Patent Application
20230250754
The present disclosure relates generally to turbine engines and aircraft engines, and more specifically to turbo expander systems for use when employing hydrogen fuel systems and related systems with turbine and aircraft engines.
Gas turbine engines, such as those utilized in commercial and military aircraft, include a compressor section that compresses air, a combustor section in which the compressed air is mixed with a fuel and ignited, and a turbine section across which the resultant combustion products are expanded. The expansion of the combustion products drives the turbine section to rotate. As the turbine section is connected to the compressor section via a shaft, the rotation of the turbine section drives the compressor section to rotate. In some configurations, a fan is also connected to the shaft and is driven to rotate via rotation of the turbine.
Typically, hydrocarbon-based fuel is employed for combustion onboard an aircraft, in the gas turbine engine. The liquid fuel has conventionally been a hydrocarbon-based fuel. Alternative fuels have been considered, but suffer from various challenges for implementation, particularly on aircraft. Hydrogen-based and/or methane-based fuels are viable effective alternatives which may not generate the same combustion byproducts as conventional hydrocarbon-based fuels. The use of hydrogen and/or methane, as a gas turbine fuel source, may require very high efficiency propulsion, in order to keep the volume of the fuel low enough to feasibly carry on an aircraft. That is, because of the added weight associated with such liquid/compressed/supercritical fuels, such as related to vessels/containers and the amount (volume) of fuel required, improved efficiencies associated with operation of the gas turbine engine may be necessary.
According to some embodiments, aircraft propulsion systems are provided The aircraft propulsion systems include an aircraft system comprising at least one hydrogen tank and an aircraft-system heat exchanger, and an engine system comprising at least a main engine core, a high pressure pump, a hydrogen-air heat exchanger, and a turbo expander assembly, wherein the main engine core comprises a compressor section, a combustor section having a burner, and a turbine section. Fuel is supplied from the at least one fuel tank through a fuel flow path, passing through the aircraft-system heat exchanger, the high pressure pump, the hydrogen-air heat exchanger, and selectively through the turbo expander assembly, prior to being injected into the burner for combustion. The turbo expander assembly is operably coupled to at least two load sources through a selective coupler and configured to selectively drive operation of the at least two load sources.
In addition to one or more of the features described above, or as an alternative, embodiments of the aircraft propulsion systems may include that the turbo expander assembly consists of a single turbo expander operably coupled to the at least two load sources and the selective coupler comprises a clutch assembly configured to selectively connect the turbo expander to a first load source and a second load source of the at least two load sources.
In addition to one or more of the features described above, or as an alternative, embodiments of the aircraft propulsion systems may include that the first load source is continuously driven by the turbo expander and the second load source is decouplable from the turbo expander by a clutch.
In addition to one or more of the features described above, or as an alternative, embodiments of the aircraft propulsion systems may include that the turbo expander assembly comprises a first turbo expander operably coupled to a first load source and a second turbo expander and operably coupled to a second load source.
In addition to one or more of the features described above, or as an alternative, embodiments of the aircraft propulsion systems may include that the second turbo expander is configured to receive a flow of fuel that is expanded within the second turbo expander and drives operation of the second load source.
In addition to one or more of the features described above, or as an alternative, embodiments of the aircraft propulsion systems may include that the first load source and the second load source are electric generators having a combined power output of at least 300 kW of power.
In addition to one or more of the features described above, or as an alternative, embodiments of the aircraft propulsion systems may include that each of the first load source and the second load source are electric generators each configured to generate at least 300 kW of power.
In addition to one or more of the features described above, or as an alternative, embodiments of the aircraft propulsion systems may include a valve assembly arranged upstream of each of the first turbo expander and the second turbo expander, wherein the valve assembly is configured to selectively direct fuel to each of the first turbo expander and the second turbo expander.
In addition to one or more of the features described above, or as an alternative, embodiments of the aircraft propulsion systems may include a valve assembly arranged upstream of each of the first turbo expander and the second turbo expander, wherein the valve assembly is configured to always supply fuel to the first turbo expander and to selectively direct fuel to the second turbo expander.
In addition to one or more of the features described above, or as an alternative, embodiments of the aircraft propulsion systems may include a first gear box arranged between the first turbo expander and the first load source and a second gear box arranged between the second turbo expander and the second load source.
In addition to one or more of the features described above, or as an alternative, embodiments of the aircraft propulsion systems may include that the at least two load sources are selected from electric generators, pumps, and actuators.
In addition to one or more of the features described above, or as an alternative, embodiments of the aircraft propulsion systems may include that the turbo expander assembly comprises a plurality of turbo expanders and the at least two load sources comprises a number of load sources equal to the number of turbo expanders, wherein each turbo expander is selectively coupled to a respective load source.
In addition to one or more of the features described above, or as an alternative, embodiments of the aircraft propulsion systems may include that the selective coupler is a valve assembly.
