Patent Description:
Within the compressor of an aircraft engine, such as a gas turbine engine, air is typically channelled through circumferential rows of vanes and blades that pressurize the air in stages. Variable guide vanes ("VGV" or "VGVs") are sometimes used within compressors, and provide vanes which are rotatable such that the angle of attack they define with the incoming flow may be varied. A control system adjusts an angle of the VGVs as a function of operating conditions of the engine. Improvements are sought.

Documents <CIT> and <CIT> relate to known variable guide vane control systems.

According to an aspect of the present invention, there is provided a method of operating a variable guide vane assembly of an aircraft engine, the variable guide vane assembly including guide vanes rotatable about respective spanwise axes and circumferentially distributed about a central axis, the method comprising: obtaining a target exit flow angle defined between a direction of a flow exiting the guide vanes and the central axis; predicting an exit flow angle as a function of at least a geometric angle, the exit flow angle defined between the direction of the flow exiting the guide vanes and the central axis, the geometric angle defined between the guide vanes and the central axis; and when a difference between the exit flow angle and the target exit flow angle is above a threshold, modulating the guide vanes to modify the geometric angle until the difference between the exit flow angle and the target exit flow angle is at or below the threshold.

The method defined above and described herein may also include one or more of the following features.

In an embodiment of the above, the predicting of the exit flow angle as a function of at least the geometric angle includes determining the exit flow angle as a function of at least the geometric angle and as a function of operating parameters of the aircraft engine.

In another embodiment according to any of the previous embodiments, the operating parameters include aircraft parameters including one or more of a spatial orientation of the aircraft engine, a location of the aircraft engine in relation to an aircraft equipped with the aircraft engine, and a configuration of an inlet of the aircraft engine.

In another embodiment according to any of the previous embodiments, the operating parameters include engine parameters including one or more of a power output of the aircraft engine and a rotational speed of a shaft of the aircraft engine.

In another embodiment according to any of the previous embodiments, the operating parameters include flight parameters including one or more of an altitude of the aircraft engine, an airspeed of the aircraft engine, and a temperature of air entering the aircraft engine.

In another embodiment according to any of the previous embodiments, the predicting of the exit flow angle includes obtaining the exit flow angle from a lookup table correlating geometric angle values and operating parameters values with exit flow angles values.

In another embodiment according to any of the previous embodiments, the modulating of the guide vanes to modify the geometric angle includes: a) causing rotation of the guide vanes from a first position to a second position; b) determining an updated geometric angle of the guide vanes at the second position; c) determining, based on at least the updated geometric angle of the guide vanes, an updated exit flow angle; d) repeating steps a) to c) if a difference between the target exit flow angle and the updated exit flow angle is greater than the threshold until the difference is less than the threshold.

In another embodiment according to any of the previous embodiments, the obtaining of the target exit flow angle includes determining the target exit flow angle as a function of operating parameters of the aircraft engine.

In another embodiment according to any of the previous embodiments, the operating parameters include one or more of engine parameters and flight parameters, the engine parameters including one or more of a power output of the aircraft engine and a rotational speed of a shaft of the aircraft engine, the flight parameters including one or more of an altitude of the aircraft engine, an airspeed of the aircraft engine, and a temperature of air entering the aircraft engine.

In another embodiment according to any of the previous embodiments, the determining of the target exit flow angle includes determining the target exit flow angle from a lookup table correlating operating parameters values with target exit flow angles values.

According to another aspect of the present invention, there is provided an aircraft engine comprising: a variable guide vane assembly including guide vanes rotatable about respective spanwise axes and circumferentially distributed about a central axis; an actuator drivingly engaged to the guide vanes for rotating the guide vanes about the respective spanwise axes; a sensor operable to send a signal indicative of a geometric angle of the guide vanes, the geometric angle defined between the guide vanes and the central axis; and a controller operatively connected to the actuator and to the sensor, the controller having a processing unit and a computer-readable medium operatively connected to the processing unit and having instructions stored thereon configured to cause the processing unit to: obtain a target exit flow angle defined between a direction of a flow exiting the guide vanes and the central axis; predicting an exit flow angle as a function of at least a geometric angle, the exit flow angle defined between the direction of the flow exiting the guide vanes and the central axis, the geometric angle determined based on the signal from the sensor; and when a difference between the exit flow angle and the target exit flow angle is above a threshold, power the actuator to modulate the guide vanes to modify the geometric angle until the difference between the predicted exit flow angle and the target exit flow angle is at or below the threshold.

