Patent Description:
The present subject matter relates generally to controlling a fuel flow demand of a gas turbine engine.

Some gas turbine engines include, in serial flow communication, a gas generator compressor, a combustor, a gas generator turbine, and a power turbine. The combustor generates combustion gases that are channeled to the gas generator turbine where they are expanded to drive the gas generator turbine. Then, the combustion gases are channeled to the power turbine where they further expand to drive the power turbine. The gas generator turbine is coupled to the gas generator compressor via a gas generator shaft, and the power turbine is coupled to an output shaft via a power turbine shaft. The output shaft may be coupled to a load, such as a main rotor of a helicopter.

Gas turbine engines typically include an engine controller to determine an amount of fuel (e.g., fuel flow demand) the gas turbine engine requires in order to produce a desired power. In operation, the engine controller can execute control logic in order to output a fuel flow demand that can be used to control fuel flow to the engine. The desired output of the load can be achieved by controlling the fuel flow to the engine. It would be welcomed in the art to provide improved control of an engine in response to disturbances or changes in desired power from a load coupled with the engine. <CIT> discloses a fuel control system for gas turbine engines that accounts for real-time thermodynamic engine effects.

A full and enabling disclosure of the present subject matter, including the best mode thereof, directed to one of ordinary skill in the art, is set forth in the specification, which makes reference to the appended figures, in which:.

Reference will now be made in detail to present embodiments of the subject matter, one or more examples of which are illustrated in the accompanying drawings. Each example is provided by way of explanation of the subject matter, not limitation of the subject matter. In fact, it will be apparent to those skilled in the art that modifications and variations can be made in the present subject matter without departing from the scope of the claims. For instance, features illustrated or described as part of one embodiment may be used on another embodiment to yield a still further embodiment. Thus, it is intended that the present subject matter covers such modifications and variations as come within the scope of the claims.

Like or similar designations in the drawings and description have been used to refer to like or similar parts of the subject matter, and identical numerals indicate the same elements throughout the drawings. As used herein, the terms "first", "second", and "third" may be used interchangeably to distinguish one component from another and are not intended to signify location or relative importance of the individual components.

For example, the approximating language may refer to being within a <NUM>, <NUM>, <NUM>, or <NUM> percent margin.

The present disclosure is generally directed to controlling fuel flow to a gas turbine engine using both feedforward and feedback control in response to disturbances associated with a load mechanically coupled with the gas turbine engine. Particularly, control logic is provided that seeks to maintain a constant power turbine speed stably and subtly in response to relatively small disturbances associated with the load and aggressively in response to relatively large disturbances associated with the load, as well as smooth transitions between the responses.

In one example aspect, a gas turbine engine mechanically coupled with a load of an aircraft is provided. For instance, the gas turbine engine can be a turboshaft engine and the load can be a main rotor of a helicopter. The gas turbine engine can include a controller having one or more processors configured to execute various operations, including turboshaft speed control logic to maintain a constant speed of the power turbine of the gas turbine engine despite disturbances associated with the rotor. The turboshaft speed control logic includes a feedforward governing module and a feedback governing module.

The feedforward governing module includes a feedforward module that translates an aircraft input rate into a first fuel flow demand. For instance, the aircraft input rate can correspond to a rate of change of the pitch angle of the main rotor of a helicopter in response to manipulation of a collective control device. The feedforward module may not be rate limited, which may allow for quick initial acceleration. The feedback governing module includes an aggressive control module and a power turbine governor module. The power turbine governor module translates a power turbine speed error into a third fuel flow demand while the aggressive control module calculates a system error based on the power turbine speed error, a power turbine speed error rate derived from the power turbine speed error, and a bandwidth of the one or more processors executing the feedback governing module. The aggressive control module translates the system error into a second fuel flow demand. Generally, the aggressive control module is configured to apply a quick and forceful "kick" when there is a relatively large disturbance with the rotor load, e.g., a relatively large increase or decrease in demanded lift.

The first, second, and third fuel flow demands are summed, and a composite fuel flow demand is determined. The fuel flow to the gas turbine engine can be controlled based on the composite fuel flow demand. Ultimately, the feedforward and feedback governing modules collectively allow the power turbine to maintain constant speed stably in response to relatively small disturbances in rotor load and aggressively in response to relatively large disturbances in rotor load. The turboshaft control logic synthesizes these two objectives into one cohesive control scheme.

The turboshaft speed control logic disclosed herein may provide a number of technical effects, advantages, and benefits. For instance, as noted above, the feedforward module of the present disclosure may not need to be rate limited. Moreover, the feedforward module of the present disclosure utilizes the rate of change of aircraft inputs as opposed to direct inputs, which reduces the physical modeling of the rotor system to an estimation of the partial derivative, rather than an exact calculation. This may streamline processing times and may free up processing resources. Further, the feedback governing module of the present disclosure allows for stable governing in response to relatively small rotor load disturbances by way of the power turbine governor module and quickly and forcefully in response to relatively large rotor load disturbances by way of the aggressive control module. Transition to or from aggressive control can occur smoothly as the "kick" provided by the aggressive control module is not filtered through the power turbine governor module. It will be appreciated that the inventive aspects of the present disclosure may provide other benefits and advantages in addition to those expressly noted herein.

