Patent Description:
Gas turbine engines are known, and typically include a fan delivering air into a bypass duct as propulsion air. The fan also delivers air to a compressor section. Air compressed by the compressor is delivered into a combustor where it is mixed with fuel and ignited. Products of this combustion pass downstream over turbine rotors, driving them to rotate.

As known, gas turbine engines have several sections that become quite hot during operation. As examples, the air compressed by the compressor will reach high temperatures. The turbine section will see very high temperatures downstream of the combustor. It is known to bleed compressor air and then utilize that bleed air to cool components in the compressor and/or turbine section.

In a known system, the air bled from the compressor section moves into a bleed chamber, and extends directly across the bleed chamber to an outer housing, where it impacts the outer housing. This provides heat stress on the outer housing.

<CIT> discloses a prior art gas turbine engine according to the preamble of claim <NUM>.

In a featured embodiment, a gas turbine engine is provided according to claim <NUM>.

In another embodiment according to the previous embodiment, the blocking portion extends at an angle that is within <NUM> degrees of being parallel to the axis of rotation.

In another embodiment according to any of the previous embodiments, the blocking portion extends at a right angle from the mount portion.

In another embodiment according to any of the previous embodiments, the flow discourager is bolted to the mount housing.

In another embodiment according to any of the previous embodiments, the blocking portion extends axially forwardly of the mount portion to a blocking portion axially forwardmost extent. The bleed holes have a bleed hole axially forwardmost extent. The bleed hole axially forwardmost extent is axially rearward of the blocking portion axially forwardmost extent.

In another embodiment according to any of the previous embodiments, the blocking portion has wrench slots which are open, and are associated with each of the bolt holes to facilitate tightening of a bolt in the bolt hole.

In another embodiment according to any of the previous embodiments, air from the bleed chamber is directed to cool components in a turbine section.

In another embodiment according to any of the previous embodiments, an intermediate heat exchanger cools the air from the bleed chamber before it reaches the turbine section.

In another embodiment according to any of the previous embodiments, there is a downstream most compressor blade that defines a highest pressure point in the compressor section and lesser pressure points upstream of the highest pressure point. The bleed holes are downstream of a compressor blade which is at a lesser pressure point of the compressor section.

The fan section <NUM> may include a single-stage fan <NUM> having a plurality of fan blades <NUM>. The fan blades <NUM> may have a fixed stagger angle or may have a variable pitch to direct incoming airflow from an engine inlet. The fan <NUM> drives air along a bypass flow path B in a bypass duct <NUM> defined within a housing <NUM> such as a fan case or nacelle, and also drives air along a core flow path C for compression and communication into the combustor section <NUM> then expansion through the turbine section <NUM>. A splitter <NUM> aft of the fan <NUM> divides the air between the bypass flow path B and the core flow path C. The housing <NUM> may surround the fan <NUM> to establish an outer diameter of the bypass duct <NUM>. The splitter <NUM> may establish an inner diameter of the bypass duct <NUM>. Although depicted as a two-spool turbofan gas turbine engine in the disclosed nonlimiting 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. The engine <NUM> may incorporate a variable area nozzle for varying an exit area of the bypass flow path B and/or a thrust reverser for generating reverse thrust.

The inner shaft <NUM> is connected to the fan <NUM> through a speed change mechanism, which in the exemplary gas turbine engine <NUM> is illustrated as a geared architecture <NUM> to drive the fan <NUM> at a lower speed than the low speed spool <NUM>. The inner shaft <NUM> may interconnect the low pressure compressor <NUM> and low pressure turbine <NUM> such that the low pressure compressor <NUM> and low pressure turbine <NUM> are rotatable at a common speed and in a common direction. In other embodiments, the low pressure turbine <NUM> drives both the fan <NUM> and low pressure compressor <NUM> through the geared architecture <NUM> such that the fan <NUM> and low pressure compressor <NUM> are rotatable at a common speed. Although this application discloses geared architecture <NUM>, its teaching may benefit direct drive engines having no geared architecture.

The low pressure compressor <NUM>, high pressure compressor <NUM>, high pressure turbine <NUM> and low pressure turbine <NUM> each include one or more stages having a row of rotatable airfoils. Each stage may include a row of vanes adjacent the rotatable airfoils. The rotatable airfoils are schematically indicated at <NUM>, and the vanes are schematically indicated at <NUM>.

