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 combustion section where it is mixed with fuel and ignited to generate a high-energy exhaust gas flow. The high-energy 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 compressor and turbine sections include airfoils supported on rotors. The airfoils may be separate parts assembled to a rotor or may also be integrally formed as part of the rotor. Forming the rotor and airfoils as a single part reduces the number of parts and eliminates the need for fastening systems for securing airfoils to the rotor.

Turbine engine manufacturers continue to seek improvements to turbine engines including improvements in assembly, manufacture, engine performance and propulsive efficiencies.

A prior art integrally bladed rotor for a gas turbine engine having the features of the preamble of claims <NUM> and <NUM> is disclosed in <CIT>. Further prior art gas turbine engines are disclosed in <CIT>, <CIT>, <CIT>, <CIT>, <CIT>, <CIT>, <CIT> and <CIT>.

In an aspect of the present invention, an integrally bladed rotor for a gas turbine engine is provided as set forth in claim <NUM>.

In an embodiment, the patch portion includes a first fillet providing a smooth transition from a peripheral surface.

In another embodiment according to any of the previous embodiments, the patch portion is disposed on the suction side at the leading edge.

In another aspect of the present invention, a method of fabricating an integrally bladed rotor for a gas turbine engine is provided as set forth in claim <NUM>.

In another embodiment according to any of the previous embodiments, forming the patch portion includes forming the patch portion on the suction side at the leading edge and spaced apart from the trailing edge.

In another embodiment according to any of the previous embodiments, forming the first thickness includes forming a first fillet between the peripheral surface and at least one of the pressure side and suction side of the airfoil to have a first radius and forming a second fillet radially outward from the first fillet to have a second radius smaller than the first radius.

In another embodiment according to any of the previous embodiments, forming the first fillet to extend radially from the peripheral surface a first distance to an interface and forming the second fillet to begin at the interface.

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>.

"Low corrected fan tip speed" is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram °R) / (<NUM> °R)]<NUM> where °R = K × <NUM>/<NUM>.

The example gas turbine engine includes the fan <NUM> that comprises in one non-limiting embodiment less than about twenty-six fan blades. In another non-limiting embodiment, the fan section <NUM> includes less than about twenty fan blades. Moreover, in one disclosed embodiment the low pressure turbine <NUM> includes no more than about six turbine rotors schematically indicated at <NUM>. In another non-limiting example embodiment the low pressure turbine <NUM> includes about three turbine rotors. A ratio between the number of fan blades <NUM> and the number of low pressure turbine rotors is between about <NUM> and about <NUM>. The example low pressure turbine <NUM> provides the driving power to rotate the fan section <NUM> and therefore the relationship between the number of turbine rotors <NUM> in the low pressure turbine <NUM> and the number of blades <NUM> in the fan section <NUM> disclose an example gas turbine engine <NUM> with increased power transfer efficiency.

Referring to <FIG> with continued reference to <FIG>, the example gas turbine engine <NUM> includes the compressor section <NUM> that includes a plurality of integrally bladed rotors <NUM> (IBR). Each of the IBRs <NUM> includes a rotor portion <NUM> defining a peripheral surface <NUM>. A plurality of airfoils <NUM> extend upward from the peripheral surface <NUM>. The IBR <NUM> is a one-piece part with portions that define, among other features, the rotor portion <NUM>, peripheral surface <NUM> and the airfoils <NUM>.

Referring to <FIG>, which shows an example outside the wording of the claims, with continued reference to <FIG>, each of the plurality of airfoils <NUM> includes a leading edge <NUM>, a trailing edge <NUM>, a pressure side <NUM> and a suction side <NUM>. Each of the airfoils <NUM> extends radially outward from the peripheral surface <NUM> defined in the IBR <NUM>. The airfoils <NUM> extend from a root portion <NUM> to a tip portion <NUM>. The root portion <NUM> is defined at the peripheral surface <NUM> of the rotor portion <NUM>. A thickness <NUM> is disposed near the root <NUM> and extends between the peripheral surface <NUM> and side surfaces of the airfoil <NUM>.

Referring to <FIG>, which shows an example outside the wording of the claims, with continued reference to <FIG>, the example thickness <NUM> is disposed at the root portion <NUM> about the airfoil <NUM> and provides a boundary to prevent crack propagation from the airfoil <NUM> into the rotor portion <NUM> of the IBR <NUM>. In this example, the thickness <NUM> extends outward from an airfoil thickness <NUM> defined between the pressure side <NUM> and the suction side <NUM>.

Referring to <FIG>, which shows an example outside the wording of the claims, with continued reference to <FIG>, the first thickness <NUM> includes a first fillet <NUM> that extends from the rotor peripheral surface <NUM> to an interface <NUM> and a second fillet <NUM> that extends radially outward from the interface <NUM>. The first fillet <NUM> includes a transition surface <NUM> that extends a distance <NUM> from the peripheral surface <NUM> to the interface <NUM>. The interface <NUM> is disposed a distance <NUM> above the peripheral surface <NUM> and a width <NUM> away from the pressure side <NUM> and suction side <NUM> of the airfoil <NUM>. The width <NUM> is less than the distance <NUM>. The disclosed interface <NUM> is the interface between the first fillet <NUM> and the second fillet and is the location where the surface <NUM> transitions from a first radius <NUM> to a second radius <NUM>. Accordingly, the second fillet <NUM> begins at the interface <NUM> and extends upward radially at the second radius <NUM> into a smooth transition that merges with the pressure and suction surfaces of the airfoil <NUM>.