Turning now to the drawings, <FIG> provides a perspective view of an aircraft <NUM> in accordance with one example embodiment of the present disclosure. In <FIG>, the aircraft <NUM> is a rotorcraft, and more specifically, a helicopter. The aircraft <NUM> defines an orthogonal coordinate system, including three orthogonal coordinate axes. More specifically, the three orthogonal coordinate axes include a lateral axis L, a longitudinal axis T, and a vertical axis V. In operation, the aircraft <NUM> may move along or around at least one of the lateral axis L, the longitudinal axis T, and the vertical axis V.

In the embodiment illustrated in <FIG>, the aircraft <NUM> includes an airframe <NUM> defining a cockpit <NUM>. The cockpit <NUM> includes, among other things, a collective pitch input device <NUM>, a cyclic pitch input device <NUM>, a tail rotor input device <NUM>, a first throttle input device <NUM>, a second throttle input device <NUM>, and an instrument panel <NUM>. The aircraft <NUM> further includes a main rotor assembly <NUM> and a tail rotor assembly <NUM>. The main rotor assembly <NUM> includes a main rotor hub <NUM> and a plurality of main rotor blades <NUM>. As shown, each main rotor blade <NUM> extends outwardly from the main rotor hub <NUM>. The tail rotor section <NUM> includes a tail rotor hub <NUM> and a plurality of tail rotor blades <NUM>. Each tail rotor blade <NUM> extends outwardly from the tail rotor hub <NUM>.

In addition, the aircraft <NUM> includes a first gas turbine engine <NUM> and a second gas turbine engine <NUM>. The first and second gas turbine engines <NUM>, <NUM> generate and transmit power to drive rotation of the main rotor blades <NUM> and the tail rotor blades <NUM>. Rotation of the main rotor blades <NUM> generates lift for the aircraft <NUM>, while rotation of the tail rotor blades <NUM> generates sideward thrust at the tail rotor section <NUM> and counteracts torque exerted on the airframe <NUM> by the main rotor blades <NUM>.

The collective pitch input device <NUM> adjusts the pitch angle of the main rotor blades <NUM> collectively (i.e., all at the same time) to increase or decrease the amount of lift the aircraft <NUM> derives from the main rotor blades <NUM> at a given rotor speed. More specifically, manipulating the collective pitch input device <NUM> causes the aircraft <NUM> to move in one of two opposing directions along the vertical direction V, or in other instances, to maintain a hover maneuver. Manipulating the collective pitch input device <NUM> can also be used to anticipate the amount of power the first and second gas turbine engines <NUM>, <NUM> provide the main rotor assembly <NUM> to generate the desired lift of the aircraft <NUM>. The collective pitch input device <NUM> may include an input device <NUM> configured to set a reference speed for the first and second gas turbine engines <NUM>, <NUM>. In one exemplary embodiment, the input device <NUM> may be a switch configured to set the reference speed for both the first and second gas turbine engines <NUM>, <NUM>.

The cyclic pitch input device <NUM> controls movement of the aircraft <NUM> around the longitudinal axis T and around the lateral axis L. In particular, the cyclic pitch input device <NUM> adjusts an angle of the aircraft <NUM> thereby allowing the aircraft <NUM> to move forward or backwards along the longitudinal direction T or sideways in the lateral direction L. Additionally, the tail rotor input device <NUM> controls a pitch angle of the tail rotor blades <NUM>. In operation, manipulating the tail rotor input device <NUM> may cause the tail rotor section <NUM> to move along the lateral direction L, which changes the orientation of the aircraft <NUM>.

The first and second throttle input devices <NUM>, <NUM> may be moved to an on position at the start of a flight and kept in the on position for the duration of the flight. For example, the first and second throttle input devices <NUM>, <NUM> may be moved to a FLY position at the start of a flight and may remain in this position through the duration of the flight. In some instances, the first and/or second throttle input devices <NUM>, <NUM> may be moved to a different position.

Although the aircraft <NUM> is shown and described herein as having a main/tail rotor configuration, it will be appreciated that the teachings of the present disclosure can apply to other types of aircrafts and vehicles more generally (see <FIG>). For example, the aircraft <NUM> can be any aircraft or vehicle, including but not limited to coaxial rotor helicopters, tandem rotor helicopters, side-by-side rotor helicopters, twin intermeshing rotor helicopters, tilt-rotor aircrafts, an Unmanned Aerial Vehicle (UAV) of an Unmanned Aircraft System (UAS), fixed-wing aircrafts, amphibious vehicles, hovercrafts, land vehicles, other turbine driven vehicles, etc..

