Airfoil for a third stage nozzle guide vane

The present invention provides an airfoil for a third stage nozzle guide vane having an external surface with first and second sides. The external surface extends spanwise between a hub and a tip and streamwise between a leading edge and a trailing edge of the airfoil. The external surface includes a contour substantially defined by Table 1 as listed in the specification.

FIELD OF THE INVENTION

The present invention relates to improved airfoil geometry, and more particularly to a high efficiency airfoil for a turbine nozzle guide vane in a gas turbine engine.

BACKGROUND

Gas turbine engine designers continuously work to improve engine efficiency, to reduce operating costs of the engine, and to reduce specific exhaust gas emissions such as NOx, CO2, CO, unburnt hydrocarbons, and particulate matter. The specific fuel consumption (SFC) of an engine is inversely proportional to the overall thermal efficiency of the engine, thus, as the SFC decreases the fuel efficiency of the engine increases. Furthermore, specific exhaust gas emissions typically decrease as the engine becomes more efficient. The thermal efficiency of the engine is a function of component efficiencies, cycle pressure ratio and turbine inlet temperature. The present invention contemplates increased thermal efficiency for a gas turbine engine by improving turbine efficiency through a new aerodynamic design of an airfoil for a third stage turbine nozzle guide vane.

SUMMARY

The present invention provides an airfoil with an external surface having first and second sides, the external surface extends spanwise between a hub and a tip and streamwise between a leading edge and a trailing edge. The external surface has a contour substantially defined by Table 1 as listed in the specification.

In another aspect of the present invention, a turbine nozzle guide vane assembly for a gas turbine engine can include an inner shroud having an upper surface and a lower surface, the upper surface of the inner shroud partially defining an inner flow path wall. An airfoil can extend radially outward from the upper surface of the inner shroud relative to an axis of rotation of the gas turbine engine. The airfoil includes first and second three-dimensional external surfaces that extend between a hub and a tip in a spanwise direction and between a leading edge and a trailing edge in a streamwise direction. A Cartesian coordinate array having X, Y and Z axis coordinates listed in Table 1 of the specification defines the first and second external surfaces of the airfoil.

Another aspect of the present invention provides for a method of forming an airfoil for a turbine nozzle guide vane. The method includes forming a contoured three-dimensional external surface of an airfoil defined by Cartesian (X, Y and Z) coordinates listed in the specification as Table 1, wherein the Z axis coordinates are generally measured in a radial direction from a longitudinal axis, the X axis coordinates are generally measured normal to the Z axis in a streamwise direction, and the Y axis coordinates are generally measured normal to the Z axis and normal to the X axis.

Another aspect of the present invention provides for a method of forming an airfoil for a turbine nozzle guide vane. The method includes forming a contoured three-dimensional external surface of an airfoil defined by Cartesian (X, Y and Z) coordinates listed in the specification as Table 1, wherein the Z axis coordinates are generally measured from an engine centerline axis, the X axis coordinates are generally measured normal to the Z axis in a streamwise direction, and the Y axis coordinates are generally measured normal to the Z axis and normal to the X axis.

DETAILED DESCRIPTION

Referring toFIG. 1, a schematic view of a gas turbine engine10is depicted. While the gas turbine engine10is illustrated with one spool (i.e. one shaft connecting a turbine and a compressor), it should be understood that the present invention is not limited to any particular engine design or configuration and as such may be used in multi spool engines of the aero or power generation type. The gas turbine engine10will be described generally, however significant details regarding general gas turbine engines will not be presented herein as it is believed that the theory of operation and general parameters of gas turbine engines are well known to those of ordinary skill in the art.

The gas turbine engine10includes an inlet section12, a compressor section14, a combustor section16, a turbine section18, and an exhaust section20. In operation, air is drawn in through the inlet12and compressed to a high pressure relative to ambient pressure in the compressor section14. The air is mixed with fuel in the combustor section16wherein the fuel/air mixture burns and produces a high temperature and pressure working fluid from which the turbine section18extracts power. The turbine section18is mechanically coupled to the compressor section14via a shaft22. The shaft22rotates about a centerline axis24that extends axially along the longitudinal axis of the engine10, such that as the turbine section18rotates due to the forces generated by the high pressure working fluid, the compressor section14is rotatingly driven by the turbine section18to produce compressed air. A portion of the power extracted from the turbine section18can be utilized to drive a secondary device26, which in one embodiment can be an electrical generator. The electrical generator can be run at a substantially constant speed that is appropriate for a desired power grid frequency; a non-limiting example being 50 or 60 Hz. Alternatively the secondary device26can be in the form of a compressor or pump for use in fluid pipelines such as oil or natural gas lines.

