HK1179585B - A method for detecting performance of an aircraft based on a customized message and method for maintenance of an aircraft - Google Patents

Abstract

The present invention relates to a method for detecting the performance of the aircraft comprising: collecting data reflecting operation status of the aircraft; generating the customized message based on the collected data, wherein, the customized message is configured to include one or more main parameters related to the operation status of the aircraft; storing or transmitting the customized message; and detecting the performance of the aircraft based on the customized message.

Description

Airplane performance detection method based on customized message and airplane maintenance method

Technical Field

The invention relates to a method for detecting airplane performance, in particular to a method for detecting airplane performance based on customized messages.

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 of these components may still be good when they are repaired and replaced on a hard time basis, which is very wasteful. In addition, for some special cases, the performance of certain components on the aircraft may deteriorate quickly. In this case, even in the case of hard-time maintenance, delay and stop of the aircraft cannot be completely avoided.

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 aircraft performance based on customized messages is provided, including: collecting data reflecting the operation state of the airplane; generating a customized message based on the collected data, wherein the customized message is customized to include one or more primary parameters related to an aircraft operating state; storing or forwarding the customized message; and detecting the performance of the aircraft based on the customized message.

According to another aspect of the invention, there is provided a method of servicing an aircraft, comprising: detecting the performance of the airplane by using the method; and performing maintenance on the aircraft in response to a performance fault of the aircraft; or in response to a degradation in performance of the aircraft, scheduling an appropriate time for maintenance of the aircraft.

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 flow diagram of a customized message-based aircraft performance detection method according to one embodiment of the invention;

FIG. 2 is a schematic diagram of a crew oxygen system performance profile;

FIG. 3 is a flow diagram of a method of detecting crew oxygen system performance according to one embodiment of the present invention;

FIG. 4 is an example of a customized group oxygen message according to one embodiment of the invention;

FIG. 5 is a graphical illustration of normalized pressure of oxygen in an oxygen cylinder of a crew oxygen system versus measurement time, in accordance with one embodiment of the present invention;

FIG. 6 is a graphical illustration of normalized pressure of oxygen in an oxygen cylinder of a crew oxygen system versus measurement time, in accordance with one embodiment of the present invention;

FIG. 7 is a graph illustrating a 24 hour 3 day rolling average leak rate of a crew oxygen system versus measurement time, according to the embodiment of FIG. 6;

FIG. 8 is a flow chart of a method of servicing an aircraft crew oxygen system in accordance with an embodiment of the present invention;

FIG. 9 is a flow diagram of a method for detecting aircraft landing quality according to one embodiment of the invention;

FIG. 10 is a flow diagram of a method for generating a landing message using an aircraft ACMS system according to one embodiment of the invention;

fig. 11 is a schematic diagram of a trigger relationship for generating a landing short message in an ACMS system according to an embodiment of the present invention;

FIG. 12 is a flow diagram of a method for generating a landing message using an aircraft ACMS system in accordance with another embodiment of the invention;

fig. 13 is a schematic diagram of a trigger relationship for generating a landing long message in an ACMS system according to an embodiment of the present invention;

FIG. 14 is an example of a customized landing short message according to one embodiment of the invention; and

FIG. 15 is an example of a customized landing length message according to another embodiment of the invention.

Detailed Description

FIG. 1 is a flow diagram of a customized message-based aircraft performance detection method according to one embodiment of the invention. As shown in fig. 1, the method 100 includes: at step 120, data reflecting the operational status of the aircraft is collected. According to one embodiment of the invention, the data acquisition of the aircraft operating state is carried out using an aircraft data system. As aircraft systems become more complex, aircraft data systems have evolved greatly. For example, the airbus flight status Monitoring System (ACMS) System and the boeing Aircraft health Monitoring System (AHM) System.

Taking the ACMS system of an air passenger as an example, the ACMS system monitors the performance of a number of important components on an aircraft, including: engines, crew, onboard Auxiliary Power Units (APUs), and passenger cabins. The ACMS system also has functions of Monitoring important Aircraft Performance Monitoring (Aircraft Performance Monitoring), data Recording (Date Recording), Special Investigation and troubleshooting (Special Investigation & troubling). The ACMS system can monitor more than 13,000 flight data in real time.

Referring to fig. 1, the method 100 further comprises: at step 140, a customized message is generated based on the data collected. This step can also be accomplished using a flight data system. One function of ACMS and AHM systems is that they can automatically generate messages containing specific data based on real-time monitored data when certain trigger conditions are met. According to one embodiment of the invention, the ACMS system or the AHM system is utilized to generate the customized message.

Taking the airbus ACMS System as an example, the ACMS System includes the flight Integrated Data System (AIDS). And a Data Management Unit (DMU) is the core of the AIDS system. The DMU has two very important functions:

-acquiring, processing and recording a number of parameters on board the aircraft. Including data from black boxes. These parameters are stored in the DMU's internal non-volatile memory or in an external Recorder, such as the AIDS Digital Recorder (DAR);

-generating a system message. And triggering to generate a specific message when the state or the system parameter of the airplane meets the triggering condition of the message.

Referring to fig. 1, the method 100 further comprises: at step 160, the customized message is stored or forwarded. According to one embodiment of the invention, the customization messages may be stored in a non-volatile memory of the DMU.

According to one embodiment of the invention, the customized message may be forwarded via an Aircraft Communication Addressing and Reporting System (ACARS). ACARS is a digital data link system for transmitting messages (i.e., short messages) between an aircraft and a ground station via radio or satellite, and provides services for large-traffic data communication between the air and ground of an airline company, thereby realizing the exchange of various information.

The ACARS system consists of an avionics computer called ACARS Management Unit (MU) and a Control Display Unit (CDU). The MU is used to transmit and receive VHF radio digital messages from the ground. On the ground, the ACARS system consists of a network of ground stations 410 with radio transceiver mechanisms, which can receive or transmit messages (data chain messages). These ground stations are typically owned by various service providers that distribute received messages to servers of different airlines on the network.

