Patent Application
20130121824
This disclosure relates to a gas turbine engine having a core auxiliary duct passage for diverting a portion of a core airflow from the core engine of the gas turbine engine.
In an aircraft gas turbine engine, such as a turbofan engine, air is pressurized in a compressor section and mixed with fuel in a combustor section for generating hot combustion gases. The hot combustion gases flow downstream through a turbine section that extracts energy from the gases. The turbine section powers a compressor section and a fan section disposed upstream of the compressor section.
Fan bypass airflow is communicated through a fan bypass passage that extends between a nacelle assembly and a core engine. The fan bypass airflow is communicated through an annular fan exhaust nozzle defined at least partially by the nacelle assembly surrounding the core engine. A majority of propulsion thrust is provided by the pressurized fan air that is discharged through the fan exhaust nozzle. The combustion gases are discharged through a core exhaust nozzle to provide additional thrust.
Mixed flow turbofan engines are known that include a mixer positioned between the nacelle assembly and the core engine at a position downstream from a turbine exit guide vane. The mixer typically includes a plurality of petals. The mixer drives core airflow from the core engine radially outward and into the petals of the mixer, and drives the fan airflow from the fan bypass passage radially inward to fill the petals of the mixer. The two airflow streams are co-mingled in the mixer and are subsequently communicated as a mixed stream through the exhaust nozzles of the gas turbine engine at a relatively equal velocity.
Mixed flow turbofans are known to provide noise reductions and improved propulsion efficiency of gas turbine engines. However, noise and efficiency issues remain a common area of concern in the field of gas turbine engines. Attempts have been made to increase the beneficial results achieved by mixed flow turbofan engines. Disadvantageously, these attempts have not been successful.
Accordingly, it is desirable to provide a gas turbine engine that achieves improved efficiency and noise reductions in a relatively inexpensive and non-complex manner.
A gas turbine engine system according to an exemplary aspect of the present disclosure includes, among other things, a nacelle assembly defined about an axis and a core engine positioned radially inward from the nacelle assembly and having a core passage and at least one core auxiliary duct passage. The at least one core auxiliary duct passage includes an inlet for receiving a portion of a core airflow from the core engine and an outlet for discharging the portion of the core airflow. At least one of the inlet and the outlet are selectively translatable to divert the portion of the core airflow into the at least one core auxiliary duct passage and a mixer disposed between the nacelle assembly and the core engine.
In a further non-limiting embodiment of the foregoing gas turbine engine system, the inlet is positioned upstream from the mixer.
In a further non-limiting embodiment of either of the foregoing gas turbine engine systems, the outlet is positioned downstream from the mixer.
In a further non-limiting embodiment of any of the foregoing gas turbine engine systems, the inlet includes at least one door and a translating ring that selectively translates the at least one door.
In a further non-limiting embodiment of any of the foregoing gas turbine engine systems, the outlet includes at least one door pivotable about a pivot.
In a further non-limiting embodiment of any of the foregoing gas turbine engine systems, a fan bypass passage is disposed between the nacelle assembly and the core engine.
In a further non-limiting embodiment of any of the foregoing gas turbine engine systems, a fan exhaust nozzle is positioned near a downstream end of the nacelle assembly and a core exhaust nozzle is positioned near a downstream end of the core engine.
In a further non-limiting embodiment of any of the foregoing gas turbine engine systems, the at least one core auxiliary duct passage extends circumferentially about the core engine.
In a further non-limiting embodiment of any of the foregoing gas turbine engine systems, the mixer includes a plurality of petals.
In a further non-limiting embodiment of any of the foregoing gas turbine engine systems, the at least one core auxiliary duct passage is positioned radially inward of the core engine.
A gas turbine engine system according to an exemplary aspect of the present disclosure includes, among other things, a nacelle assembly defined about an axis and a core engine positioned at least partially within the nacelle assembly and including at least one compressor section, a combustor section and at least one turbine section. The core engine includes a core passage and at least one core auxiliary duct passage having an inlet for receiving a portion of a core airflow from the core engine and an outlet for discharging the portion of the core airflow. A mixer is disposed between the nacelle assembly and the core engine. A controller produces a signal in response to detecting an operability condition and selectively translates the inlet and the outlet in response to the operability condition.
In a further non-limiting embodiment of the foregoing gas turbine engine system, the operability condition includes a take-off condition.
In a further non-limiting embodiment of either of the foregoing gas turbine engine systems, the inlet and the outlet are selectively translatable between a first position and a second position.
In a further non-limiting embodiment of any of the foregoing gas turbine engine systems, the first position is a closed position and the second position is an open position.
In a further non-limiting embodiment of any of the foregoing gas turbine engine systems, the system comprises a sensor that communicates with the controller.
In a further non-limiting embodiment of any of the foregoing gas turbine engine systems, the inlet and the outlet are selectively moveable between a plurality of positions, and each of the plurality of positions allows a different amount of the core airflow to be communicated through the at least one core auxiliary duct passage.
The various features and advantages of this disclosure will become apparent to those skilled in the art from the following detailed description. The drawings that accompany the detailed description can be briefly described as follows.