The aircraft engine defined above and described herein may also include one or more of the following features.

In an embodiment of the above, the instructions are configured to cause the processing unit to predict the exit flow angle as a function of at least the geometric angle by causing the processing unit to determine the exit flow angle as a function of at least the geometric angle and as a function of operating parameters of the aircraft engine.

In another embodiment according to any of the previous embodiments, the instructions are configured to cause the processing unit to determine the exit flow angle by causing the processing unit to obtain the exit flow angle from a lookup table correlating geometric angle values and operating parameters values with exit flow angles values.

In another embodiment according to any of the previous embodiments, the instructions are configured to cause the processing unit to the modulate the guide vanes to modify the geometric angle by causing the processing unit to: a) cause rotation of the guide vanes from a first position to a second position; b) determine an updated geometric angle of the guide vanes at the second position; c) determine, based on at least the updated geometric angle of the guide vanes, an updated exit flow angle; d) repeat steps a) to c) if a difference between the target exit flow angle and the updated exit flow angle is greater than the threshold until the difference is less than the threshold.

The following disclosure relates generally to gas turbine engines, and more particularly to assemblies including one or more struts and variable orientation guide vanes as may be present in a compressor section of a gas turbine engine. In some embodiments, the assemblies and methods disclosed herein may promote better performance of gas turbine engines, such as by improving flow conditions in the compressor section in some operating conditions, improving the operable range of the compressor, reducing energy losses and aerodynamic loading on rotors.

<FIG> illustrates an aircraft engine depicted as a gas turbine engine <NUM> (in this case, a turboprop) of a type preferably provided for use in subsonic flight, and in driving engagement with a rotatable load, which is depicted as a propeller <NUM>. The gas turbine engine has in serial flow communication a compressor section <NUM> for pressurizing the air, a combustor <NUM> in which the compressed air is mixed with fuel and ignited for generating an annular stream of hot combustion gases, and a turbine section <NUM> for extracting energy from the combustion gases.

It should be noted that the terms "upstream" and "downstream" used herein refer to the direction of an air/gas flow passing through an annular gaspath <NUM> of the gas turbine engine <NUM>. It should also be noted that the term "axial", "radial", "angular" and "circumferential" are used with respect to a central axis <NUM> of the gaspath <NUM>, which may also be a central axis of gas turbine engine <NUM>. The gas turbine engine <NUM> is depicted as a reverse-flow engine in which the air flows in the annular gaspath <NUM> from a rear of the engine <NUM> to a front of the engine <NUM>, relative to a direction of travel T of the engine <NUM>. This is opposite than a through-flow engine in which the air flows within the gaspath in a direction opposite the direction of travel T, from the front of the engine towards the rear of the engine <NUM>. The principles of the present disclosure can be applied to both reverse-flow and through-flow engines and to any other gas turbine engines, such as a turbofan engine and a turboprop engine.

<FIG> illustrates an aircraft <NUM> equipped with the gas turbine engine <NUM> of <FIG>. The aircraft <NUM> is shown equipped with two of the gas turbine engine <NUM> of <FIG>, but may be equipped with only one, or more than two engines in other embodiments. The aircraft <NUM> includes a fuselage <NUM>, a left wing <NUM> secured to the fuselage and a right wing <NUM> secured to the fuselage <NUM>. As shown, the gas turbine engines <NUM> are mounted to the wings <NUM>, <NUM>. The gas turbine engines may alternatively be turbofan engines mounted to the fuselage <NUM> as depicted with dashed lines.

