JP4523693B2 - Control device for aircraft gas turbine engine - Google Patents

Description

[0001]
BACKGROUND OF THE INVENTION
The present invention relates to an aircraft gas turbine engine control apparatus.
[0002]
[Prior art]
In an aircraft engine control device, generally, a command value for the flow rate of fuel supplied to a gas turbine engine is calculated so that the turbine speed matches a target value that achieves the thrust required by the operator. A stationary mode control system (or thrust control system) and an unsteady mode control system such as an acceleration / deceleration mode control system that calculates the fuel flow rate command value based on the acceleration / deceleration rate of the turbine speed are arranged in parallel. The minimum value or the maximum value of the command value thus selected is selected, and the fuel flow rate is determined based on the selected command value. As an example, there can be mentioned “Technology report 6307 (1992)“ Study prototype of high performance engine control system (Part 2) ”.
[0003]
[Problems to be solved by the invention]
In the prior art described above, the minimum or maximum value of the calculated command value is selected, in other words, when switching the control mode, it is necessary to reset the integrator in the switched control system, and a complicated algorithm is required. There is a case where control hunting is caused by resetting.
[0004]
Referring to FIG. 12, FIG. 12 (a) is a block diagram showing the operation of the steady mode control system in the prior art. In the operation of the steady mode control system in the prior art, the flow rate of fuel supplied to the gas turbine engine so that the deviation between the detected turbine rotational speed N1 and the target rotational speed N1com that realizes the driver's required thrust is reduced. (More specifically, the command value) Wf is calculated.
[0005]
In the control system according to this prior art, generally, as shown in FIG. 12 (a), a proportional element is used so that the deviation between the detected turbine speed N1 and the target speed N1com that realizes the driver's required thrust is reduced. Using KP and the integral element 1 / S (S: Laplace operator) (or using a differential element), the manipulated variable Wf (more specifically, manipulated variable Wfn1 based on the turbine speed N1) is calculated.
[0006]
In the illustrated example, the operation amount is calculated based on a preset engine model, that is, when a plant transfer function is given, the operation amount is set so that the closed-loop transfer function partially matches the model. The operation amount is calculated by using a (partial) model matching method to be calculated.
[0007]
In the prior art shown in FIG. 12A, when switching to an unsteady mode control system such as an acceleration / deceleration mode control system, the integrator (1 / S) needs to be reset and a complicated algorithm is required. In addition, there is a disadvantage that control hunting is caused by resetting.
[0008]
Accordingly, an object of the present invention is to eliminate the above-described problems. In a control device for a gas turbine engine in which a plurality of control systems are arranged in parallel, when the control mode is switched, an integrator reset is not required. It is an object of the present invention to provide an aircraft gas turbine engine control apparatus that simplifies the above.
[0009]
[Means for Solving the Problems]
To achieve the above object, according to claim 1, in an aircraft gas turbine engine having at least one turbine, an operating state of the gas turbine engine including at least the turbine speed is detected. Operating state detection means, pilot demand thrust detection means for detecting pilot demand thrust required by the pilot for the gas turbine engine, at least the detected turbine speed and the detected pilot demand thrust A steady mode control system for calculating a command value for the flow rate of fuel supplied to the gas turbine engine using at least an integral element so that the deviation of the target turbine speed for realizing In order to reduce the deviation between the acceleration / deceleration rate and the target acceleration / deceleration rate, at least the integration element is used to increase the gas turbine energy. A non-stationary mode control system including at least an acceleration / deceleration mode control system for calculating a command value of a fuel flow rate supplied to the gin, a command value calculated by the steady mode control system, and a command value calculated by the acceleration / deceleration mode control system Command value selection means for selecting any one of the above, fuel flow rate calculation means for calculating the flow rate of fuel supplied to the gas turbine engine based on the selected command value, and the gas based on the calculated fuel flow rate In a control apparatus for a gas turbine engine having fuel supply means for supplying fuel to a turbine engine, an integral element used in the steady mode control system and the unsteady mode control system is removed, and the fuel flow rate calculation means An integral element is arranged, so that the fuel flow rate calculation means is based on the command value selected using the integral element. Te is configured to calculate the fuel flow rate In the acceleration / deceleration mode control system, a schedule value corresponding to the detected turbine speed is added. It was.
