Fan blade interplatform seal

Abstract: An interplatform seal 86 for a fan 14 in an axial gas turbine engine 10 comprises a plurality of elastomeric layers (94) which sandwich a stiffener (98) therebetween. The stiffener may comprise stainless steel mesh. The seal also has a raised triangular portion (88) at its forward end for disposal within an enlarged region of the gap formed by truncating the leading edge (64) of the blade platform, so as to provide a continuous flow surface. The seal is particularly suited for use with modern impact resistant fan blades which have larger interplatform gaps.

Description: The present invention relates to gas turbine engines, and more particularly, to seals interposed between platforms of blades for a fan in the engine. A gas turbine engine, such as a turbofan engine for an aircraft, includes a fan section, a compression section, a combustion section, and a turbine section. An axis of the engine is centrally disposed within the engine, and extends longitudinally through these sections. A primary flow path for working medium flow gases extends axially through the sections of the engine. A secondary flow path for working medium gases extends parallel to and radially outward of the primary flow path. The fan section includes a rotor assembly and a stator assembly. The rotor assembly of the fan includes a rotor disk and a plurality of outwardly extending rotor blades. Each rotor blade includes an airfoil portion, a dovetailed root portion, and a platform. The airfoil portion extends through the flow path and interacts with the working medium gases to transfer energy between the rotor blade and working medium gases. The dovetailed root portion engages the attachment means of the rotor disk. The platform typically extends circumferentially from the rotor blade to a platform of an adjacent rotor blade. The platform is disposed radially between the airfoil portion and the root portion. The stator assembly includes a fan case, which circumscribes the rotor assembly in close proximity to the tips of the rotor blades. During operation, the fan draws the working medium gases, more particularly air, into the engine. The fan raises the pressure of the air drawn along the secondary flow path, thus producing useful thrust. The air drawn along the primary flow path into the compressor section is compressed. The compressed air is channeled to the combustor section, where fuel is added to the compressed air, and the air-fuel mixture is burned. The products of combustion are discharged to the turbine section. The turbine section extracts work from these products to power the fan and the compressor. Any energy from the products of combustion not needed to drive the fan and compressor contributes to useful thrust. Improvements in fan performance depend in many cases in reducing fluid flow leakage at any points in the fan. One of these places is between adjacent blade platforms. A gap typically exists between adjacent blade platforms which may result in fan blade air loss therethrough if an appropriate seal is not provided. The interplatform gap that exists between fan blades is normally a narrow space that must be sealed to prevent leakage recirculation from the blade trailing edge forward and up through the gap into the fan flow path. The seal is typically a thin and narrow rubber strip with one side portion of the seal attached to the underside of one of the fan blade platforms. The other side portion of the seal hangs loose under the gap between an adjacent platform so that when the fan starts to rotate, the seal is urged radially outwardly against the gap by centrifugal force, thereby providing an effective seal. While such seals may be generally effective, they may be unsatisfactory in certain applications. For example, certain fan blades such as those associated with modem impact resistant fan blades, have larger interplatform gaps therebetween. Similarly, other blading configurations may exist which require increased local gaps between platforms. Prior art seals would be unable to effectively seal against leakage or bridge the large gap due to centrifugal forces. The seals would be pushed through the gap and thus be ineffective in sealing. The resultant leakage of fluid flow due to ineffective sealing would enter the fan flow path. This mixing of leakage fluid flow with the fluid in the fan flow path contributes to fan inefficiency. In addition, the interplatform seals have to allow for fan blade maintenance. The seals have to accommodate radial and circumferential motion during assembly and disassembly of fan blades. Thus, the best type of seal is one that allows for ease of maintenance of associated fan blades and prevents leakage recirculation from the interplatform gaps. Prior art seals, though flexible, are not rigid enough to bridge the relatively large interplatform gaps associated with modem impact resistant