Referring now to <FIG>, an enlarged view of a portion of the compressor section <NUM> is shown. The compressor section <NUM> includes a plurality of stages, namely three in the embodiment shown although more or less than three stages is contemplated, each stage including a stator <NUM> and a rotor <NUM>. The rotors <NUM> are rotatable relative to the stators <NUM> about the central axis <NUM>. Each of the stators <NUM> includes a plurality of vanes <NUM> circumferentially distributed about the central axis <NUM> and extending into the gaspath <NUM>. Each of the rotors <NUM> also includes a plurality of blades <NUM> circumferentially distributed around the central axis <NUM> and extending into the gaspath <NUM>, the rotors <NUM> and thus the blades <NUM> thereof rotating about the central axis <NUM>. As will be seen in further detail below, at least one of the stators <NUM> includes vanes <NUM> which are variable guide vanes (VGVs) and thus includes a variable guide vane assembly as will be described.

In the depicted embodiment, the gaspath <NUM> is defined radially between an outer wall or casing <NUM> and an inner wall or casing <NUM>. The vanes <NUM> and the blades <NUM> extend radially relative to the central axis <NUM> between the outer and inner casings <NUM>, <NUM>. "Extending radially" as used herein does not necessarily imply extending perfectly radially along a ray perfectly perpendicular to the central axis <NUM>, but is intended to encompass a direction of extension that has a radial component relative to the central axis <NUM>. The vanes <NUM> can be fixed orientation or variable orientation guide vanes (referred hereinafter as VGVs). Examples of rotors include fans, compressor rotors (e.g. impellers), and turbine rotors (e.g. those downstream of the combustion chamber). Other orientations of the vanes (e.g., axial) are contemplated.

Although illustrated as a turboprop engine, the gas turbine engine <NUM> may alternatively be another type of engine, for example a turbofan engine or a turboshaft engine, also generally comprising in serial flow communication a compressor section, a combustor, and a turbine section, and a fan through which ambient air is propelled.

Referring to <FIG>, an example of a variable guide vane (VGV) assembly of a stator <NUM> of the engine <NUM> is shown at <NUM>. Any of the stators <NUM> of the compressor section <NUM> depicted in <FIG> may be embodied as a variable guide vane (VGV) assembly <NUM>. It will be appreciated that, in some cases, the VGV assembly <NUM> may be used as a stator of the turbine section <NUM> of the engine <NUM> without departing from the scope of the present disclosure. The VGV assembly <NUM> may be located at an upstream most location L1 (<FIG>) of the compressor section <NUM>. That is, the VGV assembly <NUM> may be a variable inlet guide vane assembly.

The VGV assembly <NUM> includes a plurality of guide vanes <NUM> circumferentially distributed about the central axis <NUM> and extending radially between the inner casing <NUM> and the outer casing <NUM>. In the present embodiment, the guide vanes <NUM> are rotatably supported at both of their ends by the inner casing <NUM> and the outer casing <NUM>. Particularly, each of the guide vanes <NUM> has an airfoil having a leading edge and a trailing edge both extending along a span of the airfoil. Each of the guide vanes <NUM> has an inner stem <NUM>, also referred to as an inner shaft portion, secured to an inner end of the airfoil and an outer stem <NUM>, also referred to as an outer shaft portion, secured to an outer end of the airfoil. The guide vanes <NUM> are rotatable about respective spanwise axes S1. One of the guide vanes <NUM>, which may be referred to as a master guide vane, has its outer stem <NUM> engaged by a vane arm <NUM>, which is itself drivingly engaged by an actuator <NUM> for pivoting the master vane about it spanwise axis S1. In the present embodiment, the vanes have gears <NUM> secured to the inner stems <NUM>. The gears <NUM> are meshed with a unison gear <NUM>, which is rollingly engaged to the inner casing <NUM>. Upon rotation of the master vane about its spanwise axis S1 via the actuator <NUM> engaged to the vane arm <NUM>, the gear <NUM> of the master vane rotates thereby induces rotation of the unison gear <NUM>, which extends annularly around the central axis <NUM>. Rotation of the unison gear <NUM> induces rotation of each of the other gears <NUM> and, consequently, of the other guide vanes <NUM>, which may be referred to as slave vanes, about their respective spanwise axes S1. Therefore, the unison gear <NUM> ensures that the rotation of all the guide vanes <NUM> is synchronized. Any suitable means for rotating the guide vanes <NUM> about their respective spanwise axes S1 are contemplated. The unison gear <NUM> may be located radially outwardly of the outer casing <NUM> in another embodiment. The unison gear may be replaced by any suitable unison member without departing from the scope of the present disclosure.

