Cooled cooling air for blade air seal through outer chamber

Abstract

A gas turbine engine according to an example of the present disclosure include a compressor section, a combustor, and a turbine section. The combustor has a radially outer surface that defines a diffuser chamber radially outwardly of the combustor. The turbine section has a high pressure turbine first stage blade that has an outer tip, and a blade outer air seal positioned radially outwardly of the outer tip. A tap for tapping air has been compressed by the compressor and is passed through a heat exchanger. The air downstream of the heat exchanger passes through at least one pipe and into a manifold radially outward of the blade outer air seal, and then passes across the blade outer air seal to cool the blade outer air seal.

Claims

1. A gas turbine engine comprising: a compressor section, a combustor, and a turbine section, said combustor having a radially outer surface defining a diffuser chamber radially outwardly of said combustor; said turbine section including a high pressure turbine first stage blade having an outer tip, and a blade outer air seal positioned radially outwardly of said outer tip; a tap for tapping air having been compressed by said compressor section being passed through a heat exchanger; a first portion of said air downstream of said heat exchanger passing through a plurality of pipes and into a manifold radially outward of said blade outer air seal, and then passing across said blade outer air seal to cool said blade outer air seal; wherein a second portion of said air downstream of said heat exchanger passing into a mixing chamber without passing through said manifold, and is mixed with higher temperature air from a diffuser chamber outwardly of said combustor, and mixed air passed to cool a first stage blade row in a high pressure turbine; wherein said blade outer air seal includes at least two components having different thermal coefficients of expansion to provide clearance control between said outer tip and an inner periphery of said seal portion; wherein a valve controls the air passing to said blade outer air seal, and allows control of at least one of an amount, a pressure or a temperature of the air being delivered to the blade outer air seal; wherein said diffuser chamber has an outer boundary defined by an outer core housing and said pipes are radially outward of said outer core housing; wherein said manifold is also outwardly of said outer core housing and communicates with passages passing through said outer core housing to said blade outer air seal; wherein said air flowing upstream of said blade outer air seal being routed through holes in a seal portion of said blade outer air seal to cool adjacent a leading edge of said blade outer air seal and the air passing downstream of said blade outer air seal passing through holes in said seal portion of said blade outer air seal to cool adjacent a trailing edge of said blade outer air seal; and wherein said mixing chamber is radially outward of a compressor diffuser defined downstream of a downstream most location in a high pressure compressor section and said mixed air from said mixing chamber passing through vanes in said compressor diffuser.

2. A gas turbine engine comprising: a compressor section, a combustor, and a turbine section, said combustor having a radially outer surface defining a diffuser chamber radially outwardly of said combustor; said turbine section including a high pressure turbine first stage blade having an outer tip, and a blade outer air seal positioned radially outwardly of said outer tip; a tap for tapping air having been compressed by said compressor section being passed through a heat exchanger; a first portion of said air downstream of said heat exchanger passing through a plurality of pipes and into a manifold radially outward of said blade outer air seal, and then passing across said blade outer air seal to cool said blade outer air seal; wherein a second portion of said air downstream of said heat exchanger passing into a mixing chamber without passing through said manifold, and is mixed with higher temperature air from a diffuser chamber outwardly of said combustor, and mixed air passed to cool a first stage blade row in a high pressure turbine; wherein said blade outer air seal includes at least two components having different thermal coefficients of expansion to provide clearance control between said outer tip and an inner periphery of said seal portion; wherein a valve controls the air passing to said blade outer air seal, and allows control of at least one of an amount, a pressure or a temperature of the air being delivered to the blade outer air seal; and wherein said mixing chamber is radially outward of a compressor diffuser defined downstream of a downstream most location in a high pressure compressor section and said mixed air from said mixing chamber passing through vanes in said compressor diffuser.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

(1) FIG. 1 schematically shows a gas turbine engine.

(2) FIG. 2 shows a cooling system.

(3) FIG. 3 shows details of a blade outer air seal cooling system.

DETAILED DESCRIPTION

(4) FIG. 1 schematically illustrates a gas turbine engine 20. The gas turbine engine 20 is disclosed herein as a two-spool turbofan that generally incorporates a fan section 22, a compressor section 24, a combustor section 26 and a turbine section 28. Alternative engines might include an augmentor section (not shown) among other systems or features. The fan section 22 drives air along a bypass flow path B in a bypass duct defined within a nacelle 15, while the compressor section 24 drives air along a core flow path C for compression and communication into the combustor section 26 then expansion through the turbine section 28. 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.

(5) The exemplary engine 20 generally includes a low speed spool 30 and a high speed spool 32 mounted for rotation about an engine central longitudinal axis A relative to an engine static structure 36 via several bearing systems 38. It should be understood that various bearing systems 38 at various locations may alternatively or additionally be provided, and the location of bearing systems 38 may be varied as appropriate to the application.

(6) The low speed spool 30 generally includes an inner shaft 40 that interconnects a fan 42, a first (or low) pressure compressor 44 and a first (or low) pressure turbine 46. The inner shaft 40 is connected to the fan 42 through a speed change mechanism, which in exemplary gas turbine engine 20 is illustrated as a geared architecture 48 to drive the fan 42 at a lower speed than the low speed spool 30. The high speed spool 32 includes an outer shaft 50 that interconnects a second (or high) pressure compressor 52 and a second (or high) pressure turbine 54. A combustor 56 is arranged in exemplary gas turbine 20 between the high pressure compressor 52 and the high pressure turbine 54. A mid-turbine frame 57 of the engine static structure 36 is arranged generally between the high pressure turbine 54 and the low pressure turbine 46. The mid-turbine frame 57 further supports bearing systems 38 in the turbine section 28. The inner shaft 40 and the outer shaft 50 are concentric and rotate via bearing systems 38 about the engine central longitudinal axis A which is collinear with their longitudinal axes.

