Gas turbine engine with variable pitch fan and variable pitch compressor geometry

A gas turbine engine includes a fan and an engine core that includes a compressor, a combustor, and a turbine. The fan and the compressor include variable pitch geometry. The gas turbine engine further includes a control system configured to adjust the variable pitch geometry of the fan and the compressor to optimize a performance characteristic of the gas turbine engine.

FIELD OF THE DISCLOSURE

The present disclosure relates generally to gas turbine engines, and more specifically to gas turbine engines having variable pitch fan and variable pitch compressor geometry.

BACKGROUND

SUMMARY

A gas turbine engine may include a variable pitch fan, an engine core, and a control system. The variable pitch fan may be mounted for rotation about a center axis. The variable pitch fan may include a plurality of variable-pitch fan blades that extend radially outward relative to the center axis along a plurality of fan-pitch axes that correspond with the plurality of variable-pitch fan blades. Each of the plurality of variable-pitch fan blades may be configured to rotate selectively about the corresponding fan-pitch axis. The engine core may include a compressor, a combustor, and a turbine. The compressor may include a plurality of variable-pitch compressor vanes that extend radially outward relative to the center axis along a plurality of compressor-pitch axes that correspond with the plurality of variable-pitch compressor vanes. Each of the plurality of variable-pitch compressor vanes may be configured to rotate selectively about the corresponding compressor-pitch axis.

The control system may be configured to rotate the plurality of variable-pitch fan blades to specific fan-pitch angles and to rotate the plurality of variable-pitch compressor vanes to specific vane-pitch angles to optimize a selected at least one of a plurality of engine performance characteristics of the gas turbine engine during operation of the gas turbine engine at a given operating condition of the gas turbine engine. The control system may be configured to select the specific fan-pitch angles from a range of fan-pitch angles that are bounded by preset operating limits of the gas turbine engine at the given operating condition and to select the specific vane-pitch angles from a range of vane-pitch angles that are bounded by the preset operating limits of the gas turbine engine at the given operating condition.

In some embodiments, the selected at least one of the plurality of engine performance characteristics may be at least one of specific fuel consumption of the gas turbine engine, thrust produced by the gas turbine engine, an inlet temperature of the turbine, noise of the gas turbine engine, and emissions of the gas turbine engine. In some embodiments, the given operating condition includes at least one of an altitude of the gas turbine engine, a flight Mach number of the gas turbine engine, throttle setting, and ambient temperature.

In some embodiments, the preset operating limits of the gas turbine engine at the given operating condition may include at least one of surge limits of the compressor, surge limits of the fan, torque limits of the turbine, and temperature limits of the engine core.

In some embodiments, the control system may be configured to rotate the plurality of variable-pitch fan blades about the plurality of fan-pitch axes independent of the plurality of variable-pitch compressor vanes. The control system may be configured to rotate the plurality of variable-pitch compressor vanes about the plurality of compressor-pitch axes independent of the plurality of variable-pitch fan blades.

In some embodiments, the turbine may include at least a high pressure stage turbine and a low pressure stage turbine that is rotatable independent of the high pressure stage turbine. The compressor may include an intermediate pressure compressor and a high pressure stage compressor that is rotatable independent of the intermediate pressure stage compressor, and the low pressure turbine drives the intermediate pressure stage compressor and the variable pitch fan.

In some embodiments, the selected at least one of the plurality of engine performance characteristics may be thrust produced by the gas turbine engine. The range of fan-pitch angles and the range of vane-pitch angles for the given operating condition of the gas turbine engine may be based on a surge limit of the compressor and a surge limit of the variable-pitch fan.

In some embodiments, the selected at least one of the plurality of engine performance characteristics being optimized may be based on a phase of a flight cycle the gas turbine engine is operating within. In some embodiments, the preset operating limits of the gas turbine engine at the given operating condition may be included in a look-up table stored in the control system.

In some embodiments, the control system may be configured to receive data indicative of the given operating condition of the gas turbine engine and the data may be indicative of another gas turbine engine being inoperable.

In some embodiments, the selected at least one of the plurality of engine performance characteristics may be a minimal turbine inlet temperature. In some embodiments, the selected at least one of the plurality of engine performance characteristics may be an emissions of the gas turbine engine.

In some embodiments, the selected at least one of the plurality of engine performance characteristics may be minimal engine noise. The given operating condition of the gas turbine engine may be one of taxi, take-off, and landing.

