Patent Application
20180230845
The present subject matter relates generally to a propulsion system for an aircraft, and an aircraft including the same.
A conventional aircraft generally includes a fuselage, a pair of wings, and a propulsion system that provides thrust. The propulsion system typically includes at least two aircraft engines, such as turbofan jet engines. Each turbofan jet engine is mounted to a respective one of the wings of the aircraft, such as in a suspended position beneath the wing.
Additionally, turbofan jet engines are typically designed to maintain a maximum internal operating temperature below a certain threshold when operated at ground-level conditions (i.e., when ingesting air at ambient, ground-level temperatures). However, once an aircraft reaches cruise altitudes, the turbofan jet engines are ingesting air at temperatures much lower than ambient ground-level temperatures. Accordingly, with at least certain turbofan jet engines, there is room to increase an internal operating temperature, and thus to increase an overall pressure ratio of the engine, when operating at such cruise altitudes.
Accordingly, a propulsion system that may more fully utilize an operability range of a gas turbine engine during cruise operating modes would be useful.
Aspects and advantages of the invention will be set forth in part in the following description, or may be obvious from the description, or may be learned through practice of the invention.
In one exemplary embodiment of the present disclosure, a propulsion system for an aircraft is provided. The propulsion system includes a gas turbine engine including a turbomachine and a primary fan. The turbomachine includes a first turbine and a second turbine, and the primary fan is driven by the second turbine. The propulsion system also includes an electric machine operable with at least one of the first turbine or the second turbine. The propulsion system additionally includes an auxiliary propulsor assembly configured to be mounted at a location away from the gas turbine engine. The electric machine is in electrical communication with the auxiliary propulsor assembly for transferring power with the auxiliary propulsor assembly.
In another exemplary embodiment of the present disclosure an aircraft is provided. The aircraft includes a first propulsion system including a gas turbine engine having a primary fan and a turbomachine, an electric generator and an auxiliary propulsor assembly. The turbomachine is drivingly connected to the electric generator, and the electric generator is electrically coupled to the auxiliary propulsor assembly for driving the auxiliary propulsor assembly. The aircraft additionally includes a second propulsion system including a gas turbine engine having a primary fan and a turbomachine, an electric generator and an auxiliary propulsor assembly. The turbomachine is drivingly connected to the electric generator, and the electric generator is electrically coupled to the auxiliary propulsor assembly for driving the auxiliary propulsor assembly.
In an exemplary aspect of the present disclosure a method for operating a propulsion system for an aircraft is provided. The propulsion system includes a gas turbine engine, an electric generator and an auxiliary propulsor assembly, the gas turbine engine drivingly connected to the electric generator and the electric generator electrically coupled to the auxiliary propulsor assembly for driving the auxiliary propulsor assembly. The method includes operating the gas turbine engine in a takeoff operating mode such that a turbomachine of the gas turbine engine defines a first overall pressure ratio and provides the auxiliary propulsor assembly with a first amount of electric power through the electric generator. The method also includes operating the gas turbine engine in a cruise operating mode such that the turbomachine of the gas turbine engine defines a second overall pressure ratio and provides the auxiliary propulsor assembly with a second amount of electric power through the electric generator. The second overall pressure ratio is greater than the first overall pressure ratio.
These and other features, aspects and advantages of the present invention will become better understood with reference to the following description and appended claims. The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate embodiments of the invention and, together with the description, serve to explain the principles of the invention.
A full and enabling disclosure of the present invention, including the best mode thereof, directed to one of ordinary skill in the art, is set forth in the specification, which makes reference to the appended figures, in which:
Reference will now be made in detail to present embodiments of the invention, one or more examples of which are illustrated in the accompanying drawings. The detailed description uses numerical and letter designations to refer to features in the drawings. Like or similar designations in the drawings and description have been used to refer to like or similar parts of the invention. As used herein, the terms “first”, “second”, and “third” may be used interchangeably to distinguish one component from another and are not intended to signify location or importance of the individual components.
Referring now to the drawings, wherein identical numerals indicate the same elements throughout the figures,
Each of the wings 20, 22 for the exemplary embodiment depicted includes one or more leading edge flaps 28 and one or more trailing edge flaps 30. The aircraft 10 further includes a vertical stabilizer 32 having a rudder flap (not shown) for yaw control, and a pair of horizontal stabilizers 34, each having an elevator flap 36 for pitch control. The fuselage 12 additionally includes an outer surface or skin 38. It should be appreciated however, that in other exemplary embodiments of the present disclosure, the aircraft 10 may additionally or alternatively include any other suitable configuration. For example, in other embodiments, the aircraft 10 may include any other configuration of stabilizer.
