Patent Publication Number: US-10309249-B2

Title: Control apparatus for a gas-turbine aeroengine

Description:
BACKGROUND OF THE INVENTION 
     Field of the Invention 
     An embodiment of this invention relates to control apparatus for a gas-turbine aeroengine. 
     Description of the Related Art 
     A gas-turbine aeroengine is typically equipped with at least a high-pressure turbine rotated by injection of high-pressure gas produced upon ignition and combustion of an air-fuel mixture in a combustion chamber and with a low-pressure turbine located downstream of the high-pressure turbine to be rotated by low-pressure gas exiting the high-pressure turbine. Such a gas-turbine aeroengine is provided with sensors or detectors for detecting numerous operating parameters used to control the engine, including a low-pressure turbine rotational speed N1, a high-pressure turbine rotational speed N2, and an outlet pressure P3 of a high-pressure compressor connected to the high-pressure turbine. 
     As the control is disturbed by abnormalities arising in these sensors, each or a relatively important one of the sensors is preferably monitored for malfunctioning by estimating (calculating) the operating parameter based on the output(s) of the other sensor(s) and comparing the estimated operating parameter with the outputs of the sensor(s). 
     Therefore, as taught by Japanese Laid-Open Patent Application No. 2006-9684 (Patent Document 1), it has been proposed to use the relationship between the outputs of the high-pressure turbine rotational speed sensor and an intake air temperature sensor to calculate an estimated value of the low-pressure turbine rotational speed N1 as an operating parameter and to discriminate the normality of the low-pressure turbine rotational speed sensor by comparing the calculated operating parameter with the output of the low-pressure turbine rotational speed sensor. 
     SUMMARY OF THE INVENTION 
     The technique set forth in Patent Document 1 can calculate an estimated value of the low-pressure turbine rotational speed N1. However, when the low-pressure turbine rotational speed sensor fails in a situation where breakage of a fan blade in the engine or other such mishap has occurred, a risk of low-pressure turbine overspeed arises, making it essential to discriminate the normality of the low-pressure turbine rotational speed sensor and prevent low-pressure turbine overspeed. 
     Therefore, an object of this invention is to resolve the aforesaid issue by providing a control apparatus for a gas-turbine aeroengine which discriminates normality of a low-pressure turbine rotational speed sensor and prevents low-pressure turbine overspeed even when the sensor is abnormal. 
     In order to achieve the object, this invention provides in its first aspect an apparatus for controlling a gas-turbine aeroengine mounted on an aircraft and having at least a high-pressure turbine rotated by injection of high-pressure gas produced upon ignition and combustion of an air-fuel mixture in a combustion chamber, and a low-pressure turbine located downstream of the high-pressure turbine to be rotated by low-pressure gas exiting the high-pressure turbine, comprising: a low-pressure turbine rotational speed sensor adapted to detect a rotational speed of the low-pressure turbine; a high-pressure turbine rotational speed sensor adapted to detect a rotational speed of the high-pressure turbine; a low-pressure turbine rotational speed sensor normality discriminator that discriminates whether or not the low-pressure turbine rotational speed sensor is normal; and a controller that establishes a first value as an upper limit value of the rotational speed of the high-pressure turbine and controls the rotational speed of the high-pressure turbine based on the established upper limit value; wherein the upper limit value changer changes the upper limit value to a second value that is lower than the first value, when the low-pressure turbine rotational speed sensor normality discriminator discriminates that the low-pressure turbine rotational speed sensor is not normal. 
     In order to achieve the object, this invention provides in its second aspect a method for controlling a gas-turbine aeroengine mounted on an aircraft and having at least a high-pressure turbine rotated by injection of high-pressure gas produced upon ignition and combustion of an air-fuel mixture in a combustion chamber, a low-pressure turbine located downstream of the high-pressure turbine to be rotated by low-pressure gas exiting the high-pressure turbine, a low-pressure turbine rotational speed sensor adapted to detect a rotational speed of the low-pressure turbine; and a high-pressure turbine rotational speed sensor adapted to detect a rotational speed of the high-pressure turbine; comprising the steps of: discriminating whether or not the low-pressure turbine rotational speed sensor is normal; and establishing a first value as an upper limit value of the rotational speed of the high-pressure turbine and controlling the rotational speed of the high-pressure turbine based on the established upper limit value; wherein the step of controlling changes the upper limit value to a second value that is lower than the first value, when the step of low-pressure turbine rotational speed sensor normality discriminating discriminates that the low-pressure turbine rotational speed sensor is not normal. 
