Patent Description:
The output power of a gas turbine engine may be set in accordance with a reference power, and the reference power may be selected as a minimum from a thermal limit, a mechanical limit, and any other limit that may affect the power available for the engine. Fixed power schedules are associated with each limit, based on a plurality of engine and aircraft parameters such as altitude, ambient temperature, aircraft speed, and the like. While this approach to setting the output power is suitable for its purposes, improvements are desired.

<CIT> discloses a prior art system and method for operating a turbine engine. The system includes a controller configured to receive feedback parameters indicative of a temperature of the at least one hot gas path component, estimate a remaining life of a hot gas path component based on the received feedback parameters, determine a desired power output of the turbine system based on the estimate remaining life of the at least one hot gas path component and a cooling capacity of the inlet cooling system, and control operation of the turbine system to cause the turbine system to generate the desired power output.

According to aspects of the present invention, there are provided methods for operating an engine as set forth in claims <NUM>, <NUM>, <NUM> and <NUM>.

In an embodiment according to any of the previous embodiments, setting the output power of the engine in accordance with the reference power based on non-thermal limits of the engine comprises selecting a minimum between an output power based on a mechanical engine limit and an output power based on an additional engine limit.

In an embodiment according to any of the previous embodiments, monitoring the engine core temperature comprises establishing the level of deterioration of the engine based on a reference condition of the engine.

In an embodiment according to any of the previous embodiments, the engine is an aircraft engine and the reference condition of the engine is a flight take-off condition.

In an embodiment according to any of the previous embodiments, establishing the level of deterioration of the engine based on a reference condition of the engine comprises, during a take-off phase, recording a take-off value for an engine core temperature, and computing the level of deterioration of the engine by comparing the take-off value for the engine core temperature to an expected take-off value for an engine core temperature of a new engine without deterioration at a same flight condition.

In an embodiment according to any of the previous embodiments, establishing the level of deterioration of the engine based on a reference condition of the engine comprises accounting for a change in performance of the engine over time.

According to a further aspect of the present invention, there is provided a system comprising a controller configured to be communicatively couple to a gas turbine engine as set forth in claim <NUM>.

Features of the systems, devices and methods described herein may be used in various combinations, in accordance with the embodiments described herein.

The present disclosure is directed to methods and systems for setting the output power of an engine. The power management approach considers the various constraints affecting the output power of the engine along the lines of thermal and non-thermal limits. Non-thermal limits are used to set the output power of the engine in accordance with a reference power, and a thermal limit is used to concurrently monitor an engine core temperature and intervene in power management only in certain circumstances, as will be explained in more detail below.

The power management approach is applicable to various types of engines, such as gas turbine engines, hybrid engines, electric motors, and the like. In some embodiments, the power management approach as described herein is applicable to auxiliary power units. With reference to <FIG>, an example gas turbine engine <NUM> of a type preferably provided for use in subsonic flight is illustrated, generally comprising in serial flow communication a fan <NUM> through which ambient air is propelled, 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, which exit via an exhaust <NUM>. High-pressure rotor(s) of the turbine section <NUM> (referred to as "HP turbine rotor(s) <NUM>") are drivingly engaged to high-pressure rotor(s) of the compressor section <NUM> (referred to as "HP compressor rotor(s) <NUM>") through a high-pressure shaft <NUM>. The turbine section <NUM> includes a vane <NUM> between the combustor <NUM> and the HP turbine rotor(s) <NUM>. Low-pressure rotor(s) of the turbine section <NUM> (referred to as "LP turbine rotor(s) <NUM>") are drivingly engaged to the fan rotor <NUM> and to low-pressure rotor(s) of the compressor section <NUM> (referred to as "LP compressor rotor(s) <NUM>") through a low-pressure shaft <NUM> extending within the high-pressure shaft <NUM> and rotating independently therefrom.

Although illustrated as a turbofan engine, the gas turbine engine <NUM> may alternatively be another type of engine, for example 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. A turboprop engine may also apply. In addition, although the engine <NUM> is described herein for flight applications, it should be understood that other uses, such as industrial or the like, may apply. According to the illustrated example, the engine <NUM> is provided in the form of a multi-spool engine having a high-pressure spool and a low pressure (LP) spool independently rotatable about axis <NUM>. However, it is understood that a multi-spool engine could have more than two spools. It should also be noted that the embodiments described herein also consider the use of single-spool engines.

