Structural stator airfoil

An airfoil comprises a leading edge, a trailing edge, and pressure and suction surfaces defined therebetween. An inner platform is coupled to a root section of the airfoil, with a curvilinear gusset extending along a lower surface of the inner platform, opposite the root section. An outer platform is coupled to a tip section of the airfoil, with a multi-lobed gusset extending along an upper surface of the inner platform, opposite the tip section.

BACKGROUND

This invention relates generally to turbomachinery, and specifically to stator vanes for the compressor, turbine or fan section of a gas turbine engine. In particular, the invention concerns a stator airfoil with structural features.

Gas turbine engines provide reliable, efficient power for a wide range of applications, including aviation and industrial power generation. The turbine engine is built around a power core made up of a compressor, combustor and turbine, arranged in flow series with an upstream inlet and downstream exhaust.

The compressor compresses air from the inlet, which is mixed with fuel in the combustor and ignited to generate hot combustion gas. The turbine extracts energy from the expanding combustion gas, and drives the compressor via a common shaft. Energy is delivered in the form of rotational energy in the shaft, reactive thrust from the exhaust, or both.

Small-scale gas turbines generally utilize a one-spool design, with co-rotating compressor and turbine sections. Larger-scale combustion turbines, jet engines and industrial gas turbines (IGTs) are typically arranged into a number of coaxially nested spools, which operate at different pressures and temperatures, and rotate at different speeds.

The individual compressor and turbine sections in each spool are subdivided into a number of stages, which are formed of alternating rows of rotor blade and stator vane airfoils. The airfoils are shaped to turn, accelerate and compress the working fluid flow, and to generate lift for conversion to rotational energy in the turbine.

Aviation applications include turbojet, turbofan, turboprop and turboshaft configurations. Turbojets are an older design, in which thrust is generated primarily from the exhaust. In turbofan and turboprop engines, the typical configurations for modern fixed-wing aircraft, the low pressure spool is coupled to a propulsion fan or propeller. Turboshaft engines are used on rotary-wing aircraft, including helicopters.

In turbofan engines, the fan rotor typically operates as a first stage compressor, or as the pre-compressor stage for a low-pressure compressor or booster module. This design poses additional structural constraints on the engine, because the fan is coupled to both the core and bypass flowpaths, and the fan duct must be rigidly supported from the power core.

SUMMARY

This invention concerns a stator vane for a gas turbine engine. The vane includes an airfoil section coupled to inner and outer platforms. The airfoil section extends in a chordwise direction from a leading edge to a trailing edge, and in a spanwise direction from a root portion adjacent the inner platform to a tip portion adjacent the outer platform.

The inner platform includes a curvilinear gusset extending opposite the root portion of the airfoil. The outer platform includes a multi-lobed or multi-legged gusset, extending opposite the tip portion of the airfoil.

DETAILED DESCRIPTION

FIG. 1is a cross-sectional view of gas turbine engine10, in a two-spool turbofan configuration for use as a propulsion engine. In this particular example, low spool12includes low pressure compressor (LPC)14and low pressure turbine (LPT)16, rotationally coupled via low pressure shaft18. High spool20includes high pressure compressor (HPC)22and high pressure turbine (HPT)24, rotationally coupled via high pressure shaft26. Combustor28is arranged in flow series between high pressure compressor22and high pressure turbine24, with low and high spools12and20coaxially oriented about engine centerline (or turbine axis) CLof gas turbine engine10.

Nacelle30is coaxially oriented about the forward end of the power core, with fan casing31A extending from fan stage32to fan exhaust duct31B, downstream of fan exit guide vane (FEGV)33. Fan shaft34is rotationally coupled to fan stage (or fan rotor)32, generating propulsive flow P through fan duct (or bypass duct)35. In advanced engine designs, fan drive gear system36is used to couple fan shaft34to low spool12, with ring, sun, star and planetary gear mechanisms38,40and42to provide independent fan speed control for reduced noise and improved operating efficiency.

As shown inFIG. 1, fan exit guide vanes33provide structural support for nacelle30, fan casing31A and fan exhaust duct31B. Fan exit guide vanes33also remove swirl from propulsive flow P downstream of fan rotor32, increasing thrust. This approach eliminates the need for separate aerodynamic struts, reducing engine size and weight for improved performance and efficiency. The configuration of vanes33also allows gas turbine engine10to be smaller and lighter, reducing fuel use and environmental impact.

At the same time, airfoil loading is high during operation of gas turbine engine10, imposing substantial structural demands on vanes33. This issue is addressed via a multi-function part design, as applied to fan exit guide vanes33and other circumferentially arranged stator vane assemblies for gas turbine engine10.

