Gas turbine engine flow control

Abstract

A method of controlling a gas turbine engine including receiving an instantaneous thrust demand for current operation of the engine, determining the inlet flow rate and/or the pressure ratio within the compressor of the engine and determining whether the inlet flow rate and/or the pressure ratio match the working line for the compressor. The angle of one or more vane of the compressor is adjusted according to a closed control loop if the inlet flow rate and/or pressure ratio lie outside said desired range in order to adjust the inlet inflow rate and/or pressure ratio to meet the working line. The fuel flow to the engine combustor is adjusted concurrently in order to meet the thrust demand.

Claims

1. A method of controlling a gas turbine engine, the method comprising: receiving an instantaneous thrust demand for current operation of the gas turbine engine; determining a flow rate and/or pressure ratio for a compressor of the gas turbine engine; determining whether the flow rate and/or the pressure ratio meets a predetermined operating condition for the compressor; and adjusting an angle of one or more aerofoil of the compressor according to a closed control loop if the flow rate and/or pressure ratio do not meet said predetermined operating condition in order to adjust the flow rate and/or pressure ratio towards said predetermined operating condition; and concurrently adjusting a fuel flow to a combustor of the gas turbine engine in order to meet the thrust demand by mathematically decoupling terms for adjusting the angle of the one or more airfoil of the compressor and terms for adjusting the fuel flow to the combustor of the gas turbine engine in order to adjust the angle of the one or more airfoil of the compressor and to adjust the fuel flow to the combustor of the gas turbine engine at a same time.

2. The method of claim 1, where the aerofoil comprises a stator vane and the method further comprises adjusting an angle of an array of stator vanes of the compressor.

3. The method of claim 1, where the flow rate and/or the pressure ratio are directly measured.

4. The method of claim 1, where the flow rate is estimated by measuring the compressor speed.

5. The method of claim 1, where the predetermined operating condition comprises a working line of a gas turbine engine and/or a threshold above the working line.

6. The method of claim 1, where a first closed control loop is implemented for adjusting the angle of the one or more aerofoil and a second closed control loop is implemented for adjustment of the fuel flow, wherein a common multivariable controller determines the adjustments for both the first and second closed control loops.

7. The method of claim 6, where the first and second control loops are operated in parallel based on the outputs of the common multivariable controller.

8. The method of claim 1, where concurrent control of the fuel flow is determined according to a control algorithm which comprises de-coupling of an aerofoil angle variable and a fuel flow variable.

9. The method of claim 8, where the de-coupling is achieved using input state linearization, the angle and the fuel flow comprising respective inputs in the control algorithm.

10. The method of claim 8, where the control algorithm perturbates the fuel flow and the angle of the aerofoil individually and measures the respective rates of response.

11. The method of claim 8, where the control algorithm implements the expression: Where: $\frac{d}{\mathrm{dt}}\left(\begin{array}{c}\mathrm{NH}\\ Z\end{array}\right)=\left(\begin{array}{cc}{g}_{11}& g_{12}\\ g_{21}& g_{22}\end{array}\right)\left(\begin{array}{c}\Delta W_f\\ \Delta \mathrm{VSV}\end{array}\right)$ NH is a determined spool speed Z is a determined excursion, where the excursion is the magnitude of the pressure ratio and inlet flow rate away from the predetermined range, g.sub.ii, g.sub.12, g.sub.21 and g.sub.22 are to be determined non-linear functions ΔW.sub.f is change in fuel flow ΔVSV is a change in aerofoil angle.

12. The method of claim 1, where a current thrust value is determined according to a spool speed of the engine.

13. A gas turbine engine for an aircraft comprising a control system arranged to operate according to the method of claim 1.

