Patent Description:
Gas turbine engines, such as those that power modern commercial and military aircraft, generally include a compressor to pressurize an airflow, a combustor to burn a hydrocarbon fuel in the presence of the pressurized air, and a turbine to extract energy from the resultant combustion gases. The compressor and turbine sections include rotatable blade and stationary vane arrays. The blades and vanes typically include low and high-pressure airfoils, vanes, vane rings, shrouds, and nozzle segments.

The stationary vane arrays are typically assembled between outer and inner shrouds, or rings, in a variety of manners. Although the actual elements may vary in their configuration and construction, one similarity is that the vanes are typically constructed to allow for thermal expansion. The thermal expansion is typically accommodated through assembly of the vanes relatively loosely in the inner and outer shrouds. Although effective, such assembly may result in various stresses.

<CIT> discloses a prior art airfoil fairing shell according to the preamble of claim <NUM>.

<CIT> discloses a prior art mounting for guide vanes.

<CIT> discloses a prior art diaphragm and blades for turbomachinery.

According to an aspect of the present invention, there is provided a vane ring as set forth in claim <NUM>.

In an embodiment of the above, the structural support includes an interface for attachment to an engine case structure.

In a further embodiment of any of the foregoing embodiments of the present disclosure, the structural support is an arcuate segment.

In a further embodiment of any of the foregoing embodiments of the present disclosure, the structural support is a full ring.

In a further embodiment of any of the foregoing embodiments of the present disclosure, the lug extends transverse to the mateface of the vane endwall.

In a further embodiment of any of the foregoing embodiments of the present disclosure, the axial attachment face, the pressure side tangential attachment face, and the suction side tangential attachment face are generally perpendicular to the radial attachment face.

In a further embodiment of any of the foregoing embodiments of the present disclosure, the axial attachment face, is generally perpendicular to the pressure side tangential attachment face and the suction side tangential attachment face.

In a further embodiment of any of the foregoing embodiments of the present disclosure, the suction side tangential attachment face and the axial attachment face are downstream of an aerodynamic center of a resultant aerodynamic load generated by the airfoil fairing shell such that the in plane loading to the reaction forces on these faces is compressive.

In a further embodiment of any of the foregoing embodiments of the present disclosure, the suction side tangential attachment face is parallel to the tangential attachment face.

In a further embodiment of any of the foregoing embodiments of the present disclosure, the suction side tangential attachment face and the pressure side tangential attachment face are non-parallel to the inner vane endwall.

In a further embodiment of any of the foregoing embodiments of the present disclosure, the suction side tangential attachment face and the pressure side tangential attachment face are non-parallel.

In a further embodiment of any of the foregoing embodiments of the present disclosure, the suction side tangential attachment face is generally perpendicular to a resultant aerodynamic load generated by the airfoil.

In a further embodiment of any of the foregoing embodiments of the present disclosure, the suction side tangential attachment face is aligned with respect to a resultant aerodynamic load generated by the airfoil.

The gas turbine engine <NUM> is disclosed herein as a two-spool turbo fan that generally incorporates a fan section <NUM>, a compressor section <NUM>, a combustor section <NUM>, and a turbine section <NUM>. Alternative engine architectures <NUM> might include an augmentor section <NUM>, an exhaust duct section <NUM>, and a nozzle section <NUM> (<FIG>) among other systems or features. The fan section <NUM> drives air along a bypass flowpath and into the compressor section <NUM> to drive core air along a core flowpath. The core air is compressed then communicated into the combustor section <NUM> for downstream expansion through the turbine section <NUM>. Although depicted as a turbofan in the disclosed non-limiting embodiment, it should be appreciated that the concepts described herein are not limited, only to turbofans as the teachings may be applied to other types of turbine engine architectures such as turbojets, turboshafts, and three-spool (plus fan) turbofans.

The engine <NUM> generally includes a low spool <NUM> and a high spool <NUM> mounted for rotation about an engine central longitudinal axis A relative to an engine case structure <NUM> via several bearing compartments <NUM>. The low spool <NUM> generally includes an inner shaft <NUM> that interconnects a fan <NUM>, a low pressure compressor ("LPC") <NUM> and a low pressure turbine ("LPT") <NUM>. The inner shaft <NUM> drives the fan <NUM> directly or through a geared architecture <NUM> to drive the fan <NUM> at a lower speed than the low spool <NUM>. An exemplary reduction transmission is an epicyclic transmission, namely a planetary or star gear system.

