Patent Publication Number: US-2009238683-A1

Title: Vane with integral inner air seal

Description:
BACKGROUND 
     The present invention relates generally to gas turbine engines, and more specifically, to a turbine vane segment with an integral air seal and inner platform. 
     In a gas turbine engine, air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases. Energy is extracted from the combustion gases in the turbine stages, and extracted for powering the compressor and producing work, which is the thrust to power the engine in aircraft applications. 
     The high pressure turbine of the engine includes stationary stators, or vanes, at the combustor exhaust which channels the combustion gases into an axially adjacent rotor, or blades mounted to a supporting rotor disk which in turn drives the compressor during operation. The stators include a row of aerodynamic airfoils extending radially between an inner platform and outer shroud. 
     The stators are supported in the engine either at the outer shrouds, or at the inner platform, and are typically formed in circumferential segments for accommodating thermal expansion and contraction as the hot gases, or working fluid, are discharged from the combustor and between the vanes. The stator components are cooled during operation by using a portion of the pressurized air bled from the compressor in various cooling circuits. The cooling circuits are contained outboard of the outer shroud, or inboard of the inner platform. 
     It is important that the cooling fluid and working fluids do not mix. To prevent mixing in the stators, various seals are incorporated. In current designs, air seals are fastened to the inner platform to prevent any crossover leakage of the working fluid and cooling fluids. The inner platform typically includes radially extending flanges or similar structures which cooperate with adjoining components of the engine for both mounting and sealing the turbine nozzle therewith. Although the structures are not directly exposed to the hot combustion gases of the turbine flowpath, they provide additional weight and thermal mass which affect performance of the engine. Weight is the paramount design feature in an aircraft engine and must be minimized for maximizing efficiency of the engine. Thermal mass affects thermal stresses generated during operation, and also affects the durability and life of the turbine nozzle. 
     SUMMARY 
     A stator vane segment for a gas turbine engine includes at least one airfoil joined to an outer shroud and an inner platform. A sealing element having a first platform land radially inward of the inner platform and an abradable material covering at least a portion of the first platform is integrally joined to the inner platform. 
     In another embodiment, a stator for a gas turbine engine includes a plurality of circumferentially spaced vane segments. Each individual vane segment constitutes at least one airfoil that has a concave pressure side and convex suction side extending between a leading edge and a trailing edge joined to an outer shroud and an inner platform. The outer shroud is secured to a case of the gas turbine engine. Each individual vane segment also constitutes an integral sealing element that has at least one platform radially inward from the inner platform joined to the inner platform. An abradable material covers at least a portion of the at least one platform, and a plurality of knife edges engage the abradable material. 
     In an alternate embodiment, a stator vane segment for a gas turbine engine incorporates at least one airfoil joined to an outer shroud and a first platform. A seal land is integrally joined to the first platform. The sealing element contains a second platform radially inward from the first platform and a third platform radially inward of the second platform. The second platform is connected to the first platform by a forward radial support and an aft radial support, while the third platform is connected to the second platform by a radially extending support. An abradable material covers at least a portion of the second and third platforms. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  is a schematic of a cross-section of a gas turbine engine. 
         FIG. 2  is a partial cross section of the turbine section of the gas turbine engine of  FIG. 1 . 
         FIG. 3  is detailed sectional view of an integral air seal and inner platform of a second stage vane in the turbine section of  FIG. 2 . 
         FIG. 4  is a detailed sectional view of a prior art second stage vane having an inner platform and a separate inner air seal. 
         FIG. 5  is a forward perspective view of a stator vane segment. 
         FIG. 6  is an aft perspective view of the stator vane illustrated in  FIG. 5 . 
         FIG. 7  is a perspective view of an alternate embodiment of a vane segment. 
