Patent Description:
The amount of thrust produced by a engine generally varies based upon the performance and environmental demands placed on the engine as the engine operates under different conditions. It is therefore desirable to be able to generate an approximate value of the thrust from an engine in order to aid in the understanding of current engine operating conditions and provide predictability for performance demand requirements. While existing techniques for estimating thrust from an engine are suitable for their purposes, improvements are desired.

<CIT> relates to a gas turbine engine for aviation use that is controlled by a digital computer supplied with engine condition and pilot command signals to vary blower blade pitch nozzle cross-section. <CIT> relates to determining operating parameters for controlling gas turbine engines.

According to a first aspect of the invention, there is provided a method as claimed in claim <NUM>.

According to another aspect of the invention, there is provided a system as claimed in claim <NUM>.

Embodiments of the invention are as claimed in the dependent claims thereof.

It will be noticed that throughout the appended drawings, like features are identified by like reference numerals.

<FIG> illustrates an example gas turbine engine <NUM> to which the systems and methods described herein may be applied. In the illustrated embodiment, the engine <NUM> is a turbofan engine that generally comprises, in serial flow communication, a fan <NUM> through which ambient air is propelled toward an inlet <NUM>, 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 rotors of the compressor section <NUM> (referred to as "HP compressor rotor(s) <NUM>") through a high pressure shaft <NUM> that rotates about axis 'A'. Axis 'A' defines an axial direction of the engine <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 about axis 'A'.

In one embodiment, the engine <NUM> includes a gas generator case <NUM> which surrounds and contains the combustor <NUM>. The gas generator case <NUM> generally includes inner and outer portions (not shown), the outer portion of the case <NUM> defining an outer wall of a combustor cavity <NUM> containing the combustor <NUM>. 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.

Control of the operation of the engine <NUM> can be effected by one or more control systems, for example an engine controller <NUM>, which is communicatively coupled to the engine <NUM>. The engine controller <NUM> can adjust a fuel flow provided to the engine <NUM>, the position and orientation of variable geometry mechanisms within the engine <NUM>, a bleed level of the engine <NUM>, and the like, based on predetermined schedules or algorithms. In some embodiments, the engine controller <NUM> may be implemented as part of one or more full-authority digital engine controls (FADECs) or other similar device(s), including electronic engine controller(s) (EEC(s)), engine control unit(s) (ECU(s)), or the like, that are programmed to control the operation of 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 engine 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>.

With additional reference to <FIG>, the engine <NUM> is illustrated schematically as having multiple elements forming a gas path along which gas flows from the inlet <NUM> to the exhaust <NUM> of the engine <NUM>. The engine <NUM> illustrated in <FIG> includes two spools, namely a low-pressure spool <NUM>, and a high-pressure spool <NUM>. The low-pressure spool <NUM> includes a low-pressure compressor stage <NUM><NUM>, which includes the LP compressor rotor(s) <NUM>, and a low-pressure turbine <NUM><NUM>, which includes the LP turbine rotors(s) <NUM>. It should however be understood that the engine <NUM> may include more than two spools (e.g., three spools). In other embodiments of the engine <NUM>, the low-pressure spool <NUM> can include more than one compressor stage. In the illustrated embodiment, the high-pressure spool <NUM> includes two high-pressure compressor stages <NUM><NUM> and <NUM><NUM> which include the HP compressor rotor(s) <NUM>, and a high-pressure turbine <NUM><NUM>, which includes the HP turbine rotor(s) <NUM>. In other embodiments of the engine <NUM>, the high-pressure spool <NUM> can include only one compressor stage, or more than two (e.g., three) compressor stages. In the illustrated embodiment, an inter-compressor case (ICC) <NUM> is disposed between the low-pressure compressor stage <NUM><NUM> and the high-pressure compressor stage <NUM><NUM>.

