Patent Grant
9890790
This application is related by subject matter to the non-provisional patent application Ser. No. 14/087,946 entitled “ADJUSTED STATIONARY AIRFOIL” and assigned to the same assignee.
Not Applicable.
The present invention relates generally to gas turbine engines and more specifically to an airfoil profile having an improved design.
A gas turbine engine typically comprises a multi-stage compressor that takes air, which has been drawn into the engine, and compresses it into a higher pressure and temperature. A majority of this air passes to the combustion system, which mixes the compressed and heated air with fuel and contains the resulting reaction that generates the hot combustion gases. These gases then pass through a multi-stage turbine, which, in turn drives the compressor, and possibly a shaft of an electrical generator. Exhaust from the turbine can also be channeled to provide thrust for propulsion of a vehicle.
Typical compressors and turbines comprise a plurality of alternating rows of rotating and stationary airfoils. The stationary airfoils, or vanes, direct the flow of air in a compressor or hot combustion gases in a turbine onto a subsequent row of rotating airfoils, or blades, at the proper orientation in order to maximize the output of the compressor or turbine. The performance of the gas turbine engine is dependent on the mass of air entering the engine. Generally, the greater the amount of air that enters the engine, the more power that is produced.
The present invention is defined by the claims below. Embodiments of the present invention solve at least the above problems by providing a system and method for, among other things, increasing airflow throughout a plurality of assemblies in a gas turbine engine.
In accordance with the present invention, there is provided a novel and improved airfoil for a compressor component having a redefined airfoil profile. The surface area of the rotor blade is adjusted to allow for increased air flow. The chord length of the rotor blade is increased at the root with the amount of increase tapering towards the tip. By increasing the surface area of the rotor blade, more air may be captured and harnessed by the airfoil, thus increasing the performance of the compressor and the gas turbine engine.
In an embodiment of the present invention, a compressor component having an attachment, a platform, and an airfoil extending radially outward from the platform is disclosed. The airfoil has an uncoated profile substantially in accordance with Cartesian coordinate values of X and Z, for each distance Y in inches as set forth in Table 1, carried to three decimal places.
In another embodiment, an airfoil for a compressor blade is disclosed having an uncoated profile substantially in accordance with Cartesian coordinate values X, Y, and Z as set forth in Table 1, carried to three decimal places, where Y is a distance measured in inches, the X and Z coordinate values being joined in smooth continuing splines to form airfoil sections and the airfoil sections joined smoothly to form the profile.
In another embodiment, a compressor is disclosed in which the compressor comprises a compressor disk having a plurality of compressor blades extending radially outward from the compressor disk. The compressor blades each have an airfoil with an uncoated nominal profile substantially in accordance with Cartesian coordinate values of X, Y and Z, set forth in inches in Table 1, with the Y coordinate values at perpendicular distances from planes normal to a radius from an engine centerline, wherein airfoil sections are defined at each distance Y by connecting the X and Z coordinate values with smooth continuing splines, and the airfoil sections are joined smoothly to form the airfoil profile, where the compressor blades are rotating blades located adjacent to inlet guide vanes of the compressor, the inlet guide vanes being shaped to compliment the profile of the compressor blades.
Although disclosed as an airfoil that is uncoated, it is envisioned that an alternate embodiment of the present invention can include an airfoil that is at least partially coated with an erosion resistant coating, corrosion resistant coating, or a combination thereof. In this case, the coordinates of the airfoil as listed in Table 1 would be prior to a coating being applied to any portion of the airfoil.
Additional advantages and features of the present invention will be set forth in part in a description which follows, and in part will become apparent to those skilled in the art upon examination of the following, or may be learned from practice of the invention. The instant invention will now be described with particular reference to the accompanying drawings.
The present invention is described in detail below with reference to the attached drawing figures, wherein:
The subject matter of the present invention is described with specificity herein to meet statutory requirements. However, the description itself is not intended to limit the scope of this patent. Rather, the inventors have contemplated that the claimed subject matter might also be embodied in other ways, to include different components, combinations of components, steps, or combinations of steps similar to the ones described in this document, in conjunction with other present or future technologies.
Referring initially to
Extending radially outward from the platform 106 is an airfoil 108 having a tip 112, with the tip 112 located at an end of the airfoil 108 opposite of the platform 106. For the compressor blade 100, the airfoil is solid, and fabricated from a material such as a martenestic steel alloy. The airfoil has an uncoated profile substantially in accordance with Cartesian coordinate values of X and Z, for each distance Y, in inches, as set forth in Table 1 below, and carried to three decimal places. The distance Y is measured from the engine centerline, shown in
A plurality of airfoil sections 110 are established by applying smooth continuing splines between the X, Z coordinate values at each Y height. Smoothly joining each of the airfoil section 110 together form the profile of the airfoil 108. The airfoil 108 can be fabricated by a variety of manufacturing techniques such as forging, casting, milling, and electro-chemical machining (ECM). As such, the airfoil has a series of manufacturing tolerance for the position, profile, twist, and chord that can cause the airfoil 108 to vary by as much as approximately +/−0.012 inches from a nominal state.
