Mechanically driven air carrier thermal management device

Abstract

The present disclosure is directed to an aircraft power generation system, which includes a reverse Brayton cycle system, a gas turbine engine and a gearbox. The gas turbine engine includes a compressor section, a turbine section and an engine shaft. The compressor section and the turbine section are arranged in a series flow arrangement. The engine shaft can rotate with at least a portion of the compressor section and at least a portion of the turbine section. The reverse Brayton cycle system includes a compressor, a drive shaft, a turbine and a first exchanger. The drive shaft can rotate with the compressor or the turbine, and the compressor, the first heat exchanger and the turbine are in a series flow arrangement. The gearbox is configured to receive mechanical energy from the engine shaft and transmit the mechanical energy to the reverse Brayton cycle system through the drive shaft.

Description

Technical Field

[0001] The present invention generally relates to powering aircraft systems using gas turbine engines. More specifically, the present subject matter relates to a thermal management system mechanically driven by a gas turbine engine. Background Art

[0002] Aircraft gas turbine engines transmit propulsion thrust for aircraft motion and provide energy for aircraft systems such as thermal management systems (TMS), which may include environmental control systems (ECS). TMS also includes aircraft propulsion and electrical power generation units, fluid heat exchangers, electronic cooling devices, and management of energy and heat transfer in environmental control systems (ECS), which provide conditioned air to aircraft cabins, storage devices, and avionics. ECS generally includes a reverse Brayton cycle system (RBC), which is pneumatically driven by compressed air transferred from the propulsion engine and ducts and exhaust ports to deliver compressed air from the propulsion engine to the RBC. Compressed air is transferred from a low pressure (LP) compressor or a high pressure (HP) compressor depending on the thrust setting of the engine. Under low thrust settings, compressed air is usually transferred from the HP compressor because the pressure is too low to be transferred from the LP compressor. Under high thrust settings, compressed air is usually transferred from the LP compressor to reduce the harmful effects of the high pressure / high temperature air discharged from the propulsion engine thermodynamic cycle. In some embodiments, compressed air comes from a fan and is further energized for use with the ECS.

[0003] However, whether the air is transferred from a fan, LP compressor, or HP compressor, the air is typically at too high a pressure and temperature to be transferred to the RBC without further decompression and cooling through a heat exchanger system and pressure regulator before reaching the RBC for use in the ECS. Thus, a large amount of useful energy is wasted by using compressed air from the propulsion engines to power the RBC.

[0004] Furthermore, diverting compressed air from the propulsion engine interrupts the propulsion engine's thermodynamic cycle, thereby reducing its efficiency, as energy from the compressed air is diverted away from the propulsion engine. Furthermore, the pressure regulators and heat exchangers used to remove energy from the compressed air before it reaches the RBC add weight and complexity to the propulsion engine and aircraft, thereby further reducing engine and aircraft efficiency.

[0005] Therefore, a need exists to provide energy for a reverse Brayton cycle of a thermal management system while reducing deleterious effects on the efficiency of the propulsion engine and the aircraft. Summary of the invention

[0006] 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.

[0007] An aircraft power generation system comprising a reverse Brayton cycle system, a gas turbine engine and a gearbox (sometimes also referred to as a gearbox) is generally provided. The gas turbine engine includes a compressor section, a turbine section and an engine shaft. The compressor section and the turbine section are arranged in a series flow arrangement. The engine shaft can rotate with at least a portion of the compressor section and with at least a portion of the turbine section. The reverse Brayton cycle system includes a compressor, a drive shaft, a turbine, and a first exchanger. The drive shaft can rotate with the compressor or the turbine, and the compressor, the first heat exchanger, and the turbine are in a series flow arrangement. The gearbox is configured to receive mechanical energy from the engine shaft and transfer the mechanical energy to the reverse Brayton cycle through the drive shaft.

[0008] 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. BRIEF DESCRIPTION OF THE DRAWINGS

[0009] A full and enabling disclosure of the present invention, including the best mode thereof, to one of ordinary skill in the art is set forth in the specification with reference to the accompanying drawings, in which:

[0010] Figure 1 is a schematic cross-sectional view of an exemplary gas turbine engine according to various embodiments of the present subject matter;

[0011] Figure 2 is a schematic diagram of an exemplary aircraft system including a gas turbine engine, a gearbox, a reverse Brayton cycle, an electrical powertrain, and a propulsion device;

[0012] Figure 3 is a schematic diagram of another exemplary aircraft system including a gas turbine engine, a gearbox, a reverse Brayton cycle, an electric powertrain, and a propulsion device;

[0013] Figure 4 is a schematic diagram of an exemplary aircraft system including a gas turbine engine, a gearbox, a reverse Brayton cycle, a second gearbox, and a propulsion device;

[0014] Figure 5is a schematic diagram of another exemplary aircraft system including a gas turbine engine, a gearbox, a reverse Brayton cycle, a second gearbox, and a propulsion device;

[0015] Figure 6 is a schematic diagram of an exemplary aircraft system including a gas turbine engine, a gearbox, a reverse Brayton cycle, and a gear drive system;

[0016] Figure 7 is a schematic diagram of an exemplary aircraft system including a gas turbine engine, a gearbox, a reverse Brayton cycle, a second gearbox, and a gear drive system; and

[0017] Figure 8 is a schematic diagram of an exemplary aircraft system including a gas turbine engine, a reverse Brayton cycle, and an environmental control system.

