• 专利附录
  • 专利说明
  • 权利要求
  • 类似技术
  • 同族专利
  • 引用文献
  • 法律状态
  • 优先权
    • 专利摘要:
    The invention relates to a new logic aimed, when a loss of power is detected on a motor (S_Fail (1)), to generate an alert in the form of a single message displayed on a cockpit display screen, to indicate whether the level of damage to the engine is critical or not. The logic implemented is based on alarm signals (S_Vib (1) N (Y); S_EGT (1)) transmitted by a central unit of the motor but also on alarm signals (S_Fuelleak; S_Hydroleak; S_Current ( 1); S_Smoke (1); S_Pressurecabine; S_Cool) issued by an on-board aircraft diagnostic device in order to take into account both the engine situation, but also the situation of the surrounding engine systems which may be affected by damage to an engine.
    公开号:FR3052747A1
    申请号:FR1655650
    申请日:2016-06-17
    公开日:2017-12-22
    发明作者:Antoine Pilon
    申请人:Airbus Operations SAS；
    IPC主号:
    专利说明:

    The invention relates to a method of monitoring the engines of an aircraft for informing a crew of the state of the engines of said aircraft.
    Commercial aircraft include an electronic centralized aircraft monitor (ECAM) system connected to various sensors installed on aircraft systems / engines to inform pilots via messages (a message by type of alarm) displayed on a cockpit alarm screen, the existence of malfunctions of said systems / engines. When several alarm messages concerning an engine are displayed, the pilots of the aircraft must quickly evaluate the operational situation by reading the various messages displayed on the alarm screen and decide whether or not the engine is stopped (in the case of a stopped engine, the pilots will be able to use the propulsion provided by the other engine (s) to make land the aircraft). The invention aims to improve the decision-making aid as to the stopping or not of an engine. To this end, the invention relates to a method of monitoring the engines of an aircraft comprising at least two engines, a central unit associated with each engine and configured to measure the temperature of the exhaust gas of said engine and a vibration amplitude of at least one rotor of said engine and the power of said engine, a plurality of on-board systems, an aircraft system diagnostic device configured to measure the value of at least one parameter of at least one onboard system of the an aircraft, a display screen and a monitoring device connected to each of the central units and to the diagnostic device and to said screen, said method being implemented by the monitoring device, for each engine, the monitoring device generates, if a loss of power of the engine is detected and if one of two conditions A or B is true, a signal indicative of a serious damage of said engine, a message according to which a motor has suffered serious damage being displayed on a display screen when the monitoring device sends the signal indicative of serious damage to said engine, with: condition A is true if at least: the value of at least one parameter of at least one aircraftborne system is outside a permissible range of values; and condition B is true if at least: the temperature of the exhaust gas of said engine is above a predetermined value AND a vibration amplitude of at least one rotor of the engine is beyond a predetermined value .
    Thanks to the invention, the pilots can, by reading a single message, instead of a plurality of successive messages, quickly assess the state of an engine and measure the extent of a malfunction of said engine .
    The characteristics of the invention mentioned above, as well as others, will emerge more clearly on reading the following description of exemplary embodiments, said description being given in relation with the attached FIG. 1: FIG. is a schematic view of an aircraft according to one embodiment of the invention, said aircraft comprising two engines and a monitoring device connected to a central unit of each of the engines and to a system diagnostic device of the aircraft; FIG. 2 is a diagram of a monitoring logic of an engine implemented by the monitoring device according to one embodiment of the invention; and FIG. 3 is a view similar to FIG. 2, illustrating a monitoring logic of an engine implemented by the monitoring device according to another embodiment of the invention.
    With reference to FIG. 1, the aircraft A comprises two engines 1.2 which are each arranged in a nacelle 1a, 2a fixed for example under a wing 4,5 of the aircraft, a central unit 10,20 (of type FADEC for full authority digital engine computer: engine interface system) associated with each engine 1.2 to control said engine, a diagnostic device of the onboard systems (out of engine) of the aircraft 40, said diagnostic device, a device monitoring system 50 (for example of the ECAM or EICAS type) dedicated to the generation of alerts to the crew in the event of failure of one of the engines of the aircraft and which is connected to the central unit 10, 20 of each of the motors 1,2 and the diagnostic device 40, and finally, a cockpit P comprising at least one display screen 80 for displaying the alerts, in the form of messages, generated by the monitoring device, and a man-machine interface 70 of engine control connected to the central units 10,20.
