Patent Application
20160169120
This disclosure generally relates to gas turbine engines and, more particularly, relates to a fuel schedule for supplying a combustor.
Many modern aircraft, as well as other vehicles and industrial processes, employ gas turbine engines for generating energy and propulsion. Such engines include a fan, compressor, combustor and turbine provided in serial fashion, forming an engine core and arranged along a central longitudinal axis. Air enters the gas turbine engine through the fan and is pressurized in the compressor. This pressurized air is mixed with fuel in the combustor. The fuel-air mixture is then ignited, generating hot combustion gases that flow downstream to the turbine. The turbine is driven by the exhaust gases and mechanically powers the compressor and fan via a central rotating shaft. Energy from the combustion gases not used by the turbine is discharged through an exhaust nozzle, producing thrust to power the aircraft.
Gas turbine engines contain an engine core and fan surrounded by a fan case, forming part of a nacelle. The nacelle is a housing that contains the engine. The fan is positioned forward of the engine core and within the fan case. The engine core is surrounded by an engine core cowl and the area between the nacelle and the engine core cowl is functionally defined as a fan duct. The fan duct is substantially annular in shape to accommodate the airflow from the fan and around the engine core cowl. The airflow through the fan duct, known as bypass air, travels the length of the fan duct and exits at the aft end of the fan duct at an exhaust nozzle.
In addition to thrust generated by combustion gasses, the fan of gas turbine engines also produces thrust by accelerating and discharging ambient air through the exhaust nozzle. Various parts of the gas turbine engine generate heat while operating, including the compressor, combustor, turbine, central rotating shaft and fan. To maintain proper operational temperatures, excess heat is often removed from the engine via oil coolant loops, including air/oil or fuel/oil heat exchangers, and dumped into the bypass airflow for removal from the system.
In operation, the gas turbine engine receives fuel from a fuel supply. The fuel is pressurized by a fuel pump, and injected into the combustor of the gas turbine engine via a fuel line. A gas turbine engine may include a plurality of fuel lines. Each of these fuel lines may inject fuel into the combustor at a particular location and orientation.
Each fuel line may inject fuel into the combustor using an atomizer or a jet. The atomizer may provide fuel to the combustor in a form sufficiently atomized for gas turbine transition between steady states in all conditions. To ensure successful gas turbine engine transition in all conditions, differing flow pressures for each fuel line may be called for.
Accordingly, there is a need for an improved fuel schedule for a gas turbine engine.
To meet the needs described above and others, the present disclosure provides a fuel injection system for a gas turbine engine, that may include a primary fuel line, a secondary fuel line, a fuel pump, a fuel supply, a control system adapted to operate in a first fuel delivery mode including a first delta pressure condition between the primary fuel line and the secondary fuel line, the control system adapted to further operate in a second fuel delivery mode including a second delta pressure condition between the primary fuel line and the secondary fuel line, a fuel schedule operated by the control system and employing the first fuel delivery mode during a transition between steady states and subsequently employing the second fuel delivery mode before the gas turbine engine enters a steady state.
The first delta pressure condition may have a larger pressure difference than the second delta pressure condition. The second fuel delivery mode may be employed during the transition. A flow divider valve may be employed to change between the first fuel delivery mode and the second fuel delivery mode, and the flow divider valve may be a pressure modulator having more than one position and may be used to alter a fuel pressure in the secondary fuel line. The transition between steady states may be between an inactive state and a ground idle power state, or may be between a ground idle power state and a higher power state.
The primary fuel line may include an atomizer for injecting a fuel into a combustor, while the secondary fuel line may include a jet for injecting fuel into the combustor. A dual passage injector may include the atomizer of the primary fuel line and the jet of the secondary fuel line for injecting fuel into the combustor.
