Patent Application
20090003997
This invention generally relates to a gas turbine engine, and more particularly to a gas turbine engine having a variable shape inlet section.
In an aircraft gas turbine engine, such as a turbofan engine, air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases. The hot combustion gases flow downstream through turbine stages which extract energy from the hot combustion gases. A fan section supplies air to the compressor.
Combustion gases are discharged from the turbofan engine through a core exhaust nozzle and a quantity of fan air is discharged through an annular fan exhaust nozzle defined at least partially by a nacelle assembly surrounding the core engine. A majority of propulsion thrust is provided by the pressurized fan air which is discharged through the fan exhaust nozzle, while the remaining thrust is provided by the combustion gases discharged through the core exhaust nozzle.
It is known in the field of aircraft gas turbine engines that the performance of a turbofan engine varies during diversified operability conditions experienced by the aircraft. An inlet lip section located at the foremost end of the turbofan nacelle assembly is typically designed to enable operation of the turbofan engine and reduce separation of airflow from the internal and external flow surfaces of the inlet lip section during these diversified conditions. For example, the nacelle assembly requires a “thick” inlet lip section to support operation of the engine during specific flight conditions, such as crosswind conditions, take-off conditions and the like. Disadvantageously, the “thick” inlet lip section may reduce the efficiency of the turbofan engine during normal cruise conditions of the aircraft, for example. As a result, the maximum diameter of the nacelle assembly is approximately 10-20% larger than required during cruise conditions. Since aircraft typically operate in cruise conditions for extended periods, turbofan efficiency gains can lead to substantially reduced fuel burn and emissions.
Accordingly, it is desirable to provide a nacelle assembly having an adaptive structure to improve the performance of a turbofan gas turbine engine during diversified operability conditions.
A nacelle assembly includes an inlet section having a plurality of discrete sections. Each of the plurality of discrete sections includes an adaptive structure. A thickness of each of the plurality of discrete sections is selectively adjustable between a first position and a second position to influence the adaptive structure of each of the plurality of discrete sections.
A gas turbine engine includes a compressor section, a combustor section, a turbine section, and a nacelle assembly which at least partially surrounds at least one of the compressor section, the combustor section and the turbine section. The nacelle assembly includes a plurality of discrete sections each having an adaptive structure. A leading edge and a thickness of each of the plurality of discrete sections are selectively adjustable to influence the adaptive structure of each of the plurality of discrete sections. A controller identifies an operability condition and selectively commands adjustment of each of the leading edge and the thickness of each of the plurality of discrete sections in response to sensing the operability condition.
The various features and advantages of this invention 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.
In a two-spool gas turbine engine architecture, the high pressure turbine 20 utilizes the energy extracted from the hot combustion gases to power the high pressure compressor 16 through a high speed shaft 19, and the low pressure turbine 22 utilizes the energy extracted from the hot combustion gases to power the low pressure compressor 15 and the fan section 14 though a low speed shaft 21. However, the invention is not limited to the two-spool gas turbine engine architecture described and may be used with other architectures, such as a single-spool axial design, a three-spool axial design and other architectures. That is, the present invention is applicable to any gas turbine engine, and to any application.
The example gas turbine engine 10 is in the form of a high bypass ratio turbofan engine mounted within a nacelle assembly 26, in which a significant amount of air pressurized by the fan section 14 bypasses the core engine 39 for the generation of propulsion thrust. The nacelle assembly 26 partially surrounds an engine casing 31 that houses the core engine 39 and its components. The airflow entering the fan section 14 may bypass the core engine 39 via a fan bypass passage 30 which extends between the nacelle assembly 26 and the engine casing 31 for receiving and communicating a discharge airflow F1. The high bypass flow arrangement provides a significant amount of thrust for powering the aircraft.
The engine 10 may include a geartrain 23 that controls the speed of the rotating fan section 14. The geartrain 23 can be any known gear system, such as a planetary gear system with orbiting planet gears, a planetary gear system with non-orbiting planet gears or other type of gear system. In the disclosed example, the geartrain 23 has a constant gear ratio. It should be understood, however, that the above parameters are only examples of a contemplated geared turbofan engine 10. That is, the invention is applicable to traditional turbofan engines as well as other engine architectures.