Control of the operation of the engine <NUM> can be effected by one or more control systems, for example a controller <NUM>, which is communicatively coupled to the engine <NUM>. The operation of the engine <NUM> can be controlled by way of one or more actuators, mechanical linkages, hydraulic systems, and the like. The controller <NUM> can be coupled to the actuators, mechanical linkages, hydraulic systems, and the like, in any suitable fashion for effecting control of the engine <NUM>. The controller <NUM> can modulate the position and orientation of variable geometry mechanisms within the engine <NUM>, the bleed level of the engine <NUM>, and fuel flow, based on predetermined schedules or algorithms. In some embodiments, the controller <NUM> includes one or more FADEC(s), electronic engine controller(s) (EEC(s)), or the like, that are programmed to control the operation of the engine <NUM>.

The controller <NUM> is configured for monitoring an engine core temperature and comparing it to a thermal limit of the engine. The thermal limit is adjusted for a level of deterioration of the engine. In this manner, the thermal limit used for power management of the engine gets corrected over the life of the engine, such that an older engine can output as much power as a new engine. In some embodiments, the thermal limit is adjusted by applying a deterioration bias to an actual thermal limit, as follows: <MAT>.

The deterioration bias is associated with a level of deterioration of the engine and pushes the adjusted thermal limit higher on older engines to account for a change in performance of the engine over time. Any known or other method of determining a deterioration level of the engine may be used, such as but not limited to testing, simulations, modeling (e.g. state variable model (SVM)), and the like. A corresponding deterioration bias for the level of deterioration as determined is applied to the actual thermal limit to obtain the adjusted thermal limit.

In some embodiments, the deterioration bias corresponds to a difference between a reference engine core temperature and an expected engine core temperature, as exemplified in the diagram of <FIG>. The reference core temperature <NUM> is an actual engine core temperature at a given reference condition, and is compared to the expected core temperature <NUM> at the same reference condition to obtain the deterioration bias <NUM>. The expected core temperature <NUM> may be selected from a plurality of possible expected core temperatures <NUM>, as a function of various parameters such as flight conditions (e.g. altitude, pressure, temperature, etc.), engine power, engine thrust, and the like. The expected core temperature <NUM> is the temperature expected for an engine without deterioration. The reference condition may be, for example at take-off, climb, cruise, idle, etc. The expected core temperature <NUM> also corresponds to a core temperature expected for the reference condition, and the possible expected core temperatures <NUM> may also vary as a function of the reference condition at which the actual core temperature is obtained.

In some embodiments, the actual core temperature is established via an engine power assurance check (EPAC). by using a state variable model, or any various other forms of trend monitoring. In one specific and non-limiting example, the reference condition used is a flight take-off condition. During the aircraft's take-off phase of the first flight of the day, engine data upon lift-off is recorded, and comprises the engine core temperature at the take-off phase. The controller <NUM> may use the aircraft and/or engine parameters at the take-off phase to select the expected core temperature of the engine and compute the deterioration bias accordingly.

In some embodiments, the value of the deterioration bias may be adjusted over time to account for various transient effects, for instance those associated with the setting of engine component clearances. Additionally or alternatively, the value of the deterioration bias may be scaled to take into account the fact that the deterioration recorded at, for instance, a take-off condition may not apply to all conditions across the flight envelope in the same way. For example, the deterioration bias at a take-off condition for an outside temperature of <NUM> degrees Celsius may be equivalent to the deterioration bias at a max take-off condition for an outside temperature of <NUM> degrees Celsius. Other scale factors may be contemplated as well.

In some embodiments, to account for variability and data scatter when recording deterioration bias values, the deterioration bias may be confirmed by calculating deterioration bias values at multiple points during the aircraft's flight. For instance, deterioration bias values may be calculated at a take-off condition, during a climbing condition, and at a cruising condition. Other conditions for calculating the deterioration bias may be contemplated as well. A rolling average of recently-calculated deterioration bias values, for instance the ten most-recent deterioration bias values, may be used to limit variations.

Referring to <FIG>, there is shown an exemplary method <NUM> for operating the engine <NUM> to set the engine's output power. Such method may be performed by the controller <NUM> in part or in whole.