[0018] Repeat use of reference characters in the present specification and drawings is intended to represent same or analogous features or elements of the invention. DETAILED DESCRIPTION

[0019] Reference will now be made in detail to embodiments of the present invention, one or more examples of which are illustrated in the accompanying drawings. Each embodiment is provided by way of explaining the present invention, rather than limiting the present invention. In fact, it will be apparent to those skilled in the art that various modifications and variations may be made in the present invention without departing from the scope or spirit of the present invention. For example, a feature shown or described as a part of one embodiment may be used together with another embodiment to produce yet another embodiment. Therefore, it is intended that the present invention covers such modifications and variations within the scope of the appended claims and their equivalents.

[0020] As used herein, the terms “first,” “second,” and “third” are used interchangeably to distinguish one component from another and are not intended to indicate the position and importance of the individual components.

[0021] The terms "upstream" and "downstream" refer to relative directions with respect to the flow of a fluid in a fluid path. For example, "upstream" refers to the direction from which a fluid is flowing, and "downstream" refers to the direction to which a fluid is flowing.

[0022] A system is generally provided for powering a reverse Brayton cycle (RBC) by mechanically coupling a shaft of a gas turbine engine to the reverse Brayton cycle with the aid of a gearbox. The RBC is operated by mechanical energy from the gas turbine engine. A shaft of the engine (e.g., a low pressure, intermediate pressure, or high pressure shaft) is mechanically coupled to the gearbox and transfers mechanical energy from the engine shaft to the gearbox. The gearbox outputs mechanical energy from the engine shaft to the RBC via a drive shaft that is mechanically coupled to the RBC and rotatable relative to the RBC. The gearbox outputs the same or proportional or constant rotational speed relative to the engine shaft drive shaft, so that the RBC rotational speed is operable to operate closer to its optimal efficiency independent of the engine thrust setting.

[0023] In a closed loop configuration, the RBC operates using any compressible fluid as the working fluid. The working fluid is in thermal communication with the air entering the ECS to achieve the desired temperature for the system. In an open loop configuration, the RBC obtains outside air as the working fluid, which ultimately becomes conditioned air for the cabin, storage, and avionics.

[0024] In an open loop configuration, using mechanical energy to power the RBC rather than pneumatically obtaining energy from compressed air from the engine can provide a working fluid at pressure and temperature conditions more suitable for a thermal management system (TMS). The mechanical operation of the RBC allows the RBC to operate based on the requirements of the TMS rather than based on the output of the engine. In addition, the RBC can be scaled (sometimes also referred to as proportional increase or decrease) in size to more accurately meet the needs of the TMS and other aircraft energy management systems, including systems such as environmental control systems (ECS), turbine active clearance control (ACC), aircraft anti-icing systems, cabin anti-icing systems, hydraulic and pneumatic pumps, fluid heat exchangers, and aircraft electronic and electrical systems.

[0025] In addition, mechanically driven RBCs are more exergy-efficient (sometimes also referred to as reflex-capacity-efficient) than pneumatically driven RBCs using compressed air from an engine. In addition to analyzing energy efficiency, where energy is the sum of useful energy (exergy) and useless energy (anergy), RBCs driven by mechanical energy from an engine shaft are more exergy-efficient by limiting the inefficiency caused by the increase in ineffective energy due to thermodynamics and energy conversion. Thermodynamic losses are generated by removing the exergy from the propulsion engine in the form of high-pressure / high-temperature compressed air. In addition, thermodynamic losses are derived from the high pressure and high temperature of the compressed air being processed into exergy by a heat exchanger and a pressure regulator to drive the ineffective energy of the RBC. By mechanically coupling the engine to the RBC, when used to drive the RBC, the exergy in the form of high-pressure / high-temperature compressed air in the engine is maintained rather than being considered as ineffective energy.