Referring now also to
Referring particularly to
As shown in
The exemplary core turbine engine 102 depicted generally includes a substantially tubular outer casing 106 that defines an annular inlet 108. The outer casing 106 encases, in serial flow relationship, a compressor section including a booster or low pressure (LP) compressor 110 and a high pressure (HP) compressor 112; a combustion section 114; a turbine section including a first, high pressure (HP) turbine 116 and a second, low pressure (LP) turbine 118; and a jet exhaust nozzle section 120.
The exemplary core turbine engine 102 of the turbofan 100 additionally includes one or more shafts rotatable with at least a portion of the turbine section and, for the embodiment depicted, at least a portion of the compressor section. More particularly, for the embodiment depicted, the turbofan 100 includes a high pressure (HP) shaft or spool 122, which drivingly connects the HP turbine 116 to the HP compressor 112. Additionally, the exemplary turbofan 100 includes a low pressure (LP) shaft or spool 124, which drivingly connects the LP turbine 118 to the LP compressor 110.
As stated, the primary fan of the first propulsor assembly 52 is configured as the fan 104 for the embodiment depicted. Further, the exemplary fan 104 depicted is configured as a variable pitch fan having a plurality of fan blades 128 coupled to a disk 130 in a spaced apart manner. The fan blades 128 extend outwardly from disk 130 generally along the radial direction R1. Each fan blade 128 is rotatable relative to the disk 130 about a respective pitch axis P by virtue of the fan blades 128 being operatively coupled to a suitable actuation member 132 configured to collectively vary the pitch of the fan blades 128. The fan 104 is mechanically coupled to the LP shaft 124. More particularly, the fan 104, including the fan blades 128, disk 130, and actuation member 132, is mechanically coupled to the LP shaft 124 through a power gearbox 134, and is rotatable about the longitudinal axis 101 by the LP shaft 124 across the power gear box 134. The power gear box 134 includes a plurality of gears for stepping down the rotational speed of the LP shaft 124 to a more efficient rotational fan speed. Accordingly, the fan 104 is powered by an LP system (including the LP turbine 118) of the core turbine engine 102.
Referring still to the exemplary embodiment of
Referring still to
Referring still to
It should be appreciated, however, that in other embodiments, the electric generator 56 may be positioned in any other suitable location within the core turbine engine 102, or elsewhere. For example, the electric generator 56 may be, in other embodiments, mounted coaxially with the HP shaft 122 within the turbine section, or alternatively may be offset from the HP turbine 122 and driven through a suitable gear train. Additionally, or alternatively, the electric generator 56 may be driven by both the LP system (e.g., the LP shaft 124) and the HP system (e.g., the HP shaft 122) via a dual drive system. For example, a gear assembly, such as an epicyclic gear assembly, may be provided to allow both the LP shaft 124 and HP shaft 122 to drive the electric generator 56. Additionally, or alternatively still, in various other exemplary embodiments, the electric machine/electric generator 56 may instead be operable with just the LP system. For example, referring briefly to
It should further be appreciated that the exemplary turbofan engine 100 depicted in
Referring now particularly to
Referring particularly to
In certain exemplary embodiments, the plurality of fan blades 208 may be attached in a fixed manner to the fan shaft 210, or alternatively, the plurality of fan blades 208 may be rotatable relative to the fan shaft 210, such as in the embodiment depicted. For example, the plurality of fan blades 208 each define a respective pitch axis P2, and are attached to the fan shaft 210 such that a pitch of each of the plurality of fan blades 208 may be changed, e.g., in unison, by a pitch change mechanism 211. Changing the pitch of the plurality of fan blades 208 may increase an efficiency of the second propulsor assembly 54 and/or may allow the second propulsor assembly 54 to achieve a desired thrust profile. With such an exemplary embodiment, the fan 204 may be referred to as a variable pitch fan.
Moreover, for the embodiment depicted, the auxiliary propulsor assembly 200 depicted additionally include a fan casing or outer nacelle 212, attached to a core 214 of the auxiliary propulsor assembly 200 through one or more struts or outlet guide vanes 216. For the embodiment depicted, the outer nacelle 212 substantially completely surrounds the fan 204, and particularly the plurality of fan blades 208. Accordingly, for the embodiment depicted, the auxiliary propulsor assembly 200 may be referred to as a ducted electric fan assembly.