    
    
     
       BRIEF DESCRIPTION OF DRAWINGS 
       The above and other objects and advantages of the invention will be more apparent from the following description and drawings in which: 
         FIG. 1  is an overall schematic view of a control apparatus for a gas-turbine aeroengine; 
         FIG. 2  is a flowchart for explaining operation of the apparatus; 
         FIG. 3  is a graph showing an aspect of characteristics (shown in  FIG. 2 ) expressing high-pressure turbine rotational speed relative to the flight altitude of the aircraft; and 
         FIG. 4  is a graph similarly showing an aspect of the characteristics expressing high-pressure turbine rotational speed relative to the flight speed of the aircraft. 
     
    
    
     DESCRIPTION OF EMBODIMENT 
     An embodiment of a control apparatus for a gas-turbine aeroengine according to the present invention will now be explained with reference to the attached drawings. 
       FIG. 1  is an overall schematic view of the control apparatus for a gas-turbine aeroengine. 
     Four types of gas-turbine aeroengines are known: the turbojet engine, turbofan engine, turboprop engine and turboshaft engine. A two-shaft turbofan engine will be taken as an example in the following explanation. 
     In  FIG. 1 , reference numeral  10  designates the turbofan engine (gas turbine engine; hereinafter referred to as “engine”). Reference numeral  10   a  designates a main engine unit. Two of the engines  10  are installed, one on either side of an aircraft (whose airframe is not shown). 
     The engine  10  is equipped with a fan (fan blades)  12  that sucks in external air while rotating rapidly. A rotor  12   a  is formed integrally with the fan  12 . The rotor  12   a  and a stator  14  facing it together form a low-pressure compressor  16  that compresses the sucked-in air and pumps it rearward. 
     A duct (bypass)  22  is formed in the vicinity of the fan  12  by a separator  20 . Most of the air pulled in passes through the duct  22  to be jetted rearward of the engine without being burned at a later stage (in the core). The wind from the fan  12  produces a force of reaction that acts on the airframe (not shown) on which the engine  10  is mounted as a propulsive force (thrust). Most of the propulsion is produced by the air flow from the fan. 
     The air compressed by the low-pressure compressor  16  flows rearward to a high-pressure compressor  24  where it is further compressed by a rotor  24   a  and stator  24   b  and then flows rearward to a combustion chamber  26 . 
     The combustion chamber  26  is equipped with a fuel nozzle  28  that is supplied with pressurized fuel metered by an FCU (fuel control unit)  30 . The FCU  30  is equipped with a fuel metering valve (FMV)  32 . Fuel pumped by a fuel pump  34  from a fuel tank  36  located at an appropriate part of the airframe is metered by the fuel metering valve  32  and supplied to the fuel nozzle  28  through a fuel supply line  38 . 
     The fuel metering valve  32  is connected to a torque motor  32   a  to be opened/closed thereby. The position of the fuel metering valve  32  is detected by a nearby valve position sensor  32   b . A fuel shutoff valve (SOV)  38   a  is interposed in the fuel supply line  38 . The fuel shutoff valve  38   a  is connected to an electromagnetic solenoid  38   b  to be opened/closed thereby. 
     The fuel nozzle  28  sprays the fuel supplied through the fuel supply line  38 . 
     The fuel sprayed from the fuel nozzle  28  and compressed air supplied from the high-pressure compressor  24  are mixed in the combustion chamber  26  and the air-fuel mixture is burned after being ignited at engine starting by an ignition unit (not shown) comprising an exciter and a sparkplug. Once the air-fuel mixture begins to burn, the air-fuel mixture composed of compressed air and fuel is continuously supplied and burned. 
