Dual anti surge and anti rotation feature on first vane support

A gas turbine engine includes a vane and a combustor housing that are supported relative to an engine static structure. A retaining assembly clamps the combustor housing and the vane to one another in an axial direction. A circumferential load transfer assembly circumferentially affixes the vane relative to the engine static structure. The retaining assembly is secured to the circumferential load transfer assembly.

BACKGROUND

This disclosure relates to first stage turbine vanes and associated mounting arrangement.

A gas turbine engine typically includes a fan section, a compressor section, a combustor section and a turbine section. Core flow 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-speed exhaust gas flow. The combustor section includes a combustor housing with a flange used to mount the combustor housing with respect to the engine's static structure. The high-speed exhaust gas flow expands through the turbine section to drive the compressor and the fan section.

First stage turbine vanes are arranged immediately downstream from the combustor section to efficiently communicate the core flow into the first stage of turbine blades. Prior technology for the first stage turbine vanes employs two separate features to complete two separate tasks, affixing the vanes circumferentially and supporting the combustor in the event of a compressor surge condition.

Typically an array of separate vanes or clusters of vanes are mounted with respect to the engines static structure. The engine static structure includes a circumferential load transfer assembly having a circumferential array of tabs, which are used to interface with a fork on each of the first vanes to affix the vanes circumferentially. The engine static structure also includes a boss separate from the tabs to which a retainer is bolted to provide a retaining assembly. The retaining assembly secures the combustor flange to the engine static structure via the vanes and holds the flange in place in case of a compressor surge condition. These two features are separate from one another and located circumferentially between each other around the engine static structure.

SUMMARY

In one exemplary embodiment, a gas turbine engine includes a vane and a combustor housing that are supported relative to an engine static structure. A retaining assembly clamps the combustor housing and the vane to one another in an axial direction. A circumferential load transfer assembly circumferentially affixes the vane relative to the engine static structure. The retaining assembly is secured to the circumferential load transfer assembly.

In a further embodiment of the above, the engine static structure includes a vane support. The vane support includes one of a tab and a fork. The vane includes the other of the tab and the fork. The tab is received in the fork. The tab and the fork provide the circumferential load transfer assembly.

In a further embodiment of any of the above, the fork is provided on an outer platform of the vane. The tab is provided on the vane support.

In a further embodiment of any of the above, the retainer assembly includes a retainer secured to the tab. The combustor housing is arranged axially between the retainer and the vane.

In a further embodiment of any of the above, a seal ring is engaged between the combustor housing and the vane.

In a further embodiment of any of the above, the retainer includes a finger. The combustor housing includes an annular protrusion that extends radially outward from the combustor housing. The finger engages the annular protrusion.

In a further embodiment of any of the above, the combustor housing includes an edge that engages the seal ring.

In a further embodiment of any of the above, the retainer and tab include holes. A fastener extends through the holes to secure the retainer to the vane support.

In a further embodiment of any of the above, there is a circumferential array of a number of vanes. Each vane includes a fork and a number of tabs with holes being less than the number of vanes.

In a further embodiment of any of the above, the retaining assembly is provided by discrete retainers circumferentially spaced from one another.

In a further embodiment of any of the above, the retaining assembly is provided by an annular ring that is secured to multiple tabs.

In a further embodiment of any of the above, the annular ring includes lightening holes.

In another exemplary embodiment, a gas turbine engine includes an engine static structure that includes a vane support having a tab. A vane includes a fork that has a notch that receives the tab to circumferentially affix the vane to the engine static structure. A retainer is secured to the tab and mounts a combustor housing to the vane in an axial direction.

In a further embodiment of the above, a seal ring is engaged between the combustor housing and the vane.

In a further embodiment of any of the above, the retainer includes a finger. The combustor housing includes an annular protrusion that extends radially outward from the combustor housing. The finger engages the annular protrusion.

In a further embodiment of any of the above, the combustor housing includes an edge that engages the seal ring.

In a further embodiment of any of the above, the retainer and tab include holes. A fastener extends through the holes to secure the retainer to the vane support.

In a further embodiment of any of the above, there is a circumferential array of a number of vanes. Each vane includes a fork and a number of tabs with holes being less than the number of vanes.

In a further embodiment of any of the above, discrete retainers are circumferentially spaced from one another.

In a further embodiment of any of the above, the retainer is provided by an annular ring secured to multiple tabs.

DETAILED DESCRIPTION

FIG. 1schematically illustrates a gas turbine engine20. Although commercial engine embodiment is shown, the disclosed vane mounting arrangement may also be used in military engine applications. The gas turbine engine20is disclosed herein as a two-spool turbofan that generally incorporates a fan section22, a compressor section24, a combustor section26and a turbine section28. Alternative engines might include an augmentor section (not shown) among other systems or features.