At step <NUM>, an engine core temperature is monitored. In some embodiments, monitoring the engine core temperature comprises measuring the actual core temperature using one or more real or virtual sensing device. Example temperature measurement locations of the engine <NUM> are illustrated in <FIG>. T0, taken upstream of the inlet <NUM>, refers to an ambient temperature of the environment surrounding the engine <NUM>. Although illustrated here as being captured upstream of the inlet <NUM>, it should be understood that the ambient temperature T0 can be captured at any suitable location in the environment in which the engine <NUM> is operating. T1 refers to an inlet temperature, taken at the inlet <NUM> of the engine <NUM>, just as the air from the environment enters through the fan rotor <NUM>. T2 refers to a low-pressure compressor inlet temperature, taken before the LP turbine rotor(s) <NUM> of the low-pressure compressor stage <NUM><NUM>. <NUM> refers to a high-pressure compressor temperature, taken between the ICC <NUM> and the high-pressure compressor stage <NUM><NUM>. T3 refers to a high-pressure compressor delivery temperature, taken after the high-pressure compressor stages <NUM><NUM> and <NUM><NUM>, for instance after the HP compressor rotor(s) <NUM>. T4 refers to a combustor outlet temperature, taken before the HP turbine rotor(s) <NUM>, and after the combustor <NUM>. <NUM> refers to a temperature taken at or near an entry to the high-pressure turbine <NUM><NUM>. Measurements for T4. <NUM> can serve as a proxy for T4 because the exit of the combustor (where T4 is taken) and the entry to the high-pressure turbine <NUM><NUM> (where T4. <NUM> is taken) are connected to one another. <NUM> refers to a temperature taken between the high-pressure turbine <NUM><NUM> and the low-pressure turbine <NUM><NUM>. T5 refers to the turbine outlet temperature, taken after the LP turbine rotor(s) <NUM> of the low-pressure turbine <NUM><NUM>. T6 refers to an exhaust gas temperature, taken between the low-pressure turbine <NUM><NUM> and the exhaust <NUM>. T8 refers to an exhaust gas temperature, taken at the outlet of the exhaust <NUM>.

In some embodiments, the temperature used for the actual engine core temperature is the maximum temperature of the engine. The maximum temperature usually occurs at location T4 or at location T4. <NUM>, which may be difficult to measure in at least some engines due to possible instrumentation and material temperature limitations. One approach to overcoming such difficulties is deriving the temperature at location T4 based on a temperature measured downstream from location T4, where the temperature is cooler, and where instrumentation and material temperature limitations are lowered. One example includes measuring the temperature at location T4. <NUM> is sometimes referred to as an inter-turbine or indicated turbine temperature (ITT) and in this embodiment is taken between the HP turbine rotor(s) <NUM> and LP turbine rotor(s) <NUM>. A relationship between the temperatures at locations T4 to T4. <NUM>, used for deriving the temperature at location T4, can be determined during the development phase of the engine <NUM>. The relationship can be provided to the engine controller <NUM> to derive the T4 temperature as may be required for operation of the engine <NUM>.

In some embodiments, monitoring the engine core temperature comprises estimating or deriving the actual engine core temperature. For example, an iterative method for predicting T4, as described in <CIT> may be used. Other methods or algorithms for deriving or estimating the engine core temperature may be used.

Referring back to <FIG>, and as part of step <NUM>, it is determined if the engine core temperature is below an adjusted engine thermal limit. When the engine core temperature is below the adjusted thermal limit, the output power of the engine <NUM> is set in accordance with a reference power based on non-thermal limits of the engine, at step <NUM>. In various embodiments, such non-thermal limits may comprise mechanical limits, for instance based on the speed of fan <NUM> or gearbox <NUM>, or other engine <NUM> limits such as gas generator pressure and mechanical or corrected rotor speeds, among others. The reference power may be selected as a minimum from a plurality of non-thermal limits. Fixed power schedules may be associated with each non-thermal limit, based on a plurality of engine and aircraft parameters such as altitude, ambient temperature, aircraft speed, and the like.

When the engine core temperature is above the adjusted thermal limit, the output power of the engine <NUM> is set to a value lower than the reference power to reduce the engine core temperature, at step <NUM>. Therefore, instead of using a fixed thermal limit to be considered with the mechanical limit and any other limit of the engine to set engine power, engine core temperature is used as a monitoring value in setting the engine power. The power management approach also concurrently manages engine temperature, and considers engine deterioration in temperature and power management.

Step <NUM> is performed before the core temperature is above the adjusted thermal limit. For example, if a trend of increasing engine core temperature is detected, step <NUM> may be performed when the engine core temperature is near or approaches the adjusted thermal limit. Various triggers may be used to cause step <NUM> to be performed prior to reaching the adjusted thermal limit. In one embodiment, a trigger such as a given number (X) of consecutive increases in core temperature and a difference between the core temperature and the adjusted thermal limit of less than a given difference threshold (Tdiff) causes the method <NUM> to move to step <NUM> before the adjusted thermal limit is reached by the core temperature:<MAT>.