In addition to one or more of the features described above, or as an alternative, embodiments of the aircraft propulsion systems may include that the selective coupler is a clutch assembly.
In addition to one or more of the features described above, or as an alternative, embodiments of the aircraft propulsion systems may include that the fuel is hydrogen.
In addition to one or more of the features described above, or as an alternative, embodiments of the aircraft propulsion systems may include a controller operably connected to the selective couple and configuration to control operation thereof.
In addition to one or more of the features described above, or as an alternative, embodiments of the aircraft propulsion systems may include that the turbo expander assembly comprises a first turbo expander operably coupled to a first load source through a first gear box and a second turbo expander operably coupled to a second load source through a second gear box, and the selective coupler comprises a clutch arranged between the second turbo expander and the second load source.
In addition to one or more of the features described above, or as an alternative, embodiments of the aircraft propulsion systems may include a third turbo expander operably coupled to a third load source through a third gear box and the selective coupler comprises a second clutch arranged between the third turbo expander and the third gear box.
In addition to one or more of the features described above, or as an alternative, embodiments of the aircraft propulsion systems may include that a first turbo expander of the turbo expander assembly is selectively coupled to a first load source by a clutch of the selective coupler and a second turbo expander of the turbo expander assembly is selectively coupled to a source of fuel through a valve assembly of the selective coupler.
In addition to one or more of the features described above, or as an alternative, embodiments of the aircraft propulsion systems may include that a first load source of the at least two load sources is an electric generator configured to generate a first power output and a second load source of the at least two load sources is an electric generator configured to generate a second power output that is different from the first power output.
According to some embodiments, methods of operating an aircraft propulsion system are provided. The methods include coupling a turbo expander assembly to at least two load sources through a selective coupler, supplying, from at least one fuel tank, fuel through a fuel flow path comprising passing the fuel through an aircraft-system heat exchanger, a high pressure pump, a fuel-air heat exchanger, and selectively through a turbo expander assembly comprising at least one turbo expander, prior to injecting the fuel into a burner for combustion thereof, wherein the at least one turbo expander is rotationally driven by the fuel passing therethrough, detecting a throughflow of fuel through the turbo expander assembly, and controlling operation of the selective coupler to at least one of drive operation of the at least one turbo expander and selectively control operation of one or more of the operably connected load sources based on the detected throughflow.
In addition to one or more of the features described above, or as an alternative, embodiments of the methods may include that the turbo expander assembly consists of a single turbo expander operably coupled to the at least two load sources and the methods further include controlling the selective coupler to operate a clutch assembly to selectively connect the turbo expander assembly to a first load source and a second load source of the at least two load sources based on the detected throughflow.
In addition to one or more of the features described above, or as an alternative, embodiments of the methods may include that the turbo expander assembly comprises a first turbo expander operably coupled to a first load source of the at least two load sources and a second turbo expander and operably coupled to a second load source of the at least two load sources, and the methods further include controlling the selective coupler to operate a valve assembly to selectively direct fuel to one or both of the first load source and the second load source based on the detected throughflow.
In addition to one or more of the features described above, or as an alternative, embodiments of the methods may include that the turbo expander assembly comprises a first turbo expander operably coupled to a first load source of the at least two load sources and a second turbo expander and operably coupled to a second load source of the at least two load sources, and the methods further include controlling the selective coupler to operate a clutch assembly to selectively couple one or both of the first turbo expander and the second turbo expander to the respective first load source and second load source based on the detected throughflow.
The foregoing features and elements may be executed or utilized in various combinations without exclusivity, unless expressly indicated otherwise. These features and elements as well as the operation thereof will become more apparent in light of the following description and the accompanying drawings. It should be understood, however, that the following description and drawings are intended to be illustrative and explanatory in nature and non-limiting.
The subject matter is particularly pointed out and distinctly claimed at the conclusion of the specification. The foregoing and other features, and advantages of the present disclosure are apparent from the following detailed description taken in conjunction with the accompanying drawings in which:
In this two-spool configuration, the gas turbine engine 20 includes a low speed spool 30 and a high speed spool 32 mounted for rotation about an engine central longitudinal axis A relative to an engine static structure 36 via one or more bearing systems 38. It should be understood that various bearing systems 38 at various locations may be provided, and the location of bearing systems 38 may be varied as appropriate to a particular application and/or engine configuration.
The low speed spool 30 includes an inner shaft 40 that interconnects the fan 42 of the fan section 22, a first (or low) pressure compressor 44, and a first (or low) pressure turbine 46. The inner shaft 40 is connected to the fan 42 through a speed change mechanism, which, in this illustrative gas turbine engine 20, is as a geared architecture 48 to drive the fan 42 at a lower speed than the low speed spool 30. The high speed spool 32 includes an outer shaft 50 that interconnects a second (or high) pressure compressor 52 and a second (or high) pressure turbine 54. A combustor 56 is arranged in the combustor section 26 between the high pressure compressor 52 and the high pressure turbine 54. A mid-turbine frame 57 of the engine static structure 36 is arranged between the high pressure turbine 54 and the low pressure turbine 46. The mid-turbine frame 57 may be configured to support one or more of the bearing systems 38 in the turbine section 28. The inner shaft 40 and the outer shaft 50 are concentric and rotate via the bearing systems 38 about the engine central longitudinal axis A which is collinear with their longitudinal axes.
