Patent Description:
Gas turbine engines are well known devices, having a plurality of rotating elements including a fan, a compressor, and a turbine, as well as a combustor and other components. The fan is rotated to draw ambient air into the engine and accelerate that air, while the compressor is rotated to compress the air entering the engine. The compressed air is then mixed with fuel and combusted in the combustor creating an exhaust which exits the engine as it expands through the turbine. As the exhaust expands through the turbine, the turbine rotates. This rotational motion is transferred via an engine shaft to the compressor and fan causing both to also rotate.

In an effort to reduce noise, emissions, and fuel burn new aircraft and engines have been developed. These new aircraft and engines designs, however, increase the ingested distortion in a boundary layer of low energy air formed on the surfaces of the aircraft or engine. These distortions are caused by an increase in thickness of the typically thin boundary layer or by an increase in the amount of this low energy air that enters the engine due to placement of the engines on the aircraft. This low energy air reduces engine efficiency and thus is undesirable. These distortions can either occur throughout the operation of the engine or aircraft, or under specific circumstances such as, but not limited to, take-off, landing, or where wind is moving in a lateral direction to that of the engine. Therefore, new features must be developed to address issues arising from this increased ingested boundary layer distortion, including high cycle fatigue and resonant stresses.

The <CIT> describes a typical fan blade utilized by gas turbine engines. While effective, such fan blades do little to mitigate the new boundary layer ingested distortions, and thus new fan blades that address these difficulties are needed.

<CIT> discloses an airfoil having a reduced thickness portion near the blade tip.

In accordance with one aspect of the disclosure, an airfoil is provided as defined by claim <NUM>.

In another refinement, the root of the blade may include about twenty-five percent of a radial height of the blade.

In another refinement, the airfoil may further include a transition zone between the tip and the root of the blade. The transition zone may be aerodynamically smooth.

In accordance with another aspect of the present disclosure a fan of a gas turbine engine is provided as defined by claim <NUM>.

In accordance with yet another aspect of the present disclosure, a gas turbine engine is provided as defined by claim <NUM>.

These and other aspects and features of the present disclosure will be better understood in light of the following detailed description when read in light of the accompanying drawings.

It should be understood that the drawings are not necessarily to scale and that the disclosed embodiments are sometimes illustrated diagrammatically and in partial views. In certain instances, details which are not necessary for an understanding of this disclosure or which render other details difficult to perceive may have been omitted. It should be understood, of course, that this disclosure is not limited to the particular embodiments illustrated herein.

Referring now to the drawings and with specific reference to <FIG>, a gas turbine engine and, more specifically, a turbofan type gas turbine engine is depicted and generally referred to by a reference numeral <NUM>. While <FIG> is depicted as a turbofan engine for an aircraft, it should be understood that this is in no way limiting, but only for ease of illustration, in that any gas turbine engine is possible. Further, the structure and function of a gas turbine engine <NUM> are well known in the art and as such only a limited description will be provided herein.

The engine <NUM> is depicted in <FIG> as including a plurality of components axially aligned along a central axis <NUM>. Such components include a fan <NUM>, a dual-spool compressor <NUM> downstream, that is to say next in line with respect to the flow of air through the engine <NUM>, from the fan <NUM>, a combustor <NUM> downstream from the compressor <NUM>, and a dual-spool turbine <NUM> downstream from the combustor <NUM>. While the engine <NUM> of <FIG> is depicted as a dual-spool engine, it is to be understood that any configuration is possible, such as, but not limited to, single or triple spool configurations.

The fan <NUM> includes a plurality of airfoils <NUM> engaged with a hub <NUM> of the fan <NUM> and rotating about the central axis <NUM> and is surrounded by a fan case <NUM>. As best shown in <FIG>, each of the airfoils <NUM> has a platform <NUM> and a blade <NUM> radially extending outward from the platform <NUM>. The blade <NUM> of the airfoil <NUM> has a root <NUM> proximate the platform <NUM> of the airfoil <NUM> and a tip <NUM> radially outward from the root <NUM> of the blade <NUM>. The blade <NUM> also has a leading edge <NUM>, which interacts with incoming air before other surfaces of the blade <NUM>, and a trailing edge <NUM>, which interacts with outgoing air before the air enters the compressor <NUM> or by-passes the rest of the engine <NUM>. As can be seen in <FIG>, the root <NUM> may have a thickness T1, while a typical cross-section radially outward from the root <NUM>, such as at a quarter-span or greater, may have a thickness T2, with T2 being less than T1. The blade <NUM>, as seen in <FIG>, may have an aerodynamically smooth transition zone <NUM> from the tip <NUM> to the root <NUM> as well.