14. A controller for a gas turbine engine comprising machine readable instructions for controlling operation of the gas turbine engine by: receiving an instantaneous thrust demand for current operation of the gas turbine engine; determining a flow rate and/or a pressure ratio for a compressor of the gas turbine engine; determining whether the flow rate and/or the pressure ratio meet a predetermined operating condition of the compressor; and adjusting the angle of one or more aerofoil of the compressor according to a closed control loop if the flow rate and/or pressure ratio do not meet the predetermined operating condition in order to adjust the flow rate and/or pressure ratio towards said condition; and concurrently adjusting fuel flow to a combustor of the gas turbine engine in order to meet the thrust demand by mathematically decoupling terms for adjusting the angle of the one or more airfoil of the compressor and terms for adjusting the fuel flow to the combustor of the gas turbine engine in order to adjust the angle of the one or more aerofoil of the compressor and to adjust the fuel flow to the combustor of the gas turbine engine at a same time.

15. The controller of claim 14, wherein the controller further comprises a multivariable controller that is common to both the closed control loop for adjusting the angle of the one or more aerofoil and a further closed control loop for adjustment of the fuel flow.

16. A gas turbine engine for an aircraft comprising the controller of claim 14.

17. The gas turbine engine of claim 16, further comprising: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft.

18. The gas turbine engine of claim 17, wherein: the turbine is a first turbine, the compressor is a first compressor, and the core shaft is a first core shaft; the engine core further comprises a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor; and the second turbine, second compressor, and second core shaft are arranged to rotate at a higher rotational speed than the first core shaft.

Description

BRIEF DESCRIPTION

(1) Embodiments will now be described by way of example only, with reference to the Figures, in which:

(2) FIG. 1 shows a generic plot of pressure ratio against inlet flow for a conventional high pressure compressor of a gas turbine engine;

(3) FIG. 2 is a sectional side view of a gas turbine engine;

(4) FIG. 3 is a close up sectional side view of an upstream portion of a gas turbine engine;

(5) FIG. 4 is a partially cut-away view of a gearbox for a gas turbine engine;

(6) FIG. 5 shows an example control scheme according to the present disclosure;

(7) FIG. 6 shows a transient excursion in a plot of pressure ratio against inlet flow;

(8) FIG. 7 shows a transient excursion of reduced magnitude in a plot of pressure ratio against inlet flow;

(9) FIG. 8 shows a plot of pressure ratio against inlet flow with a modified working line.

DETAILED DESCRIPTION

(10) FIG. 2 illustrates a gas turbine engine 10 having a principal rotational axis 9. The engine 10 comprises an air intake 12 and a propulsive fan 23 that generates two airflows: a core airflow A and a bypass airflow B. The gas turbine engine 10 comprises a core 11 that receives the core airflow A. The engine core 11 comprises, in axial flow series, a low pressure compressor 14, a high-pressure compressor 15, combustion equipment 16, a high-pressure turbine 17, a low pressure turbine 19 and a core exhaust nozzle 20. A nacelle 21 surrounds the gas turbine engine 10 and defines a bypass duct 22 and a bypass exhaust nozzle 18. The bypass airflow B flows through the bypass duct 22. The fan 23 is attached to and driven by the low pressure turbine 19 via a shaft 26 and an epicyclic gearbox 30.

(11) In use, the core airflow A is accelerated and compressed by the low pressure compressor 14 and directed into the high pressure compressor 15 where further compression takes place. The compressed air exhausted from the high pressure compressor 15 is directed into the combustion equipment 16 where it is mixed with fuel and the mixture is combusted. The resultant hot combustion products then expand through, and thereby drive, the high pressure and low pressure turbines 17, 19 before being exhausted through the core exhaust nozzle 20 to provide some propulsive thrust. The high pressure turbine 17 drives the high pressure compressor 15 by a suitable interconnecting shaft 27. The fan 23 generally provides the majority of the propulsive thrust. The epicyclic gearbox 30 is a reduction gearbox.

(12) An exemplary arrangement for a geared fan gas turbine engine 10 is shown in FIG. 3. The low pressure turbine 19 (see FIG. 2) drives the shaft 26, which is coupled to a sun wheel, or sun gear, 28 of the epicyclic gear arrangement 30. Radially outwardly of the sun gear 28 and intermeshing therewith is a plurality of planet gears 32 that are coupled together by a planet carrier 34. The planet carrier 34 constrains the planet gears 32 to precess around the sun gear 28 in synchronicity whilst enabling each planet gear 32 to rotate about its own axis. The planet carrier 34 is coupled via linkages 36 to the fan 23 in order to drive its rotation about the engine axis 9. Radially outwardly of the planet gears 32 and intermeshing therewith is an annulus or ring gear 38 that is coupled, via linkages 40, to a stationary supporting structure 24.