On one hand, the ACARS can enable a flying airplane to automatically provide real-time data information such as flying dynamic and engine parameters to an airline ground workstation without intervention of crew members, and can also transmit other various information to the ground, so that an airline operation control center obtains real-time and uninterrupted large amount of flight data and related information of the airplane on an application system of the airline operation control center, the dynamic of the airplane of the airline company is mastered in time, the real-time monitoring of the airplane is realized, and the requirements of management of various related departments such as aviation service, operation and airplane service are met; on the other hand, the ground can provide various services such as meteorological information, airway conditions, air emergency fault troubleshooting measures and the like for the airplane flying in the air, and the flight safety guarantee capability and the service level for passengers are improved. Under the conditions that the common VHF ground-air communication channel is increasingly saturated, the information transmission quantity is small, and the speed is low, the bidirectional data communication system can remarkably improve and enhance the ground-air communication guarantee capability.

According to an embodiment of the invention, the customized message may also be forwarded based on a communication device or system of the Aviation Telecommunication Network (ATN).

Referring to fig. 1, the method 100 further comprises: at step 180, the performance of the aircraft is detected based on the customized message. The values of the parameters reflecting the operating state of the aircraft can be obtained by the customized messages. The performance of the airplane can be detected through the values of the parameters, and the 'optional' maintenance of the airplane is realized.

The concept of "optional" repair has been proposed for many years. However, there has been no good application in aircraft performance testing and aircraft maintenance. One important aspect of the present invention is that the decoding of flight data is costly and cannot be performed as a routine maintenance. The invention solves the problem through customized messages, and makes real-time monitoring of the airplane state possible. The invention fully utilizes the function of the airplane data system for acquiring the airplane running state data, can flexibly generate customized messages according to the requirements, and the messages can be conveniently used for detecting the airplane performance after being stored and forwarded.

In order to obtain a more accurate reflection of the state of the aircraft, the values of the parameters obtained directly need to be corrected. According to one embodiment of the invention, the customized message includes a primary parameter that directly reflects the status of the aircraft, and an auxiliary parameter that is used to modify the primary parameter.

When acquiring data of flight state, the time which can reflect the state of the airplane most needs to be selected to acquire the values of the main parameters or the auxiliary parameters of the customized message. Moreover, for a certain determined moment, the same parameter can be measured by adopting a mode of measuring for multiple times and averaging to obtain a more accurate result. Alternatively, the maximum and minimum values over a certain period of time are recorded to reflect the limit values of the main parameter, the latter auxiliary parameter. Accordingly, by responding to one or more trigger conditions, the values of the primary and secondary parameters at one or more time instants are obtained.

According to one embodiment of the invention, the values of the main parameters and the auxiliary parameters are converted into values under specified conditions to facilitate the detection of the performance of the aircraft.

How to select an appropriate main parameter from a plurality of airplane state parameters to form a customized message is a difficult problem, particularly for a relatively complex system, the number of parameters to be selected is large.

According to one embodiment of the invention, a plurality of parameters relating to the operating state of the aircraft are acquired, and then a plurality of primary parameters are determined according to the physical meaning represented by the plurality of parameters. After obtaining the plurality of principal parameters, some of the principal parameters may be highly correlated. In this case, a change in one primary parameter may be representative of a change in another primary parameter. Therefore, the main parameters of the highly correlated parts can be removed by checking the correlation degree. According to one embodiment of the invention, a correlation between a plurality of main parameters is calculated; and removing one or more of the plurality of main parameters based on the correlation between the plurality of main parameters. By this step, a plurality of relatively independent main parameters reflecting the operating state of the aircraft component are obtained.

According to one embodiment of the invention, the main parameters can be determined by data mining by using the existing data of a plurality of parameters related to the state of the aircraft when the fault event of the aircraft component occurs, and determining which parameters are highly related to the fault event of the aircraft.

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. Calculating the correlation between the parameter associated with the aircraft state 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 parameter related to the aircraft state and the parameter representing the fault event. Various partial correlation analysis methods in statistics can be applied to the present embodiment. 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 main parameter. By verifying all parameters related to the aircraft state in this way, the main parameters reflecting the aircraft state can be obtained.

The value of the threshold determines the final main parameter 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 of which is not highly correlated, 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.

Thus, the main parameters of a customized message may be determined by the following steps: obtaining a plurality of parameters related to the aircraft operating state; and correlating the parameters with the aircraft fault event, calculating the correlation between the parameters and the aircraft fault event through data mining, and determining a plurality of main parameters. Likewise, the degree of correlation between the plurality of main parameters may be calculated; then, one or more of the plurality of primary parameters are removed based on the correlation between the plurality of primary parameters.

The following describes the determination of the main parameters and the auxiliary parameters of the customized message by a specific example:

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.

According to one embodiment of the invention, the operation state of the APU can be detected by generating a customized APU message. When determining the main parameters of the customized APU message, a plurality of parameters are related to the operation condition of the APU due to the relative complexity of the APU system. For example, the parameters during the engine start-up phase include EGT temperature, IGV opening angle, compressor inlet pressure, load compressor inlet temperature, bleed air flow, bleed air pressure, oil temperature, APU generator load. The parameters of the APU during starting comprise starting time, EGT peak value, rotating speed at the EGT peak value, inlet temperature of a load compressor and the like.

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. 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.

The generation of customized messages may also be customized. According to one embodiment of the invention, first a first process is initiated in the aircraft data system for monitoring whether the trigger condition is met. And if the triggering condition is met, the aircraft data system enters a customized message task processing state to complete the related processing of the customized message. One benefit of such a design is that the data system of the aircraft need not monitor the task of the customized message in the non-triggered state, to conserve system resources. When the trigger condition is met, the task of processing the customized message is started.