One benefit of providing the root <NUM> with an increased thickness is that it reduces stress on the airfoil <NUM> as a whole, such as static, vibration, and tensile stresses, associated with the increased boundary layer ingested distortion produced by new gas turbine engine <NUM> and aircraft designs. This is accomplished by distributing the stresses over a larger area provided by the greater thickness T1 of the root <NUM>. The smooth transition zone <NUM> also reduces air drag associated with the greater thickness T1 of the root <NUM> as opposed to a right angle transition from the greater thickness T1 of the root <NUM> to the lesser thickness T2 at around a quarter-span of the airfoil <NUM>.

In the embodiment presented in <FIG>, the root <NUM> is about twenty-five percent of a radial height <NUM> of the blade <NUM>. However, this is only one exemplary embodiment, and the root <NUM> of the blade <NUM> may be modified to be a greater or lesser percentage of the radial height <NUM> of the blade <NUM> to alter the stress reduction qualities of the root <NUM>, modify the mass distribution of the blade <NUM>, as well as to further alter the air flow distribution on the blade <NUM> as desired.

The overall stress capacity of the blade <NUM> and the air flow distribution on the blade <NUM> may be modified by modifying the thickness T1 of the root <NUM>. In one embodiment, shown in <FIG>, the thickness T1 of the root <NUM> is about twenty percent larger than the thickness T2 at about a quarter-span of the blade <NUM>. However, this is only one exemplary embodiment, and the thickness T1 of the root <NUM> may be greater or less than twenty percent larger than the thickness T2 at about a quarter-span of the blade <NUM>.

Referring now to <FIG>, the root <NUM> may also have a fillet <NUM> where the blade <NUM> joins the platform <NUM>, as opposed to a right-angle transition between the platform <NUM> and the blade <NUM> as with the airfoil <NUM> of <FIG>. The fillet <NUM> of <FIG> further increases the ability of the airfoil <NUM> to resist stresses, tensile and bending stresses for example, from the increased boundary layer ingested distortion by providing a supporting structure proximate the platform <NUM>. A fillet <NUM> having an elliptical shape may provide increased reductions in the bending stresses, however, other shapes for the fillet <NUM> are also possible. The fillet <NUM> also provides an aerodynamically smooth transition zone between the platform <NUM> and blade <NUM> of the airfoil <NUM> reducing air drag on the airfoil <NUM>.

While the fillet <NUM> may have any desired circumferential width <NUM> and radial height <NUM>, the exemplary embodiment presented in <FIG> illustrates the fillet <NUM> with an elliptical shape where the height <NUM> is half of a major axis and the width <NUM> is half of a minor axis. <FIG> illustrates another embodiment where the fillet <NUM> has a width <NUM> that varies with an axial length <NUM> of the blade <NUM>. As can be seen from <FIG>, the leading edge <NUM> of the blade <NUM> is located at the line marked A-A. The width <NUM> of the fillet <NUM> is at a minimum at the leading edge <NUM> and gradually increases in a leading portion <NUM> of the blade <NUM> to a maximum at a line marked B-B. The width <NUM> of the fillet <NUM> may remain constant in a central portion <NUM> of the blade <NUM> from the line B-B to a line marked C-C. In a trailing portion <NUM> of the blade <NUM> extending from the line C-C to the trailing edge <NUM> at the line D-D the width <NUM> of the fillet <NUM> gradually decreases until reaching a minimum at the trailing edge <NUM>. The narrower width <NUM> at the leading portion <NUM> may allow the air to be split around the blade <NUM> easier than a wider width <NUM> or a blunt surface. The constant larger width <NUM> of the central portion <NUM> may allow the air to flow smoothly along the joint between the blade <NUM> and platform <NUM>. The narrower width <NUM> of the trailing portion <NUM> may allow the air to recombine in controlled manner without creating pockets of swirling air downstream from the airfoil <NUM>. Modifying this geometry may allow for the splitting and recombining features of the fillet <NUM> to be modified to provide a desired air flow profile as well as to further modify the enhancements provided by the greater thickness T1 of the root <NUM>.

