Patent Application
20160003152
This disclosure relates to a gas turbine engine, and more particularly, to turbine vane platform cooling arrangements that may be incorporated into a gas turbine engine.
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.
High pressure turbine vanes are subjected to high operating temperatures and frequently require film cooling on flow path surfaces. Film cooling is facilitated by machining cooling holes through an exterior surface to an internal cooling passage.
Multi-vaned stators include multiple airfoils supported by a common platform. In some applications, a thermal barrier coating, such as ceramic, is applied to the exterior surface to provide thermal protection to the stator. At least some of the surfaces of the airfoils are obstructed by the adjacent airfoils. As a result, each airfoil may exhibit a different cooling characteristic than the adjacent airfoil.
In one exemplary embodiment, a stator for a gas turbine engine has a platform supporting multiple vanes that includes first and second vanes respectively. First and second regions are arranged at the same location on the first and second vanes. The first and second regions respectively include first and second cooling hole configurations that are different than one another.
In a further embodiment of the above, the first and second cooling hole configurations correspond to cooling hole size.
In a further embodiment of any of the above, the first and second cooling hole configurations corresponding to cooling hole shape.
In a further embodiment of any of the above, the first cooling hole configuration include an oblong exit, and the second cooling hole configuration includes a conical exit.
In a further embodiment of any of the above, the first and second cooling hole configurations corresponding to cooling hole density.
In a further embodiment of any of the above, the first and second regions are the same size. The first and second cooling configurations each include a different number of cooling holes.
In a further embodiment of any of the above, the first and second regions are provided on airfoils.
In a further embodiment of any of the above, the first and second regions are provided on pressure sides.
In a further embodiment of any of the above, the first and second regions are provided on suction sides.
In a further embodiment of any of the above, the first cooling hole configuration includes a cooling hole that has a cooling hole axis that provides a line of sight that is obstructed by the second vane.
In a further embodiment of any of the above, the platform corresponds to an outer platform. An inner platform supports the multiple vanes.
In a further embodiment of any of the above, the number of multiple vanes is two.
In one exemplary embodiment, a gas turbine engine includes a combustor that is in fluid communication with the compressor section. A turbine section in fluid communication with the combustor. The turbine section includes a platform supporting multiple vanes including first and second vanes respectively including first and second regions. The first and second regions are arranged at the same location on the first and second vanes. The first and second regions respectively include first and second cooling hole configurations that are different than one another.
In a further embodiment of any of the above, the first and second cooling hole configurations corresponding to cooling hole size.
In a further embodiment of any of the above, the first and second cooling hole configurations corresponding to cooling hole shape.
In a further embodiment of any of the above, the first and second cooling hole configurations corresponding to cooling hole density.
The disclosure can be further understood by reference to the following detailed description when considered in connection with the accompanying drawings wherein:
Although the disclosed non-limiting embodiment depicts a turbofan gas turbine engine, it should be understood that the concepts described herein are not limited to use with turbofans as the teachings may be applied to other types of turbine engines; for example a turbine engine including a three-spool architecture in which three spools concentrically rotate about a common axis and where a low spool enables a low pressure turbine to drive a fan via a gearbox, an intermediate spool that enables an intermediate pressure turbine to drive a first compressor of the compressor section, and a high spool that enables a high pressure turbine to drive a high pressure compressor of the compressor section.
The example 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.
The low speed spool 30 generally includes an inner shaft 40 that connects a fan 42 and a low pressure (or first) compressor section 44 to a low pressure (or second) turbine section 46. The inner shaft 40 drives the fan 42 through a speed change device, such 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 (or first) compressor section 52 and a high pressure (or second) turbine section 54. The inner shaft 40 and the outer shaft 50 are concentric and rotate via the bearing systems 38 about the engine central longitudinal axis A.
A combustor 56 is arranged between the high pressure compressor 52 and the high pressure turbine 54. In one example, the high pressure turbine 54 includes at least two stages to provide a double stage high pressure turbine 54. In another example, the high pressure turbine 54 includes only a single stage. As used herein, a “high pressure” compressor or turbine experiences a higher pressure than a corresponding “low pressure” compressor or turbine.
The example low pressure turbine 46 has a pressure ratio that is greater than about 5. The pressure ratio of the example low pressure turbine 46 is measured prior to an inlet of the low pressure turbine 46 as related to the pressure measured at the outlet of the low pressure turbine 46 prior to an exhaust nozzle.