The present disclosure also provides a gas turbine engine, that may include a compressor for compressing an airflow, a combustor downstream of the compressor, a primary fuel line, a secondary fuel line, a fuel pump, a fuel supply, a control system adapted to operate in a first fuel delivery mode including a first delta pressure condition between the primary fuel line and the secondary fuel line, the control system adapted to further operate in a second fuel delivery mode including a second delta pressure condition between the primary fuel line and the secondary fuel line, a fuel schedule operated by the control system and employing the first fuel delivery mode during a transition between steady states and subsequently employing the second fuel delivery mode before the gas turbine engine enters a steady state, and a turbine downstream of the combustor.
The first delta pressure condition may have a larger pressure difference than the second delta pressure condition. The second fuel delivery mode may be employed during the transition. A flow divider valve may be employed to change between the first fuel delivery mode and the second fuel delivery mode, and the flow divider valve may be a pressure modulator having more than one position and may be used to alter a fuel pressure in the secondary fuel line. The transition between steady states may be between an inactive state and a ground idle power state or may be between a ground idle power state and a higher power state.
The primary fuel line may include an atomizer for injecting a fuel into the combustor, while the secondary fuel line may include a jet for injecting fuel into the combustor. A dual passage injector may include the atomizer of the primary fuel line and the jet of the secondary fuel line for injecting fuel into the combustor.
The present disclosure further provides a method of scheduling fuel delivery in a gas turbine engine, that may comprise operating the gas turbine engine in a first fuel delivery mode during a transition between steady states, a first delta pressure condition existing between a primary fuel line and a secondary fuel line in the first fuel delivery mode, and subsequently operating the gas turbine engine in a second fuel delivery mode before the gas turbine engine enters a steady state, a second delta pressure condition existing between the primary fuel line and the secondary fuel line in the second fuel delivery mode, the first delta pressure condition having a larger pressure difference than the second delta pressure condition.
The method may include changing between the first fuel delivery mode and the second fuel delivery mode using a flow divider valve, the flow divider valve being a pressure modulator on the secondary fuel line having more than one position.
These, and other aspects and features of the present disclosure, will be better understood upon reading the following detailed description when taken in conjunction with the accompanying drawings.
For further understanding of the disclosed concepts and embodiments, reference may be made to the following detailed description, read in connection with the drawings, wherein like elements are numbered alike, and in which:
It is to be noted that the appended drawings illustrate only exemplary embodiments and are therefore not to be considered limiting with respect to the scope of the disclosure or claims. Rather, the concepts of the present disclosure may apply within other equally effective embodiments. Moreover, the drawings are not necessarily to scale, emphasis generally being placed upon illustrating the principles of certain embodiments.
The exemplary engine 20 generally 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 several bearing systems 38. It should be understood that various bearing systems 38 at various locations may alternatively or additionally be provided, and the location of bearing systems 38 may be varied as appropriate to the application.
The low speed spool 30 generally includes an inner shaft 40 that interconnects a fan 42, a low pressure compressor 44 and a low pressure turbine 46. The inner shaft 40 is connected to the fan 42 through a speed change mechanism, which in exemplary gas turbine engine 20 is illustrated 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 high pressure compressor 52 and high pressure turbine 54. A combustor 56 is arranged in exemplary gas turbine 20 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 generally between the high pressure turbine 54 and the low pressure turbine 46. The mid-turbine frame 57 further supports bearing systems 38 in the turbine section 28. The inner shaft 40 and the outer shaft 50 are concentric and rotate via bearing systems 38 about the engine central longitudinal axis A which is collinear with their longitudinal axes.
The core airflow 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 which are 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. It will be appreciated that each of the positions of the fan section 22, compressor section 24, combustor section 26, turbine section 28, and fan drive gear system 48 may be varied. For example, gear system 48 may be located aft of combustor section 26 or even aft of turbine section 28, and fan section 22 may be positioned forward or aft of the location of gear system 48.
The gas turbine engine 20 in one example is a high-bypass geared aircraft engine. In a further example, the gas turbine engine 20 bypass ratio is greater than about six (6), with an example embodiment being greater than about ten (10), 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. In one disclosed embodiment, the gas turbine engine 20 bypass ratio is greater than about ten (10:1), the fan diameter 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. 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. 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 exemplary of one embodiment of a geared architecture engine and that the present invention is applicable to other gas turbine engines including direct drive turbofans.