The variable guide vane assembly <NUM> is used to properly orient the flow before it meets blades of a rotor located downstream of the variable guide vane assembly <NUM>. Put differently, the flow is redirected by the variable guide vane assembly <NUM> so that an incidence angle between the flow and the downstream blades is optimal. This incidence angle varies with operating parameters of the gas turbine engine <NUM>. Namely, flight parameters, such as altitude, airspeed, air temperature, and engine parameters, such as power and speed, are expected to influence the incidence angle at which the flow should meet the blades.

A control system may be used to continuously adjust a geometric angle of the guide vanes <NUM> as a function of those parameters. The geometric angle is defined between the guide vanes <NUM>, namely their chord lines, and the central axis <NUM> of the gas turbine engine <NUM>. The control system may be supplied with the various parameters and compute a target angle for the guide vanes <NUM>. The control system then modulates the rotation of the guide vanes <NUM> and a sensor sends a signal to the control system; the signal indicative of the geometric angle of the guide vanes <NUM>. The control system continues to module the rotation of the guide vanes <NUM> until the signal provided by the sensor is indicative that the geometric angle of the guide vanes <NUM> corresponds to the target angle.

Referring now to <FIG>, in some conditions, an incoming airflow F0 extends along a direction that is devoid of a circumferential component relative to the central axis <NUM> of the gas turbine engine <NUM>. The guide vanes <NUM> deviates the flow and an outgoing airflow F1 exits flow passages defined between the guide vanes <NUM> along a direction that comprises a circumferential component relative to the central axis <NUM>. In this case, a difference between a geometric angle T1 of the guide vanes <NUM>, which is an angle defined between the central axis <NUM> and a chord line CL of the guide vanes <NUM>, and an exit flow angle T2, which is an angle defined between the central axis <NUM> and a direction of the outgoing airflow F1, is expected to be known or at least be considered constant for a given flow velocity and density.

Referring now to <FIG>, it has been observed that, in some cases, the incoming airflow F0' may extend along a direction defining a swirl angle T3. The swirl angle T3 is defined between the direction of the incoming airflow F0' and the central axis <NUM>. The swirl angle T3 represents a circumferential component of the incoming airflow F0' and may be caused by inlet flow distortion and other parameters as will be discussed below. The swirl angle F3 may affect the compressor operating line. The guide vanes <NUM> may at least partially alleviate this swirl angle T3 by redirecting the flow in the desirable direction, but might not be sufficient to totally cancel the swirl angle T3. In some cases, a sever inlet flow distortion may be sufficient to alter the compressor stability and efficiency. It has been observed that the swirl angle T3 has for effect of decreasing the exit flow angle T2' compared to the exit flow angle T2 of the configuration of <FIG> where the incoming flow F0 is devoid of a swirl angle. Thus, for the same geometric angle T1 of the guide vanes <NUM>, the exit flow angle T2 of the configuration of <FIG> where no swirl angle is present may differ from the exit flow angle T2' of the configuration of <FIG> where the incoming airflow F0' present the swirl angle T3. A relation between the geometric angle T1 and the exit flow angle T2' where the incoming flow F0' presents the swirl angle T3 may vary with many parameters.

Typical variable guide vane control systems target a given geometric angle T1, which is scheduled in function of one or more engine parameters, such as power and/or corrected compressor speed for instance. The approach consists of anticipating the maximum compressor stability margin loss associated to the highest possible encountered inlet flow distortion during the engine operation. This approach does not take into consideration the swirl angle F3 and may result in non-optimized performance of the downstream rotor (e.g., compressor rotor). Put differently, rotating the guide vanes <NUM> until their geometric angle T1 matches a target geometric angle does not factor in the effect the swirl angle T3.

Referring now to <FIG>, a control system for the variable guide vane assembly <NUM> is shown at <NUM> and may at least partially alleviate the aforementioned drawbacks.