In one embodiment, the trailing and leading portions <NUM>, <NUM> are each twenty percent of the axial length <NUM> of the blade <NUM>. However, other values are also possible and may be used to create a preferential air flow distribution on the blade <NUM> of the airfoil <NUM>. Modifying this percentage may allow tailoring of the growth rate of the width <NUM> of the fillet <NUM> further modifying the splitting and recombining features provided by the fillet <NUM>.

In another embodiment the width <NUM> of the fillet <NUM> has a minimum which is a third of the maximum. However, this is only one exemplary embodiment, and the minimum width may be altered to further modify the root <NUM> of the blade <NUM> to create a preferential air flow distribution on the blade <NUM>. This ratio between the minimum and maximum width of the fillet <NUM> may be modified to further enhance the splitting and recombining features provided by the fillet <NUM>. A smaller ratio may increase the air splitting and recombining potential of the fillet <NUM> but may structurally weaken the fillet <NUM> as the fillet becomes to small at the leading and trailing edge <NUM>, <NUM>. A larger ratio, on the other hand, may increase the structural integrity of the leading and trailing edges <NUM>, <NUM> but may reduce the air splitting and recombining potential of the fillet <NUM> as the leading and trailing edges <NUM>, <NUM> become blunt.

While the preceding description has been directed towards an airfoil for a fan, one skilled in the art will see that the present invention may also be used in conjunction with any other airfoil such as, but not limited to, rotor or stator airfoils for compressors or turbines.

From the foregoing, it can be seen that the technology disclosed herein has industrial applicability in a variety of settings such as, but not limited to, reducing the effects of increased boundary layer ingested distortion caused by new gas turbine engine and aircraft technology. This may be accomplished by increasing the thickness of the root of the blade with respect to the tip of the blade of each airfoil, specifically for fan airfoils. Additionally, the radial length of the root may be modified to further reduce these stresses and increase redistribution of the air flow on the blade. A fillet may also be formed between the blade and the platform for further reductions and to decrease air drag on the airfoil. These modifications and additions to the root of airfoil is also applicable to compressor and turbine rotor and stator airfoils as well.

Claim 1:
An airfoil (<NUM>), comprising:
a platform (<NUM>); and
a blade (<NUM>) extending from the platform (<NUM>), wherein the blade (<NUM>) has a root (<NUM>) and a tip (<NUM>) each radially outward from the platform (<NUM>), the root (<NUM>) being located between the tip (<NUM>) and the platform (<NUM>) and the blade (<NUM>) has a radial height (<NUM>) from the platform (<NUM>), wherein a cross-sectional thickness (T1) of the blade (<NUM>) at the root (<NUM>) is greater than a cross-sectional thickness (T2) of the blade (<NUM>) along at least <NUM>% of the radial height (<NUM>) of the blade (<NUM>); and
further including a fillet (<NUM>) joining the blade (<NUM>) with the platform (<NUM>) of the airfoil (<NUM>), wherein the fillet (<NUM>) has a width (<NUM>) that varies along an axial length (<NUM>) of the blade (<NUM>); and
wherein the blade (<NUM>) includes a leading edge (<NUM>), a central portion (<NUM>), and a trailing edge (<NUM>); the leading edge (<NUM>) interacting with incoming airflow before other surfaces of the blade (<NUM>), the trailing edge (<NUM>) interacting with outgoing airflow, and the central portion (<NUM>) extending between the leading and trailing edges (<NUM>,<NUM>); wherein the width of the fillet (<NUM>) at the leading edge (<NUM>) and the trailing edge (<NUM>) of the blade is less than the width of the fillet (<NUM>) at the central portion (<NUM>).