In one embodiment, specific locations of the engine <NUM> may be identified using station numbering. While station numbering is described herein with reference to the engine schematic diagram illustrated in <FIG>, this is for example purposes only. The station numbering may be applied to other types of engines than the engine of <FIG>, and station numbering of one or more standards or industry conventions may apply. Free stream conditions are identified as station number <NUM>, where free stream refers to the air upstream of the engine <NUM>. Engine intake front flange or leading edge is identified as station number <NUM>. A first (e.g., LP) compressor entry is identified as station number <NUM>. An intermediate (e.g., HP) compressor entry is identified as station number <NUM>. A last (e.g., HP) compressor exit (referred to herein as an exit of the compressor section <NUM> or "compressor exit") is identified as station number <NUM>. A combustor entry (or inlet) is identified as station number <NUM>. A combustor exit is identified as station number <NUM>. The last turbine exit is identified as station number <NUM>. The flow conditions upstream of a mixer occur at station number <NUM>. Station number <NUM> is at an entry of an exhaust nozzle or thrust reverser, station number <NUM> is at a nozzle throat and station number <NUM> is downstream of the nozzle throat or at an exhaust nozzle exit. The station numbers <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM> and <NUM> may be referred to as fundamental station numbers. Station numbers between the fundamental station numbers may be referred to as intermediate station numbers. Intermediate station numbers may be denoted using a second digit suffixed to a fundamental station number, such as <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, <NUM>, and the like. It should be understood that intermediate station numbers may also be denoted using decimal numbers, such as <NUM>.

As working fluids, for instance a gas mixture, pass through the engine <NUM>, the working fluids undergo numerous pressure and temperature changes. Temperature and/or pressure measurements (e.g., working fluid temperature and/or working fluid pressure measurements) may be obtained for specific locations of the engine <NUM>. The temperature and/or pressure measurements may be identified based on station numbering. As illustrated in <FIG>, temperature (e.g., total or static temperature) measured at a specific location of the engine <NUM> may be denoted using a station number suffixed to the letter T. Similarly, pressure (e.g., total or static pressure) measured at a specific location of the engine <NUM> may be denoted using a station number suffixed to the letter P. Example temperature and pressure measurement locations, T0 to T8 and P0 to P8, for the flow of the gas mixture along the gas path <NUM> are illustrated in <FIG>.

T0, taken upstream of the inlet <NUM>, refers to free stream temperature, i.e. an ambient temperature of the environment surrounding the engine <NUM> and P0 refers to free stream pressure, i.e. an ambient pressure. Although illustrated as being captured upstream of the inlet <NUM>, it should be understood that the ambient temperature T0 and the ambient pressure P0 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 ambient air enters through the fan rotor <NUM>, and P1 refers to an inlet pressure.

T2 refers to a low-pressure compressor inlet temperature, taken before the LP compressor rotor(s) <NUM> of the low-pressure compressor stage <NUM><NUM>, and P2 refers to a low-pressure compressor inlet pressure. T25 refers to a high-pressure compressor temperature, taken between the ICC <NUM> and the high-pressure compressor stage <NUM><NUM> (i.e. taken at the inlet of the high pressure compressor stage <NUM><NUM>), and P25 refers to a high-pressure compressor entry pressure.

T3 refers to a high-pressure compressor delivery (or discharge) temperature, taken after (i.e. at an outlet of) the high-pressure compressor stages <NUM><NUM> and <NUM><NUM>, and P3 refers to a high-pressure compressor delivery pressure. T3 and P3 may be taken at the last compressor exit face), for instance after the HP compressor rotor(s) <NUM> for a high-pressure spool <NUM> including a single compressor stage. T31 refers to a combustor intake temperature and P31 refers to a combustor intake pressure. Measurements for T31 (or P31) can serve as a proxy for T3 (or P3) because the last compressor exit (where T3 or P3 is taken) and the entry to the combustor <NUM> (where T31 or P31 is taken) are in close proximity.