The core airflow through core airflow path C is compressed by the low pressure compressor 44 then the high pressure compressor 52, mixed and burned with fuel in the combustor 56, then expanded over the high pressure turbine 54 and low pressure turbine 46. The mid-turbine frame 57 includes airfoils 59 (e.g., vanes) which are arranged in the core airflow path C. The turbines 46, 54 rotationally drive the respective low speed spool 30 and high speed spool 32 in response to the expansion of the core airflow. It will be appreciated that each of the positions of the fan section 22, the compressor section 24, the combustor section 26, the turbine section 28, and geared architecture 48 or other fan drive gear system may be varied. For example, in some embodiments, the geared architecture 48 may be located aft of the combustor section 26 or even aft of the turbine section 28, and the fan section 22 may be positioned forward or aft of the location of the geared architecture 48.
The gas turbine engine 20 in one example is a high-bypass geared aircraft engine. In some such examples, the engine 20 has a bypass ratio that is greater than about six (6), with an example embodiment being greater than about ten (10). In some embodiments, the geared architecture 48 is an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3 and the low pressure turbine 46 has a pressure ratio that is greater than about five (5). In one non-limiting embodiment, the bypass ratio of the gas turbine engine 20 is greater than about ten (10:1), a diameter of the fan 42 is significantly larger than that of the low pressure compressor 44, and the low pressure turbine 46 has a pressure ratio that is greater than about five (5:1). The low pressure turbine 46 pressure ratio is pressure measured prior to inlet of low pressure turbine 46 as related to the pressure at the outlet of the low pressure turbine 46 prior to an exhaust nozzle. In some embodiments, the geared architecture 48 may be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3:1. It should be understood, however, that the above parameters are only for example and explanatory of one non-limiting embodiment of a geared architecture engine and that the present disclosure is applicable to other gas turbine engines including turbojets or direct drive turbofans, turboshafts, or turboprops.
A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The fan section 22 of the gas turbine engine 20 is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet (10,668 meters). The flight condition of 0.8 Mach and 35,000 ft (10,668 meters), with the engine at its best fuel consumption—also known as “bucket cruise Thrust Specific Fuel Consumption (‘TSFC’)”—is the industry standard parameter of lbm of fuel being burned divided by lbf of thrust the engine produces at that minimum point. “Low fan pressure ratio” is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.45. “Low corrected fan tip speed” is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram ° R)/(518.7° R)]{circumflex over ( )}0.5. The “Low corrected fan tip speed” as disclosed herein according to one non-limiting embodiment is less than about 1150 ft/second (350.5 meters/second).
Gas turbine engines generate substantial amounts of heat that is exhausted from the turbine section 28 into a surrounding atmosphere. This expelled exhaust heat represents wasted energy and can be a large source of inefficiency in gas turbine engines. Further, transitioning away from hydrocarbon-based engines may be significant advantages, as described herein.
Turning now to
As shown, the turbine engine system 200 includes a hydrogen fuel system 222. The hydrogen fuel system 222 is configured to supply a hydrogen fuel from a hydrogen fuel tank 224 to the combustor 210 for combustion thereof. In this illustrative embodiment, the hydrogen fuel may be supplied from the hydrogen fuel tank 224 to the combustor 210 through a fuel supply line 226. The fuel supply line 226 may be controlled by a flow controller 228 (e.g., pump(s), valve(s), or the like). The flow controller 228 may be configured to control a flow through the fuel supply line 226 based on various criteria as will be appreciated by those of skill in the art. For example, various control criteria can include, without limitation, target flow rates, target turbine output, cooling demands at one or more heat exchangers, target flight envelopes, etc.
As shown, between the cryogenic fuel tank 224 and the flow controller 228 may be one or more heat exchangers 230, which can be configured to provide cooling to various systems onboard an aircraft by using the hydrogen as a cold-sink. Such hydrogen heat exchangers 230 may be configured to warm the hydrogen and aid in a transition from a liquid state to a supercritical fluid or gaseous state for combustion within the combustor 210. The heat exchangers 230 may receive the hydrogen fuel directly from the hydrogen fuel tank 224 as a first working fluid and a component-working fluid for a different onboard system. For example, the heat exchanger 230 may be configured to provide cooling to power electronics of the turbine engine system 200 (or other aircraft power electronics). In other embodiments, the arrangement of the heat exchanger 230 and the flow controller 228 (or a flow controller element, such as a pump) may be reversed. In some such embodiments, a pump, or other means to increase a pressure of the hydrogen sourced from the hydrogen fuel tank 224 may be arranged upstream of the heat exchanger 230. This pumping or pressure increase may be provided to pump the hydrogen to high pressure as a liquid (low power). It will be appreciated that other configurations and arrangements are possible without departing from the scope of the present disclosure.