The engine <NUM> may be a high-bypass geared aircraft engine. The bypass ratio can be greater than or equal to <NUM> and less than or equal to about <NUM>, or more narrowly can be less than or equal to <NUM>. The geared architecture <NUM> may be an epicyclic gear train, such as a planetary gear system or a star gear system. The epicyclic gear train may include a sun gear, a ring gear, a plurality of intermediate gears meshing with the sun gear and ring gear, and a carrier that supports the intermediate gears. The sun gear may provide an input to the gear train. The ring gear (e.g., star gear system) or carrier (e.g., planetary gear system) may provide an output of the gear train to drive the fan <NUM>. A gear reduction ratio may be greater than or equal to <NUM>, or more narrowly greater than or equal to <NUM>, and in some embodiments the gear reduction ratio is greater than or equal to <NUM>. The fan diameter is significantly larger than that of the low pressure compressor <NUM>. The low pressure turbine <NUM> can have a pressure ratio that is greater than or equal to <NUM> and in some embodiments is greater than or equal to <NUM>. All of these parameters are measured at the cruise condition described below.

The engine parameters described above, and those in the next paragraph are measured at this condition unless otherwise specified.

"Fan pressure ratio" is the pressure ratio across the fan blade <NUM> alone, without a Fan Exit Guide Vane ("FEGV") system. A distance is established in a radial direction between the inner and outer diameters of the bypass duct <NUM> at an axial position corresponding to a leading edge of the splitter <NUM> relative to the engine central longitudinal axis A. The fan pressure ratio is a spanwise average of the pressure ratios measured across the fan blade <NUM> alone over radial positions corresponding to the distance. The fan pressure ratio can be less than or equal to <NUM>, or more narrowly greater than or equal to <NUM>, such as between <NUM> and <NUM>. "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 x <NUM>/<NUM>). The corrected fan tip speed can be less than or equal to <NUM> ft / second (<NUM> meters/second), and can be greater than or equal to <NUM> ft / second (<NUM> meters/second).

<FIG> shows a compressor section <NUM> rotating compressor blades <NUM> and <NUM>. Blade <NUM> may be a downstream most compressor blade in the compressor section. Thus, it could be said that blade <NUM> is at a location which is upstream of the highest compression point <NUM> in the compressor section <NUM>.

Housing members <NUM> are associated with stators <NUM>, with one intermediate blades <NUM> and <NUM> and one downstream of the blade <NUM>. A housing member <NUM> is associated with the seal mount <NUM>. Housing <NUM> is connected to an outer housing <NUM> which defines a bleed chamber <NUM> in combination with the housing <NUM> and an inner housing <NUM> which defines an outer wall to the compressor section <NUM>.

Bolts <NUM> connect a housing member <NUM> to the stator mounts <NUM>. An outer housing <NUM> is positioned outwardly of the housing member <NUM>.

A bleed hole <NUM> allows air from a location which is upstream of the most downstream point <NUM> in the compressor section <NUM> to move into the bleed chamber <NUM>. As can be seen, bleed chamber <NUM> is much larger than bleed hole <NUM>. That air will be at high temperature and pressure. In the prior art the air moved across the bleed chamber <NUM> generally directly radially outwardly and against a wall portion <NUM> of the housing <NUM>. Thus, the portion <NUM> which is impacted by the air from the bleed hole <NUM> would increase in temperature, and there are thermal stresses between it and another housing location <NUM> which is spaced radially outwardly.

To address this thermal stress, a flow discourager <NUM> is positioned in the path of the bleed air flow. The flow discourager <NUM> is generally L-shaped. That is, there is a mount section <NUM> which receives the bolt <NUM> to secure the flow discourager <NUM> to the housings <NUM> and <NUM>. A blocking portion <NUM> extends from the mount portion <NUM> such that the blocking portion is within <NUM> degrees of being parallel to the axis of rotation of the compressor section <NUM>. In embodiments, the blocking portion <NUM> is within <NUM> degrees of being parallel. In one embodiment, the blocking section is at a right angle relative to the mount portion <NUM>, and thus the flow discourager <NUM> could be said to be L-shaped.

Now, when air leaves the bleed port <NUM> it impacts on blocking section <NUM>. The air then swirls outwardly around the blocking section <NUM> such that it does not directly impact the wall portion <NUM>.