The example gas turbine engine 10 is in the form of a high bypass ratio engine mounted within a nacelle assembly 26, in which most of the air pressurized by the fan section 14 bypasses the core engine 28 for generating propulsion thrust. The nacelle assembly 26 partially surrounds the core engine 28. The airflow entering the fan section 14 may bypass the core engine 28 via a fan bypass passage 27 that extends between the nacelle assembly 26 and the core engine 28 for receiving and communicating a discharge airflow F1. The high bypass flow arrangement provides a significant amount of thrust for powering the aircraft.
The discharge airflow F1 is discharged from the engine through a fan exhaust nozzle 30 positioned adjacent a downstream end 32 of the nacelle assembly 26. Meanwhile, core airflow F2 is communicated through a core passage 34 of the core engine 28. Core airflow F2 is discharged from the core engine 28 through a core exhaust nozzle 36 that is defined between the core engine 28 and a tail cone 38 disposed coaxially therein around the longitudinal centerline axis A of the gas turbine engine 10. A bypass ratio is defined that represents the ratio of the fan discharge airflow F1 relative to the core airflow F2.
The core auxiliary duct passage 44 includes an inlet 46 and an outlet 48. In one example, the inlet 46 is positioned upstream from the mixer section 40. In another example, the inlet 46 is positioned on the core engine 28 between a turbine exit guide vane 45 and the mixer section 40. The outlet 48 is positioned downstream from the mixer section 40, in this example. However, it should be understood that the inlet and outlet 46, 48 may be positioned at other locations of the gas turbine engine 10 and that these locations may vary depending upon design specific parameters including, but not limited to, the efficiency and noise requirements of the gas turbine engine 10.
The inlet 46 of the core auxiliary duct passage 44 selectively receives a portion F3 of the core airflow F2 that is communicated through the core passage 34 of the core engine 28 in response to specific operability conditions. The portion F3 of the core airflow F2 is communicated through the core auxiliary duct passage 44 and is discharged via the outlet 48.
Diverting a portion F3 of the core airflow F2 through the core auxiliary duct passage 44 increases the gas turbine engine 10 bypass ratio and thereby improves overall engine efficiency and reduces engine noise. Specifically, communicating airflow through the core auxiliary duct passage 44 enables an increased core airflow F2 through the core passage 34 and reduces any backpressure (e.g., pressure losses that result in reductions in engine efficiency) experienced by the low pressure turbine 22. In addition, diverting core airflow F2 away from the mixer section 40 enables the fan bypass airflow F1 to increase, thereby improving engine efficiency.
The inlet 46 and the outlet 48 are selectively translated to divert the portion F3 of the core airflow F2 into the core auxiliary duct passage 44. For example, opening the inlet 46 and the outlet 48 permits an airflow F3 to enter the core auxiliary duct passage 44, and closing the inlet 46 and the outlet 48 blocks any airflow
F3 from entering the core auxiliary duct passage 44. The inlet 46 and the outlet 48 are selectively moveable between a first position X (i.e., a closed position, represented by phantom lines) to a second position X′ (an open position, represented by solid lines) in response to detecting an operability condition of a gas turbine engine 10, for example. In another example, the inlet 46 and the outlet 48 are selectively moveable between a plurality of positions, each allowing a different amount of airflow F3 to enter the core auxiliary duct passage 44.
In one example, the operability condition includes a takeoff condition. However, the inlet 46 and the outlet 48 may be selectively opened to the second position X′, or to any intermediate position between the first position X and the second position X′, in response to any known operability condition. In one example, a sensor 52 detects the operability condition and communicates a signal to a controller 54 to move the inlet 46 and the outlet 48 between the first positions X and the second positions X′ via an actuator assembly 56. Of course, this view is highly schematic.
It should be understood that the sensor 52 and the controller 54 may be programmed to detect any known operability condition. Also, the sensor 52 can be replaced by any control associated with the gas turbine engine 10 or an associated aircraft. In fact, the controller 54 itself can generate the signal to cause the actuation of the inlet 46 and the outlet 48. The actuator assembly 56 returns the inlet 46 and the outlet 48 to the first position X during normal cruise operation (e.g., a generally constant speed at a generally constant, elevated altitude), in one example. The actuator assembly 56 may include any known type of actuator or combination of actuators that include hydraulic and electric actuation systems. In another example, the inlet 46 and the outlet 48 are returned to the first position X in response to detecting a climb condition.
Although the different examples have the specific components shown in the illustrations, embodiments of this invention are not limited to those particular combinations. It is possible to use some of the components or features from one of the examples in combination with features or components from another one of the examples.
The foregoing description shall be interpreted as illustrative and not in any limiting sense. A worker of ordinary skill in the art would recognize that certain modifications would come within the scope of this disclosure. For that reason, the following claims should be studied to determine the true scope and content of this disclosure.
This application is a continuation application of U.S. patent application Ser. No. 11/866,547, filed Oct. 3, 2007.
| Number | Date | Country | |
|---|---|---|---|
| Parent | 11866547 | Oct 2007 | US |
| Child | 13735345 | US |