<FIG> provides a schematic cross-sectional view of an exemplary gas turbine engine <NUM> in accordance with one embodiment of the present disclosure. As shown in <FIG>, the gas turbine engine <NUM> defines a longitudinal or centerline axis <NUM> extending therethrough for reference. The gas turbine engine <NUM> may generally include a substantially tubular outer casing <NUM> that defines an annular inlet <NUM>. The outer casing <NUM> may be formed from a single casing or multiple casings. The outer casing <NUM> encloses, in serial flow relationship, a gas generator compressor <NUM>, a combustion section <NUM>, a turbine section <NUM>, and an exhaust section <NUM>. The gas generator compressor <NUM> includes an annular array of inlet guide vanes <NUM>, one or more sequential stages of compressor blades <NUM>, one or more sequential stages of stationary and/or variable guide vanes <NUM>, and a centrifugal compressor <NUM>. Collectively, the compressor blades <NUM>, vanes <NUM>, and the centrifugal compressor <NUM> define a compressed air path <NUM>.

The combustion section <NUM> includes a combustor defining a combustion chamber <NUM> and one or more fuel nozzles <NUM> extending into the combustion chamber <NUM>. The fuel nozzles <NUM> supply fuel to mix with compressed air entering the combustion chamber <NUM>. A mixture of fuel and compressed air combust within the combustion chamber <NUM> to form combustion gases <NUM>. As will be described below in more detail, the combustion gases <NUM> drive the turbine <NUM>, which in turn drives the gas generator compressor <NUM>.

The turbine section <NUM> includes a gas generator turbine <NUM> and a power turbine <NUM>. The gas generator turbine <NUM> includes one or more sequential stages of turbine rotor blades <NUM> and one or more sequential stages of stator vanes <NUM>. Likewise, the power turbine <NUM> includes one or more sequential stages of turbine rotor blades <NUM> and one or more sequential stages of stator vanes <NUM>. Additionally, the gas generator turbine <NUM> drives the gas generator compressor <NUM> via a gas generator shaft <NUM>, and the power turbine <NUM> drives an output shaft <NUM> via a power turbine shaft <NUM>.

More specifically, as shown in the embodiment illustrated in <FIG>, the gas generator compressor <NUM> and the gas generator turbine <NUM> are coupled to one another via the gas generator shaft <NUM>, and the power turbine <NUM> and the output shaft <NUM> are coupled to one another via the power turbine shaft <NUM>. In operation, the combustion gases <NUM> drive both the gas generator turbine <NUM> and the power turbine <NUM>. As the gas generator turbine <NUM> rotates around the centerline axis <NUM>, the gas generator compressor <NUM> and the gas generator shaft <NUM> both also rotate around the centerline axis <NUM>. Further, as the power turbine <NUM> rotates, the power turbine shaft <NUM> rotates and transfers rotational energy to the output shaft <NUM>. As an example, the gas turbine engine <NUM> may be the first and second gas turbine engines <NUM>, <NUM> of <FIG>, and the output shaft <NUM> may rotate both the main and tail rotor blades <NUM>, <NUM> of the aircraft <NUM>.

Still referring to <FIG>, the gas turbine engine <NUM> also includes a first sensor <NUM> and a second sensor <NUM>. In one embodiment, the first sensor <NUM> may be configured to sense information indicative of a rotational speed NP of the power turbine shaft <NUM>. However, in alternative embodiments, the first sensor <NUM> may be configured to sensor information indicative of a rotational speed NR of the output shaft <NUM>. The second sensor <NUM> may be configured as at least one of a pressure sensor or a temperature sensor. For example, in one exemplary embodiment, the second sensor <NUM> may be a temperature sensor configured to sense information indicative of a turbine gas temperature T<NUM> of the gas turbine engine <NUM>. Alternatively, or in addition to, the second sensor <NUM> may be a pressure sensor configured to sense information indicative of a compressor discharge pressure PS3 of the gas turbine engine <NUM>.

Referring briefly now to <FIG> and <FIG>, it should be appreciated, that in at least certain exemplary embodiments, one or both of the first and second gas turbine engines <NUM>, <NUM> of the aircraft <NUM> in <FIG> may be configured in substantially the same manner as the gas turbine engine <NUM> depicted in <FIG>. In addition, the first and second gas turbine engines <NUM>, <NUM> may be mechanically coupled to one another such that the first and second gas turbine engines <NUM>, <NUM> operate together. For example, the first and second gas turbine engines <NUM>, <NUM> may be ganged together in a gearbox by, e.g., differentials and one-way clutches (such as sprag clutches), such that they operate together.

It should be appreciated, however, that in other exemplary embodiments, the gas turbine engine of <FIG> may instead have any other suitable configuration. For example, in other exemplary embodiments, the combustion section <NUM> may include a reverse flow combustor. Additionally, in still other exemplary embodiments, the gas turbine engine <NUM> may not be configured as a dual spool machine, and instead may include a common shaft configured to couple the compressor, the turbine, and the output shaft.

