Patent ID: 12215628

DETAILED DESCRIPTION

FIG.1illustrates a system20for an aircraft. This aircraft may be an airplane, a helicopter, a drone (e.g., an unmanned aerial vehicle (UAV)), a spacecraft or any other manned or unmanned aerial vehicle. The aircraft system20ofFIG.1includes a gas turbine engine22, an external compressor24and a gearbox26. Herein, the term “external” may describe a component separate from, outside of and/or otherwise discrete from the gas turbine engine22. The aircraft system20ofFIG.1also includes one or more fluid components28and/or one or more nozzles30.

The gas turbine engine22ofFIG.1includes at least (or only) a compressor section32, a combustor section34, a turbine section36and an (e.g., annular) internal engine flowpath38. The engine flowpath38extends from an upstream airflow inlet40into the gas turbine engine22, sequentially through the compressor section32, the combustor section34and the turbine section36, to a combustion products exhaust42from the gas turbine engine22.

The combustor section34and the turbine section36each include a respective bladed rotor44and46. Each of these bladed rotors44,46includes a plurality of rotor blades distributed circumferentially around and connected to one or more respective rotor disks and/or hubs. The rotor blades, for example, may be formed integral with or mechanically fastened, welded, brazed, adhered and/or otherwise attached to the respective rotor disk(s) and/or hub(s).

The compressor section rotor44is connected to the turbine section rotor46through an engine shaft48. The compressor section rotor44, the turbine section rotor46and the engine shaft48collectively form or are otherwise included in an internal rotating structure50of the gas turbine engine22. This rotating structure50and its elements44,46and48are rotatable about an engine rotational axis52, which engine rotational axis52may also be an axial centerline of the gas turbine engine22.

The external compressor24may be configured as or otherwise include an axial flow compressor, a (e.g., one piece rotor assembly) centrifugal flow compressor or a combination of an axial flow and a centrifugal flow compressor. This external compressor24may be a single stage compressor or a multi-stage compressor. The external compressor24ofFIG.1includes a bladed compressor rotor54rotatable about a compressor rotational axis56, which compressor rotational axis56may also be an axial centerline of the external compressor24. The compressor rotational axis56may be parallel with, but laterally offset from/laterally spaced from, the engine rotational axis52. Of course, in other embodiments, the compressor rotational axis56may alternatively be angularly offset from the engine rotational axis52by an angle greater than zero degrees (0°) and less than one-hundred and eighty degrees (180°). The compressor rotor54includes a plurality of compressor blades distributed circumferentially around and connected to one or more respective disks and/or hubs. The compressor blades, for example, may be formed integral with or mechanically fastened, welded, brazed, adhered and/or otherwise attached to the respective disk(s) and/or hub(s).

The gearbox26operatively couples the rotating structure50to the compressor rotor54. The gearbox26ofFIG.1, for example, includes a gearbox input58(e.g., a shaft, a gear, a coupling, etc.) and a gearbox output60(e.g., a shaft, a gear, a coupling, etc.). The gearbox input58is coupled to and rotatable with the rotating structure50. The gearbox output60is coupled to and rotatable with the compressor rotor54. The gearbox output60is also coupled to the gearbox input58through an internal power transmission system61within the gearbox26; e.g., a gear system, etc.

The gearbox26may be configured as a constant speed transmission. With such a configuration, a ratio between an input speed into the gearbox26(e.g., a rotational speed of the gearbox input58, a rotational speed of the rotating structure50) and an output speed from the gearbox26(e.g., a rotational speed of the gearbox output60, a rotational speed of the compressor rotor54) remains constant. The gearbox26, however, may alternatively be configured as a variable speed transmission. With such a configuration, the ratio between the gearbox input speed and the gearbox output speed may change; e.g., selectively go up or down.

During operation of the aircraft system20, (e.g., outside, ambient) air enters the gas turbine engine22and its engine flowpath38through the airflow inlet40. This air is compressed by the compressor section rotor44and directed into a (e.g., annular) combustion chamber62of a combustor64in the combustor section34. Fuel is injected into the combustion chamber62by one or more fuel injectors66, and the fuel is mixed with the compressed air to provide a fuel-air mixture. The fuel-air mixture is ignited and combustion products thereof flow through and cause the turbine section rotor46to rotate. The rotation of the turbine section rotor46drives rotation of the compressor section rotor44and, thus, compression of the air received from the airflow inlet40. The rotation of the turbine section rotor46and, more generally, rotation of the rotating structure50also drives rotation of the compressor rotor54through the gearbox26. The rotation of the compressor rotor54compresses (e.g., outside, ambient) air entering the external compressor24through an inlet68into the external compressor24. This compressed air is directed out of the external compressor24through an outlet70from the external compressor24and provided to the fluid component(s)28to facilitate operation of the fluid component(s)28. Examples of the fluid component(s)28include, but are not limited to, downstream flow device(s), fluid control device(s), flow conversion component(s) or nozzle(s) configured for providing thrust and/or aerodynamic aircraft controls, aircraft surface(s) and aircraft wing(s).

