Patent Application
20190071975
The application relates generally to a blade airfoil and, more particularly, to an airfoil profile suited for use in a power turbine stage of a gas turbine engine.
Every stage of a gas turbine engine must meet a plurality of design criteria to assure the best possible overall engine efficiency. The design goals dictate specific thermal and mechanical requirements that must be met pertaining to heat loading, parts life and manufacturing, use of combustion gases, throat area, vectoring, the interaction between stages to name a few. The design criteria for each stage is constantly being re-evaluated and improved upon. Each airfoil is subject to flow regimes which lend themselves easily to flow separation, which tend to limit the amount of work transferred to the compressor, and hence the total thrust or power capability of the engine. The blades of a power turbine are also subject to harsh temperatures and pressures, which require a solid balance between aerodynamic and structural optimization. Therefore, improvements in airfoil design are sought.
In one aspect, the present application provides a turbine blade for a gas turbine engine having a gaspath, the blade comprising an airfoil having an intermediate portion contained within the gaspath and defined by a nominal profile substantially in accordance with Cartesian coordinate values of X, Y, and Z of Sections 2 to 11 set forth in Table 2, wherein the point of origin of the orthogonally related axes X, Y and Z is located at an intersection of a centerline of the gas turbine engine and a stacking line of the turbine blade, the Z values are radial distances measured along the stacking line, the X and Y are coordinate values defining the profile at each distance Z.
In another aspect, the present application provides a turbine blade for a gas turbine engine having a gaspath, the turbine blade having a cold uncoated intermediate airfoil portion contained within the gaspath and defined by a nominal profile substantially in accordance with Cartesian coordinate values of X, Y, and Z of Sections 2 to 11 set forth in Table 2, wherein the point of origin of the orthogonally related axes X, Y and Z is located at an intersection of a centerline of the gas turbine engine and a stacking line of the turbine blade, the Z values are radial distances measured along the stacking line, the X and Y are coordinate values defining the profile at each distance Z.
In another aspect, the present application provides a turbine rotor assembly for a gas turbine engine having a gaspath, the assembly comprising a plurality of blades, each blade including an airfoil having an intermediate portion contained with the gaspath of the engine and defined by an un-coated nominal profile substantially in accordance with Cartesian coordinate values of X, Y, and Z of Sections 2 to 11 set forth in Table 2, wherein the point of origin of the orthogonally related axes X, Y and Z is located at an intersection of a centerline of the gas turbine engine and a stacking line of the turbine blade, the Z values are radial distances measured along the stacking line, the X and Y are coordinate values defining the profile at each distance Z.
In a still further aspect of the present application, there is provided a power turbine blade comprising at least one airfoil having a surface lying substantially on the points of Table 2, the airfoil extending from a platform defined generally by some of the ID gaspath coordinates given in Table 1, wherein a fillet radius is applied around the airfoil between the airfoil and the platform.
Further details of these and other aspects of the present application will be apparent from the detailed description and figures included below.
Reference is now made to the accompanying figures depicting aspects of the present invention, in which:
The power turbine section 18a has two stages located in the gaspath downstream of the combustor 16. Referring to
More specifically, the rotor assemblies 36, 38 each include a plurality of circumferentially distributed blade 42a and 42b respectively which extend radially across the hot gaspath 27.
The novel airfoil shape of each second stage power turbine blade 42b is defined by a set of X-Y-Z points in space. This set of points represents a novel and unique solution to the target design criteria discussed above, and are well-adapted for use in a two-stage power turbine design. This blade design provides the following features: tip vortex control; reduced airfoil count for high lift design; and non-axisymmetric endwall contouring. The set of points are defined in a Cartesian coordinate system which has mutually orthogonal X, Y and Z axes. The X axis extends axially along the turbine rotor centerline 29, i.e., the rotary axis. The positive X direction is axially towards the aft of the turbine engine 10. The Z axis extends along the blade stacking line 50 of each respective blade 42a in a generally radial direction and intersects the X axis. The positive Z direction is radially outwardly toward the outer shroud 62 of the blade. The Y axis extends tangentially with the positive Y direction being in the direction of rotation of the rotor assembly 38. Therefore, the origin of the X, Y and Z axes is defined at the point of intersection of all three orthogonally-related axes: that is the point (0,0,0) at the intersection of the center of rotation of the turbine engine 10 and the stacking line 50.