The control system <NUM> aims at controlling the rotation of the variable guide vanes <NUM> until the exit flow angle T2' reaches a target exit flow angle. Put differently, the control system <NUM> focuses on the exit flow angle T2' rather than the geometric angle T1. The control system <NUM> takes into consideration the swirl angle T3 of the incoming flow F0' and changes the geometric angle T1 until a predicted exit flow angle reaches the target exit flow angle.

The control system <NUM> includes a controller <NUM> operatively connected to the actuator <NUM> and to a sensor <NUM> operable to send a signal to the controller <NUM>; the signal indicative of the geometric angle T1 of the guide vanes <NUM>. The controller <NUM> includes a target angle calculator <NUM> and an exit flow angle predictor <NUM>. The controller <NUM> receives operating parameters <NUM>, which includes one or more of flight parameters 55A and engine parameters 55B. The flight parameters 55A includes one or more of an altitude of the aircraft <NUM> (<FIG>) equipped with the gas turbine engine <NUM>, an airspeed of the aircraft <NUM>, a temperature of ambient air, and so on. The engine parameters 55B includes one or more of a power generated by the gas turbine engine <NUM>, a rotational speed of the gas turbine engine <NUM>, and so on. The controller <NUM> feeds the one or more of the flight parameters 55A and the engine parameters 55B to the target angle calculator <NUM> that computes a target exit flow angle as a function of one or more of the flight parameters 55A and the engine parameters 55B. The target exit flow angle is defined between a direction of the flow exiting the guide vanes <NUM> and the central axis <NUM>. The target exit flow angle may correspond to an angle at which the flow exiting the guide vanes <NUM> should meet a compressor face of a downstream rotor of the compressor section <NUM>. This angle may be computed to yield the maximum efficiency of the downstream rotor. The compressor face may be defined by leading edges of blades of rotor located downstream of the VGV assembly <NUM>.

As shown in <FIG>, the target exit flow angle may be determined from a lookup table that correlates target exit flow angle values with operating parameter values, more specifically, with one or more of flight parameter values and engine parameter values. The target angle calculator <NUM> may receive the one or more of the flight parameter values and the engine parameter values from the controller <NUM>, either inputted manually by a pilot, provided by one or more sensors, and so on. Based on these parameters, the target angle calculator <NUM> may interpolate within the lookup table of <FIG> to determine the target exit flow angle at which the flow exiting the guide vanes <NUM>, namely exiting flow passages defined between the guide vanes <NUM>, should meet the compressor face.

Referring back to <FIG>, as aforementioned, the relation between the geometric angle T1 of the guide vanes <NUM> and the exit flow angle is not straightforward. Consequently, the controller <NUM> feeds at least the geometric angle T1 to the exit flow angle predictor <NUM> that determines a predicted exit flow angle as a function of at least the geometric angle T1. The predicted exit flow angle corresponds to a prediction or an estimate of the angle at which the outgoing flow is expected to exit the guide vanes <NUM>. In the embodiment shown, the exit flow angle predictor <NUM> determines the predicted exit flow angle as a function of at least the geometric angle T1 and as a function of the operating parameters <NUM> of the gas turbine engine <NUM>. The operating parameters <NUM> may include one or more of the flight parameters 55A, the engine parameters 55B, and aircraft parameters 55C.

The aircraft parameters 55C may include, for instance, a spatial orientation of the aircraft such as a yaw angle, a pitch angle, and a roll angle. The aircraft parameters 55C may further include information regarding an inlet of the gas turbine engine <NUM>. For instance, if the gas turbine engine <NUM> is a turbofan, the swirl angle T3 is expected to be minimal. However, if the gas turbine engine <NUM> is a turboprop as illustrated in <FIG>, the air entering an inlet <NUM> is deviated from being oriented substantially radially to being oriented substantially axial as it meets the VGV assembly <NUM>. In some cases, the gas turbine engine <NUM> may be an auxiliary power unit (APU) and the inlet may be S-duct. The air then navigates through a series of elbows before reaching the VGV assembly <NUM>. A filter may be provided upstream of the VGV assembly <NUM>. A scoop may be used to drawn air from an environment into the gas turbine engine <NUM>. The aircraft parameters 55C may further includes information regarding the position of the gas turbine engine <NUM>. For instance, and as depicted in <FIG>, the gas turbine engine <NUM> may be mounted on a right wing, on a left wing, on a right side of a fuselage, on a left side of the fuselage, on a top of the fuselage, within a tail of the aircraft, and so on. All of these different configurations are expected to create distortion into the incoming flow F0' and affect the swirl angle T3. The distortion may thus be regarded as a non-asymmetry of a velocity/pressure field of the air as it meets the VGV assembly <NUM>.