[0010]
Since the integral element used in the steady mode control system is removed, the integral element is arranged in the fuel flow rate calculation means, and the fuel flow rate is calculated using the integral element. In other words, the control mode is Since the integration is performed after the switching, the integrator can be simplified by eliminating the need for the integrator reset, and the above-described inconvenience can be solved.
[0012]
further Since the acceleration / deceleration mode control system is configured to add a schedule value according to the detected turbine speed , System The accuracy can be improved. Moreover, it is not affected by the heat mass of the gas turbine engine.
[0013]
DETAILED DESCRIPTION OF THE INVENTION
An aircraft gas turbine engine control apparatus according to an embodiment of the present invention will be described below with reference to the accompanying drawings.
[0014]
FIG. 1 is a schematic view showing the entire apparatus.
[0015]
There are four known types of aircraft gas turbine engines: turbojet engines, turbofan engines, turboprop engines, and turboshaft engines. The following are examples of 2-axis turbofan engines. I will explain to you.
[0016]
In FIG. 1, reference numeral 10 indicates a turbofan engine (hereinafter referred to as “engine”), and reference numeral 10a indicates an engine body. The engine 10 is mounted at an appropriate position on the airframe (not shown).
[0017]
The engine 10 includes a fan (fan blade) 12 and the fan 12 sucks air from outside air while rotating at high speed. The fan 12 is integrally formed with a rotor 12a, and the rotor 12a constitutes a low-pressure compressor 16 together with a stator 14 disposed so as to face the rotor 12, and the sucked air is compressed and sent backward.
[0018]
A duct (bypass) 22 is formed in the vicinity of the fan 12 by the separator 20, and most of the sucked air is not burned in the rear stage (core side) and is ejected through the duct 22 to the rear of the engine. It is done. The fan exhaust produces thrust (thrust) in the airframe (not shown) on which the engine 10 is mounted as a reaction. Most of the thrust is generated by this fan exhaust.
[0019]
The air compressed by the low-pressure compressor 16 is sent to the subsequent high-pressure compressor 24, where it is further compressed by the rotor 24 a and the stator 24 b and then sent to the subsequent combustor 26.
[0020]
The combustor 26 includes a fuel nozzle 28, and fuel metered by an FCU (Fuel Control Unit) 30 is pumped to the fuel nozzle 28. That is, the FCU 30 is provided with a fuel metering valve 32, and fuel pumped up from a fuel tank 36 disposed at an appropriate position of the machine body by a fuel pump (gear pump) 34 is metered by the fuel metering valve 32, and then fuel The fuel is supplied to the fuel nozzle 28 through the supply passage 38.
[0021]
The sprayed fuel is mixed with the compressed air pumped from the high-pressure compressor 24, and is ignited and burned by an exciter (not shown in FIG. 1) and a spark plug (not shown) when the engine is started. Once the air-fuel mixture is ignited and starts to burn, the air-fuel mixture comprising compressed air and fuel is continuously supplied to continue the combustion.
[0022]
The high-temperature high-pressure gas generated by the combustion is sent to the high-pressure turbine 40 and rotates the high-pressure turbine 40 at a high speed. The high-pressure turbine 40 is connected to the rotor 24a of the high-pressure compressor described above via a high-pressure turbine shaft 40a, and rotates the rotor 24a.
[0023]
The high-temperature high-pressure gas rotates the high-pressure turbine 40 and is then sent to the low-pressure turbine 42 to rotate the low-pressure turbine 42 at a relatively low speed. The low-pressure turbine 42 is connected to the rotor 12a of the low-pressure compressor 16 through a low-pressure turbine shaft 42a (a biaxial structure concentric with the shaft 40a), and rotates the rotor 12a and the fan 12.
[0024]
The high-temperature high-pressure gas (turbine exhaust) that has passed through the low-pressure turbine 42 is mixed with the fan exhaust that is discharged as it is through the duct 22 and is ejected from the jet nozzle 44 to the rear of the engine.
[0025]
An accessory drive gearbox (hereinafter referred to as “gearbox”) 50 is attached via a stay 50a near the front side of the outer lower surface of the engine body 10a, and is integrally formed at the front end of the gearbox 50. A starter and a generator (hereinafter collectively referred to as “starter”) 52 are attached. The FCU 30 is disposed at the rear end of the gear box 50.