    According to the invention, the monitoring device 50 (of the central unit type) implements a logic aimed, when a loss of power is detected on an engine, in generating a message displayed on the display screen 80 indicating whether the level of damage to the engine is critical or not. The logic implemented is based on alarms issued by the engine CPU 10,20 in order to take into account the state of the engine, but also on alarms issued by the diagnostic device 40 in order to take also account the state of embedded systems that can also be damaged due to a problem encountered on the engine (example: blade debris thrown out of the engine and damaging various surrounding embedded systems to the engine).
    In a known manner, the central unit 10,20 associated with each engine 1, 2 controls the performance of the engine by controlling the operation of the engine as a function of values of the crew setpoints and engine parameter values measured by a network. sensors installed on the engine or its components installed in the platform 1a, 2a (oil pumps, fuel pump ..).
    The sensor network associated with a motor 1,2 or its components notably comprises vibration amplitude sensors of different engine parts such as compressor or turbine rotors, and exhaust gas temperature sensors (known as temperature). EGT). The central unit 10,20 of each motor 1,2 sends, for each parameter measured, a status signal to the monitoring device 50.
    This signal is for example a boolean whose setting to 1 indicates an abnormal state (alarm). For example, for the temperature of the exhaust gases, the central unit 10, 20 of an engine X (where X is equal to 1, for the engine 1, or X = 2, for the engine 2) emits, to the monitoring device 50, a Boolean signal S_EGT (X) set to 0 when the exhaust gas temperature of the engine X is within a range of allowable values, and set to 1 when said temperature is greater than the terminal upper of said range. As for the amplitudes of vibrations, the central unit 10,20 of the motor X emits, for each monitored rotor, a Boolean signal S_Vib (X) N (Y) (with Y the number of the rotor in the motor X, for example Y equal 2 for the rotor N2 of the compressor) set to 0 when the amplitudes of the vibrations are less than or equal to a maximum allowable value, and set to 1 when said amplitudes are greater than said admissible value.
    The diagnostic device 40 collects the values of various parameters resulting from measurements made by sensors arranged on on-board systems of the aircraft A, and compares these values with acceptable values in order to deduce the state of health of said systems. The on-board systems monitored by the diagnostic device are those whose operation can be impacted by an incident on a motor 1,2, for example because of an engine fire or an engine bursting projecting blade debris or debris. blades. The embedded systems are for example: the nacelles, the fuel system, the hydraulic circuit, the cabin, the pneumatic generation circuit of the aircraft, an electronic power circuit of the aircraft, etc. For example, the parameters taken into account for these systems are: • for the fuel or hydraulic system: a flow rate or a liquid level measured by mechanical sensors and making it possible to detect a liquid leak such as kerosene or else hydraulic fluid); such leakage may be due to the drilling of a pipe or tank by engine debris as a result of a problem on an engine; • for a power electronics circuit: voltages or currents measured by electronic sensors. Measurements of voltages and intensities make it possible, for example, to detect short circuits in electrical circuits that may have been impacted by a fire or whose components are damaged following the bursting of a blade / blade of an engine. ; For the nacelles 1a, 2a: temperatures at the inside of the nacelle 1a, 2a of an engine, measured by temperature sensors. The temperature measurement makes it possible to detect a fire causing a sudden rise in temperature, the fire being able to be triggered following a problem on an engine (for example: leakage of gasoline igniting due to a short circuit caused by engine debris or crankcase drilling); • for the cabin: the pressure measured by pressure sensors. A loss of pressure that may result from an engine burst with debris piercing the skin of the fuselage; and for the pneumatic generation circuit the aircraft: temperatures at the pneumatic generation circuit of the aircraft, measured by temperature sensors. Such a temperature measurement makes it possible to detect a leak of a hot air circuit of a motor or of the aircraft that may have been caused by a debris having damaged said circuit.
    These parameters are given by way of examples, but other parameters of other systems such as, for example, indications of the position of the flaps or the edge of attack, or indications of the position of the thrust reversers of an engine. can be taken into account.
    The diagnostic device 40 sends, for each parameter measured, a status signal to the monitoring device 50. This signal is for example a Boolean.
    For example: • for leaks of liquids, the diagnostic device 40 emits a Boolean signal S_Fuelleak set to a 0 when no kerosene leak is detected, and set to 1 otherwise, and a Boolean signal S Hydroleak set to a 0 when no hydraulic fluid leak is detected, set to 1 otherwise; For the intensity / voltage measurements, the diagnostic device transmits a Boolean signal S_Current (X) set to a 1 when a short circuit is detected in electronic circuits of a predetermined zone situated around the motor X, and set to 0 otherwise; For the temperature measurements at the level of the nacelle of the engine X, the diagnostic device 40 emits a Boolean signal S_Smoke (X) set to a 0 when the temperature at the nacelle of the aircraft has a value included in a range of acceptable value, and set to 1 otherwise; • for the cabin pressure measurements, the diagnostic device transmits a B S Pressurecabine signal set to 0 when the cabin pressure has a value within an acceptable range (varying according to the altitude), and set to 1 otherwise. For the temperature measurements at the pneumatic generation circuit of the aircraft, the diagnostic device transmits a Boolean signal S Cool set to 0 when the temperature at said circuit has a value within an acceptable value range, and set to 1 otherwise.
