Abstract:
An example turbomachine system includes a first variable outer air seal including at least one channel. The first variable outer air seal configured to selectively communicate a fluid in response to movement of a second variable outer air seal relative to the first variable outer air seal. An example fluid control method includes selectively covering a channel inlet using a variable outer air seal to control flow through the channel.

Description:
BACKGROUND 
     This disclosure relates to a blade outer air seal (BOAS) assembly for a turbomachine and, more particularly, to a BOAS assembly having segments that are moved relative to each other to selectively communicate fluid. 
     Turbomachines, such as gas turbine engines, typically include a fan section, a compression section, a combustion section, and a turbine section. Turbomachines may employ a geared architecture connecting portions of the compression section to the fan section. 
     BOAS circumscribe arrays of blades in the compression section, turbine section, or both. Turbomachines have developed passive and active systems for controlling clearances of the gap between the outer air seal and the tip of the turbine blade. Significant and varied thermal energy levels may be concentrated in these areas. Cooling these areas is often difficult. Specific and dedicated components are used to provide flow and cooling, which adds weight and cost. 
     SUMMARY 
     A turbomachine system according to an exemplary aspect of the present disclosure includes, among other things, a first variable outer air seal including at least one channel. The first variable outer air seal is configured to selectively communicate fluid in response to movement of a second variable outer air seal relative to the first variable outer air seal. 
     In a further non-limiting embodiment of the foregoing turbomachine system, the at least one channel may extend from a radially outward facing surface to a radially inward facing surface. 
     In a further non-limiting embodiment of either of the foregoing turbomachine systems, the at least one channel may extend to a circumferentially facing surface. 
     In a further non-limiting embodiment of any of the foregoing turbomachine systems, the first and second variable outer air seal may be circumferentially adjacent. 
     In a further non-limiting embodiment of any of the foregoing turbomachine systems, the at least one channel has an inlet, and the second variable outer air seal may move relative to inlet between positions that permit flow through the inlet and positions that limit flow through the inlet to selectively communicate flow. 
     In a further non-limiting embodiment of any of the foregoing turbomachine systems, the first variable outer air seal may include an inclined surface, and the second variable outer air seal may move across the inclined surface to selectively communicate fluid. 
     In a further non-limiting embodiment of any of the foregoing turbomachine systems, the first and second variable outer air seals may have a shiplapped configuration. 
     In a further non-limiting embodiment of any of the foregoing turbomachine systems, the first and second variable outer air seals may be moveable relative to each other between a first position and a second position to selectively control fluid flow through at least one channel. The first and second variable outer air seals may circumferentially overlap each other when in the first position more than when in the second position. 
     In a further non-limiting embodiment of any of the foregoing turbomachine systems, the fluid may be cooling air. 
     In a further non-limiting embodiment of any of the foregoing turbomachine systems, the first and second variable outer air seals are blade outer air seals. 
     A method of turbomachine fluid control according to another exemplary aspect of the present disclosure includes, among other things, selectively covering a channel inlet using a variable outer air seal to control flow through the channel. 
     In a further non-limiting embodiment of the foregoing method of turbomachine fluid control, the channel may be a cooling channel. 
     In a further non-limiting embodiment of either of the foregoing methods of turbomachine fluid control, the variable outer air seal is a first outer air seal, and the channel may be provided by a second variable outer air seal. 
     In a further non-limiting embodiment of any of the foregoing methods of turbomachine fluid control, the variable outer air seal is a first outer air seal, and a second outer air seal may selectively cover the channel inlet to control flow through the channel. 
     A method of turbomachine fluid control according to yet another exemplary aspect of the present disclosure includes, among other things, moving a first variable outer air seal relative to a second variable outer air seal to control flow of a cooling fluid. 
     In a further non-limiting embodiment of the foregoing method of turbomachine fluid control, the moving may comprise moving the first and second variable outer air seals circumferentially relative to each other. 
    
    
     
       DESCRIPTION OF THE FIGURES 
       The various features and advantages of the disclosed examples will become apparent to those skilled in the art from the detailed description. The figures that accompany the detailed description can be briefly described as follows: 
         FIG. 1  is a cross-sectional view of an example turbomachine. 
         FIG. 2  shows a cross-sectional view of the high-pressure turbine of the turbomachine of  FIG. 1 . 
         FIG. 3  shows a perspective view of a variable area outer air seal fluid control system. 
         FIG. 4  shows a close up view of two variable area outer air seals of the system of  FIG. 3  in a first position. 
         FIG. 5  shows the two variable area outer air seals of  FIG. 4  in second position where the seals are more overlapped than when in the first position. 
         FIG. 6  shows a section view of one of the variable area outer air seals of  FIG. 4 . 
         FIG. 7  shows a section view another example variable area outer air seal. 
     