In some non-limiting embodiments, an optional secondary fluid circuit may be provided for cooling one or more aircraft loads. In this secondary fluid circuit, a secondary fluid may be configured to deliver heat from the one or more aircraft loads to one or more liquid hydrogen heat exchanger. As such, heating of the hydrogen and cooling of the secondary fluid may be achieved. The above described configurations and variations thereof may serve to begin raising a temperature of the hydrogen fuel to a desired temperature for efficient combustion in the combustor 210.
The hydrogen may then pass through an optional supplemental heating heat exchanger 236. The supplemental heating heat exchanger 236 may be configured to receive hydrogen as a first working fluid and as the second working fluid may receive one or more aircraft system fluids, such as, without limitation, engine oil, environmental control system fluids, pneumatic off-takes, or cooled cooling air fluids. As such, the hydrogen will be heated, and the other fluid may be cooled. The hydrogen will then be injected into the combustor 210 through one or more hydrogen injectors, as will be appreciated by those of skill in the art.
When the hydrogen is directed along the flow supply line 226, the hydrogen can pass through a core flow path heat exchanger 232 (e.g., an exhaust waste heat recovery heat exchanger) or other type of heat exchanger. In this embodiment, the core flow path heat exchanger 232 is arranged in the core flow path downstream of the combustor 210, and in some embodiments, downstream of the low pressure turbine 214. In this illustrative embodiment, the core flow path heat exchanger 232 is arranged downstream of the low pressure turbine 214 and at or proximate the core nozzle 216 upstream of the outlet 218. As the hydrogen passes through the core flow path heat exchanger 232, the hydrogen will pick up heat from the exhaust of the turbine engine system 200. As such, the temperature of the hydrogen will be increased.
The heated hydrogen may then be passed into an expansion turbine 234. As the hydrogen passes through the expansion turbine 234 the hydrogen will be expanded. The process of passing the hydrogen through the expansion turbine 234 cools the hydrogen and extracts useful power through the expansion process. Because the hydrogen is heated from a cryogenic or liquid state in the hydrogen fuel tank 224 through the various mechanisms along the flow supply line 226, engine thermals may be improved.
Turning now to
The engine system 302 may include the components shown and described above, including, without limitation, a fan section, a compressor section, a combustor section, a turbine section, and an exhaust section. In this schematic illustration, without limitation, the engine system 302 include an engine oil system 306, an air cooling system 308, a burner 310 (e.g., part of a combustion section), an anti-ice system 312, and a generator system 314. Those of skill in the art will appreciate that other systems, components, and devices may be incorporated into the engine system 302, and the illustrative embodiment is merely for explanatory and illustrative purposes. In this illustrative embodiment, a hydrogen high pressure pump 316 and an oil pump 318 are arranged as part of the generator system 314. The generator system 314 further includes a turbo expander 320, an engine-side generator 322, and a hydrogen-air heat exchanger 324. An air turbine starter 326 is provided within the engine system 302. The anti-ice system 312 of the engine system 302 includes an engine bleed system 328 that is configured to supply warm air to a cowl anti-ice system 330 to prevent ice buildup on an engine cowl.
The aircraft system 304 include various features installed and present that are separate from but may be operably or otherwise connected to one or more of the engine system 302. In this illustrative, non-limiting configuration, the aircraft system 304 include one or more hydrogen tanks 332 configured to store liquid hydrogen onboard the aircraft, such as in tanks that are wing-mounted or arranged within the aircraft fuselage. The aircraft system 304 include a cabin air cooling system 334, a wing anti-ice system 336, flight controls 338, one or more aircraft-side generators 340, and aircraft power systems 342.
The schematic diagram in
Referring to the hydrogen flow path 344, liquid hydrogen may be sourced or supplied from the hydrogen tanks 332. One or more pumps 354 may be arranged to boost a pressure of the hydrogen as it is supplied from the hydrogen tanks 332. In some configurations, the pumps 354 may be low pressure pumps, providing an increase in pressure of about 20 psid to 50 psid, for example. The hydrogen may be supplied to one or more combustion systems. For example, a portion of the hydrogen may be supplied to an auxiliary power source 356, such as an auxiliary power unit having a hydrogen burner or a hydrogen-based fuel cell. The auxiliary power source 356 may be part of the aircraft systems 304 and may be configured to direct air to the air turbine starter 326 along a section of an air flow path 346. Further, the auxiliary power source 356 may be configured to generate power at the generator 340 to supply power to the aircraft power system 342, the hydraulic pump 350, and/or the cabin air cooling system 334 and other ECS systems or other onboard electrically powered systems of the aircraft.
