the plurality of parameters is associated with a failure event of the aircraft component; and determining the detection parameter based on the association of the plurality of parameters with the fault event comprises:

calculating a degree of correlation of data changes of the plurality of parameters to the fault event of the aircraft component; and

and taking one or more parameters of the plurality of parameters, of which the association degree is greater than a threshold value, as the detection parameters.

4. The method of claim 3, wherein the threshold value is in the range of 0.3-0.5.

5. The method of claim 3, wherein the threshold value is in the range of 0.5-0.7.

6. The method of claim 3, further comprising:

calculating the correlation degree among the plurality of detection parameters; and

removing one or more of the plurality of detection parameters based on the correlation between the plurality of detection parameters.

7. The method of claim 1, wherein the step of obtaining one or more parameters reflecting the operational status of the aircraft component comprises:

obtaining a plurality of parameters related to the operating state of the aircraft component; and

determining the detection parameters according to the physical meaning represented by the plurality of parameters.

8. The method of claim 1, wherein the measured values of the one or more detection parameters are replaced with values at a reduced designated state.

9. The method of claim 1, wherein based on the results of the comparison, the step of assessing whether aircraft component performance enters a decline period comprises determining whether the measured values of the one or more detected parameters approach or exceed the corresponding limit values.

10. The method of claim 1, wherein the predetermined value is a reference slope term of a trend of change of the one or more detection parameters;

wherein the step of comparing the data of the one or more detection parameters with respective predetermined values comprises: and calculating a slope term of the change trend of the measured values of the one or more detection parameters in a period of time.

11. The method of claim 10, further comprising: and comparing the slope term of the change trend of the measured values of the one or more detection parameters in a period of time with the reference slope term to determine whether the two change significantly.

12. The method of claim 11, wherein the step of evaluating whether aircraft component performance enters a decline period based on the results of the comparison comprises: determining that the performance of the aircraft component enters a decline period in response to a slope term of a trend of the measured values of the one or more detection parameters over a period of time changing significantly relative to a reference slope term.

13. The method of claim 10, wherein the reference slope term is a slope term of a trend of change over a period of time after initial installation of the aircraft component.

14. The method of claim 10, wherein the reference slope term is a slope term of a trend of change over time of the aircraft component operating well on other aircraft of the same model.

15. The method of claim 10, further comprising replacing the measured values of the one or more detection parameters with values at a reduced designated state.

16. The method of claim 10, further comprising smoothing the measured values of the one or more detection parameters.

17. The method of claim 16, wherein the smoothing process uses a rolling average of multiple point averages.

18. The method of claim 16, wherein the smoothing process employs the following equation:

X_new＝C1X_smooth+C2X_old

wherein, X_oldIs a measured value; x_newIs the smoothed value; x_smoothIs the smoothed value of the neighboring points or the average value of several nearby points; c1 and C2 are weight values.

19. The method of claim 1, wherein the step of comparing the data of the one or more detection parameters with respective predetermined values comprises:

taking the measured values of the one or more detection parameters over a period of time as a sample;

taking the measured values of the one or more detection parameters in an equal time before the time period as a reference sample; and

determining whether a significant change has occurred between the sample and the reference sample based on an independent sample test.

20. The method of claim 19, wherein the step of evaluating whether aircraft component performance enters a decline period based on the results of the comparison comprises: determining that the performance of the aircraft component enters a decline period in response to a significant change between a sample of the measured values of the one or more test parameters and the corresponding reference sample.

21. The method of claim 19, further comprising replacing the measured values of the one or more detection parameters with values at a reduced designated state.

22. A method of servicing an aircraft, comprising:

the method of any of claims 1-21, determining whether performance of an aircraft component of the aircraft enters a decline period;

scheduling a maintenance plan for the aircraft in response to the performance of the aircraft component entering a decline period; and

performing a repair on the aircraft component of the aircraft.