[0026]
When the engine 10 is started, when the shaft 56 is rotated by the starter 52, the rotation is transmitted to the high-pressure turbine shaft 40a via the drive shaft 58 (and a gear mechanism such as a bevel gear (not shown)), and air necessary for combustion is transferred. It is sent.
[0027]
On the other hand, the rotation of the shaft 56 is transmitted to a PMA (permanent magnet alternator) 60 and the fuel pump 34 to drive the fuel pump 34 and spray the fuel through the fuel nozzle 28 as described above. The resulting air-fuel mixture is ignited and starts to burn.
[0028]
When the engine 10 reaches the self-sustained operation speed, the rotation of the high-pressure turbine shaft 40a is transmitted to the shaft 56 via the drive shaft 58 (and a gear mechanism such as a bevel gear (not shown)) to drive the fuel pump 34, The PMA 60 and the starter 52 are driven. Thereby, the PMA 60 generates power, and the starter 52 supplies power to the airframe.
[0029]
In the engine 10, an N1 sensor (rotational speed sensor) 62 is disposed in the vicinity of the low pressure turbine shaft 42a, and outputs a signal proportional to the low pressure turbine rotational speed (the rotational speed of the low pressure turbine shaft 42a) N1. An N2 sensor (rotational speed sensor) 64 is disposed in the vicinity, and outputs a signal proportional to the high-pressure turbine rotational speed (the rotational speed of the high-pressure turbine shaft 40a) N2.
[0030]
A T1 (temperature sensor) sensor 68 and a P1 sensor (pressure sensor) 70 are disposed in the vicinity of the air intake 66 on the front surface of the engine body 10a, and output signals proportional to the temperature T1 and pressure P1 of the inflowing air. ECU described later
A P0 sensor (pressure sensor) 72 is provided inside (Electronic Control Unit) and outputs a signal proportional to the atmospheric pressure P0.
[0031]
A P3 sensor (pressure sensor) 74 is disposed downstream of the rotor 24a to output a signal proportional to the output pressure P3 of the high-pressure compressor 24, and at an appropriate position between the high-pressure turbine 40 and the low-pressure turbine 42. A sensor (temperature sensor) 76 is arranged, and outputs a signal proportional to the temperature (engine representative temperature) ITT of the part.
[0032]
The above-described ECU (indicated by reference numeral 80) is accommodated at the upper end position of the engine body 10a. The output of the sensor group described above is sent to the ECU 80.
[0033]
FIG. 2 is a block diagram generally showing the configuration of the ECU 80 and the FCU 30 described above, particularly the configuration of the FCU 30.
[0034]
In addition to the sensor group described above, a TLA sensor (thrust lever position sensor) 84 is disposed in the vicinity of a thrust lever (throttle lever) 82 installed in the vicinity of the aircraft control seat (cockpit, not shown). A signal proportional to the thrust lever position (operator request thrust) TLA input is output. The output of the TLA sensor 84 is also input to the ECU 80. 2 and 3, each sensor (P0 sensor, TLA sensor, etc.) is indicated by its detection target parameter name (P0, TLA, etc.).
[0035]
Further, an FMVP sensor (valve position sensor; not shown in FIG. 2) is provided at an appropriate position of the FCU 30 and outputs a signal proportional to the valve position FMVP of the fuel metering valve 32. The output of the FMVP sensor is also input to the ECU 80.
[0036]
Further, the ECU 80 receives, via the communication interface unit 88, a pilot selection command 90 for devices other than the thrust lever 82 described above, data from the airframe computer (Air Data Computer or ADC) 92 (for example, Mach number Mn, pressure altitude). ALT, outside air temperature (more specifically, total temperature TAT, true air temperature SAT), and data from ECU 94 of the second engine (not shown) are input (or output), and a display in the cockpit It is connected to 96 and the data of ECU80 is displayed.
[0037]
Based on the input value, the ECU 80 supplies the engine 10 so that the deviation between the low-pressure turbine shaft rotational speed (low-pressure turbine rotational speed) N1 and the target rotational speed N1com decreases according to the thrust lever position (operator request output) TLA. The command value (operation amount) Wf for the fuel flow should be calculated as the energization current command value for the torque motor 98 and sent to the FCU 30.