T4 refers to a combustor outlet temperature, taken before the HP turbine rotor(s) <NUM>, and after the combustor <NUM>, and P4 refers to a combustor outlet pressure. T41 refers to a temperature taken at or near an entry to the high-pressure turbine <NUM><NUM>, and P41 refers to pressure taken at the same location. Measurements for T41 (or P41) can serve as a proxy for T4 (or P4) because the exit of the combustor (where T4 or P4 is taken) and the entry to the high-pressure turbine <NUM><NUM> (where T41 or P41 is taken) are in close proximity. T45 refers to a temperature taken between the high-pressure turbine <NUM><NUM> and the low-pressure turbine <NUM><NUM>, and P45 corresponds to the pressure taken at the same location.

Located at an intermediate point between the combustor <NUM> and the high-pressure turbine <NUM><NUM> is a vane <NUM>. The vane <NUM> directs the gas mixture passing through the engine <NUM> toward the high-pressure turbine <NUM><NUM>. The geometry of the vane <NUM> defines a vane throat, which is referred to hereinafter as a high-pressure turbine (HPT) vane throat <NUM>. The HPT vane throat <NUM> is a narrowing at the exit of the combustor <NUM> formed by the vane <NUM>. For the purposes of the present disclosure, temperature values, pressure values, or other values which are said to be evaluated at the exit of the compressor section <NUM> may be evaluated at an outlet of the high-pressure compressor stages <NUM><NUM> and <NUM><NUM> (i.e. at station number <NUM>), at an inlet of the combustor <NUM> (i.e. at station number <NUM>), or at any other suitable location. Also, temperature values, pressure values, or other values which are said to be evaluated upstream of the exit of the compressor section <NUM> may be evaluated at an inlet of the engine <NUM> (i.e. at station number <NUM>), at an inlet of a first high-pressure compressor stage <NUM><NUM> (i.e. at station number <NUM>), or at any other suitable location upstream of the exit of the compressor section <NUM>.

T5 refers to the turbine outlet temperature and P5 refers to the turbine outlet pressure, taken after the LP turbine rotor(s) <NUM> of the low-pressure turbine <NUM><NUM>. T6 refers to an exhaust gas temperature and P6 refers to an exhaust gas pressure, taken between the low-pressure turbine <NUM><NUM> and the exhaust <NUM>. T8 refers to an exhaust gas temperature and P8 refers to an exhaust gas pressure, taken at the outlet of the exhaust <NUM>, taken at the same location.

It should be noted that the above description of <FIG> pertains to an embodiment of the engine <NUM> which includes multiple spools, namely the low- and high-pressure spools <NUM>, <NUM>. The present disclosure may be applied to other types of engines, including engines with only one spool, or with more than two spools, as appropriate. Additionally, it should be understood that the foregoing disclosure relating to temperatures measurable within the engine <NUM> is not exhaustive, and various physical and/or virtual sensors may be deployed within the engine <NUM> to assess other temperature values for other locations within the engine <NUM>.

Referring now to <FIG> in addition to <FIG>, an example system <NUM> for synthesizing thrust, i.e. for generating an approximate or synthesized value (referred to herein as "synthesizing") of thrust from a turbofan engine, such as the engine <NUM>, will now be described. The engine <NUM> is configured to produce thrust, which is a mechanical force that moves an aircraft (not shown) the engine <NUM> is provided on. Thrust results from unbalanced momentum and pressure forces created within the engine <NUM>. The engine controller <NUM>, which can be electrically and/or mechanically coupled to the engine <NUM> in any suitable fashion, is configured to monitor the operating parameters of the engine <NUM> and to control at least part of the operation of the engine <NUM>. As will be described further below, the controller <NUM> is configured to synthesize thrust from the engine <NUM> based on the monitored engine parameter(s), and more specifically based on pressure measured at an exit of the compressor section <NUM> (referred to herein as "compressor exit pressure") and on temperature measured upstream of the exit of the compressor section <NUM> (i.e. at a location upstream of the outlet of the last high-pressure compressor stage, e.g. compressor stage <NUM><NUM>). The synthesizes value of thrust (also referred to herein as "synthesized thrust") may aid in the understanding of current operating conditions of the engine <NUM>, as well as enable calculation of a current performance state of the engine <NUM>. The synthesized value of thrust may, in some embodiments, be used by the controller <NUM> to control operation of the engine <NUM>.

