JP4831586B2 - Control device for aircraft gas turbine engine - Google Patents

Description

  The present invention relates to a control device for an aircraft gas turbine engine, and in particular, has a misfire determination means for determining whether or not a combustor is in a misfire state. When the misfire determination means determines that a misfire has occurred, a required reignition is performed. The present invention relates to a control device for an aircraft gas turbine engine that performs an operation.

  In an aircraft gas turbine engine, if a combustor misfires during flight, changes in the engine speed, gas temperature, etc. appear, so it is possible to determine whether or not a misfire has occurred based on these measured values. (See Patent Documents 1 and 2).

JP 58-79627 A JP 59-18241

  However, the engine speed and gas temperature, which serve as indicators for misfire determination, vary greatly depending on flight conditions (altitude and speed) and engine operating conditions (rotational speed), making it difficult to distinguish between normal operating conditions, especially deceleration. It is difficult to make misfire determination accurately in a short time.

  The present invention has been devised to solve such problems of the prior art, and its purpose is to perform misfire determination accurately in a short time regardless of flight conditions and engine operating conditions. An object of the present invention is to provide a control device for an aircraft gas turbine engine configured to be capable of performing the above.

The first invention made to solve the above-described problem has misfire determination means (3) for determining whether or not the combustor (2) is in a misfire state, and the misfire determination means determines that misfire has occurred. Then, in the control device for an aircraft gas turbine engine (1) that performs a predetermined reignition operation, the misfire determination means is preset with operating states represented by two characteristic values of gas temperature and engine speed. there rows determined by whether the misfire region characteristic value of the gas temperature, the rate of change of the gas temperature engine inlet temperature, a configuration obtained by modifying a highly and machine speed.

  According to this, since it is possible to accurately determine the misfire state and the deceleration state by a simple procedure by performing the determination in the operation state represented by the two types of characteristic values of temperature and rotation speed, The time required can be shortened and at the same time the determination accuracy can be increased.

A second invention has been made in order to solve the above problems, in the first invention, the engine speed characteristic value, a configuration obtained by modifying the engine speed at the engine inlet temperature.

  Thus, according to the present invention, the temperature characteristic value obtained by correcting the rate of change (derivative) of the gas temperature with the engine inlet temperature, altitude and speed, and the engine rotational speed with the engine inlet temperature are obtained. Therefore, the misfire determination can be performed with high accuracy in a short time regardless of the flight condition and the operating state of the engine.

The block diagram which shows the control apparatus of the gas turbine engine for aircrafts by this invention. The graph which shows the area | region used as the reference | standard of misfire determination by this invention. The graph which shows the area | region used as the reference | standard of reignition condition determination by this invention. The flowchart which shows the procedure of control by this invention.

  Hereinafter, embodiments of the present invention will be described with reference to the drawings.

  FIG. 1 shows a control apparatus for a gas turbine engine to which the present invention is applied. Here, it is determined in the misfire determination block (misfire determination means) 3 whether or not the combustor 2 of the gas turbine engine 1 is in a misfire state, and if it is determined that there is a misfire here, a reignition condition determination block (reignition) Condition determining means) 4 determines a re-ignitable condition and performs a predetermined re-ignition operation according to the re-ignition condition obtained here.

Gas turbine engine 1 comprises a high-pressure shaft 8 for connecting the compressor 6 and the high-pressure turbine 7, and a low pressure shaft 11 which connects the low pressure turbine 9 and the fan 10, the engine inlet temperature T ₁ is the engine inlet temperature detecting unit 12 The engine inlet pressure P ₁ is detected by the engine inlet pressure detector 13, the rotational speed (engine speed) N of the high-pressure shaft 8 is detected by the engine rotational speed detector 14, and the inlet of the low-pressure turbine 9 is detected. The temperature (gas temperature) T of the portion is detected by the gas temperature detector 15. In addition, the atmospheric pressure P ₀ is detected by the atmospheric pressure detector 16, the altitude ALT is detected by the altitude detector 17, and the aircraft speed Mn is detected by the aircraft speed detector 18. Incidentally, altitude ALT from atmospheric pressure P _0, aircraft speed Mn may be calculated respectively from the atmospheric pressure P ₀ and the engine inlet pressure P _1.