Notably, the fan 204 of the auxiliary propulsor assembly 200 may define a fan pressure ratio. The fan pressure ratio may generally refer to a ratio of a fan discharge pressure to a fan inlet pressure. As will be described in greater detail below, the propulsion system 50 may be operated during certain operations such that the auxiliary propulsor assembly provides relatively efficient thrust. For example, during, e.g., cruise operations of the propulsion system 50, the fan 204 of the auxiliary propulsor assembly 200 may define a fan pressure ratio of less than about 1.4:1. More specifically, in certain exemplary embodiments, the auxiliary fan 204 of the auxiliary propulsor assembly 200 may define, during cruise operations, a fan pressure ratio of less than about 1.3:1, such as less than about 1.2:1. It should be appreciated, that as used herein, terms of approximation, such as “about” or “approximately,” refer to being within a 10% margin of error. Additionally, the term “cruise operations” generally refers to a level flight segment that occurs between an ascent phase and a descent phase of the flight, at which the aircraft is designed for optimum performance.
Referring still particularly to
Notably, however, in certain exemplary embodiments, the electric motor 206 may be configured as a motor/generator. Accordingly, during, e.g., emergency operations, the auxiliary propulsor assembly 200 may operate as a ram air turbine, such that inlet air to the auxiliary propulsor assembly 200 rotates the plurality of fan blades 208 of the fan 204, in turn rotating the electric motor/generator, allowing the electric motor/generator to operate as an electric generator providing electrical power to the power bus 58. Notably, with such an exemplary embodiment, the electric generator 56 of the turbofan engine 100 of
The fan shaft 210 is supported by one or more bearings 218, such as the one or more roller bearings, ball bearings, or any other suitable bearings. Additionally, the electric motor 206 may be an inrunner electric motor (i.e., including a rotor positioned radially inward of a stator), or alternatively may be an outrunner electric motor (i.e., including a stator positioned radially inward of a rotor). As briefly noted above, the electric generator 56 of the propulsion system 50 is in electrical communication with the auxiliary propulsor assembly 200 for powering the auxiliary propulsor assembly 200. More particularly, the electric motor 206 of the auxiliary propulsor assembly 200 is in electrical communication with the electrical power bus 58, which for the embodiment depicted includes one or more electrical lines 60 electrically connected to the electric motor 206. Accordingly, the electric motor 206 is more particularly in electrical communication with the electrical power bus 58 through one or more electrical lines 60 of the electrical power bus 58, and the electrical power bus 58 may deliver power to the electric motor 206 for driving the electric motor 206, and in turn driving the fan 204. Notably, for the embodiment depicted, the electrical power bus 58 further includes one or more electrical disconnects 61, such that the electrical power bus 58 may isolate one or more components in the event of an electrical failure of one or more components. The one or more electrical disconnects 61 may be manually operated, or alternatively, may be automatically triggered in the event of an electrical failure.
Referring again briefly to
It should be appreciated, however, that in other embodiments, the auxiliary propulsor assembly 200 may have any other suitable configuration. For example, referring now to
For example, the exemplary auxiliary propulsor assembly 200 of
Additionally, an outer nacelle 212 and an electric motor 206 are provided, with the outer nacelle 212 surrounding the auxiliary fan 204. However, for the embodiment depicted, the electric motor 206 is not configured to drive the auxiliary fan 204 through fan shaft 210. Instead, the electric motor 206 is at least partially integrated into the tips 220 of one or more of the plurality of fan blades 208 for rotating the plurality of fan blades 208 directly. More specifically, the exemplary electric motor 206 of
A propulsion system in accordance with an exemplary embodiment of the present disclosure may provide for more efficient propulsion for an aircraft. For example, typically an overall pressure ratio of a gas turbine engine is limited by a temperature limit certain components of the gas turbine engine may withstand during takeoff operations (i.e., at sea level and typical sea level ambient temperatures). However, once an aircraft incorporating such a gas turbine engine reaches cruising altitudes, the ambient temperature of air ingested by the gas turbine engine is greatly reduced. Accordingly, the temperatures within such a gas turbine engine are also greatly reduced. When the gas turbine engine is operated during cruise operations to simply provide thrust directly and through, e.g., a fan (such as in a turbofan configuration), the gas turbine engine may not be utilizing its full potential. However, with the present disclosure, energy may be extracted in certain exemplary aspects from both an LP system (through the fan) and an HP system (through an electric generator) of the gas turbine engine, requiring the engine to operate at increased overall pressure ratios, and thus increased temperatures. Specifically, extracting energy through both the LP system (i.e., through a primary fan) and the HP system (i.e., through an electric generator) requires an increased amount of energy to be generated by the turbomachine, which in turn requires an increased overall pressure ratio. Turbomachinery operating at an increased overall pressure ratio generally operates more efficiently. Notably, as discussed above, the energy extracted from the HP system may be transferred to an auxiliary fan of an auxiliary propulsor assembly to generate additional thrust relatively efficiently.
