Patent Description:
An airplane or other vehicle may include a propulsion system having one or more gas turbine engines for generating an amount of thrust, or for generating power to be transferred to a thrust generating device. The gas turbine engine generally includes turbomachinery. The turbomachinery, in turn, generally includes a compressor section, a combustion section, a turbine section, and an exhaust section.

During operation of the gas turbine engine, air is provided to an inlet of the compressor section where one or more axial compressors progressively compress the air until it reaches the combustion section. Fuel is mixed with the compressed air and burned within the combustion section to provide combustion gases, which are routed from the combustion section to the turbine section. The flow of combustion gasses through the turbine section drives the turbine section and is then routed through the exhaust section, e.g., to atmosphere.

<CIT> discloses prior art methods and apparatus to determine airflow conditions at an inlet of an engine.

According to a first aspect of the present invention, there is provided an optically-based measurement system as set forth in claim <NUM>. The optically-based measurement system includes an additional feature wherein the measurement controller is configured to calculate a thrust force of the gas turbine engine while the aircraft is in flight based at least in part on the calculated first mass flow and the calculated second mass flow.

The optically-based measurement system includes an additional feature wherein the measurement controller calculates a thrust force of the gas turbine engine while the aircraft is in flight based at least in part on the calculated first mass flow and the calculated second mass flow.

The optically-based measurement system includes an additional feature wherein the first imaging system comprises a first energy source configured to direct first energy at the first target area and a first sensor configured to detect a first energy spectrum at the first target area resulting from the first energy, and wherein the second imaging system comprises a second energy source configured to direct second energy at the second target area and a second sensor configured to detect a second energy spectrum at the second target area resulting from the second energy.

The optically-based measurement system includes an additional feature wherein the first energy source is coupled to a body of the aircraft and is remotely located from the gas turbine engine.

The optically-based measurement system includes an additional feature wherein the first energy source is disposed within an inlet of the gas turbine engine.

According to a further aspect of the present invention, there is provided a method of monitoring a gas turbine engine during flight of an aircraft as set forth in claim <NUM>.

The method may include additional operations comprising calculating, via the measurement controller, a thrust force of the gas turbine engine while the aircraft is in flight based at least in part on the calculated first mass flow and the calculated second mass flow.

The method may include additional operations comprising directing, via a first energy source, first energy at the first target area, sensing, via a first sensor, a first energy spectrum at the first target area resulting from the first energy, directing, via a second energy source, second energy at the second target area, and sensing, via a second sensor, a second energy spectrum at the second target area resulting from the second energy.

The method may include an additional feature, wherein the first energy source is coupled to a body of the aircraft and is remotely located from the gas turbine engine.

The method may include an additional feature, wherein the first energy source is disposed within an inlet of the gas turbine engine.

An amount of thrust provided by a gas turbine engine is typically determined according to several estimated values of the gas turbine engine rather than in-flight measured parameters. However, such a determination method may result in relatively inaccurate thrust information. Further, it may be beneficial for a control system of the gas turbine engine or vehicle to receive and/or use relatively accurate information regarding an amount of thrust in order to more appropriately control various operations of the gas turbine engine.

When quantifying the performance of gas turbine engines, there is a need to ascertain the ingested air mass flow and net thrust in flight. Altitude test chambers are available for engine thrust measurement, but are extremely expensive to maintain and operate. Current methods for estimating mass flow and net thrust rely upon extrapolations from ground-based measurements, whereas direct measurement would provide performance information useful for improving the integration of gas turbine engines with airframes.

Various non-limiting embodiments described herein provide an optically-based propulsion mass flow and thrust measurement system capable of performing a direct, non-intrusive measurement of thrust and mass flow of an installed propulsion engine of an aircraft while in flight. In one or more non-limiting embodiments, the measurement system includes one or more lasers that probe the in-flow and out-flow planes and spectrally-sensitive cameras that image the laser probe planes to obtain velocity and density measurements from the spectrum of light scattered by flow gas molecules. The scattered spectrum of light is commonly referred to as "Rayleigh scattering", "Filter Rayleigh scattering" (FRS), or "Rayleigh/Mie scattering effect", which occurs when light photons interact with local molecules or particles, respectively. The interaction between the photons and the molecules and particles produces an elastic scattering of light, which can be detected by an optical sensor.