The exit flow angle predictor <NUM> is able to provide a prediction of the exit flow angle T2' at which the outgoing flow F1' is expected to meet the compressor face. As illustrated in <FIG>, the exit flow angle predictor <NUM> may comprise a lookup table correlating predicted exit flow angle values with geometric angle values and operating parameter values, such as one or more of flight parameter values, engine parameter values, aircraft parameter values. The lookup table may include one or more multidimensional tables. The values of the lookup table of <FIG> may be generated by experimental testing and/or by computational fluid dynamics (CFD) simulations. For instance, for each possible permutations of the geometric angle T1 of the guide vanes <NUM>, the flight parameters 55A, the engine parameters 55B, and the aircraft parameters 55C, a simulation, either experimental or numerical, is performed to compute the resulting exit flow angle. These values are tabulated and the exit flow angle predictor <NUM> may be able to interpolate to output the predicted exit flow angle. In an alternate embodiment, the exit flow angle predictor <NUM> may comprise a linearized model using partial derivate or an embarked full CFD model that may be solved in real-time.

Referring back to <FIG> and <FIG>, in use, the engine parameters 55B and the flight parameters 55A are fed to the target angle calculator <NUM> that outputs a target exit flow angle at which the outgoing flow F1' should meet the compressor face. The exit flow angle predictor <NUM> receives the geometric angle T1, which may be supplied by the sensor <NUM>, and, in some embodiments, one or more of the flight parameters 55A, the engine parameters 55B, and the aircraft parameters 55C, and outputs the predicted exit flow angle. If the predicted exit flow angle differs from the target exit flow angle outputted by the target angle calculator <NUM>, the controller <NUM> sends a signal to the actuator <NUM> to pivot the guide vanes <NUM> about their spanwise axes S1. This process continues until the predicted exit flow angle is sufficiently close the target exit flow angle. A control gain may be supplied for operation of the actuator <NUM>.

Referring now to <FIG>, a method of operating the variable guide vane assembly of <FIG> is shown at <NUM>. The method <NUM> includes obtaining the target exit flow angle defined between the direction of a flow exiting the guide vanes <NUM> and the central axis <NUM> at <NUM>; predicting the exit flow angle as a function of at least the geometric angle T1 at <NUM>; and when a difference between the exit flow angle and the target exit flow angle is above a threshold, modulating the guide vanes <NUM> to modify the geometric angle T1 until the difference between the exit flow angle and the target exit flow angle is below the threshold at <NUM>. This threshold may be from <NUM> to +/-<NUM> degrees for instance.

In the embodiment shown, the predicting of the exit flow angle as a function of at least the geometric angle includes determining the exit flow angle as a function of at least the geometric angle and as a function of operating parameters of the aircraft engine. As aforementioned, the operating parameters include aircraft parameters including one or more of a spatial orientation of the aircraft engine, a location of the aircraft engine in relation to an aircraft equipped with the aircraft engine, and a configuration of an inlet of the aircraft engine. The operating parameters may include engine parameters including one or more of a power output of the aircraft engine and a rotational speed of a shaft of the aircraft engine. The operating parameters may include flight parameters including one or more of an altitude of the aircraft engine, an airspeed of the aircraft engine, and a temperature of air entering the aircraft engine.

In the present embodiment, the predicting of the exit flow angle includes obtaining the exit flow angle from the lookup table of <FIG>. This table correlates geometric angle values and operating parameters values with exit flow angles values.

The modulating of the guide vanes <NUM> to modify the geometric angle T1 may be an iterative process. This process may include: a) causing rotation of the guide vanes <NUM> from a first position to a second position; b) determining an updated geometric angle of the guide vanes <NUM> at the second position; c) determining, based on at least the updated geometric angle of the guide vanes, an updated exit flow angle; and d) repeating steps a) to c) if a difference between the target exit flow angle and the updated exit flow angle is greater than the threshold until the difference is less than the threshold.