Additionally, as used herein, “overall pressure ratio” of a gas turbine engine refers to a pressure ratio of a compressor section of a turbomachine of a gas turbine engine (e.g., for the embodiment of
Moreover, it should also be appreciated, that in still other exemplary embodiments, the exemplary propulsion system 50 described above with reference to
Additionally, in still other embodiments, the exemplary propulsion system may be integrated into an aircraft 10 in any other suitable manner. For example, referring now to
For example, the exemplary aircraft 10 of
However, for the embodiment depicted, the aircraft 10 further includes a second propulsion system 258, the second propulsion system 258 also configured in accordance with an exemplary embodiment of the present disclosure. For example, the second propulsion system 258 may also be configured in substantially the same manner as exemplary propulsion system 50 described above with reference to one or more of
Specifically speaking, the gas turbine engine 252 of the first propulsion system 250 is mounted to the port side wing 20 and drives the auxiliary propulsor assembly 254 mounted to the starboard side wing 22. Additionally, the gas turbine engine 260 of the second propulsion system 258 is mounted to the starboard side wing 22 and drives the auxiliary propulsor assembly 262 mounted to the port side wing 20.
More generally, for the embodiment depicted, the gas turbine engine 252 of the first propulsion system 250 is mounted to one of the port side wing 20 or starboard side wing 22 and the auxiliary propulsor assembly 254 of the first propulsion system 250 is mounted to the other of the port side wing 20 or starboard side wing 22. Similarly, the gas turbine engine 260 of the second propulsion system 258 is mounted to one of the port side wing 20 or the starboard side wing 22 and the auxiliary propulsor assembly 262 of the second propulsion system 258 is mounted to the other of the port side wing 20 or the starboard side wing 22.
Specifically, for the embodiment depicted, the gas turbine engine 252 of the first propulsion system 250 is mounted to the port side wing 20 and the auxiliary propulsor assembly 254 of the first propulsion system 250 is mounted to the starboard side wing 22. By contrast, the gas turbine engine 260 of the second propulsion system 258 is mounted to the starboard side wing 22 and the auxiliary propulsor assembly 262 of the second propulsion system 258 is mounted to the port side wing 20.
An aircraft in accordance with the exemplary embodiment of
It should be appreciated, however, that in still other exemplary embodiments, the exemplary propulsion system may be integrated into an aircraft 10 in any other suitable manner. For example, referring now to
For example, the exemplary aircraft 10 of
Additionally, the exemplary aircraft 10 of
Moreover, as with the embodiment of
Notably, referring back briefly to
As will be appreciated, such a configuration may provide for a more equally balanced aircraft 10 (from a weight standpoint). For example, as is depicted in
It should be appreciated, however, that the exemplary propulsion systems described above, and the exemplary aircraft configurations incorporating such propulsion systems described above, are provided for exemplary purposes only. In other exemplary embodiments, the propulsion systems and/or aircraft may have any other suitable configuration. For example, in other exemplary embodiments, a propulsion system in accordance with the present disclosure may include a plurality of auxiliary propulsor assemblies mounted in any suitable configuration, each driven by an electric generator of the propulsion system. Additionally, in still other embodiments, one or more of the propulsion systems may include energy storage devices (such as batteries). Further, for example, referring now to
For example, the exemplary aircraft 10 of
However, for the exemplary embodiment of
Additionally, in still other exemplary embodiments of the present disclosure, the propulsion systems, and the aircraft incorporating such propulsion systems, may be configured in still other suitable manners. For example, in still other exemplary embodiments, one or both of a gas turbine engine and an auxiliary propulsor assembly of a propulsion system may be mounted to a fuselage of the aircraft proximate a tail end of the aircraft. For example, the propulsion system may include the gas turbine engine mounted to one side of the fuselage and the auxiliary propulsor assembly mounted to an opposite side of the fuselage (e.g., the gas turbine engine mounted to one of a starboard side 26 or port side 24 of a fuselage 12 of an aircraft 10 proximate a tail end 18, and the auxiliary propulsor assembly mounted to the other of the starboard side 26 or port side 24 of the fuselage 12 of the aircraft 10 proximate the tail end 18). Moreover, in still other exemplary embodiments, one or both of the gas turbine engine and the auxiliary propulsor assembly of the propulsion system may be mounted to a stabilizer of the aircraft, such as to a vertical stabilizer of the aircraft. In either of these embodiments, the propulsion systems may further include one of the gas turbine engine or auxiliary propulsor assembly mounted in an under-wing configuration as well.