The high spool <NUM> includes an outer shaft <NUM> that interconnects a high pressure compressor ("HPC") <NUM> a and high pressure turbine ("HPT") <NUM>. A combustor <NUM> is arranged between the HPC <NUM> and the HPT <NUM>. Core airflow is compressed by the LPC <NUM>, then the HPC <NUM>, mixed with the fuel and burned in the combustor <NUM>, then expanded over the HPT <NUM> and the LPT <NUM> which rotationally drive the respective low spool <NUM> and high spool <NUM> in response to the expansion.

With reference to <FIG>, an enlarged schematic view of a portion of the HPT <NUM> is shown by way of example; however, other engine sections will also benefit herefrom. A shroud assembly <NUM> mounted to the engine case structure <NUM> supports a Blade Outer Air Seal (BOAS) assembly <NUM> with a multiple of circumferentially distributed BOAS <NUM> proximate to a rotor assembly <NUM> (one schematically shown).

The shroud assembly <NUM> and the BOAS assembly <NUM> are axially disposed between a forward stationary vane ring <NUM> and an aft stationary vane ring <NUM>. The rotor assembly <NUM> includes an array of blades <NUM> circumferentially disposed around a disk <NUM>. Each blade <NUM> includes a root <NUM>, a platform <NUM>, and an airfoil <NUM>. The blade roots <NUM> are received within a rim <NUM> of the disk <NUM> and the airfoils <NUM> extend radially outward such that a tip <NUM> of each airfoil <NUM> adjacent to the blade outer air seal (BOAS) assembly <NUM>. The platform <NUM> separates a gas path side inclusive of the airfoil <NUM> and a non-gas path side inclusive of the root <NUM>.

With reference to <FIG>, the forward stationary vane ring <NUM> will be described as a clamped stator assembly, however, it should be appreciated that the aft stationary vane ring <NUM> as well as other vane rings in the turbine section and the compressor section will also benefit herefrom. In the example forward stationary vane ring <NUM>, each airfoil <NUM> extends between a respective inner vane endwall <NUM> and an outer vane endwall <NUM> to form an airfoil fairing shell <NUM>. Each airfoil fairing shell <NUM> is respectively clamped between an outer structural support <NUM>, and an inner structural support <NUM> (also shown in <FIG>).

The outer structural support <NUM>, and the inner structural support <NUM> may be full rings or circumferentially segmented structures that are mounted within or a portion of the engine case structure <NUM>, or attached thereto via fasteners, clamping, pins, or other such interface <NUM>, <NUM> (illustrated schematically). That is, the airfoil fairing shells <NUM> are clamped into a full ring or circumferentially segmented outer and inner structural support <NUM>, <NUM> that are, in turn, formed, fastened or otherwise located in the engine case structure <NUM> (<FIG>). In the example circumferentially segmented structure, each segment of the outer structural support <NUM> and/or the inner structural support <NUM> may support a cluster of one or more airfoil fairing shells <NUM> (three shown in <FIG>).

With reference to <FIG>, each airfoil <NUM> defines a blade chord between a leading edge <NUM>, which may include various forward and/or aft sweep configurations, and a trailing edge <NUM>. A first airfoil sidewall <NUM> that may be convex to define a suction side, and a second airfoil sidewall <NUM> that may be concave to define a pressure side, are joined at the leading edge <NUM> and at the axially spaced trailing edge <NUM>. An aerodynamic center "C" of the airfoil is located at about a quarter chord position, however, such aerodynamic centers may vary dependent upon the airfoil.

The inner vane endwall <NUM> and the outer vane endwall <NUM> are generally a parallelogram, chevron, arc, or other shape when viewed from the top and generally includes a respective forward edge <NUM>, <NUM>, an aft edge <NUM>, <NUM> and a mateface 138A, 138B, 140A, 140B therebetween. The endwalls <NUM>, <NUM> may be cylindrical, conical, arbitrary axisymmetric, or non-axisymmetric when viewed in cross-section. The non-gaspath face of the platform may be any of these as well. That is, the airfoil <NUM>, the inner vane endwall <NUM>, and the outer vane endwall <NUM> form the airfoil fairing shell <NUM> that is radially clamped by the outer structural support <NUM>, and the inner structural support <NUM>.