The discharge airflow F1 is discharged from the engine 10 through a fan exhaust nozzle 33. Core exhaust gases C are discharged from the core engine 39 through a core exhaust nozzle 32 disposed between the engine casing 31 and a center plug 34 disposed coaxially around a longitudinal centerline axis A of the gas turbine engine 10.
The inlet lip section 38 defines a contraction ratio. The contraction ratio represents a relative thickness of the inlet lip section 38 of the nacelle assembly 26 and is represented by the ratio of a highlight area Ha (ring shaped area defined by highlight diameter Dh) and a throat area Ta (ring shaped area defined by throat diameter Dt). Currently industry standards typically require a contraction ratio of approximately 1.33 to reduce the separation of oncoming airflow F2 from the outer and inner flow surfaces of the inlet lip section 38 during engine operation, but other contraction ratios may be feasible. “Thick” inlet lip section designs, which are associated with large contraction ratios, increase the maximum diameter Dmax and increase the weight and drag penalties associated with the nacelle assembly 26. In addition, a desired ratio of the maximum diameter Dmax relative to the highlight diameter Dh is typically less than or equal to about 1.5, for example. A person of ordinary skill in the art would understand that other ratios of the maximum diameter Dmax relative to the highlight diameter Dh are possible and will vary depending upon design specific parameters.
Referring to
In one example, the discrete sections 40 are comprised of an aluminum alloy. In another example, the discrete sections are comprised of a titanium alloy. It should be understood that any deformable material may be utilized to form the discrete sections 40. A person of ordinary skill in the art having the benefit of this description would be able to choose an appropriate material for the example discrete sections 40 of the inlet lip section 38.
Influencing the adaptive structure of the inlet lip section 38 during specific flight conditions to achieve a desired shape change increases the amount of airflow communicated through the gas turbine engine 10 and reduces the internal and external drag experienced by the inlet lip section 38. In one example, the adaptive structure of the inlet lip section 38 is influenced by adjusting the shape of the leading edge 42 of each discrete section 40 (see
At least one linkage assembly 48 is provided within each discrete section 40 and includes a plurality of linkage arms 52 and a plurality of pivot points 54. The rotary actuator 46 pivots, toggles, extends and/or flexes the linkage arms 52 of the linkage assembly 48 about the pivot points 54 to move the leading edge 42 between the “thin”, first position X and the “blunt”, second position X′. Although the present example is illustrated with a rotary actuator and linkage arms connected via pivot points, other mechanisms may be utilized to move the leading edges 42 of the discrete sections 40 between the first position X and the second position X′, including but not limited to linear actuators, bell cranks, etc. A person of ordinary skill in the art having the benefit of this disclosure will be able to implement an appropriate actuator assembly to manipulate the leading edge 42 of each discrete section 40. In addition, it should be understood that the leading edge 42 is moveable to any position between the first position X and second position X′.
The adaptive structure of the inlet lip section 38 is influenced by moving the leading edge 42 of each discrete section 40 between the first position X and the second position X′ in response to detecting an operability condition of the gas turbine engine 10. In one example, the operability condition includes a take-off condition. In another example, the operability condition includes a climb condition. In yet another example, the operability condition includes a landing condition. In still another example, the operability condition includes a high angle of attack condition. It should be understood that the adaptive structure of the inlet lip section 38 is adjustable in response to any operability condition experienced by the aircraft. Each leading edge 42 is positioned at/returned to the first position X during normal cruise conditions of the aircraft.
A sensor 61 detects the operability condition and communicates with a controller 62 to translate the leading edge 42 between the first position X and the second position X′ and influence the adaptive structure of the inlet lip section 38. Of course, this view is highly schematic. In addition, the illustrations of the movement of the inlet lip section 38 are shown exaggerated to better illustrate the adaptive structure thereof. A person of ordinary skill in the art would understand the distances the leading edge 42 should be moved between the position X and the second position X′ in response to sensing a specific operability condition.