[0038]
Further, the ECU 80 monitors and detects whether any of the detected values of the low-pressure turbine rotational speed N1 and the high-pressure turbine rotational speed N2 exceeds a limit value (for example, a value corresponding to 107% of each maximum rotational speed). When either the low-pressure turbine speed N1 or the high-pressure turbine speed N2 exceeds the limit value, it is determined that the speed is overspeed, and the fuel flow to be supplied to the engine 10 is a predetermined value, more specifically, zero or minimum. Thus, an energization current command value for the torque motor 98 is determined and sent to the FCU 30.
[0039]
Further, the ECU 80 instructs the fuel flow rate to be supplied to the engine 10 so that the deviation between the detected change rate N2 dots (the differential value of N2; acceleration / deceleration rate) of the high-pressure turbine rotation speed N2 and the target acceleration / deceleration rate N2 dots com decreases. A value Wf, more specifically, an energization current command value for the torque motor 98 is determined and sent to the FCU 30.
[0040]
The FCU 30 includes a low-pressure fuel pump 100, and the fuel pumped up from the fuel tank 36 (not shown in FIG. 2) is increased in pressure by the fuel pump 34 through the filter (and oil cooler) 102, and the fuel metering valve 32. Sent to. The torque motor 98 is connected to the fuel metering valve 32 and determines its spool position. Therefore, the fuel pressure-fed through the high-pressure pump 34 is adjusted (metered) to a flow rate corresponding to the spool position by the fuel metering valve 32.
[0041]
The metered fuel is supplied to the fuel nozzle 28 through the shut-off valve 104, the drain valve 106, and the shut-off mechanism 108.
[0042]
An emergency stop switch 110 is connected to the low-pressure turbine shaft 42a. The low-pressure turbine shaft 42a is turned on when the low-pressure turbine shaft 42a is displaced for some reason, and the fuel supply to the fuel nozzle 28 is mechanically operated by operating the shut-off mechanism 108. Cut off. Similarly, a solenoid 112 is provided, and the shutoff valve 104 is operated according to the pilot selection command 90 to shut off the fuel supply to the fuel nozzle 28.
[0043]
Next, the operation of the control device for the gas turbine engine according to the embodiment, more specifically, the operation of the ECU 80 will be specifically described.
[0044]
FIG. 3 is a block diagram specifically showing the operation of the ECU 80.
[0045]
As shown in the figure, the ECU 80 includes a maximum N1 calculation unit (shown as “Max N1 on Condition” in the figure) 800. The maximum N1 calculation unit 800 includes a Mach number Mn, a pressure altitude ALT, a true atmospheric temperature SAT, an Anti-Ice, and the like. The maximum value N1 * maxcom of the low-pressure turbine speed N1 is calculated based on the bleed state A / I from the engine.
[0046]
Further, the ECU 80 includes an idle speed calculation unit (shown as “Idle (Ground / Flight)”) 802. The idle speed calculation unit 802 includes the Mach number Mn, the pressure altitude ALT, the true atmospheric temperature SAT, A / Based on I, WOW (Weight-on-Wheel, wheel load signal), the minimum value N1 * mincom of the low-pressure turbine rotational speed N1 and the high-pressure turbine rotational speed N2 * idle during idling are calculated.
[0047]
The calculated maximum value N1 * maxcom of the low-pressure turbine rotation speed N1, the minimum value N1 * mincom of the low-pressure turbine rotation speed N1, and the high-pressure turbine rotation speed N2 * idle during idling are calculated by a thrust calculation unit (“TLA Power Setting” in the figure). It is sent to 804. The thrust calculation unit 804 calculates the low-pressure turbine target rotational speed N1com and the high-pressure turbine target rotational speed N2com based on the input values, the detected thrust lever position TLA and the inflow air temperature T1 that are separately input.
[0048]
The calculated low-pressure turbine target rotational speed N1com and high-pressure turbine target rotational speed N2com are indicated by a rotational speed feedback control unit (shown as “Speed Feedback Control (N1 / N2)” in the figure, corresponding to the above-described steady mode control system) 806 Sent.