The compressor blade 100 is generally fabricated from a steel alloy such as 15-5PH, which is a precipitation-hardened, martensitic stainless steel alloy that is used on parts requiring corrosion resistance and high strength at temperatures up to approximately 600 deg. F. While other alloys could be used, it is preferred that a high-temperature steel alloy be selected because of the operating conditions. Although the compressor blade has been discussed as having an attachment, a platform, and an airfoil, it is to be understood that all of these features of the blade are typically fabricated from the same material and are most likely integral with one another.
In addition to manufacturing tolerances affecting the overall size of the airfoil 108, it is also possible to scale the airfoil 108 to a larger or smaller airfoil size. However, in order to maintain the benefits of this airfoil shape and size, in terms of stiffness and stress, as will be discussed further below, it is necessary to scale the airfoil uniformly in X and Z directions, but Y direction may be scaled separately.
As previously discussed, the profile generated by the X, Y, and Z coordinates of Table 1 is an uncoated profile. While an embodiment of the present invention is an uncoated compressor blade 100, it is possible to add a coating to at least a portion of the airfoil 108 in an alternate embodiment. This coating would have a thickness of up to approximately 0.010 inches. Such coatings can be applied to the airfoil to improve resistance to erosion or to increase temperature capability.
Referring to
Depending on the blade configuration, it is possible that a second platform can be positioned proximate the tip 112 of the airfoil 108. A second platform located at the tip 112, is commonly referred to as a shroud and interlocks with a shroud of an adjacent blade. The shrouds provide an outer air path seal that increases efficiency by preventing air from passing over the blade tip 112 and also serves to reduce the vibration of the airfoils 108. The use of a second platform, or a shroud, is common in airfoils having a relatively long radial length.
In an alternate embodiment of the invention, a compressor comprises at least one compressor disk (not depicted) having a plurality of compressor blades 100 that extend radially outward from the compressor disk. As one skilled in the art understands, a compressor typically comprises a plurality of alternating stages of rotating and stationary airfoils that raise the pressure and temperature of a fluid passing through. While the compressor blade 100 having the airfoil 108 can operate in a variety of locations within a compressor, depending on the compressor size, one such location that suits this blade, is adjacent an inlet of the compressor.
For compressor blades in this location, a common durability issue exhibited by prior art blades is erosion of the blade leading edge. The leading edge of the blade (see 114 in
As one skilled in the art understands, as a compressor blade is rotated by a compressor disk, and the weight of the blade pulls radially outward on the disk. However, because of blade design issues such as desired compression of the airflow, blade materials, and compressor size, rarely is the only load a truly radial pulling load. For large unshrouded blades there is usually a substantial amount of blade twist from airfoil root to airfoil tip. Due to the blade's pulling load, the airfoil will tend to untwist or try to straighten itself out. The compression of the airflow also creates load on the airfoil that tries to bend the blade where the airfoil attaches to the platform. Blade pull, untwist, and aero loading result in concentrated steady stress that can occur near the blade's airfoil root leading edge and the blade attachment, as seen with blades of prior art. Airfoil unsteady stress can occur due to the vibratory nature of the blade. Specific vibratory shapes for the blade result in stress concentrations on the airfoil. Blade failure can occur when the blade steady and unsteady stress concentrations occur together. If erosion forms at a location of high steady and unsteady stress then the chance of blade failure is increased.
For a compressor blade, increasing the surface area near the root of the compressor blade may allow for the compressor blade to take in a larger amount of air than the prior art. By taking in a larger amount of air, more air may be compressed and consequently, more power may be produced by the engine. However, due to the fixed geometry of the compressor case, where an airfoil of a compressor blade increases in axial length, there generally must also be a corresponding decrease in the axial length of an adjacent vane, thereby reducing the surface area of the vane. Decreasing the surface area of the stator vane allows for the stator vane to clear the compressor blade.
Referring to
While the invention has been described in what is known as presently the preferred embodiment, it is to be understood that the invention is not to be limited to the disclosed embodiment but, on the contrary, is intended to cover various modifications and equivalent arrangements within the scope of the following claims. The present invention has been described in relation to particular embodiments, which are intended in all respects to be illustrative rather than restrictive.
From the foregoing, it will be seen that this invention is well adapted to attain all the ends and objects set forth above, together with other advantages which are obvious and inherent to the system and method. It will be understood that certain features and sub-combinations are of utility and may be employed without reference to other features and sub-combinations. This is contemplated by and within the scope of the claims.
| Number | Name | Date | Kind |
|---|---|---|---|
| 7566202 | Noshi | Jul 2009 | B2 |
| Number | Date | Country | |
|---|---|---|---|
| 20150147182 A1 | May 2015 | US |