fan blades. US 4,422,827A discloses thin strips of silicone rubber which are used to seal any gaps between the abutting edges of adjacent rotor blades in a gas turbine engine. GB 1,450,730 discloses a sealing device for use in a high temperature environment such as a coke oven, the device comprising a central core of a powder or fibre material, one or more longitudinal reinforcing elements extending through the core and a covering of an elastomeric material. According to the present invention, there is provided a seal stiffened to reduce fluid flow through large gaps between adjacent blade platforms for a fan in an axial flow gas turbine engine. Large interplatform gaps are associated with modem impact resistant fan blades. Due to the increased gaps, the seal has to withstand centrifugal forces across the large sealing surfaces when the fan rotates. In particular there is provided a seal for reducing fluid flow through the gap between adjacent blade platforms of circumferentially adjacent blades in the fan of an axial flow gas turbine engine, said seal comprising multiple layers of elastomer sandwiching a stiffener therebetween. A primary feature of the present invention is a seal adapted to seal a large gap between platforms of adjacent blades. The seal includes a laminate of materials which strengthens the seal. In accordance with one particular embodiment of the invention, the seal comprises a plurality of layers of an elastomer such as silicone reinforced with fiberglass fabric. Another feature is a seal which includes a stiffening material sandwiched between the elastomeric layers. One or more layers of stiffening material may be used. The stiffening material may be a mesh material, for example a metallic mesh. In accordance with one particular embodiment of the invention, the stiffening comprises a plurality of stainless steel mesh layers. Another feature of preferred aspects of the invention is a seal including a raised portion. A primary advantage of the present invention is the reduction in fluid flow through the interplatform gap between circumferentially adjacent fan blade platforms. Another advantage is the flexibility of the blade platform seal which is non-interfering during radial blade disassembly and assembly. This facilitates fan blade maintenance. Some preferred embodiments of the present invention will now be described, by way of example only, with reference to the accompanying drawings in which: FIG. 1 is a perspective view of an axial flow, turbofan gas turbine engine. FIG. 2 is an isometric view of a blade of prior art for a fan in the engine of FIG. 1 . FIG. 3 is an isometric view of a modified blade for a fan in the engine shown in FIG. 1. FIG. 4. is an isometric view showing the fan blade with an associated seal in accordance with the invention FIG. 5 . is an isometric view of the seal being adapted between two adjacent fan blades. FIG. 6 . is an exploded view of the seal of FIG. 4 . Referring to FIG. 1 , an axial flow, turbofan gas turbine engine 10 comprises of a fan section 14, a compressor section 16, a combustor section 18 and a turbine section 20. An axis of the engine A r is centrally disposed within the engine and extends longitudinally through these sections. A primary flow path 22 for working medium gases extends longitudinally along the axis A r . The secondary flow path 24 for working medium gases extends parallel to and radially outward of the primary flow path 22. The fan section 14 includes a stator assembly 27 and a rotor assembly 28 . The stator assembly has a longitudinally extending fan case 30 which forms the outer wall of the secondary flow path 24. The fan case has an outer surface 31. The rotor assembly 28 includes a rotor disk 32 and a plurality of rotor blades 34 . Each rotor blade 34 extends outwardly from the rotor disk 32 across the working medium flow paths 22 and 24 into proximity with the fan case 30 . Each rotor blade 34 has a root portion 36 , an opposed tip 38 , and a midspan portion 40 extending therebetween. FIG. 2 shows a blade of prior art for a fan in the axial flow gas turbine engine 10 shown in FIG. 1 . The fan blade 34 includes a root portion 44 , a platform portion 46 , and an airfoil portion 48 . Referring to FIG. 3 , an impact resistant fan blade 34 with which the use of a seal of the present invention will be described further below includes a root portion 44, a platform 46 and an airfoil portion 48. The airfoil portion has a leading edge 50 , a trailing edge 52 , a pressure side 54 and a suction side 56 . The airfoil portion is adapted to extend across the flow paths 22, 24 for the working medium gases. The root portion 44 is disposed radially inward of the airfoil portion 48 and it includes a dovetail neck 60 and a dovetail attachment 62. The platform 46 is disposed radially between the airfoil portion 48 and root portion 44. The platform 46 extends circumferentially