[0049]
The rotational speed feedback control unit 806, based on these input values, separately detected low-pressure turbine rotational speed N1 and high-pressure turbine rotational speed N2, the target rotational speed that realizes the detected turbine rotational speed N1 and the driver's required thrust. A manipulated variable (fuel flow rate command value) Wf dot n1 supplied to the engine 10 is calculated so that the deviation of N1com is reduced, and the target rotational speed N2com that realizes the detected turbine rotational speed N2 and the driver's required thrust. The manipulated variable Wf dot n2 is calculated so that the deviation of. In this way, the operation amount Wf is calculated as a differential value (denoted as a dot) as will be described later.
[0050]
In addition, the ECU 80 includes an acceleration control unit (shown as “Acceleration Control” in the figure, which corresponds to the above-described acceleration / deceleration mode control system) 808, and the acceleration control unit 808 includes the high-pressure turbine rotation speed N2, the inflow air temperature T1, the inflow air. The command value of the fuel flow rate supplied to the engine 10 so that the deviation between the acceleration rate N2 dots of the high-pressure turbine rotational speed N2 detected based on the pressures P1, A / I and the target acceleration rate N2 dots com decreases. (Operation amount) Wf dot acc is calculated.
[0051]
In addition, the ECU 80 includes a P3 pressure control unit (shown as “P3 Limit Control” in the figure, which corresponds to the above-described unsteady mode control system) 810, and the P3 pressure control unit 810 is connected to the outlet pressure P3 and PR of the high-pressure compressor 24. Based on (Power-Reserve signal), a command value (operation amount) Wf dot p3 of the fuel flow rate supplied to the engine 10 corresponding to the limit value of the outlet pressure P3 is calculated.
[0052]
Further, the ECU 80 includes an ITT temperature control unit (shown as “ITT Limit Control” in the figure. Equivalent to the above-described non-steady mode control system) 812. Based on the above, a command value (operation amount) Wf dot itt for the fuel flow rate supplied to the engine 10 corresponding to the limit value of the turbine temperature ITT is calculated.
[0053]
In addition, the ECU 80 includes a deceleration control unit (shown as “Deceleration Control (Dec-limit)” in the figure. Equivalent to the above-described acceleration / deceleration mode control system) 814. The deceleration control unit 814 includes the high-pressure turbine rotation speed N2, the inflow air. Fuel supplied to the engine 10 so that the deviation between the deceleration rate N2 dot of the turbine speed N2 detected based on the temperature T1, the inflow air pressure P1, and A / I and the target target deceleration rate N2 dot com decreases. The flow rate command value (operation amount) Wf dot dec is calculated.
[0054]
All the command values (operation amounts) Wf of the calculated fuel flow rate are sent to a selection circuit (shown as “Select Logic” in the figure. This corresponds to the command value selection means described above) 820, from which the selection circuit 820 According to the determined logic, one value is selected and set as an output Wfcom.
[0055]
The ECU 80 includes a torque motor driver 826. The torque motor driver 826 determines the valve command value FMMVCMD (more specifically, the torque motor energization current command value) from the input maximum value Wfcom and the detected valve position FMVP of the fuel metering valve 32. Is calculated and output.
[0056]
FIG. 4 is a partial block diagram of FIG. 3 showing an operation operation inside the selection circuit 820 described above in FIG.
[0057]
As shown in the figure, among the fuel flow rate command values sent to the selection circuit 820, all command values except the fuel flow rate command value Wf dot dec from the deceleration control unit 814 are selected by the selection circuit 821 ("Select / The minimum value is selected.
[0058]
Next, the minimum value and the command value Wf dot dec of the fuel flow rate from the deceleration control unit 814 are sent to the selection circuit 822 (shown as “Select / High” in the figure), and the maximum value is selected. Next, the previous value LastWf is added to the maximum value.
[0059]
Here, before continuing the description of the figure, the operation of the apparatus according to this embodiment will be described with reference to FIG. 12 again.
[0060]
In the prior art shown in FIG. 12 (a), the operation amount Wf is calculated using the proportional element KP and the integral element 1 / S so that the deviation between the target value N1com and the detected value N1 is reduced. As a result, the control mode is switched. In some cases, it is necessary to reset the integrator (1 / S), which requires a complicated algorithm, and resetting the control causes hunting.
[0061]
Therefore, in this embodiment, as shown in FIG. 12B, the integral element 1 / S is moved downstream of a selection point (shown as “Select” in the drawing) by the selection circuit 824 indicated by a broken line. The fuel flow rate) is calculated. Since the integral element is moved downstream of the selection point, the differential value Wf dot is used for the operation amount (fuel flow rate command value) upstream thereof.
