Patent Description:
A gas turbine engine includes a compressor section with multiple rows or stages of stator vanes and rotor blades. In a gas turbine engine, the turbine rotor blades drive the compressor and an electric generator to generate electrical power. Some gas turbine engines include a fan positioned forward of the entrance to the compressor. This fan can provide additional propulsion to the gas turbine engine.

During operation, the fan rotates in order to provide propulsion. The fan may create high turning of the airflow and may create tangential or circumferential air flow. As the fan is positioned forward of the compressor, the air will not be flowing in an axial direction into the compressor from the fan. A set of stator blades may be provided at the inlet to the compressor in order to turn the air exiting the fan to an intended direction. This set of stator blades may be referred to as a fan exit stator.

A prior art airfoil is disclosed in <CIT>. <CIT> discloses another prior art airfoil.

From one aspect, the present invention provides an airfoil in accordance with claim <NUM>.

Other features of specific embodiments are disclosed in the dependent claims.

A more complete understanding of the present invention, however, may best be obtained by referring to the detailed description and claims when considered in connection with the drawing figures, wherein like numerals denote like elements.

As used herein, the "forward" and "aft" directions are defined in reference to the predominate flow direction through a gas turbine engine, with air generally flowing from the forward direction toward the aft direction.

In various embodiments and with reference to <FIG>, a gas turbine engine <NUM> is provided. Gas turbine engine <NUM> may be a two-spool turbofan that generally incorporates a fan section <NUM>, a compressor section <NUM>, a combustor section <NUM> and a turbine section <NUM>. Alternative engines may include, for example, an augmentor section among other systems or features. In operation, fan section <NUM> can drive air along a bypass flow-path B while compressor section <NUM> can drive air along a core flow-path C for compression and communication into combustor section <NUM> then expansion through turbine section <NUM>. Although depicted as a turbofan gas turbine engine <NUM> herein, it should be understood that the concepts described herein are not limited to use with turbofans as the teachings may be applied to other types of turbine engines including one-, two- and three-spool architectures.

Gas turbine engine <NUM> may generally comprise a low speed spool <NUM> and a high speed spool <NUM> mounted for rotation about an engine central longitudinal axis A-A' relative to an engine static structure <NUM> via several bearing systems <NUM>, <NUM>-<NUM>, and <NUM>-<NUM>. It should be understood that various bearing systems <NUM> at various locations may alternatively or additionally be provided, including for example, bearing system <NUM>, bearing system <NUM>-<NUM>, and bearing system <NUM>-<NUM>.

Low speed spool <NUM> may generally comprise an inner shaft <NUM> that interconnects a fan <NUM>, a low pressure (or first) compressor section <NUM> and a low pressure (or first) turbine section <NUM>. Inner shaft <NUM> may be connected to fan <NUM> through a geared architecture <NUM> that can drive fan <NUM> at a lower speed than low speed spool <NUM>. Geared architecture <NUM> may comprise a gear assembly <NUM> enclosed within a gear housing <NUM>. Gear assembly <NUM> couples inner shaft <NUM> to a rotating fan structure. High speed spool <NUM> may comprise an outer shaft <NUM> that interconnects a high pressure (or second) compressor section <NUM> and high pressure (or second) turbine section <NUM>. A combustor <NUM> may be located between high pressure compressor <NUM> and high pressure turbine <NUM>. A mid-turbine frame <NUM> of engine static structure <NUM> may be located generally between high pressure turbine <NUM> and low pressure turbine <NUM>. Mid-turbine frame <NUM> may support one or more bearing systems <NUM> in turbine section <NUM>. Inner shaft <NUM> and outer shaft <NUM> may be concentric and rotate via bearing systems <NUM> about the engine central longitudinal axis A-A', which is collinear with their longitudinal axes. As used herein, a "high pressure" compressor or turbine experiences a higher pressure than a corresponding "low pressure" compressor or turbine.

Positioned between fan <NUM> and low pressure compressor <NUM> is a fan exit stator <NUM>. Fan exit stator <NUM> receives air from fan <NUM> and turns the air so that it flows towards low pressure compressor <NUM>. Fan exit stator <NUM> includes at least one airfoil <NUM> stacked around axis Z. Generally, a stator airfoil <NUM> is stationary and does not rotate about axis A-A'. Airfoil <NUM> may be made from, for example, stainless steel, an austenitic nickel-chromium-based alloy such as Inconel® which is available from Special Metals Corporation of New Hartford, New York, USA, titanium, composite materials, and other suitable materials or the like.

