Patent Description:
A gas turbine engine typically includes a fan section, a compressor section, a combustor section and a turbine section. Air entering the compressor section is compressed and delivered into the combustor section where it is mixed with fuel and ignited to generate a high-speed exhaust gas flow. The high-speed exhaust gas flow expands through the turbine section to drive the compressor and the fan section. The compressor section typically includes low and high pressure compressors, and the turbine section includes low and high pressure turbines.

The shape of an airfoil designed for turbomachinery applications is an important characteristic. It is often a result of multidisciplinary considerations including aerodynamics, durability, structures and manufacturability. However, recent advances in the design of aerodynamically high-performing, high-pressure turbine blades, particularly at the tip, have caused increased difficulties in the design of blades.

<CIT> discloses a prior art turbine blade as set forth in the preamble of claim <NUM>.

<CIT> discloses a prior art fan disk and gas turbine engine.

<CIT> discloses a prior art high order shaped curve region for an airfoil.

<CIT> discloses a prior art turbine airfoil and steam turbine.

<CIT> discloses prior art blading for turbomachines.

<CIT> discloses a prior art gas turbine blade.

In one aspect, there is provided a turbine blade for a gas turbine engine as recited in claim <NUM>.

In another aspect, there is provided a gas turbine engine as recited in claim <NUM>.

A feature of an embodiment of the invention is set forth in the dependent claim.

Features described in connection with one embodiment are applicable to all embodiments, unless such features are incompatible.

The geared architecture <NUM> may be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about <NUM>:<NUM>.

Referring to <FIG>, a cross-sectional view through a high pressure turbine section <NUM> is illustrated. In the example high pressure turbine section <NUM>, first and second arrays of circumferentially spaced fixed vanes <NUM>, <NUM> are axially spaced apart from one another. A first stage array of circumferentially spaced turbine blades <NUM>, mounted to a rotor disk <NUM>, is arranged axially between the first and second fixed vane arrays. A second stage array of circumferentially spaced turbine blades <NUM> is arranged aft of the second array of fixed vanes <NUM>. It should be understood that any number of stages may be used.

The turbine blades each include a tip <NUM> adjacent to a blade outer air seal <NUM> of a case structure <NUM>, which provides an outer flow path. The first and second stage arrays of turbine vanes and first and second stage arrays of turbine blades are arranged within a core flow path C and are operatively connected to a spool <NUM>, for example.

Each blade <NUM> includes an inner platform <NUM> respectively defining inner flow path. The inner platform <NUM> supports an airfoil <NUM> that extends in a radial direction R, as shown in <FIG>. It should be understood that the turbine vanes <NUM>, <NUM> may be discrete from one another or arranged in integrated clusters. The airfoil <NUM> includes a leading edge <NUM> and a trailing edge <NUM>.

The airfoil <NUM> is provided between pressure side <NUM> (predominantly concave) and suction side (predominantly convex) <NUM> in an airfoil thickness direction (<FIG>), which is generally perpendicular to a chord-wise direction provided between the leading and trailing edges <NUM>, <NUM>. Multiple turbine blades <NUM> are arranged in a circumferentially spaced apart manner in a circumferential direction Y (<FIG>). The airfoil <NUM> includes multiple film cooling holes <NUM>, <NUM> respectively schematically illustrated on the leading edge <NUM> and the pressure side <NUM> (<FIG>).

The turbine blades <NUM> are constructed from a high strength, heat resistant material such as a nickel-based or cobalt-based superalloy, or of a high temperature, stress resistant ceramic or composite material. In cooled configurations, internal fluid passages and external cooling apertures provide for a combination of impingement and film cooling. Other cooling approaches may be used such as trip strips, pedestals or other convective cooling techniques. In addition, one or more thermal barrier coatings, abrasion-resistant coatings or other protective coatings may be applied to the turbine vanes <NUM>.

<FIG> schematically illustrate an airfoil including pressure and suction sides joined at leading and trailing edges <NUM>, <NUM>. An attachment or root <NUM> supports the platform <NUM>. The root <NUM> may include a fir tree that is received in a correspondingly shaped slot in the rotor disk <NUM>, as is known. The airfoil <NUM> extends a span from a support, such as an inner platform <NUM> to an end, such as a tip <NUM> in a radial direction R from a radially outer side of the platform <NUM>. The <NUM>% span and the <NUM>% span positions, respectively, correspond to the radial airfoil positions at the support and the end. The leading and trailing edges <NUM>, <NUM> are spaced apart from one another and an axial chord bx length (<FIG>) extends in the axial direction X.

