Patent Application
20160230581
This disclosure relates to a gas turbine engine seal used in an airfoil component array. More particularly, the disclosure relates to a cooling hole provided in the seal arranged at an airfoil component platform.
Gas turbine engines typically include a compressor section, a combustor section and a turbine section. During operation, air is pressurized in the compressor section and is mixed with fuel and burned in the combustor section to generate hot combustion gases. The hot combustion gases are communicated through the turbine section, which extracts energy from the hot combustion gases to power the compressor section and other gas turbine engine loads.
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. The turbine vanes, which generally do not rotate, guide the airflow and prepare it for the next set of blades.
Circumferential seals are used between adjacent airfoil components, such as turbine vanes and blades. The seals are provided at an inner gas flow path to seal between adjacent platforms of the airfoil components. Gaps are typically provided circumferentially between adjacent lateral faces of adjoining platforms to accommodate thermal growth during engine operation. Round holes have been provided in the seal that are in communication with the gap. The holes are normal to the sealing surface of the seal.
In one exemplary embodiment, a gas turbine engine component array includes first and second components each having a platform. The platforms are arranged adjacent to one another and provide a gap. A seal is arranged circumferentially between the first and second components and in engagement with the platforms to obstruct the gap. A cooling hole is provided in the seal and is in fluid communication with the gap. The cooling hole has an increasing taper toward the gap.
In a further embodiment of the above, the seal includes a gas path flow side that engages the platform. Another side is arranged opposite the gas path flow side and facing a cavity provided between the first and second components. The cooling hole is configured to fluidly connect the cavity to the gap.
In a further embodiment of any of the above, the seal is a damper seal arranged in the pocket.
In a further embodiment of any of the above, the seal extends axially lateral faces of the platform to obstruct the gap.
In a further embodiment of any of the above, a slot is provided in each of the lateral faces. The seal is a feather seal arranged within the slots.
In a further embodiment of any of the above, the lateral faces overlap the cooling hole in a circumferential direction.
In a further embodiment of any of the above, the cooling hole is arranged at an acute angle with respect to the gas path flow side. The cooling hole extends generally in a lenghtwise direction of the gap.
In a further embodiment of any of the above, the cooling hole includes a metering portion and a diffuser portion.
In a further embodiment of any of the above, the metering portion has a length L and a diameter D. The metering portion has an L/D ratio of greater than 1.
In a further embodiment of any of the above, the L/D ratio is greater than 3.
In a further embodiment of any of the above, the diffuser portion has a width greater than height. Width is arranged in a circumferential direction.
In a further embodiment of any of the above, the first and second components are blade outer airseals or turbine blades.
In another exemplary embodiment, a gas turbine engine component seal includes a wall that has a gas flow path side and another side opposite the gas flow path side. A cooling hole is provided in the wall and extends in a direction at an acute angle relative to the gas path flow side. The cooling hole has an increasing taper toward the gas path flow side.
In a further embodiment of the above, the cooling hole includes a metering portion and a diffuser portion. The metering portion has a length L and a diameter D. The metering portion includes an L/D ratio of greater than 1.
In a further embodiment of any of the above, the L/D ratio is greater than 3.
In a further embodiment of any of the above, the diffuser portion has a width greater than height. The wall has a length and a circumferential width that is less than the length. The width of the diffuser portion is oriented in generally the same direction as the circumferential width.
In another exemplary embodiment, a method of cooling a gas turbine engine component array includes the steps of providing cooling fluid to a cavity between adjacent components, flowing cooling fluid from the cavity through a cooling hole in a seal provided between the adjacent components, and diffusing the cooling fluid through the cooling hole on a gas path flow side of the seal opposite the cavity to create a cooling film in a gap provided between adjacent platforms of the components.
In a further embodiment of the above, the cooling hole is arranged at an acute angle with respect to the gas path flow side. The cooling hole extends generally in a lengthwise direction of the gap. The cooling hole includes a metering portion and a diffuser portion.