[0026] Now referring to the accompanying drawings, Figure 1 is a schematic cross-sectional view of a gas turbine engine according to an exemplary embodiment of the present disclosure. More specifically, Figure 1 In the embodiment of the present invention, the gas turbine engine is a high bypass turbofan jet engine 10, referred to herein as a "turbofan engine 10". Although further described below generally with reference to the turbofan engine 10, the present disclosure is also applicable to turbomachinery, generally including turbojets, turboprops and turboshaft turbine engines, including industrial and marine turbine engines and auxiliary power units. Figure 1 As shown in , the turbofan engine 10 defines an axial direction A (extending parallel to a longitudinal centerline 12 provided for reference) and a radial direction R. Generally, the turbofan 10 includes a fan section 14 and a core turbine engine 16 disposed downstream of the fan section 14 .

[0027] The depicted exemplary core turbine engine 16 generally includes a generally tubular outer casing 18 defining an annular inlet 20. The outer casing 18 surrounds a compressor section 21, which includes a supercharger or low pressure (LP) compressor 22 and a high pressure (HP) compressor 24, a combustion section 26, a turbine section 31, which includes a high pressure (HP) turbine 28 and a low pressure (LP) turbine 30, and a jet exhaust nozzle section 32 in series flow relationship. A high pressure (HP) shaft or spool 34 drivingly connects the HP turbine 28 to the HP compressor 24. A low pressure (LP) shaft or spool 36 drivingly connects the LP turbine 30 to the LP compressor 22. The compressor section 21, the combustion section 26, the turbine section 31, and the nozzle section 32 together define a core air flow path 37.

[0028] For the depicted embodiment, the fan section 14 includes a variable pitch fan 38 having a plurality of fan blades 40 coupled to a disk 42 in a spaced apart manner. As depicted, the fan blades 40 generally extend outwardly from the disk 42 in a radial direction R. Each fan blade 40 is rotatable relative to the disk 42 about a pitch axis P by means of a fan blade 40 operably coupled to a suitable actuation member 44 configured to collectively and uniformly change the pitch of the fan blade 40. The fan blades 40, disk 42, and actuation member 44 are rotatable together about the longitudinal axis 12 via the LP shaft 36 across a power gearbox 46. The power gearbox 46 includes a plurality of gears for adjusting the rotational speed of the fan 38 relative to the LP shaft to a more efficient rotational fan speed.

[0029] The power gearbox 46 may include an epicyclical gear train including at least two gear types, wherein the sun gear is located at the center of the gear train and the planet gears are in an epicyclic relationship relative to the sun gear. In other words, when the sun gear rotates on its fixed axis, the center of at least one planet gear revolves around the center of the sun gear. The planet gears are supported by a carrier, which can help transfer torque from the sun gear to the planet gears. The planet gears are surrounded by a fixed ring gear, wherein the teeth of the planet gears ride on the ring gear and the sun gear. In other embodiments, the second planetary gear set is radially placed between the first planetary gear set and the ring gear. In another embodiment, the epicyclic gear train has a star gear configuration, wherein the center of each planet gear is fixed so that the planet gears rotate on a fixed axis relative to the sun gear. The enclosed ring gear rotates and transfers torque from the input power transmitted to the sun gear. The carrier acts as a spacer between the sun gear and the planet gears and fixes the axis of each planet gear.

[0030] Still reference Figure 1 In an exemplary embodiment of the present invention, the disk 42 is covered by a rotatable forward hub 48 that is aerodynamically contoured to lift airflow through the plurality of fan blades 40. In addition, the exemplary fan section 14 includes an annular fan case or outer nacelle 50 that circumferentially surrounds the fan 38 and / or at least a portion of the core turbine engine 16. It should be appreciated that the nacelle 50 may be configured to be supported relative to the core turbine engine 16 by a plurality of circumferentially spaced outlet guide vanes 52. In addition, a downstream section 54 of the nacelle 50 may extend over an outer portion of the core turbine engine 16 to define a bypass airflow passage 56 therebetween.

[0031] During operation of the turbofan engine 10, a volume of air 58 enters the turbofan 10 and / or fan section 14 through an associated inlet 60 of the nacelle 50. As the volume of air 58 passes over the fan blades 40, a first portion of the air 58, as indicated by arrows 62, is directed or directed into the bypass airflow passage 56, and a second portion of the air 58, as indicated by arrows 64, is directed or directed into the core air flow path 37, or more specifically into the LP compressor 22. The ratio between the first portion 62 of air and the second portion 64 of air is generally referred to as the bypass ratio. Then, when the second portion 64 of air is directed through the high pressure (HP) compressor 24 and into the combustion section 26 (where it is mixed with fuel and combusted to provide combustion gases 66), the pressure of the second portion 64 of air increases.

