The present disclosure relates to a gas turbine engine and, more particularly, to the protection of turbine vanes from particulate blockage of airfoil cooling circuits.
Gas turbine engines typically include a compressor section to pressurize airflow, a combustor section to burn a hydrocarbon fuel in the presence of the pressurized air, and a turbine section to extract energy from the resultant combustion gases. The combustion gases commonly exceed 2000 degrees F. (1093 degrees C.).
Cooling of engine components such as the high pressure turbine vane may be complicated by the presence of entrained particulates in the secondary cooling air that are carried through the engine. During engine operation a single point feed passage to each airfoil cooling circuit may be prone to blockage by foreign object particles. If these single source feed apertures become blocked, the associated downstream airfoil cooling circuit is starved of cooling air which may result in airfoil distress.
A vane ring for a gas turbine engine component according to one disclosed non-limiting embodiment of the present disclosure includes a multiple of vanes that extend between the inner vane platform and the outer vane platform, each of the multiple of vanes contains an airfoil cooling circuit that receives cooing airflow through a respective one of a multiple of feed passages; a multiple of metering passages in the outer vane platform, each of the multiple of metering passages in communication with the respective one of the multiple of feed passages; and a multiple of secondary passages in the outer vane platform, each of the multiple of secondary passages in communication with the respective one of the multiple of metering passages.
A further embodiment of any of the foregoing embodiments of the present disclosure includes that each of the multiple of secondary passages is a branch from one of the multiple of metering passages
A further embodiment of any of the foregoing embodiments of the present disclosure includes that the secondary passage diameter is equivalent to the metering passage diameter.
A further embodiment of any of the foregoing embodiments of the present disclosure includes that the secondary passage is at an angle to the metering passage.
A further embodiment of any of the foregoing embodiments of the present disclosure includes that the metering passage is circular in cross-section.
A further embodiment of any of the foregoing embodiments of the present disclosure includes that the secondary passage is circular in cross section.
A further embodiment of any of the foregoing embodiments of the present disclosure includes that the multiple of metering passages and the multiple of secondary passages are located within a rail of the outer vane platform.
A further embodiment of any of the foregoing embodiments of the present disclosure includes that the multiple of metering passages and the multiple of secondary passages are located within a hooked rail of the outer vane platform.
A further embodiment of any of the foregoing embodiments of the present disclosure includes that the metering passage and the secondary passage are formed in a surface transverse to the axis.
A further embodiment of any of the foregoing embodiments of the present disclosure includes that the airflow scrubs along the surface.
A vane ring for a gas turbine engine component according to one disclosed non-limiting embodiment of the present disclosure includes an inner vane platform around an axis; a multiple of vanes that extend between the inner vane platform and the outer vane platform, each of the multiple of vanes contains an airfoil cooling circuit that receives cooing airflow through a respective one of a multiple of feed passages; a hooked rail that extends from the outer vane platform; a multiple of metering passages in the hooked rail, each of the multiple of metering passages in communication with one of the multiple of feed passages; and a multiple of secondary passages recessed in the hooked rail, each of the multiple of secondary passages in communication with one of the multiple of metering passages.
A further embodiment of any of the foregoing embodiments of the present disclosure includes that each of the multiple of the metering passages and each of the multiple of secondary passages are formed in a surface of the hooked rail transverse to the axis.
A further embodiment of any of the foregoing embodiments of the present disclosure includes that the airflow scrubs along the surface of the hooked rail.
A further embodiment of any of the foregoing embodiments of the present disclosure includes that the airflow is a cooling airflow.
A further embodiment of any of the foregoing embodiments of the present disclosure includes that the vane ring is in a second turbine stage.
A further embodiment of any of the foregoing embodiments of the present disclosure includes that each of the multiple of secondary passages is a branch from one of the multiple of metering passages
A further embodiment of any of the foregoing embodiments of the present disclosure includes that the secondary passage diameter is equivalent to the metering passage diameter.
A further embodiment of any of the foregoing embodiments of the present disclosure includes that the secondary passage is at an angle to the metering passage.
A method of communicating airflow into an airfoil cooling circuit of each of a multiple of vanes though a respective feed passage of a gas turbine engine component according to one disclosed non-limiting embodiment of the present disclosure includes flowing a cooling airflow through an entrance in a surface of a hooked rail to a secondary passage, then to a metering passage in communication with a feed passage to an airfoil cooling circuit for each of the multiple of vanes.
A further embodiment of any of the foregoing embodiments of the present disclosure includes that the airflow scrubs along the surface.
A further embodiment of any of the foregoing embodiments of the present disclosure includes that the airflow is a cooling airflow.
A further embodiment of any of the foregoing embodiments of the present disclosure includes that the secondary passage is angled with respect to the metering passage.