     
    
    
     DETAILED DESCRIPTION 
       FIG. 1  shows a schematic gas turbine engine  10 , in which a vane segment with an integral inner air seal can be used. Gas turbine engine  10  comprises fan  12 , low pressure compressor (LPC)  14 , high pressure compressor (HPC)  16 , combustor section  18 , high pressure turbine (HPT)  20  and low pressure turbine (LPT)  22 , which are each concentrically disposed around longitudinal engine centerline CL. Fan  12  is enclosed at its outer diameter within fan case  23 A. Likewise, the other engine components are correspondingly enclosed at their outer diameters within various engine casings, including LPC case  23 B, HPC case  23 C, HPT case  23 D and LPT case  23 E such that an annular air flow path is formed around centerline CL. 
     Ambient air A enters fan  12  and it is divided into streams of primary air AP and secondary air AS after it passes through fan  12 . Fan  12  is rotated by low pressure turbine  22  through shaft  24  to accelerate secondary air AS (also known as bypass air) through exit guide vanes  26 , thereby producing a major portion of the thrust output of engine  10 . Shaft  24  is supported within engine  10  at ball bearing  25 A, ball bearing  25 B and roller bearing  25 C. Primary air AP (also known as gas path air) is directed first into low pressure compressor (LPC)  14  and then into high pressure compressor (HPC)  16 . LPC  14  and HPC  16  work together to incrementally step up the pressure of primary air AP. HPC  16  is rotated by HPT  20  through shaft  28  to provide pressurized air to combustor section  18 . Shaft  28  is supported within engine  10  at ball bearing  25 D and roller bearing  25 E. The pressurized air is delivered to annular combustor  18 , illustrated as  18 A and  18 B, along with fuel through multiple injectors, illustrated as  30 A and  30 B, such that a combustion process can be carried out to produce the high energy gases necessary to turn turbines  20  and  22 . Primary air AP continues through turbine sections  20  and  22  whereby it is typically passed through an exhaust nozzle to further produce thrust. 
     HPT  20  includes a plurality of stator stages  32 ,  34 , and a plurality of rotor stages  36 ,  38 . Stator stages  32 ,  34  contain vanes that are cantilevered from HPT casing  23 D, and extend radially inward. Rotor stages  36 ,  38  contain blades that are connected to a rotatable rotor and extend radially outward from shaft  28 . Combustion gases exiting combustor  18  are directed by the shape and angle of HPT turbine vanes to impart force against the rotor blades, thus causing rotor stages  36 ,  38  and shaft  28  to rotate. 
       FIG. 2  is a partial cross section of the HPT  20 , which comprises first stage stator  32 , first rotor stage  36 , second stage stator  34 , and second stage rotor  38 . First stage stator  32  and second stage stator  34  contain first stage vane  40  and second stage vane  44 , respectively, to receive the flow of combustion gases and direct them to first stage blade  42  and second stage blade  46  of rotor stages  36 ,  38 . 
     Second stage stator  34  has second stage vane  44  with an airfoil portion contained between outer shroud  48  and integral air seal and inner platform  58 . Outer shroud  48  is integral with second stage vane  44 , and secures vane  44  to casing  52  in a hook and slot arrangement. Outer shroud  48  may be further supported by adjacent shrouds  54 ,  56  that are attached to casing  52  and lie radially spaced from the edges of first stage blade  42  and second stage blade  46 , respectively. A plurality of like structures extend circumferentially around the centerline CL of the engine to create second stage stator  44 , also commonly referred to as a second stage nozzle, of gas turbine engine  10 . 
     First stage blade  42  has inner platform  60  that extends axially rearward to be adjacent integral platform  58  of second stage vane  44 , and is attached to rotor disk  62 . Similarly, second stage blade  46  has inner platform  64  that extends axially forward to be adjacent integral platform  58  of second stage vane  44 , and is attached to second stage rotor disk  66 . Blades  42  and  46  are secured to rotor disks  62  and  66 , respectively, in part by knife seal mounts  68  and  70 . Knife seal mounts  68  and  70  contain knife edges that interact with corresponding abradable surfaces attached to integral platform  58 . The interaction of the knife edges and abradable surfaces creates a seal that prevents leakage of compressed cooling air disposed radially below the working fluid flowpath of the turbine into the working fluid. 