As mentioned above, the obtaining of the target exit flow angle may include determining the target exit flow angle as a function of operating parameters of the aircraft engine. The operating parameters include one or more of engine parameters and flight parameters, the engine parameters including one or more of a power output of the aircraft engine and a rotational speed of a shaft of the aircraft engine, the flight parameters including one or more of an altitude of the aircraft engine, an airspeed of the aircraft engine, and a temperature of air entering the aircraft engine. The determining of the target exit flow angle may include determining the target exit flow angle from the lookup table of <FIG> that correlates operating parameters values with target exit flow angles values.

The present disclosure may provide a system and method to mitigate the negative impact of the distortion induced swirl flow angle. The proposed approach consists of including into the control loop a predictor of the flow angle at the compressor face, namely the "effective guided vane angle" or the "predicted exit flow angle". Rather than controlling a "geometric guided vane angle" as done in the traditional method, the proposed method controls the exit flow angle. For any type of possible engine inlet flow distortion encountered, the disclosed control system may allow the commanded geometric actuation angle to match the desirable effective flow swirl angle in front of the compressor face.

With reference to <FIG>, an example of a computing device <NUM> is illustrated. For simplicity only one computing device <NUM> is shown but the system may include more computing devices <NUM> operable to exchange data. The computing devices <NUM> may be the same or different types of devices. The controller <NUM> may be implemented with one or more computing devices <NUM>. Note that the controller <NUM> can be implemented as part of a full-authority digital engine controls (FADEC) or other similar device, including electronic engine control (EEC), engine control unit (ECU), electronic propeller control, propeller control unit, and the like. In some embodiments, the controller <NUM> is implemented as a Flight Data Acquisition Storage and Transmission system, such as a FAST™ system. The controller <NUM> may be implemented in part in the FAST™ system and in part in the EEC. Other embodiments may also apply.

The methods and systems for operating the variable guide vane assembly <NUM> described herein may be implemented in a high level procedural or object oriented programming or scripting language, or a combination thereof, to communicate with or assist in the operation of a computer system, for example the computing device <NUM>. Alternatively, the methods and systems for operating the variable guide vane assembly <NUM> may be implemented in assembly or machine language. The language may be a compiled or interpreted language. Program code for implementing the methods and systems for operating the variable guide vane assembly <NUM> may be stored on a storage media or a device, for example a ROM, a magnetic disk, an optical disc, a flash drive, or any other suitable storage media or device. The program code may be readable by a general or special-purpose programmable computer for configuring and operating the computer when the storage media or device is read by the computer to perform the procedures described herein. Embodiments of the methods and systems for operating the variable guide vane assembly <NUM> may also be considered to be implemented by way of a non-transitory computer-readable storage medium having a computer program stored thereon. The computer program may comprise computer-readable instructions which cause a computer, or more specifically the processing unit <NUM> of the computing device <NUM>, to operate in a specific and predefined manner to perform the functions described herein, for example those described in the method <NUM>.

Claim 1:
A method of operating a variable guide vane assembly (<NUM>) of an aircraft engine (<NUM>), the variable guide vane assembly (<NUM>) including guide vanes (<NUM>) rotatable about respective spanwise axes (S1) and circumferentially distributed about a central axis (<NUM>), the method comprising:
obtaining a target exit flow angle defined between a direction of a flow exiting the guide vanes (<NUM>) and the central axis (<NUM>);
predicting an exit flow angle (T2; T2') as a function of at least a geometric angle (T1), the exit flow angle (T2; T2') defined between the direction of the flow (F1; F1') exiting the guide vanes (<NUM>) and the central axis (<NUM>), the geometric angle (T1) defined between the guide vanes (<NUM>) and the central axis (<NUM>); and
when a difference between the exit flow angle (T2; T2') and the target exit flow angle is above a threshold, modulating the guide vanes (<NUM>) to modify the geometric angle (T1) until the difference between the exit flow angle (T2; T2') and the target exit flow angle is at or below the threshold.