Air downstream of the chamber <NUM> reaches an optional heat exchanger <NUM>, and then is delivered at <NUM> to a use such as cooling components in a turbine section.

<FIG> shows details of the flow discourager <NUM>. The mount section <NUM> is illustrated at a right angle relative to the blocking section <NUM>. A bolt hole <NUM> extends through the mount portion <NUM>.

<FIG> shows that the flow discourager <NUM> extends generally circumferentially. In embodiments, the flow discourager <NUM> will extend over <NUM> degrees relative to the rotational axis of the gas turbine engine. In one embodiment, there are circumferential two sub-portions to the flow discourager <NUM>. As shown in <FIG>, associated with each bolt hole <NUM> is an open slot <NUM> in the blocking portion <NUM>. This allows wrench access to a bolt <NUM> received in a bolt hole <NUM>.

As shown in <FIG>, the flow discourager <NUM> has the blocking portion <NUM> having an axially forwardmost end <NUM> received in the chamber <NUM>, and which is axially forward of an axially forwardmost point <NUM> of the bleed hole <NUM>. Thus, the vast majority of the air leaving the bleed hole <NUM> will impact upon the blocking portion <NUM>.

<FIG> shows another feature. There are a plurality of bleed holes <NUM> spaced circumferentially. In one embodiment there are eight bleed holes. The bolt holes <NUM>, and hence the slots <NUM>, are circumferentially intermediate adjacent ones of the bleed holes <NUM>. Thus, little or no air leaving the bleed holes <NUM> will impact directly through the slot <NUM>.

The proposed flow discourager has been shown to provide great improvement in reducing the thermal load at the wall portion <NUM>.

The flow discourager may be formed of a sheet metal.

A gas turbine engine under this disclosure could be said to include a compressor section, a turbine section and an intermediate combustor. The compressor section has a plurality of rotating compressor blades rotating about an axis of rotation, and static stator vanes positioned axially intermediate the rotating compressor blades. There are a plurality of bleed holes extending through a compressor outer housing positioned radially outwardly of the rotating compressor blades, and allowing air compressed by the compressor blades to move into a bleed chamber. There is a bleed chamber outer housing positioned radially outwardly of the compressor outer housing, and defining the bleed chamber in part in combination with the compressor outer housing. A flow discourager is positioned radially intermediate the bleed holes and the bleed outer housing, such that air leaving the bleed hole impacts upon the flow discourager before reaching the bleed chamber outer housing.

Claim 1:
A gas turbine engine (<NUM>) comprising:
a compressor section (<NUM>, <NUM>), a turbine section (<NUM>) and an intermediate combustor (<NUM>);
said compressor section (<NUM>, <NUM>) having a plurality of rotating compressor blades (<NUM>, <NUM>) rotating about an axis of rotation, and static stator vanes (<NUM>) positioned axially intermediate the rotating compressor blades (<NUM>, <NUM>);
there being a plurality of bleed holes (<NUM>) extending through a compressor outer housing (<NUM>) positioned radially outwardly of said rotating compressor blades (<NUM>, <NUM>), and allowing air compressed by said compressor blades (<NUM>, <NUM>) to move into a bleed chamber (<NUM>);
there being a bleed chamber outer housing (<NUM>) positioned radially outwardly of said compressor outer housing (<NUM>), and defining said bleed chamber (<NUM>) in part in combination with said compressor outer housing (<NUM>); and
a flow discourager (<NUM>) positioned radially intermediate said bleed holes (<NUM>) and said bleed chamber outer housing (<NUM>), such that air leaving said bleed hole (<NUM>) impacts upon said flow discourager (<NUM>) before reaching said bleed chamber outer housing (<NUM>),
wherein said flow discourager (<NUM>) has a mount portion (<NUM>) which is connected to a mount housing, and a blocking portion (<NUM>) extending from said mount portion (<NUM>) at an angle that is within <NUM> degrees of being parallel to the axis of rotation of the compressor blades (<NUM>, <NUM>);
characterised in that:
there are a plurality of circumferentially spaced bleed holes (<NUM>), and a plurality of fastener or bolt holes (<NUM>) in said mount portion (<NUM>); and
said fastener or bolt holes (<NUM>) are formed to be circumferentially offset and intermediate adjacent ones of said bleed holes (<NUM>).