       FIG. 3  is a detailed sectional view of integral air seal and inner platform  58  of a second stage vane of a gas turbine engine.  FIG. 4  is a detailed sectional view of a prior art second stage vane having an inner platform  58 A and separate inner air seal  59 . 
     In the prior art illustrated in  FIG. 4 , inner air seal  59  has a lower platform  72  connected to radial support  74 , which is attached to another platform radially outward of inner platform  72  and has aft side  76  and forward side  78 . Forward side  78  contains a hook  80  that attaches inner air seal  59  to a corresponding counter-hook  82  that extends radially inward from the bottom of inner platform  58 A. Radial support  74  extends past forward and aft sides  78  and  76  to mate with extension  86 . Extension  86  is generally centrally located on platform  58 A and extends radially inward. Both radial support  74  and extension  86  contain apertures for securing the inner air seal  59  to inner platform  58 A with a fastener  84 , which is commonly a bolt and corresponding nut. An enclosed pocket  88  is created between inner air seal  59  and inner platform  58 A between the bottom surface of shroud  58 A, radial support  74 , forward side  78 , and hook  80  and counter-hook  82 . The radial inner sides of forward and aft sides  78  and  76 , as well as lower platform  72 , contain abradable lands  90 . Both inner air seal  59  and inner platform  58 A contain slots  92  for the reception of a leaf seal (not illustrated) that bridge the gap between adjacent, circumferential segments. 
     Integral air seal and inner platform  58 , illustrated in  FIG. 3 , are integrally connected by upper forward and aft radial supports  94  and  98 , respectively. Aft radial support  98  eliminates extension  86  and fastener  84  contained in the prior art, thus eliminating excess cost, weight and complexity in the engine. Also, the integral design of  FIG. 3  allows for moving the aforementioned support structures (extension  86 , fastener  84 , and radially outer portion of radial support  74 ) afterward from the generally central location resulting in aft radial support  98 . 
     Forward side  96  and aft side  100  of radially upper platform  104  extend past forward and aft radial supports  94  and  98 . Radial support  74  remains generally located near the centers of radially upper platform  104  and lower platform  72 . Radially upper platform  104 , the bottom of the inner platform, and forward and aft radial supports  98  and  94  create a large pocket  102 . With the design of integral air seal and inner platform  58 , additional material and weight can be removed from radially upper platform  104  of each segment compared to the radially outer platform illustrated in  FIG. 4 , as is represented by  102 A. 
     A fluid passage (not illustrated) may extends through inner platform  58  to pocket  102 . The fluid passage allows for cavity on-board injection (COBI). That is, secondary air flow inside pocket  102  is discharged into the turbine section in a direction that is in substantial alignment with the fluid flow path exiting the vane trailing edge. 
     Abradable lands  90  are located on the radial inner surface of lower platform  72 , and on forward and aft portions of the radial inner surface of radially upper platform  104 . Abradable lands  90  may be an annular honeycomb structure that interacts and cooperates with the knife edges of knife seal mounts  68  and  70  to form the seal. The honeycomb is made of a plurality of open cells that face radially inward. The knife edges define with the cells of abradable strip  90  a labyrinth seal which is effective for maintaining the differential pressure across the second stage nozzle. The face of integral air seal and inner platform  58  also contains slots  92 . Flat sealing structures constructed from sheet metal, often referred to as leaf seals or feather seals, bridge the gap between adjacent stator vane segments and are captured by slots  92 . The leaf seals minimize fluid leakage between adjacent stator vane segments. 