(13) Note that the terms “low pressure turbine” and “low pressure compressor” as used herein may be taken to mean the lowest pressure turbine stages and lowest pressure compressor stages (i.e. not including the fan 23) respectively and/or the turbine and compressor stages that are connected together by the interconnecting shaft 26 with the lowest rotational speed in the engine (i.e. not including the gearbox output shaft that drives the fan 23). In some literature, the “low pressure turbine” and “low pressure compressor” referred to herein may alternatively be known as the “intermediate pressure turbine” and “intermediate pressure compressor”. Where such alternative nomenclature is used, the fan 23 may be referred to as a first, or lowest pressure, compression stage.

(14) The epicyclic gearbox 30 is shown by way of example in greater detail in FIG. 4. Each of the sun gear 28, planet gears 32 and ring gear 38 comprise teeth about their periphery to intermesh with the other gears. However, for clarity only exemplary portions of the teeth are illustrated in FIG. 4. There are four planet gears 32 illustrated, although it will be apparent to the skilled reader that more or fewer planet gears 32 may be provided within the scope of the claimed invention. Practical applications of a planetary epicyclic gearbox 30 generally comprise at least three planet gears 32.

(15) The epicyclic gearbox 30 illustrated by way of example in FIGS. 3 and 4 is of the planetary type, in that the planet carrier 34 is coupled to an output shaft via linkages 36, with the ring gear 38 fixed. However, any other suitable type of epicyclic gearbox 30 may be used. By way of further example, the epicyclic gearbox 30 may be a star arrangement, in which the planet carrier 34 is held fixed, with the ring (or annulus) gear 38 allowed to rotate. In such an arrangement the fan 23 is driven by the ring gear 38. By way of further alternative example, the gearbox 30 may be a differential gearbox in which the ring gear 38 and the planet carrier 34 are both allowed to rotate.

(16) It will be appreciated that the arrangement shown in FIGS. 3 and 4 is by way of example only, and various alternatives are within the scope of the present disclosure. Purely by way of example, any suitable arrangement may be used for locating the gearbox 30 in the engine 10 and/or for connecting the gearbox 30 to the engine 10. By way of further example, the connections (such as the linkages 36, 40 in the FIG. 2 example) between the gearbox 30 and other parts of the engine 10 (such as the input shaft 26, the output shaft and the fixed structure 24) may have any desired degree of stiffness or flexibility. By way of further example, any suitable arrangement of the bearings between rotating and stationary parts of the engine (for example between the input and output shafts from the gearbox and the fixed structures, such as the gearbox casing) may be used, and the disclosure is not limited to the exemplary arrangement of FIG. 3. For example, where the gearbox 30 has a star arrangement (described above), the skilled person would readily understand that the arrangement of output and support linkages and bearing locations would typically be different to that shown by way of example in FIG. 3.

(17) Accordingly, the present disclosure extends to a gas turbine engine having any arrangement of gearbox styles (for example star or planetary), support structures, input and output shaft arrangement, and bearing locations.

(18) Optionally, the gearbox may drive additional and/or alternative components (e.g. the intermediate pressure compressor and/or a booster compressor).

(19) Other gas turbine engines to which the present disclosure may be applied may have alternative configurations. For example, such engines may have an alternative number of compressors and/or turbines and/or an alternative number of interconnecting shafts. By way of further example, the gas turbine engine shown in FIG. 2 has a split flow nozzle 18, 20 meaning that the flow through the bypass duct 22 has its own nozzle 18 that is separate to and radially outside the core exhaust nozzle 20. However, this is not limiting, and any aspect of the present disclosure may also apply to engines in which the flow through the bypass duct 22 and the flow through the core 11 are mixed, or combined, before (or upstream of) a single nozzle, which may be referred to as a mixed flow nozzle. One or both nozzles (whether mixed or split flow) may have a fixed or variable area. Whilst the described example relates to a turbofan engine, the disclosure may apply, for example, to any type of gas turbine engine, such as an open rotor (in which the fan stage is not surrounded by a nacelle) or turboprop engine, for example. In some arrangements, the gas turbine engine 10 may not comprise a gearbox 30.

