Patent Description:
Both the compressor and turbine sections may include alternating series of rotating blades and stationary vanes that extend into the core flow path of the gas turbine engine. For example, in the turbine section, turbine blades rotate and extract energy from the hot combustion gases that are communicated along the core flow path of the gas turbine engine. Turbine blades are known to include an airfoil section, over which the hot combustion gases flow, and a root attached to a rotatable disc. A support, or platform, is typically rigidly attached (e.g., bolted) adjacent to the turbine blade near the root, or is integrally formed with the turbine blade (e.g., by casting or molding).

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The present invention provides a gas turbine engine as defined in claim <NUM>.

In a further embodiment of any of the above, the turbine section includes a high pressure section and a low pressure section, the turbine blade being provided in the high pressure section.

In a further embodiment of any of the above, the turbine blade is made of a ceramic matrix composite (CMC) material.

In a further embodiment of any of the above, the second portion is at least partially moveable relative to the first portion of the component during operation of the gas turbine engine.

The present invention further provides a method of forming a component for a gas turbine engine as defined in claim <NUM>.

In a further embodiment of any of the above, the airfoil section and root are formed by molding ceramic matrix composite (CMC) material.

In a further embodiment of any of the above, the platform is formed by molding ceramic matrix composite (CMC) material.

These and other features of the present disclosure can be best understood from the following drawings and detailed description.

The drawings can be briefly described as follows:.

<FIG> schematically illustrates an example gas turbine engine <NUM> that includes a fan section <NUM>, a compressor section <NUM>, a combustor section <NUM> and a turbine section <NUM>. Alternative engines might include an augmenter section (not shown) among other systems or features. The fan section <NUM> drives air along a bypass flow path B while the compressor section <NUM> draws air in along a core flow path C where air is compressed and communicated to a combustor section <NUM>. In the combustor section <NUM>, air is mixed with fuel and ignited to generate a high pressure exhaust gas stream that expands through the turbine section <NUM> where energy is extracted and utilized to drive the fan section <NUM> and the compressor section <NUM>.

Although the disclosed non-limiting embodiment depicts a turbofan gas turbine engine, 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; for example a turbine engine including a three-spool architecture in which three spools concentrically rotate about a common axis and where a low spool enables a low pressure turbine to drive a fan via a gearbox, an intermediate spool that enables an intermediate pressure turbine to drive a first compressor of the compressor section, and a high spool that enables a high pressure turbine to drive a high pressure compressor of the compressor section. The concepts disclosed herein can further be applied outside of gas turbine engines, such as in the context of wind turbines.

The example engine <NUM> generally includes a low speed spool <NUM> and a high speed spool <NUM> mounted for rotation about an engine central longitudinal axis X relative to an engine static structure <NUM> via several bearing systems <NUM>. It should be understood that various bearing systems <NUM> at various locations may alternatively or additionally be provided.

The low speed spool <NUM> generally includes an inner shaft <NUM> that connects a fan <NUM> and a low pressure (or first) compressor section <NUM> to a low pressure (or first) turbine section <NUM>. The inner shaft <NUM> drives the fan <NUM> through a speed change device, such as a geared architecture <NUM>, to drive the fan <NUM> at a lower speed than the low speed spool <NUM>. The high-speed spool <NUM> includes an outer shaft <NUM> that interconnects a high pressure (or second) compressor section <NUM> and a high pressure (or second) turbine section <NUM>. The inner shaft <NUM> and the outer shaft <NUM> are concentric and rotate via the bearing systems <NUM> about the engine central longitudinal axis X.

A combustor <NUM> is arranged between the high pressure compressor <NUM> and the high pressure turbine <NUM>. In one example, the high pressure turbine <NUM> includes at least two stages to provide a double stage high pressure turbine <NUM>. In another example, the high pressure turbine <NUM> includes only a single stage. As used herein, a "high pressure" compressor or turbine experiences a higher pressure than a corresponding "low pressure" compressor or turbine.

The example low pressure turbine <NUM> has a pressure ratio that is greater than about five (<NUM>). The pressure ratio of the example low pressure turbine <NUM> is measured prior to an inlet of the low pressure turbine <NUM> as related to the pressure measured at the outlet of the low pressure turbine <NUM> prior to an exhaust nozzle.