     The hot high-pressure gas produced by the combustion is sent to a high-pressure turbine  40  to rotate it at high speed. The high-pressure turbine  40  is connected to the rotor  24   a  of the high-pressure compressor  24  through a high-pressure turbine shaft  40   a  to rotate the rotor  24   a.    
     After driving the high-pressure turbine  40 , the hot high-pressure gas is sent to a low-pressure turbine  42  to rotate it at relatively low speed. The low-pressure turbine  42  is connected to the rotor  12   a  of the low-pressure compressor  16  through a low-pressure turbine shaft  42   a  (in a dual concentric structure with the shaft  40   a ), so as to rotate the rotor  12   a  and fan  12 . The gas having passed through the high-pressure turbine  40  is lower in pressure than gas jetted from the combustion chamber  26 . 
     The exhaust gas exiting the low-pressure turbine  42  (turbine exhaust gas) is mixed with the fan exhaust air passing as is through the duct  22  and jetted together rearward of the engine  10  through a jet nozzle  44 . 
     An accessory drive gearbox (hereinafter referred to as “gearbox”)  46  is attached to the outer undersurface at the front end of the main engine unit  10   a  through a stay  46   a . An integrated starter/generator (hereinafter called “starter”)  50  is attached to the front of the gearbox  46 . The FCU  30  is located at the rear of the gearbox  46 . 
     At starting of the engine  10 , the starter  50  rotates a shaft  52  whose rotation is transmitted through a drive shaft  54  (and a gear mechanism including a bevel gear etc. (not shown)) to the high-pressure turbine shaft  40   a  to generate compressed air. The generated compressed air is supplied to the combustion chamber  26 , as mentioned above. 
     The rotation of the shaft  52  is also transmitted to a PMA (permanent magnet alternator)  56  and the (high-pressure) fuel pump  34 , whereby, as explained above, the fuel pump  34  is driven to supply metered fuel to the fuel nozzle  28  so as to be mixed with compressed air and atomized. The resulting air-fuel mixture is ignited to start combustion. 
     When the engine  10  reaches self-sustaining operating speed, the rotation of the high-pressure turbine shaft  40   a  is transmitted back to the shaft  52  through the drive shaft  54  (and the gear mechanism including the bevel gear etc. (not shown)) to drive the fuel pump  34  and also drive the PMA  56  and starter  50 . 
     As a result, the PMA  56  generates electricity and the starter  50  also generates electricity to be supplied to the airframe. Therefore, particularly when the electrical load on the airframe side increases, power generated by the starter  50  increases and rotational load on the high-pressure turbine shaft increases, thereby affecting the high-pressure turbine rotational speed, as will be explained later. 
     An ECU (Electronic Control Unit)  60  is installed at an upward location of the main engine unit  10   a . The ECU  60  is equipped with a microcomputer comprising a CPU, ROM, RAM, I/O etc. (none of which are shown) and is housed in a container for mounting at the upward position. 
     An N1 sensor (rotational speed sensor)  62  is installed near the low-pressure turbine shaft  42   a  of the engine  10  and outputs a signal indicating the rotational speed of the low-pressure turbine (rotational speed of the low-pressure turbine shaft  42   a ) N1 (so as to detect the speed N1), and an N2 sensor (rotational speed sensor)  64  is installed near the shaft  52  and outputs a signal indicating the rotational speed of the high-pressure turbine (rotational speed of the high-pressure turbine shaft  40   a ) (so as to detect the speed N2). 
     A T1 sensor (temperature sensor)  70  installed near an air intake  66  at the front of the main engine unit  10   a  outputs a signal indicating the engine inlet temperature (ambient or intake temperature) T1 (so as to detect the temperature the temperature T1). An EGT sensor (exhaust gas temperature sensor)  72  installed at a suitable location downstream of the low-pressure turbine  42  outputs a signal indicating the exhaust gas temperature (low-pressure turbine outlet temperature) EGT (so as to detect the temperature EGT). 
     A P0 sensor (pressure sensor)  74  installed inside the container that houses the ECU  60  outputs a signal indicating atmospheric pressure P0 (so as to detect the pressure P0), and a P1 sensor (pressure sensor)  76  installed near the air intake  66  outputs a signal indicating engine inlet pressure (air intake pressure) P1 (so as to detect the pressure P1). In addition, a P3 sensor  78  installed downstream of the high-pressure compressor  24  outputs a signal indicating compressor outlet pressure (outlet pressure of the high-pressure compressor  24 ) P3 (so as to detect the pressure P3). 