As shown schematically in <FIG>, the gas turbine engine <NUM> can include a controller <NUM>. In general, the controller <NUM> may correspond to any suitable processor-based device. For instance, the controller <NUM> can include one or more processors and one or more memory devices. The one or more processors can be configured to perform a variety of computer-implemented functions (e.g., performing the operations and the like disclosed herein). As used herein, the term "processor" refers not only to integrated circuits referred to in the art as being included in a computer, but also refers to a controller, microcontroller, a microcomputer, a programmable logic controller (PLC), an application specific integrated circuit (ASIC), a Field Programmable Gate Array (FPGA), and other programmable circuits. Additionally, the one or more memory devices can include various memory element(s) including, but not limited to, computer readable medium (e.g., random access memory (RAM)), computer readable non-volatile medium (e.g., flash memory), a compact disc-read only memory (CD-ROM), a magneto-optical disk (MOD), a digital versatile disc (DVD) and/or other suitable memory elements or combinations thereof. The memory <NUM> may store computer-executable instructions that, when executed by the one or more processors, cause the one or more processors to perform operations. The controller <NUM> can be an Electronic Engine Controller (EEC) or a Digital Engine Controller (DEC), for example. The controller <NUM> can be part of a Full Authority Digital Engine Control (FADEC) system. Moreover, the controller <NUM> can be communicatively coupled via one or more wired and/or wireless connections with one or more input devices within the cockpit <NUM>, one or more controllable devices onboard the gas turbine engine <NUM>, one or more sensors, such as sensors <NUM>, <NUM>, among other devices and elements. The one or more controllable devices within the gas turbine engine <NUM>, can include, without limitation, fuel metering or control valves, fuel pumps, other fuel control units, variable geometry elements, etc. The input devices can include, without limitation, the collective pitch input device <NUM>, the cyclic pitch input device <NUM>, the tail rotor input device <NUM>, the first throttle input device <NUM>, the second throttle input device <NUM>, and the instrument panel <NUM>, among other devices.

<FIG> provides a control logic diagram in accordance with one example embodiment of the present disclosure. Particularly, the control logic diagram of <FIG> depicts turboshaft speed control logic <NUM> that, when executed by one or more processors, seeks to maintain a constant power turbine speed stably in response to small disturbances in rotor load and aggressively in response to large disturbances in rotor load. The turboshaft speed control logic <NUM> synthesizes these two objectives into one cohesive control scheme. In general, the turboshaft speed control logic <NUM> will be described with reference to the aircraft <NUM> and the gas turbine engine <NUM> described above with reference to <FIG> and <FIG>. However, in other embodiments, the turboshaft speed control logic <NUM> may be implemented or used in association with any other aircraft and/or suitable gas turbine engine.

The turboshaft speed control logic <NUM> includes a feedforward governing module <NUM> and a feedback governing module <NUM>. As depicted in <FIG>, the turboshaft speed control logic <NUM> includes a feedforward module <NUM>, an aggressive control module <NUM>, and a power turbine governor module <NUM>. The feedforward module <NUM> is a component of the feedforward governing module <NUM> while the aggressive control module <NUM> and the power turbine governor module <NUM> are components of the feedback governing module <NUM>. For this embodiment, the aggressive control module <NUM> and the power turbine governor module <NUM> are separate modules and are arranged in parallel with respect to one another in the feedback governing module <NUM>.

Regarding the feedforward governing module <NUM>, the feedforward module <NUM>, when executed, translates a rate of change of aircraft inputs using one or more physical models <NUM> of the rotor system of the aircraft <NUM> into a first fuel flow demand Ẇf<NUM>. Stated another way, the one or more processors can determine, by executing the feedforward module <NUM>, the first fuel flow demand Ẇf<NUM> based at least in part on a power demand rate <MAT> associated with a rotor system of the aircraft <NUM>, e.g., the main rotor <NUM> of the aircraft <NUM> of <FIG>. The first fuel flow demand Ẇf<NUM> output from the feedforward module <NUM> is routed to a summation block <NUM>.

Generally, the feedforward module <NUM> is tuned to ensure that first fuel flow demands output therefrom will not force the engine to accelerate or decelerate in the wrong direction. Moreover, the feedforward module <NUM> is tuned to the rotor system without regard for the engine capability, which allows for aircraft handling qualities to be specifically targeted and tuned. Further, the feedforward module <NUM> is not rate limited, which ultimately allows for quick initial acceleration or deceleration of the rotor system. In addition, using the rate of change of aircraft inputs as opposed to direct inputs may reduce the physical modeling of the rotor system to an estimation of the partial derivative, rather than an exact calculation.

The rate of change of aircraft inputs can be derived from operator manipulation of an operator-manipulated input device (positioned onboard the aircraft <NUM> or offboard the aircraft <NUM> at a remote pilot station). The operator-manipulated input device can be at least one of the collective pitch input device <NUM>, the cyclic pitch input device <NUM>, and the tail rotor input device <NUM> depicted in <FIG>, for example. Additionally or alternatively, the rate of change of aircraft inputs can be derived from an automated flight system manipulating the power demand associated with a rotor system of the aircraft <NUM>.