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. The flight condition of 0.8 Mach and 35,000 ft, 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)]0.5. The “Low corrected fan tip speed” as disclosed herein according to one non-limiting embodiment is less than about 1150 ft/second. The gas turbine engine 20 may further include an RPM sensor 62 for monitoring a rotational speed.
A fuel injection system 70 may be used to supply fuel to the combustor 56 as shown in
The fuel injection system 70 may include a fuel injector tip 82 located near the forward end of the combustor 56. The fuel injector tip 82 may include various means for injecting fuel into the combustor 56. The primary fuel line 76 and the secondary fuel line 78 may each travel through, or near, the fuel injector tip 82.
The primary fuel line 76 may include an atomizer 84 for injecting atomized fuel into the combustor 56. The secondary fuel line 78 may include a jet 86 for injecting fuel into the combustor 56. The primary fuel line 76 and the secondary fuel line 78 may terminate at the atomizer 84 and the jet 86, respectively. As will be described in detail below, a flow divider valve 90 may also be located on the secondary fuel line 78.
The atomizer 84 and jet 86 may both be housed in a dual passage injector 96, as shown in
For efficient continued combustion, various properties of the injected fuel may be controlled. One of these properties is the size of a fuel droplet injected into the combustor 56. Smaller droplets enable an increased total fuel surface area, aiding combustion efficiency. However, reducing the size of the droplets may introduce disadvantages including an increased pressure requirement, reduced fuel flow or more expensive components. Accordingly, a balance of larger and smaller droplets may be injected to the combustor 56.
During operation, a gas turbine engine 20 may transition between steady states. Steady states may include an inactive state, a ground idle power state and a higher power state. The higher power state may be a bleed air power state, a takeoff power state, a steady flight power state or another steady power state. Alternatively, the gas turbine engine 20 may transition among higher power states. The inactive state may involve the combustor 56 not continually combusting a fuel-air mixture, whereas ground idle power and higher power states may involve the combustor 56 continually combusting a fuel-air mixture.
The smaller droplets may provide fuel to the combustor 56 in a form sufficiently atomized for gas turbine engine 20 transition in all conditions. The atomizer 84 may inject fuel having this smaller droplet size into the combustor 56.
Alternatively, larger droplets may provide fuel to the combustor 56 that requires an interaction with the airflow 98 before being sufficiently atomized for gas turbine engine 20 transitioning in all conditions. This airflow 98, which may come from an air passage 92, may be insufficient prior to transition to guarantee a successful transition. The air passage 92 may be a swirler 94, or another type of vent, hole or passage. The jet 86 may inject fuel into the combustor 56 having this larger droplet size.
To ensure successful gas turbine engine 20 transitioning in all conditions, differing pressures for each fuel line 76, 78 may be called for. One example may involve limiting the pressure in the secondary fuel line 78 during transition, but before the gas turbine engine 20 reaches a steady state.
In limiting the pressure in the secondary fuel line 78, a greater percentage of total fuel injected into the combustor 56 may come through the primary fuel line 76 and atomizer 84 than through the secondary fuel line 78 and jet 86. The greater percentage may be sufficient to allow gas turbine engine 20 transition during all conditions. This fuel delivery configuration may be called a first fuel delivery mode, and it may be understood that this first fuel delivery mode employs a first delta pressure condition as the fuel pressure in the primary and secondary fuel lines 76, 78 differs by a first delta. The flow divider valve 90 may be used to change a fuel pressure in the secondary fuel line 78. The flow divider valve 90 may be a pressure modulator having more than one position, and may comprise a solenoid valve, or other type of valve or pressure modulator.