[0062]
That is, the integration is performed after switching the control mode, and the integration element in the control mode is removed. This eliminates the need for resetting the integrator, simplifies the algorithm, and prevents the above-described control hunting from occurring.
[0063]
In the configuration shown in FIG. 12B as well, the operation amount is calculated based on a preset engine model. More specifically, when a plant transfer function is given, the closed loop transfer function is included in the model. The operation amount is calculated by using a (partial) model matching method for calculating the operation amount so as to be consistent with each other.
[0064]
Returning to the description of FIG. 4, the addition of the previous value Lastwf corresponds to this integration operation. Next, the added value is multiplied by the outlet pressure P3 of the high-pressure compressor 24. This is because the value obtained by dividing the fuel flow rate command value by the outlet pressure P3 in FIG. 3 is actually used as the fuel flow rate command value.
[0065]
Further, the corrected value is selected by comparing with a predetermined value in another selection circuit 823, 824, and the selected value is sent to the torque motor driver 826 as described above. Since it is in the calculation of the operation amount after the selection within 820, in other words, the control mode (control system) is switched, the detailed description is omitted.
[0066]
The same applies to the operation of the acceleration control unit 808 described with reference to FIG.
[0067]
Referring to FIG. 5, (a) shows the acceleration control in the prior art, in which the proportional element Kp, the differential elements KD and 2 so that the deviation between the target value N2 dots com and the detected value N2 dots is reduced. The manipulated variable Wfacc is calculated using the integral elements 1 / S. Therefore, when switching to this control mode, it is necessary to reset the two integrators (1 / S), which requires a complicated algorithm, and there is a disadvantage that control hunting is caused by resetting.
[0068]
Therefore, in this embodiment, as shown in FIG. 5B, the integral element 1 / S is moved downstream of the selection point indicated by the broken line to calculate the operation amount, as in FIG. 12B. .
[0069]
That is, the integration is performed after switching the control mode, and the integration element in the control mode is removed. This eliminates the need for resetting the integrator, simplifies the algorithm, and prevents the above-described control hunting from occurring.
[0070]
Since the integration element is moved downstream of the selection point, the differential value Wf dot acc of the operation amount (fuel flow rate command value) is used upstream thereof, which is not different from FIG.
[0071]
Further, the remaining one integral element is removed from the acceleration control according to the prior art shown in FIG. 5A, and a schedule of a correction value N2c (correction value based on the high-pressure turbine rotational speed N2) is added instead. I made it.
[0072]
FIG. 6 is a partial block diagram of FIG. 3 showing details of the acceleration control unit 808.
[0073]
That is, the product of the deviation between the target value and the detected value and the proportional gain Kp, the value obtained by multiplying the differential value of the deviation by the differential gain KD, and the Wf dot as the N2c schedule value are added. As shown in the figure, the Wf dot is in accordance with the high-pressure turbine rotational speed N2 (more specifically, the value obtained by correcting N2 with the square root value of the atmospheric temperature ratio θ (= incoming air temperature T1 / predetermined temperature TSTD (both absolute temperatures)) Calculated. Here, the reason why the schedule value is calculated from the vicinity exceeding N2 square root θ = 28000 is that the influence of the steady error due to the removal of the integral element from this vicinity becomes large.
[0074]
As described above, in the acceleration control, an integral element is originally required in the acceleration mode control system in order to improve the control accuracy. However, the integral element is removed, and in this embodiment, a schedule is replaced instead. Since the configuration is such that the values are added, the same effect as described with reference to FIG. 12B can be obtained, and the control accuracy can be improved. In addition, the engine 10 is not affected by the heat mass.
[0075]
FIG. 7 is a data diagram showing test results conducted by the inventors.
[0076]
In the figure, symbol a indicates the result when the proportional gain KP and the differential gain KD having the configurations shown in FIGS. 5B and 6 are used, that is, when the schedule value is not added.
[0077]
A symbol b indicates a result when the proportional gain Kp and the differential gain KD are increased as compared with the case of the symbol a. Symbol c indicates the result when the same proportional gain Kp and differential gain KD as in the case of symbol a are used, and the schedule values shown in FIG. 5 (and FIG. 6) are added.