The primary functions of fan exit guide vanes33include directing flow in bypass duct35, and providing mechanical connections to transfer loads between adjacent engine structures. In particular, vanes33provide a basic load path for all operating conditions of gas turbine engine10, including static structural loads, flight-induced loads and vibratory loads on nacelle30, fan casing31A and fan exhaust duct31B, and on the airfoil sections of fan exit guide vanes33themselves.

Vanes33also provide integrated features that serve multiple functions, including position control for nacelle30to maintain clearance between fan casing31A and fan rotor32during flight operations, and thrust reverser support and alignment in the downstream section of fan exhaust duct31B. Fan exit guide vanes33also provide an electrically conducting flow path between nacelle30and the power core, and include mounting structures designed for more efficient maintenance of gas turbine engine10, as described below.

FIG. 2is perspective view of vane50for a gas turbine engine, for example fan exit guide vane33for turbofan engine10ofFIG. 1. Alternatively, vane50operates as a stator airfoil for the fan, compressor or turbine section of a different gas turbine engine, in any of the designs and configurations described above.

Vane50includes airfoil portion52with root section54coupled to inner platform56at fillet F. Airfoil52extends in a spanwise direction from root section54to tip section58, with tip section58coupled to outer platform60.

Inner and outer platforms56and60are used to mechanically attach or mount vane50within an annular flow structure, for example a fan duct, compressor duct, turbine duct or transition duct. When vane50is mounted for operation in a gas turbine engine, ID platform56is positioned at a radially inner diameter (ID), and attached using bolts or similar devices intended for that purpose. Outer platform60is positioned at a radially outer diameter (OD), and attached with additional bolts or mechanical elements.

Alternatively, vane50is configured as a shrouded rotor blade, with rotating inner and outer platforms56and60. In these designs, vane50is typically referred to as a blade, and airfoil portion52is referred to as a blade airfoil. Vane50may also be configured as an unshrouded rotor blade or cantilevered stator vane, with root section54attached to a rotating or stationary inner or outer platform56or60, and extending to an unshrouded tip section58. Thus, the terms blade and vane are both be used, depending on whether airfoil section52is configured for operation as a stationary (stator vane) element, or a rotating (rotor blade) element.

Longitudinal axis L extends in a substantially radial or spanwise direction along the blade midchord (half of axial chord S), from root section54to tip section58of airfoil52. As shown inFIG. 2, longitudinal axis L is canted or swept at angle θ with respect to turbine axis CL, and angled respect to the planes of attachment defined along ID and OD platforms56and60, at the inner and outer limits of longitudinal axis L.

Note that longitudinal axis L may or may not fall along one or more maximum camber points M, as defined for individual camber lines C. Camber lines C give the curvature of airfoil52, as a function of span along longitudinal axis L. Each camber line C extends in an axial or chordwise sense between leading edge62and trailing edge63, along the mean line of the blade profile, halfway between concave (or pressure) surface64and convex (or suction) surface65, on the front and back sides of airfoil52, respectively.

The orientation of vane50(including cant angle θ) serves at least two functions. First, cant angle θ varies the rotor/stator spacing between the radial extremes of longitudinal axis L; that is, between the upstream rotor stage and root section54of vane airfoil52, at ID platform56, as compared to tip section58, at OD platform60. Second, cant angle θ improves resistance to damage caused by objects ingested or released into the gas flow path, in any form. This includes ingestion of foreign objects at the inlet, leading to foreign object damage or FOD events, and propagation of released objects along the gas flowpath, leading to domestic object damage or DOD events.

Cant angle θ is generally measured as a simple angle from a perpendicular to the engine centerline; that is, in a plane from the radial direction (r) toward turbine axis CL, up to the projection of longitudinal axis L in the radial-axial plane of radius r and turbine axis CL. Alternatively, cant angle θ is measured as a compound angle; that is, from radius r to longitudinal axis L extending out of the radial-axial plane, including exit angle variations ±φ. Typical values for cant angle θ range from 5° to 30°, for example at least 20° to at least 25°, or about 30°. These values of cant angle θ are selected for structural purposes, as described above, and to produce sufficient variation in rotor/stator spacing to attenuate acoustic impulses emitted from the main thrust producing fan or compressor stage, or other upstream compressor or turbine rotor stage.

Cant angle θ also orients the longitudinal flow surfaces formed by ID and OD platforms56and60at an angle relative to longitudinal axis L of airfoil52. This directs working fluid flow axially downstream through the propulsion engine, contributing to increased performance and thrust.

To produce vane50, airfoil portion52and platforms56and60may be machined or manufactured from one or more metal forgings, sheets or stock materials, for example using aluminum or titanium alloy, or a high temperature metal such as a nickel-based or cobalt-based superalloy. Alternatively, vane50can be made from a combination of materials, for example metal and graphite-based materials or other composites.

