Patent Application
20150361826
This disclosure relates to an array of airfoils within a stage of a gas turbine engine, for example, a turbine stator vane array.
A gas turbine engine typically includes a fan section, a compressor section, a combustor section and a turbine section. Air entering the compressor section is compressed and delivered into the combustor section where it is mixed with fuel and ignited to generate a high-speed exhaust gas flow. The high-speed exhaust gas flow expands through the turbine section to drive the compressor and the fan section. The compressor section typically includes low and high pressure compressors, and the turbine section includes low and high pressure turbines.
The compressor and turbine section includes circumferential arrangements of fixed and rotating stages. Structural vibratory coupling between adjacent airfoils can occur during engine operation. For rotating stages of the engine, blade mistuning has been used in which there are two sets of blades are arranged in circumferentially alternating relationship to provide an even numbered blade array. One set of blades has a different characteristic than the other set of blades to provide two different resonant frequencies. For fixed stages, vanes have been mistuned by providing different sets of vanes in adjacent quadrants of the array.
A significant driver for turbine blade vibration is the pressure distortion produced by the interaction of the array of blades and the axially adjacent array of turbine vanes. As the blades pass the vanes, a pressure fluctuation occurs at the frequency at which the blades pass the vanes. Depending upon the distribution of pressure on the blades, certain blade vibratory modes may be excited if the blade/vane passing frequency matches the normal mode of the blade.
In one exemplary embodiment, a gas turbine engine includes an array of airfoils. Each airfoil includes a first circumferentially extreme position. The first circumferentially extreme positions of the airfoils are circumferentially spaced apart from one another a pitch. Each airfoil includes a second circumferentially extreme position circumferentially spaced from the first circumferentially extreme position in an angular spacing that is at least one half the pitch.
In a further embodiment of the above, the first circumferentially extreme position is one radial end of the airfoil.
In a further embodiment of any of the above, the second circumferentially extreme position is another radial end of the airfoil opposite the one radial end.
In a further embodiment of any of the above, the airfoil includes another radial end of the airfoil opposite the one radial end. The second circumferentially extreme position is located radially between the one radial end and the other radial end.
In a further embodiment of any of the above, the angular spacing is approximately the pitch.
In a further embodiment of any of the above, the angular spacing is greater than the pitch.
In a further embodiment of any of the above, the gas turbine engine includes a rotating stage and a fixed stage of stator vanes. The array of airfoils provides the fixed stage of stator vanes.
In a further embodiment of any of the above, the stator vanes include inner and outer platforms. The airfoil extends radially between and joins the inner and outer platforms.
In a further embodiment of any of the above, the first and second circumferentially extreme positions are provided by the airfoil at the inner and outer platforms.
In a further embodiment of any of the above, there is a turbine section in which the array is arranged in the turbine section.
In a further embodiment of any of the above, the array of airfoils is configured to produce a radial alternating pressure amplitude and phase distribution on the blades in the rotating stage that reduces the response of potentially damaging vibratory modes for the blade.
In another exemplary embodiment, an array of airfoils includes airfoils. The first circumferentially extreme positions of the airfoils are circumferentially spaced apart from one another a pitch. Each airfoil includes a second circumferentially extreme position circumferentially spaced from the first circumferentially extreme position an angular spacing that is at least one half the pitch.
In a further embodiment of the above, the first circumferentially extreme position is one radial end of the airfoil.
In a further embodiment of any of the above, the second circumferentially extreme position is another radial end of the airfoil opposite the one radial end.
In a further embodiment of any of the above, the airfoil includes another radial end of the airfoil opposite the one radial end. The second circumferentially extreme position is located radially between the one radial end and the other radial end.
In a further embodiment of any of the above, the angular spacing is approximately the pitch.
In a further embodiment of any of the above, the angular spacing is greater than the pitch.
In a further embodiment of any of the above, the array is a fixed stage of stator vanes.
In a further embodiment of any of the above, the stator vanes include inner and outer platforms. The airfoil extends radially between and joining the inner and outer platforms.
In a further embodiment of any of the above, the first and second circumferentially extreme positions are provided by the airfoil at the inner and outer platforms.
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 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.
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 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 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.
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.
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 (‘TSFC’)”—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)]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).
Referring to
The section includes a fixed stage 60 that provides a circumferential array 61 (
The array 61 of airfoils 70 in the fixed stage 60 is schematically shown in
The disclosed array of airfoil uses the geometry of the vanes to tailor the pressure distribution that the downstream array of blades will encounter. The radial distribution of the phase of maximum alternating pressure that the blade encounters each rotation is varied. The pressure-phase distribution is tailored to not excite a particular blade mode shape and make the “modal efficiency” of the excitation low. The result is that potentially damaging vibratory modes of the blade are not efficiently excited. This is achieved by curving the airfoil such that at least a portion of the airfoil in the circumferential direction has an amplitude or angular spacing greater than at least half of the pitch, and in another example, approximately the pitch. In another example, the angular spacing between the first and second circumferentially extreme positions is greater than one pitch.
An example vane cluster 76 is shown in
The second circumferentially extreme position corresponds to a location radially between the location at which the airfoil connects to the inner and outer platform. In the example, the second circumferentially extreme position is at approximately an intermediate location.
In this example, the first circumferentially extreme position corresponds to a location of the airfoil at the outer platform. As shown in
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 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.
This application claims priority to U.S. Provisional Application No. 62/001,915, which was filed on May 22, 2014 and is incorporated herein by reference.
This invention was made with government support under Contract No. FA8650-09-D-2923-0021, awarded by the Air Force. The Government has certain rights in this invention.
| Number | Date | Country | |
|---|---|---|---|
| 62001915 | May 2014 | US |