[0032] The combustion gases 66 are directed through the HP turbine 28, where a portion of the thermal and / or kinetic energy from the combustion gases 66 is extracted via successive stages of HP turbine stator vanes 68 coupled to the outer casing 18 and HP turbine rotor blades 70 coupled to the HP shaft or spool 34, thereby causing the HP shaft or spool 34 to rotate, thereby supporting operation of the HP compressor 24. The combustion gases 66 are then directed through the LP turbine 30, where a second portion of the thermal and kinetic energy is extracted from the combustion gases 66 via successive stages of LP turbine stator vanes 72 coupled to the outer casing 18 and LP turbine rotor blades 74 coupled to the LP shaft or spool 36, thereby causing the LP shaft or spool 36 to rotate, thereby supporting operation of the LP compressor 22 and / or rotation of the fan 38.

[0033] The combustion gases 66 are then directed through the jet exhaust nozzle section 32 of the core turbine engine 16 to provide propulsive thrust. At the same time, the pressure of the first portion of air 62 is substantially increased as the first portion of air 62 is directed through the bypass airflow passage 56 before it is discharged from the fan nozzle exhaust section 76 of the turbofan 10, also providing propulsive thrust. The HP turbine 28, the LP turbine 30, and the jet exhaust nozzle section 32 at least partially define a hot gas path 78 for directing the combustion gases 66 through the core turbine engine 16.

[0034] The system of gears 162 is mechanically coupled to the HP shaft 34 and the tower shaft 160, and transmits mechanical energy from the rotation of the HP shaft 34 through the tower shaft 160 to at least one accessory gearbox (AGB) 45. In addition, the gears 162 transform the direction of transmission of mechanical energy, such as to an angle perpendicular or parallel to the centerline 12, or to an angle oblique to the centerline 12. Suitable types of gears 162 may include bevel gears (such as, but not limited to, helical line gears, zero-degree bevel gears, hypoid gears, or miter gears), helical gears, spur gears, ring gears, helical gears, or worm gears, or combinations thereof. The type of gear 162 employed depends on the angle of the tower shaft 160 relative to the HP shaft 34 and the output speed and torque transmission from the engine shaft 35 to the tower shaft 160. The tower shaft 160 extends at an oblique angle to the HP shaft 34, or in another embodiment, may extend substantially perpendicularly (in the radial direction R) from the HP shaft 34. The tower shaft 160 extends radially outward (in the radial direction R) through the compressor center casing vanes 103 in the core flow path 37 and through the outer casing 18 .

[0035] In one embodiment, the AGB 45 is positioned within the outer casing 18 and includes gears that receive mechanical energy from the tower shaft 160 and output mechanical energy to drive auxiliary equipment for the engine 10 and the aircraft, such as fuel and oil pumps, hydraulic pumps, fluid coolers, power generators, and fuel and oil heat exchangers. In one embodiment, the AGB 45 is also mechanically coupled to a starter motor that provides mechanical energy to initially drive the HP shaft 34 and rotate the HP compressor 24 to generate sufficient air flow 64 before introducing fuel and igniting the fuel / air mixture in the combustion section 26. The generated combustion gases 66 flow into the HP turbine 28, thereby rotating the HP shaft 34 and thus supporting the operation of the gas turbine engine 10 as previously mentioned. In another embodiment, the AGB 45 is positioned radially outward (in the direction R) of the outer casing 18 or radially outboard of the fan case 50. In yet another embodiment, the AGB 45 is positioned on an aircraft wing or pylon (sometimes also referred to as a boom).

[0036] Figure 2A schematic diagram of an embodiment including a gas turbine engine 10, a gearbox 47, and a reverse Brayton cycle (RBC) system 104 is shown. The engine 10 includes a compressor section 21, a combustion section 26, and a turbine section 31 arranged in series flow. The engine 10 includes an engine shaft 35 that can rotate with at least a portion of the compressor section 21 and with at least a portion of the turbine section 31. A portion of the compressor section 21 includes at least one compressor, such as an LP compressor 22, or an HP compressor 24, or an intermediate compressor. A portion of the turbine section 31 includes at least one turbine, such as an LP turbine 28, or an HP turbine 30, or an intermediate turbine. In one embodiment, the engine shaft 35 is an HP shaft 34. In another embodiment, the engine shaft 35 is an LP shaft 36. In yet another embodiment, the engine shaft 35 is an intermediate shaft.

[0037] Gearbox 47 is configured to receive mechanical energy from engine shaft 35 and transfer the mechanical energy to drive shaft 168 to operate RBC 104. Drive shaft 168 is mechanically coupled to gearbox 47, compressor 110, and turbine 112 of the RBC, and is rotatable with RBC 104.

[0038] Still reference Figure 2 , the RBC 104 is in an operable closed loop configuration in which a working fluid 122 circulates from the compressor 110 through the first heat exchanger 106 and then through the turbine 112 and through the second heat exchanger 107 and back to the compressor 110. In the operable closed loop configuration, the working fluid 122 is any compressible fluid including, as non-limiting examples, carbon dioxide, oxygen, nitrogen, air, or mixtures thereof.