According to one embodiment of the invention, the first process initiates a second process for monitoring whether a condition for generating a customized message is satisfied, and a third process for collecting values of the main parameter and the auxiliary parameter required for the customized message. And when the second process determines that the condition for generating the customized message is satisfied, generating the customized message by using the values of the main parameter and the auxiliary parameter collected by the third process. In this way, flexible control of the conditions under which customized messages are generated can be achieved, so that customized messages are generated only if specific conditions are met.

This customized message generation is illustrated below by a specific example:

the engine bleed system is the prerequisite of guaranteeing safe and reliable work of systems such as aircraft air conditioner, pressure boost, big wing anti-icing, hydraulic pressure. However, the high occurrence rate, high repetition rate, long removal time and great difficulty have long been a problem which plagues the maintenance of the aircraft. At present, the conventional method cannot timely and accurately process data, so that fault elimination is not timely, flight interruption is possibly caused, and even major unsafe events in flight occur.

By means of the method and the system, the bleed air system of the engine can be detected in a manner of generating the customized bleed air message. In order to reflect the performance of the engine bleed air system, it is necessary to detect the outlet temperatures of the left and right engine precoolers. If the temperature is too high or too low, this is an indication that the engine bleed air system may be malfunctioning. Thus, the main parameters of the customized bleed air messages include: the outlet temperatures of the left and right engine precoolers are above 220 degrees or below 155 degrees for a time duration and the outlet temperatures of the left and right engine precoolers during the time duration or the maximum or minimum of the outlet temperatures of the left and right engine precoolers during the time duration. The auxiliary parameters of the customized bleed air messages may include: the altitude of the aircraft and the ambient temperature.

According to one embodiment of the invention, a first process is initiated in an aircraft system to determine whether an aircraft is in a takeoff, climb or descent phase. If the aircraft is judged to be in the takeoff, climbing or descending stage, the second process and the third process are started. Wherein the second process judges whether the outlet temperature of the left and right engine precoolers is higher than 220 ℃ or lower than 155 ℃ and lasts for more than 5 seconds; wherein the third process collects the outlet temperature of the left and right engine precoolers, the altitude of the aircraft and the ambient temperature each second.

If the outlet temperatures of the left and right engine precoolers are higher than 220 degrees or lower than 155 degrees and the duration is longer than 5 seconds, the customized bleed air messages are generated according to the outlet temperatures of the left and right engine precoolers, the altitude of the aircraft and the outside temperature collected by the third process and the corresponding duration information.

The customized bleed air messages may be stored in the DMU for review and printing by flight crew or maintenance personnel. The customized bleed air messages can also be transmitted to a server of an airline company through the ACARS system, so that the performance of the bleed air system of the aircraft can be monitored in real time. The airline on the ground can make decisions based on the detected performance of the engine bleed air system and even request maintenance for the aircraft to land.

According to one embodiment of the invention, the triggering condition or the message generation condition of a customized message is modifiable. For example, the trigger condition of the customized message is configured such that each flight of the aircraft generates a customized message. One benefit of doing so is the ease of obtaining a large amount of selected flight data. For many performance testing or maintenance models, a large amount of real data is required for training and learning. And customized messages are the best way to provide these training data.

After the customized messages of multiple flights are collected, the flight state data provided by the customized messages of multiple flights is utilized, and the performance of the airplane can be detected based on an actual physical model, a characteristic evolution model or an intelligent model.

The actual physical model is a model created by using actual physical characteristics of an aircraft component. The model can actually reflect the actual condition of the aircraft performance.

The characteristic evolution model is a model reflecting the condition of the performance of the airplane through the degradation rate of the performance of the airplane. The model is a model built using known failure modes and may also substantially truly reflect the performance of the aircraft.

By intelligent model is meant a "smart" model that does not require precise mathematical and physical models, but rather is formed through the learning or training of large amounts of data. Neural network models are common intelligent models.

Different models may be built for different components of the aircraft to reflect the state of those components. These models can be facilitated using customized messages. Moreover, the data of the customized message can be judged based on the models, so that the detection of the airplane performance is realized.

According to an embodiment of the invention, after the performance of the airplane is detected by using the airplane state detection method of the above embodiment of the invention; if the detection indicates that the performance of the aircraft has failed, the aircraft may be serviced immediately. If the detection shows that the performance of the airplane only enters the decline period, the airplane is maintained at a proper time, and accordingly 'optional' maintenance of the airplane is achieved.

The customized message-based aircraft performance detection method of the present invention is further illustrated by three specific examples below.

Application example of unit oxygen system：

FIG. 2 is a schematic diagram 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 achieved according to one embodiment of the present invention.

For a crew oxygen system, the primary parameters are relatively easy to obtain. The oxygen pressure of the oxygen cylinder in the crew oxygen system is the best main 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 k₂Can be obtained by actual measurement or by statistical analysis.

According to one embodiment of the present invention, k =1 may be taken. Equation (2) can be rewritten as:

while the oxygen temperature thus derived may not be as accurate as equations (2) and (3), 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 20 ℃. Of course, other temperatures may be used.

FIG. 3 is a flow chart of a method of detecting crew oxygen system performance according to one embodiment of the present invention. In a method 300 of detecting performance of a crew oxygen system as shown in fig. 3, oxygen pressure data, atmospheric temperature, and cabin temperature of oxygen cylinders in the crew oxygen system are obtained at step 310. In step 320, 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 330, the generated crew oxygen message is transmitted to a server for processing the crew oxygen message. In step 340, the server converts the oxygen pressure of the oxygen cylinder 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 20 ℃. Of course, other temperatures may be used.

As shown in FIG. 3, in step 350, multiple sets of normalized pressure data of the unit oxygen system at different times are obtained in the manner of step 310 and step 340. After a plurality of sets of standard state pressures of oxygen in oxygen bottles in the crew oxygen system at different times at standard temperatures are obtained, the performance of the crew oxygen system can be determined by processing and evaluating the data. FIG. 4 is an example of a customized group oxygen message according to one embodiment of the invention.