    
    
     DETAILED DESCRIPTION 
       FIG. 1  schematically illustrates an example turbomachine, which is a gas turbine engine  20  in this example. The gas turbine engine  20  is a two-spool turbofan gas turbine engine that generally includes a fan section  22 , a compression section  24 , a combustion section  26 , and a turbine section  28 . 
     Although depicted as a two-spool turbofan gas turbine engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to use with turbofans. That is, the teachings may be applied to other types of turbomachines and turbine engines including three-spool architectures. Further, the concepts described herein could be used in environments other than a turbomachine environment and in applications other than aerospace applications. 
     In the example engine  20 , flow moves from the fan section  22  to a bypass flowpath. Flow from the bypass flowpath generates thrust. The compression section  24  drives air along a core flowpath. Compressed air from the compression section  24  communicates through the combustion section  26 . The products of combustion expand through the turbine section  28 . 
     The example engine  20  generally includes a low-speed spool  30  and a high-speed spool  32  mounted for rotation about an engine central axis A. The low-speed spool  30  and the high-speed spool  32  are rotatably supported by several bearing systems  38 . It should be understood that various bearing systems  38  at various locations may alternatively, or additionally, be provided. 
     The low-speed spool  30  generally includes a shaft  40  that interconnects a fan  42 , a low-pressure compressor  44 , and a low-pressure turbine  46 . The shaft  40  is connected to the fan  42  through 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 a shaft  50  that interconnects a high-pressure compressor  52  and high-pressure turbine  54 . 
     The shaft  40  and the shaft  50  are concentric and rotate via bearing systems  38  about the engine central longitudinal axis A, which is collinear with the longitudinal axes of the shaft  40  and the shaft  50 . 
     The combustion section  26  includes a circumferentially distributed array of fuel nozzles within an annular combustor  56  that is generally arranged axially between the high-pressure compressor  52  and the high-pressure turbine  54 . 
     In some non-limiting examples, the engine  20  is a high-bypass geared aircraft engine. In a further example, the engine  20  bypass ratio is greater than about six (6 to 1). 
     The geared architecture  48  of the example engine  20  includes an epicyclic gear train, such as a planetary gear system or other gear system. The example epicyclic gear train has a gear reduction ratio of greater than about 2.3 (2.3 to 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 of the engine  20 . In one non-limiting embodiment, the bypass ratio of the engine  20  is greater than about ten (10 to 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 5 (5 to 1). The geared architecture  48  of this embodiment is an epicyclic gear train with a gear reduction ratio of greater than about 2.5 (2.5 to 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 disclosure is applicable to other gas turbine engines including direct drive turbofans. 
     In this embodiment of the example engine  20 , a significant amount of thrust is provided by the bypass flow 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. This flight condition, with the engine  20  at its best fuel consumption, is also known as “Bucket Cruise” Thrust Specific Fuel Consumption (TSFC). TSFC is an industry standard parameter of fuel consumption per unit of thrust. 
     Fan Pressure Ratio is the pressure ratio across a blade of the fan section  22  without the use of a Fan Exit Guide Vane system. The low Fan Pressure Ratio according to one non-limiting embodiment of the example engine  20  is less than 1.45 (1.45 to 1). 
     “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 Temperature represents the ambient temperature in degrees Rankine. The Low Corrected Fan Tip Speed according to one non-limiting embodiment of the example engine  20  is less than about 1150 fps (351 m/s). 
     Referring to  FIGS. 2 to 4 , the turbine section  28  of the engine  20  includes a blade outer air seal (“BOAS”) assembly  60  disposed between a plurality of circumferentially distributed rotor blades  62  of a rotor stage  64 , and an annular outer engine case  66 . In one embodiment, the BOAS  60  is adapted to limit air leakage between blade tips  68  and the engine case  66 . The example BOAS  60  is supported by rails  70  and  72  attached to the engine case  66 . BOAS  60  is also connected to an actuator  74  through a rod  76 . The actuator  74  may connect to a main digital control. In some examples, the actuator  74  may be wired to a control system via a cable  78 . In other examples, the actuator  74  attaches the main digital electronic control of the engine  20  in another ways. 
     The BOAS  60  includes multiple variable outer air seal segments  80  distributed annularly about the axis A. In this example, each segment has radially inwardly facing surfaces  82  and radially outwardly facing surfaces  84 . The segments  82  each include an inclined surface  86  attached to a base portion  88 . The inclined surface  86  is one of the radially outwardly facing surfaces  84  in this example. An extension  90  extends radially outward from the base portion  88 . The extension  90  may be a stanchion, tab, lug, or some other structure. The extension  90  has an aperture  92  for receiving a connector pin  94 . 
     Each segment  80  is connected to a circumferentially adjacent segment through a link  96  attached with the pin  94 . Some of the segments,  80   a  and  80   b  are attached to a single circumferentially adjacent segment  80 . Segment  80   b  is attached to the actuating rod  76 . Actuating rod  76  is directly coupled to the actuator  74 . Actuator  74  is attached to a control system  100  via the cable  78 . 