The mid-turbine frame <NUM> further supports bearing systems <NUM> in the turbine section <NUM> as well as setting airflow entering the low pressure turbine <NUM>.

The core airflow C is compressed by the low pressure compressor <NUM> then by the high pressure compressor <NUM> mixed with fuel and ignited in the combustor <NUM> to produce high speed exhaust gases that are then expanded through the high pressure turbine <NUM> and low pressure turbine <NUM>. The mid-turbine frame <NUM> includes vanes <NUM>, which are in the core airflow path and function as an inlet guide vane for the low pressure turbine <NUM>. Utilizing the vane <NUM> of the mid-turbine frame <NUM> as the inlet guide vane for low pressure turbine <NUM> decreases the length of the low pressure turbine <NUM> without increasing the axial length of the mid-turbine frame <NUM>. Reducing or eliminating the number of vanes in the low pressure turbine <NUM> shortens the axial length of the turbine section <NUM>. Thus, the compactness of the gas turbine engine <NUM> is increased and a higher power density may be achieved.

The disclosed gas turbine engine <NUM> in one example is a high-bypass geared aircraft engine. In a further example, the gas turbine engine <NUM> includes a bypass ratio greater than about six (<NUM>), with an example embodiment being greater than about ten (<NUM>). The example geared architecture <NUM> is an epicyclical gear train, such as a planetary gear system, star gear system or other known gear system, with a gear reduction ratio of greater than about <NUM>.

In one disclosed embodiment, the gas turbine engine <NUM> includes a bypass ratio greater than about ten (<NUM>:<NUM>) and the fan diameter is significantly larger than an outer diameter of the low pressure compressor <NUM>. It should be understood, however, that the above parameters are only exemplary of one embodiment of a gas turbine engine including a geared architecture and that the present disclosure is applicable to other gas turbine engines.

The fan section <NUM> of the engine <NUM> is designed for a particular flight condition-typically cruise at about <NUM> Mach and about <NUM> (<NUM>,<NUM> feet). The flight condition of <NUM> Mach and <NUM> (<NUM>,<NUM> ft. ), with the engine at its best fuel consumption-also known as "bucket cruise Thrust Specific Fuel Consumption ('TSFC')"-is the industry standard parameter of pound-mass (lbm) of fuel per hour being burned divided by pound-force (lbf) of thrust the engine produces at that minimum point.

In another non-limiting embodiment the low fan pressure ratio is less than about <NUM>.

"Low corrected fan tip speed" is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram °C) / (<NUM>)]<NUM> ([(Tram °R) / (<NUM> °R)]<NUM>). The "Low corrected fan tip speed," as disclosed herein according to one non-limiting embodiment, is less than about <NUM>/s (<NUM> ft/second).

Turning now to <FIG>, the disclosed assembly can be used in various portions of a gas turbine engine. For exemplary purposes, a turbine blade <NUM> is described. It should be understood that this disclosure could apply to compressor blades, as well as stator vanes and fan blades.

Referring to <FIG>, an example turbine blade <NUM> includes an airfoil section <NUM>, a root <NUM>, and a support, here a platform, <NUM>. In this example the root <NUM> is a dove tail root. Other example roots, such as fir tree roots, could be used, however. The airfoil section <NUM> includes a pressure side <NUM> and a suction side <NUM> provided between a leading edge <NUM> and a trailing edge <NUM>. In this example, the platform <NUM> is provided over at least a portion of the root <NUM>. As will be explained in detail below, the platform <NUM> is a one-piece structure and is formed separately from the airfoil section <NUM> and the root <NUM>.

<FIG> illustrates a sectional view taken along line <NUM>-<NUM>, showing the turbine blade <NUM> arranged relative to a disc <NUM> of a turbine section. As illustrated, an intermediate layer <NUM> is provided between the root <NUM> and the platform <NUM>. During operation of the gas turbine engine <NUM>, it is possible that a portion of the intermediate layer <NUM> will be worn away. <FIG> illustrates an initial condition before any wearing of the intermediate layer <NUM> has occurred. In the initial condition the intermediate layer <NUM> prevents the platform <NUM> from directly contacting the root <NUM>.