The airfoil fairing shell <NUM> may include passages "P" (three shown) for cooling airflow and or electrical conduits, may be solid, may be hollow, or combinations thereof. Such an arrangement facilitates manufacture of metallic or non-metallic airfoil fairing shells, particularly but not exclusively those of low-ductility and/or low coefficient of thermal expansion materials, that are readily assembled to the outer structural support and the inner structural support which, in turn, are manufactured of metallic or non-metallic material and which may be manufactured from the same material as or dissimilar material to the airfoil fairing shells.

In this disclosed non-limiting embodiment, the outer vane endwall <NUM> will be described, however, it should be appreciated that the inner vane endwall <NUM> as well as other vane rings will also benefit herefrom. The outer vane endwall <NUM> generally includes a radial attachment face <NUM>, a suction side tangential attachment face <NUM>, a pressure side tangential attachment face <NUM>, and an axial attachment face <NUM>. The attachment faces <NUM>, <NUM>, <NUM>, <NUM> transmit axial and tangential aerodynamic loads from the airfoil fairing shell <NUM> into the structural supports and transmit clamping load through the fairing shell.

The airfoil fairing shell <NUM> includes cylindrical, conical, arbitrary axisymmetric or planar radial attachment faces through which the spanwise clamping load is generally transmitted, and two pairs of orthogonal planar attachment faces through which aerodynamic loads and retention loads are generally transmitted, and which are quasi-orthogonal to the radial direction at the airfoil's circumferential station. It should be appreciated that one attachment face may be the primary attachment face while another attachment face is a secondary attachment face with respect configurations where the faces are rotated with respect to the engine axis. Axial and tangential oriented primary and secondary attachment faces are one disclosed non-limiting embodiment. More generally, primary and secondary attachment faces, where axial and tangentially aligned are one specific type, where primary aligned with the resultant load is another specific type, and where primary aligned with the platform mateface edge is a third type. The primary and secondary are orthogonal to one another, and are quasi-orthogonal to the radial direction at the airfoil's circumferential station.

The attachment faces <NUM>, <NUM>, <NUM>, <NUM> are generally formed by thickened areas of the outer vane endwall <NUM> or other features such as tabs that are arranged to form these faces and interface with respective attachment faces formed by the associated outer structural support <NUM> (<FIG>). In this disclosed non-limiting embodiment, the outer structural support <NUM> includes a series of lugs <NUM> provide tangential reaction faces that may be angled with respect to the engine axis A and extend transverse to a respective circumferential interface <NUM> between each airfoil fairing shell <NUM> (<FIG>, <FIG>, <FIG>). Alternatively, the series of lugs <NUM> provide tangential reaction faces is parallel to the engine axis (<FIG>).

The attachment faces <NUM>, <NUM>, <NUM>, <NUM> are arranged to transmit loads between the respective structural supports <NUM>, <NUM> and the airfoil fairing shell <NUM>. The surface of the thickened area of the airfoil fairing shell <NUM> forms the radial attachment face <NUM> that is clamped between the outer and inner structural support <NUM>, <NUM>.

The attachment faces <NUM>, <NUM>, <NUM> react in-plane loads formed by the step transitions of the thickened areas of the outer vane endwall <NUM>. The attachment faces <NUM>, <NUM>, <NUM> may be aligned with the axial and tangential directions or rotated to an arbitrary angle, such as that which presents a large face perpendicular to the resultant aerodynamic load (<FIG>). The suction side tangential attachment face <NUM> is tasked with reacting aerodynamic in-plane loads and may be located downstream, and to the suction side of the aerodynamic center "C" of the airfoil such that aerodynamic loads tend to create compressive rather than tensile loads in the material that are in the plane with the platform between the aerodynamic center "C" to the attachment faces. That is, the suction side tangential attachment face <NUM> interfaces with one of the series of lugs <NUM> of the structural support <NUM> to provide a primary reaction load. The axial attachment face <NUM> also interfaces with the structural support <NUM> to provide a secondary reaction load.