In one example embodiment, the operator manipulated input device may be the collective pitch input device <NUM> of <FIG>. As such, at a first timestep, in response to manipulation of the collective pitch input device <NUM>, the collective pitch input device <NUM> or other sensor onboard the aircraft <NUM> may be configured to generate a first signal, e.g., indicating an increase or decrease in the vertical lift demanded of the aircraft <NUM>. Then, at a second timestep, in response to manipulation of the collective pitch input device <NUM>, the collective pitch input device <NUM> or other sensor onboard the aircraft <NUM> may be configured to generate a second signal, e.g., indicating an increase or decrease in the vertical lift demanded of the aircraft <NUM>. Based on the first signal and the second signal, as will be appreciated, a power demand rate <MAT> associated with the main rotor <NUM>.

Regarding the feedback governing module <NUM>, the power turbine governor module <NUM>, when executed, translates a power turbine speed error Np Error into a third fuel flow demand Ẇf<NUM>. The power turbine speed error Np Error indicates a speed error between a reference speed of the power turbine <NUM> and the actual speed of the power turbine <NUM>. When executed, the power turbine governor module <NUM> can translate the power turbine speed error Np Error into the third fuel flow demand Ẇf<NUM>, e.g., using one or more models, lookup tables, a combination thereof, etc. The third fuel flow demand Ẇf<NUM> output from the power turbine governor module <NUM> is routed to the summation block <NUM>.

The aggressive control module <NUM>, when executed, utilizes the power turbine speed error Np Error as well. Specifically, when executed, the aggressive control module <NUM> calculates a system error s based at least in part on a relationship between the power turbine speed error Np Error, a power turbine speed error rate <MAT>, and a bandwidth λ of the one or more processors executing the aggressive control module <NUM>. Particularly, the system error s is defined by the following equation: <MAT> wherein s is the system error, NpError is the power turbine speed error, <MAT> is the power turbine speed error rate, and λ is a bandwidth of the one or more processors executing the aggressive control module <NUM>. When executed, the aggressive control module <NUM> translates the system error s into a second fuel flow demand Ẇf<NUM>. In this way, the second fuel flow demand Ẇf<NUM> is determined based at least in part on the determined system error s. The second fuel flow demand Ẇf<NUM> output from the aggressive control module <NUM> is routed to the summation block <NUM>.

Execution of the aggressive control module <NUM> will now be described in further detail. As depicted in <FIG>, the Np Error is input into the aggressive control module <NUM>. As noted, the power turbine speed error Np Error indicates a speed error between a reference speed of the power turbine <NUM> and the actual speed of the power turbine <NUM>. The power turbine speed error rate <MAT> can be derived from the power turbine speed error Np Error. For instance, the power turbine speed error Np Error at a first timestep and the power turbine speed error Np Error at a second timestep can be used to derive the rate of change of the power turbine speed error or power turbine speed error rate <MAT>, wherein the second timestep occurs later in time than the first timestep. The bandwidth λ of the one or more processors executing the aggressive control module <NUM> is set to attenuate drivetrain resonant frequencies and account for the bandwidth capability of the power turbine governor module <NUM>.

With the power turbine speed error Np Error, the power turbine speed error rate <MAT>, and the bandwidth λ of the one or more processors executing the aggressive control module <NUM> calculated or known, the system error s can be determined by the one or more processors executing the aggressive control module <NUM>. As shown in <FIG>, the second fuel flow demand Ẇf<NUM> can be scheduled as a function of the system error s.

As depicted, the aggressive control module <NUM> includes a dead-band ϕ that indicates a range or band of system errors in which the second fuel demand Ẇf<NUM> is determined as being zero. That is, the dead-band ϕ indicates a band of system errors that, when the system error s is determined to be within, the second fuel demand Ẇf<NUM> is determined to be at or about zero. The dead-band ϕ ranges between a first bound -ϕ and a second bound +ϕ as depicted in <FIG>. Notably, when the system error s is determined to be within the dead-band ϕ, the second fuel demand Ẇf<NUM> is scheduled as being zero. In this regard, the aggressive control module <NUM> is not active when the system error s is determined as being within the dead-band ϕ. In contrast, when the system error s is determined to be not within the dead-band ϕ, the second fuel demand Ẇf<NUM> is determined as being not zero. Accordingly, the aggressive control module <NUM> is active when the system error s is determined as being not within the dead-band ϕ.

The first bound -ϕ and a second bound +ϕ of the dead-band ϕ can be tuned so that the aggressive control module <NUM> is active only when there are large disturbances with the rotor system. For instance, system errors that are left of the first bound -ϕ and system errors that are right of the second bound +ϕ correspond to relatively large disturbances while system errors within the dead-band ϕ (or that are both right of the first bound -ϕ and left of the second bound +ϕ) correspond to relatively small disturbances associated with the rotor system.