In another embodiment, the trigger is that the difference between the core temperature and the adjusted thermal limit is less than the difference threshold (Tdiff) for a duration greater than a given time threshold (Ttime):<MAT>.

In yet another embodiment the trigger is that the difference between the core temperature and the adjusted thermal limit is less than the difference threshold (Tdiff) and the last increase in temperature was for a value greater than an increase threshold (Tincr):<MAT>.

In yet another embodiment, a prediction algorithm, such as those implemented by a proportional-integral (PI) or proportional-integral-derivative (PID) controller, or a machine-learning algorithm, is used to detect a trend or predict a crossing of the adjusted thermal limit by the engine core temperature and trigger step <NUM> prior to reaching the adjusted thermal limit.

Referring to <FIG>, an exemplary control scheme <NUM> for setting the output power of the engine <NUM> is shown. As discussed above, non-thermal limits <NUM>, <NUM> are used to set the output power of the engine <NUM> in accordance with a reference power, and a thermal limit <NUM> is used to concurrently monitor an engine core temperature and intervene in power management only in certain circumstances. The thermal limit <NUM> of the engine <NUM> may take into account, for instance, the engine's current rating (e.g. max takeoff, max continuous) and power lever angle (PLA). The thermal limit is then adjusted for the level of deterioration of the engine, for instance by applying the deterioration bias to the actual thermal limit to obtain an adjusted thermal limit. In some embodiments, the deterioration bias may be determined via the method exemplified in the diagram of <FIG>, as discussed above. Other methods for determining the deterioration bias may be contemplated as well.

Non-thermal limits <NUM>, <NUM> are used to set the output power of the engine <NUM> in accordance with a reference power. At <NUM>, a flat rating, i.e. a mechanical limit of the engine <NUM>, is determined. Such mechanical limits may be based on, for instance, the rotational speed of fan <NUM> or the rotational speed of gearbox <NUM>. The flat rating may take into account, for instance, the engine's current rating, PLA and flight condition(s) of the aircraft. Similarly, at <NUM>, other engine limits, for instance such as gas generator pressure and mechanical or corrected rotor speeds, are determined. Such other engine limits may take into account, for instance, the engine's current rating, PLA and flight condition(s) of the aircraft. Fixed power schedules may be associated with each non-thermal limit, based on a plurality of engine and aircraft parameters such as altitude, ambient temperature, aircraft speed, and the like. At <NUM>, a minimum between the flat rating and the other engine limits may be selected and outputted as the reference power.

At <NUM>, steps <NUM>, <NUM> and <NUM> of method <NUM> may be implemented to set the engine's output power. Various engine sensors <NUM> positioned throughout engine <NUM> may monitor or detect a current engine core temperature. Such sensors may be real or virtual sensing devices, as discussed above. Sensors <NUM> may also monitor or detect an actual engine output power. As such, at <NUM> the engine power output is set to the reference power. The engine core temperature is concurrently monitored to ensure that it does not exceed the adjusted thermal limit. If the engine core temperature is near or exceeds the adjusted thermal limit, the output power of the engine <NUM> is set to a value lower than the reference power to reduce the engine core temperature. In various embodiments, the output power of the engine <NUM> may be set to a value lower than the reference power until the engine core temperature drops below the engine thermal limit adjusted based on the level of deterioration of the engine <NUM>.

In some embodiments, the controller <NUM> is implemented with a computing device <NUM>, an example of which is illustrated in <FIG>. For simplicity only one computing device <NUM> is shown but the controller <NUM> may include more computing devices <NUM> operable to exchange data. The computing devices <NUM> may be the same or different types of devices. 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), and the like.

The methods and systems for operating the engine <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 engine <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 engine <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 engine <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 for operating an engine (<NUM>), the method comprising:
monitoring an engine core temperature;
when the engine core temperature is below an engine thermal limit adjusted for a level of deterioration of the engine (<NUM>), setting an output power of the engine (<NUM>) in accordance with a reference power based on non-thermal limits of the engine (<NUM>); and
when a given number (X) of consecutive increases in engine core temperature is reached, and a difference between the engine core temperature and the adjusted thermal limit is less than a given difference threshold (Tdiff), setting the output power of the engine (<NUM>) to a value lower than the reference power based on non-thermal limits of the engine (<NUM>) to reduce the engine core temperature.