(7) The core airflow is compressed by the low pressure compressor 44 then the high pressure compressor 52, mixed and burned with fuel in the combustor 56, then expanded through the high pressure turbine 54 and low pressure turbine 46. The mid-turbine frame 57 includes airfoils 59 which are in the core airflow path C. The turbines 46, 54 rotationally drive the respective low speed spool 30 and high speed spool 32 in response to the expansion. It will be appreciated that each of the positions of the fan section 22, compressor section 24, combustor section 26, turbine section 28, and fan drive gear system 48 may be varied. For example, gear system 48 may be located aft of combustor section 26 or even aft of turbine section 28, and fan section 22 may be positioned forward or aft of the location of gear system 48.

(8) The engine 20 in one example is a high-bypass geared aircraft engine. In a further example, the engine 20 bypass ratio is greater than about six (6), with an example embodiment being greater than about ten (10), the geared architecture 48 is an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3 and the low pressure turbine 46 has a pressure ratio that is greater than about five. In one disclosed embodiment, the engine 20 bypass ratio is greater than about ten (10:1), the fan diameter is significantly larger than that of the low pressure compressor 44, and the low pressure turbine 46 has a pressure ratio that is greater than about five 5:1. Low pressure turbine 46 pressure ratio is pressure measured prior to inlet of low pressure turbine 46 as related to the pressure at the outlet of the low pressure turbine 46 prior to an exhaust nozzle. The geared architecture 48 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 2.3:1. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention is applicable to other gas turbine engines including direct drive turbofans.

(9) A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The fan section 22 of the engine 20 is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet (10,668 meters). The flight condition of 0.8 Mach and 35,000 ft (10,668 meters), with the engine at its best fuel consumption—also known as “bucket cruise Thrust Specific Fuel Consumption (‘TSFCT’)”—is the industry standard parameter of lbm of fuel being burned divided by lbf of thrust the engine produces at that minimum point. “Low fan pressure ratio” is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.45. “Low corrected fan tip speed” is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram °R)/(518.7°R)].sup.0.5. The “Low corrected fan tip speed” as disclosed herein according to one non-limiting embodiment is less than about 1150 ft/second (350.5 meters/second).

(10) FIG. 2 shows a cooling system 100 for cooling turbine components. As shown, a compressor section 101 is provided with a tap 102 for tapping pressurized air.

(11) The tap 102 may be at a location upstream from a downstream most portion of a high pressure compressor, in which case, it is typically provided with a boost compressor to raise its pressure. Alternatively, the air can be tapped from a location 99 where it has been fully compressed by the high pressure compressor.

(12) In either case, pressurized air passes through a heat exchanger 104 where it is cooled, such as by air. In one embodiment, the heat exchanger 104 may be in the bypass duct as described in FIG. 1. From heat exchanger 104, air passes into conduit 106.

(13) From the conduit 106, the air passes into a mixing chamber 108, which may be outward of a compressor diffuser 109. The air passes through vanes in the compressor diffuser 109, such that it is separate from the air downstream of a downstream most compression point 99. The air passes, as shown at 116, to cool a turbine blade 118. In the mixing chamber 108, hot air is shown at 110 mixing with the cool high pressure air from the conduit 106. This air is from a diffuser chamber 112, and is at the pressure downstream of the downstream most point 99. As such, it mixes easily with the air in the mixing chamber such that the air delivered at 116 is not unduly cool.

(14) The chamber 112 is outward of a combustion chamber 114. An outer core housing 113 is positioned outwardly of the chamber 112.

(15) A plurality of pipes 120 (only one of which is shown) tap air from the conduit 106 upstream of the mixing chamber 108. As such, this air is entirely the cooled high pressure cooling air. The air from the plurality of pipes is delivered into a manifold 122 which extends circumferentially over more than 270° about an axis of rotation of the engine. In embodiments, the manifold 122 extends over 360° about the axis of rotation. That air then passes through a plurality of pipes 124 to cool a blade outer air seal 126.

(16) A valve 150 is shown schematically. The valve 150 may be controlled by a control 151 to control the cooling air being sent to the blade outer air seal 126. As an example, the valve may control the amount, pressure or temperature of the air being delivered to the blade outer air seal 126. An optional line 152 may selectively bypass the heat exchanger 104 to allow temperature control, as an example.

(17) FIG. 3 shows details. As shown, air from the pipes 120 enters manifold 122 and then flows through pipe 124 to the blade outer air seal 126. The blade outer air seal 126 is shown to have components 128 and 130 which are formed of materials having distinct coefficients of thermal expansion. These components expand at different rates in response to exposure to heat and provide clearance control for a clearance between an inner portion of a seal 132 and an outer tip of the first stage high pressure turbine blade 118 as known.

(18) As shown, the air flows at 134 upstream of the blade outer air seal components 128 and 130 and through holes 137 to cool a leading edge 136 of the seal portion 132. This cooling air drives, or controls, the expansion of the components and thus the clearance control.

(19) Similarly, the air flows at 138 downstream of the blade outer air seal and through holes 142 to cool a trailing edge 140 of the blade outer air seal.

(20) Although an embodiment of this invention has been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of this invention. For that reason, the following claims should be studied to determine the true scope and content of this invention.