For system errors left of the first bound -ϕ, the scheduled second fuel demand Ẇf<NUM> corresponds to an increased demand in fuel to account for the relatively large disturbance associated with the rotor system. For instance, as one example, additional and significant lift may be demanded so that the aircraft <NUM> may perform a vertical climb. Thus, the rotor system is subjected to a relatively large disturbance. To account for the demanded additional and significant lift, the aggressive control module <NUM> reacts quickly and forcefully to apply a "kick" to rapidly increase the fuel flow to the engine <NUM>. Accordingly, the second fuel demand Ẇf<NUM> output by the aggressive control module <NUM> corresponds to a demanded fuel flow increase.

For system errors right of the second bound +ϕ, the scheduled second fuel demand Ẇf<NUM> corresponds to a decreased demand in fuel to account for the relatively large disturbance associated with the rotor system. For instance, as one example, significantly reduced lift may be demanded. Thus, the rotor system is subjected to a relatively large disturbance. To account for the demanded reduced lift, the aggressive control module <NUM> reacts quickly and forcefully to apply a "kick" to rapidly decrease the fuel flow to the engine <NUM>. Accordingly, the second fuel demand Ẇf<NUM> output by the aggressive control module <NUM> corresponds to a demanded fuel flow decrease.

As the aggressive control module <NUM> is only activated when there is relatively large system error, the aggressive control module <NUM> does not need to be held to the same stability requirements as the power turbine governor module <NUM> and does not need to attenuate modes in the rotor system. Therefore, as noted, the aggressive control module <NUM> can react much more quickly and forcefully than gain kickers of traditional power turbine governors.

Further, the dead-band ϕ, or rather the first and second bounds -ϕ, +ϕ thereof, can be tuned to manage the transition between the stable power turbine governor module <NUM> and the aggressive control module <NUM> for optimal or otherwise improved system response. This may allow for the outputs of the feedback governing module <NUM> to smoothly transition between governing with the power turbine governor module <NUM> when the system error s is relatively small and with kick using the aggressive control module <NUM> in addition to the power turbine governor module <NUM> when the system error s is relatively large. The dead-band ϕ can be tuned automatically, e.g., by an autotuning loop, or can be tuned manually. Generally, the aggressive control module <NUM> will be less active the larger the dead-band ϕ, and conversely, the aggressive control module <NUM> will be more active the smaller the dead-band ϕ.

As depicted in <FIG>, a schedule associated with the third fuel flow demand Ẇf<NUM> is shown as a function of the system error s to illustrate the difference in how the feedback governing module <NUM> reacts to small disturbances compared to large disturbances. Notably, the slope of the schedule associated with the third fuel flow demand Ẇf<NUM> is smaller or less steep than the slope of the non-dead-band portions of the schedule of associated with the second fuel flow demand Ẇf<NUM>. In this regard, when the system error s is relatively small, the power turbine governor module <NUM> reacts in a relatively conservative manner to increase or decrease fuel flow and the aggressive control module <NUM> is inactive. Indeed, a feedback schedule corresponding to a composite feedback fuel flow demand Ẇf<NUM>-<NUM> traces directly along the schedule associated with the third fuel flow demand Ẇf<NUM> when the system error s is within the dead-band ϕ. The feedback schedule associated with the composite feedback fuel flow demand Ẇf<NUM>-<NUM> represents a combination of the schedule associated with the second fuel flow demand Ẇf<NUM> and the third fuel flow demand Ẇf<NUM>. When the system error s transitions from a small disturbance to a large disturbance, e.g., by the system error s being determined to be not within the dead-band ϕ, the feedback schedule associated with the composite feedback fuel flow demand Ẇf<NUM>-<NUM> transitions relatively smoothly to incorporating the "kick" from the aggressive control module <NUM> compared to gain kickers of traditional power turbine governor modules. The relative gain of the aggressive control module <NUM> is placed to smoothly transition between the aggressive control and non-aggressive control.

Referring again to <FIG>, the one or more processors can determine a composite fuel flow demand Ẇf. Particularly, the one or more processors can determine a composite fuel flow demand Ẇf based at least in part on the first fuel flow demand Ẇf<NUM>, the second fuel flow demand Ẇf<NUM>, and the third fuel flow demand Ẇf<NUM>. To determine the composite fuel flow demand Ẇf, the one or more processors can execute the summation block <NUM> to sum the first fuel flow demand Ẇf<NUM>, the second fuel flow demand Ẇf<NUM>, and the third fuel flow demand Ẇf<NUM>. In this way, the composite fuel flow demand Ẇf can be a summation of the first fuel flow demand Ẇf<NUM>, the second fuel flow demand Ẇf<NUM>, and the third fuel flow demand Ẇf<NUM>. As will be appreciated from the teachings above, when the system error s is within the dead-band ϕ, the second fuel flow demand Ẇf<NUM> equates to zero. When the system error s is not within the dead-band ϕ, the second fuel flow demand Ẇf<NUM> does not equate to zero, and consequently, the aggressive control module <NUM> provides a "kick" to quickly and forcefully respond to large disturbances associated with the rotor system. As will be appreciated, the one or more processors can control a fuel flow to the gas turbine engine <NUM> based at least in part on the composite fuel flow demand Ẇf, e.g., by controlling one or more controllable devices that, when actuated, cause more or less fuel to be provided to the gas turbine engine <NUM> or a combustor thereof.