In this disclosed non-limiting embodiment, the primary reaction load and the secondary reaction load are non-parallel to the resultant aerodynamic load from the aerodynamic center "C" as generated by the airfoil <NUM>, and the suction side tangential attachment face <NUM> is non-parallel to the matefaces 138A, 138B. Here, the suction side tangential attachment face <NUM> and the pressure side tangential attachment face <NUM> are perpendicular to the forward edge <NUM> and the aft edge <NUM>. In this disclosed non-limiting embodiment, a corner <NUM> at the interface between the suction side tangential attachment face <NUM> and the axial attachment face <NUM> is chamfered. The chamfered corner <NUM>. The purpose of such a chamfer allows the load bearing faces on the stator shell to clear the edges of the complementary faces' fillet at their junction on the structural platform to prevent contact stress concentration (<FIG>).

In another disclosed non-limiting embodiment, the primary reaction load is parallel to a resultant aerodynamic load from an aerodynamic center as generated by the airfoil and the suction side tangential attachment face <NUM> is non-parallel to the matefaces 138A, 138B (<FIG>). Notably, the primary and secondary load transmitting faces (alternatively called the axial and tangential faces when these faces are not angled with respect to the engine axis) are always perpendicular to each other, regardless of whether or not one of them is aligned to the resultant loading direction.

In another disclosed non-limiting embodiment, the suction side tangential attachment face <NUM> and the pressure side tangential attachment face <NUM> are generally parallel to the matefaces 138A, 138B of the outer vane endwall <NUM> (<FIG>). Such arrangements may facilitate assembly into the engine <NUM>.

The attachment faces <NUM>, <NUM>, <NUM>, <NUM> enable the use of separate structural platforms for a turbine stator which itself has an airfoil and endwalls while limiting tensile stresses in the plane of the platform.

The use of the terms "a," "an," "the," and similar references in the context of description (especially in the context of the following claims) are to be construed to cover both the singular and the plural, unless otherwise indicated herein or specifically contradicted by context. The modifier "about" used in connection with a quantity is inclusive of the stated value and has the meaning dictated by the context (e.g., it includes the degree of error associated with measurement of the particular quantity). All ranges disclosed herein are inclusive of the endpoints, and the endpoints are independently combinable with each other.

Although the different non-limiting embodiments have specific illustrated components, the embodiments of this invention are not limited to those particular combinations. It is possible to use some of the components or features from any of the non-limiting embodiments in combination with features or components from any of the other non-limiting embodiments.

It should be appreciated that like reference numerals identify corresponding or similar elements throughout the several drawings. It should also be appreciated that although a particular component arrangement is disclosed in the illustrated embodiment, other arrangements will benefit herefrom.

Claim 1:
A vane ring (<NUM>) for a gas turbine engine, comprising:
a multiple of airfoil fairing shells (<NUM>) each with an airfoil section (<NUM>) between an outer vane endwall (<NUM>) and an inner vane endwall (<NUM>),
at least one of said outer vane endwall (<NUM>) and said inner vane endwall (<NUM>) of each of the multiple of airfoil fairing shells (<NUM>) forms a mateface (<NUM>), each of said multiple of airfoil fairing shells (<NUM>) adjacent to another one of said multiple of airfoil fairing shells (<NUM>) at said mateface; and
a structural support (<NUM>, <NUM>),
characterised in that:
at least one of said outer vane endwall (<NUM>) and said inner vane endwall (<NUM>) of each of the multiple of airfoil fairing shells (<NUM>) includes a thickened region comprising step transitions, the thickened region forming a radial attachment face (<NUM>), a suction side tangential attachment face (<NUM>), a pressure side tangential attachment face (<NUM>), and an axial attachment face (<NUM>) configured to react in-plane loads formed by the step transitions;
said suction side tangential attachment face (<NUM>) downstream of an aerodynamic center (C) of a resultant aerodynamic load generated by said airfoil fairing shell (<NUM>) such that the in-plane loading to the reaction forces on said suction side tangential attachment faces (<NUM>) is compressive; and
the structural support (<NUM>, <NUM>) includes a multiple of lugs (<NUM>), the suction side tangential attachment face (<NUM>) of each of said multiple of airfoil fairing shells (<NUM>) interfacing with one of the multiple of lugs (<NUM>) to provide a primary reaction load, and the axial attachment face (<NUM>) of each of said multiple of airfoil fairing shells (<NUM>) interfacing with the structural support (<NUM>, <NUM>) to provide a secondary reaction load.