from the blade. The platform 46 includes a leading edge portion 64 which is forward of the airfoil portion leading edge 50 , a trailing edge portion 66 which is aft of the airfoil portion trailing edge 52. The platform 46 also includes an outer surface 68 defining a flow surface of the flow path and an inner surface 70 which is radially inward of the outer surface. The fan blade 34 includes an undercut 72 which defines a recessed area so that if the fan blade fractures the fracture is located within the dovetail neck 60 . The undercut 72 is located in the inner surface 70 of the platform and extends into the dovetail neck 60 in the root portion 44 . This undercut 72 moves the fillet radius between the inner surface 70 of the platform 46 and the dovetail neck 60 circumferentially away from the following blade. As a result, when the platform 46 fractures, the edge of the fracture is located within the dovetail neck 60 in the root portion 44. The fan blade 34 as illustrated in FIG. 3 also includes a groove 74 on the outer surface 68 of the platform 46 which is axially and circumferentially coincident with the fillet radius between the inner surface 70 of the platform 46 and dovetail neck 60 within the undercut 72. The groove 74 is a weakened area which ensures that the fracture of the platform 46 occurs at the groove 74. In addition, the leading edge of the dovetail neck 60 in the root portion 44 includes a spanwise chamfer 76 which blunts the forward corner of the dovetail neck 60. The chamfer 76 provides for a blunted comer that upon impact on the leading edge of the following blade airfoil 50 will not cause damage to the airfoil 48 . Referring to FIG. 3 , the leading edge 64 of the platform is truncated 78 to provide for a blunt comer. The truncation 78 further minimizes the risk of damage to the leading edge 50 of the following blade airfoil 48 in the event the leading edge comer impacts the airfoil 48 . In addition, the platform 46 is circumferentially dimensioned to define, with an adjacent platform, a large gap. This gap defines the proximity of adjacent blade platforms. An increased gap reduces the possibility of platform edges of the following adjacent blade contacting those of the released blade during a blade loss condition. The contact between adjacent platform edges causes damage to the platforms 46 which can result in fracturing the following blade platform 46 . Further, the airfoil leading edge 50 is thickened at a radial distance from the platform where the airfoil portion 48 is most likely to be impacted by a disassociated blade. The enhanced thickness is defined by a recess 51 in the leading edge at a radially inner location which provides for a stronger leading edge. FIG. 4 illustrates a seal 86 associated with the fan blade 34. The seal 86 is generally elastomeric. The seal is adapted to seal the locally large gap between platforms of adjacent blades. The seal includes an upstanding or raised portion 88 which is adapted to seal the gap defined by the truncation 78 in the leading edge 64 of the platform 46. Referring to FIG. 5 , the seal 86 is interposed between two adjacent fan blade platforms 46. The seal has a radially outer major surface. The outer surface includes two opposed side portions. One side portion of the elastomeric seal 86 is fixed to the inner surface 70 of one platform 46 such as by adhesive bonding. The second side portion of the seal 86 hangs loose in the interplatform gap defined by the space between two adjacent fan blade platforms 46. FIG. 6 shows an exploded view of the seal 86 of the present invention shown in FIG. 4. The seal has a forward portion 90 and longitudinal aft portion 92. The forward portion 90 seals the leading edge region 64 of the platform 46. The longitudinal aft portion 92 seals the remaining interplatform gap. The forward portion 90 comprises of a plurality of layers of silicone rubber 94 reinforced with fiberglass fabric. Sandwiched between the elastomeric layers is a plurality of layers of stainless steel mesh 98 . The particular embodiment shown in FIG. 6 includes four (4) layers of silicone rubber 94 reinforced with fiberglass fabric and two (2) layers of stainless steel mesh 98 embedded therebetween. The longitudinal aft portion 92 of the seal is comprised of a plurality of layers of silicone rubber 94 reinforced with fiberglass fabric. The particular embodiment shown in FIG. 6 includes two (2) layers of silicone rubber 94 reinforced with fiberglass fabric. During operation of the gas turbine engine, the working medium gases are compressed in the fan section 14 and the compressor section 16 . The gases are burned with fuel in the combustion section 18 to add energy to the gases. The hot, high pressure gases are expanded through the turbine section 20 to produce thrust in useful work. The work done by expanding gases drives rotor assemblies in the engines, such as the rotor