In step 360, the multiple sets of normalized pressure data at different times are analyzed to determine if the crew oxygen system performance is degraded. Alternatively, in step 370, 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 100PSI, 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 （6）

wherein, P_sIs a standard pressure，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. In order to reduce the influence of the accuracy of the outside temperature and the accuracy of the pressure sensor, according to one embodiment of the invention, the air section 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 may be used, for exampleA time interval of greater than 12 or 36 hours. 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 are calculated over a 3 day period_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 crew oxygen performance occurs over a period of time, the data from the set of time periods may be taken as a set of samplesThen, the process is carried out; 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. Suppose 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 cylinder 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 cylinder 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. 5 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. 5, the broken lines represent the standard state pressures of the actual sampling conversion, and the straight lines represent the lines regressing from the standard state pressures of oxygen and the measurement time, respectively. The formula (9) of the slope detection method is adopted for detection, and the leakage rate of the oxygen system of the unit is too large, the slope is-0.024929, and the slope 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. 6 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. 6 represent the normalized pressure for the actual sample transition. Fig. 7 is a graph illustrating the rolling average leakage rate of the crew oxygen system for 24 hours and3 days as a function of the measurement time according to the embodiment of fig. 6. 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. 5 to fig. 7, 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. 8 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 800 of aircraft crew oxygen system maintenance as shown in fig. 7, oxygen pressure data, atmospheric temperature, and cockpit temperature of oxygen cylinders in the crew oxygen system are obtained at step 810. In step 820, 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 830, the generated crew oxygen message is transmitted to the server. In step 840, the server processes the crew oxygen messages to obtain the standard pressure of the oxygen cylinders in the crew oxygen system at the oxygen standard temperature. In step 850, it is determined whether the crew oxygen system performance is degraded based on the sets of standard pressure data at different times. In step 860, if the crew oxygen system performance is degraded, a suitable time is scheduled for maintenance of the crew oxygen system.

Application example of airplane landing quality detection

The term "Heavy Landing" or "Hard Landing" refers to a Landing event in which the speed or acceleration of the aircraft in the vertical direction exceeds a threshold value when the aircraft lands, wherein Heavy Landing refers to an overrun when the aircraft Landing weight is greater than the maximum Landing weight, and Hard Landing refers to an overrun when the aircraft Landing weight is equal to or less than the maximum Landing weight. Heavy or hard landings can cause strong impacts and vibrations to the structure of the aircraft, particularly to the aircraft components such as the wings, landing gear, and engines that are subjected to large loads, causing damage to the aircraft structure. Therefore, once a heavy landing occurs, the airline must perform a strict safety check on the aircraft to ensure aviation safety.

According to the rules of the aircraft manufacturing company, the responsible subject for reporting the occurrence of a hard or hard landing event on an aircraft is the flight crew. However, the heavy or hard landing event reported by the flight crew has a large uncertainty. The final outcome of the majority of hard or hard landing events reported by the flight crew is "no hard or hard landing occurred". However, the entire process results in aircraft outages and a substantial waste of maintenance resources.

Thus, in the prior art, once a flight crew reports a hard or heavy landing event, the maintenance crew can only provide the raw flight data to the aircraft manufacturer for analysis. This approach is not only costly, but also has a long latency that affects the normal flight of the aircraft.

According to one embodiment of the invention, a heavy landing event of an aircraft may be detected via customized landing messages. The landing data in the customized landing message includes, but is not limited to, the following:

1. RALT (radio altitude ft), RALR (vertical velocity ft/sec), PTCH (pitch angle deg), PTCR (pitch rate deg/sec), ROLL (ROLL angle deg), ROLR (ROLL rate deg/sec), and YAW (YAW rate deg/sec) values 1 second before the aircraft is grounded;

2. RALT (radio altitude ft), RALR (vertical velocity ft/sec), PTCH (pitch angle deg), PTCR (pitch rate deg/sec), ROLL (ROLL angle deg), ROLR (ROLL rate deg/sec), and YAW (YAW rate deg/sec) values when the aircraft is grounded;

3. maximum and minimum values of VRTA (vertical load), LONA (longitudinal load), and LATA (lateral load) during the period from 1 second before grounding; and

4. maximum and minimum values of VRTA (vertical load), LONA (longitudinal load), LATA (lateral load) within 1 second before grounding to 3 seconds after grounding.

It should be noted that the data acquired by the ACMS system is measured immediately and stored in the data cache. When triggered by a set trigger condition, it is fully possible and can be realized to acquire relevant data before the trigger condition from the data cache.

FIG. 9 is a flow diagram of a method for detecting aircraft landing quality according to one embodiment of the invention. As shown in the figure, the method 900 for detecting the landing quality of an aircraft of the present embodiment includes: at step 910, it is determined whether the vertical velocity of the aircraft at ground exceeds a predetermined value. If the predetermined value is not exceeded, at step 920, no landing message need be generated.

By setting the appropriate predetermined value of vertical velocity at step 920, it can be ensured that all data of suspected re-landing or hard landing events are recorded. According to one embodiment of the invention, the absolute value of the predetermined value of vertical velocity is less than or equal to 0.5 ft/s. The setting of the preset vertical speed value can ensure that the airplane collects and generates the landing message every time the airplane lands, even if the airplane lands normally at the moment.

Another advantage of setting the predetermined value of the vertical rate is that the trigger condition for generating the landing message can be flexibly changed, and the user can collect and record the landing state of the airplane according to the actual need, rather than collecting and recording only the data related to heavy landing or hard landing or collecting and recording the data each time the airplane lands. For example, a predetermined value of the vertical velocity may be reduced, for example 20% -40% below the limit value of the vertical velocity, so that as long as the landing is heavy, data is collected and recorded, resulting in a landing message.