The fan section22drives air along a bypass flowpath B while the compressor section24drives air along a core flowpath C (as shown inFIG. 2) for compression and communication into the combustor section26then expansion through the turbine section28. Although depicted as a two-spool turbofan gas turbine engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to use with two-spool turbofans as the teachings may be applied to other types of turbine engines including three-spool architectures.

An area of the combustor section26is shown in more detail inFIG. 2. The combustor section26includes a combustor56having a combustor housing60. An injector62is arranged at a forward end of the combustor housing60and is configured to provide fuel to the combustor housing60where it is ignited to produce hot gases that expand through the turbine section54.

A diffuser case64is secured to the combustor housing60and forms a diffuser plenum surrounding the combustor housing60. The diffuser plenum may receive a diffuser flow D for diffusing flow from the compressor section52into the combustor section56. The diffuser case64and the combustor housing60are fixed relative to the engine static structure36(FIG. 1), illustrated as elements36aand36binFIG. 2.

In one example, an array of vanes72of a first stage of turbine stator vanes includes an inner portion that is partially supported by the diffuser case64. One typical mounting method for first stage turbine vanes is to provide a radially inwardly extending flange84that includes a hole86(shown inFIG. 4A). A pin (not shown) is received in the hole to secure the flange84at a joint88(shown inFIG. 2).

With continuing reference toFIG. 2, the diffuser case64includes a portion arranged downstream from the compressor section52and upstream from the combustor section26that is sometimes referred to as a “pre-diffuser”66. A bleed source68, such as fluid from a compressor stage, provides cooling fluid through the pre-diffuser66to various locations interiorly of the diffuser case64. A heat exchanger (not shown) may be used to cool the cooling fluid before entering the pre-diffuser66.

The compressor section52includes a compressor rotor70supported for rotation relative to the engine static structure36bby the bearing38. The bearing38is arranged within a bearing compartment74that is buffered using a buffer flow R. The turbine section54includes a turbine rotor76arranged downstream from a tangential on-board injector module78, or “TOBI.” The TOBI78provides cooling flow T to the turbine rotor76.

Referring toFIGS. 3A-4B, the vanes72include an outer portion that is supported by the engine static structure36ausing a vane support92, which is provided by a unitary annular structure, however, it should be understood that the vane support92may instead be constructed from multiple segments. In one example, the vane support92is grounded to an outer case of the engine static structure using teeth93. The vanes72may be provided as multiple arcuate segments. In one example, each vane72is provided a doublet having a pair of airfoils joined between radially spaced apart inner and outer platforms80,82.

The outer platform82includes radially extending circumferentially spaced structures providing a fork90that defines a notch85. The vane support92includes a radially inwardly extending tab94that is received circumferentially within the fork90in the notch85to provide a circumferential load transfer assembly. In one example, at least one fork is provided on each vane. This fork and tab arrangement circumferentially locates the vanes72and transfers the circumferential load from the vanes72during engine operation to the engine static structure36avia the vane support92.

Referring toFIGS. 5 and 6, the tab94includes a hole96to which a retainer108is secured to provide a retaining assembly. In one example, up to twenty retaining assemblies may be provided circumferentially, which may be less than the number of vanes72. The retaining assembly clamps the combustor housing60to the vane72and holds the assembly together, in particular, during compressor surge conditions.

In one example, a ring seal98is arranged axially between an aft end of the combustor housing60and a forward face100of the outer platform82. An edge104of the combustor housing60urges a sealing face102of the ring seal98into engagement with the forward face100. A radially inwardly extending finger110of the retainer108engages an annular protrusion106that extends radially outwardly from the combustor housing60. A fastener114received in the hole96and a hole112in the retainer108is used to apply a clamping load to seal the combustor60relative to the vane72.

For vanes172(FIG. 3A) that do not have a retaining assembly, for example, tabs194, which without a hole95to accommodate the retainer108, the fork190and its notch185may be narrower since there is no need to accommodate a fastener through the tab.

Another example retainer208is illustrated inFIG. 7A-7B. Unlike the discrete retainer108illustrated inFIGS. 5 and 6, the retainer208may be a continuous annular ring or arcuate segments that provide multiple of fingers210. Lightning holes118may be provided on the ring to reduce the weight of the retainer208.

The retaining assembly is secured to the circumferential load transfer assembly. Integrating the retaining assembly with the circumferential load transfer assembly provides a significant weight savings. The disclosed arrangement uses a single bolted on feature at several circumferential locations, which prevents circumferential movement of the vanes and prevents the combustor from moving forward in a surge condition.