Patent Application
20180223684
Control of the radial clearance between the tips of rotating blades and the surrounding annular shroud in axial flow gas turbine engines improves engine efficiency. For example, by reducing the blade tip to shroud clearance, designers can reduce the quantity of turbine working fluid which bypasses the blades, thereby increasing engine power output for a given fuel or other engine input. On the other hand, blade tip to shroud contact leads to friction losses and wearing of parts. “Active clearance control” refers to clearance control arrangements wherein a quantity of working fluid, such as air, is employed by the clearance control system to regulate the thermal expansion of engine structures, thereby controlling the blade tip to shroud clearance.
Disclosed is an active tip clearance control system (ATCCS) for a gas turbine engine, including an electronically controlled regulating valve directing cooling airflow to a turbine case, and an engine electronic control (EEC), controlling the electronically controlled regulating valve, wherein the EEC controls the electronically controlled regulating valve to regulate cooling airflow according to a selected target blade tip clearance schedule, and wherein the selected target blade tip clearance schedule is selected before or after an engine cycle, from a plurality of target blade tip clearance schedules, each correlating to one of a plurality of thrust rating applications for the engine.
In addition to one or more of the features described above, or as an alternative, further embodiments may include that each of the target blade tip clearance schedules regulates cooling airflow for each phase of flight and for throttle excursions within and between each phase of flight.
In addition to one or more of the features described above, or as an alternative, further embodiments may include that the EEC is a full authority digital engine control (FADEC).
In addition to one or more of the features described above, or as an alternative, further embodiments may include a turbine case, a bladed rotary component supported by a spool, a shroud disposed radially within and fixedly supported by the turbine case, wherein blade tips are radially within and proximate to the shroud.
In addition to one or more of the features described above, or as an alternative, further embodiments may include that the electronically controlled regulating valve is exterior to the turbine case, and cooling airflow is directed therefrom toward a radially exterior side of the turbine case, and against thermally exposed portions of the turbine case and shroud connectors.
Also disclosed is a gas turbine engine including a turbine, the turbine including a bladed rotary component supported by a spool, a turbine case, and the active tip clearance control system (ATCCS).
Also disclosed is a method for providing active tip clearance control to a gas turbine engine, the method including selecting, by a computer processor, before or after an engine cycle of the gas turbine engine, a thrust rating application for a next engine cycle that differs from a currently selected thrust rating application, obtaining, by the computer processor, a target blade tip clearance schedule from of a plurality of target blade tip clearance schedules, each of the plurality of target blade tip clearance schedules correlating to one of a plurality of thrust rating applications for the engine, and forwarding cooling airflow toward a turbine case by controlling an electronically controlled regulating valve pursuant to the selected target blade tip clearance schedule.
The following descriptions should not be considered limiting in any way. With reference to the accompanying drawings, like elements are numbered alike:
A detailed description of one or more embodiments of the disclosed apparatus and method are presented herein by way of exemplification and not limitation with reference to the Figures.
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. An engine static structure 36 is arranged generally between the high pressure turbine 54 and the low pressure turbine 46. The engine static structure 36 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 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 engine 20 in one example is a high-bypass geared aircraft engine. In a further example, the 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 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 engine 20 is designed for a particular flight condition--typically cruise at about 0.8 Mach and about 35,000 feet (10,688 meters). The flight condition of 0.8 Mach and 35,000 ft (10,688 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)]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 m/sec).
Referring to
Similar to the engine 20 illustrated in
As illustrated in
A quantity of cooling airflow may be introduced via the active tip clearance control system 114 from the atmosphere, from, e.g., ram air, or bled from the compressor stage of the gas turbine engine 110 and into an aperture 113, illustrated in
The apertures 124 in the shielding plate 122 permit cooling airflow to impinge the case 126. As illustrated in
Cooling airflow, supplied through the active tip clearance control system 114, is funneled through an electronically controlled regulating valve 140, illustrated schematically in
Engines, such as engine 110, are designed to be used with different aircrafts requiring different levels of thrust, commonly referred to as thrust ratings. For each engine, the amount of cooling airflow needed, in order to provide the preferred blade tip clearance control, changes based on the aircraft thrust rating. Placing the engine 110 in an aircraft with a relatively higher rating will expose the engine 110 to greater thermal stresses, and therefore greater thermal expansions, requiring more cooling airflow to achieve preferred blade tip clearance control.