In one example, the intermediate layer has a thickness of between <NUM> to <NUM> inches (<NUM> to <NUM>). Accordingly, the intermediate layer <NUM> is relatively thin such that the dimensions of the platform <NUM> substantially correspond to those of the root <NUM>.

<FIG> illustrates the assembly of <FIG> during operation of the gas turbine engine <NUM>. During operation, the airfoil section <NUM> and the root <NUM> undergo a radial pull force, illustrated at R. This radial pull force R causes the airfoil section <NUM> to contract in direction D1. In at least some circumstances, however, the platform <NUM> does not experience the same radial pull force R, and thus does not contract in the same way as the airfoil section <NUM>. Because the platform <NUM> is formed separately from the airfoil section <NUM> and the root <NUM>, the airfoil section <NUM> and the root <NUM> are allowed to move relative to the platform <NUM>. Thus, a gap, illustrated at <NUM>, can be formed between the platform <NUM> and the root <NUM>. Even though the gap <NUM> exists, the root <NUM> is still sufficiently supported by the slot <NUM> of the disc <NUM>, however. In the illustrated example, an inner face 86I of the slot <NUM> directly abuts the outer face 78F of the platform <NUM>. The platform <NUM>, in turn, abuts the root <NUM> either directly or by way of the intermediate layer <NUM> (if present). The turbine blade <NUM> is thus loaded into the disc as illustrated at <NUM>.

Because the platform <NUM> does not experience the same forces as the remainder of the turbine blade <NUM>, this relative movement reduces the stress on the platform <NUM>, and thus in turn reduces deflections of the platform <NUM>. Further, cracks that may form in the platform <NUM> tend not to propagate to the root <NUM>, which tends to prevent damage to the airfoil section <NUM> and the root <NUM>.

The intermediate layer <NUM> may be made of carbon, boron nitride, silicon, or another suitable applied material. The airfoil section <NUM>, the root <NUM>, and the platform section <NUM> can be made from ceramic matrix composite (CMC) materials. As is known in the art, a CMC material is one that includes fibers (such as carbon, silicon carbide or glass fibers, as example) supported within a ceramic matrix.

As mentioned above, the airfoil section <NUM> and the root <NUM> are formed separately from the platform section <NUM> to allow for relative movement therebetween. In one example method, the airfoil section <NUM> and root <NUM> are first formed by molding a CMC material (e.g., a plurality of CMC fabric sheets) into a desired shape. Next, the intermediate layer <NUM> is provided over the root <NUM>. In one example, the intermediate layer <NUM> is sprayed onto the root <NUM>. The platform <NUM>, also formed of a CMC material (e.g., a plurality of CMC fabric sheets) is then provided over the intermediate layer <NUM>. The entire assembly, including the formed airfoil <NUM> section and root <NUM>, the intermediate layer <NUM>, and the platform <NUM> is then molded. The final molded product can be further machined as necessary. Because the turbine blade <NUM> is formed of a CMC material, it is suitable for use in the high pressure turbine section <NUM>, however this disclosure is not limited to a particular section of the gas turbine engine <NUM>.

The intermediate layer <NUM> is selected of a material (e.g., a glass) such that the intermediate layer <NUM> is de-bonded from the root <NUM> during molding process.

Claim 1:
A gas turbine engine (<NUM>), comprising:
a compressor section (<NUM>), a combustor section (<NUM>), and a turbine section (<NUM>), the turbine section (<NUM>) including a stationary stage and a rotating stage; and
a turbine blade (<NUM>) provided in the rotating stage, the turbine blade (<NUM>) including a first portion (<NUM>, <NUM>), including an airfoil section (<NUM>) and a root (<NUM>), a second portion (<NUM>) formed separately from the first portion (<NUM>, <NUM>), wherein the second portion is a platform (<NUM>), and an intermediate layer (<NUM>) provided between the root (<NUM>) and the second portion (<NUM>),
characterized in that:
the intermediate layer (<NUM>) is selected of a material such that the intermediate layer (<NUM>) is de-bonded from the root (<NUM>) during a molding process of the airfoil section (<NUM>), the root (<NUM>), the intermediate layer (<NUM>) and the platform (<NUM>); and
an outer face (78F) of the platform (<NUM>) directly abuts an inner face (86I) of a slot (<NUM>) of a rotor disc (<NUM>).