The foregoing features and elements may be combined in various combinations without exclusivity, unless expressly indicated otherwise. These features and elements as well as the operation thereof will become more apparent in light of the following description and the accompanying drawings. It should be appreciated; however, the following description and drawings are intended to be exemplary in nature and non-limiting.
Various features will become apparent to those skilled in the art from the following detailed description of the disclosed non-limiting embodiments. The drawings that accompany the detailed description can be briefly described as follows:
The engine 20 generally includes a low spool 30 and a high spool 32 mounted for rotation about an engine central longitudinal axis A relative to an engine case structure 36 via several bearing structures 38. The low spool 30 generally includes an inner shaft 40 that interconnects a fan 42, a low pressure compressor (“LPC”) 44 and a low pressure turbine (“LPT”) 46. The inner shaft 40 drives the fan 42 directly or through a geared architecture 48 to drive the fan 42 at a lower speed than the low spool 30. An exemplary reduction transmission is an epicyclic transmission, namely a planetary or star gear system.
The high spool 32 includes an outer shaft 50 that interconnects a high pressure compressor (“HPC”) 52 and high pressure turbine (“HPT”) 54. A combustor 56 is arranged between the high pressure compressor 52 and the high pressure turbine 54. The inner shaft 40 and the outer shaft 50 are concentric and rotate about the engine central longitudinal axis A which is collinear with their longitudinal axes.
Core airflow is compressed by the LPC 44 then the HPC 52, mixed with the fuel and burned in the combustor 56, then the combustion gasses are expanded over the HPT 54 and the LPT 46. The turbines 46, 54 rotationally drive the respective low spool 30 and high spool 32 in response to the expansion. The main engine shafts 40, 50 are supported at a plurality of points by bearing assemblies 38 within the engine case structure 36.
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The blade outer air seal (BOAS) assembly 62 is affixed to the engine case structure 36 to form an annular chamber between the blade outer air seal (BOAS) assembly 62 and the engine case structure 36. The blade outer air seal (BOAS) assembly 62 bounds the working medium combustion gas flow in a primary flow path 94. The vane rings 68, 70 align the flow of the working medium combustion gas flow while the rotor blades 90 collect the energy of the working medium combustion gas flow to drive the turbine section 28 which in turn drives the compressor section 24.
The forward stationary vane ring 68 is mounted to the engine case structure 36 upstream of the blade outer air seal (BOAS) assembly 62 by a vane support 96. The vane support 96, for example, may include a rail 97 that extends from the outer vane platform 80 that is fastened to the engine case structure 36. The rail 97 includes a multitude of apertures 99 spaced therearound to communicate cooling air “C” into the vanes 72 as well as downstream thereof. Cooling air “C”, also referred to as secondary airflow, often contains foreign object particulates (such as sand). As only a specific quantity of cooling air “C” is required, the cooling air “C” is usually metered to minimally affect engine efficiency.
The aft stationary vane ring 70 is mounted to the engine case structure 36 downstream of the blade outer air seal (BOAS) assembly 62 by a vane support 98. The vane support 98 extends from the outer vane platform 82 and may include an annular hooked rail 84 (also shown in
The annular hooked rail 84 includes a feed passage 100 (also shown in
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Cooling airflow within the plenum 120 adjacent the outer vane platform 80, 82 generally scrubs along the surface 122 such that foreign object particles therein have a lessened tendency to enter the metering passage 132 and the secondary passages 134, 136, 138, 140 as they are displaced from the surface 122. Nonetheless, should one passage become blocked, the other passages permit unobstructed flow into the airfoil cooling circuit 102 within the vane 74.
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During operation of the engine, cooling flow “C” from the high pressure compressor flows around the combustor and into the first vane cavity 102. This cooling air has particulates entrained in it. These particulates are present in the working medium flow path as ingested from the environment by the engine. The majority of the particulates are very fine in size, thus they are carried through the sections of the engine as the working medium gases flow axially downstream. Should a particle be of a size to block the metering passage, the secondary flow passages necessarily permit communication of at least a portion of the cooling air which significantly reduces the risk of damage to the airfoil and increases component field life.
Although particular step sequences are shown, described, and claimed, it should be appreciated that steps may be performed in any order, separated or combined unless otherwise indicated and will still benefit from the present disclosure.
The foregoing description is exemplary rather than defined by the limitations within. Various non-limiting embodiments are disclosed herein, however, one of ordinary skill in the art would recognize that various modifications and variations in light of the above teachings will fall within the scope of the appended claims. It is therefore to be appreciated that within the scope of the appended claims, the disclosure may be practiced other than as specifically described. For that reason, the appended claims should be studied to determine true scope and content.