Generally, the flow of air travels from A to A', so fan <NUM> is upstream from low pressure compressor <NUM>, high pressure compressor <NUM> is downstream from low pressure compressor <NUM>, etc. Additionally, the direction towards A from A' may be referred to as forward and the direction towards A' from A may be referred to as aft.

The core airflow C may be compressed by low pressure compressor section <NUM> then high pressure compressor <NUM>, mixed and burned with fuel in combustor <NUM>, then expanded over high pressure turbine <NUM> and low pressure turbine <NUM>. Mid-turbine frame <NUM> includes airfoils <NUM> which are in the core airflow path. Turbines <NUM>, <NUM> rotationally drive the respective low speed spool <NUM> and high speed spool <NUM> in response to the expansion.

Gas turbine engine <NUM> may be, for example, a high-bypass geared aircraft engine. In various embodiments, the bypass ratio of gas turbine engine <NUM> may be greater than about six (<NUM>). In various embodiments, the bypass ratio of gas turbine engine <NUM> may be greater than ten (<NUM>). In various embodiments, geared architecture <NUM> may be an epicyclic gear train, such as a star gear system (sun gear in meshing engagement with a plurality of star gears supported by a carrier and in meshing engagement with a ring gear) or other gear system. Gear architecture <NUM> may have a gear reduction ratio of greater than about <NUM> and low pressure turbine <NUM> may have a pressure ratio that is greater than about <NUM>. In various embodiments, the bypass ratio of gas turbine engine <NUM> is greater than about ten (<NUM>:<NUM>). In various embodiments, the diameter of fan <NUM> may be significantly larger than that of the low pressure compressor section <NUM>, and the low pressure turbine <NUM> may have a pressure ratio that is greater than about <NUM>:<NUM>. Low pressure turbine <NUM> pressure ratio may be measured prior to inlet of low pressure turbine <NUM> as related to the pressure at the outlet of low pressure turbine <NUM> prior to an exhaust nozzle. It should be understood, however, that the above parameters are exemplary of various embodiments of a suitable geared architecture engine and that the present disclosure contemplates other turbine engines including direct drive turbofans and turbo shafts.

<FIG> is a cross-sectional view of the portion of <FIG> labeled <NUM>. As illustrated, a front center body duct <NUM> is downstream from fan exit stator <NUM>. Downstream of front center body duct <NUM> and before low pressure compressor <NUM> is an inlet guide vane <NUM>.

Air enters fan exit stator <NUM> from fan <NUM>. Fan exit stator <NUM> turns the air so that it has reduced tangential flow (swirl). Air flows around front center body duct struts and then around inlet guide vanes <NUM> prior to the air entering into low pressure compressor <NUM>.

Fan exit stator <NUM> may include a plurality of airfoils. The airfoils may circumferentially surround the longitudinal axis A-A' illustrated in <FIG>. The airfoils may be designed with high camber to impart high turning of the air - that is, the airfoils may be designed to turn a received airflow at a significant swirl angle, such as, for example, at, near or above <NUM> degrees. The airflow received at fan exit stator <NUM> may have a tangential component which the airfoils turn so that the air flows in the intended downstream direction. The intended direction may be the axial direction (i.e., along the longitudinal axis A-A').

<FIG> illustrates a planar section <NUM> of an airfoil <NUM> of fan exit stator <NUM>. Illustrated on the planar section <NUM> is a suction side <NUM>, a pressure side <NUM>, a leading edge <NUM> and a trailing edge <NUM>. An axis Z' between leading edge <NUM> and trailing edge <NUM>, indicates the chord line. Angle θ indicates an angle between the chordwise direction Z' and the axis of rotation Z.

Planar section <NUM> has a centroid <NUM> that is the center of mass for planar section <NUM>. Centroid <NUM> may be, for example, a center of gravity. Planar section <NUM> may be positioned in space by the three dimensional location of centroid <NUM>. A traditional coordinate system may be used to position section <NUM>, where the Z axis is parallel to the axis of rotation (A-A'), the X axis (illustrated in <FIG>) is the radial direction relative to the Z axis and the Y axis is tangential to the circumference of rotation. The X axis is also referred to as the stacking axis.