(20) The geometry of the gas turbine engine 10, and components thereof, is defined by a conventional axis system, comprising an axial direction (which is aligned with the rotational axis 9), a radial direction (in the bottom-to-top direction in FIG. 2), and a circumferential direction (perpendicular to the page in the FIG. 2 view). The axial, radial and circumferential directions are mutually perpendicular.

(21) FIG. 5 shows a control system 48 according to the present disclosure for operation of the engine 10.

(22) The control system 48 receives a thrust instruction signal 50 to specify a specific required thrust output from the engine. The thrust instruction signal 50 is generated by a pilot throttle actuator (e.g. by the angular position thereof) onboard the aircraft. In other examples, the thrust instruction signal 50 is generated by a control system on board the aircraft and/or the engine as is conventional.

(23) The control system 48 receives air data 51 relating to the properties of the air entering the engine 10, and/or flowing through the compressor(s), during the current operating conditions. This may include the pressure and/or temperature at one or more locations (e.g. at the engine intake 12, downstream of the fan 23, at the high pressure compressor 15 and/or at the low pressure compressor 14). The flow readings may be taken at one or more compressor stage as required.

(24) The current operating conditions may be directly measured using conventional means, for example, using temperature, pressure and/or flow sensors. In other embodiments, the current operating conditions are estimated, for example, the rotational speed of the compressor etc.

(25) The thrust instruction signal 50 is received by a first controller 52. The first controller 52 is configured to determine an instantaneous thrust demand for the engine according to the received signal 50.

(26) The thrust demand is received by a first control loop 54 configured to control the angle of one or more aerofoil of the compressor, and a second control loop 56 to control the fuel flow into a combustor of the engine. The following disclosure concerns how those control loops can be performed, e.g. concurrently and/or commonly, using multivariate control.

(27) The first control loop 54 comprises a second controller 58 to determine desired/optimum pressure and/or flow conditions in the compressor according to the thrust demand provided by the first controller 52 (e.g. the working line pressure, pressure ratio and/or flow rate).

(28) The desired working line condition may lie within a predetermined pressure ratio and/or flow rate range for a predetermined thrust, thereby providing a margin about the working line for which a transient event is considered acceptable.

(29) The second controller 58 determines a difference in the desired working line pressure and/or airflow at a given thrust demand and the current operating pressure and/or airflow. The second controller 58 therefore outputs a ΔP target 60 indicating a difference between desired and current pressures in the engine. Additionally or alternatively, the second controller 58 outputs a ΔV target (i.e. a flow rate change target) indicating the difference in desired and current airflow through the engine. The/each target may seek to return the compressor operation to the working line, or within an acceptable/small divergence therefrom.

(30) A multi-variable controller 62 receives the outputs from the second controller 58. The multi-variable controller 62 then calculates the appropriate angle for the variable stator vanes of the compressor, to minimise the difference in pressure/airflow between the working line and the current operating conditions.

(31) The angle signal is then sent to a valve system (e.g. servo valve 64 and/or fuel-draulic valve 66) to move the variable stator vanes 70 to the appropriate angle. The valve position is monitored by a position sensor within a position feedback loop 68 to ensure the vanes are in the correct position, and adjustment is performed as necessary.

(32) The second control loop 56 comprises a series of controllers 72, 74 configured to determine a difference between the desired thrust determined by the thrust demand and the current thrust produced by the engine. The controllers therefore output a “thrust target” 75, e.g. comprising a change in the current thrust needed to meet the thrust demand.

(33) Each controller may utilise an actuator to regulate its control parameters but conscious of the fact that cross coupling exists between both loops—which in this disclosure is addressed though de-coupling of this interaction in a multi-variable controller.