[0039] The working fluid 122 circulates through the compressor 110 and exits the compressor 110 as a high pressure, high temperature (HP / HT) fluid 130. The HP / HT fluid 130 enters the first heat exchanger 106 and exits as a high pressure, low temperature (HP / T1) fluid 131. The HP / T1 fluid 131 enters the turbine 112 and expands and exits as a low pressure, low temperature (LP / T2) fluid 132. The LP / T2 fluid 132 enters the second heat exchanger 107 and exits as a low pressure, low temperature (LP / T3) fluid 133. When the RBC 104 and the working fluid 122 are in a closed loop configuration, the LP / T3 fluid 133 enters the compressor 110 and the aforementioned cycle is repeated.

[0040] As the closed loop system experiences leaks or losses, the closed loop system may be operated to periodically "recharge" the RBC 104 with the working fluid 122. In one embodiment, the recharging may also regulate the working fluid 122 pressure and adjust the performance of the RBC 104 based on inputs from and outputs to the control system. In one embodiment of a closed loop configuration, where the working fluid 122 is air, a bleed valve from the compressor section 21 of the engine 10 is operated to provide compressed air 64 from the compressor section 21 of the engine 10 to the compressor 110. In other embodiments where the working fluid 122 is a compressible fluid other than air, the system is recharged by a source of carbon dioxide, nitrogen, or oxygen, as non-limiting examples, rather than by air 64 from the compressor section 21 of the engine 10.

[0041] In one embodiment, the heat exchangers 106, 107 of the RBC 104 are in thermal communication with a thermal management system (TMS) 108. The TMS 108 includes an environmental control system (ECS). The ECS typically includes dehumidifiers, water extractors, sensors, pressure and temperature regulators, pressurization and distribution systems, and air, which is conditioned by the TMS 108 for use in aircraft cabins, storage devices, avionics, and anti-icing systems. In other embodiments, the TMS 108 may also include thermal management of fuel, oil, hydraulic and pneumatic fluids, electronic and electrical systems, emergency systems, and thermal communication with propulsion units and auxiliary power units (APU) subsystems. The air used for the TMS 108 can be external atmospheric air or compressed air 64 from the compressor section 21 of the engine 10. The external air ranges in temperature from those temperatures on the ground (e.g., up to about 60°C) to those temperatures at an altitude of about 16,200 meters (e.g., about -58°C). The outside air is placed in thermal communication with the heat exchangers 106 , 107 until the temperature is suitable for use by the TMS 108 .

[0042] In one embodiment, the gearbox 47 is configured to output an operable constant rotational speed to the drive shaft 168 such that the RBC 104 operates at a constant rotational speed that is functionally independent of the rotational speed of the motor shaft 35. In other words, after the motor shaft 35 rotational speed is greater than a certain threshold, the drive shaft 168 rotational speed is different from the motor shaft 35 rotational speed, regardless of the magnitude or change in the rotational speed of the motor shaft 35.

[0043] However, in another embodiment, the gearbox 47 is configured to output a proportional rotational speed to the drive shaft 168 so that the RBC 104 operates at a fraction of the rotational speed of the motor shaft 35. For example, the drive shaft 168 rotational speed may be, for example, half, or one-third, or one-quarter, or one-sixth, or one-tenth, etc., of the rotational speed of the motor shaft 35.

[0044] In yet another embodiment, the gearbox 47 can be configured to output a combination of a proportional rotational speed and an operable constant rotational speed to the drive shaft 168 relative to the rotational speed of the engine shaft 35. For example, the power output state of the engine 10 can cover from zero to ground idle (GIDLE), to flight idle (FIDLE), and to maximum take-off (MTO) state in increasing order of thrust output, wherein MTO is considered to be the maximum power output of the engine 10 and GIDLE is considered to be the minimum power output of the engine 10. The rotational speeds of the drive shaft 168 and the engine shaft 35 are approximately proportional or the same from zero to GIDLE. From state GIDLE to MTO, the drive shaft 168 has a different and generally constant rotational speed than the engine shaft 35, independent of the change in the rotational speed of the engine shaft 35 between the GIDLE and MTO states. Alternatively, in another embodiment, the rotational speeds of the drive shaft 168 and the engine shaft 35 are approximately proportional or the same from zero to FIDLE. From state FIDLE to MTO, the drive shaft 168 has a rotational speed that is different and generally constant from the engine shaft 35, independent of the change in the rotational speed of the engine shaft 35 between the FIDLE and MTO states. In yet another embodiment, between GIDLE and FIDLE, the drive shaft 168 rotational speed is generally constant at a first value and different from the rotational speed of the engine shaft 35. Between the FIDLE and MTO states, the drive shaft 168 rotational speed is generally constant at a second value and different from the rotational speed of the engine shaft 35.