    The monitoring device 50 is an electronic device of the central unit type, for example installed in the avionics bay (not shown) of the aircraft A. The monitoring device 50 implements a similar logic for the two engines 1,2 using as inputs the signals transmitted by the central units 10,20 of the motors and by the diagnostic device 40.
    It should be noted that certain signals emitted by the diagnostic device 40, such as the signals relating to the pressure in the cabin or the temperature-related signals of the pneumatic generation circuit of the aircraft, or the signals relating to the leakage of liquids, do not occur. are not related to a particular engine. Such signals are then used in the logic implemented by the monitoring device 50 for one or the other of the motors 1,2. On the other hand, the signals linked to a particular engine, such as for example the signal relating to the temperature at the nacelle of the engine, are only used for the logic set by the monitoring device 50 for this engine.
    With reference to FIG. 2, and for the engine 1, an example of Boolean logic implemented by the monitoring device 50 is as follows: • the monitoring device 50 receives a so-called fault signal, S_Fail (1), transmitted by the central unit 10 of the engine 1. This signal is set when said unit 10 detects a power loss of the engine 1, to 0 otherwise. By loss of power of a motor 1,2, it means a thrust delivered by the engine which is below a set value (entered via the man-machine interface 70) thrust desired by the crew. A loss of power will be detected by the central unit of an engine, for example when the engine 1,2 produces only 20% of power while the thrust reference value of the crew is greater than 20%, by example 60%. The monitoring device 50 receives, from the central unit 10, the signals S_EGT (1) and S_Vib (1) N (Y) respectively relating to the temperature of the exhaust gases of the engine 1 and to the amplitude of the vibrations at least one rotor of the motor 1. An S_Eng (1) signal is formed as the output of an AND logic gate 200 having as inputs the signal S_Vib (1) N (Y) (whatever Y) and the signal S_EGT (1)). The alarm signal S_Eng (1) is thus set when the amplitude values of the vibrations of at least one rotor and the value of the temperature of the exhaust gases are abnormal. The monitoring device 50 receives, from the diagnostic device 40, the signals S Fuelleak, S_Hydroleak; S Current (1), S_Smoke (1), S_Presscabine, S_Cool. An alarm signal S Aircraft (1) is formed as the output of an OR logic gate 210 having the inputs S_Fuelleak, S Hydroleak; S_Current (1), S_Smoke (1), S_Presscabine, S_Cool. This signal S Aircraft (1) is set to 1 if at least one of the signals transmitted by the diagnostic device 40 to the monitoring device 50 is set to 1.
    The output signal of the logic implemented by the monitoring device 50 for the engine 1 is a state signal S Sevdamage (1) of the engine 1 which is indicative, when set to 1, of a situation of serious damage. of the motor 1. The signal SevDamage (1) is the output of an AND logic gate 230 receiving as input the signal S Fail (1) and a signal S_Pb (1), the signal S_Pb (1) being the output of a logic gate OR 220 receiving as input the signal S_ Aircraft (1) and the signal S_ Eng (1).
    The signal S_Shdamage (1) is thus set when the central unit 10 detects a power loss of the engine 1 and when the central unit 10 or diagnostic device 40 emits an alarm on one of the monitored parameters. When the signal S_Sdamage (1) is set to 1, the monitoring device 50 sends a control signal (instructions) to the display screen 80 so that the latter indicates, via a message, that the motor 1 has undergone a serious damage.
    Note that the same logic, with modification of the engine code, is implemented for the engine 2, although this is not shown.
    Thanks to the invention, the pilots can, by reading a single message, instead of a plurality of successive messages, quickly appreciate the situation of a motor 1,2 at the occurrence of a loss of power and make a quick decision on whether or not to shut down the engine.
    In an alternative embodiment of the invention, the logic as described can be modified so that the central unit 10,20 of each motor 1,2 takes into account parameters other than the vibration amplitudes of the rotors or the temperature engine exhaust. Thus, as illustrated in FIG. 3, the central unit 20 of the engine 1 sends, to the monitoring device 50, a signal S Mecafail (1) which is indicative, when it is set to 1, of a mechanical anomaly encountered. by the motor 1. This signal S_Mecafail (1) is for example an indicator of a defect on a transmission shaft or a damage to the rotor of the blower of the engine 1 due to an impact with a bird.
