Patent Publication Number: US-10316684-B2

Title: Rapid response clearance control system for gas turbine engine

Description:
This application claims priority to PCT Patent Application No. PCT/US14/15083 filed Feb. 6, 2014, which claims priority to U.S. Patent Appln. No. 61/811,533 filed Apr. 12, 2013. 
    
    
     STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH OR DEVELOPMENT 
     This disclosure was made with Government support under FA-8650-09-D-2923 0021 awarded by The United States Air Force. The Government has certain rights in this disclosure. 
    
    
     BACKGROUND 
     The present disclosure relates to a gas turbine engine and, more particularly, to a blade tip rapid response active clearance control (RRACC) system therefor. 
     Gas turbine engines, such as those that power modern commercial and military aircraft, generally include a compressor to pressurize an airflow, a combustor to burn a hydrocarbon fuel in the presence of the pressurized air, and a turbine to extract energy from the resultant combustion gases. The compressor and turbine sections include rotatable blade and stationary vane arrays. Within an engine case structure, the radial outermost tips of each blade array are positioned in close proximity to a shroud assembly. Blade Outer Air Seals (BOAS) supported by the shroud assembly are located adjacent to the blade tips such that a radial tip clearance is defined therebetween. 
     When in operation, the thermal environment in the engine varies and may cause thermal expansion and contraction such that the radial tip clearance varies. The radial tip clearance is typically designed so that the blade tips do not rub against the BOAS under high power operations when the blade disk and blades expand as a result of thermal expansion and centrifugal loads. When engine power is reduced, the radial tip clearance increases. To facilitate engine performance, it is operationally advantageous to maintain a close radial tip clearance through the various engine operational conditions. 
     SUMMARY 
     An active clearance control system for a gas turbine engine according to one disclosed non-limiting embodiment of the present disclosure includes a puller engageable with an air seal segment. 
     A further embodiment of the present disclosure includes, wherein the puller is not rigidly mounted to the air seal segment. 
     A further embodiment of any of the foregoing embodiments of the present disclosure includes, wherein the puller includes a plate configured to engage a forward hook and an aft hook of the air seal segment. 
     A further embodiment of any of the foregoing embodiments of the present disclosure includes, wherein the plate is X-shaped. 
     A further embodiment of any of the foregoing embodiments of the present disclosure includes, further comprising a rod affixed to the plate. 
     A further embodiment of any of the foregoing embodiments of the present disclosure includes an actuator mounted to the rod to drive the puller in response to a control. 
     A gas turbine engine according to another disclosed non-limiting embodiment of the present disclosure includes a full-hoop thermal control ring. A multiple of air seal segments movably mounted to the full-hoop thermal control ring and a multiple of pullers, each of the multiple of pullers engageable with one of the multiple of air seal segments. 
     A further embodiment of any of the foregoing embodiments of the present disclosure includes, wherein each of the multiple of pullers is not rigidly mounted to the respective one of the multiple of air seal segments. 
     A further embodiment of any of the foregoing embodiments of the present disclosure includes, wherein each of the multiple of air seal segments includes a forward hook and an aft hook engageable with the full-hoop thermal control ring. 
     A further embodiment of any of the foregoing embodiments of the present disclosure includes, wherein the puller includes a plate configured to engage the forward hook and the aft hook of each of the multiple of air seal segments. 
     A further embodiment of any of the foregoing embodiments of the present disclosure includes, wherein the plate is X-shaped. 
     A method of active blade tip clearance control for a gas turbine engine, according to another disclosed non-limiting embodiment of the present disclosure includes selectively engaging a puller with each of a multiple of air seal segments to selectively extend and retract each of the multiple of air seal segments with the puller not being rigidly mounted to the air seal segment. 
     A further embodiment of any of the foregoing embodiments of the present disclosure includes at least partially supporting each of the multiple of air seal segments with a full-hoop thermal control ring. 
     A further embodiment of any of the foregoing embodiments of the present disclosure includes engaging a forward hook and an aft hook of each of the multiple of air seal segments with the full-hoop thermal control ring. 
     A further embodiment of any of the foregoing embodiments of the present disclosure includes engaging a plate of the puller with the forward hook and the aft hook of each of the multiple of air seal segments. 
     The foregoing features and elements may be combined 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, the following description and drawings are intended to be exemplary in nature and non-limiting. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       Various features will become apparent to those skilled in the art from the following detailed description of the disclosed non-limiting embodiment. The drawings that accompany the detailed description can be briefly described as follows: 
         FIG. 1  is a schematic cross-section of one example aero gas turbine engine; 
         FIG. 2  is an is an enlarged partial sectional schematic view of a portion of a rapid response active clearance control system according to one disclosed non-limiting embodiment; 
         FIG. 3  is an enlarged top view of one of a multiple of air seal segments of the rapid response active clearance control system; 
         FIG. 4  is an enlarged partial sectional schematic view of one of a multiple of air seal segments taken along line  4 . 5 - 4 . 5  in  FIG. 3  with the rapid response active clearance control system in an extended position; and 
         FIG. 5  is an enlarged partial sectional schematic view of one of a multiple of air seal segments taken along line  4 . 5 - 4 . 5  in  FIG. 3  with the rapid response active clearance control system in an extended position. 
     