     The outputs of the foregoing sensors indicating the operating condition of the engine  10  are sent to the ECU  60 . 
     On the airframe side are installed a flight altitude sensor  80  that produces an output indicating the flight altitude ALT of the aircraft (so as to detect the flight altitude ALT) and a flight speed sensor  82  that produces an output indicating the flight speed Mn (Mach Number) of the aircraft (so as to detect the speed Mn). The outputs of these sensors are also sent to the ECU  60  comprising a computer on the airframe side. 
     The high-pressure compressor  24  is equipped with a BOV (Bleed Off Valve)  84  at a location of its front stage. During starting, low-speed operation and the like of the engine  10 , some of the compressed air flowing through a compression passage of the high-pressure compressor  24  is bled off through the first BOV  84  and discharged into the duct  22 . 
     The BOV  84  is opened and closed by an electromagnetic solenoid valve operated by commands from the ECU  60 . A BOV position sensor  86  installed near the BOV  84  to produce and send to the ECU  60  a signal indicating the amount of air bled through the BOV  84  based on the position (opening angle) of the BOV  84  (so as to detect the bleed air amount). 
     In addition, the high-pressure compressor  24  is equipped with another BOV (Bleed Off Valve)  90  at a location downstream of the BOV  84 , and some of the compressed air flowing through a compression passage of the high-pressure compressor  24  is bled off through the BOV  90  and sent to the cabin etc. on the airframe side for airframe cabin pressurization, air conditioning, wing de-icing, air sealing and other purposes. The BOV  90  is opened and closed by an electromagnetic solenoid valve in response to manual operation of a switch by a pilot seated in the cockpit of the airframe. 
     Further, the ECU  60  is responsive to the position of a thrust lever operated by the pilot for controlling the operation of the torque motor  32   a  to open/close the fuel metering valve  32  and for energizing/de-energizing the electromagnetic solenoid  38   b  to open/close the fuel shutoff valve  38   a  and control supply of fuel to the fuel nozzle  28 . 
       FIG. 2  is a block diagram for functionally explaining such operation (processing) of the apparatus, more specifically the ECU  60 . The illustrated processing is executed at predetermined time intervals. 
     Explaining this, the ECU  60  has an N1 sensor normality discriminating block (discriminator)  60   a  and a control block (controller)  60   b.    
     The N1 sensor normality discriminating block  60   a  discriminates or determines whether or not the N1 sensor  62  is normal, more specifically the output of the N 1  sensor  62  is normal by performing a range check, performing disconnection/short-circuiting checks and comparing outputs between channels (in other ECUs not shown), and also by comparison with estimated values obtained by the technique of Patent Document 1 and/or appropriately established reference values. 
     The control block  60   b  is connected with a first value setting block  60   b   1  that sets a first value as an upper limit value of the rotational speed N2 of the high-pressure turbine  40 . Upon receiving the first value set by the value defining block  60   b   1  as input, the control block  60   b  establishes or defines the first value as the upper limit value of the rotational speed N2 of the high-pressure turbine  40  and controls the rotational speed N2 of the high-pressure turbine  40  based on the established first value such that the rotational speed N1 of the low-pressure turbine  42  is within a permissible rotational speed. 
     The control block  60   b  is further equipped with an upper limit value changing block (changer)  60   b   2 . The upper limit value changing block  60   b   2  receives the outputs of the flight altitude sensor  80  and flight speed sensor  82  as inputs, retrieves characteristics ( 3 D mapped data)  60   b   21  by the inputted values ALT and Mn to establish or define a second value, and outputs the second value to a selection circuit  60   b   22 . 
       FIG. 3  is a graph showing an aspect of the characteristics  60   b   21  expressing high-pressure turbine rotational speed N2 (corresponding to the second value) relative to the flight altitude ALT, and  FIG. 4  is a graph showing an aspect of the characteristics  60   b   21  expressing high-pressure turbine rotational speed N2 (also corresponding to the second value) relative to the flight speed Mn. Characteristic curves for a number of flight speeds Mn are shown in  FIG. 3  and for a number of flight altitudes ALT in  FIG. 4 . 