(34) The “thrust target” 75 is received by the multi-variable controller 62. The multi-variable controller 62 then calculates the appropriate fuel flow to maintain the desired thrust according to the current and/or desired engine parameters. This provides a measure of the required volumetric fuel flow into the combustor that can be used to control fuel delivery.

(35) The calculation of the fuel flow rate and the angle of the stator vanes are performed concurrently/simultaneously, such that adjustment of the fuel flow and the angle of the stator vanes is performed in concert to maintain the desired thrust output. Both the engine thrust and the compressor pressure ratio are controlled in a multi-variable manner. The control laws governing the control of the fuel flow and the angle of the stator vanes will be described below.

(36) The multi-variable controller 62 outputs a signal to a fuel control system 76 (e.g. comprising a fuel limiter) to control fuel flow to the engine accordingly.

(37) The fuel control system sends a signal to a fuel metering system (e.g. comprising torque motors 78 and/or a hydromechanical metering valve 80) to provide an adjusted fuel flow rate 82. The hardware of the fuel metering system may be conventional save for the manner in which it is controlled according to the present disclosure. The fuel metering system is monitored by a fuel feedback loop 84 to ensure the valves are in the correct position to provide the appropriate fuel flow, and adjustment is performed as necessary. The feedback loop 84 may make use of suitable sensors, e.g. measuring fuel flow rate and/or an operational condition of one or more components of the metering system. A pump or valve condition/position may be monitored, e.g. a fuel metering valve position.

(38) An output thrust of the engine is measured using conventional techniques. For example, this may be achieved by measuring one or more of: low pressure shaft speed; engine Pressure Ratio (EPR); or Turbofan Power Ratio (TPR).

(39) The output thrust is then monitored by a thrust feedback loop, e.g. comprising a controller 88. The thrust feedback monitors the thrust produced by the engine and forms part of the fuel control loop 56, i.e. feeding the thrust values back into the start of fuel control loop 56. The measured thrust is compared with the thrust demand and the fuel flow and/or the stator vane angle are adjusted accordingly.

(40) As described herein, both the fuel flow and the angle of the compressor vanes are controlled in a closed loop manner.

(41) Whilst individual ‘controllers’ are described herein for different stages/aspects of the control system 48, it will be appreciated by the skilled person that such controllers may or may not comprise separate computational algorithms. For example, controllers may comprise control modules of common or different computational processors. The order/sequence of steps within each control loop is as shown in FIG. 5 although it will be appreciated that the different control loops 54 and 56, including the feedback loops 68 and 84 thereof, can operate concurrently and/or in parallel.

(42) A control law governing the concurrent operation of the one or more compressor aerofoil angle and the fuel flow will now be described according to an example of the present disclosure.

(43) Concurrent/simultaneous control of the aerofoil angle and the fuel flow is achieved through mathematical de-coupling of the algorithms used by the multi-variable controller 62. As shown in FIG. 6, during a transient event 45 (dashed line), the pressure ratio and/or the inlet flow deviate from the working line 42. This deviation is defined as an excursion Z. The control law therefore attempts to minimise the excursion Z.

(44) Firstly, an algorithm is derived to identify a first order characteristic of the fuel flow to thrust relationship and/or the vane angle to pressure ratio/flow characteristic relationship. Additionally, the algorithm determines the cross-coupling terms between these two loops (e.g. between fuel flow and vane angle characteristics). The algorithm then mathematically de-couples these characteristics and allows fuel flow and vane angle actuation to be applied simultaneously to deliver the required trust demand and compressor pressure ratio/flow characteristic.

(45) According to a Multiple Input Multiple Output (MIMO) control law, the following system of equations (1) can be defined:
{dot over (x)}=f(x)+g(x)u
y=h(x)  (1)

(46) Where x is the internal state of the system, u is a controlled, input parameter of the system, y is the system output variable and f, g and h are to be determined dynamic response variables.

(47) A 2×2 MIMO system can be defined where:

(48) $x=\left(\begin{array}{c}\mathrm{NH}\\ Z\end{array}\right),$
where NH is the spool speed and Z is the excursion.