[0045] In yet another embodiment, the gearbox 47 is configured to output a variable rotational speed to the drive shaft 168 relative to the rotational speed of the engine shaft 35. The gearbox 47 may include a controller and communications means to receive operating inputs, communicate load requirements, and govern changes in the rotational speed of the drive shaft 168. For example, where the engine 10 may be operated at a constant rotational speed between FIDLE and MTO during aircraft cruising, the gearbox 47 may output a variable rotational speed to the drive shaft 168 based on the thermal load requirements for the TMS 108.

[0046] Still reference Figure 2In the embodiment shown in FIG. 1 , the gearbox 47 is mechanically coupled to the electric power train 156 in a parallel arrangement. The electric power train 156 converts the mechanical energy transmitted from the gearbox 47 into electrical energy to drive the propulsion device 170. In one embodiment, the propulsion device 170 is the fan section 14 of the turbofan engine 10. In another embodiment, the propulsion device 170 is a propeller of a turboprop or turboshaft engine. In still other embodiments, the electric power train 156 may also include a power conditioner (e.g., a rectifier or transformer) and a communication system to distribute electrical energy to aircraft systems other than the propulsion device 170, such as avionics, lighting, emergency systems, hydraulic and pneumatic pumps, actuators, and sensors.

[0047] In other embodiments of the gearbox 47, where the gearbox 47 transfers mechanical energy from the engine 10 to the RBC 104 and another parallel-arranged system, the gearbox 47 can be configured to output different rotational speeds to the RBC 104 and another parallel-arranged system. For example, the gearbox 47 can be configured to output an operational constant rotational speed to the RBC 104, while the gearbox 47 is also configured to output a proportional rotational speed (relative to the engine shaft 35) to the electric powertrain 156.

[0048] In one embodiment, the gearbox 47 is configured as a power gearbox 46, wherein the power gearbox 46 is mechanically coupled to the LP shaft 36 and is also configured to transfer mechanical energy to the RBC 104. In another embodiment, the gearbox 47 is configured as an AGB 45, wherein the AGB 45 is mechanically coupled to the HP shaft 34 and is also configured to transfer mechanical energy to the RBC 104. In yet another embodiment, the gearbox 47 is separate from the AGB 45 or the power gearbox 46, wherein the gearbox 47 is configured to transfer mechanical energy from the HP shaft 36, the LP shaft 34, or an intermediate shaft to the RBC 104 via a drive shaft 168 that is mechanically coupled to the RBC 104 and rotatable relative to the RBC 104.

[0049] Figure 3 An embodiment is shown in which the gearbox 47, RBC 104, electric powertrain 156, and propulsion device 170 are arranged in series. Drive shaft 168 is mechanically coupled to RBC 104 and can rotate relative to RBC 104 and transmit mechanical energy to electric powertrain 156. Electric powertrain 156 is coupled to propulsion device 170. Electric powertrain 156 can be configured to output a rotational speed to propulsion device 170 that is independent of the rotational speed of drive shaft 168 or engine shaft 35. Figure 3 As shown in , the TMS 108 can be configured to be in thermal communication with the electric power train 156. The TMS 108 can also be configured to communicate the RBC 104 output according to the thermal requirements for the electric power train 156.

[0050] Figure 4 and Figure 5 Each shows an embodiment in which the propulsion device 170 is mechanically driven by a second gearbox 172 arranged in parallel or in series, respectively. The second gearbox 172 is configured to receive mechanical energy from the gearbox 47. Similar to the gearbox 47, the second gearbox 172 can be configured to output the same rotational speed as the engine shaft 35, or a rotational speed proportional to the engine shaft 35, or an operable constant drive speed, or a variable drive speed, or a combination thereof to the propulsion device 170.

[0051] Figure 6 An embodiment is shown in which gearbox 47 transfers mechanical energy from engine 10 to a gear drive system 174 arranged in parallel with drive shaft 168 and RBC 104. Gear drive system 174 may include, for example, auxiliary equipment such as, but not limited to, a hydraulic pump, a pneumatic pump, or a heat exchanger.

[0052] Figure 7 An embodiment is shown in which a gear drive system 174 is arranged in series with the gearbox 47, the RBC 104, and the second gearbox 172. The gearbox 47 transfers mechanical energy from the engine 10 to the drive shaft 168, where the drive shaft 168 is mechanically coupled to the RBC 104 and the second gearbox 172. Similar to the gearbox 47, the second gearbox 172 can output the same rotational speed as the drive shaft 168 or the engine shaft 35, or output a proportional rotational speed, or a constant drive speed, or a combination thereof to the gear drive system 174. In another embodiment, the second gearbox 172 is an AGB 45.