[0078]
From FIG. 7, it can be understood that the addition of the schedule values shown in FIGS. 5B and 6 is effective.
[0079]
8 and 9 are partial block diagrams of FIG. 3 showing the operation of the above-described rotation speed feedback control unit 806 in detail. As shown in the figure, the integral element is removed and the fuel flow rate command values Wf dot n1 and Wf dot n2 corresponding to the low pressure turbine target rotational speed N1com and the high pressure turbine target rotational speed N2com are obtained by multiplying the deviation by the integral gain KI. Is calculated.
[0080]
FIGS. 10 and 11 are partial block diagrams of FIG. 3 showing the operations of the P3 pressure control unit 810 and the ITT temperature control unit 812 in detail. As shown in the figure, the integral element is removed, and the fuel flow rate command value corresponding to the limit value of the outlet pressure P3 and the fuel flow rate command corresponding to the limit value of the turbine temperature ITT are obtained by multiplying the deviation by the integral gain KI. The value Wf dot p3 and Wf dot itt are calculated.
[0081]
In this embodiment, as described above, since the integral element is removed in all the control systems, when the control mode is switched, it is not necessary to reset the integrator, and the algorithm can be simplified. At the same time, the above-described control hunting can be prevented from occurring.
[0082]
In the acceleration mode control system, the schedule value is added instead of the second integration element, so that the control accuracy can be improved.
[0083]
As described above, in this embodiment, an aircraft gas turbine engine (turbofan engine 10) having at least one turbine (more specifically, a low pressure turbine (42) and a high pressure turbine (40)). ), Operating state detecting means (N1 sensor 62, N2 sensor 64, ECU 80) for detecting the operating state of the gas turbine engine including at least the turbine speed (low pressure turbine speed N1, high pressure turbine speed N2), Pilot required thrust detection means (TLA sensor 84, ECU 80) for detecting a pilot required thrust (thrust lever position TLA) required by the pilot for the gas turbine engine, at least the detected turbine speed (N1) ) And the target turbine speed (N1) that realizes the detected driver's required thrust om), a steady mode control system (rotational speed feedback control unit 806) for calculating a command value of the flow rate of fuel supplied to the gas turbine engine using at least an integral element (1 / S) so as to reduce the deviation of om). The gas turbine engine uses at least an integral element (1 / S) so that the deviation between the detected acceleration / deceleration rate (N2 dot) and the target acceleration / deceleration rate (N2 dot com) of the turbine rotation speed (N2) decreases. An unsteady mode control system (acceleration control unit 808, P3 pressure control unit 810, ITT temperature control unit 812, deceleration control unit 814) including at least an acceleration / deceleration mode control system for calculating a command value of a fuel flow rate supplied to Select either the command value (Wf) calculated by the mode control system or the command value (Wf) calculated by the acceleration / deceleration mode control system Command value selection means (selection circuit 820), fuel flow rate calculation means (ECU 80) for calculating the flow rate of fuel to be supplied to the gas turbine engine (command value Wf) based on the selected command value, and the calculation In the control apparatus for a gas turbine engine provided with fuel supply means (FCU 30, fuel metering valve 32, torque motor 98, etc.) for supplying fuel to the gas turbine engine based on the fuel flow rate, the steady mode control system And an integral element used in the non-stationary mode control system is removed, and an integral element is arranged in the fuel flow rate calculation means, so that the fuel flow rate calculation means is based on the command value selected using the integral element. Configured to calculate the fuel flow rate. Further, the acceleration / deceleration mode control system is configured to add a schedule value corresponding to the detected turbine speed, as shown in FIG. It was.
[0085]
In the above-described embodiment, a turbofan engine is used as an example of an aircraft gas turbine engine. However, a turbojet engine, a turbofan engine, a turboprop engine, a turboshaft engine, and the like may be used. May be.
[0086]
【The invention's effect】
According to the first aspect, the integral element used in the steady mode control system or the like is removed, the integral element is arranged in the fuel flow rate calculation means, and the fuel flow rate is calculated using the integral element. Therefore, in other words, since the integration is performed after the control mode is switched, the integrator can be simplified by eliminating the need for resetting the integrator, and the inconveniences described above can be eliminated.