As shown in <FIG>, a radially outer end of the airfoil <NUM> includes a bowed tip portion <NUM>. The bowed tip portion <NUM> is bowed in a direction perpendicular to a mid-camber line <NUM> (<FIG>) of the airfoil <NUM> such that the bowed tip portion <NUM> extends towards the suction side <NUM> in a circumferential and towards the trailing edge <NUM> in an axial downstream direction. The geometry of the bowed tip portion <NUM> results in a non-pointed or rounded bowed tip leading edge <NUM> that reduces vulnerability of a leading edge of the bowed tip portion <NUM> to damaging vibrations and oxidation.

As shown in <FIG>, in one example, the bowed tip portion <NUM> begins at <NUM>% of the span of the airfoil <NUM> and continues to the tip <NUM>. In accordance with the claims, the bowed tip portion <NUM> begins between <NUM>% and <NUM>% of the bowed tip portion.

As shown in <FIG>, the bowed tip portion <NUM> bends in a circumferential direction towards the suction side <NUM> at an angle α to the radial direction. In accordance with the claims, the angle α is <NUM> degrees at the radially innermost part of the bowed tip portion <NUM> and <NUM> degrees at the tip <NUM>. In accordance with the claims, the angle α is between <NUM> and <NUM> degrees. Similarly, as shown in <FIG>, the bowed tip portion <NUM> bends in an axially downstream direction towards at an angle β to the radial direction. In accordance with the claims, the angle β is between <NUM> and <NUM> degrees. Moreover, in accordance with the claims, the bowed tip portion <NUM> is bowed by the same degree in both the circumferential direction and the axially downstream direction. In accordance with the claims, the bowed tip portion <NUM> follows a curvilinear profile spanwise beginning at <NUM> degrees and increasing to <NUM> degrees at the tip <NUM>. The angles α and β are measured at either of the leading and trailing edges of the bowed tip portion <NUM>.

It should also be understood that although a particular component arrangement is disclosed in the illustrated embodiment, other arrangements will benefit herefrom. Although particular step sequences are shown, described, and claimed, it should be understood that steps may be performed in any order, separated or combined unless otherwise indicated and will still benefit from the present invention.

Although the different examples have specific components shown in the illustrations, embodiments of this invention are not limited to those particular combinations.

Claim 1:
A turbine blade (<NUM>) for a gas turbine engine (<NUM>), the turbine blade (<NUM>) comprising:
a platform (<NUM>) having a radially inner side and a radially outer side;
a root portion (<NUM>) extending from the radially inner side of the platform (<NUM>); and
an airfoil (<NUM>) extending from the radially outer side of the platform (<NUM>), the airfoil (<NUM>) including:
a pressure side (<NUM>) extending between a leading edge (<NUM>) and a trailing edge (<NUM>);
a suction side (<NUM>) extending between the leading edge (<NUM>) and the trailing edge (<NUM>),
and a bowed tip portion (<NUM>) extending perpendicular to a mid-camber line (<NUM>) of the airfoil (<NUM>), such that the bowed tip portion (<NUM>) is bowed to extend towards the suction side (<NUM>) in a circumferential direction and towards the trailing edge (<NUM>) in an axial downstream direction,
characterised in that:
the bowed tip portion (<NUM>) begins at between <NUM>% and <NUM>% of a span of the airfoil (<NUM>);
said bowed tip portion (<NUM>) is bowed at a leading edge (<NUM>) of the bowed tip portion (<NUM>) or a trailing edge of the bowed tip portion (<NUM>) between <NUM> degrees at a radially innermost part of the bowed tip portion (<NUM>) and <NUM> degrees at a tip (<NUM>) of the airfoil (<NUM>) relative to a radial direction (R) perpendicular to the mid-camber line (<NUM>) in the circumferential direction (Y);
the bowed tip portion (<NUM>) is bowed at a leading edge (<NUM>) of the bowed tip portion (<NUM>) or a trailing edge of the bowed tip portion (<NUM>) at an angle (β) between <NUM> degrees at a radially innermost part of the bowed tip portion (<NUM>) and <NUM> degrees at a tip (<NUM>) of the airfoil (<NUM>) relative to the radial direction (R) perpendicular to the mid-camber line (<NUM>) in the axial direction (X);
the bowed tip portion (<NUM>) is bowed at a leading edge (<NUM>) of the bowed tip portion (<NUM>) or a trailing edge of the bowed tip portion (<NUM>) by the same degree relative to the radial direction (R) in both the circumferential direction (Y) and the axial direction (X); and
the bowed tip portion (<NUM>) follows a curvilinear profile at a leading edge (<NUM>) of the bowed tip portion (<NUM>) or a trailing edge of the bowed tip portion (<NUM>) beginning at <NUM> degrees and increasing to <NUM> degrees relative to a radial direction (R) at a tip (<NUM>) of the airfoil (<NUM>) in the axial direction (X).