Patent Application
20160108735
The present disclosure relates to rotor blades, such as rotor blades in gas turbine engines. Traditionally, gas turbine engines can include multiple stages of rotor blades and stator vanes to condition and guide fluid flow through the compressor and/or turbine sections. Stages in the high pressure compressor section can include alternating rotor blade stages and stator vane stages. Each vane in a stator vane stage can interface with a seal on the rotor disk, for example, a knife edge seal. The knife edge seals can be one source of increased temperature in the high-pressure compressor due to windage heat-up. Increased temperatures can reduce the durability of aerospace components, specifically those in the last stages of the high pressure compressor.
Such conventional methods and systems have generally been considered satisfactory for their intended purpose. However, there is still a need in the art for improved gas turbine engines.
A gas turbine engine includes a compressor section and a compressor case with a low pressure compressor (LPC) and a high pressure compressor (HPC). The HPC is aft of the LPC. The compressor case defines a centerline axis. The compressor section also includes a rotor disk defined between the compressor case and the centerline axis. A plurality of stages defined radially inward relative to the compressor case. The plurality of stages includes at least one tandem blade stage. The tandem blade stage includes a plurality of blade pairs. Each blade pair is circumferentially spaced apart from the other blade pairs, and is operatively connected to the rotor disk. Each blade pair includes a forward blade and an aft blade. The aft blade is configured to further condition air flow with respect to the forward blade without an intervening stator vane stage shrouded cavity therebetween.
In certain embodiments, a leading edge of each aft blade can be defined forward of a trailing edge of a respective forward blade with respect to the centerline axis. The gas turbine engine can also include a plurality of circumferentially disposed blade platforms defined radially between the rotor disk and the blade pairs. Each blade pair can be integrally formed with a respective one of the blade platforms. The gas turbine engine can include an exit guide vane stage aft of the tandem blade stage. The exit guide vane stage can define the end of the compressor section.
In another aspect, the plurality of stages can include at least one forward stator vane stage forward of the tandem blade stage. The forward stator vane stage can include a plurality of circumferentially disposed stator vanes. Each stator vane can extend from a vane root to a vane tip along a respective vane axis and can be operatively connected to a forward shrouded cavity disposed radially between each respective vane root and the rotor disk. A forward knife edge seal can be between the rotor disk and an inner diameter surface of the forward shrouded cavity. The forward stator vane stage and the tandem blade stage can define the last two sequential stages before the exit guide vane stage.
It is contemplated that the gas turbine engine can include a tandem stator vane stage aft of the tandem blade stage. The tandem stator vane stage can include at least one stator vane pair extending radially between the compressor case and the centerline axis. Each stator vane pair can include a forward stator vane and an aft stator vane. A leading edge of each aft stator vane can be defined forward of a trailing edge of its respective forward stator vane with respect to the centerline axis. The tandem stator vane stage can define the end of the compressor section and the tandem blade stage and the tandem stator vane stage can define the last two sequential stages in the compressor section. In another aspect, a turbomachine can include a stator vane stage and a tandem blade stage aft of the stator vane stage, similar to stator vane and tandem blade stages described above.
These and other features of the systems and methods of the subject disclosure will become more readily apparent to those skilled in the art from the following detailed description of the preferred embodiments taken in conjunction with the drawings.
So that those skilled in the art to which the subject disclosure appertains will readily understand how to make and use the devices and methods of the subject disclosure without undue experimentation, preferred embodiments thereof will be described in detail herein below with reference to certain figures, wherein:
Reference will now be made to the drawings wherein like reference numerals identify similar structural features or aspects of the subject disclosure. For purposes of explanation and illustration, and not limitation, a cross-sectional view of an exemplary embodiment of the gas turbine engine 100 constructed in accordance with the disclosure is shown in
As shown in
Gas turbine engine 10 also includes an inner shaft 30 that interconnects a fan 32, a LPC 34 and a low pressure turbine 36. Inner shaft 30 is connected to fan 32 through a speed change mechanism, which in exemplary gas turbine engine 10 is illustrated as a geared architecture 38. An outer shaft 40 interconnects a HPC 42 and high pressure turbine 44. A combustor 46 is arranged between HPC 42 and high pressure turbine 44. The core airflow is compressed by LPC 34 then HPC 42, mixed and burned with fuel in combustor 46, then expanded over the high pressure turbine 44 and low pressure turbine 36.
With continued reference to
Tandem blade stage 24 combines two, typically discrete, blade stages into a single stage. For example, a traditional compressor configuration generally has the last stages in the pattern of stator stage, rotor stage, stator stage, rotor stage, and exit guide vane stage. Embodiments described herein have the pattern of stator stage 28, tandem rotor stage 24, and exit guide vane stage 26 or a tandem stator stage, described below. Tandem rotor stage 24 does more work than a traditional single blade stage, providing additional pressure-ratio and also reducing the need for a traditional stator vane stage that typically separates two traditional single blade stages. By removing one of the stator vane stages, respective shrouded cavities that are typically associated with each vane in the stator vane stage, are no longer needed. Shrouded cavities tend to increase metal temperatures because of the interface between a seal, typically a knife edge seal, and the rotor disk. The increased temperatures at the knife edge seal cause increased overall temperatures as part of windage heat-up. By removing one of the shrouded cavities, the windage heat-up is reduced and temperatures of other engine components in the last stages of the HPC are also reduced.
Those skilled in the art will readily appreciate that by reducing the temperatures, the component life can be improved. For example, by removing the intervening stator vane stage and its knife edge seal, the remaining knife edge seals can be approximately ten to fifteen percent of compressor discharge temperature cooler than they would be if the traditional intervening stator stage and knife edge seal was included. Not only does this potentially increase the life of the remaining seals, it also increases the life of the surrounding engine components due to the reduced windage heat-up temperature. On the other hand, the overall operating temperatures can be increased in order to increase the pressure ratio while still remaining within the traditional temperature tolerances of the engine components. Reducing the need for a traditional stator vane stage by using a tandem blade stage also reduces the length of the compressor since gaps between stages can be removed, and/or tandem rotor blades can overlap each other in the axial direction.
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Now with reference to
With continued reference to
The methods and systems of the present disclosure, as described above and shown in the drawings, provide for gas turbine engines with superior properties including improved control over fluid flow properties through the engine and reduced windage heat up. While the apparatus and methods of the subject disclosure have been shown and described with reference to preferred embodiments, those skilled in the art will readily appreciate that changes and/or modifications may be made thereto without departing from the scope of the subject disclosure.
This application claims the benefit of U.S. Provisional Patent Application Ser. No. 62/064,536 filed Oct. 16, 2014, the entire contents of which are incorporated herein by reference thereto.
| Number | Date | Country | |
|---|---|---|---|
| 62064536 | Oct 2014 | US |