       FIG. 5  is a forward perspective view of a typical stator vane segment  44 .  FIG. 6  is an aft perspective view of the stator vane segment  44  illustrated in  FIG. 5 . Stator vane  44  has two airfoils  108 ,  110  contained between outer shroud  48  and integral air seal and inner platform  58 . Airfoils  108 ,  110  each have a generally concave pressure side and generally convex suction side, and contain a leading edge that first encounters the working fluid and a trailing edge adjacent the working fluid exhaust. Airfoils  108 ,  110  may contain simple or compound curvatures for receiving the working fluid. The bottom surface of outer shroud  48  and upper surface of and integral air seal and inner platform  58  create the working fluid flowpath in which airfoils  108 ,  110  are contained. A plurality of like segments may be placed adjacent one another to create second stage stator  34  (see  FIGS. 1 and 2 ). 
     Integral air seal and inner platform  58  has flow path platform  112 , upper platform  104 , and lower platform  72 , which are all generally parallel to one another. Platforms  112 ,  104 , and  72  also contain a slight radial curve to facilitate formation of a ring centered about a centerline CL of gas turbine engine  10  as like segments are placed adjacent one another. The bottom surface of outer shroud  48  is likewise curved and parallel to platforms  112 ,  104 , and  72 , with each decreasing in length from the outer shroud  48  towards the centerline CL of gas turbine engine  10  to create a generally angled profile of stator vane segment  44 . 
     Radial support  74  extends between radially upper platform  104  and lower platform  72 , and in the embodiment illustrated is located generally in the center of the axial direction of platforms  72  and  104 . In an alternate embodiment, not illustrated, radial support  74  may be offset to one side of center in the axial direction of platforms  72  and  104 . 
     The bottom surface of flow path platform  112  is connected to radially upper platform  104  by forward radial support  94  and aft radial support  98 . In this embodiment, forward radial support  94  and aft radial support  98  are located directly below the leading edge and trailing edges, respectively, of airfoils  108  and  110 . This arrangement assures structural integrity of vane segment  44 . Radial support  74  is generally parallel to forward radial support  94  and aft radial support  98 , and radial support  74  is located medially between forward radial support  94  and aft radial support  98 . Pocket  102  is formed by forward radial support  94  and aft radial support  98 , upper platform  104 , and flow path platform  112 . Pocket  102  is a hollow area that may receive additional fluid flow, such as for cooling of vane segment  44 . 
     The underside of lower platform  72  is covered with abradable lands  90 , such as a honeycomb seal material. Similarly, a portion of the underside of upper platform  104  has abradable lands  90  secured thereto, including a first land that covers forward side  96  and aft side  100  of radially upper platform  104 . The geometry of abradable lands  90  may vary, so long as they are able to interact with the knife edges of knife seals  68  and  70  (See  FIG. 3 ). Both outer radial surfaces of integral air seal and inner platform  58  of vane segment  44  also contain slots  92  for the insertion of sealing elements between adjacent vane segments. 
       FIG. 7  is a perspective view of an alternate embodiment of vane segment  44 . In this embodiment, single airfoil  114  is contained between outer shroud  48  and integral air seal and inner platform  58 . Flow direction apertures  116  are present in outer shroud  48  to allow for directing a cooling fluid into airfoil  114 . The stators and rotors of the HPT operate in some of the highest temperature environments within gas turbine engine  10  and thus cooling of the surfaces exposed to the working fluid flow path is necessary. A portion of air from the compressor is bled from the compressor system to provide cooling fluid to the HPT. In one embodiment, some of the cooling fluid will exhaust into pocket  102 . 
     Integral air seal and inner platform  58  has lower platform  72  connected to upper platform  104  by radial support  74 . Upper platform  104  is connected to the bottom surface of inner platform  112  through forward radial support  94  and aft radial support  98 . In this embodiment, forward radial support  94  and aft radial support  98  are angled axially inward towards radial support  74 . Forward side  96  and aft side  100  of radially upper platform  104  terminate prior to reaching the leading edge and trailing edge of airfoil  114 . Lower platform  72  has a greater axial length than radially upper platform  104 . In the embodiment illustrated, lower platform  72  is approximately equal in axial length to inner platform  112 . In an alternate embodiment, lower platform  72  has a different axial length than inner platform  112 . 