Another fuel delivery configuration could also limit the pressure in the secondary fuel line 78, and a greater percentage of total fuel injected into the combustor 56 may come through the primary fuel line 76 and atomizer 84 than through the secondary fuel line 78 and jet 86. However, this fuel delivery configuration may be called a second fuel delivery mode, and it may be understood that this second fuel delivery mode employs a second delta pressure condition as the fuel pressure in the primary and secondary fuel lines 76, 78 differs by a second delta. The first delta pressure condition may have a larger pressure difference than the second delta pressure condition. As before, the flow divider valve 90 may be used to change a fuel pressure in the secondary fuel line 78, and the flow divider valve 90 may also be employed to change between the first fuel delivery mode and the second fuel delivery mode.
In another embodiment, the second fuel delivery mode could include equal fuel pressures for the primary and secondary fuel lines 76, 78. That is, the second fuel delivery mode may employ a second delta pressure condition where the pressure in the primary and secondary fuel lines 76, 78 differs by a second delta equal to zero.
A control system 100, which may include a microprocessor 102 and a memory 104, may be in electronic communication with the RPM sensor 62, fuel pump 74, flow meter 80 and flow divider valve 90, as shown in
When operating in the first fuel delivery mode, the decreased pressure in the secondary fuel line 78 may not allow enough fuel flow to sufficiently fill the secondary fuel line 78. If the secondary fuel line 78 is not sufficiently filled, and an increased fuel flow rate is commanded by the control system 100, the additional fuel supplied will simply fill the secondary fuel line 78 rather than be combusted in the gas turbine engine 20. This may produce an inconsistency between the actual gas turbine engine 20 power output and the expected gas turbine engine 20 power output for the amount of fuel being injected, particularly if a steady state were entered while operating in the first fuel delivery mode. In response to the lack of actual power being produced, the control system 100 may command a still higher fuel flow rate, possibly leading to improper fuel regulation.
However, operating in the second fuel delivery mode may allow the secondary fuel line 78 to be sufficiently filled, eliminating the filling of the secondary fuel line 78 in response to an increased fuel flow, as shown in
Although not shown in
The present disclosure allows for the elimination of individual fuel injector valves, while providing a system and process for successfully transitioning a gas turbine engine 20 in all conditions. The present disclosure also helps avoid operation during a steady state, or increasing power mode, with insufficiently filled secondary fuel line 78. The elimination of individual fuel injector valves may reduce the total complexity and number of parts of the fuel injection system 70. In turn, this reduction may lead to decreased build, acquisition and maintenance costs, reduced system weight and improved system packaging.
While the present disclosure has shown and described details of exemplary embodiments, it will be understood by one skilled in the art that various changes in detail may be effected therein without departing from the spirit and scope of the disclosure as defined by claims supported by the written description and drawings. Further, where these exemplary embodiments (and other related derivations) are described with reference to a certain number of elements it will be understood that other exemplary embodiments may be practiced utilizing either less than or more than the certain number of elements.
In operation, the present disclosure sets forth a fuel delivery schedule for a gas turbine engine which can find industrial applicability in a variety of settings. For example, the disclosure may be advantageously employed by gas turbine engines 10 in aviation, naval and industrial settings. More specifically, the fuel delivery schedule for a gas turbine engine can be used to enable successful transition in all conditions, but avoid the possibility of steady state, or increasing power mode, operations with an insufficiently filled secondary fuel line 78.
The present disclosure allows for the elimination of individual fuel injector valves, while providing a system and process for successfully transitioning a gas turbine engine 20 in all conditions. The elimination of individual fuel injector valves may reduce the total complexity and number of parts of the fuel injection system 70. In turn, this reduction may lead to decreased build, acquisition and maintenance costs, reduced system weight and improved system packaging.
The fuel delivery schedule for a gas turbine engine of the present disclosure contributes to a gas turbine engine's 10 continued and efficient operation. The disclosed system may be original equipment on new gas turbine engines 10, or added as a retrofit to existing gas turbine engines 10.
This patent application claims priority under the 35 USC §119(e) to U.S. Provisional Patent Application Ser. No. 62/090,225, filed on Dec. 10, 2014.
| Number | Date | Country | |
|---|---|---|---|
| 62090225 | Dec 2014 | US |