    According to this variant, the signal S_Eng (1) is the output of an OR logic gate 205 having for its inputs the signal S Mecafail (1) and the output of an AND logic gate 200 having as inputs the signal S_Vib (1) N (Y) (does not matter Y) and signal S EGT (1)). The signal S_Eng (1) is thus set when, for the engine 1, the amplitude values of the vibrations of a rotor and of the temperature of the exhaust gases are abnormal OR when an anomaly on a mechanical part is detected .
    The diagnostic device 40 has been described as collecting the values of various parameters resulting from measurements made by sensors associated with various onboard systems of the aircraft A. Without departing from the scope of the present invention, the diagnostic device 40 could collect only the value of a single parameter resulting from a single measurement made by a sensor associated with an onboard system of the aircraft A.
    The above description relates to an exemplary implementation of the invention for a twin-engine aircraft. Without departing from the scope of the present invention, the invention could be implemented in an aircraft comprising more engines, such as for example a four-engine aircraft.
    权利要求:
    Claims (2)
    [1" id="c-fr-0001]
    1) A method of monitoring the engines of an aircraft (A) comprising at least two motors (1,2), a central unit (10,20) associated with each engine and configured to measure the temperature of the exhaust gas of said engine and a vibration amplitude of at least one rotor of said engine and the power of said engine, a plurality of embedded systems, a diagnostic device (40) of the aircraft's onboard systems configured to measure the value of at least one parameter of at least one on-board system of the aircraft (A), a display screen (80) and a monitoring device (50) connected to each of the central units and the diagnostic device (40) and to said screen, said method being implemented by the monitoring device (50), characterized in that for each motor (1,2), the monitoring device (50) generates, if a loss of power of the motor is detected (S_Fail (l )) and if one of two conditions A (S_A ircraft (l)) or B (S_Eng (l)) is true, a signal (S_SevDamage (l)) indicative of serious damage to said engine, a message that a motor (1,2) has suffered serious damage being displayed on a display screen (80) on transmission, by the monitoring device (50), of the signal (S SevDamage (l)) indicative of a serious damage of said engine, with: condition A is true if at least: the value of at least one parameter of at least one aircraftborne system (A) is outside a permissible range of values; and condition B is true if at least: the temperature of the exhaust gas of said engine is above a predetermined value AND a vibration amplitude of at least one rotor of the engine is beyond a predetermined value .
    [0002]
    2) Monitoring method according to claim 1, characterized in that each central unit (10,20) associated with a motor is configured to detect a mechanical anomaly of said engine (1,2), the condition B is true if: the temperature exhaust gas from said engine (1,2) is above a predetermined value AND a vibration amplitude of at least one rotor of the engine is above a predetermined value OR a mechanical engine abnormality ( 1,2) is detected.
    类似技术:
    公开号 | 公开日 | 专利标题
    EP2202500B1|2012-11-14|Operations support systems and methods for engine diagnostic
    EP2202147B1|2017-04-26|Operations support systems and methods for power management
    FR3052747A1|2017-12-22|METHOD FOR MONITORING THE ENGINES OF AN AIRCRAFT
    AU2013206339B2|2017-02-23|Systems and methods for diagnosing an engine
    EP2207072A2|2010-07-14|Operations support systems and methods with engine diagnostics
    US8850876B2|2014-10-07|Methods and systems for monitoring engine oil temperature of an operating engine
    US10006337B2|2018-06-26|Systems and methods for diagnosing an engine
    CA2896695C|2017-03-28|Method and device for detecting frost in the air intake of a turbine engine
    EP2820246A2|2015-01-07|Debris detection in turbomachinery and gas turbine engines
    FR2942001A1|2010-08-13|SYSTEM FOR MONITORING THE HEALTH STATUS OF EQUIPMENT INVOLVED IN STARTER CAPACITY OF A TURBOREACTOR