assembly 28 extending to the fan section 14 about the axis of rotation A r . The gases flow along the working medium flow path at high velocities into the rotor assembly in the fan section. As the rotor assembly is rotated at high velocities, the fan blades travel at high velocities about the axis of rotation and the working medium gases are compressed in the fan flow path. As a result, the pressure at the aft trailing edge 66 of the fan blade platforms is higher than that at the forward leading edge 64 . The fluid flow from the blade platform trailing edge 66 recirculates forward and up through the interplatform gap into the fan flow path. This recirculation is minimized by the interplatform gap seal 86 of the present invention. One side portion of the radially outer surface of the seal is bonded to the inner surface 70 of a platform 46 . During fan operation, the second opposed side portion of the radially outer surface of the seal is urged radially outwardly against the gap between an adjacent platform, thereby providing an effective interplatform gap seal. The seal is effective for exaggerated interplatform gaps associated with modem impact resistant fan blades having relatively narrow platforms. In the preferred embodiment, the gap between platforms can be increased up to 19 mm (0.75 inches). This represents an increase in the interplatform gap of up to twelve (12) times over that of prior art gaps for a given radial location of seal and fan rotational speed. The measure of seal capability is related to how big a gap the seal has to bridge and therefore seal for a given centrifugal force. The aforementioned radial location of seal and fan rotational speed provide for a measure of the centrifugal forces the seal has to withstand. The stiffening material, such as the stainless steel mesh 98 in the preferred embodiment reinforces the seal. This is important when the interplatform gap is increased as the seal is able to withstand the centrifugal forces due to fan operation. In addition, the stainless steel mesh 98 will not damage the engine in the unlikely event the seal disassociates from a blade platform. In addition, the stainless steel mesh provides for the flexibility required by the seal to facilitate the assembly and disassembly of fan blades. The seals accommodate radial and circumferential motion during fan blade maintenance. Although the invention has been shown and described with respect to detailed embodiments thereof, it should be understood by those skilled in the art that various changes in form and detail thereof may be made without departing from the scope of the claimed invention.

Claims: A seal (86) for reducing fluid flow through the gap between adjacent blade platforms (46) of circumferentially adjacent blades (34) in the fan (14) of an axial flow gas turbine engine (10), characterised in that said seal (86) comprises multiple layers (94) of elastomer sandwiching a stiffener (98) therebetween. A seal according to claim 1 , wherein the stiffener (98) comprises a plurality of layers of stainless steel mesh. A seal according to claim 1 or 2 , wherein each of the elastomeric layers (94) is reinforced with fiberglass fabric embedded therein. A seal according to any preceding claim wherein the stiffener (98) is provided at a forward end only of the seal (86). A seal according to any preceding claim wherein the forward end of the seal (86) has a raised portion. A seal as claimed in any preceding claim, said seal further comprising a raised triangular portion (88) at its forward end for disposal within an enlarged region of said gap defined by truncating a said platform leading edge (64) such that when said seal is attached to said platform (46) the seal provides for a continuous surface for fluid flow in the outer surface (68) of the platform (46). A gas turbine fan (14) comprising a seal as claimed in any preceding claim. A gas turbine fan (14) according to claim 7 wherein the seal is attached to the underside of the platform (46) of one blade (34) and extends under the platform (46) of an adjacent blade (34) so as to be forced into sealing contact with that surface as said fan rotates in use A gas turbine fan (14) according to claim 7 or 8, the fan including a blade array with a plurality of radially extending and circumferentially spaced blades (34), each blade (34) having a platform (46) including an outer surface (68) defining a surface for fluid flowing thereover, an inner surface (70) radially inwardly of the outer surface (68) , adjacent platforms defining the said gaps therebetween, each platform including a truncated leading edge (64).

Inventors: KASPROW ROBERT F, KURZ PHYLLIS L, LESHANE JEFFREY S, KASPROW, ROBERT F., KURZ, PHYLLIS L., LESHANE, JEFFREY S.
Applicants: UNITED TECHNOLOGIES CORP, UNITED TECHNOLOGIES CORPORATION
Application year: 1998
Main classification: F01D11
IPC codes: F04D29, F02C7, F02K3, F01D11