Referring now toFIG. 2, a partial cross section of the turbine section18is shown therein. As the working fluid exits the combustor section16, the working fluid is constrained between an inner flow path wall31and an outer flow path wall33as it flows through the turbine section18. The turbine section18includes a turbine inlet or first stage nozzle guide vane (NGV) assembly30. The first stage NGV assembly30includes a plurality of static vanes or airfoils32positioned circumferentially around a flow path annulus of the engine10. The first stage NGV assembly30is operable for accelerating and turning the flow of working fluid to a desired direction, as the working fluid exits the combustor section16and enters the turbine section18.

Each airfoil32of the first stage NGV assembly30extends between a leading edge34and a trailing edge36in the stream wise direction and between an inner shroud38and an outer shroud40in the spanwise direction. It should be understood that the terms leading edge and trailing edge are defined relative to the general flow path of the working fluid, such that the working fluid first passes the leading edge and subsequently passes the trailing edge of a particular airfoil. The inner and outer shrouds38,40form a portion of the inner and outer flow path walls31,33respectively at that location in the engine10.

The turbine section18further includes a first stage turbine assembly42positioned downstream of the first stage NGV assembly30. The first stage turbine assembly42includes a first turbine wheel44which is comprised of a first turbine disk46having a plurality of first stage turbine blades48coupled thereto. It should be noted here that in one preferred embodiment the turbine blades48and the disk46can be separate components, but that the present invention contemplates other forms such as a turbine wheel having the blades and disk integrally formed together. This type of component is commonly called a “BLISK,” short for a “Bladed Disk,” by those working in the gas turbine engine industry.

Each turbine blade48includes an airfoil50that rotates with the turbine disk46. Each airfoil50extends between a leading edge52and a trailing edge54in the stream wise direction and between an inner shroud or platform56and an outer shroud58in the spanwise direction. The disk46may include one or more seals60extending forward or aft in the streamwise direction. The seals60, sometimes called rotating knife seals, limit the leakage of working fluid from the desired flowpath. The first stage turbine assembly42is operable for extracting energy from the working fluid via the airfoils50which in turn cause the turbine wheel44to rotate and drive the shaft22.

Directly downstream of the first stage turbine assembly42is a second stage nozzle guide vane (NGV) assembly70. The second stage NGV assembly70includes a plurality of static vanes or airfoils72positioned circumferentially around the flow path of the engine10. The airfoils72of the second stage NGV assembly70are operable for accelerating and turning the working fluid flow to a desired direction as the working fluid exits the second stage NGV assembly70. Each airfoil72extends between a leading edge74and a trailing edge76in the stream wise direction and between an inner shroud78and an outer shroud80in the spanwise direction. The inner and outer shrouds78,80form a portion of the inner and outer flow path walls31,33respectively at that location in the engine10.

A second stage turbine assembly82is positioned downstream of the second stage NGV assembly70. The second stage turbine assembly82includes a second turbine wheel84which is comprised of a second turbine disk86having a plurality of second stage turbine blades88coupled thereto. Each turbine blade88includes an airfoil90that rotates with the turbine disk86when the engine10is running. Each airfoil90extends between a leading edge92and a trailing edge94in the stream wise direction and between an inner shroud or platform96and an outer shroud98in the spanwise direction. The disk86may include one or more seals100extending forward or aft in the streamwise direction. In this particular embodiment of the invention, the second stage turbine assembly82is connected to the first stage turbine assembly42and therefore increases the power delivered to the shaft22.

A third stage nozzle guide vane (NGV) assembly110is located downstream of the second stage turbine assembly82. The third stage NGV assembly110includes a plurality of static vanes or airfoils112positioned circumferentially around the flowpath of the engine10. The airfoils112of the third stage NGV assembly110are operable for accelerating and turning the working fluid flow to a desired direction as the working fluid exits the third stage NGV assembly110. Each airfoil112extends between a leading edge114and a trailing edge116in the streamwise direction and between an inner shroud118and an outer shroud120in the spanwise direction. The inner and outer shrouds118,120form a portion of the inner and outer flow path walls31,33respectively at that location in the engine10. The third stage nozzle guide vane (NGV) assembly110will be described in more detail below.

A third stage turbine assembly130is positioned downstream of the third stage NGV110. The third stage turbine assembly130includes a third turbine wheel132which is comprised of a third turbine disk134having a plurality of third stage turbine blades136coupled thereto. Each turbine blade136includes an airfoil138that rotates with the turbine disk134when the engine10is running. Each airfoil138extends between a leading edge140and a trailing edge142in the stream wise direction and between an inner shroud or platform144and an outer shroud146in the spanwise direction. The third disk134may also include one or more seals148extending forward or aft of the disk134in the streamwise direction. Similar to the second stage turbine assembly82, the third stage turbine assembly130is also connected to the first stage turbine assembly42and therefore further increases the power delivered to the shaft22.