According to one or more embodiments, the detected scattered spectrum of light can be analyzed according to optical filter spectroscopy during in flight of the aircraft. The measurement system utilizes field measurements of flow density and velocity obtained from the optical filter spectroscopy analysis to compute mass and momentum flux at planes upstream and downstream of the engine (e.g., at the front and rear of the engine) to evaluate the rigorous integral conservation equations for mass flow and thrust. Accordingly, the ability to accurately and reliably measure installed engine thrust in flight as provided by the measurement system described herein supports both engine manufacturers and airframe manufacturers in determining the delivered thrust level.

The exemplary gas turbine engine <NUM> is a two-spool turbofan engine that generally incorporates a fan section <NUM>, a compressor section <NUM>, a combustor section <NUM>, and a turbine section <NUM>. The fan section <NUM> drives air along a bypass flow path B, while the compressor section <NUM> drives air along a core flow path C for compression and communication into the combustor section <NUM>. Hot combustion gases generated in the combustor section <NUM> are expanded through the turbine section <NUM>. Although depicted as a turbofan gas turbine engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to turbofan engines and these teachings could extend to other types of engines.

The gas turbine engine <NUM> generally includes a low-speed spool <NUM> and a high-speed spool <NUM> mounted for rotation about an engine centerline longitudinal axis A. The low-speed spool <NUM> and the high-speed spool <NUM> may be mounted relative to an engine static structure <NUM> via several bearing systems <NUM>. It should be understood that other bearing systems <NUM> may alternatively or additionally be provided.

The low-speed spool <NUM> generally includes an inner shaft <NUM> that interconnects a fan <NUM>, a low-pressure compressor <NUM> and a low-pressure turbine <NUM>. The inner shaft <NUM> can be connected to the fan <NUM> through a geared architecture <NUM> to drive the fan <NUM> at a lower speed than the low-speed spool <NUM>. The high-speed spool <NUM> includes an outer shaft <NUM> that interconnects a high-pressure compressor <NUM> and a high-pressure turbine <NUM>. In this embodiment, the inner shaft <NUM> and the outer shaft <NUM> are supported at various axial locations by bearing systems <NUM> positioned within the engine static structure <NUM>.

A combustor <NUM> is arranged between the high-pressure compressor <NUM> and the high-pressure turbine <NUM>. A mid-turbine frame <NUM> may be arranged generally between the high-pressure turbine <NUM> and the low-pressure turbine <NUM>. The mid-turbine frame <NUM> can support one or more bearing systems <NUM> of the turbine section <NUM>. The mid-turbine frame <NUM> may include one or more airfoils <NUM> that extend within the core flow path C.

The inner shaft <NUM> and the outer shaft <NUM> are concentric and rotate via the bearing systems <NUM> about the engine centerline longitudinal axis A, which is co-linear with their longitudinal axes. The core airflow is compressed by the low-pressure compressor <NUM> and the high-pressure compressor <NUM>, is mixed with fuel and burned in the combustor <NUM>, and is then expanded over the high-pressure turbine <NUM> and the low-pressure turbine <NUM>. The high-pressure turbine <NUM> and the low-pressure turbine <NUM> rotationally drive the respective high-speed spool <NUM> and the low-speed spool <NUM> in response to the expansion.