A method of optimizing performance of a gas turbine engine having a fan with variable pitch fan blades and a compressor with variable pitch vanes in accordance with the present disclosure may include a number of steps. The method may include providing the fan with variable pitch fan blades and the compressor with variable pitch vanes, identifying an operating condition of the gas turbine engine, identifying an engine performance characteristic of the gas turbine engine to be optimized, and varying a pitch angle of the variable pitch fan blades while independently varying a pitch angle of the variable pitch vanes to configure the variable pitch fan blades and the variable pitch vanes to optimize the engine performance characteristic for the operating condition.

In some embodiments, the method may include selecting a specific fan blade pitch angle from a range of fan blade pitch angles that are bounded by preset operating limits of the gas turbine engine at the operating condition. The method may include selecting a specific compressor vane pitch angle from a range of compressor vane pitch angles that are bounded by the preset operating limits of the gas turbine engine at the operating condition.

In some embodiments, the engine performance characteristic may be selected based on a phase of a flight cycle in which the gas turbine engine is operating. In some embodiments, the method may include rotatably coupling the fan and an intermediate pressure stage included in the compressor.

In some embodiments, the engine performance characteristic may be a specific fuel consumption of the gas turbine engine which is optimized by minimizing the specific fuel consumption and the operating condition is one of climb and cruise. In some embodiments, the engine performance characteristic may be an inlet temperature of a turbine and the operating condition is one of takeoff and climb.

In some embodiments, the engine performance characteristic may be at least one of specific fuel consumption of the gas turbine engine, thrust produced by the gas turbine engine, an inlet temperature of the turbine, noise of the gas turbine engine, and emissions of the gas turbine engine. In some embodiments, the operating condition may include one of an altitude of the gas turbine engine, flight Mach number of the gas turbine engine, throttle setting, and ambient temperature.

DETAILED DESCRIPTION OF THE DRAWINGS

Recognizing that the pitch angle of the variable pitch fan blades and variable guide vanes of the compressor may both be adjusted to optimize one or more engine performance characteristics, the present disclosure provides a control system and methods for optimizing the engine operation for improved performance by controlling fan blade pitch while simultaneously controlling compressor pitch to thereby optimize a selected particular engine performance characteristic for a given engine condition. As such, an engine10of the present disclosure may be optimized for one engine performance characteristic during one phase of a flight cycle while optimizing a different characteristic during a different phase of the flight cycle.

A gas turbine engine10in accordance with the present disclosure is shown inFIG. 1. The gas turbine engine10includes a fan12, a compressor14, a combustor16, and a turbine18. The compressor14, combustor16, and turbine18cooperate to form an engine core21of the gas turbine engine10. The fan12is driven by the turbine18and provides thrust for propelling an aircraft. The compressor14compresses and delivers air to the combustor16. It should be understood that the compressor14may include more than one stage, such as a low pressure compressor (LPC), an intermediate pressure compressor (IPC), and a high pressure compressor (HPC). The discussion of the compressor14herein is directed to the compressor14in general, but may also be applicable to a specific stage of the compressor14. The combustor16mixes fuel with the compressed air received from the compressor14and ignites the fuel. The hot, high-pressure products of the combustion reaction in the combustor16are directed into the turbine18to cause the turbine18to rotate about a central axis11of the gas turbine engine10and drive the compressor14and the fan12.

The illustrative fan12is a variable-pitch fan12that includes a plurality of fan blades28and a pitch controller22both mounted to rotate about the central axis11as shown inFIGS. 1 and 2. The fan blades28are arranged circumferentially about the central axis11and are configured to rotate about corresponding radially extending fan-blade pivot axes30to change a pitch (sometimes called an incident angle) of the fan blades28as suggested inFIGS. 2 and 3. The control system22is configured to vary and set the pitch of the fan blades28for different operating conditions during use of the gas turbine engine10. As one example, the pitch of the fan blades28may be varied to optimize fuel burn throughout a flight mission.

The illustrative fan rotor20includes a fan disk26and a plurality of fan blades28as shown inFIGS. 2 and 4. The fan blades28extend radially outward away from the fan disk26relative to the central axis11along radially extending fan-blade pivot axes30as shown inFIG. 2. Each fan blade28is configured to rotate selectively about its corresponding fan-blade pivot axis30between and including a reverse thrust position32and a closed position34to vary the pitch of the fan blade28as suggested inFIG. 3. The fan-blade pivot axis30is perpendicular to the central axis11. The fan blades28are rotatable between a plurality of positions between the reverse thrust position32and the closed position34, including, for example, a feather position36, a max takeoff (MTO) position38, and a top of climb (TOC) position40as suggested inFIG. 3. The position of the fan blades28for different flight stages and engine operating conditions, such as max takeoff and top of climb, may be selected to optimize one or more of the selected engine performance characteristics.