Axis Y' is an axis normal to the chord line Z' in the radial direction. Therefore, angle θ exists between axis Y' and axis Y. As utilized herein, geometric dihedral is a lean of airfoil <NUM> along axis Y'. Furthermore, bow is airfoil lean in the tangential direction, i.e., along axis Y. In other words, bow is defined as the angle between the airfoil stacking and the radial direction, in the tangential direction. Positive bow leans a blade toward the airfoil suction surface and can improve the radial pressure distribution and reduce secondary flow effects.

<FIG> illustrates a cross-sectional view of airfoil <NUM> for the purposes of illustrating aerodynamic sweep. As shown in <FIG>, the sweep angle σ at any arbitrary radius is the acute angle between a line <NUM> tangent to leading edge <NUM> of airfoil <NUM> and a plane <NUM> perpendicular to the relative velocity vector Vr. The sweep angle is measured in plane <NUM> which contains both the relative velocity vector Vr and the tangent line and is perpendicular to the plane <NUM>.

Airfoil <NUM> discussed herein includes modifications that increase the robustness of airfoil <NUM> to variations in inlet flow from an upstream fan <NUM> into low pressure compressor <NUM>. Air flow is directed towards fan exit stator <NUM> from fan <NUM>. This airflow is received at fan exit stator <NUM> in a direction significantly different from the desired direction, so fan exit stator <NUM> is a high turning stator. Generally, fan exit stator <NUM> turns air at least <NUM> degrees. However, various embodiments of the present disclosure can be applied to a stator having a lower or higher turn profile. The modifications to airfoil <NUM> can also be applied to rotor blades or stator vanes positioned anywhere else in gas turbine engine <NUM>.

When the air is received from fan <NUM> at fan exit stator <NUM>, air flow can have a component of flow in the tangential direction. Various embodiments of the present disclosure address this substantially tangential air flow so that the air flow is turned to an intended, often axial, direction with minimal losses into low pressure compressor <NUM>. This improvement is achieved by addressing, in various embodiments, the forward sweep of airfoil <NUM>, the increased chord of airfoil <NUM> and/or the bow of airfoil <NUM>. These features can alter the incoming air flow so that tangential airflow may be directed towards the inner diameter edge <NUM> or outer diameter edge <NUM>.

After passing through front center body duct <NUM>, the air flows through inlet guide vane <NUM> and into low pressure compressor <NUM>. The rotors of low pressure compressor <NUM> may be counter rotating. With counter rotating rotors, inlet guide vane <NUM> may turn the air at a large angle in order to account for the counter rotation. Using the fan exit stator <NUM> disclosed herein, fan exit stator <NUM> may turn the air such that less turning is performed by inlet guide vane <NUM>. This can reduce the pressure losses through inlet guide vane <NUM> as inlet guide vane <NUM> will turn air to a lesser degree.

<FIG> illustrates planar section <NUM> of airfoil <NUM> at outer diameter edge <NUM>, midspan portion <NUM> and inner diameter edge <NUM>. <FIG> illustrates two different geometric sweep positions of inner diameter edge <NUM> of airfoil <NUM>. The embodiment illustrated by inner diameter edge <NUM> illustrates an aft or rearward sweep of approximately <NUM> degrees while the embodiment illustrated by inner diameter edge <NUM> illustrates an aft sweep of approximately <NUM> degrees. As illustrated, airfoil <NUM> can be moved along the chord line Z' towards the airflow (indicated by arrow <NUM>) in order to increase sweep. Sweep can also be increased by positioning leading edge <NUM> farther upstream while leaving trailing edge <NUM> in the same position, increasing the chord of airfoil <NUM> as well as the geometric sweep.

<FIG> illustrates a meridional view of airfoil <NUM>. Cross section <NUM> corresponds to inner diameter edge <NUM> with the increased sweep of <NUM> degrees. Cross section <NUM> corresponds to inner diameter edge <NUM> with an aerodynamic sweep of <NUM> degrees.