For propulsion onboard the aircraft, a portion of the hydrogen may be supplied from the hydrogen tanks 332 along the hydrogen flow path 344 to an aircraft-system heat exchanger 358 which may include a hydrogen-air heat exchanger to cool air. The aircraft-system heat exchanger 358 may be part of the aircraft system 304. One or more low pressure pumps 354 may be arranged to boost a pressure of the hydrogen and thus heat the hydrogen before entering the aircraft-system heat exchanger 358. In some embodiments, the aircraft-system heat exchanger 358 may be part of an environmental control system (ECS) of the aircraft. The cooled air may be supplied, for example, to the cabin air cooling system 334. As this air is cooled, the hydrogen will be warmed within the aircraft-system heat exchanger 358. The warmed hydrogen may then be passed from the aircraft system 304 to the engine system 302. As shown, the hydrogen may flow through a portion of the hydrogen flow path 344 to the hydrogen high pressure pump 316. The hydrogen high pressure pump 316 is configured to increase the pressure of the warmed hydrogen to maintain a pressure above a combustor pressure and/or above a critical pressure in order to avoid a phase change to gas in the plumbing, piping, flow path, or heat exchangers, for example.
The boosted pressure hydrogen may then be conveyed to a second heat exchanger. In this configuration, the second heat exchanger is the hydrogen-air heat exchanger 324 of the generator system 314. The second heat exchanger 324 of this embodiment may be a hydrogen-air heat exchanger arranged proximate an exit or nozzle of the engine system 302 (e.g., exhaust air heat exchanger). In the second heat exchanger 324, the temperature of the hydrogen is further raised. Next, the hydrogen may be passed through the turbo expander 320 of the generator system 314. As the hydrogen is expanded through the turbo expander 320, a turbine may be driven to generate power at the engine-side generator 322. In one non-limiting example, the aircraft-side generator 340 may be configured to generate about 120 kW whereas the engine-side generator 322 may be configured to generate about 300 kW at cruise and about 1 MW at takeoff. That is, in accordance with some embodiments of the present disclosure, the engine-side generator 322 may be configured to generate more power than the aircraft-side generator 340. The expanded hydrogen may then be directed into (e.g., injected) the burner 310, with such supply of hydrogen to the burner 310 controlled by a valve 360. In some embodiments, and as shown, an electric compressor actuator 362 may be included within the engine system 302. The electric compressor actuator 362 may be configured to boost a pressure of the hydrogen prior to injection into the burner 310.
The engine system 302 may further include one or more heat exchanges 364, 366 configured to provide heat exchange onboard the engine. These additional heat exchanges may not be part of the hydrogen flow path 344. For example, an air-oil heat exchanger 364 and an air-air heat exchanger 366 may be arranged for appropriate cooling (or heating) as will be appreciated by those of skill in the art. In some non-limiting embodiments, a post-expander hydrogen-air heat exchanger 368 may be arranged between the turbo expander 320 and the burner 310 and may be used for cooled cooling air, for example.
Using the architecture illustrated in
In operation, the hydrogen high pressure pump 316, the hydrogen-air heat exchanger 324, and the turbo expander 320 may be configured to employ the full heat capacity of the hydrogen. For example, the hydrogen may be heated to, but not exceed, an auto-ignition temperature. To achieve this, the hydrogen high pressure pump 316 and the hydrogen-air heat exchanger 324 of the generator system 314 may be sized and configured to increase the temperature of the hydrogen such that it is near the auto-ignition temperature as it passes through the turbo expander 320. The increased pressure and temperature of the hydrogen results in an overheated and/or over pressurized hydrogen that is passed into the turbo expander 320. As such, the engine-side generator 322 may extract the most work from the hydrogen and generate electrical power within the engine system 302.
Turning now to
The output rotor shaft 410 is operably coupled to the gear box 404 at an input side 416 of the gear box 404. A gear system 418 of the gear box 404 can be configured to change a rotational speed of an output from an output side 420 of the gear box 404. An input rotor shaft 422 is operably coupled to the gear box 404 at the output side 420 of the gear box 404. In some embodiments, and without limitation, the gear box 404 may be configured with a 4:1 gearing ratio. It will be appreciated that other gearing ratios may be employed without departing from the scope of the present disclosure. The input rotor shaft 422 is operably coupled to a rotor system 424 of the generator 406. The rotor system 424 may include magnets, permanent magnets, or the like, which are rotationally driven relative to a stator system 426 of the generator 406. The generator 406 may thus generate electrical power for use onboard an aircraft, for example as described above. In some embodiments, the generator 406 may be configured to generate 300 kW, 500 kW, 1 MW, etc. of electrical power or greater.