    
    
     DETAILED DESCRIPTION 
       FIG. 1  schematically illustrates a gas turbine engine  20 . The gas turbine engine  20  is disclosed herein as a two-spool low-bypass augmented turbofan that generally incorporates a fan section  22 , a compressor section  24 , a combustor section  26 , a turbine section  28 , an augmenter section  30 , an exhaust duct section  32 , and a nozzle system  34  along a central longitudinal engine axis A. Although depicted as an augmented low bypass turbofan in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are applicable to other gas turbine engines including non-augmented engines, geared architecture engines, direct drive turbofans, turbojet, turboshaft, multi-stream variable cycle adaptive engines and other engine architectures. Variable cycle gas turbine engines power aircraft over a range of operating conditions and essentially alters a bypass ratio during flight to achieve countervailing objectives such as high specific thrust for high-energy maneuvers yet optimizes fuel efficiency for cruise and loiter operational modes. 
     An engine case static structure  36  defines a generally annular secondary airflow path  40  around a core airflow path  42 . Various case static structures and modules may define the engine case static structure  36  which essentially defines an exoskeleton to support the rotational hardware. 
     Air that enters the fan section  22  is divided between a core airflow through the core airflow path  42  and a secondary airflow through a secondary airflow path  40 . The core airflow passes through the combustor section  26 , the turbine section  28 , then the augmentor section  30  where fuel may be selectively injected and burned to generate additional thrust through the nozzle system  34 . It should be appreciated that additional airflow streams such as third stream airflow typical of variable cycle engine architectures may additionally be sourced from the fan section  22 . 
     The secondary airflow may be utilized for a multiple of purposes to include, for example, cooling and pressurization. The secondary airflow as defined herein may be any airflow different from the core airflow. The secondary airflow may ultimately be at least partially injected into the core airflow path  42  adjacent to the exhaust duct section  32  and the nozzle system  34 . 
     The exhaust duct section  32  may be circular in cross-section as typical of an axisymmetric augmented low bypass turbofan or may be non-axisymmetric in cross-section to include, but not be limited to, a serpentine shape to block direct view to the turbine section  28 . In addition to the various cross-sections and the various longitudinal shapes, the exhaust duct section  32  may terminate in a Convergent/Divergent (C/D) nozzle system, a non-axisymmetric two-dimensional (2D) C/D vectorable nozzle system, a flattened slot nozzle of high aspect ratio or other nozzle arrangement. 
     With reference to  FIG. 2 , a blade tip rapid response active clearance control (RRACC) system  58  includes a radially adjustable blade outer air seal system  60  that operates to control blade tip clearances inside for example, the turbine section  28 , however, other sections such as the compressor section  24  may also benefit herefrom. The radially adjustable blade outer air seal system  60  may be arranged around each or particular stages within the gas turbine engine  20 . That is, each rotor stage may have an associated radially adjustable blade outer air seal system  60  of the blade tip rapid response active clearance control system  58 . 
     Each radially adjustable blade outer air seal system  60  is subdivided into a multiple of circumferential segments  62 , each with a respective air seal segment  64 , a drive link  66  and a puller  68  (also shown in  FIG. 3 ). In one disclosed non-limiting embodiment, each circumferential segment  62  may extend circumferentially for about nine (9) degrees. It should be appreciated that any number of circumferential segments  62  may be and various other components may alternatively or additionally be provided. 
     Each of the multiple of air seal segments  64  is at least partially supported by a generally fixed full-hoop thermal control ring  70 . That is, the full-hoop thermal control ring  70  is mounted to, or forms a portion of, the engine case static structure  36 . It should be appreciated that various static structures may additionally or alternatively be provided to at least partially support the multiple of air seal segments  64  yet permits relative radial movement therebetween. 
     Each air seal segment  64  may be manufactured of an abradable material to accommodate potential interaction with the rotating blade tips  28 T within the turbine section  28 . Each air seal segment  64  also includes numerous cooling air passages  64 P to permit secondary airflow therethrough. 
     A radially extending forward hook  72  and an aft hook  74  of each air seal segment  64  respectively cooperates with a forward hook  76  and an aft hook  78  of the full-hoop thermal control ring  70 . The forward hook  76  and the aft hook  78  of the full-hoop thermal control ring  70  may be segmented ( FIG. 3 ) or otherwise configured for assembly of the corresponding respective air seal segment  64  thereto. The forward hook  72  may extend axially aft and the aft hook  74  may extend axially forward (shown); vice-versa or both may extend axially forward or aft within the engine to engage the reciprocally directed forward hook  76  and aft hook  78  of the full-hoop thermal control ring  70 . 