As illustrated, to increase sweep between inner diameter edge <NUM> and midspan portion <NUM> (aft sweep), the planar sections closer to inner diameter edge <NUM> can be moved forward (upstream) while leaving the portion between midspan portion <NUM> and outer diameter edge <NUM> alone. Also, to increase sweep between inner diameter edge <NUM> and midspan portion <NUM>, leading edge <NUM> can be extended forward (i.e., increased chord) while leaving the portion between midspan portion <NUM> and outer diameter edge <NUM> alone. Increasing the chord length reduces airfoil loading, thus reducing the likelihood of airfoil flow separation under high incidence conditions.

Returning to <FIG>, an increased chord is illustrated near inner diameter edge <NUM> of fan exit stator <NUM>. By increasing the chord, loading is reduced.

<FIG> illustrates axial stacking of airfoil <NUM> across the span of airfoil <NUM>. The axis labeled "% span" represents the spanwise distribution of airfoil <NUM>. Zero percent (<NUM>%) represents inner diameter edge <NUM> of airfoil <NUM> and <NUM>% represents outer diameter edge <NUM> of airfoil <NUM>.

<FIG> illustrates stacking along axis Z (illustrated in <FIG>, <FIG> and <FIG>). In order to determine axial stacking, centroid <NUM> is disposed along axis Z in either the positive or negative direction. Each planar section <NUM> can be stacked based on the coordinate of its centroid <NUM>. The graph illustrated in <FIG> illustrates the position of centroid <NUM> along the Z axis. As illustrated, the axial stacking of airfoil <NUM> increases throughout the span of airfoil <NUM>.

<FIG> illustrates geometric sweep of airfoil <NUM> across the span of airfoil <NUM>. Geometric sweep of airfoil <NUM> is illustrated by the positioning of centroid <NUM> along the Z' axis (parallel to the chord line of airfoil <NUM>). Of note in <FIG>, the geometric sweep stacking almost doubles as the span increases from inner diameter edge <NUM> to outer diameter edge <NUM>. This represents that leading edge <NUM> near inner diameter edge <NUM> is extended, increasing the chord of airfoil <NUM>. Towards outer diameter edge <NUM>, the chord becomes more constant, which is illustrated by the flatter representation of geometric sweep stacking in <FIG>.

<FIG> illustrates tangential stacking, or bow, across the span of airfoil <NUM>. Bow is illustrated by the positioning of centroid <NUM> along the Y axis. Bow represents stacking of airfoil in the tangential direction (Y).

As illustrated in <FIG>, the bow of airfoil <NUM> is not symmetric about the <NUM>% span line. Instead, the bow is decreasing from the inner diameter edge <NUM> to about the <NUM>% span line. Here bow is weighted to be higher near inner diameter edge <NUM> and lower near outer diameter edge <NUM>. Airfoil <NUM> thus would likely induce higher radial flow movement towards inner diameter edge <NUM> than outer diameter edge <NUM>.

<FIG> illustrates geometric dihedral across the span of airfoil <NUM>. Dihedral is illustrated by the positioning of centroid <NUM> along the Y' axis. Dihedral represents lean of airfoil <NUM> in a direction normal to geometric sweep. As illustrated, the dihedral of airfoil <NUM> is decreasing throughout the span, starting slightly above the zero point and gradually decreasing.

<FIG> illustrates the aerodynamic sweep angle of airfoil <NUM>. The aerodynamic sweep angle is determined based on stacking across the Z' axis illustrated in <FIG> relative to air flow stream surface <NUM>. By increasing the chord at leading edge <NUM>, as well as moving airfoil <NUM> planar sections forward in the sweep direction near inner diameter edge <NUM>, the position of leading edge <NUM> of airfoil <NUM> is moved to create aft sweep. The forward positioning contributes to a higher aerodynamic sweep angle. In <FIG>, the aerodynamic sweep angle of airfoil <NUM> is positive throughout the span of airfoil <NUM>. Positive aerodynamic sweep redistributes flow towards inner edge <NUM>.

<FIG> illustrates the total chord length across the span of airfoil <NUM>. As illustrated, the chord of airfoil <NUM> decreases from inner diameter edge <NUM> to outer diameter edge <NUM>. For example, the chord of airfoil <NUM> may be nearly <NUM>% larger at inner diameter edge <NUM> than at outer diameter edge <NUM>. This represents a significant increase in chord at inner diameter edge <NUM> as compared to outer diameter edge <NUM>. It is shown how this change in total chord affects airfoil <NUM> on <FIG>. As illustrated, trailing edge <NUM> of airfoil <NUM> does not necessarily mirror the forward positional shift of leading edge <NUM> from inner diameter edge <NUM> to outer diameter edge <NUM>.