In a further embodiment of any of the above, the metering portion has a length L and a diameter D and an L/D ratio of greater than 1. The diffuser portion has a width greater than height. The width is arranged in a circumferential direction.
In a further embodiment of any of the above, the adjacent components are blade outer airseals or turbine blades.
The disclosure can be further understood by reference to the following detailed description when considered in connection with the accompanying drawings wherein:
The embodiments, examples and alternatives of the preceding paragraphs, the claims, or the following description and drawings, including any of their various aspects or respective individual features, may be taken independently or in any combination. Features described in connection with one embodiment are applicable to all embodiments, unless such features are incompatible.
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.
The low speed spool 30 generally includes an inner shaft 40 that interconnects a fan 42, a low pressure compressor 44 and a 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 high pressure compressor 52 and 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 supports one or more 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.
The core airflow C 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 over 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. 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.
The disclosed serpentine cooling passage may be used in various gas turbine engine components. For exemplary purposes, a turbine blade 64 is described. It should be understood that the cooling passage may also be used in vanes, blade outer air seals, and turbine platforms, for example.
Referring to
With continuing reference to
The airfoil 78 of
The airfoil 78 includes a cooling passage 90 provided between the pressure and suction walls 86, 88. The exterior airfoil surface may include multiple film cooling holes (not shown) in fluid communication with the cooling passage 90. Typically, the cooling fluid source that provides cooling fluid to the cooling passage 90 as provides some fluid into a pocket 75.
Adjoining platforms 76 provide a gap 104 circumferentially between lateral faces 81 of the adjoining platform 76. A seal 77 is arranged between the airfoils 78 and are in engagement with the platform 76. The seal 77 is positioned to obstruct the gap 104 in the radial direction to prevent gas path flow from entering the pockets 75.
Referring to
The pressure of fluid within the cavity 94 is greater than the pressure of fluid at the inner gas flow path. The seal 77 includes one or more cooling holes 102 (only one shown for clarity) that extends from the other side 98 to the gas path flow side 96 to provide fluid communication from the cavity 94 to the gap 104. In the example, the cooling hole 102 includes an increasing taper toward the gap 104 to diffuse the fluid flow through the cooling hole 102 as it exits the seal 77 into the gap 104. In this manner, the velocity of the cooling fluid through the hole 102 is slowed, such that the cooling fluid will linger within the gap 104 forming a boundary layer of cooling film.
Referring to
The metering portion 108 has a diameter D (hydraulic diameter D if the cross-sectional area of the metering section 108 is not circular) and a length L that provides an L/D ratio of greater than 1, and in one example, greater than 3.
The cooling hole 102 can be any suitable shape and may be drilled or electro-discharge machined into the seal 77.
Referring to
It should also be understood that although a particular component arrangement is disclosed in the illustrated embodiment, other arrangements will benefit herefrom. Although particular step sequences are shown, described, and claimed, it should be understood that steps may be performed in any order, separated or combined unless otherwise indicated and will still benefit from the present invention.
Although the different examples have specific components shown in the illustrations, embodiments of this invention are not limited to those particular combinations. It is possible to use some of the components or features from one of the examples in combination with features or components from another one of the examples.
Although example embodiments have been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of the claims. For that and other reasons, the following claims should be studied to determine their true scope and content.
This application claims priority to U.S. Provisional Application No. 61/879,009, which was filed on Sep. 17, 2013 and is incorporated herein by reference.
This invention was made with government support under Contract No. FA8650-09-D-29230021 awarded by the United States Air Force. The Government has certain rights in this invention.
| Filing Document | Filing Date | Country | Kind |
|---|---|---|---|
| PCT/US2014/055193 | 9/11/2014 | WO | 00 |
| Number | Date | Country | |
|---|---|---|---|
| 61879009 | Sep 2013 | US |