[0087]
further Since the acceleration / deceleration mode control system is configured to add a schedule value according to the detected turbine speed , System The accuracy can be improved. Moreover, it is not affected by the heat mass of the gas turbine engine.
[Brief description of the drawings]
FIG. 1 is a schematic view showing an overall control apparatus for an aircraft gas turbine engine according to an embodiment of the present invention.
2 is an explanatory block diagram showing a configuration of an ECU and an FCU in the apparatus of FIG. 1. FIG.
FIG. 3 is a block diagram specifically showing the operation of the ECU of FIG. 2;
4 is a partial block diagram of FIG. 3 showing an operation operation inside the selection circuit 820 in FIG. 3;
5 is a block diagram showing an operation of an acceleration control unit in FIG. 3. FIG.
6 is a partial block diagram of FIG. 3 showing details of the acceleration control unit in FIG. 3;
7 is a data diagram showing test results conducted by the inventors for the schedule value of the acceleration control unit in FIG. 3; FIG.
8 is a partial block diagram of FIG. 3 showing details of a thrust calculation unit in FIG. 3;
9 is a partial block diagram of FIG. 3 showing the details of the thrust calculation unit in FIG. 3 as well.
10 is a partial block diagram of FIG. 3 showing details of the P3 pressure control unit in FIG. 3;
11 is a partial block diagram of FIG. 3 showing details of the ITT temperature control unit in FIG. 3;
12 is a block diagram showing the operation of the rotation speed feedback control unit (steady control mode system) in FIG. 3 in comparison with the prior art.
[Explanation of symbols]
10. Aircraft gas turbine engine (turbofan engine)
12 fans
12a rotor
14 Stator
16 Low pressure compressor
24 High pressure compressor
24a rotor
24b Stator
26 Combustor
28 Fuel nozzle
30 FCU (Fuel Control Unit)
32 Fuel metering valve
40 High-pressure turbine
40a High-pressure turbine shaft
42 Low pressure turbine
42a Low pressure turbine shaft
62 N1 sensor (low-pressure turbine rotation speed detection means)
64 N2 sensor (high-pressure turbine rotation speed detection means)
68 T1 sensor
70 P1 sensor
72 P0 sensor
74 P3 sensor
76 ITT sensor
80 ECU (Electronic Control Unit)
82 Thrust lever (throttle lever)
92 Aircraft Computer (ADC)
98 torque motor
806 Speed feedback controller (steady control mode system)
808 Acceleration control unit (acceleration / deceleration mode control system)
810 P3 pressure controller (unsteady mode control system)
812 ITT temperature controller (unsteady mode control system)
814 Deceleration control unit (acceleration / deceleration mode control system)
820 selection circuit (command value selection means)

Claims (1)

In an aircraft gas turbine engine having at least one turbine,
a. An operating state detecting means for detecting an operating state of the gas turbine engine including at least the turbine rotational speed;
b. A pilot request thrust detecting means for detecting a pilot request thrust required by the pilot for the gas turbine engine;
c. A command for a fuel flow rate to be supplied to the gas turbine engine using at least an integral element so that a deviation between at least the detected turbine speed and a target turbine speed that realizes the detected driver's required thrust is reduced. Steady mode control system that calculates values,
d. An acceleration / deceleration mode control system for calculating a command value of a fuel flow rate to be supplied to the gas turbine engine using at least an integral element so that a deviation between the detected acceleration / deceleration rate of the turbine rotation speed and a target acceleration / deceleration rate is reduced; A non-stationary mode control system, including at least
e. Command value selection means for selecting either a command value calculated by the steady mode control system or a command value calculated by the acceleration / deceleration mode control system;
f. Fuel flow rate calculation means for calculating a flow rate of fuel supplied to the gas turbine engine based on the selected command value;
And g. Fuel supply means for supplying fuel to the gas turbine engine based on the calculated fuel flow rate;
A control device for a gas turbine engine comprising: an integral element used in the steady mode control system and the unsteady mode control system is removed, and an integral element is disposed in the fuel flow rate calculation means, and thus the fuel flow rate calculation The means is configured to calculate the fuel flow rate based on the selected command value using the integration element , and further in the acceleration / deceleration mode control system, a schedule value corresponding to the detected turbine speed control system for a gas turbine aeroengine, characterized by being configured to sum a.

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