     The control system  100 , in this example, includes a sensor  102 , for example a thermocouple, which may be positioned to sense a gas path temperature at a particular location along a core flow path of the engine. In one example, the sensor  102  extends through a turbine case to measure a temperature approximate location T 4  at the entrance to the high-pressure turbine section, where airfoils and other components are particularly susceptible to thermal damage due to peaking gas temperatures. In another example, temperature sensor  102  may be positioned approximate another stage of the high-pressure turbine  54 , or within the low-pressure turbine  46 , or a compression section  24 . In other examples, a number of temperature probes are positioned in different locations within the engine  20  to measure multiple gas path temperatures along flowpaths of the engine  20 . 
     The control system  100  includes a flight controller  104  having a flight condition module, a thrust control, and other related engine functions. Depending on the embodiment, the flight controller  104  may comprise additional flight, engine, and navigational systems utilizing other control, sensor, and processor components located throughout the engine  20 , and in other regions of the engine. 
     Flight controller  104  includes a combination of software and hardware components configured to determine and report flight conditions relevant to the operation of engine  20 . In general, flight controller  104  includes a number of individual flight modules, which determine a range of different flight conditions based on a combination of pressure, temperature and spool speed measurements and additional data such as attitude and control surface positions. 
     Flight controller  104  may include a control law (CLW) configured to direct actuator  74  to adjust the modulated BOAS  60 . The CLW directs actuator  74  based on the sensed inputs from sensor  102 , the flight conditions determined by flight module, and other parameters, such as core flow gas path temperatures TC. 
     The flight controller  104  may direct the actuator  74  to adjust rod  76  in order to regulate the gap between the blade tips and radially inward facing surfaces  82  of the segments  80 . The linkage design connected to modulated BOAS  60  is designed such that if pushed in one direction, linkages are pulled in tension, thus increasing the diameter of the modulated BOAS  60 , while movement in the other direction creates compression within the linkages and decreases the overall diameter of modulated BOAS  60 . 
     Referring to  FIGS. 5 and 6  with continuing reference to  FIGS. 2 to 4 , adjacent ones of the segments  80  are moveable to shiplapped positions. When shiplapped, portions of circumferentially adjacent segments  80  overlap each other. The flight controller  104  may direct the actuator  74  to adjust rod  76  to move circumferentially adjacent segments  80 ′ and  80 ″ ( FIGS. 4 and 5 ) between the less shiplapped position of  FIG. 4  and the more shiplapped position of  FIG. 5 . In some examples, the actuator  74  may be configured to move the circumferentially adjacent segments  80 ′ and  80 ″ to positions where no portion of circumferentially adjacent segments  80 ′ and  80 ″ overlap. 
     The example segments  80 ′ and  80 ″ include channels  110  extending from the inclined surface  86  to a radially inward facing surface  82 . The channels  110  deliver a fluid, such as cooling air from a supply  112  to an interface between the radially inward facing surface  82  and the blade tip  68 . The supply  112  is radially outside the segments  80 ′ and  80 ″ in this example. 
     The flight controller  104  may direct the actuator  74  to adjust rod  76  in order to regulate flow of fluid through the channels  110 . The fluid cools the interface. The flow is regulated by selectively blocking flow entering an inlet  120  of the channels  110 . For example, the segment  80 ′ is used to selectively block the flow through channels  110  in the segment  80 ″. 
     The segment  80 ′ blocks flow through the channels  110  in the segment  80 ″ by covering some or all of the inlets  120  in the segment  80 ″. In this example, in circumferential Region R, increasing the circumferential overlap between the segments  80 ′ and  80 ″ increases the amount of blocked flow and reduces the amount of flow moving through channels  110 . The amount of blocked flow may thus be controlled by varying the amount of overlap between the segment  80  and the inlets  120 . 
     The example channels  110  are shown as being entirely within a single one of the segments  80 ′ or  80 ″. In other examples, the channels  110  may be defined partially by one of the segments  80 ′ or  80 ″, such as if the channels  110  were notches in a side of one of the segments  80 ′ and  80 ″. 
     The example channels  110  deliver fluid to the radially inward facing surfaces  82  interacting with the blade tip  68 . In other examples, the channels  110  may instead, or in addition to, deliver fluid to other areas, such as to a circumferentially facing surface  116  of the segments  80  ( FIG. 7 ). The size, angles, and positions of the channels  110  are adjustable according to specific cycle requirements, method or control, etc. 
     The preceding description is exemplary rather than limiting in nature. Variations and modifications to the disclosed examples may become apparent to those skilled in the art that do not necessarily depart from the essence of this disclosure. Thus, the scope of legal protection given to this disclosure can only be determined by studying the following claims.