If the vertical velocity at landing exceeds a predetermined value, at step 930, landing data is collected. Next, in step 940, a landing message is generated according to the collected landing data. At step 930, landing data may be collected using the ACMS system of the aircraft. The DMU of the ACMS initiates a corresponding landing data collection procedure according to a specific trigger condition. After the data acquisition is completed, in step 940, a landing message is generated according to the acquired landing data.

At step 950, the landing message is stored or forwarded. At step 960, it is determined whether a hard or heavy landing has occurred while the aircraft is landing, based on the landing data in the landing message.

According to one embodiment of the invention, whether a heavy or hard landing has occurred is determined by the speed or acceleration of the aircraft in the vertical direction when landing exceeding a limit value. The limit value of the vertical velocity of the aircraft is related to the landing weight of the aircraft, considering the limit of the structural strength of the aircraft. In the determination of whether the vertical rate (RALR) is exceeded, a comparison is made according to the landing weight of the aircraft. According to one embodiment of the invention, the limit is-9 ft/sec when the aircraft landing weight is less than the maximum landing weight. In the case where the aircraft landing weight is greater than the maximum landing weight, the limit is-6 feet/second. The above is merely an example and the limit values may be different for different aircraft landing weights less than or greater than the maximum landing weight.

As with the logic for determining the vertical rate overrun, the limit value for the aircraft vertical load is also related to the landing weight of the aircraft in determining whether the vertical load VRTA is overrun. According to one embodiment of the invention, in the case where the aircraft landing weight is less than the maximum landing weight, the limit value is 2.6G; in the case where the aircraft landing weight is greater than the maximum landing weight, the limit value is 1.7G. The above is merely an example and the limit values may be different for different aircraft landing weights less than or greater than the maximum landing weight.

Whether heavy landing or hard landing occurs can be directly judged when the vertical speed and the vertical acceleration of the aircraft during landing exceed or approach the limit values. Even if the judgment cannot be directly made, the method can provide a very valuable reference for judging whether the aircraft is in the heavy landing or the hard landing, and if the method is combined with the report of the flight personnel and other factors, whether the aircraft is in the heavy landing or the hard landing can be determined without sending the original data to the airline company for processing.

FIG. 10 is a flow diagram of a method for generating a landing message using an aircraft ACMS system according to one embodiment of the invention. As shown in the figure, the landing data collecting method 1000 of the embodiment includes: at step 1010, a determination is made as to whether the aircraft is grounded. According to one embodiment of the invention, whether the aircraft has grounded is determined by detecting whether the left and/or right main landing gear shock strut of the aircraft has transitioned from an extended state to a compressed state.

If the aircraft is already grounded; then at step 1020 it is determined whether the vertical velocity and vertical acceleration of the aircraft upon landing exceed threshold values. Meanwhile, in step 1030, the landing data of the aircraft before the ground is reached and the landing data of the aircraft at the ground are collected, and the landing data from 1 second before the ground is reached to 3 seconds after the ground is reached. In step 1040, if any one of the vertical velocity and the vertical acceleration exceeds the threshold value, all the collected landing data is formatted to generate a landing short message. Otherwise, no landing message is generated.

Fig. 11 is a schematic diagram of a trigger relationship for generating a landing short message in an ACMS system according to an embodiment of the present invention. The flip-flop shown in fig. 11 may be used in the method shown in fig. 9. As shown in fig. 11, in the DMU, top-level service TOPSERV is a trigger of system reservation, which is equivalent to a main thread of a processor or a basic service in an operating system. All other triggers are initiated or activated by TOPSERV. Immediately before the aircraft LANDs, the TOPSERV in the DMU activates the trigger lan d1 to monitor whether the aircraft has grounded during the FINAL APPR phase when the slat has been paid out at greater than 5 degrees and the altitude is less than 10000 feet.

When LAND1 detects a change in the status of the compression approaching electric door on any of the left and right main landing gears, the flag "aircraft grounded". While LAND1 activates the flip-flops LAND2 or LAND2B, and LAND3 and LAND 4. Both LAND2 and LAND2B are used to determine whether the vertical velocity (RALR) and vertical acceleration (VRTA) of the aircraft ground exceed threshold values. LAND1 activated LAND3 and LAND4 recorded landing data.

After the LAND4 is executed, all parameters in the short message are acquired, then the format of the parameters is converted, so that the parameters are convenient to print and read, and finally the landing short message is generated.

In accordance with one embodiment of the present invention, the LAND1 is operable to read the position of the left and right main landing gear shock strut adjacent the electric door during landing of the LAND1 aircraft. The detection frequency was 32 times/second to detect whether the variation occurred within 1/32 seconds. If the value of the parameter indicative of the position state changes from 0 to 1, this indicates that any one of the shock strut is returning from the extended position to the compressed position. Thus, it is judged that the aircraft has landed. This is exactly the starting point for the moment when the aircraft lands.

According to one embodiment of the present invention, LAND2, LAND2B determine whether the vertical velocity (RALR) and vertical acceleration (VRTA) of the aircraft ground exceed threshold values in the following manner. In order to reflect the landing state of the aircraft more accurately, it is necessary to determine whether RALR and VRTA within 0.5 second before and after the landing time exceed threshold values.

In this embodiment, LAND2 is first activated. LAND1 outputs a landing time value T₀，T₀Is an integer between 0 and 32. LAND2 reaction of T₀Comparing with a fine tuning parameter CHK with a value range of 0-5, if T is₀/2-CHK<0, which indicates that the grounding time is too close to the parameter measurement time, and there is a possibility that the data change due to grounding is not reflected in the measured parameter, LAND2B is activated, and it is determined whether RALR and VRTA exceed the threshold value within the next second of the grounding time, and LAND2 is ended. T is₀/2-CHK>0, LAND2 determines whether the RALR and VRTA at the landing time exceed the threshold value. If not, LAND2 will T₀Comparing with 16, judging whether T is₀-16>0. If T is₀-16<And 0, in order to more accurately reflect the landing condition of the airplane, activating LAND2B, judging whether the RALR and the VRTA within the next second of landing time exceed threshold values, and ending LAND 2. If the RALR and the VRTA at the landing moment exceed the threshold value, which is found by comparing the LAND2 with the LAND2B, it is said that the landing condition of the airplane is consistent with the situation of generating the landing message.