Although not shown in each of the drawings it should be understood the airfoils for both the turbine blades and turbine nozzle guide vanes may include internal cooling flow passages and apertures extending through portions of the external surfaces of the airfoil. Pressurized cooling fluid can then flow from the internal passages through the apertures to cool the external surface of the airfoils as would be known to those skilled in the art. In this manner, the engine10may be run at the higher turbine inlet temperatures, and thus produce higher thermal efficiencies while still providing adequate component life as measured by such parameters as high cycle fatigue limits, low cycle fatigue limits, and creep, etc.

It should be further noted that the airfoils may include coatings to increase component life. The coatings can be of the thermal barrier type and/or the radiation barrier type. Thermal barrier coatings have relatively low heat transfer coefficients which help to reduce the heat load that the cooling fluid is required to dissipate. Thermal barrier coatings are typically ceramic based and can include mullite and zirconia based composites, although other types of coatings are contemplated herein. Radiation barrier coatings operate to reduce radiation heat transfer to the coated component by having highly reflective external surfaces such that radiation emanating from the high temperature exhaust gas is at least partially reflected away and not absorbed by the component. Radiation barrier coatings can include materials from high temperature chromium based alloys as is known to those skilled in the art. The radiation barrier coatings and thermal barrier coatings can be used to coat the entire airfoil, but alternate embodiments include a partial coating and/or a coating with intermittent discontinuities formed therein.

Referring now toFIGS. 3 through 9, the third stage nozzle guide vane (NGV) assembly110will be described in more detail. As partially described previously, the nozzle guide vane assembly110includes an inner shroud118wherein an outer surface350of the shroud defines a portion of the inner flow path wall31at that particular location in the engine10. A plurality of airfoils112extend radially outward from the outer surface350of the shroud118from a hub352toward a tip354. Each airfoil112is attached to the shroud118proximate the hub352of the airfoil112. The airfoils112can be integrally formed with the shroud118through a casting process or the like or alternatively may be mechanically joined via welding, brazing or by any other joining method known to those skilled in the art. An outer shroud120can be attached to the airfoil112proximate the tip354of the airfoil112. The outer shroud120includes an inner surface156which forms a portion of the outer flow path33in the turbine section18. The third stage nozzle guide vane assembly110further includes an outer lug360positioned above each airfoil112. The lugs360are operable for reacting torque loads produced by the airfoils112through the frame (not shown) of the aircraft engine10.

Each airfoil112includes an external surface361having first and second sides363,365sometimes called pressure and suction sides respectively. The external surface361of the airfoil112provides aerodynamic control of the flow of working fluid so as to optimize efficiency in the turbine section18. The external surface361of the airfoil112is substantially defined by Table1listed below. Table 1 lists data points in Cartesian coordinates that define the external surface of the airfoil112at discrete locations. The Z axis coordinates are generally measured radially outward from a reference location. In one form the reference location is the engine centerline axis, and in another form the reference location is the inner shroud118from the hub352. The Z axis defines an imaginary stacking axis from which the contoured external surface is formed. The term “stacking axis,” as it is typically used by aerodynamic design engineers, is nominally defined normal to the inner shroud118or in a radial direction from the axis of rotation, but in practice can “lean” or “tilt” in a desired direction to satisfy mechanical design criteria as is known to those skilled in the art. The lean or tilt angle is typically within 10°-25° of the normal plane in any direction relative to the inner shroud118, but can vary with larger angles as necessary to satisfy design criteria. The X axis coordinates are generally measured normal to the stacking axis in a streamwise direction. The Y axis coordinates are generally measured normal to the stacking axis and normal to the X axis. The airfoil112defined by Table 1 improves the third stage turbine nozzle guide vane efficiency by 2.03% over prior art designs.

While the external surface of airfoil112is defined by discrete points, the surface can be “smoothed” between these discrete points by parametric spline fit techniques and the like. One such method called numerical uniform rational B-spline (NURB-S) is employed by software run on Unigraphics® computer aided design workstations. The data splines can be formed in the streamwise direction and or the spanwise direction of the airfoil112. Other surface smoothing techniques known to those skilled in the art are also contemplated by the present invention.

The airfoils of the present invention can be formed from any manufacturing process known to those skilled in the art. One such process is an investment casting method whereby the entire third stage NGV assembly110is integrally cast as a one-piece component. Alternatively the third stage NGV assembly110can be formed in multiple pieces and bonded together. In another form the third stage NGV assembly110can be formed from wrought material and finished machined to a desired specification.

The present invention includes airfoils having an external surface formed within a manufacturing tolerance of +/−0.025 inches with respect to any particular point in Table 1 or spline curve between discrete points. Furthermore, if the airfoil of the present invention has a material coating applied, the tolerance band can be increased to +/−0.050 inches.