It should be understood that the sensor 61 and the controller 62 may be programmed to detect any known operability condition and that each operability condition may be associated with a distinct position of the leading edge 42 of the inlet lip section 38. That is, the sensor 61 and the controller 62 are operable to situate the leading edge 42 of each discrete section 40 at a position which corresponds to the operability condition that is detected. Also, the sensor can be replaced by any controller associated with the gas turbine engine 10 or an associated aircraft. In fact, the controller 62 itself can include the “sensor” and generate the signal to adjust the contour of the inlet lip section 38.
The thickness T adjustment of each body panel portion 44 is achieved via a linear actuator 56 and a linkage assembly 58. The linear actuator 56 and the linkage assembly 58 are received in the cavity 50 of each discrete section 40. Although the present example is illustrated with a linear actuator and linkage arms connected via pivot points, other mechanisms may be utilized to adjust the thickness T of each body panel portion 44.
The linear actuator 56 includes an actuator arm 60 which is moveable in a R or L direction to move the linkage assembly 58 and thereby adjust the thickness of the body panel portion 44. The linkage assembly 58 includes a plurality of linkages 64 and a plurality of pivot points 66. The linear actuator 56 adjusts the thickness T of each body panel portion 44 by retracting, pivoting, toggling, extending and/or flexing the linkages 64 about each pivot point 66. In one example, the actuator arm 60 of the linear actuator 56 moves in a R direction to retract the outer skin (i.e., move the outer skin in the Z direction) of the body panel portion 44 and provide a “thin” inlet lip section 38. In another example, the actuator arm 60 of the linear actuator 56 is moved in a L direction to expand the outer skin (i.e., move the outer skin in the Y direction) of the body panel portion 44 and provide a “thick” inlet lip section 38. That is, the thickness T of each body panel portion 44 is adjusted either radially outwardly or radially inwardly to provide a “thick” inlet lip section or a “thin” inlet lip section, respectively.
The thickness of each discrete section 40 is adjusted in response to detecting an operability condition. In one example, the operability condition includes a take-off condition. In another example, the operating condition includes a climb condition. In another example, the operability condition includes a high angle of attack condition. In still another example, the operability condition includes a landing condition. It should be understood that the thickness of the body panel portion 44 may be adjusted to influence the adaptive structure of the inlet lip section 38 in response to any operability condition experienced by the aircraft. The thickness T is adjusted/returned to a “thin” position at cruise conditions of the aircraft.
A sensor 61, as is shown in
It should be understood that the sensor 61 and the controller 62 may be programmed to detect any known operability condition and that each operability condition may be associated with a distinct thickness T of the body panel portions 44 of the discrete sections 40. That is, the sensor 61 and the controller 62 are operable to adjust the thickness T of each discrete section 40 to a position which corresponds to the operability condition that is detected. The thickness T of each discrete section 40 may be adjusted uniformly or differently about the circumference. In some instances, such as operating during strong cross-winds, for example, only certain discrete sections 40 may be adjusted, while other discrete sections 40 are left unchanged. Also, the sensor can be replaced by any controller associated with the gas turbine engine 10 or an associated aircraft. In fact, the controller 62 itself can generate the signal to adjust the contour of the inlet lip section 38.
Although illustrated in
Influencing the adaptive structure of the inlet lip section 38 may also be achieved during diverse operating conditions by “drooping” a portion of the inlet lip section 38 relative to a remaining portion of the inlet lip section 38 (See
The adaptive inlet lip section 38 improves aerodynamic performance of the gas turbine engine 10 during all operability conditions experienced by the aircraft. In addition, because of the shape changing capabilities of the inlet lip section 38, the aircraft may be designed having a “thin” inlet lip section 38 (i.e., a slim line nacelle having a reduced contraction ratio is achieved). As a result, the nacelle assembly 26 is designed for specific cruise conditions of the aircraft. A reduced maximum diameter of the nacelle assembly 26 may therefore be achieved while reducing weight, reducing drag, reducing fuel burn and increasing the overall efficiency of the gas turbine engine 10.
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 invention. For that reason, the following claims should be studied to determine the true scope and content of this invention.