In the misfire determination block (misfire determination means) 3, as shown in FIG. 2, the determination is made based on whether or not the operation state represented by the two characteristic values of the temperature and the rotational speed is within a preset misfire region. The temperature characteristic value TC is obtained by correcting the derivative (rate of change) Td of the low-pressure turbine inlet temperature (gas temperature) T with the engine inlet temperature T ₁ , the altitude ALT and the machine speed Mn in the temperature characteristic value calculator 20. The obtained engine speed characteristic value NC is obtained by correcting the engine speed N with the engine inlet temperature T ₁ in the engine speed characteristic value calculation block 21.

ここで、Ｔ₁／２８８．１５はエンジン入口温度、Ｐ₀／１．０３３は高度、Ｐ₁／Ｐ₀は機速に基づく変量である。また第２項は次式で得られる。
Ｃ（Ｍｎ，Ｔｄ）＝（−０．５９３４×Ｍｎ²−（２．３７２Ｅ^-17）×Ｍｎ）×Ｔｄ＋１３．５２×Ｍｎ² （式２） Specifically, the temperature characteristic value TC is obtained by the following equation.

Here, T ₁ /288.15 is an engine inlet temperature, P ₀ /1.033 is an altitude, and P ₁ / P ₀ is a variable based on the aircraft speed. The second term is obtained by the following equation.
C (Mn, Td) = (− 0.5934 × Mn ² − (2.372E ⁻¹⁷ ) × Mn) × Td + 13.52 × Mn ² (Formula 2)

The rotational speed characteristic value NC is obtained by the following equation.

The temperature characteristic value calculation unit 20 generates a turbine inlet temperature derivative Td every 40 ms control cycle based on the low pressure turbine inlet temperature T obtained by the gas temperature detection unit 15 in the derivative calculation block 22. In the temperature characteristic value first term calculation block 23, the turbine inlet temperature derivative Td obtained in the derivative calculation block 22, the engine inlet temperature T ₁ from the engine inlet temperature detector 12, and the engine inlet from the engine inlet pressure detector 13. Based on the pressure P ₁ and the atmospheric pressure P ₀ from the atmospheric pressure detector 16, the first term of the temperature characteristic value represented by Equation 1 is calculated. In the temperature characteristic value second term calculation block 24, the second term of the temperature characteristic value represented by Equation 1 is calculated based on the machine speed Mn and the turbine inlet temperature derivative Td. At the addition point 25, the temperature characteristic value TC is obtained by adding the value of the first term obtained by the temperature characteristic value first term calculation block 23 and the value of the second term obtained by the temperature characteristic value second term calculation block 24. calculate.

In the rotational speed characteristic value calculation block 21, the rotational speed characteristic value NC is obtained from Equation 3 based on the engine rotational speed N obtained by the engine rotational speed detection unit 14 and the engine inlet temperature T ₁ obtained by the engine inlet temperature detection unit 12. Is calculated.

  The misfire determination line L1 shown in FIG. 2 can be obtained by simulation. When simulation is performed under various engine inlet conditions, both characteristic values of temperature and rotational speed are gathered on the misfire line L2 in the case of misfire, and both characteristic values are gathered on the deceleration line L3 in the case of deceleration. By setting the misfire determination line L1 at an intermediate position between the misfire and deceleration lines L2 and L3, it is possible to accurately determine misfire and deceleration.

  A re-ignition condition determination block (re-ignition condition determination means) 4 determines a re-ignitable condition based on the speed Mn, the altitude ALT, and the engine speed N. Here, as shown in FIG. 3, a determination is made based on which of the three states preset in the flight envelope the flight state represented by the speed Mn and the altitude ALT. Is a speed limit area where the engine speed is likely to ignite at a high altitude, a starter assist area that requires a starter assist operation to assist the starter at a low speed and increase the engine speed, and a high speed. This is a windmill area that does not require starter assist operation.

  The determination line L4 that becomes the boundary between the rotation speed limit region, the windmill region, and the starter assist region is an engine that affects the ignition condition of the combustor, that is, the inlet condition (inlet pressure and inlet air temperature) of the combustor and the air flow rate. The engine speed N is determined uniformly from the engine speed, and the engine speed N influences the possibility of ignition in an altitude area higher than the determination line L4 (the engine speed limit area), and the engine speed N falls below a predetermined value. Cannot be re-ignited. In a region at an altitude lower than the determination line L4 (wind mill region and starter assist region), the combustor alone can always ignite.