Each of the compressor section <NUM> and the turbine section <NUM> may include alternating rows of rotor assemblies and vane assemblies (shown schematically) that carry airfoils that extend into the core flow path C. For example, the rotor assemblies can carry a plurality of rotating blades <NUM>, while each vane assembly can carry a plurality of vanes <NUM> that extend into the core flow path C. The blades <NUM> of the rotor assemblies add or extract energy from the core airflow that is communicated through the gas turbine engine <NUM> along the core flow path C. The vanes <NUM> of the vane assemblies direct the core airflow to the blades <NUM> to either add or extract energy.

Various components of a gas turbine engine <NUM>, including but not limited to the airfoils of the blades <NUM> and the vanes <NUM> of the compressor section <NUM> and the turbine section <NUM>, may be subjected to repetitive thermal cycling under widely ranging temperatures and pressures. The hardware of the turbine section <NUM> is particularly subjected to relatively extreme operating conditions. Therefore, some components may require internal cooling circuits for cooling the parts during engine operation. Example cooling circuits that include features such as airflow bleed ports are discussed below.

Although a specific architecture for a gas turbine engine is depicted in the disclosed non-limiting example embodiment, it should be understood that the concepts described herein are not limited to use with the shown and described configuration. For example, the teachings provided herein may be applied to other types of engines. Some such example alternative engines may include, without limitation, turbojets, turboshafts, and other turbofan configurations (e.g., wherein an intermediate spool includes an intermediate pressure compressor ("IPC") between a low-pressure compressor ("LPC") and a high-pressure compressor ("HPC"), and an intermediate pressure turbine ("IPT") between the high-pressure turbine ("HPT") and the low-pressure turbine ("LPT").

Turning now to <FIG>, an optically-based propulsion mass flow and thrust measurement system <NUM> (hereinafter referred to as "the measurement system") is illustrated according to a non-limiting embodiment of the present disclosure. The measurement system <NUM> includes a front energy source <NUM>, a forward sensor <NUM>, a rear energy source <NUM>, and a rear sensor <NUM>. The front energy source <NUM> and the forward sensor <NUM> can operate together to establish a first imaging system. Similarly, the rear energy source <NUM> and the rear sensor <NUM> can operate together to establish a second imaging system. Each of the front energy source <NUM>, forward sensor <NUM>, rear energy source <NUM>, and rear sensor <NUM> are in signal communication with a measurement controller <NUM>, which facilitates control and analysis of the measurement system <NUM> as described in greater detail below.

According to a non-limiting embodiment illustrated in <FIG>, the front energy source <NUM> includes a laser unit <NUM> that is coupled to the aircraft <NUM> and is configured to direct frontal laser energy <NUM> to a targeted first region <NUM> (e.g., a front region <NUM>) of a gas turbine engine <NUM> of an aircraft <NUM>. Although a single front energy source <NUM> is illustrated, it should be appreciated that additional front energy sources, potentially useful for improving spatial coverage of the energy sheet at the targeted front region <NUM> or improving sensitivity to velocity, temperature, or density, can be implemented without departing from the scope of the present disclosure. In addition, although a front laser unit <NUM> will be described going forward, it should be appreciated that other types of energy sources capable of directing energy that can be sensed thereat can be employed without departing from the scope of the invention.

According to another non-limiting embodiment, the front laser unit <NUM> is arranged within the inlet of the gas turbine engine <NUM> (see <FIG>). In this manner, the frontal laser energy <NUM> can impinge directly on the inner surface of the gas turbine engine <NUM>. Accordingly, the targeted front region <NUM> can be focused on the inner surface (e.g., a first control surface) of the engine <NUM> and imaging can be performed from the inner engine nacelle.

The measurement controller <NUM> outputs a control signal that drives the front laser unit <NUM>. For example, the measurement controller <NUM> can output a control signal that drives the laser unit <NUM> to output the frontal laser energy <NUM> according to a set frequency and/or wavelength. During flight of the aircraft <NUM>, the frontal laser energy <NUM> (e.g. photons) interact with particles of the airflow input to the engine <NUM> to produce an inflow Rayleigh/Mie scattering effect occurring at the targeted front region <NUM> of a gas turbine engine <NUM>.