     With continued reference to  FIG. 2 , the forward hook  76  and the aft hook  78  also interact with the puller  68  which permits the respective air seal segment  64  to be radially positioned between an extended radially contracted position ( FIG. 4 ) and a retracted radially expanded position ( FIG. 5 ) with respect to the full-hoop thermal control ring  70  by the puller  68 . In the retracted radially expanded position ( FIG. 5 ) when the air seal segments  64  are retracted, the air seal segments  64  are pinned against the thermal control ring  70  by the puller  68  but movement of the puller  68  is not radially restricted by the thermal control ring  70 . 
     The puller  68  generally includes a plate  80  and a rod  82 . The plate  80  may be X-shaped or otherwise configured to engage the forward hook  72  and the aft hook  74  of the respective air seal segment  64  ( FIG. 3 ). It should be appreciated that other configurations may alternatively be provided. The rod  82  is rigidly mounted to the plate  80 , e.g., fastened, bolted, welded, brazed, etc. such that movement of the rod  82  moves the plate  80  which then radially positions the respective air seal segment  64 . 
     The puller  68  provides actuation of the respective air seal segment  64  yet permits the effective use of legacy cooling schemes. That is, the plate  80  is engageable with the respective air seal segment  64  but because the plate  80  is not rigidly mounted directly to the retractable air seal segment  64 , the puller  80  has minimal—if any—effect upon the numerous cooling air passages  64 P. The plate  80  interfaces with the respective air seal segment  64  and also reduces the radial tolerance stack to permit the puller  68  to support at least a portion of a radial load when the respective air seal segment  64  are in the circumferentially contracted position ( FIG. 4 ). 
     Each rod  82  may extend through an engine case  84  to an actuator  86  (illustrated schematically) that operates in response to a control  88  (illustrated schematically). The actuator  86  may include a mechanical, electrical and/or pneumatic drive that operates to move the rod  82  along a rod axis W so as to contract and expand the radially adjustable blade outer air seal system  60 . It should be appreciated that various other control components such as sensors, actuators and other subsystems may be utilized herewith. 
     The control  88  generally includes a control module that executes radial tip clearance control logic to thereby control the radial tip clearance relative the rotating blade tips. The control module typically includes a processor, a memory, and an interface. The processor may be any type of known microprocessor having desired performance characteristics. The memory may be any computer readable medium which stores data and control algorithms such as logic as described herein. The interface facilitates communication with other components such as a thermocouple, and the actuator  86 . In one non-limiting embodiment, the control module may be a portion of a flight control computer, a portion of a Full Authority Digital Engine Control (FADEC), a stand-alone unit or other system. 
     In operation, the blade tip rapid response active clearance control system  58  may utilize, for example, an actuator  86  that provides about 1200-1400 pounds (544-635 kilogram) of force to provide a radial displacement capability for the array of air seal segments  64  of about 0.040″ (40 thousandths; 1 mm) in one disclosed non-limiting embodiment. The radial displacement may, at least partially, be a function of the engine core size and the dynamic conditions of the particular engine architecture. 
     The puller  68  of the rapid response active clearance control system  58  provides thermal and aerodynamic isolation from the respective air seal segment  64  and facilitates the use of legacy BOAS cooling schemes. 
     The use of the terms “a” and “an” and “the” and similar references in the context of description (especially in the context of the following claims) are to be construed to cover both the singular and the plural, unless otherwise indicated herein or specifically contradicted by context. The modifier “about” used in connection with a quantity is inclusive of the stated value and has the meaning dictated by the context (e.g., it includes the degree of error associated with measurement of the particular quantity). All ranges disclosed herein are inclusive of the endpoints, and the endpoints are independently combinable with each other. It should be appreciated that relative positional terms such as “forward,” “aft,” “upper,” “lower,” “above,” “below,” and the like are with reference to the normal operational attitude of the vehicle and should not be considered otherwise limiting. 
     Although the different non-limiting embodiments have specific illustrated components, the embodiments of this invention are not limited to those particular combinations. It is possible to use some of the components or features from any of the non-limiting embodiments in combination with features or components from any of the other non-limiting embodiments. 
     It should be appreciated that like reference numerals identify corresponding or similar elements throughout the several drawings. It should also be appreciated that although a particular component arrangement is disclosed in the illustrated embodiment, other arrangements will benefit herefrom. 
     The foregoing description is exemplary rather than defined by the limitations within. Various non-limiting embodiments are disclosed herein, however, one of ordinary skill in the art would recognize that various modifications and variations in light of the above teachings will fall within the scope of the appended claims. It is therefore to be appreciated that within the scope of the appended claims, the disclosure may be practiced other than as specifically described. For that reason the appended claims should be studied to determine true scope and content.