Referring now to
For the exemplary aspect depicted, the exemplary method (300) includes at (302) operating the gas turbine engine in a takeoff operating mode, such that the turbomachine defines a first overall pressure ratio and provides the auxiliary propulsor assembly with a first amount of electric power from the electric generator. Additionally, at (304) the method (300) includes operating the gas turbine engine in a cruise operating mode, such that the turbomachine defines a second overall pressure ratio and provides the auxiliary propulsor assembly with a second amount of electric power from the electric generator.
For the exemplary aspect depicted, the second overall pressure ratio defined by the turbomachine of the gas turbine engine at (304) is greater than the first overall pressure ratio defined by the turbomachine of the gas turbine engine at (302). More specifically, in at least certain exemplary aspects, the second overall pressure ratio may be at least about 5% greater, at least about 10% greater, or at least about 20% greater than the first overall pressure ratio. Further, in at least certain exemplary aspects, the second amount of power provided to the auxiliary propulsor assembly at (304) may be substantially equal to the first amount of power provided to the auxiliary propulsor assembly at (302). Accordingly, the method 300 may include providing a substantially constant amount of power to the auxiliary propulsor assembly during takeoff and cruise. For example, in certain exemplary aspects, the method 300 may include providing a substantially constant amount of power to the auxiliary propulsor throughout a flight envelope.
As will be appreciated, an inlet air temperature and density for the turbomachine of the gas turbine engine during takeoff operating conditions is generally higher than an inlet air temperature and density for the turbomachine of the gas turbine engine during cruise operating conditions. Such allows for the turbomachine of the gas turbine engine to produce more power during takeoff as compared to during cruise. Accordingly, providing a substantially constant amount of power to the auxiliary propulsor during takeoff and cruise operating conditions results in a smaller fraction of power extraction at takeoff as compared to the relatively larger fraction of power extraction at cruise. The fraction of power extraction refers to a ratio of power provided to the auxiliary propulsor from the turbomachine to a total amount of power generated by the turbomachine.
The relatively larger fraction of power extraction at cruise is accomplished, at least in part, by increasing a core speed of the turbomachine, and accordingly increasing an overall pressure ratio of the turbomachine. Notably, a decreased inlet air temperature allows for the increase in overall pressure ratio of the turbomachine during cruise operations. For example, the turbomachine is typically operated to a maximum compressor discharge temperature and/or exhaust temperature. By reducing an inlet air temperature, a greater pressure increase is allowed across a compressor section of the turbomachine, while maintaining the compressor discharge temperature and/or exhaust temperature at or below the maximum compressor discharge temperature and/or exhaust temperature.
Further, it will be appreciated, that by increasing the core speed and overall pressure ratio of the turbomachine during cruise operations, an increased marginal fuel efficiency is accomplished for the turbomachine of the gas turbine engine. For example, the energy extracted by the generator of the propulsion system drives the auxiliary propulsor, effectively increasing an overall bypass ratio of the gas turbine engine, and thus increasing its overall propulsive efficiency, while simultaneously increasing the turbomachine efficiency due to the higher overall core speed and overall pressure ratio of the turbomachine.
Additionally, as is depicted schematically in
With such an exemplary aspect, the method (300) may further include at (306) operating the second gas turbine engine of the second propulsion system in a takeoff operating mode such that the second turbomachine of the second gas turbine engine defines a third overall pressure ratio and provides the second auxiliary propulsor assembly of the second propulsion system with a third amount of electric power through the second electric generator of the second propulsion system. Moreover, the method (300) may further include at (308) operating the second gas turbine engine in a cruise operating mode such that the second turbomachine of the second gas turbine engine defines a fourth overall pressure ratio and provides the second auxiliary propulsor assembly with a fourth amount of electric power through the second electric generator. The fourth overall pressure ratio may be greater than the third overall pressure ratio (such as at least about 5% greater), and in certain exemplary aspects, the fourth amount of electric power may be substantially equal to the third amount of electric power.
This written description uses examples to disclose the invention, including the best mode, and also to enable any person skilled in the art to practice the invention, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the invention is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they include structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal languages of the claims.