The front sensor <NUM> is coupled to the aircraft <NUM> and has a front field of view (FOV) <NUM> that captures the targeted front region <NUM> of the gas turbine engine <NUM>. Although a single front sensor <NUM> is illustrated, it should be appreciated that additional front sensors, which may provide improved spatial coverage or sensitivity of the measurement, can be implemented without departing from the scope of the present disclosure. The front sensor <NUM> is configured to detect laser scattering of molecules caused by an inflow Rayleigh/Mie scattering effect occurring at the targeted front region <NUM> of a gas turbine engine <NUM> and produce an inflow Rayleigh scattering distribution. In one or more non-limiting embodiments, the front sensor <NUM> includes a front sensor filter (not shown) that filters the detected inflow Rayleigh scattering spectrum to define the targeted inflow spectra, also referred to as a "spectral distribution". The targeted inflow spectra can be utilized to determine input mass flow associated with the gas turbine engine <NUM>.

The rear energy source <NUM> is coupled to the aircraft <NUM> and is configured to direct rear energy <NUM> to a targeted second region <NUM> (e.g., a rear region <NUM>) of the gas turbine engine <NUM>. Although a single rear energy source <NUM> is illustrated, it should be appreciated that additional rear energy sources, potentially useful for improving spatial coverage of the energy sheet at the targeted second region <NUM> or improving sensitivity to velocity, temperature, or density, can be implemented without departing from the scope of the present disclosure. In addition, although a rear laser unit <NUM> will be described going forward, it should be appreciated that other types of energy sources capable of directing energy that can be sensed thereat can be employed without departing from the scope of the invention.

The distance from the targeted front region <NUM> to the targeted rear region <NUM> defines a "relaxation distance" (d) such that pressure variations at region <NUM> are reduced for mitigating uncertainties due to pressure contribution to thrust. Accordingly, the location of the targeted rear region <NUM> can set the relaxation distance, which in turn varies the contribution of pressure on the overall calculated thrust. In one or more non-limiting embodiments, the location of the targeted rear region <NUM> can be selected so as to minimize the contribution of pressure on the overall calculated thrust force.

With continued reference to <FIG>, the measurement controller <NUM> is configured to output a control signal that drives the rear laser unit <NUM>. For example, the measurement controller <NUM> can output a control signal that drives the rear laser unit <NUM> to output the rear energy <NUM> according to a set frequency and/or wavelength. During flight of the aircraft <NUM>, the rear laser energy <NUM> (e.g. photons) interact with particles of the exhaust output from the engine <NUM> to produce an outflow Rayleigh/Mie scattering effect occurring at the targeted rear region <NUM> of a gas turbine engine <NUM>.

The rear sensor <NUM> is coupled to the aircraft <NUM> and has a rear FOV <NUM> that captures the targeted rear region <NUM> of the gas turbine engine <NUM>. Although a single rear sensor <NUM> is illustrated, it should be appreciated that additional rear sensors can be implemented without departing from the scope of the present disclosure. The rear sensor <NUM> is configured to detect laser scattering of molecules caused by a rear Rayleigh/Mie scattering effect occurring at the targeted rear region <NUM> of a gas turbine engine <NUM> and produce an outflow Rayleigh scattering distribution. In one or more non-limiting embodiments, the rear sensor <NUM> includes a rear sensor filter (not shown) that filters the detected outflow Rayleigh scattering spectrum to define the targeted outflow spectra. The targeted outflow spectra can be utilized to determine an exhaust momentum flux with the gas turbine engine <NUM>.