[0053] Figure 8 A schematic diagram of an embodiment including a gas turbine engine 10 and an RBC 104 is shown, wherein the RBC 104 is in an open loop configuration and in fluid communication with a TMS 108. In the open loop configuration, the working fluid 122 in the RBC 104 is outside air 118. When the aircraft is at altitude or on the ground, the outside air 118 is received from atmospheric conditions through one or more plenums in the aircraft. More specifically, "at altitude" refers to all altitudes except on the ground, such as, for example, once the aircraft has taken off.

[0054] exist Figure 8In the embodiment shown in FIG. 1 , the working fluid 122 enters the compressor 110 and is compressed into a high pressure, high temperature (HP / HT) fluid 130. The HP / HT fluid 130 enters the first heat exchanger 106 and becomes a high pressure, low temperature (HP / T1) fluid 131. The HP / T1 fluid 131 enters the turbine 112, where the fluid 131 expands, thereby reducing the pressure and temperature, and becomes a low pressure, low temperature (LP / T2) fluid 132. The LP / T2 fluid 132 enters the second heat exchanger 107 and exits as a low pressure, low temperature (LP / T3) fluid 133. The HP / HT fluid 130 may be partially directed to the LP / T3 fluid 133 and mixed with the LP / T3 fluid 133. The LP / T3 fluid 133 or its mixture with the HP / HT fluid 130 is regulated by the control valve 124. The LP / T3 fluid 133 or its mixture with the HP / HT fluid 130 is directed to the TMS 108 as the ECS air 123. Furthermore, the HP / HT fluid 130 exiting the compressor 110 may be directed in part to other systems of the TMS 108 , such as, but not limited to, an aircraft wing anti-icing system or a nacelle anti-icing system.

[0055] In other embodiments, the second heat exchanger 107 may be excluded and only a single heat exchanger 106 may be employed, depending on the heat exchanger performance and specific system requirements relative to the size and type of aircraft. In still other embodiments, the ECS air 123 is a mixture of the LP / LT fluid 133 and the HP / T1 fluid 131. Furthermore, other embodiments do not include the second heat exchanger 107, and the LP / T2 fluid 132 is either the ECS air 123 or mixed with the HP / HT fluid 130 or the HP / T1 fluid 131.

[0056] like Figure 8 As shown in , a portion of the working fluid 117 from the RBC working fluid 122 can be directed to the turbine section 31 of the engine 10 as cooling air for the active clearance control (ACC) system 151. The portion of the working fluid 117 that enters the ACC 151 is regulated by the ACC valve 116. The portion of the working fluid 117 in the ACC 151 that goes onto the turbine shell reduces the radial distance between the turbine shell and the turbine blades during engine operation. The combustion gases 66 that pass in the area between the blade tip and the shell do not substantially drive the turbine blades. Therefore, the radial clearance between the turbine shell and the turbine blades is a source of engine inefficiency.

[0057] As the engine operates and increases in power, a combination of factors contributes to the radial distance between the turbine blade tips and the surrounding turbine casing during engine operation ( Figure 1These factors include, but are not limited to, centrifugal forces on the turbine blades and disks, the effects of hot combustion gases 66 radially expanding the turbine blades and disks, and pressure changes on the turbine casing due to the combustion gases 66. The ACC 151 operates by directing cooler fluid from the compressor 110 into thermal communication with the hot outer casing 18 at the turbine section 31, thereby causing the hot outer casing 18 at the turbine section 31 to contract toward the turbine blades and reduce the radial area between the blades and the casing.

[0058] The present disclosure may increase gas turbine engine and aircraft efficiency by providing cooling capabilities without extracting bleed air (sometimes referred to as bleed air) from engine 10 and thereby avoiding inefficiencies due to interruption of the thermodynamic cycle of engine 10. Additionally, the present disclosure avoids inefficiencies due to power conversion losses in converting mechanical energy to electrical energy and back to mechanical energy to drive a mechanical system (e.g., a reverse Brayton cycle system).

[0059] 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 language of the claims.

Claims

1. An aircraft power generation system, comprising: a gas turbine engine comprising a compressor section, a turbine section, and an engine shaft, the compressor section being arranged in a series flow arrangement with the turbine section, and the engine shaft being rotatable with at least a portion of the compressor section and with at least a portion of the turbine section; a reverse Brayton cycle system comprising a compressor, a drive shaft, a turbine, a first heat exchanger, and a second heat exchanger downstream of the turbine and upstream of the turbine section, the second heat exchanger receiving fluid from the turbine, and the drive shaft being rotatable with the compressor or the turbine, and wherein the compressor, the first heat exchanger, the second heat exchanger, and the turbine are in a series flow arrangement; a gearbox, wherein the gearbox is configured to receive mechanical energy from the engine shaft and transmit the mechanical energy to the reverse Brayton cycle system through the drive shaft; and An electric powertrain is operably coupled in series with the drive shaft on a side of the compressor and the turbine opposite the gearbox.