The controller <NUM> illustratively comprises an input module <NUM>, a thrust synthesizing module <NUM> communicatively coupled to the input module <NUM>, and an output module <NUM> communicatively coupled to the thrust synthesizing module <NUM>. One or more sensors <NUM> are provided on the engine <NUM> and configured to measure one or more parameters of the engine <NUM>. In some embodiments, the sensor(s) <NUM> are pre-existing sensors of the engine <NUM>. The sensor(s) <NUM> may comprise one or more of a temperature sensor, pressure sensor, altimeter, fuel flow sensor (or meter) and/or any other suitable sensor. The engine parameter(s) may include, but are not limited to, temperature of the engine <NUM>, pressure of the engine <NUM>, ambient air temperature, ambient air pressure, altitude, and fuel flow rate (Wf) to the engine <NUM>. The engine parameter(s) may be measured continuously or at predetermined time intervals, and the measurements may be recorded in memory or any suitable storage (not shown) accessible by the controller <NUM>. In some embodiments, the controller <NUM> may be configured to trigger an action to cause the sensor(s) <NUM> to acquire the engine parameter measurements.

In other embodiments, the values of the engine parameter(s) may be provided by an engine computer (e.g., the controller <NUM>) or an aircraft computer (not shown), which may be configured to synthesize one or more of the engine parameter(s). In some embodiments, the value of a first engine parameter (e.g., pressure) may be obtained directly from the sensor(s) <NUM> and the value of a second engine parameter (e.g., temperature) may be provided by the engine computer or the aircraft computer, or vice versa. Other embodiments may apply. The engine computer or the aircraft computer may be configured to calculate one or more engine parameters from an arithmetic function of one or more engine parameters. For example, the one or more engine parameters may be the summation, delta, product, quotient, exponent or other arithmetic function of multiple engine parameters.

It is proposed herein to synthesize the thrust of the engine <NUM> as a function of compressor exit pressure and of temperature measured upstream of the compressor exit, e.g., as measured using sensor(s) <NUM>. In this manner, equipment (e.g., sensor(s) <NUM>) provided in a cold section (reference <NUM> in <FIG>) of the engine <NUM> may be used to synthesize thrust, thus alleviating the need for equipment or instrumentation provided in a hot section (reference <NUM> in <FIG>) of the engine <NUM>. In this manner, by using equipment provided in the cold section of the engine <NUM>, the overall life of the engine components may be increased.

In one embodiment, the synthesized value of thrust is generated by the thrust synthesizing module <NUM> based on the high-pressure compressor delivery pressure (P3), i.e. on pressure measurement(s) acquired by sensor(s) <NUM> located at engine station number <NUM>. It should however be understood that, in other embodiments, the synthesized value of thrust may be generated based on the combustor intake pressure (P31) taken at engine station number <NUM> since the combustor intake pressure can serve as a proxy for the high-pressure compressor delivery pressure, as previously noted.

In some embodiments, the thrust synthesizing module <NUM> may alternatively be configured to synthesize the thrust of the engine <NUM> as a function of a pressure ratio computed based on the compressor exit pressure. In one embodiment, the pressure ratio across one or more compressor stages of the engine <NUM> may be used. For example, the thrust synthesizing module <NUM> may be configured to compute a ratio of the compressor exit pressure (e.g., P3 or P31) to the high-pressure compressor pressure (P25). In other words, the ratio P3/P25 or P31/P25 is computed and used to synthesize engine thrust. In another embodiment, the thrust synthesizing module <NUM> may be configured to compute a ratio of the compressor exit pressure (e.g., P3 or P31) to the engine inlet pressure (P1). In other words, the ratio P3/P1 or P31/P1 is computed and used to synthesize engine thrust.