To reduce weight while retaining strength and structural integrity, internal voids or cavities66may be formed by machining pressure surface64or suction surface65of airfoil52between ribs67. Depending on design, cavities66may be filled with a lightweight material for improved impact resistance, for example aluminum foam, or left hollow. Alternatively, cavities66may be provided with cooling fluid for internal or external cooling, for example in a compressor or turbine section where airfoil section52is subject to high temperature flow.

Cavities66are then covered with one or more panels on either the front or back of airfoil52, in order to define aerodynamically smooth pressure and suction surfaces64and65. Alternatively, cavities66are formed as an insert, which is bonded between pressure surface64and suction surface65as a unit.

Internal cavities66and ribs67are configured to prevent local distortion under severe load and vibratory excitations, while at the same time reducing mass. The configuration of vanes50with cant or sweep angle θ also provides additional attachment area for ID and OD platforms56and60, creating inner and outer flow surfaces with aerodynamic integrity. Vanes50also reduce the effect of forced vibratory excitations, and contribute to the load carrying capacity of the main engine structure.

Airfoil52thus forms a load-carrying aerodynamic structure between pressure surface64and suction surface65, connecting ID platform56to OD platform60along longitudinal axis L. This allows vane50to provide overall structural and vibration stability, while at the same time redirecting flow to improve engine operations, either by tangential action on the fan exhaust to improve thrust output, or by other compressor or turbine flow turning distribution.

To transfer operational and structural loads across the load-bearing structure of airfoil52, ID and OD platforms56and60include ID and OD gussets68and70, respectively. In the particular configuration ofFIG. 2, for example, ID platform56includes curvilinear gusset68, which follows camber line C as defined along root section54of airfoil52, opposite curvilinear gusset68on ID platform56. OD platform60includes multi-lobed or multi-legged OD gusset70, with nexus N defined on the upper surface of OD platform60, opposite maximum camber point M as defined for tip section58of airfoil52.

For each camber line C, maximum camber point M is defined by the greatest (perpendicular) distance between camber line C and axial chord S, as shown inFIG. 2, and as known in the art. Individual camber lines C and maximum camber points M vary between root section54to tip section58, reflecting blade twist and the load distribution along the span of airfoil52, between ID platform56and OD platform60.

The position of maximum camber point M also depends upon variation ±φ in the trailing edge angle of airfoil52, as measured in a circumferential sense about turbine axis CL, at trailing edge63of airfoil52in tip section58. The trailing edge angle varies according to the desired turning provided by vane (or vane class50), and is not uniform from vane to vane, in order to account for different flow turning distributions about engine axis CL.

In fan exit guide vane applications of vane50, for example, one class of vane is configured with a uniform exit angle (that is, with variation φ=0), while other classes of vanes are configured with non-uniform exit angles to direct flow around the engine pylon or other mounting structures. In these and other non-uniform turning configurations, different fan, compressor and turbine vane classes may have exit angle variations ±φ of at least 5° in magnitude, or at least 10° in magnitude or more, depending on engine design and flow configuration.

FIG. 3Ais a detail showing ID platform56of vane50, with curvilinear gusset68. Curvilinear gusset68follows the shape of camber line C along lower (radially inner, or bottom) surface71of ID platform56, where camber line C is defined for root section54of airfoil52, on the opposite side of gusset68with respect to longitudinal axis L.

Curvilinear gusset68includes web portion72and chevron joint or mount portion74. The upper (radially outer) part of web72is attached to lower surface71of ID platform56, and the lower (radially inner) part of web72is attached to chevron joint74.

Typically, web72, chevron joint74and the other components of curvilinear gusset68are attached along structural fillets F, formed either by welding or machining. In some designs, curvilinear gusset68and ID platform56are integrally formed from a stock, sheet or other workpiece, for example by machining the metal or composite material of vane50and airfoil section52, as described above. Alternatively, one or both of ID platform56and ID gusset68are formed separately, and attached via welding or other bonding process.

Axial bolt holes76are formed by drilling or machining chevron joint74, at the aft end of ID platform56on web72. Radial bolt holes78are formed by drilling or machining forward ID mount80, at the forward end of ID platform56on lower surface71.

As shown inFIG. 3A, web portion72of curvilinear ID gusset68follows camber line C of airfoil52, as defined along root section54, opposite bottom surface71of ID platform56. In some designs, web72extends along the aft portion of camber line C, as shown inFIG. 3A, and forward ID mount80is axially spaced from the forward end of ID gusset68. Alternatively, web72extends along substantially the full length of camber line C, and forward ID mount80forms the axially forward portion of ID gusset68.