<FIG> illustrates total camber of airfoil <NUM> along its span. Camber represents the change in angle from leading edge <NUM> to trailing edge <NUM>. As illustrated, the camber of airfoil <NUM> is greater than <NUM> degrees throughout the span.

Also illustrated, the camber increases as the span approaches outer diameter edge <NUM>. This can be seen in <FIG>. Outer diameter edge <NUM> has total turning from leading edge <NUM> to trailing edge <NUM> that is higher than the midspan portion <NUM> and inner diameter edge <NUM>.

<FIG> illustrates a perspective view of airfoil <NUM> from suction surface <NUM>. <FIG> also includes the Y, Z, Y' and Z' axes for reference. In <FIG>, the bow of airfoil <NUM> is not symmetric about the <NUM>% span line. The bow is greater at inner diameter edge <NUM> than it is at outer diameter edge <NUM>. This is shown by the curvature of airfoil <NUM> about suction surface <NUM> in the Y direction. Suction surface <NUM> is curved in the positive direction near inner diameter edge <NUM> for <NUM>% span and near outer diameter edge <NUM> for <NUM>% span. This illustrates the change in bow illustrated in <FIG>.

<FIG> illustrates a perspective view of airfoil <NUM> from pressure surface <NUM>. <FIG> illustrates, in an axial projection, how leading edge <NUM> positioning for sweep affects airfoil <NUM>. As illustrated, the position of leading edge <NUM> is substantially forward near inner diameter edge <NUM> for airfoil <NUM>. As illustrated, the chord of airfoil <NUM> is greater at inner diameter edge <NUM> than at outer diameter edge <NUM>. This represents the chord changes in <FIG>. This positioning of leading edge <NUM> contributes to an aerodynamic sweep that is positive throughout the span of airfoil <NUM>.

<FIG> illustrates a perspective view of airfoil <NUM> from leading edge <NUM> and outer diameter edge <NUM>. <FIG> also includes the Y, Z, Y' and Z' axis for reference. Suction surface <NUM> is illustrated. Again, the bow of airfoil <NUM> is illustrated by the curve of airfoil <NUM> from inner diameter edge <NUM> to outer diameter edge <NUM>.

<FIG> illustrates a front view of airfoil <NUM> from leading edge <NUM>, and <FIG> illustrates a rear view of airfoil <NUM> from trailing edge <NUM>. <FIG> also illustrate the bow of airfoil <NUM>.

The combination of features of airfoil <NUM>, in particular the aft sweep, the chord and the asymmetric bow, reduce pressure loss through fan exit stator <NUM> relative to conventional systems. Often, highly loaded flow separates on an endwall adjacent inner diameter edge <NUM>. By incorporating these features, flow is pulled towards inner diameter edge <NUM>, reducing pressure loss and flow defect in the inner diameter edge region.

This combination of features also improve turning in compressor section <NUM>. Without these features, a traditional airfoil would have non-optimal exit angles, resulting in higher tangential velocities. The features of airfoil <NUM> described herein provide appropriate turning of the flow. This optimal turning results in lower tangential velocities, so the flow will be approaching compressor section <NUM> with a swirl angle closer to the intended flow direction.

Claim 1:
An airfoil (<NUM>) for a fan exit stator (<NUM>) of a gas turbine engine, the airfoil (<NUM>) comprising:
an inner diameter edge (<NUM>, <NUM>, <NUM>) having a first chord length;
an outer diameter edge (<NUM>) having a second chord length, wherein the first chord length is larger than the second chord length;
a trailing edge (<NUM>); characterised by
a leading edge (<NUM>) having a positive aerodynamic sweep across substantially an entire span of the leading edge (<NUM>),
wherein a chord of the airfoil (<NUM>) continuously decreases from the inner diameter edge (<NUM>, <NUM>, <NUM>) to the outer diameter edge (<NUM>); and
wherein a bow of the airfoil (<NUM>) is not symmetric around a midspan of the airfoil (<NUM>).