With the arrangement ofFIG.1, the external compressor24is disposed outside of the gas turbine engine22and its engine flowpath38. Moreover, a compressor flowpath72through the external compressor24from the compressor inlet68to the compressor outlet70is discrete (e.g., fluidly decoupled from, parallel with, etc.) the engine flowpath38. The compressed air discharged from the external compressor24and provided to the fluid component(s)28and/or the nozzle(s)30may thereby be relatively cool, compared to the combustion products discharged from the gas turbine engine22. Furthermore, the external compressor24and its compressor rotor54may be specifically tailored for (e.g., optimally) providing the compressed air to the fluid component(s)28and/or the nozzle(s)30, whereas the compressor section32and its compressor section rotor44may be specifically tailored for (e.g., optimally) providing the compressed air to the combustor section34for combustion. The external compressor24and the compressor section32of the aircraft system20may thereby be more efficient than a single compressor providing compressed air to both a combustion section of a gas turbine engine and other components outside of that gas turbine engine.

In some embodiments, the aircraft system20may include a single external compressor24coupled to the gas turbine engine22through the gearbox26. In other embodiments, referring toFIG.2, the external compressor24may be one of a plurality of external compressors24(e.g.,24A and24B) coupled to the gas turbine engine22through the gearbox26. These external compressors24may have a common configuration. Alternatively, at least one of the external compressors24may be configured (e.g., sized, etc.) different than another one of the external compressors24. Furthermore, the external compressors24A and24B may be individually coupled to the gas turbine engine22via the gearbox26. The compressor rotor54A, for example, may be rotatably driven by the rotating structure50through the gearbox26independent of the compressor rotor54B. In addition, the compressor rotor54B may be rotatably driven by the rotating structure50through the gearbox26independent of the compressor rotor54A. However, the present disclosure is not limited to such an exemplary parallel arrangement.

In some embodiments, referring toFIGS.1and2, the gas turbine engine22may be dedicated to powering the external compressor(s)24. The gas turbine engine22, for example, may not be configured to power any other aircraft components (e.g., propulsor, generator, etc.) and/or provide (e.g., significant or any) aircraft thrust. In other embodiments, however, the combustion products exhausted from the gas turbine engine22may be utilized to provide aircraft thrust; e.g., the gas turbine engine22may be a turbojet gas turbine engine22. In addition or alternatively, referring toFIGS.3and4, the gas turbine engine22and its rotating structure50may be configured to drive rotation of at least (or only) one additional component through the gearbox26. The gearbox26ofFIG.3, for example, includes a second gearbox output60′ that is coupled to and rotatable with a driven rotor74within an electric generator76. In another example, referring toFIG.4, the second gearbox output60is coupled to and rotatable with a driven rotor of an aircraft propulsor; e.g., a bladed propulsor rotor78. The propulsor rotor78may be configured as or otherwise include an open propulsor rotor such as, but not limited to, a propeller or a helicopter rotor (e.g., a main rotor). The propulsor rotor78may alternatively be configured as or otherwise include a ducted propulsor rotor such as, but not limited to, a fan rotor. The compressor rotor54ofFIGS.3and4may be rotatably driven by the rotating structure50through the gearbox26independent of the driven rotor74,78. In addition, the driven rotor74,78ofFIGS.3and4may be rotatably driven by the rotating structure50through the gearbox26independent of the compressor rotor54. However, the present disclosure is not limited to such an exemplary parallel arrangement.