In addition, because engine thrust is influenced by the temperature of working fluids, for instance the gas mixture, passing through the engine <NUM>, the thrust synthesizing module <NUM> is also configured to synthesize the thrust of the engine <NUM> as a function of temperature measured (e.g., using the sensor(s) <NUM>) upstream of the exit of the compressor section <NUM>. More specifically and as will be discussed further below, the thrust synthesizing module <NUM> is configured to synthesize thrust as a function of pressure weighted by temperature, by computing a product of pressure and temperature values. In one embodiment, the temperature measurement used by the thrust synthesizing module <NUM> to synthesize the thrust is the engine's inlet temperature (T1), i.e. obtained from temperature measurement(s) acquired by the sensor(s) <NUM> located at engine station number <NUM>. In other embodiments, the thrust may be synthesized based on the high-pressure compressor temperature (T25) taken at engine station number <NUM>. The temperature measurement used to synthesize thrust may therefore be acquired at an intermediate compressor stage. It should be understood that other embodiments may apply and the temperature measurement may be acquired at any other suitable location upstream of the exit of the compressor section <NUM>. For instance, in some embodiments, the thrust may be synthesized based on the low-pressure compressor inlet temperature (T2) taken at engine station <NUM>, i.e. taken before the LP turbine rotor(s) <NUM> of the low-pressure compressor stage <NUM><NUM>. For example, the product of P3 and T1 (or similarly the product of P31 and T1), the product of P3 and T25 (or similarly the product of P31 and T25) or the product of P3 and T2 (or similarly the product of P31 and T2) may be used to synthesize thrust. In addition, pressure ratios computed based on compressor exit pressure (e.g., P3/P25, P31/P25, P3/P1, P31/P1, P3/P2, or P31/P2) may also be weighted by temperatures (e.g., T1, T25 or T2) to provide the synthesized thrust.

Still referring to <FIG>, the input module <NUM> is configured to receive the engine parameters (e.g., pressure and temperature) measured by the sensor(s) <NUM>, during operation of the engine <NUM>. In some embodiments, the input module <NUM> may be configured to pre-process (e.g., filter to remove noise, using any suitable filtering means such as a digital filter or the like) the signal(s) containing the engine parameter measurements received from the sensor(s) <NUM>. The signal(s), which are optionally pre-processed, are then transmitted to the thrust synthesizing module <NUM> which uses the engine parameter measurements to synthesize thrust for the engine <NUM>. While reference is made herein to the thrust synthesizing module <NUM> using measurements from the sensor(s) <NUM> to synthesize thrust from the engine <NUM>, it should be understood that the thrust synthesizing module <NUM> may synthesize thrust based on synthesized engine parameters, as described herein above.

As will be described further below, the thrust synthesizing module <NUM> is configured to compute the synthesized thrust as a function of the product of at least a first factor and a second factor, the first factor being a function of the compressor exit pressure and the second factor being a function of the temperature measurement upstream of the compressor exit. In one embodiment, not falling under the claimed matter, the synthesized thrust is computed as follows: <MAT> where Synthesized thrust is the synthesized value of the thrust from the engine <NUM>, P is the compressor exit pressure, T is the temperature upstream of the compressor exit, k and a are constants whose values are selected to produce a synthesized value of thrust that is within a desired threshold (or tolerance) of the actual (or true) thrust of the engine <NUM>, and f is a mathematical relationship between (or a mathematical function of) T and a. In equation (<NUM>), the first factor is P and the second factor is f(T, a).