2. The system according to claim 1, further comprising: Thermal management system; and A working fluid, wherein the working fluid is in the reverse Brayton cycle system, and wherein the working fluid is in fluid communication with the thermal management system.

3. The system according to claim 2, wherein: The thermal management system further includes an environmental control system, and wherein the environmental control system is in fluid communication with the reverse Brayton cycle system.

4. The system according to claim 1, further comprising: A second heat exchanger, wherein the first heat exchanger interacts with the working fluid upstream of the turbine of the reverse Brayton cycle system and the second heat exchanger interacts with the working fluid downstream of the turbine of the reverse Brayton cycle system.

5. The system according to claim 4, wherein: A first portion of the working fluid is directed through the first heat exchanger, and a second portion of the working fluid bypasses the first heat exchanger.

6. The system according to claim 5, wherein: The first portion of the working fluid and the second portion of the working fluid mix with each other downstream of the second heat exchanger.

7. The system according to claim 4, further comprising: An active clearance control system is operably coupled to a turbine section of the gas turbine engine.

8. The system according to claim 7, further comprising: An active clearance control valve is positioned downstream of the second heat exchanger and upstream of a turbine section of the gas turbine engine.

9. The system according to claim 7, wherein: The working fluid in the active clearance control system reduces a radial distance between a turbine casing and turbine blades during engine operation.

10. The system according to claim 1, wherein: The reverse Brayton cycle system further includes a working fluid, and wherein the working fluid is a compressible fluid.

11. The system of claim 1 , further comprising: A propulsion device, wherein the propulsion device is coupled to the electric powertrain.

12. A method of operating an engine, the method comprising: compressing the working fluid into a fluid at a first pressure and a first temperature in the compressor; flowing the working fluid at the first pressure and the first temperature through a first heat exchanger, causing the working fluid to retain the first pressure and second temperature, wherein the second temperature is lower than the first temperature; expanding the working fluid at the first pressure and the second temperature through a turbine, thereby changing the working fluid to a second pressure and a third temperature, wherein the second pressure is lower than the first pressure and the third temperature is lower than the second temperature; flowing the working fluid at the second pressure and the third temperature through a second heat exchanger to change the working fluid to the second pressure and a fourth temperature, wherein the fourth temperature is lower than the third temperature; routing the working fluid at the second pressure and the fourth temperature to a thermal management system; and A drive shaft is rotated, the drive shaft operably coupled to an electric powertrain operably connected in series with the compressor and the turbine on an opposite side of the compressor and the turbine from the gearbox.

13. The method according to claim 12, wherein: The working fluid is outside air received through one or several air chambers in the aircraft.

14. The method according to claim 12, further comprising: A portion of the working fluid at the second pressure and the fourth temperature is directed to a turbine section of the engine as cooling air for an active clearance control system.

15. The method according to claim 14, further comprising: The working fluid at the second pressure and the fourth temperature is regulated to a turbine section of the engine by actuation of an active clearance control valve.

16. An aircraft power generation system: a gas turbine engine comprising a compressor section, a turbine section, and an engine shaft, the compressor section being arranged in a series flow arrangement with the turbine section, and the engine shaft being rotatable with at least a portion of the compressor section and with at least a portion of the turbine section; a reverse Brayton cycle system comprising a compressor, a drive shaft, a turbine, and a first heat exchanger, wherein the drive shaft is rotatable with the compressor or the turbine, and the compressor, the first heat exchanger, and the turbine are in a series flow arrangement; a plenum configured to direct external air to a compressor section of the gas turbine engine; a gearbox, wherein the gearbox is configured to receive mechanical energy from the engine shaft and transmit the mechanical energy to the reverse Brayton cycle system through the drive shaft; and A working fluid circuit operably connected to the reverse Brayton cycle system and the turbine section, the working fluid circuit being configured to deliver a first portion of the working fluid to an active clearance control valve positioned upstream of the turbine section of the gas turbine engine and to deliver the first portion of the working fluid and a second portion of the working fluid to a thermal management system, the first portion of the working fluid having a first pressure and a first temperature and the second portion of the working fluid having a second pressure and a second temperature, the first pressure being less than the second pressure and the first temperature being less than the second temperature.

17. The aircraft power generation system of claim 16, further comprising: A thermal management system is downstream of a compressor of the reverse Brayton cycle system.

18. The aircraft power generation system of claim 17, further comprising: A control valve is positioned upstream of the thermal management system and downstream of a compressor of the reverse Brayton cycle system.

19. The aircraft power generation system of claim 16, further comprising: A second heat exchanger, wherein the first heat exchanger interacts with the working fluid upstream of the turbine of the reverse Brayton cycle system and the second heat exchanger interacts with the working fluid downstream of the turbine of the reverse Brayton cycle system.

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