  On the other hand, the determination line L5 that becomes the boundary between the windmill region and the starter assist region is determined by the discharge condition of the fuel pump and the flow rate condition of the fuel nozzle. Even if the combustor can always be ignited, if the fuel pump discharge amount is insufficient, it will be impossible to ignite, and even if the fuel flow rate can be discharged by the fuel pump, the fuel nozzle may ignite if the fuel flow rate that can be sufficiently atomized is not obtained. Can not. Since the discharge amount of the fuel pump is uniformly determined by the engine speed, the required speed is determined by the wind mill in the region where the machine speed is higher than the judgment line L5 corresponding to the engine speed which is the lower limit of the ignitable discharge amount (wind mill region). A starter that can always ignite because the engine speed can be obtained, and that the required engine speed cannot be obtained with the wind mill in the area where the machine speed is lower than the judgment line L5 (starter assist area). Assist operation is required.

The misfire determination and reignition condition determination described above and the reignition operation performed subsequently may be performed according to the procedure shown in FIG. First, the turbine inlet temperature T, the engine inlet temperature T ₁ , the engine inlet pressure P ₁ , the atmospheric pressure P ₀ , and the machine speed Mn are detected (step 1), and misfire determination is performed according to the above procedure based on these (step 2). ). If it is determined that there is no misfire, the process returns to step 1. If it is determined that there is a misfire, the fuel flow controller 26 operates the control valve 27 to block the flow of fuel into the combustor 2 (step 3). .

  Next, the machine speed Mn, the altitude ALT, and the engine speed N are detected (step 4), and based on these, the re-ignition condition is determined by the above procedure (step 5). If it is determined that re-ignition is impossible because the flight state is in the speed limit range and the engine speed N falls below a predetermined lower limit, a required operation such as displaying a warning and lowering the altitude to the pilot (Step 6).

  If it is determined in step 5 that the flight state is in the windmill region, or the flight state is in the rotational speed limit region and the engine rotational speed N exceeds a predetermined lower limit value, and it is determined that the windmill starts, the ignition controller 28 The spark plug 29 is operated (step 7), the fuel flow controller 26 operates the control valve 27 to start supplying fuel (step 8), and then the operation of the spark plug 29 is stopped (step 9). ).

  If it is determined in step 5 that the flight state is in the starter assist region and the starter assist start is determined, the starter controller 30 operates the starter 31 (step 10), and the ignition controller 28 sets the spark plug 29. (Step 11), the fuel flow controller 26 operates the control valve 27 to start supplying fuel (step 12), and then the operation of the spark plug 29 is stopped (step 13). The operation is stopped (step 14).

  Even if the flight state is in the starter assist region, the engine speed N may actually exceed a predetermined lower limit (for example, 3000 rpm). In this case, reignition may be performed without performing the starter assist operation. Since it is possible, the re-ignition operation may be performed in the same procedure as the windmill start.

1 Gas Turbine Engine 2 Combustor 3 Misfire Determination Block (Misfire Determination Means)
4 Re-ignition condition determination block (re-ignition condition determination means)
7 High-pressure turbine 8 High-pressure shaft 9 Low-pressure turbine 11 Low-pressure shaft 12 Engine inlet temperature detector 13 Engine inlet pressure detector 14 Engine speed detector 15 Gas temperature detector 20 Temperature characteristic value calculator 21 Rotational speed characteristic value calculation block 31 Starter

Claims (2)

A control device for an aircraft gas turbine engine that includes a misfire determination unit that determines whether or not a combustor is in a misfire state, and performs a predetermined reignition operation when the misfire determination unit determines a misfire,
The misfire determining means, have rows determined by whether in the misfire region operating state represented by two characteristic values of the gas temperature and the engine speed is set in advance, the characteristic value of the gas temperature A control apparatus for an aircraft gas turbine engine, which is obtained by correcting the rate of change in gas temperature with the engine inlet temperature, altitude and speed . 2. The aircraft gas turbine engine control device according to claim 1, wherein the characteristic value of the engine speed is obtained by correcting the engine speed with an engine inlet temperature .

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