The measurement controller <NUM> is configured to process the targeted inflow spectra to determine a first mass flow ingested by the engine <NUM>, and to process the targeted outflow spectra to determine a second mass flow exhausted by the engine <NUM>. Based on the first and second mass flows, the measurement controller <NUM> can generate thrust measurements during the in-flight of the aircraft <NUM>. For example, the measurement controller <NUM> can process the targeted inflow spectrum to determine an inflow temperature value (e.g., a local static fluid temperature) (T<NUM>) and an inflow density value (e.g., local static fluid density) (ρ<NUM>) at each point on the image produced according to the output of the front sensor <NUM>. The measurement controller <NUM> can further apply a Doppler shift to the targeted inflow spectrum to determine an inflow velocity magnitude value (U<NUM>) (e.g., a velocity magnitude normal to a control volume surface of the engine <NUM>) at each point on the image produced according to the front sensor <NUM> and associated with the targeted front region <NUM>. In one or more non-limiting embodiments, the measurement controller <NUM> can store one or more models indicating known temperature, density and velocities that produce a given inflow spectrum. Accordingly, the measurement controller <NUM> can process the targeted inflow spectrum by comparing it to the stored spectrum models, and then extracting the inflow temperature value (T<NUM>), the inflow density value (ρ<NUM>), and the inflow velocity magnitude value (U<NUM>) that defines a matching spectrum model.

Similarly, the measurement controller <NUM> can process the targeted outflow spectrum to determine an outflow temperature value (e.g., a local static fluid temperature) (T<NUM>) and an outflow density value (e.g., local static fluid density) (ρ<NUM>) at each point on the image produced according to the output of the rear sensor <NUM> and associated with the targeted rear region <NUM> of the gas turbine engine <NUM>. The measurement controller <NUM> can further apply a Doppler shift to the targeted output spectrum to determine an outflow velocity magnitude value (U<NUM>) (e.g., a velocity magnitude normal to a control volume surface of the engine <NUM>) at each point on the image produced according to the output of the rear sensor <NUM> associated with the targeted rear region <NUM>. As described herein, the measurement controller <NUM> can process the targeted outflow spectrum by comparing it to the stored spectrum models, and then extracting the outflow temperature value (T<NUM>), the outflow density value (ρ<NUM>), and the outflow velocity magnitude value (U<NUM>) that defines a matching spectrum model.

Based on the distribution of inflow temperature (T<NUM>) and inflow density (ρ<NUM>), the measurement controller <NUM> can calculate an inflow pressure value (P<NUM>). Similarly, the measurement controller <NUM> can calculate an outflow pressure value (P<NUM>) based on the distribution of outflow temperature (T<NUM>) and outflow density (ρ<NUM>). Both the inflow pressure value (P<NUM>) and the outflow pressure value (P<NUM>) can be calculated, for example, according to the following equation: <MAT>.

In addition, the measurement controller <NUM> can calculate a thrust force (<IMG>) of the gas turbine engine <NUM> while the aircraft <NUM> is in flight based on the inflow temperature (T<NUM>), inflow density (ρ<NUM>), and inflow velocity magnitude (U<NUM>) values along with the distribution of outflow temperature (T<NUM>), outflow density (ρ<NUM>) and outflow velocity magnitude values (U<NUM>). In one or more non-limiting embodiments, the measurement controller <NUM> first calculates an inflow integrated mass flow (ṁ<NUM>) associated with the targeted front region <NUM> and an outflow integrated mass flow (ṁ<NUM>) associated with the targeted rear region <NUM>. The inflow and outflow integrated mass flows can each be calculated based on the following equation: <MAT> where, dA is the differential area over which control surface integration occurs and "i" indicates corresponds to the ith target region or control surface (e.g., the inflow associated with the front region <NUM> or the outflow associated with the rear region <NUM>).

The thrust force (<IMG>) produced by the engine <NUM> is the difference between the momentum flux and pressure exerted on the targeted front region <NUM> (i.e., the inlet) and targeted rear region <NUM> (i.e., the outlet): <MAT>.

In one or more non-limiting embodiments, the inflow integrated mass flow (ṁ<NUM>) may be computed using Eq. <NUM> via direct measurement at the targeted front region <NUM> and standard flight instrumentation used to obtain the flight velocity (U∞) simplifying the first term in Eq. <NUM>: <MAT>.