The values of k and a may vary depending on the configuration and on operating conditions of the engine <NUM> including, but not limited to, flight conditions, altitude, airspeed, installation losses, and flight regime (e.g., takeoff, cruise, climb, descent, landing, etc.). The values of k and a may be determined and refined through development testing performed on ground and/or in flight. In one embodiment, the engine <NUM> is operated under a test environment (e.g., in a production test cell provided at a testing facility) to obtain the values of the constants (k and a) used to synthesize thrust. It should however be understood that simulation and/or modeling of the engine <NUM> may also be used (e.g., through the controller <NUM>) during a testing phase of the engine <NUM> to obtain the values of the constants k and a. Once determined, the values of k and a may be stored (in any suitable format such as a map, matrix, or lookup-table) in memory or other suitable storage accessible by the controller <NUM>.

The values of k and a are selected to ensure that the synthesized thrust is below the actual thrust delivered by the engine <NUM>. In one embodiment, the values of k and a are selected to bring a difference between the synthesized thrust and the actual thrust within a predetermined threshold (or tolerance). For example, the thrust synthesizing module <NUM> may be configured to provide a synthesized value of the engine's thrust that is below the actual thrust of the engine <NUM> by the desired tolerance (e.g., <NUM>% or the like), the tolerance depending on operating conditions on the engine <NUM>. In this manner, in operation, the engine <NUM> may deliver more thrust than the synthesized thrust, which translates in the aircraft's actual performance being better than expected. In one embodiment, the value of k may range between <NUM> and <NUM> and the value of a may range between <NUM> and <NUM>. Other embodiments may apply.

Any suitable mathematical function f may apply, depending on the configuration and on operating conditions of the engine <NUM>. The mathematical function f may be selected to improve the accuracy of the synthesized value of the thrust delivered by the engine <NUM>. In one embodiment, an exponential relationship (or function) may be used as the mathematical function f. For example, an exponential relationship (i.e. an exponentiation operation) Ta (or <MAT>) between the temperature T (or the non-dimensionalized temperature <MAT>, as discussed below with reference to equation (<NUM>)) and the constant a may apply. With the thrust being synthesized as a function of pressure measured at station <NUM> (i.e. P3) and the temperature being measured at station <NUM> (i.e. T1), equation (<NUM>) therefore becomes in this example: <MAT>.

It should however be understood that any other suitable mathematical function other than the exponential function, including, but not limited to, a quadratic function and a logarithmic function, may apply.

In some embodiments, the thrust synthesizing module <NUM> may be configured to correct (i.e. non-dimensionalize or normalize) the engine parameters in order to take into account ambient conditions. In other words, although equation (<NUM>) above makes use of the raw pressure and temperature values (e.g., as obtained from the sensor(s) <NUM> or synthesized in any suitable manner) to synthesize thrust, thrust may be synthesized using non-dimensionalized (or referred) pressure and temperature values. For this purpose, the thrust synthesizing module <NUM> may be configured to divide the compressor exit pressure by a reference pressure to obtain a non-dimensionalized (or "referred") pressure, and to divide the temperature upstream of the compressor exit by a reference temperature to obtain a non-dimensionalized (or "referred") temperature. In one embodiment, the reference pressure is <NUM> psi (pounds per square inch) (<NUM> kPa) and the reference temperature is <NUM> °R on the Rankine scale (<NUM>), at standard sea level. Other embodiments may apply.

With the thrust being synthesized as a function of the non-dimensionalized pressure and the non-dimensionalized temperature, equation (<NUM>) therefore becomes: <MAT> where Pref is the reference pressure, <MAT> is the non-dimensionalized compressor exit pressure, Tref is the reference temperature, <MAT> is the non-dimensionalized temperature upstream of the compressor exit.