In some embodiments, referring toFIG.5, the gas turbine engine22may be configured with a heat recovery circuit80. This heat recovery circuit80ofFIG.5includes a first heat exchanger82, a turbine expander84(also referred to as a turboexpander), a second heat exchanger86and a recovery circuit compressor88. The first heat exchanger82is arranged with the engine flowpath38downstream of the turbine section36. The first heat exchanger82includes an inlet90fluidly coupled with and downstream of an outlet92from the recovery circuit compressor88. The first heat exchanger82also includes an outlet94fluidly coupled with and upstream of an inlet96to the turbine expander84. The turbine expander84includes its inlet96and an outlet98fluidly coupled with and upstream of an inlet100to the second heat exchanger86. The turbine expander84also includes a bladed turbine expander rotor102configured to drive rotation of a driven rotor104of a mechanical load106. This mechanical load106may be an electric generator, a bladed propulsor rotor or another device. The second heat exchanger86includes its inlet100and an outlet108fluidly coupled with and upstream of an inlet110to the recovery circuit compressor88. The recovery circuit compressor88includes its outlet92, its inlet110and a bladed compressor rotor112, where the compressor rotor112may be coupled to and rotatably driven by the turbine expander rotor102(or another device).

During operation of the heat recovery circuit80, fluid (e.g., refrigerant) is pressurized and directed through the first heat exchanger82by the recovery circuit compressor88. This pressurized fluid absorbs heat energy from the combustion products exhausted from the turbine section36through the first heat exchanger82. The heated and pressurized fluid flows through the turbine expander84and drives rotation of the turbine expander rotor102. The fluid output from the turbine expander84is condensed by the second heat exchanger86before flowing into the recovery circuit compressor88. The rotation of the turbine expander rotor102may drive rotation of the compressor rotor112and/or the driven rotor104. The heat recovery circuit80may thereby capture waste heat energy exhausted with the combustion products from the gas turbine engine22and use that captured energy to power the mechanical load106.

In some embodiments, referring toFIG.6, the gas turbine engine22may be configured with a fuel circuit114. This fuel circuit114includes a fuel source116(e.g., a fuel reservoir and/or a fuel pump), a heat exchanger118and the fuel injector(s)66. The heat exchanger118is arranged with the engine flowpath38downstream of the turbine section36. The heat exchanger118includes an inlet120fluidly coupled with and downstream of an outlet122from the fuel source116. The heat exchanger118also includes an outlet124fluidly coupled with and upstream of inlet(s)126to the fuel injector(s)66. With this arrangement, the heat exchanger118may pre-heat the fuel provided to the fuel injector(s)66to facilitate more efficient combustion.

In some embodiments, referring toFIG.7, the gas turbine engine22may be configured with a heat exchanger128such as a recuperator/an intercooler. The heat exchanger128ofFIG.7includes a first heat exchanger path130and a second heat exchanger path132. The first heat exchanger path130is fluidly coupled with and between the compressor section32and the combustor section34. The first heat exchanger path130may thereby form a select portion of the engine flowpath38between the compressor section32and the combustor section34. The second heat exchanger path132is fluidly coupled with and between the turbine section36and the combustion products exhaust42. The second heat exchanger path132may thereby form a select portion of the engine flowpath38between the turbine section36and combustion products exhaust42. With this arrangement, the heat exchanger128may pre-heat the compressed air provided to the combustor section34to facilitate more efficient combustion.

In some embodiments, the gas turbine engine22and the external compressor(s)24may be operated through aircraft operation; e.g., during taxiing, takeoff, climb, cruise, descent and/or landing. Such continuous operation is in contrast to, for example, intermittent operation of a typical auxiliary power unit (APU). However, in other embodiments, the gas turbine engine22and/or the external compressor(s) may be selectively operated.

The aircraft system20may include various gas turbine engines other than the one described above. The aircraft system20, for example, may include a geared turbine engine or a direct drive gas turbine engine. The gas turbine engine may include a single spool (e.g., rotating structure) or multiple spools (e.g., rotating structures). The turbine engine may be configured as a turbofan engine, a turbojet engine, a turboprop engine, a turboshaft engine, a propfan engine, a pusher fan engine or any other type of turbine engine. Furthermore, the gas turbine engine may also be configured as a hybrid engine with, for example, one or more electric machines (e.g., electric motors) for selectively driving rotation of the compressor rotor(s)54through the gearbox26. The present disclosure therefore is not limited to any particular types or configurations of gas turbine engines.

While various embodiments of the present disclosure have been described, it will be apparent to those of ordinary skill in the art that many more embodiments and implementations are possible within the scope of the disclosure. For example, the present disclosure as described herein includes several aspects and embodiments that include particular features. Although these features may be described individually, it is within the scope of the present disclosure that some or all of these features may be combined with any one of the aspects and remain within the scope of the disclosure. Accordingly, the present disclosure is not to be restricted except in light of the attached claims and their equivalents.