In the embodiment, the two triggers operate in different time periods, so that whether the RALR and the VRTA which land within 0.5 second before and after the grounding point time exceed the limit or not can be accurately judged.

According to one embodiment of the invention, the determination of whether the vertical load, i.e., the vertical acceleration VRTA, is overrun is conditional. When the forward vertical rate RALR is not overrun, the trigger will further determine the overrun condition of the vertical load VRTA. And if the RALR has found that the vertical rate exceeds the limit, skipping the judgment of the exceeding of the vertical load (VRTA) and directly generating a landing short message.

According to one embodiment of the present invention, the vertical rate (RALR) is calculated in LAND2 and LAND2B as follows. On board the aircraft, the RALR sampling rate is 16 times/second. In order to reflect the true RALR more accurately, the measured RALR needs to be corrected based on the pitch, roll attitude, triaxial acceleration and constants of the aircraft, based on the vertical velocity-IVV detected by the ADIRU (air data and inertial navigation computer).

According to one embodiment of the invention, the RALR may also be calculated using the following program segment:

IVV = IVV actual sample n (current vertical velocity value)

IVV1= IVV previous sample n-1 (vertical rate previous sample value)

PTCHACC = PTCHACC actual sample n (current value of pitch acceleration, intermediate variable)

PTCHACC1= PTCHACC previous sample n-1 (previous sample value of pitch acceleration, intermediate variable)

PTCHACC2= PTCHACC previous sample n-2 (2 nd sample value before pitch acceleration, intermediate variable)

PTCR: pitch rate

PTCR 1: pitch rate previous sample value

PTCHRAW Pitch (intermediate variable)

PTCHRAW 1: previous sample value of pitch (intermediate variable)

VACC: vertical acceleration (from inertial navigation)

RALT radio altitude

PTCH pitch

Constants (constant):

d geometric correction factor for ROLRft/deg (geometric correction of roll Rate, default of "0")

DX left arm correction (X-axis) for R/ARALT ft (altitude X-axis force arm correction, 321 airplane: 28.8/320 airplane: 18/319 airplane: 18.5/318 airplane: 16.8)

DZ lever arm correction (Z-axis) for R/ARALT ft (altitude Z-axis force arm correction, 321 airplane: 7.8/320 airplane: 7.1/319 airplane: 7.2/318 airplane: 7.6)

DXTPIR left arm correction (X-axis) for PTC + H ft (pitching X-axis force arm correction, 321 airplane: 53.1/320 airplane: 39/319 airplane: 33.8/318 airplane: 29.5)

FC filter frequency Hz (Filter frequency, default "0.3")

K1 filter constant (Filter constant, Default "5.2"

K2 filter constant (Filter constant, default "25")

K3 filter constant (Filter constant, default "5")

THETA0 average PTCH at touchdown deg (average value of pitching on earth, 321 airplane: 4.5/320 airplane: 6/319 airplane: 2/318 airplane: 6)

Initializing parameters:

PTCHRAW1=0.0

PTCHACC1=0.0

PTCHACC2=0.0

PTCR1=0.0

EN1=0.0

VZN1=IVV/60.0

ZN1=RALT

PTCHRAW=(PTCR-PTCR1)/T （T=1/16）

PTCHACC=PTCHACC1+T*(2*PI*FC)*(PTCHRAW+PTCHRAW1-PTCHACCI-PTCHACC2)/2（PI=3.14159265）

NZTCOR=VACC*9.81/0.3048-DXTPIR*PTCHACC/57.3*cos(PTCH/57.3)

HRACOR=RALT+DX*(sin(PTCH/57.3)-sin(THEATA0/57.3))-DZ*(cos(PTCH/57.3)-cos(THEATA0/57.3))

EPSN =ZN1-HRACOR

EN =EN1+T*(K3*EPSN)

VZN =VZN1+T*(ZTCOR-EN-*EPSN)

VZNU =VZN-D*ABS(ROLR)

RALR =VZNU

ZN=ZN1+T*(VZN–K1*EPSN)

the above is a process of calculating 1 RALR sample value, and the rest sample values are calculated by a loop iteration method after the first sample is calculated, and the iteration method after the first sample is calculated is as follows:

EN1=EN

VZN1=VZN

ZN1=ZN

PTCHACC2=PTCHACC1

PTCHACC1=PTCHACC

PTCHRAW1=PTCHRAW

PTCR1=PTCR

according to one embodiment of the present invention, the vertical acceleration in LAND2 and LAND2B can be obtained directly from the vertical load obtained in the ACMS system.

According to one embodiment of the invention, LAND3 implements the following functions:

a) recording RALT, RALR, PTCH, PTCR, ROLL, ROLR and YAW values 1 second before the landing point;

b) landing site RALT, RALR, PTCH, PTCR, ROLL, ROLR, and YAW values are recorded.

According to one embodiment of the invention, the LAND4 runtime is 4 seconds, recording the maximum and minimum values of VRTA, LONA, LATA, and RALR from 1 second before the landing site to 3 seconds after the landing site.

FIG. 12 is a flow diagram of a method for generating a landing message using an aircraft ACMS system according to another embodiment of the invention. The ground bounce may bounce the aircraft after landing and then the aircraft may be dropped back onto the ground. This phenomenon is known as "bouncing" of the aircraft. This bounce of the landing of the aircraft may occur once or several times. The bounce of the landing of the aircraft is likely to be a heavy or hard landing and therefore needs to be detected. The landing message related to landing bounce of the airplane is a long landing message; and the common landing message is a landing short message.