     As shown in  FIG. 3 , the high-pressure turbine rotational speed N2, i.e., the second value is established or defined to decrease with increasing ALT (detected from the output of the flight altitude sensor  80 ) of the aircraft (in which the engine  10  is mounted). 
     Further, as shown in  FIG. 4 , the high-pressure turbine rotational speed N2, i.e., the second value is established or defined to decrease with decreasing flight speed Mn of the aircraft detected from the output of the flight speed sensor  82 . 
     Returning to the explanation of  FIG. 2 , as illustrated, the upper limit value changing block  60   b   2  includes a flight altitude sensor normality discriminating block (discriminator)  60   b   23  that discriminates whether or not flight altitude sensor  80 , more precisely its output is normal, and a flight speed sensor normality discriminating block (discriminator)  60   b   24  that discriminates whether or not the flight speed sensor  82 , more precisely its output is normal. 
     When at least one of the flight altitude sensor normality discriminating block  60   b   23  or the flight speed sensor normality discriminating block  60   b   24  discriminates that the associated sensor is not normal (is abnormal), the output of the block  60   b   23  or  60   b   24  is sent through an OR circuit  60   b   25  to the selection circuit  60   b   22 . 
     Further, the upper limit value changing block  60   b   2  is connected to a third value setting block  60   b   26  that sets a third value as the upper limit value of the rotational speed N2 of the high-pressure turbine  40 , and the output of the third value setting block  60   b   26  is also sent to the selection circuit  60   b   22 . The selection circuit  60   b   22  operates in response to commands from the control block  60   b.    
     In the configuration shown in  FIG. 2 , when the N1 sensor normality discriminating block  60   a  discriminates that the N1 sensor  62  is not normal (is abnormal), the control block  60   b  operates the selection circuit  60   b   22  to input the second value, thereby changing the upper limit value to the second value that is lower than the first value. 
     Further, when the sensor associated with either the flight altitude sensor normality discriminating block  60   b   23  or the flight speed sensor normality discriminating block  60   b   24  is discriminated not to be normal (to be abnormal), the control block  60   b  operates the selection circuit  60   b   22  to input the third value set by the third value setting block  60   b   26 , thereby establishing or defining the third value that is still lower than the second value as the upper limit value of the rotational speed N2 of the high-pressure turbine  40 . 
     In the configuration shown in  FIG. 2 , the control block  60   b  thus calculates and outputs a control value of the turbine rotational speed N2 of the high-pressure turbine  40  based one of the established or defined first to third values. 
     As stated above, the embodiment is configured to have an apparatus (and method) for discriminating ignition in a gas-turbine aeroengine ( 10 ) mounted on an aircraft and having at least a high-pressure turbine ( 40 ) rotated by injection of high-pressure gas produced upon ignition and combustion of an air-fuel mixture in a combustion chamber ( 26 ), and a low-pressure turbine ( 42 ) located downstream of the high-pressure turbine to be rotated by low-pressure gas exiting the high-pressure turbine, comprising: a low-pressure turbine rotational speed sensor (N1 sensor  62 ) adapted to detect a rotational speed N1 of the low-pressure turbine  42 ; a high-pressure turbine rotational speed sensor (N2 sensor  64 ) adapted to detect a rotational speed N2 of the high-pressure turbine ( 40 ); a low-pressure turbine rotational speed sensor normality discriminator (ECU  60 ;  60   a ) that discriminates whether or not the low-pressure turbine rotational speed sensor ( 62 ) is normal; and a controller (ECU  60 ,  60   b ) that establishes a first value as an upper limit value of the rotational speed of the high-pressure turbine and controls the rotational speed N2 of the high-pressure turbine ( 40 ) based on the established upper limit value; wherein the controller has an upper limit value changer (upper limit value changing block  60   b   2 ) that changes the upper limit value to a second value that is lower than the first value, when the low-pressure turbine rotational speed sensor normality discriminator discriminates that the low-pressure turbine rotational speed sensor ( 62 ), more specifically its output is not normal. 