It will be appreciated that the turboshaft speed control logic <NUM> depicted in <FIG> can be constructed in varying ways yet may still provide the advantages and benefits disclosed herein. For instance, <FIG> depicts an alternative construction of the turboshaft speed control logic <NUM>. In <FIG>, the second fuel flow demand Ẇf<NUM> from the aggressive control module <NUM> and the third fuel flow demand Ẇf<NUM> from the power turbine governor <NUM> may be summed at summation block <NUM> to render a feedback fuel flow demand Ẇf<NUM>-<NUM>. The feedback fuel flow demand Ẇf<NUM>-<NUM> may then be routed to summation block <NUM> where the feedback fuel flow demand Ẇf<NUM>-<NUM> is summed with the first fuel flow demand Ẇf<NUM> from the feedforward module <NUM>. It will be appreciated that <FIG> depicts one example alternative to the construction of the turboshaft speed control logic <NUM> of <FIG> and that other alternatives are possible.

<FIG> provides a flow diagram of an example method <NUM> of controlling a fuel flow to a gas turbine engine in response to disturbances associated with a rotor mechanically coupled thereto. The method <NUM> of <FIG> can be implemented using, for instance, the controller <NUM> and other components described herein. In some implementations, the gas turbine engine can be the turboshaft gas turbine engine <NUM> of <FIG> and the rotor can be the main rotor <NUM> of the aerial vehicle of <FIG>. <FIG> depicts actions performed in a particular order for purposes of illustration and discussion. Those of ordinary skill in the art, using the disclosure provided herein, will understand that various actions of the method <NUM> can be modified in various ways without deviating from the scope of the present disclosure.

At <NUM>, the method <NUM> includes determining, by one or more processors executing a feedforward module, a first fuel flow demand based at least in part on a power demand rate associated with a rotor of an aircraft, the rotor being mechanically coupled with a gas turbine engine, the gas turbine engine having a power turbine. For instance, the power demand rate can be a rate of change in the pitch associated with blades of a main rotor of the aircraft. The rate of change can be determined in response to manipulation of a collective pitch input device at a first timestep to a second timestep, for example.

At <NUM>, the method <NUM> includes determining, by the one or more processors executing an aggressive control module, a second fuel flow demand based at least in part on a power turbine speed error associated with the power turbine and a power turbine speed error rate derived from the power turbine speed error. In some implementations, determining at <NUM> includes calculating, by the one or more processors, a system error associated with the rotor based at least in part on a relationship between the power turbine speed error, the power turbine speed error rate, and a bandwidth of the one or more processors executing the aggressive control module. In such implementations, the second fuel flow demand is determined based at least in part on the system error. For instance, the second fuel flow demand can be scheduled as a function of system error, e.g., as shown in <FIG>. The second fuel flow demand can be scheduled as a function of the system error so that the "kick" provided by the aggressive control module can be proportional to the size of the disturbance. In this way, very large kicks can be applied to very large disturbances while less large kicks can be applied to less large disturbances. As noted, for small disturbances, the aggressive control module may not kick at all, e.g., due to a scheduled dead-band, thus allowing the feedback governing control to be handled by the power turbine governor module. The system error can be defined according to Equation <NUM> disclosed herein.

Further, in some implementations, the aggressive control module includes a dead-band that indicates a band of system errors in which, when the system error is within the dead-band, the second fuel flow demand is determined as being zero. When the system error is not within the dead-band, the second fuel flow demand is determined as being not zero. The dead-band can be bound by a first bound (e.g., a negative system bound) and a second bound (e.g., a positive system bound). The first and second bounds can be dynamically tuned or adjusted, e.g., to manage the transition between the stable response provided by the power turbine governor and the aggressive control module for optimal system response.

At <NUM>, the method <NUM> includes determining, by the one or more processors executing a power turbine governor module, a third fuel flow demand based at least in part on the power turbine speed error. For instance, the power turbine speed error can be directly translated into a third fuel flow demand, e.g., by using one or more models or lookup tables. The aggressive control module and the power turbine governor module can be modules of a feedback governing module and can be arranged in parallel with respect to one another.

At <NUM>, the method <NUM> includes determining, by the one or more processors, a composite fuel flow demand based at least in part on the first fuel flow demand, the second fuel flow demand, and the third fuel flow demand. For instance, in some implementations, determining the composite fuel flow demand at <NUM> can include summing, by the one or more processors, the first fuel flow demand, the second fuel flow demand, and the third fuel flow demand. The three fuel flow demands can be summed at a single summation block, e.g., at summation block <NUM> as shown in <FIG>. Accordingly, the composite fuel flow demand is a summation of the first fuel flow demand, the second fuel flow demand, and the third fuel flow demand.