As shown in fig. 12, the method for generating a landing length message according to this embodiment includes: at step 1210, whether the aircraft is already grounded; in step 1220, collecting landing data of the aircraft before and during grounding and landing data from 1 second before to 3 seconds after grounding; at step 1230, it is determined whether a bounce has occurred while the aircraft is landing. If no bounce occurs, in step 1240, it is determined whether the vertical velocity and vertical acceleration of the aircraft at landing exceed predetermined values; if yes, in step 1250, a landing short message is generated; otherwise, no landing message is generated.

If the aircraft bounces, acquiring landing data of the aircraft within 1 second before the aircraft is grounded again and3 seconds after the aircraft lands in step 1260; meanwhile, in step 1270, it is determined whether the vertical acceleration of the re-grounding is greater than a threshold value, and if it exceeds the limit, in step 1280, the limit value, the maximum value at the time of the exceeding, the trigger code, and the trigger reason are recorded. In step 1290, the secondary landing data is formatted to generate a landing length message.

According to one embodiment of the invention, whether the aircraft is bouncing during landing is determined by determining whether the two main landing gears have been compressed and held for a sufficient time, and then determining whether the left and right main landing gears are in an extended state again. It is further determined that the left and right main landing gears are again in the extended state for less than 10 seconds to confirm re-grounding and thus further confirm the occurrence of bounce.

Fig. 13 is a schematic diagram of a trigger relationship for generating a landing length message in an ACMS system according to an embodiment of the present invention. The flip-flop shown in fig. 13 may be used in the method shown in fig. 12. As shown in fig. 13, in the DMU, the top-level service TOPSERV is a trigger for system reservation. Immediately prior to the aircraft landing, the TOPSERV in the DMU activates the triggers bound 1 and LAND1 during the FINAL APPR phase when the slat is paid out at greater than 5 degrees and the flight height is less than 10000 feet. Bound 1 detects whether both main landing gears have compressed and remain in place for a sufficient time.

If BOUNCE1 confirms that the main landing gear has been compressed and held for a sufficient time, BOUNCE1 activates triggers BOUNCE2 and BOUNCE3 to detect whether the left and right main landing gears are again extended, respectively. Then, the bound 2 and bound 3 activate the corresponding trigger bound 4 or bound 5 to further confirm the flight status of the airplane. The BOUNCE4 and the BOUNCE5 continuously detect the extension states of the shock-absorbing struts of the left landing gear and the right landing gear, and judge that the airplane BOUNCEs when certain conditions are met.

The BOUNCE4 and the BOUNCE5 respectively activate the triggers BOUNCE6 and BOUNCE7 to search, compare and collect landing data of the airplane in the first 1 second of re-grounding and3 seconds after landing.

The bouncer 7 also searches and compares whether the vertical acceleration of the earth again after bouncing is larger than a limit value, and if the vertical acceleration exceeds the limit value, the maximum value during exceeding, a trigger code and a trigger reason are recorded in the message.

The manner in which bound 6 and bound 7 obtain landing data is similar to LAND3 and LAND4, and is not described here in detail.

The LAND1 is used to monitor whether the aircraft is grounded. If grounded, LAND1 activates the trigger BOUNCE 8. Based on whether the aircraft BOUNCEs when landing, bound 8 determines whether to generate a long landing message or a short landing message. And finally, converting the formats of the landing related parameters of the two landings so as to facilitate the reading and printing of the numerical values recorded by the message and generate a corresponding landing message.

According to one embodiment of the invention, whether a bounce has occurred while the aircraft is landing is detected as follows. The position status of the left and right main landing gear shock strut adjacent the electric door is continuously read by bound 1 at a frequency of 32 times/second to detect if there is a change in 1/32 seconds. When its state transitions from "0" to "1", bound 1 starts a counter to increment. Bound 1 activates bound 2 and bound 3 only if the counter is greater than 16. This indicates that both main landing gears have been compressed and held for at least 0.5 seconds. If the condition is not met, the counter is cleared and re-incremented.

Taking the left main landing gear as an example below, the right main landing gear may be treated in the same manner.

Following operation of BOUNCE2, the status of the compressed electric door position of the left main landing gear continues to be detected at a frequency of 32 times/second. When the parameter value is "0", the counter starts to accumulate. Bound 4 is activated only when the counter value is greater than 32. At this point, the shock strut of the left main landing gear is extended for a duration greater than 1 second. When the condition is not met, the counter is cleared and re-accumulated.

The detection principle of bound 4 is similar to bound 2, and when the parameter value is "0", the counter starts to count up. When the parameter value is "1", the accumulated value of the counter is judged. If the accumulated value of the counter is less than 320, then it is determined that the left main landing gear is bouncing. The time during which the shock strut of the left main landing gear remains extended, i.e. the hold time, is less than 10 seconds. And then again in the compressed state.

To summarize the above, the three conditions of the method for determining bounce of the present embodiment are:

1. determining whether the left and right main landing gears are in a compressed state and last for more than 0.5 second;

2. determining whether any of the left and right main landing gears are again in an extended state for more than 1 second; and

3. it is determined whether any of the left and right main landing gears are again in an extended state for less than 10 seconds.

If the above conditions are met, the aircraft is considered to bounce during landing.

According to an embodiment of the present invention, after the BOUNCE8 runs for 30 seconds, whether a long message or a short message is generated is judged according to parameter values of BOUNCED, LONGLRPT and BRPTCODE

Bound: and the state parameter marks the occurrence of bounce of the airplane. The trigger BOUNCE4 or BOUNCE5 assigns a value after detecting a BOUNCE.

LONGLRPT: status parameters, flag long messages may be generated. When the first grounding vertical load is over-limited, the trigger LAND2/2B is assigned when the vertical load is over-limited.

BRPTCODE is a message trigger code, is assigned when the secondary grounding exceeds the limit, and is assigned after the trigger BOUNCE7 detects the exceeding.