In the present disclosure, the compressor14includes variable pitch vanes15which are adjustable to operate the compressor14efficiently over its full speed range. The variable pitch vanes15are used to correct the angle of incidence of the air onto a stage of turbine blades19to angles which the turbine blades19tolerate without a break down of flow, stall or surge at relatively low compressor pressure ratios and compressor rotor speeds.

In illustrative embodiments, the turbine includes at least the high pressure stage turbine and the low pressure stage turbine (relative to the high pressure stage) that is rotatable independent of the high pressure stage turbine. The compressor includes the high pressure stage compressor and the intermediate pressure stage compressor that is rotatable independent of the high pressure stage compressor. In the illustrative embodiment, the low pressure stage turbine drives the intermediate pressure stage compressor and the variable pitch fan. As a result, the pitch angle of the fan directly correlates with the rotational speed of the intermediate stage compressor which may allow for more control of the design space when optimizing a parameter. Thus, the characteristics of the compressor may be directly altered by the fan blade pitch angles.

In some embodiments, the fan and the intermediate pressure stage compressor are connected with the low pressure turbine via a gear box. As a result, the fan and the intermediate pressure stage compressor rotate relative to each other at a fixed ratio of speeds. In other embodiments, the fan and the intermediate pressure stage compressor are coupled on the same shaft. In other embodiments, the fan and intermediate pressure stage compressor are independently rotatable relative to each other.

In some instances, the variable geometry of a fan12has been used to maintain a constant engine speed while varying the thrust developed by the fan12. In other instances, a variable pitch vane15arrangement in a compressor14has been used to maintain stability of the compressor14under different pressure ratios. In the present disclosure, the variable geometry of the fan12and the compressor14is coordinated by the control system22. Coordinating the operation of the compressor vanes15and the fan blades28allows the optimization of a selected one or more of various engine operating characteristics such as specific fuel consumption (SFC), thrust, turbine inlet temperature, engine noise, engine operability, or emissions.

Referring now toFIG. 4, the typical flight cycle54of an aircraft using an engine10is illustrated. The engine10experiences different operating phases. The taxi phase42has minimal thrust requirements, but noise or emissions may be minimized in the taxi phase42. In some embodiments, thrust is selected to be optimized and, thus, maximized at the take-off phase44. In some embodiments, the inlet temperature to the turbine is selected to be optimized and, thus, minimized at the take-off phase44. The density of the air during the taxi phase42and take-off phase44is relatively high, such that the compressor14may operate at a lower compression ratio.

During the phases of airborne flight, the air temperature and density change, as well as the needed thrust. During a climb phase46, the thrust must be maintained while the air temperature and density drops. At the cruise phase48, the engine10is operated at a relatively steady state with thin air and low air temperatures. In some embodiments, specific fuel consumption and emissions is minimized during the cruise phase48, while providing sufficient thrust to meet the flight speed expectations. In other embodiments, the specific fuel consumption may be selected to be optimized and, thus, minimized, during other phases such as, for example, climb or descent. During a descent phase50, the engine10experiences increasing air density as well as increasing air temperature, while experiencing a lower thrust load.

Finally, during a landing phase52, noise is minimized while the fan12pitch is varied to control thrust for landing and reverse thrust to help slow the aircraft on the ground in some embodiments. In some embodiments, the engine emissions are selected to be minimized during the landing phase52such as, for example, during an approach portion of the landing phase. The pitch angles may be selected for reverse thrust after the wheels of the aircraft touch ground. Thus, it can be seen that the normal flight cycle54that the demands placed on the engine10vary considerably. It should also be understood that a typical flight cycle54may include multiple climb phases46and descent phases50as the aircraft is moved through various altitudes for air traffic control and to avoid weather patterns or turbulent air.

The operation of the engine10is controlled according to predetermined criteria for the various phases42,44,46,48,50,52of the flight cycle54so as to meet various engine operating characteristics such as rating limits for safety, engine life maximization, or contractual limits imposed by a user of the engine10. For any given operating condition, i.e. altitude, flight Mach number, throttle setting, ambient temperature, etc., the variable geometry of the fan12and compressor14is set to optimize at least one operating characteristic, such as specific fuel consumption, for example. This optimization occurs within any required limits, such as mechanical and corrected speeds, stability limits, temperature limits, torque limits, and loading of various blades. The limits may be steady state limits or transient limits.