While equations (<NUM>), (<NUM>), and (<NUM>) above are described with reference to the compressor exit pressure, it should be understood that equations (<NUM>), (<NUM>), and (<NUM>) may also be used when pressure ratios are derived from the compressor exit pressure, as described above. In this case, equations (<NUM>) and (<NUM>) above become: <MAT> And: <MAT> where Pratio is the pressure ratio derived from the compressor exit pressure, y is the adiabatic index of working fluid flowing through the turbofan engine <NUM>, f<NUM> is a first mathematical relationship between (or a first mathematical function of) Pratio, γ, and a, and f<NUM> is a second mathematical relationship between (or a second mathematical function of) T and a. For example, Pratio may be P3/P1, P31/P1, P3/P25, or P31/P25. The adiabatic index γ is a constant that refers to the ratio of the heat capacity of the working fluid at constant pressure to heat capacity at constant volume. In some embodiments, the same mathematical functions (e.g., exponential relationship) may apply for f<NUM> and f<NUM>.

In one embodiment, each mathematical function f<NUM>, f<NUM> is an exponential relationship such that (when the pressure ratio is computed as a ratio of P3/P1 and the temperature T1 is measured): <MAT> And: <MAT>.

And similarly (when the pressure ratio is computed as a ratio of P3/P1 and the temperature T1 is measured): <MAT> And: <MAT>.

Since the compression process in a gas turbine engine, such as the engine <NUM>, is accompanied by aerodynamic and thermodynamic losses and singularities, the mathematical depiction of the compression process may be considered polytropic with the governing exponent being lower than the adiabatic index γ.

Thus, in one embodiment, when the compressor exit pressure is measured at station <NUM> (i.e. P3 is used) and the temperature upstream of the compressor exit is measured at station <NUM> (i.e. T1 is used), the synthesized thrust may be obtained as follows: <MAT>.

As another example, when the compressor exit pressure measured at station <NUM> (i.e. P3 is used) and the temperature upstream of the compressor exit is measured at station <NUM> (i.e. T25 is used), the synthesized thrust may be obtained as follows: <MAT>.

In still other embodiments, fuel flow rate (Wf) to the engine <NUM> may be used to synthesize thrust. In particular, the ratio of fuel flow rate to compressor exit pressure (referred to herein as a "Ratio Unit" or RU, where RU = Wf/P, where P is the compressor exit pressure) may be used to synthesize thrust. In this case, the compressor exit pressure (e.g., P3 or P31) may be derived from the fuel flow rate and the ratio unit, such that P is replaced by Wf/RU in the equation for computing the synthesized thrust.

For example, when the compressor exit pressure is measured at station <NUM> (i.e. P3 is used) and the temperature upstream of the compressor exit is measured at station <NUM> (i.e. T1 is used), the synthesized thrust may be obtained as follows: <MAT>.

The thrust synthesizing module <NUM> may then provide the synthesized thrust to the output module <NUM>. In some embodiments, the output module <NUM> may be configured to output the synthesized thrust to a suitable output device (e.g., a cockpit display or the like). In other embodiments, the output module <NUM> may be configured to generate, based on the synthesized thrust, one or more control signals for use in controlling the operation of the engine <NUM>. The output module <NUM> may then output the control signal(s) to the engine <NUM> (e.g., to the actuators, mechanical linkages, hydraulic systems, and the like coupled to the engine controller <NUM>), using any suitable means, for effecting control of the engine <NUM> based on the synthesized thrust.

Referring now to <FIG>, a flowchart illustrating an example method <NUM> for synthesizing thrust for a turbofan engine, such as the engine <NUM> of <FIG>, will now be described. At step <NUM>, the turbofan is operated. At step <NUM>, a compressor exit pressure of the turbofan engine and a temperature upstream of the compressor exit are determined during operation of the turbofan engine. As described above with reference to <FIG>, in one embodiment, the compressor exit pressure may be determined based on at least one pressure measurement obtained from at least one pressure sensor located at an outlet of at least one high compressor stage of the turbofan engine. In another embodiment, the compressor exit pressure may be determined based on at least one pressure measurement obtained from at least one pressure sensor located at an inlet of the combustor. In addition, in one embodiment, the temperature upstream of the engine's compressor exit used to synthesize the thrust is the engine inlet temperature. In other embodiments, the thrust may be synthesized based on the high-pressure compressor temperature.