In the disclosed example, the first radius <NUM> is larger than the second radius <NUM>. In one disclosed example, the first radius <NUM> is between one third and one half greater than the second radius <NUM>. In another disclosed example, the first radius <NUM> is about <NUM> inches (<NUM>) and the second radius is <NUM> inches (<NUM>). In another disclosed example the first radius is about <NUM> inches (<NUM>) and the second radius is about <NUM> inches (<NUM>). In another disclosed dimensional example, the first radius is about <NUM> inches (<NUM>) and the second radius is about <NUM> inches (<NUM>). Moreover, in one example, the width <NUM> is between about <NUM> inches (<NUM>) and <NUM> inches (<NUM>). In other disclosed example the width <NUM> is about <NUM> inches (<NUM>). It should be understood, that the disclosed dimensional example is provided by way of example and other radiuses and widths could be utilized and are within the contemplation of this disclosure.

It should be understood that although dimensional examples are disclosed by way of example, the first fillet <NUM> is larger than the second fillet <NUM>. The specific relative size between the first fillet <NUM> and the second fillet <NUM> may be different to provide a predefined stress propagation path that prevents crack propagation radially inward into the rotor portion <NUM>.

Referring to <FIG>, which shows an example outside the wording of the claims, with continued reference to <FIG>, the separation at the interface <NUM> between the first fillet <NUM> and the second fillet <NUM> defines a crack propagation boundary schematically shown at <NUM>. A potential crack schematically referred to as <NUM> is prevented from propagating radially inward toward the rotor portion <NUM> by the thicker portions of the airfoil defined by the first fillet <NUM>. Instead, the crack <NUM> propagates in direction substantially along and parallel to the interface <NUM>. Accordingly, the interface <NUM> defines the crack propagation boundary and prevents cracks from propagating into the rotor portion <NUM>.

Referring to <FIG>, an IBR <NUM> embodiment in accordance with the present invention is schematically illustrated and includes an airfoil <NUM> with a suction side <NUM> and a pressure side <NUM> that extends between a leading edge <NUM> and a trailing edge <NUM>. The airfoil <NUM> extends upward from a rotor peripheral surface <NUM> and includes a patch portion <NUM>. In this example, the patch portion <NUM> is disposed on the pressure side <NUM> and extends a distance <NUM> from the leading edge <NUM> toward the trailing edge <NUM>. The patch <NUM> is therefore spaced apart a distance <NUM> away from the trailing edge <NUM>. The patch <NUM> is disposed at a location between the airfoil <NUM> and peripheral surface <NUM> that defines a boundary that prevents crack propagation into the rotor portion <NUM>. Moreover, the area of the increased thickness <NUM> provided by the patch <NUM> is based on stress analysis of potential crack propagation and may vary in location and thickness.

The patch <NUM> is an increased thickness indicated at <NUM> (<FIG>) that is greater than the airfoil thickness <NUM>. The airfoil thickness <NUM> may vary depending on the airfoil shape between the leading edge and the trailing edge. The patch portion <NUM> includes a thickness <NUM> in addition to the airfoil thickness <NUM> of the airfoil <NUM> in a specific location near the leading edge <NUM>. The location of the thickness <NUM> is provided based on analysis of stresses inflicted on the airfoil <NUM> during operation. The increased thickness <NUM> is shown on one side of the airfoil <NUM>, but may extend to both sides of the airfoil <NUM>. Moreover, the patch <NUM> may extend different distances <NUM> toward the trailing edge <NUM> as determined to define a boundary to possible crack propagation radially inward. Moreover, the increased thickness <NUM> provided by the patch <NUM> could be spaced from the leading edge <NUM> or any position along the interface between the peripheral surface <NUM> and the airfoil <NUM> where a reduction in operating stress is required to direct crack propagation away from the rotor portion <NUM>.

Accordingly, the example disclosed IBR <NUM> includes airfoils <NUM>, <NUM> with features that define crack propagation boundaries to prevent cracks from propagating radially inward to into rotor portions.

Claim 1:
An integrally bladed rotor (<NUM>; <NUM>) for a gas turbine engine (<NUM>) comprising:
a rotor portion (<NUM>) with a peripheral surface (<NUM>);
at least one airfoil (<NUM>) including a suction side (<NUM>) and a pressure side (<NUM>) extending between a leading edge (<NUM>) and a trailing edge (<NUM>), the at least one airfoil (<NUM>) extending radially from the peripheral surface (<NUM>), the airfoil (<NUM>) including an airfoil thickness (<NUM>) between the pressure side (<NUM>) and the suction side (<NUM>); and
a patch portion (<NUM>) disposed at a location between the airfoil (<NUM>) and the peripheral surface (<NUM>) that defines a boundary that prevents crack propagation into the rotor portion (<NUM>), on one of the suction side (<NUM>) and the pressure side (<NUM>), the patch portion (<NUM>) including a first thickness (<NUM>) in addition to the airfoil thickness (<NUM>),
characterized in that:
the patch portion (<NUM>) extends partway between the leading edge (<NUM>) and the trailing edge (<NUM>) and is spaced apart from the trailing edge (<NUM>) and/or the leading edge (<NUM>).