     Lower platform  72  and radially upper platform  104  each contain abradable lands  90 A secured to a portion of the upper surfaces. Abradable strips  90 A are located towards the forward and aft sides. This geometry and arrangement of vane segment  44  allows for a different design of knife edge seals (not illustrated) wherein the knife edges extend radially inward to abradable lands  90 A, rather than radially outward as illustrated in  FIG. 4 . Further, the length of the knife edge seals can be adjusted according to the axial length of lower platform  72  and radially upper platform  104 . In the embodiment illustrated, lower platform  72  also has abradable lands  90 B attached to the radially inner surface. This allows for an additional knife edge seal element to be secured directly below vane segment  44 . 
     The shape of pocket  102  as illustrated in  FIG. 7  is generally trapezoidal as compared to the generally rectangular shape illustrated in  FIG. 3 . By adjusting the location and the angle of forward radial support  94  and aft radial support  98 , pocket  102  can be shaped to create a desired fluid flowpath for cooling fluids. 
     Vane segment  44  is made by integrally casting outer shroud  48 , airfoil(s)  108 ,  110 , or  114 , and integral air seal and inner platform  58  as a single structure. The casting is done using conventional molds or dies, and uses a process such as investment casting. In an alternate embodiment, outer shroud  48 , airfoil(s)  108 ,  110 , or  114 , and integral air seal and inner platform  58  are each cast separately. The cast parts are then assembled together, held in position by a fixture so that they can be secured to one another by brazing or welding. 
     Stator vane segment  44  incorporating integral air seal and inner platform  58  results in many advantages over the prior art containing separate structures for the air seal and inner platform. Providing an integral air seal and inner platform  58  removes the joints and other interfaces between the air seal and inner platform. This eliminates the need for pins, bolts, hooks, and similar fasteners. Also, the part count of the stator stage is reduced. Elimination of the interface with a fastener allows for creating a shorter radial distance between the radially upper platform  104  and inner fluid flowpath platform  112 . Thus, the overall height of vane segment  44  is reduced. This saves on material and overall engine weight. Additionally, this allows the abradable lands  90  and corresponding knife edge seals to be moved outboard. 
     Elimination of the interface between air seal  59  and inner platform  58 A (See  FIGS. 3 and 4 ) also allows for creation of a smoother configuration for slots  92  and corresponding feather seals. Potential ingestion of the working fluid into the inboard hardware is reduced with the integral air seal and inner platform  58  as a result of the feather seal configuration, as well as with the corresponding elimination of joints or interfaces. The location of forward radial support  94  and aft radial support  98  can be adjusted to create a desired cooling fluid flow path in pocket  102  without the worry of potential leakage of the cooling fluid. No fasteners or similar extensions into pocket  102  disrupt the flow of the cooling fluid. Similarly, this adjustable location allows for moving of extension  86  (see  FIGS. 3 and 4 ) as there is no need to worry about access to fastener  84  with tools to secure fastener  84 . 
     The overall weight of vane segment  44  is also reduced compared to a separate vane and inner air seal. The elimination of the fastener  84  and rail  86  decreases weight. By moving the position of the aft radial support  98 , additional material can be removed from pocket  102 , as represented by  102 A (see  FIG. 3 ). Material from the radially upper platform above radial support  72  can be removed without affecting the structural integrity of vane segment  44 . The reduction of weight associated with integral platform  58  helps reduce overall fuel burn. 
     Vane segment  44  containing integral air seal and inner platform can be incorporated into single airfoil segments, doublets, quads or more. The more airfoils per segment helps reduce potential leakage. Vane segment  44  designed to minimize leakage and minimize weight while assuring that all structural requirements and heat loads are adequately accounted for. 
     Although the present invention has been described with reference to preferred embodiments, workers skilled in the art will recognize that changes may be made in form and detail without departing from the spirit and scope of the invention.