(49) $u=\left(\begin{array}{c}{W}_f\\ \mathrm{VSV}\end{array}\right),$
where W.sub.f is the fuel flow and VSV is the angle of the variable stator vane.

(50) Substituting the above definitions into equation (1) for a 2×2 MIMO system, yields:

(51) $\begin{array}{cc}\frac{d}{\mathrm{dt}}\left(\begin{array}{c}\mathrm{NH}\\ Z\end{array}\right)=\left(\begin{array}{c}{f}_1\left(\mathrm{NH},Z\right)\\ f_2\left(\mathrm{NH},Z\right)\end{array}\right)+\left(\begin{array}{c}{g}_1\left(\mathrm{NH},Z\right)\\ g_2\left(\mathrm{NH},Z\right)\end{array}\right)\left(\begin{array}{c}{W}_f\\ \mathrm{VSV}\end{array}\right)& \left(2\right)\\ \left(\begin{array}{c}{y}_1\\ y_2\end{array}\right)=\left(\begin{array}{c}\mathrm{NH}\\ Z\end{array}\right)& \left(3\right)\end{array}$

(52) where

(53) $\left(\begin{array}{c}{g}_1\\ g_2\end{array}\right)\mathrm{and}\left(\begin{array}{c}{f}_1\\ f_2\end{array}\right)$
are respective to be determined non-linear functions.

(54) The multi-variable de-coupling is achieved by recognising that in some cases we can simplify the dynamics by input-state linearization. The rate of change of output variable v is chosen judiciously as follows:

(55) $\begin{array}{cc}\left(\begin{array}{c}{v}_1\\ v_2\end{array}\right)=\frac{d}{\mathrm{dt}}\left(\begin{array}{c}{y}_1\\ y_2\end{array}\right)=\frac{d}{\mathrm{dt}}\left(\begin{array}{c}\mathrm{NH}\\ Z\end{array}\right)& \left(4\right)\end{array}$

(56) which in conjunction with (2) and (3) leads to

(57) $\begin{array}{cc}\left(\begin{array}{c}{W}_f\\ \mathrm{VSV}\end{array}\right)=-{\left(\begin{array}{c}{g}_1\left(\mathrm{NH},Z\right)\\ g_2\left(\mathrm{NH},Z\right)\end{array}\right)}^{-1}\left(\begin{array}{c}{f}_1\left(\mathrm{NH},Z\right)\\ f_2\left(\mathrm{NH},Z\right)\end{array}\right)+{\left(\begin{array}{c}{g}_1\left(\mathrm{NH},Z\right)\\ g_2\left(\mathrm{NH},Z\right)\end{array}\right)}^{-1}\left(\begin{array}{c}{v}_1\\ v_2\end{array}\right)& \left(5\right)\end{array}$

(58) where f(NH,Z) and g(NH,Z) are to be determined.

(59) Using equation (2), for a perturbation about an operating point W.sub.f.sub.0 and VSV.sub.0 of ΔW.sub.f and ΔVSV, respectively, yields:

(60) $\begin{array}{cc}\frac{d}{\mathrm{dt}}\left(\begin{array}{c}\mathrm{NH}\\ Z\end{array}\right)=f\left({\mathrm{NH}}_0,Z_0\right)+g\left({\mathrm{NH}}_0,Z_0\right)\left(\begin{array}{c}{W}_{f_0}+\Delta W_f\\ {\mathrm{VSV}}_0+\Delta \mathrm{VSV}\end{array}\right)& \left(6\right)\end{array}$

(61) For small perturbations

(62) $\left(\begin{array}{c}\Delta W_f\\ \Delta \mathrm{VSV}\end{array}.fwdarw.0\right)$
at the equilibrium condition NH.sub.0, Z.sub.0, therefore:

(63) $\begin{array}{cc}0=f\left({\mathrm{NH}}_0,Z_0\right)+g\left({\mathrm{NH}}_0,Z_0\right)\left(\begin{array}{c}{W}_{f_0}\\ {\mathrm{VSV}}_0\end{array}\right)& \left(7\right)\end{array}$