In other implementations, determining the composite fuel flow demand at <NUM> can include summing, by the one or more processors, the second fuel flow demand and the third fuel flow demand to render a feedback fuel flow demand, e.g., as summation block <NUM> depicted in <FIG>. In such implementations, the method <NUM> can further include summing, by the one or more processors, the first fuel flow demand and the feedback fuel flow demand, e.g., as summation block <NUM> depicted in <FIG>.

At <NUM>, the method <NUM> includes controlling, by the one or more processors, a fuel flow to the gas turbine engine based at least in part on the composite fuel flow demand. For instance, based on the composite fuel flow demand, one or more controllable devices, e.g., a fuel metering valve, can be actuated or controlled to allow more or less fuel to the engine. In this way, the engine can better maintain constant speed despite disturbances in rotor load.

<FIG> provides a block diagram of an example computing system <NUM>. The computing system <NUM> can be used to implement the aspects disclosed herein. The computing system <NUM> can include one or more computing device(s) <NUM>. The controller <NUM> disclosed herein can be constructed and may operate in a same or similar manner as one of the computing devices <NUM>, for example.

As shown in <FIG>, the one or more computing device(s) <NUM> can each include one or more processor(s) <NUM> and one or more memory device(s) <NUM>. The one or more processor(s) <NUM> can include any suitable processing device, such as a microprocessor, microcontroller, integrated circuit, logic device, or other suitable processing device. The one or more memory device(s) <NUM> can include one or more computer-readable media, including, but not limited to, non-transitory computer-readable media, RAM, ROM, hard drives, flash drives, and other memory devices, such as one or more buffer devices.

The one or more memory device(s) <NUM> can store information accessible by the one or more processor(s) <NUM>, including computer-readable instructions <NUM> that can be executed by the one or more processor(s) <NUM>. The instructions <NUM> can be any set of instructions or control logic that when executed by the one or more processor(s) <NUM>, cause the one or more processor(s) <NUM> to perform operations. The instructions <NUM> can be software written in any suitable programming language or can be implemented in hardware. In some embodiments, the instructions <NUM> can be executed by the one or more processor(s) <NUM> to cause the one or more processor(s) <NUM> to perform operations.

The memory device(s) <NUM> can further store data <NUM> that can be accessed by the processor(s) <NUM>. For example, the data <NUM> can include sensor data such as engine parameters, model data, logic data, etc., as described herein, aircraft inputs, power demand rates, etc. The data <NUM> can include one or more table(s), function(s), algorithm(s), model(s), equation(s), etc. according to example embodiments of the present disclosure.

The one or more computing device(s) <NUM> can also include a communication interface <NUM> used to communicate, for example, with the other components of the aircraft. The communication interface <NUM> can include any suitable components for interfacing with one or more network(s), including for example, transmitters, receivers, ports, controllers, antennas, or other suitable components.

<FIG> provides example vehicles <NUM> according to example embodiments of the present disclosure. The inventive aspects of the present disclosure can be implemented on an aircraft, such as a helicopter or fixed-wing aircraft, automobile, boat, submarine, train, unmanned aerial vehicle or drone and/or on any other suitable vehicle. While the present disclosure is described herein with reference to an aircraft implementation, this is intended only to serve as an example and not to be limiting. One of ordinary skill in the art would understand that the inventive aspects of the present disclosure can be implemented on other vehicles without deviating from the scope of the present disclosure. Further, the inventive aspects can be implemented for non-vehicle applications. For instance, the inventive aspects can be applied to nuclear power applications, e.g., emergency diesel generators for nuclear reactors, and turbine power generation.

It will be appreciated that the inherent flexibility of computer-based systems allows for a great variety of possible configurations, combinations, and divisions of tasks and functionality between and among components.

Claim 1:
A gas turbine engine (<NUM>), comprising:
a power turbine (<NUM>) mechanically coupled with a load (<NUM>); and
one or more processors (<NUM>) configured to:
determine a first fuel flow demand (Ẇf<NUM>) based at least in part on a power demand rate ( <MAT>) associated with the load (<NUM>);
determine a second fuel flow demand (Ẇf<NUM>) based at least in part on a power turbine speed error (Np Error) associated with the power turbine (<NUM>) and a power turbine speed error rate ( <MAT>) derived from the power turbine speed error (Np Error);
determine a third fuel flow demand (Ẇf<NUM>) based at least in part on the power turbine speed error (Np Error);
determine a composite fuel flow demand (Ẇf) based at least in part on the first fuel flow demand (Ẇf<NUM>), the second fuel flow demand (Ẇf<NUM>), and the third fuel flow demand (Ẇf<NUM>); and
control a fuel flow to the gas turbine engine (<NUM>) based at least in part on the composite fuel flow demand (Ẇf).