The expression described in Eq. <NUM> is true due to conservation of mass and momentum for the stream of flow that enters the inlet. The measurement controller <NUM> can calculate the thrust force (<IMG>) of the gas turbine engine by direct application of Eq. <NUM> using the measurements at targeted front region <NUM> and targeted rear region <NUM> or with the simplified equation combining Eq. <NUM> and Eq. <NUM> which would carry reduced uncertainties due to the elimination of the pressure term for the inlet: <MAT>.

With continued reference to <FIG>, the aircraft <NUM> includes an aircraft controller <NUM> in signal communication with the measurement controller <NUM>. Although the measurement controller <NUM> is illustrated as being externally located from the aircraft controller <NUM>, it should be appreciated that the measurement controller <NUM> can be embedded in the aircraft controller <NUM> to provide a single controller. The aircraft controller <NUM> is configured to control various operations of the aircraft <NUM> and/or the gas turbine engine <NUM>. In one or more non-limiting embodiments, the measurement controller <NUM> can output the calculated thrust force (<IMG>), which the aircraft controller <NUM> can use to control the aircraft <NUM> and/or engine <NUM>.

For example, the aircraft controller <NUM> can utilize the calculated thrust force (<IMG>) as feedback information to control the gas turbine engine <NUM> and perform engine trimming operations aimed to minimize fuel burn. According to another example, the aircraft controller <NUM> can utilize the calculated thrust force (<IMG>) provided by the measurement controller <NUM> to control the engine <NUM> to reduce noise operations. The calculated thrust force (<IMG>) can also be utilized by the aircraft controller <NUM> to perform health monitoring operations. For example, the aircraft controller <NUM> can utilize the calculated thrust force (<IMG>) to detect unexpected changes in exhaust flow indicative of a possible engine fault.

As described herein, various non-limiting embodiments described herein provide an optically-based propulsion mass flow and thrust measurement system capable of performing a direct, non-intrusive measurement of thrust and mass flow of an installed propulsion engine of an aircraft while in flight. The ability to accurately and reliably measure installed engine thrust in flight as provided by the measurement system described herein supports both engine manufacturers and airframe manufacturers in determining the delivered thrust level, which optimizes engine operation compared to current methods for estimating mass flow and net thrust that rely upon extrapolations from ground-based measurements.

As used herein, the terms "about" and "substantially" are intended to include the degree of error associated with measurement of the particular quantity based upon the equipment available at the time of filing the application. For example, the terms may include a range of ± <NUM>%, or <NUM>%, or <NUM>% of a given value or other percentage change as will be appreciated by those of skill in the art for the particular measurement and/or dimensions referred to herein.

It should be appreciated that relative positional terms such as "forward," "aft," "upper," "lower," "above," "below," "radial," "axial," "circumferential," and the like are with reference to normal operational attitude and should not be considered otherwise limiting.

Claim 1:
An optically-based measurement system (<NUM>) comprising:
a first imaging system configured to be disposed adjacent an inlet of a gas turbine engine (<NUM>) of an aircraft (<NUM>), the first imaging system configured to perform a first imaging of a first target area (<NUM>) of the inlet of a gas turbine engine (<NUM>) of an aircraft (<NUM>), the first imaging performed while the gas turbine engine (<NUM>) operates while the aircraft (<NUM>) is in flight;
a second imaging system configured to be disposed adjacent an outlet of the gas turbine engine (<NUM>), the second imaging system configured to perform a second imaging of a second target area (<NUM>) of the outlet of the gas turbine engine (<NUM>) of an aircraft (<NUM>), the second imaging performed while the gas turbine engine operates while the aircraft (<NUM>) is in flight; and
a measurement controller (<NUM>) configured to calculate the first mass flow and the second mass flow of the gas turbine engine based at least in part on the first imaging of the inlet and the second imaging of the outlet, respectively.