     Thus, the control apparatus for a gas-turbine aeroengine has the normality discriminator that discriminates normality of the low-pressure turbine rotational speed sensor  62  for detecting a low-pressure turbine rotational speed N1 and the controller establishes the first value as the upper limit value of the high-pressure turbine rotational speed N2 and controls the high-pressure turbine rotational speed N2 based on the established upper limit value, and is configured to change the upper limit value to the second value lower than the first value when the low-pressure turbine rotational speed sensor  62  is discriminated not to be normal, whereby, by suitably defining the first value, engine output (thrust) determined by the low-pressure turbine rotational speed N1 can be controlled to a desired value while restraining the high-pressure turbine rotational speed N2 to not greater than the first value, and whereby, by changing the second value to lower than the first value, low-pressure turbine overspeed at the time of a mishap such as blades of the fan  12  breakage can be reliably prevented. 
     The apparatus further including: a flight altitude sensor ( 80 ) adapted to detects flight altitude of the aircraft; and a flight speed sensor ( 82 ) adapted to detect a flight speed of the aircraft; and wherein the upper limit value changer establishes the second value based on outputs of the flight altitude sensor ( 80 ) and the flight speed sensor ( 82 ). 
     Thus, the apparatus includes flight altitude sensor  80  for detecting aircraft flight altitude ALT and flight speed sensor  82  for detecting flight speed Mn of the aircraft, and is configured to establish the second value based on the outputs of the flight altitude sensor  80  and flight speed sensor  82  for detecting aircraft flight altitude ALT and flight speed Mn, whereby, in addition to the aforesaid effects, by defining the second value based on the flight altitude ALT and the flight speed Mn, which are major operating parameters affecting the low-pressure turbine rotational speed N1 (and high-pressure turbine rotational speed N2) that determines the engine output, it is possible to suitably define the second value and minimize engine output decline under prevailing conditions. 
     In the apparatus, the upper limit value changer ( 60   b   2 ) establishes the second value such that the second value decreases with increasing flight altitude of the aircraft. 
     With this, the apparatus is configured to define the second value to decrease with increasing aircraft flight altitude ALT detected from the output of the flight altitude sensor  80 , whereby, in addition to the aforesaid effects, engine output decline can be minimized. 
     In the apparatus, the upper limit value changer ( 60   b   2 ) establishes the second value such that the second value decreases with decreasing flight speed of the aircraft. 
     With this, the apparatus is configured to define the second value to decrease with decreasing aircraft flight speed Mn detected from the output of the flight speed sensor  82 , whereby, in addition to the aforesaid effects, engine output decrease can be prevented to a minimum extent. 
     In the apparatus, the upper limit value changer ( 60   b   2 ) includes: a flight altitude sensor normality discriminator (discriminating block  60   b   23 ) that discriminates whether or not the flight altitude sensor ( 80 ) is normal; and a flight speed sensor normality discriminator (discriminating block  60   b   24 ) that discriminates whether or not the flight speed sensor ( 82 ) is normal; and wherein the upper limit value changer ( 60   b   2 ) changes the upper limit value to a third value that is lower than the second value, when at least one of the flight altitude sensor normality discriminator and the flight speed sensor normality discriminator discriminates that at least one of the flight altitude sensor ( 80 ) and the flight speed sensor ( 82 ) is not normal. 
     Thus, the apparatus is configured to includes the normality discriminator that discriminates normality of the flight altitude sensor  80  and flight speed sensor  82 , and is configured to change the upper limit value to the third value lower than the second value when that either or both of the flight altitude sensor and the flight speed sensor are not normal, whereby, in addition to the aforesaid effects, by delaying substantial change of the upper limit value until this stage, a sharp decline in engine output (thrust) can be delayed to this stage, and by suitably defining the third value, the low-pressure turbine overspeed can be reliably prevented even when the flight altitude sensor  80  or the flight speed sensor is not in a normal state. 
     While the invention has thus been shown and described with reference to a specific embodiment, it should be noted that the invention is in no way limited to the details of the described arrangements; changes and modifications may be made without departing from the scope of the appended claims.