At step <NUM>, a synthesized value of thrust from the turbofan engine is determined (as per equations (<NUM>) to (<NUM>) above) based on a product of at least a first factor and a second factor, the first factor being a function of the compressor exit pressure and the second factor being a function of the temperature upstream of the compressor exit. In some embodiments, the thrust may be synthesized based on a ratio of the compressor exit pressure to a pressure at an inlet of the at least one high compressor stage. In other embodiments, the thrust may be synthesized based on a ratio of the compressor exit pressure to a pressure at an intake of the turbofan engine. In some embodiments, the compressor exit pressure is derived based on a fuel flow rate to the turbofan engine (determined during operation of the turbofan engine) and then used to synthesize thrust. In some embodiments, the pressure is divided by a reference pressure in order to obtain a referred (also referred to as a normalized or non-dimensionalized) pressure (used to determine the first factor), and the temperature is divided by a reference temperature in order to obtain a referred (also referred to as a normalized or non-dimensionalized) temperature (used to determine the second factor). As described herein above, according to the invention the first factor is determined as a function of the pressure of working fluid at the exit of the compressor section and an adiabatic index of the turbofan engine. The synthesized thrust may then be output at step <NUM>.

Referring now to <FIG>, a graph 500a plots in solid line a curve <NUM> of the true (or actual) thrust of a turbofan engine, such as the engine <NUM>, versus fan speed corrected to engine inlet conditions (N1C2), taken at ground conditions. A curve <NUM> (illustrated in dashed lines) is indicative of the synthesized thrust from the engine versus N1C2, at ground conditions, the synthesized thrust obtained using the systems and methods described herein. <FIG> shows as a bar chart 500b the synthesized thrust from the turbofan engine compared to the actual thrust delivered by the turbofan engine, for varying temperature conditions (i.e. internal standard atmosphere (ISA) conditions (on the left) versus hot day conditions (on the right)), for different corrected fan speed values (N1C2<NUM>, N1C2<NUM>, and N1C2<NUM>). It can be seen from <FIG> and <FIG> that the synthesized thrust (e.g., curve <NUM> in <FIG>) is equal to or below the true (or actual) thrust (e.g., curve <NUM> in <FIG>), which is indicative of the ability of the system and methods described herein to accurately synthesize engine thrust. While <FIG> illustrates the actual and synthesized thrust at ground conditions, it should be understood that the systems and methods described herein may be configured to synthesize thrust for flight conditions, at various altitudes and ambient temperatures.

With reference to <FIG>, an example of a computing device <NUM> is illustrated. For simplicity only one computing device <NUM> is shown but more computing devices <NUM> operable to exchange data may be provided. The computing devices <NUM> may be the same or different types of devices. The controller (reference <NUM> of <FIG>) and/or the method (reference <NUM> of <FIG>) may be implemented with one or more computing devices <NUM>. Other embodiments may also apply.

The methods and systems 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 described herein 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 described herein 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 described herein 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 (<NUM>) for synthesizing thrust from a turbofan engine (<NUM>), the turbofan engine (<NUM>) comprising a compressor section (<NUM>), a combustor (<NUM>), and a turbine section (<NUM>) in serial fluid flow communication, the method (<NUM>) comprising:
operating the turbofan engine (<NUM>);
determining, during the operating of the turbofan engine (<NUM>), a pressure of fluid at an exit of the compressor section (<NUM>) and a temperature of fluid at a location upstream of the exit of the compressor section (<NUM>);
determining a synthesized value of thrust from the turbofan engine (<NUM>) based on a product of a first factor determined as a first mathematical function of the pressure, a first constant, and an adiabatic index of fluid within the turbofan engine (<NUM>), and a second factor determined as a second mathematical function of the temperature and the first constant; and
outputting the synthesized value of thrust from the turbofan engine (<NUM>).