The bound 8 calls the above parameters to determine whether to generate a long message or a short message.

See the following table for details:

FIG. 14 is an example of a customized landing short message according to one embodiment of the invention. As shown, it can be seen that the vertical rate RALR is only 1.8 ft/sec during the landing. The vertical acceleration VRTA is 1.64G, also in the range of normal landing. However, the lateral acceleration may be slightly higher, 0.21G. Under the condition, even if the flight personnel report that the landing is heavier, the landing can be easily seen according to the short landing message, and the landing is normal, and the heavy landing or the hard landing does not occur.

FIG. 15 is an example of a customized landing length message according to another embodiment of the invention. As shown in the figure, it can be found that the airplane bounces in the landing process. During the first grounding, the vertical velocity RALR is 7.2 ft/sec and the vertical acceleration VRTA is 2.07G. The vertical velocity is within the normal range and the vertical acceleration is also below the threshold. During the second grounding process, the vertical velocity RALR is 1.5 ft/sec, and the vertical acceleration VRTA is 2.65G. Thus, the aircraft bounces during landing and the vertical load is exceeded during the second landing.

The aircraft maintenance personnel can obtain the landing message from the nonvolatile memory of the DMU, also can print the aircraft landing message in the aircraft cockpit, or the performance monitoring personnel can read the landing message downloaded through the air-ground data link through the ground workstation to realize the monitoring of the aircraft landing performance, thereby ensuring to timely and accurately find the abnormality of the aircraft landing performance. Therefore, a large amount of data processing and checking work after the heavy landing or the hard landing is reported can be avoided to determine whether the aircraft has the heavy landing or the hard landing, the stop time of the aircraft is saved, and the utilization rate of the aircraft is improved; meanwhile, the airplane is prevented from running under the condition of potential safety hazard, and the potential safety hazard of the running of the airplane is eliminated. The recorded data also help the flight quality monitoring department to evaluate the operation technical quality of the flight personnel.

Although the ACMS system of the airbus company is taken as an example, the application of the present invention is not limited to the airplane of the airbus company. The present invention may also be applied to boeing aircraft using the AHM 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 (16)

1. A method for detecting airplane performance based on customized messages comprises the following steps:

collecting data reflecting the operation state of the airplane;

generating a customized message based on the collected data, wherein the customized message is customized to include one or more primary parameters related to an aircraft operating state;

storing or forwarding the customized message; and

detecting the performance of the aircraft based on the customized message;

wherein the customized message further comprises one or more auxiliary parameters associated with modifying the primary parameter;

wherein the step of generating a customized message based on the collected data further comprises: acquiring values of the main parameter and the auxiliary parameter at one or more moments in time in response to a trigger condition;

wherein the method further comprises:

initiating a first process to monitor whether the trigger condition is satisfied;

initiating a second process for monitoring whether conditions for generating a customized message are met; and initiating a third process for collecting values of said primary and secondary parameters required by a customized message; and when the second process determines that the condition for generating the customized message is met, generating the customized message by using the values of the main parameters and the auxiliary parameters collected by the third process.

2. The method of claim 1, further comprising: and converting the values of the main parameters and the auxiliary parameters into the values in the specified state.

3. The method of claim 2, further comprising: and correcting the value of the main parameter according to the value of the auxiliary parameter.

4. The method of claim 1, wherein said principal parameter of said customized message is determined by the steps of:

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

determining a plurality of main parameters according to the physical meanings represented by the parameters;

calculating the correlation degree among the main parameters; and

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

5. The method of claim 1, wherein said principal parameter of said customized message is determined by the steps of:

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

correlating the plurality of parameters with an aircraft fault event, determining a plurality of primary parameters;

calculating the correlation degree among the main parameters; and

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

6. The method of claim 5, wherein the step of associating the plurality of parameters with an aircraft failure event comprises: and calculating the correlation degree of the plurality of parameters and the airplane fault event through data mining.

7. The method of claim 1, further comprising: the first process initiates the second process and the third process.

8. The method of claim 7, wherein the first process is configured to determine whether the aircraft is in a takeoff, climb, or descent phase;

wherein the second process judges whether the outlet temperature of the left and right engine precoolers is higher than 220 ℃ or lower than 155 ℃ and lasts for more than 5 seconds;

wherein the third process collects outlet temperatures of the left and right engine precoolers every second.

9. The method of claim 8, further comprising: generating a customized bleed air message, wherein the main parameters of the customized bleed air message comprise: the outlet temperatures of the left and right engine precoolers are higher than 220 degrees or lower than 155 degrees in duration, the maximum or minimum of the outlet temperatures of the left and right engine precoolers; and the auxiliary parameters of the customized bleed air messages comprise: the altitude of the aircraft and the ambient temperature.

10. The method of claim 1, wherein the customized message is a custom message other than a system message.

11. The method of claim 1 wherein said customized message is generated using an ACMS system of air passenger company or an AHM system of boeing company.

12. The method of claim 1, wherein the step of storing or forwarding the customized message comprises: and storing the customized message in a nonvolatile memory of an airplane data system or forwarding the customized message by utilizing an ACARS system or an ATN system.

13. The method of claim 1, wherein said triggering condition or said message generating condition of said customized message is modifiable.

14. The method of claim 13, wherein the trigger condition of the customized message is configured such that each flight of the aircraft generates a customized message.

15. The method of claim 14, wherein detecting the performance of the aircraft based on the customized message comprises:

collecting customized messages of multiple flights; and

and detecting the performance of the airplane by using the customized message of multiple flights based on an actual physical model, a characteristic evolution model or an intelligent model.

16. A method of servicing an aircraft, comprising:

detecting the performance of the aircraft using the method of any one of claims 1-15; and

in response to a performance fault of the aircraft, performing a repair on the aircraft; or

In response to a performance degradation of the aircraft, an appropriate opportunity is scheduled for maintenance of the aircraft.