As shown in
The use of cryogenic liquid hydrogen, or other cold fuels, for aviation fuel requires heat addition from the engine, such as from the exhaust. Having an exhaust heat exchanger opens the possibility of recovering waste exhaust heat if the heated fluid is run through turbo-expanders connected to load sources such as generators, pumps, or the like. Because the fuel flow rate from idle to maximum at takeoff varies by over ten time, it is difficult to design an expander that can operate efficiently over the full range of engine operation. As such, it may be advantageous to employ multiple turboexpanders that are used on demand and selected using coupling/decoupling assemblies (e.g., valves or clutches) to selectively activate and operate one or more of a multitude of turbo expanders. As described herein selective couplers are described for operation of one or more turbo expanders in a single system.
The inclusion of two turbo expanders within the system can achieve a wide range of turbo expander operability with respect to fuel flow that varies considerably between idle and max power. In accordance with an example operation, electrical power from the load source generators may be used in both idle and max power conditions but due to the range of fuel flow, only one expander and load source generator combination would run at idle and two would run during higher-power conditions. This approach prevents any one turbo expander from having to run significantly off-design with very insufficient flow because a smaller number will be used at lower power/fuel flow conditions via the connecting valve. Because the fuel flow can vary by approximate a factor of ten or more between maximum and minimum power, it will be appreciated that additional turbo expanders (e.g., beyond the two illustrated) to accommodate such variability in the fuel flow of the system. For example, and without limitation, in one non-limiting embodiment, a single turbo expander may be used at idle, and four or more turbo expanders may be used at maximum power. In such a scenario, multiple turbo expanders could be connected to a single load source via a gear system. In addition to accommodating higher fuel flow, such multi-turbo expander configuration provides redundancy within the system, such as a failure or issue related to one of the turbo expanders. In such a case, if one turbo expander fails or cannot provide the necessary output, an alternative turbo expander of the set can be used to ensure continued operation without substantial negative impacts.
Turning now to
In one non-limiting example, and as illustrated, the first turbo expander 502 may be always supplied with hydrogen from the valve assembly 516 (illustrated by solid lines). That is, the valve assembly 516 may be open to the first turbo expander 502 at all times and thus the fuel will always flow into and through the first turbo expander 502. As such, the first load source 504 may be operated to generate power at all times of operation. The second turbo expander 508 may be selectively operated by actuation of the valve assembly 516 which can direct all or a portion of the flow from the fluid line 514 (illustrated as dashed lines). For example, in some configurations, the second turbo expander 508 and the second load source 510 may be provided as a redundant or backup system in the event of a failure of the first turbo expander 502, the first load source 504, the first gear box 506, or components thereof. In other embodiments, the second turbo expander 508 and the second load source 510 may be used simultaneously with operation of the first turbo expander 502 and the first load source 504. In such embodiments, the secondary operation may be configured to generate additional power if required by systems of the aircraft, for example. In some configurations, the two load sources 504, 510, when configured as generators, may be typically operated at a power generation level that is less than a maximum generation level in order to minimize wear on the components of the turbo generator system 500. However, each of the turbo expanders 502, 508 and the load sources 504, 510 (e.g., generators) may be sized and configured to generate power at sufficient levels to compensate in the event that the other set of components has a failure or otherwise does not or cannot generate power. In some embodiments, such as in a redundancy configuration, each of the load source generators 504, 510 may be configured to generate at least 300 kW (e.g., 300 kW, 500 kW, 750 kW, 1 MW, etc.) of electrical power. In other embodiments, the combined output of the two load source generators 504, 510 may be at least 300 kW (e.g., 300 kW, 500 kW, 750 kW, 1 MW, etc.) of electrical power.
In the embodiment of
Turning now to
In one non-limiting example, and as illustrated, both the first turbo expander 602 and the second turbo expander 608 may be always supplied with hydrogen from the fluid line 614 (illustrated by solid lines). That is, the fluid line 614 may be open to both the first turbo expander 602 and the second turbo expander 608 at all times and thus the fuel will always flow into and through the turbo expanders 602, 608. When the selective coupler 616 has the second turbo expander 608 disengaged by the clutch 620, the second turbo expander 608 may freewheel.
In this illustrative configuration, the first load source 604 may be operated to generate power at all times of operation because the first turbo expander 602. The second turbo expander 608 may be selectively operated by actuation of the clutch 620 of the selective coupler 616. For example, in some configurations, the second turbo expander 608 and the second load source 610 may be provided as a redundant or backup system in the event of a failure of the first turbo expander 602, the first load source 604, the first gear box 606, or components thereof. In other embodiments, the second turbo expander 608 and the second load source 610 may be used simultaneously with operation of the first turbo expander 602 and the first load source 604. In such embodiments, the secondary operation may be configured to generate additional power if required by systems of the aircraft, for example. In some configurations, the two load sources 604, 610, when configured as generators, may be typically operated at a power generation level that is less than a maximum generation level in order to minimize wear on the components of the turbo generator system 600. However, each of the turbo expanders 602, 608 and the load sources 604, 610 (e.g., generators) may be sized and configured to generate power at sufficient levels to compensate in the event that the other set of components has a failure or otherwise does not or cannot generate power. In some embodiments, such as in a redundancy configuration, each of the load source generators 604, 610 may be configured to generate at least 300 kW of electrical power. In other embodiments, the combined output of the two load source generators 604, 610 may be at least 300 kW of electrical power.