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 as well as setting airflow entering the low pressure turbine 46.
The core airflow C is compressed by the low pressure compressor 44 then by the high pressure compressor 52 mixed with fuel and ignited in the combustor 56 to produce high speed exhaust gases that are then expanded through the high pressure turbine 54 and low pressure turbine 46. The mid-turbine frame 57 includes vanes 59, which are in the core airflow path and function as an inlet guide vane for the low pressure turbine 46. Utilizing the vane 59 of the mid-turbine frame 57 as the inlet guide vane for low pressure turbine 46 decreases the length of the low pressure turbine 46 without increasing the axial length of the mid-turbine frame 57. Reducing or eliminating the number of vanes in the low pressure turbine 46 shortens the axial length of the turbine section 28. Thus, the compactness of the gas turbine engine 20 is increased and a higher power density may be achieved.
The disclosed gas turbine engine 20 in one example is a high-bypass geared aircraft engine. In a further example, the gas turbine engine 20 includes a bypass ratio greater than about six (6), with an example embodiment being greater than about ten (10). The example geared architecture 48 is an epicyclical gear train, such as a planetary gear system, star gear system or other known gear system, with a gear reduction ratio of greater than about 2.3.
In one disclosed embodiment, the gas turbine engine 20 includes a bypass ratio greater than about ten (10:1) and the fan diameter is significantly larger than an outer diameter of the low pressure compressor 44. It should be understood, however, that the above parameters are only exemplary of one embodiment of a gas turbine engine including a geared architecture and that the present disclosure is applicable to other gas turbine engines.
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. The flight condition of 0.8 Mach and 35,000 ft., with the engine at its best fuel consumption—also known as “bucket cruise Thrust Specific Fuel Consumption ('TSFC')”—is the industry standard parameter of pound-mass (lbm) of fuel per hour being burned divided by pound-force (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 1.50. In another non-limiting embodiment the low fan pressure ratio 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)]0.5. The “Low corrected fan tip speed”, as disclosed herein according to one non-limiting embodiment, is less than about 1150 ft/second.
Referring to
The turbine blades each include a tip 80 adjacent to a blade outer air seal 70 of a case structure 72. The first and second stage arrays 54a, 54c of turbine vanes and first and second stage arrays 54b, 54d of turbine blades are arranged within a core flow path C and are operatively connected to a spool 32.
Each vane 60 includes an inner platform 74 and an outer platform 76 respectively defining inner and outer flow paths. The platforms 74, 76 are interconnected by an airfoil 78 extending in a radial direction R. It should be understood that the turbine vanes may be discrete from one another or arranged in integrated clusters. For example, a “doublet” vane cluster is illustrated in
The turbine vanes 60, 62 are constructed from a high strength, heat resistant material such as a nickel-based or cobalt-based superalloy, or of a high temperature, stress resistant ceramic or composite material. In cooled configurations, internal fluid passages and external cooling apertures can provide for a combination of impingement and film cooling. Other internal cooling approaches may be used such as trip strips, pedestals or other convective cooling techniques. In addition, one or more thermal barrier coatings, abrasion-resistant coatings or other protective coatings may be applied to the turbine vane 60, 62.
Referring to
Referring to
An example cooling hole 104 is shown in the first vane 78. The cooling hole 104 includes an axis having a line of sight 106 that is obstructed by the adjacent second vane 178. As a result, the coating 110 may be thinner than the coating 210. Thus, it is desirable to provide enhanced cooling on the first vane 78 where the thermal barrier coating is thinner.
Cooling holes provided in obstructed line of sight areas may be machined by EDM drilling, whereas it may be desirable to machine fully accessible cooling holes using a laser.
The different cooling hole configurations are provided at the same location, for example, on either the pressure or suctions sides, on the first and second vanes 78, 178. The different cooling hole configurations may correspond to cooling different hole sizes, cooling hole shapes and/or cooling hole densities.
Example different cooling hole sizes and shapes are shown in
The examples shown in
Although an example embodiment has 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 reason, the following claims should be studied to determine their true scope and content.
| Filing Document | Filing Date | Country | Kind |
|---|---|---|---|
| PCT/US2014/024240 | 3/12/2014 | WO | 00 |
| Number | Date | Country | |
|---|---|---|---|
| 61782685 | Mar 2013 | US |