The curved camber line configuration of ID gusset68more efficiently transfers structural, vibratory and flow-induced loads between airfoil52and ID platform56, with better resistance to stress, strain and fatigue. In particular, ID gusset68provides a primary interfacing structural joint in the aft region of ID platform56, in a curvilinear design configured for self-correction of varying load path vectors.

The curvature of ID gusset68along camber line C also reduces mechanical contacting stress, relative to previous (e.g., linear) designs, and multiplies the effective surface area for load transfer in a similar space envelope. In addition, curvilinear ID gusset68reduces mechanical stress (or distress) at the fixity point, defined here at the radially inner extreme of longitudinal axis L.

FIG. 3Bis a detail showing a bolt hole configuration for ID platform56of vane50. Operational loads are transferred to and from airfoil52by mounting vane50to an engine case or other structure, using axial bolt holes76in chevron joint74, at the axially aft end of ID platform56, and radial bolt holes78in forward ID mount80, at the axially forward end of ID platform56.

The curvilinear geometry of ID gusset68allows ID platform56and the attached structures of vane50to be rotated out of position for access. Combined with the axial/radial bolting configuration of ID platform56, this allows vanes50to be attached at either end for assembly and disassembly, including removal of the fan casing with vanes50remaining attached to the engine core, or removal of the engine core with vanes50remaining attached to the fan casing. The configuration of vane50thus reduces maintenance costs, by not requiring removal of unnecessary engine parts during routine servicing, or during engine disassembly and transport.

As shown inFIG. 3B, axial bolt holes76are uniformly positioned in a circumferential direction along adjacent ID platform56, with uniform angular spacing α1between adjacent centers, as shown inFIG. 3B. Similarly, radial bolt holes78are uniformly positioned in the circumferential direction, with uniform on-center angular spacing α2. Thus, each ID platform56is mounted with the same bolt configuration, regardless of vane class or variations in camber C. This contrast with the OD platform mounting scheme, as described below.

FIG. 4is a detail showing OD platform60for vane50, with multi-lobed OD gusset70. Multi-lobed gusset70is formed on formed on upper (radially outer, or top) surface82of OD platform60, for example in a K-shaped or X-shaped configuration with at least four lobes or legs84and86extending forward and aft from nexus N, as shown inFIG. 4.

Forward lobes (or legs)84of OD gusset70meet aft lobes (or legs)86at nexus N, opposite maximum camber point M of airfoil52. Maximum camber point M is defined along camber line C of airfoil tip section58, opposite top surface82of OD platform60along longitudinal axis L.

Forward legs84of OD gusset70extend axially forward from nexus N to forward bolt holes88, with uniform circumferential spacing α3along forward cross-member90. Aft legs86extend axially rearward from nexus N to aft bolt holes92, with variable spacing DValong aft cross-member94.

In contrast to uniform spacing α3for forward bolt holes88, variable spacing Dvfor aft bolt holes92is defined by angular variation ±φ in the trailing edge angle of airfoil52. ThusFIG. 4illustrates a plurality of vanes50, each selected from a different vane classes with a different trailing edge angle distribution, such that the trailing edge angles differ by ±φ to provide a non-axially symmetric flow turning pattern to accommodate a downstream strut, fairing or pylon, as described above.

Angular variation ±φ is defined in a circumferential sense about the turbine axis, as shown inFIG. 2, above, and is the substantially the same at trailing edge63of airfoil52in tip section58, and at aft bolt holes92on multi-lobed gusset70. In vane classes where trailing edge63is shifted by angle +φ in the circumferential direction, for example, as compared to a uniform circumferential spacing arrangement, aft bolt holes92are shifted by the same angle +φ, for example 0 to +5° or more, or up to +10° or more.

In vane classes where trailing edge63is shifted by angle −φ, aft bolt holes are shifted in the other direction, where the shift has the same typical magnitudes. Thus the relative position of aft bolt holes92with respect to forward bolt holes88is defined by the trailing edge angle of airfoil52, as defined along trailing edge63of airfoil tip section58.

This arrangement keeps aft bolt holes92circumferentially spaced on either side of camber line C for each vane class, but shifted in a circumferential sense with respect to adjacent vanes50in a different class. In addition to structural load-bearing and load transfer benefits, shifting the relative positions of aft bolt holes92thus provides an installation check or foolproof as well, because vanes of one class will not mount with the same aft bolt hole configuration as vanes in another class.

While this invention has been described with reference to exemplary 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 spirit and scope of the invention. In addition, modifications may be made to adapt a particular situation or material to the teachings of the invention, without departing from the essential scope thereof. Therefore, the invention is not limited to the particular embodiments disclosed herein, but includes all embodiments falling within the scope of the appended claims.