In the embodiment of
In the embodiments of
Turning now to
In basic operation, a fuel may flow along a fluid line 724 and into the turbo expander 702. The fuel will drive rotation of the turbo expander 702 which will output rotational energy that can be converted into work by the load sources 704a-c. The controller 710 may be configured to selectively engage or disengage one or more of the clutches 712, 714 to generate work by second and third load sources 704b, 704c. When the clutches 712, 714 are disengaged, only the first load source 704a is operated and the second and third gears 720, 722 may freewheel. However, as fluid flow through the turbo expander 702 increases, the first load source 704a may not consume all of the rotational energy, and thus downstream load sources 704b-c may be selectively operated to take advantage of any excess energy generated at the turbo expander 702.
Turning now to
In the embodiment of
If the fuel flow is at sufficient level, generation of additional work may be achieved and the valve assembly 816 may be controlled to open and direct fluid toward the downstream turbo generators 802b-d. Such direction may be incremental, such that only a second turbo generator 802b is supplied with fluid and operated to drive the second load source 804b. If the flow rates are high enough, a first downstream valve 824a may be opened to allow a portion of the fluid to pass down to the third turbo generator 802c and thus the third load source 804c may be operated. Such further expansion of the system may be achieved by opening more downstream valves 824b-c. As such, the amount of work or power extracted from the flow of fuel may be maximized.
In an alternative configuration, each of the turbo expanders 802a-d and/or load sources 804a-d may be rated at different levels, and the valve system 816, 824a-c, may be controlled for selective operation of one or more of the turbo generators 802a-d. In such configurations, the first turbo generator 802a may not be continuously operated, but may be selectively operated as done with the other turbo generators 802b-c. For example, in one non-limiting example, each of the load sources 804a-c may be an electrical power generator that is configured to generate a different amount of peak power. Based on the flow passing through the valve assembly 816, the controller 818 may be configured to open one or more valves 824b-c (and/or direct flow to the first turbo generator 802a in this illustrative configuration).
It will be appreciated that various components and elements of the above described embodiments may be combined in configurations that are not illustratively shown but captured by the present disclosure. For example, a plurality of turbo expanders may be configured with both valving and clutches for selective operation thereof. Further, in some embodiments, a single turbo generator may be coupled to multiple different load sources by clutches or the like.
In the above described embodiments, a turbo expander assembly is provided to selectively engage or disengage from one or more load sources. In some embodiments the turbo expander assembly may be formed of a single turbo expander that is selectively coupled to one or more load sources (e.g.,
In accordance with embodiments of the present disclosure, multiple expanders may be provided within systems disclosed herein to retain high efficiency and an acceptable range(s) of mechanical power (e.g., speed and load) throughout a wide range of fuel flow rates. For example, such flow rates may vary by a factor of ten or greater between idle and full power operations. In view of this, the controllers (e.g., controllers 520, 618, 710, 818) may be configured in operable communication with one or more flow sensors or operably connected or in communication with other control systems, for example. Such controllers may be configured to respond to engine thrust command(s) and/or fuel flow rates as a way to determine how many expanders and/or load source to be engaged during a particular operational envelop. As such, the controller may be configured to control flow diversions and/or operation of one or more turbo expanders and/or load sources, based on external factors, such as associated with flight control operations.
In one non-limiting example of operation of a controller, turbo expander assembly, and connected load sources onboard an aircraft, the controller may be configured to operate or direct flow to a lowest (fewest) number of turbo expanders (or load sources) at ground and decent idle conditions and a maximum number at takeoff and climb conditions. Cruise flow conditions are in between, and thus more than the lowest number and less than the maximum number may be employed or operated at cruise. In some aircraft configurations and flight conditions, cruise is at about 30% of the flow at takeoff and idle is about 5-10% of takeoff flow. As such, in a configuration with two turbo expanders (and/or two load sources), the controller may be configured to ensure operation of both devices at takeoff and climb and one at cruise and idle. However, by having more than two turbo expanders (and/or load sources) in a system, the number of currently operating turbo expanders (and/or load sources) may be adjusted by the controller based on actual flow ratios through the one or more turbo expanders.
Further, in embodiments with multiple turbo expanders and/or multiple connected load sources, the controller(s) may be configured to detect failures or issues associated with one or more of the operably connected turbo expanders/load sources. For example, in a case of a failed turbo expander, a controller may be configured to detect such failure by detecting a shaft RPM of a turbo expander that is operating at a range outside of expected or predetermined range(s). A failed turbo expander may also be detected through a controller function/operation by detecting a low or non-existent power output from a load source that is operably connected to a given turbo expander. In such instances, a valve or clutch associated with the turbo expander may be actuated or operated to disconnect the individual turbo expander. The controller, in such situations, may also accommodate the loss of generator power from that turbo expander by raising the load on other turbo expander.