Typically, an active tip clearance control system 114 controls airflow, using the EEC 142 to operate the valve 140, pursuant to a middle ground clearance schedule in all anticipated applications during the service life of the engine 110. The third curve 206 in
Having the active tip clearance control system 114 control the valve 140 pursuant to a schedule defined by curve 206 for all thrust rating applications may not be ideal. When the engine 110 is used to achieve the higher thrust rating, controlling the valve 140 pursuant to the first curve 202 may not provide enough cooling airflow. This results in a the occurrence of a certain amount of blade tip rub, friction losses, efficiency losses and a decrease in the life of engine parts. When the engine 110 is used to achieve the lower thrust rating, controlling the valve 140 pursuant to the second curve 204 may provide too much cooling airflow. This results in excessive blade tip clearance, allowing core air to escape around turbine blade edges instead of driving the turbine, reducing engine efficiencies.
In the disclosed active tip clearance control system 114, the EEC 142 may be programmed to operate the valve 140 pursuant to plural clearance target curves 202, 204, corresponding to plural anticipated thrust rating applications during the service life of the engine 110. The EEC 142 may control the electronically controlled regulating valve 140 to allow more cooling airflow to the case 126 and shroud connectors 128 under the higher thrust rating application, and less cooling airflow under the lower thrust rating application. As a result, the same engine 110 may be used in plural aircrafts, having plural thrust ratings, without resulting in the inefficiencies of the active tip clearance control system 114 operating the valve 140 pursuant to middle ground clearance target curve 206.
The EEC 142 in the active tip clearance control system 114 may be switched to control the valve 140 pursuant to any of the plural blade tip clearance target curves, any time before or after an engine cycle, i.e., before engine start or after engine shutdown. Periods for switching include prior to use in an aircraft, e.g., at or before install of the engine 110 in a nacelle mounted to an airframe, or upon a first flight after an install. The EEC 142 for the active tip clearance control system 114 may be an integral part of the FADEC, or may be provided separately from the FADEC, in which case the EEC 142 may electronically communicate blade tip clearance control data and/or thrust rating data to the FADEC. If not part of the FADEC, the EEC 142 may be located on the engine 110, elsewhere in the aircraft, or at a remote location.
This step 304 may occur proximate to engine install, such as at the time of install, or thereafter, but before a next engine run. This step 304 may occur well in advance of engine install, such after a last engine cycle in a prior application. This step 304 may include providing an automated query to persons responsible for assisting in this operation, and updating the EEC 142 based on a response. This step 304 may be automated, via an electronic communication between a specially programmed EEC 142 and the engine FADEC. To accomplish this step 304, the active tip clearance control system 114 may include an on-engine manual switch, which identifies thrust rating application options, and which electronically communicates with the EEC 142 for switching the operational parameters of the active tip clearance control system 114 to achieve the preferred target clearances.
A next step 306, includes the EEC 142 of the active tip clearance control system 114 obtaining a target blade tip clearance schedule for operating the valve 140. The schedule is obtained from of the plurality of target blade tip clearance schedules for the engine 110, each of the plurality of target blade tip clearance schedules correlating to one of the plurality of thrust rating applications for the engine 110. This step may include retrieving the preferred schedule stored within an on-board EEC, or using networked communications to receive the information from a remote data store.
A next step 308 is the EEC 142 of the active tip clearance control system 114 forwarding cooling airflow toward a turbine by controlling the electronically controlled regulating valve 140 pursuant to the selected target blade tip clearance schedule. A next step 310 is the active tip clearance control system 114, via the EEC 142, monitoring electronic communications for the engine 110 to identify when a new thrust rating application is selected, at which point the process cycles back to step 304.
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” can include a range of ±8% or 5%, or 2% of a given value.
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.
While the present disclosure has been described with reference to an exemplary embodiment or embodiments, it will be understood by those skilled in the art that various changes may be made and equivalents may be substituted for elements thereof without departing from the scope of the present disclosure. In addition, many modifications may be made to adapt a particular situation or material to the teachings of the present disclosure without departing from the essential scope thereof. Therefore, it is intended that the present disclosure not be limited to the particular embodiment disclosed as the best mode contemplated for carrying out this present disclosure, but that the present disclosure will include all embodiments falling within the scope of the claims.