    US10962448B2|2021-03-30|Method for monitoring the engines of an aircraft
    EP2202394A2|2010-06-30|Gas turbine engine operations support systems and methods with model-based torque estimates
    FR2998058A1|2014-05-16|DEVICE AND METHOD FOR PROTECTING AN AIRCRAFT TURBO-MACHINE COMPUTER AGAINST SPEED MEASUREMENT ERRORS
    FR2710026A1|1995-03-24|Device for monitoring the propulsion system of an aircraft
    CN110566289A|2019-12-13|Method of operating a turbomachine
    EP2662543A2|2013-11-13|Device for controlling the tightness of a rotorcraft gearbox by depressurisation
    EP1753939B1|2009-08-05|Overspeed limiter for turboshaft engines
    EP3421375B1|2019-08-14|Monitoring system of an aircraft turbine engine
    US10677170B2|2020-06-09|Apparatus and method for detecting a threshold vibration condition in a gas turbine engine
    US20210052999A1|2021-02-25|Method and system for monitoring a fluid system configured to operate with a filter
    EP3155232B1|2018-12-12|Method for assisting with the detection of damage to a turbojet duct
    Vega et al.2018|Prevention of Compressor Stall on Rolls Royce M250 Engines C20 Series Using Trend Monitoring Analysis
    Jagadish Babu et al.2019|Analysis and Prevention of Failures in a Turboprop Engine
    Murphy1978|Diagnostic system requirements for helicopter propulsion systems
    FR3075877A1|2019-06-28|SYSTEM FOR DETECTING A TEMPERATURE FAULT IN THE INTER-VEIN COMPARTMENT OF A DOUBLE FLOW TURBOREACTOR
    同族专利:
    公开号 | 公开日
    CN107526347A|2017-12-29|
    FR3052747B1|2018-05-25|
    US10254199B2|2019-04-09|
    CN107526347B|2019-11-05|
    US20170363514A1|2017-12-21|
    引用文献:
    公开号 | 申请日 | 公开日 | 申请人 | 专利标题
    FR2821452A1|2001-02-26|2002-08-30|Eads Airbus Sa|DEVICE FOR MONITORING A PLURALITY OF AIRCRAFT SYSTEMS, ESPECIALLY A TRANSPORT AIRCRAFT|
    US20130199204A1|2012-02-06|2013-08-08|Eurocopter|Method and a device for performing a check of the health of a turbine engine of an aircraft provided with at least one turbine engine|
    US3987279A|1975-04-22|1976-10-19|The Boeing Company|Automatic performance reserve system|
    CN102183575B|2011-02-21|2013-01-02|中国民航大学|Composite sensor and composite-sensor-based aircraft engine gas circuit fault detection and diagnosis method|
    US8626371B2|2011-09-15|2014-01-07|General Electric Company|Systems and methods for diagnosing auxiliary equipment associated with an engine|
    CN102567639A|2011-12-30|2012-07-11|南京航空航天大学|Method for evaluating reliability of aircraft engine aiming at competing failure|
    US20130197739A1|2012-01-31|2013-08-01|Gulfstream Aerospace Corporation|Methods and systems for aircraft health and trend monitoring|
    CN104700547B|2014-07-11|2017-08-25|成都飞亚航空设备应用研究所有限公司|A kind of fire alarm prior-warning device for aircraft engine|
    US20160160776A1|2014-12-08|2016-06-09|Caterpillar Inc.|Engine System and Method|
    US20170345318A1|2016-05-25|2017-11-30|General Electric Company|Aircraft control system|US10962448B2|2016-06-17|2021-03-30|Airbus Operations Sas|Method for monitoring the engines of an aircraft|
    FR3068392A1|2017-06-29|2019-01-04|Airbus Operations |DEVICE FOR MONITORING A TURBOMACHINE OF AN AIRCRAFT|
    US20200038697A1|2018-07-31|2020-02-06|Telesafe LLC|System and a method for monitoring engine conditions|
    法律状态:
    2017-06-21| PLFP| Fee payment|Year of fee payment: 2 |
    2017-12-22| PLSC| Search report ready|Effective date: 20171222 |
    2018-06-26| PLFP| Fee payment|Year of fee payment: 3 |
    2020-06-19| PLFP| Fee payment|Year of fee payment: 5 |
    2021-06-22| PLFP| Fee payment|Year of fee payment: 6 |
    优先权:
    申请号 | 申请日 | 专利标题
    FR1655650A|FR3052747B1|2016-06-17|2016-06-17|METHOD FOR MONITORING THE ENGINES OF AN AIRCRAFT|
    FR1655650|2016-06-17|FR1655650A| FR3052747B1|2016-06-17|2016-06-17|METHOD FOR MONITORING THE ENGINES OF AN AIRCRAFT|
    CN201710348117.2A| CN107526347B|2016-06-17|2017-05-17|The monitoring method of the engine of aircraft|
    US15/621,416| US10254199B2|2016-06-17|2017-06-13|Method for monitoring the engines of an aircraft|
    US16/225,285| US10962448B2|2016-06-17|2018-12-19|Method for monitoring the engines of an aircraft|
    [返回顶部]