Given the redundancy and potential need to ensure specific power generation and/or work output from associated load sources, each of the load sources and/or turbo expanders may be configured and designed to accommodate such a failure scenario. For example, as discussed above, during normal operation, one or more of the turbo expanders and/or load sources may be configured to operate at less than maximum capacity, and thus there is space to increase capacity of one or more turbo expanders and/or load sources to maintain the required power and/or work generation through such systems.
The controllers of systems of the present disclosure and associated with turbo expander assemblies with operably connected load sources, may be configured as independent electronic and/or electrical devices, or may be software or hardware incorporated into already existing control devices and systems that exist, for example, onboard an aircraft. The controllers may be operably coupled to thrust command software and/or hardware in order to respond to thrust commands. The controllers may be in electrical and/or mechanical connection and/or communication with various types of sensors, including, without limitation, flow sensors (e.g., to detect throughflow of fuel through a turbo expander assembly), pressure sensors, temperature sensors, back pressure sensors, contact sensors (e.g., to detect contact in clutch-type configurations), and the like, as will be appreciated by those of skill in the art.
In some configurations, the controller may determine which and/or how many turbo expanders receive fuel throughflow and/or are operably connected to respective load sources based on throughflow and rotational energy of the turbo expanders. In some configurations, alternatively or in combination with turbo expander selective coupling, the controllers may be configured to determine which and/or how many load sources are operably connected and driven by one or more turbo expanders of the turbo expander assemblies. Further, the controllers may be configured to detect failure of turbo expanders and/or load sources based on rotational speeds, power output, current draws, sensor-based detection of failure, or the like. Upon detection of such failure, the controller can control the selective coupler to accommodate such failures. It will be appreciated that a failure may not be the only condition requiring such selective coupling. For example, if extra power is required from an operation, the controller may be configured to engage and/or selectively couple one or more additional load sources and/or associated turbo generators to generate such excess power.
Advantageously, embodiments of the present disclosure are directed to improved turbine engine systems that employ non-hydrocarbon fuels at cryogenic temperatures. In accordance with some embodiments, the systems described herein provide for a hydrogen-burning turbine engine that may include one or more load sources that are driven using one or more turbo expanders that are arranged along a fuel flow path from a cryogenic fuel source to a burner or other consumption device (e.g., fuel cell). The turbo expanders may be configured as multi-stage, multi-portion expanders that extract work from the fuel or working fluid, while reducing the pressure and temperature of the working fluid. In accordance with embodiments of the present disclosure, the multiple load sources may be used to generate work, generate power (e.g., electric generator), perform actuation or operation (e.g., actuation/valves/pumps), or the like. By including multiple load sources that can be selectively coupled to the driving mechanism (e.g., one or more turbo expanders), an aircraft system is provided that can leverage energy from a cryogenic source in a stepped manner such that optimal or desired work extraction can be achieved. Further, through selective decoupling, when flow rates drop, one or more turbo expanders can be decoupled or transitioned to freewheel. Accordingly, a controlled load source operation may be achieved without negatively impacting the fuel systems of the aircraft.
As used herein, the term “about” is intended to include the degree of error associated with measurement of the particular quantity based upon the equipment available at the time of filing the application. For example, “about” may include a range of ±8%, or 5%, or 2% of a given value or other percentage change as will be appreciated by those of skill in the art for the particular measurement and/or dimensions referred to herein.
The terminology used herein is for the purpose of describing particular embodiments only and is not intended to be limiting of the present disclosure. As used herein, the singular forms “a,” “an,” and “the” are intended to include the plural forms as well, unless the context clearly indicates otherwise. It will be further understood that the terms “comprises” and/or “comprising,” when used in this specification, specify the presence of stated features, integers, steps, operations, elements, and/or components, but do not preclude the presence or addition of one or more other features, integers, steps, operations, element components, and/or groups thereof. It should be appreciated that relative positional terms such as “forward,” “aft,” “upper,” “lower,” “above,” “below,” “radial,” “axial,” “circumferential,” and the like are with reference to normal operational attitude and should not be considered otherwise limiting.
While the present disclosure has been described in detail in connection with only a limited number of embodiments, it should be readily understood that the present disclosure is not limited to such disclosed embodiments. Rather, the present disclosure can be modified to incorporate any number of variations, alterations, substitutions, combinations, sub-combinations, or equivalent arrangements not heretofore described, but which are commensurate with the scope of the present disclosure. Additionally, while various embodiments of the present disclosure have been described, it is to be understood that aspects of the present disclosure may include only some of the described embodiments.
Accordingly, the present disclosure is not to be seen as limited by the foregoing description but is only limited by the scope of the appended claims.