(64) Substituting equation (7) into equation (6) yields:

(65) 0$\begin{array}{cc}\frac{d}{\mathrm{dt}}\left(\begin{array}{c}\mathrm{NH}\\ Z\end{array}\right)=\left(\begin{array}{cc}{g}_{11}& g_{12}\\ g_{21}& g_{22}\end{array}\right)\left(\begin{array}{c}\Delta W_f\\ \Delta \mathrm{VSV}\end{array}\right)=\left(\begin{array}{c}{v}_1\\ v_2\end{array}\right)& \left(8\right)\end{array}$

(66) By perturbing the fuel flow W.sub.f and the angle of the variable stator VSV individually and measuring the respective rates of response, g(NH,Z) can be mapped out across the respective operating points. Using equation (7) an appropriate mapping for f(NH,Z) can be made across the respective operating points. The control law in equation (5) can then be implemented using the virtual input v.

(67) But choosing v as a set of compensated loop outputs thus

(68) $\begin{array}{cc}v=\left(\begin{array}{c}\alpha \left({\mathrm{NH}}_d-\mathrm{NH}\right)\\ \beta \left(Z_d-Z\right)\end{array}\right)& \left(11\right)\end{array}$

(69) where: α and β are constants NH.sub.d is the spool speed at a set point (e.g. the working line speed) and Z.sub.d is the excursion at a set point (e.g. the working line pressure ratio/flow rate)

(70) This provides a closed loop dynamic for NH and Z that is first order, decoupled and independent of the operating point.

(71) The above control law allows the transient excursion ‘Z’ to be actively controlled in a continuous manner throughout the transient event (e.g. engine maneuver, bird strike, ice intake etc.), whilst maintaining the predetermined thrust demand. The system therefore actively controls the pressure ratio/inlet flow within the compressor during an excursion back toward the desired working line pressure ratio/inlet flow.

(72) The controller therefore utilises an actuator to regulate the control of the fuel flow and the vane angle respectively, whilst taking into consideration cross coupling between both control loops.

(73) The system allows a change in one of the fuel flow and the vane angle without affecting the thrust output, by providing a simultaneous reactive change in the other of the fuel flow and the vane angle accordingly.

(74) As shown in the FIG. 7, the concurrent, closed loop control of the pressure ratio and/or fuel flow provide closer control of the engine thrust and reduces the magnitude of the transient excursion 45 within the engine. Therefore, for a nominal transient event (i.e. of a magnitude demonstrated in FIG. 1), the magnitude of the transient excursion 45 can be significantly reduced. This will allow the engine to operate within an increased safety margin.

(75) The duration of the transient event may also be reduced.

(76) Alternatively, as shown in FIG. 8, due to the reduced magnitude of the transient excursion 45, the surge margin 46 may be increased. This may allow the working line 42 to be operated safely at an increased pressure ratio, i.e. closer to the surge line 44, thereby increasing the efficiency of the engine.

(77) It will be understood that the invention is not limited to the embodiments above-described and various modifications and improvements can be made without departing from the concepts described herein. Except where mutually exclusive, any of the features may be employed separately or in combination with any other features and the disclosure extends to and includes all combinations and sub-combinations of one or more features described herein.

Advantages of the Invention

(78) The control method of the present disclosure may permit improved operating efficiency of a gas turbine engine.

(79) The control method of the present disclosure may allow adoption of a reduced surge margin for a gas turbine engine and/or may allow greater safety at a conventional surge margin.

(80) The present disclosure may allow control of the pressure ratio/airflow in a gas turbine engine without detriment to the operation of a primary thrust control system (i.e. without interfering with a thrust demand of an engine)

(81) The present disclosure may allow independent control of thrust and surge margin excursion in a gas turbine engine.

(82) The present disclosure may reduce the time taken for a control system to react to a transient air flow event in a gas turbine engine.

(83) The present disclosure may provide active, closed loop control of a transient excursion from a steady state operating condition.

(84) The present disclosure may allow a reduction in the specific fuel consumption of the engine.