In a particular embodiment of the second stage power turbine blade, the set of points which define the blade airfoil profile relative to the axis of rotation of the turbine engine 10 and stacking line 46 thereof are set out in Table 2 below as X, Y and Z Cartesian coordinate values. Particularly, the blade airfoil profile is defined by profile sections 66 at various locations along its height, the locations represented by Z values. For example, if the blades 42b are mounted about the rotor assembly 38 at an angle with respect to the radial direction, then the Z values are not a true representation of the height of the airfoils of the blades 42b. Furthermore, it is to be appreciated that, with respect to Table 2, Z values are not actually radial heights, per se, from the centerline but rather a height from a plane through the centerline—i.e. the sections in Table 2 are planar. The coordinate values are set forth in inches in Table 2 although other units of dimensions may be used when the values are appropriately converted.
Thus, at each Z distance, the X and Y coordinate values of the desired profile section 66 are defined at selected locations in a Z direction normal to the X, Y plane. The X and Y coordinates are given in distance dimensions, e.g., units of inches, and are joined smoothly, using appropriate curve-fitting techniques, at each Z location to form a smooth continuous airfoil cross-section. The blade airfoil profiles of the various surface locations between the distances Z are determined by smoothly connecting the adjacent profile sections 66 to one another to form the airfoil profile.
The coordinate values listed in Table 2 below represent the desired airfoil profiles in a “cold” non-operating un-coated condition (and at nominal restagger). However, the manufactured airfoil surface profile will be slightly different, as a result of manufacturing and applied coating tolerances. According to an embodiment of the present invention, the finished blade is coated with a thermal protecting layer.
The Table 2 values are generated for determining the profile of the second stage power turbine blade airfoil. However, as mentioned above, there are manufacturing tolerance issues to be addressed and, accordingly, the values for the profile given in Table 2 are for a theoretical airfoil. A profile tolerance of ±0.018 inches, measured perpendicularly to the airfoil surface is additive to the nominal values given in Table 2 below. The blade airfoil design functions well within these ranges of variation. The cold or room temperature profile is given by the X, Y and Z coordinates for manufacturing purposes. It is understood that the airfoil may deform, within acceptable limits, once entering service.
The coordinate values given in Table 2 below provide the preferred nominal second stage power turbine blade airfoil profile.
It should be understood that the finished second stage power turbine blade 42b does not necessarily include all the sections defined in Table 2. The portion of the airfoil 54 proximal to the platform 60 and the tip may not be defined by a profile section 66. It should be considered that the blade 42b airfoil profile proximal to the platform 60 may vary due to several imposed constraints. However, the blade 42b has an intermediate airfoil portion 64 defined between platform 60 and tip thereof and which has a profile defined on the basis of at least the intermediate sections of the various blade profile sections 66 defined in Table 2.
It should be appreciated that the intermediate airfoil portion 64 of the blade 42b is defined between the inner and outer gaspath walls 28 and 30 and that the platform 60 forms part of the inner gaspath wall 28. The airfoil profile physically appearing on blade 42b and fully contained in the gaspath includes Sections 2 to 11 of Table 2. The remaining sections are at least partly located outside of the gaspath 27, but are provided, in part, to fully define the airfoil surface and/or, in part, to improve curve-fitting of the airfoil at its radially distal portions. The skilled reader will appreciate that a suitable fillet radius is to be applied between the inner platform 60 and the airfoil portion of the blade as well as between the outer shroud 62 and the airfoil. The blade inner diameter and outer diameter endwall fillet is in the range of about 0.040″ to about 0.175″. The local ID/OD endwall profile tolerance is +/−0.010″.
The above description is meant to be exemplary only, and one skilled in the art will recognize that changes may be made to the embodiments described without department from the scope of the invention disclosed. All modifications which fall within the scope of the present invention will be apparent to those skilled in the art, in light of a review of this disclosure, and such modifications are intended to fall within the appended claims.