In some cases, the variable geometry is over-ruled by other requirements at any given operating condition, if necessary. For example, the variable geometry could also be used to maximize available thrust during a one-engine-out situation, minimize turbine inlet temperature during key high-temperature conditions, minimize engine noise during take-off and landing conditions. In one embodiment, the engine controller includes a real-time-model run with real-time simulation and optimization that responds to environmental inputs, engine operating information, and user inputs to match the geometry of the fan blades28and the compressor vanes15to optimize a predetermined objective function. The objective function is set or changed by the pilot in real time during operation, or set to specific parameters ahead of time, based on the expected flight conditions. In other embodiments, the operating characteristics of the engine10may be controlled by using tables that are referenced by the control system22under various flight conditions.

A number of charts are shown inFIGS. 5-11for different engine performance characteristics that may be optimized. In each chart, the x-axis corresponds to a pitch angle of the compressor vanes15in the intermediate pressure stage of the compressor14while the y-axis corresponds to the pitch angle of the fan blade28. The chart is bounded on the left by a low limit of the vane pitch angle and the right by a high limit of the vane pitch angle. In some embodiments, the range between the lower and upper limits of the vane angles is about 60 degrees. The chart is bounded on the top by a maximum blade angle and on the bottom by a lower blade angle. In some embodiments, the range between the lower and upper limits of the fan blade angles is about 45 degrees

The graph ofFIG. 5illustrates max thrust for the different ranges of fan blade angles and compressor vane angles. In the illustrative performance characteristic of max thrust, the ranges of angles for the fan blades and compressor vanes are bounded by a fan surge limit and an intermediate pressure compressor surge limit, respectively.

An optimal fan blade pitch angle and compressor vane pitch angle for max thrust at the given operating condition and limits of the engine10is indicated as point90which is the intersection of the optimal fan blade pitch angle88and the optimal compressor vane pitch angle86. The optimal pitch angels may be chosen from a small range of optimal angles around the point90in some embodiments. The max thrust performance characteristic may be selected, for example, at take-off or climb. The max thrust performance characteristic may be selected, for example, if the controller receives or determines data indicative of a one-engine out on the aircraft such that the gas turbine engine10needs to compensate for lost thrust from another engine on the aircraft.

The graph ofFIG. 6illustrates specific fuel consumption for the different ranges of fan blade angles and compressor vane angles. When being optimized, minimum specific fuel consumption is desired. The specific fuel consumption may be optimized during climb, cruise, and/or descent.

In the illustrative performance characteristic of specific fuel consumption, the ranges of angles for the fan blades and compressor vanes are bounded by the intermediate pressure compressor surge limit80, the fan surge limit84, and the high pressure compressor corrected rotational speed limit82. The max fan blade pitch angle is bounded by a fan corrected speed (Nc) limit85. An optimal fan blade pitch angle and compressor vane pitch angle for minimizing specific fuel consumption at the given operating condition and limits of the engine10is indicated as point90which is the intersection of the optimal fan blade pitch angle88and the optimal compressor vane pitch angle86. The optimal pitch angels may be chosen from a small range of optimal angles around the point90in some embodiments.

The graph ofFIG. 7illustrates turbine inlet temperature for the different ranges of fan blade angles and compressor vane angles. When being optimized, minimum turbine inlet temperature may be desired. In the illustrative performance characteristic of turbine inlet temperature, the ranges of angles for the fan blades and compressor vanes are bounded at least by the high pressure compressor rotational speed limit82. The max fan blade pitch angle is bounded by a fan corrected speed (Nc) limit85. An optimal fan blade pitch angle range and compressor vane pitch angle range for minimizing turbine inlet temperature at the given operating condition and limits of the engine10is indicated by range92. The turbine inlet temperature may be selected to be optimized and, thus, minimized, during key high-temperature conditions. The turbine inlet temperature may be optimized and, thus, minimized during takeoff or climb.

The graph ofFIG. 8illustrates engine noise for the different ranges of fan blade angles and compressor vane angles. In the illustrative performance characteristic of engine noise, the ranges of angles for the fan blades and compressor vanes are bounded by the intermediate pressure compressor surge limit80and the high pressure compressor rotational speed limit82. An optimal fan blade pitch angle and compressor vane pitch angle for reducing engine noise at the given operating condition and limits of the engine10is indicated as range92. The engine noise performance characteristic may be selected, for example, at taxi to reduce engine noise on the ground. The engine noise performance characteristic may be selected to be minimized, for example, at take-off and landing. The noise may be determined or based on a fan tip Mach number. In some embodiments, only the fan blade pitch angles are used to optimize engine noise as suggested inFIG. 8. As a result, the compressor vane pitch angles may be set to the optimal vane pitch angle for another parameter other than engine noise.

The graph ofFIG. 9illustrates engine operability as based on fan surge margin for the different ranges of fan blade angles and compressor vane angles. The ranges of angles for the fan blades and compressor vanes are bounded by the fan surge limit84and the high pressure compressor rotational speed limit82. The max fan blade pitch angle is bounded by a fan corrected speed (Nc) limit85. The fan surge limit84may be moved depending on requirements for a given flight condition, flight phase, engine operation condition, and/or known engine threat such as, for example, crosswinds, ambient temperature, altitude, etc.

The graph ofFIG. 10illustrates engine operability as based on intermediate pressure compressor surge margin for the different ranges of fan blade angles and compressor vane angles. The ranges of angles for the fan blades and compressor vanes are bounded by the compressor surge limit80and the high pressure compressor rotational speed limit82. The max fan blade pitch angle is bounded by a fan corrected speed (Nc) limit85. The compressor surge limit80may be moved depending on requirements for a given flight condition, flight phase, engine operation condition, and/or known engine threat such as, for example, crosswinds, ambient temperature, altitude, etc.

The graph ofFIG. 11illustrates emissions of the gas turbine engine10for the different ranges of fan blade angles and compressor vane angles. The ranges of angles for the fan blades and compressor vanes are bounded by the high pressure compressor corrected speed limit82along a bottom of the graph. An optimal fan blade pitch angle and compressor vane pitch angle for minimizing emissions at the given operating condition and limits of the engine10are indicated as range92. As one example, the emissions may be NOx.

The graphs ofFIGS. 5-11are examples of the manner in which the optimization of the operating parameters of gas turbine engine10may be calculated under various conditions. This information may then be stored in look-up tables used by the control system22to choose the angles of the fan blades28and the compressor vanes15to control the operation of the gas turbine engine10during the flight cycle54. In some embodiments, the calculations of the optimization of the angles may be conducted in real-time. In either case, the control system22adjusts the geometry of the fan12and compressor14to achieve the optimization. By combining the adjustability of the geometry of the fan12with the compressor14, the performance of a particular gas turbine engine10is expanded to take advantage the adjustment of the fan12and the compressor14cooperatively to compensate for the variations induced by on or the other of the fan12or compressor14.

In some embodiments, the engine performance characteristic is selected manually by the pilot. In some embodiments, the engine performance characteristic selected is determined by the controller based on the expected flight cycle. In some embodiments, the engine performance characteristic selected is based on real time engine performance, engine operating condition, and/or flight cycle which are calculated or received by the controller.

In some embodiments, two or more engine performance characteristics are selected to be optimized. As a result, optimal pitch angles of the fan blades and compressor vanes are selected around the optimal point90of the two or more engine performance characteristics such that the selected fan blade and compressor vane pitch angles are close to or as close as possible to the optimal point90of the two or more engine performance characteristics charts.

The engine in accordance with the present disclosure includes a variable pitch fan and variable compressor geometry and this disclosure provides a controller and method for scheduling of both sets of geometry. Variable geometry fans may be used to maintain a constant engine rotation speed, while the variable geometry in the compressor is used to maintain compressor stability. It is believed that, historically, engines have not been designed with variable geometry fans and variable geometry in the compressor. An engine of the present disclosure may operate at a constant fan speed and vary thrust in the fan by varying the fan-blade pitch and fuel flow. The variable pitch fan may be adjusted to control noise in certain environments.

According to the present disclosure, the schedules of the variable geometry of the fan blades and compressor are designed on the optimization of desired engine outputs for improved engine performance. For example, the engine parameters that may be optimized include specific fuel consumption, thrust, turbine inlet temperature, engine noise, engine operability, emissions, or any other suitable parameter.

For any given operating condition (i.e. altitude, flight Mach number, throttle setting, ambient temperature, etc.), the variable geometry can be set such that the specific fuel consumption is at its optimum setting, for example. This optimization would occur within any limits, such as mechanical and corrected speeds, stability limits, temperature limits, etc. Such limits could be steady state limits or transient limits. The schedule could be over-ruled by other requirements at a given operating condition in some embodiments. For example, the variable geometry could also be used to do any of the following: maximize available thrust during a One-Engine-Out situation; minimize turbine inlet temperature during key high-temperature conditions; and minimize engine noise during take-off and landing conditions.

To accomplish this scheduling of pitch angles, outputs of engine model optimization could be built into tables that would be referenced from the engine control or a real-time-model could be run with real-time simulation and optimization. The objective function could be set or changed by the pilot in real time during operation or set to specific functions ahead of time.