Patent Grant
7520727
The invention relates generally to a blade airfoil for a gas turbine engine and, more particularly, to an airfoil profile suited for use in the second stage rotor assembly of a two-stage high pressure turbine.
Where a blade airfoil is part of a two-stage turbine driving a compressor (i.e. part of a high pressure or HP turbine), the requirements for such a blade airfoil design are significantly more stringent than multiple stage airfoil designs, as the compressor relies solely on this two-stage HP turbine to deliver all the required work, as opposed to work being spread over several turbine stages. Over and above this, the 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 HP turbine is also subject to harsh temperatures and pressures, which require a solid balance between aerodynamic and structural optimization.
It is therefore an object of this invention to provide an improved blade airfoil for a two-stage high pressure turbine.
In one aspect, the present invention provides a turbine blade for a gas turbine engine comprising an airfoil having an intermediate portion defined by a nominal profile substantially in accordance with Cartesian coordinate values of X, Y, and Z of Sections 2 to 12 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 invention provides a turbine blade for a gas turbine engine comprising an airfoil having an intermediate portion at least partly defined by a nominal profile substantially in accordance with Cartesian coordinate values of X, Y, and Z of Sections 2 to 12 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 in the engine, the Z values are radial distances measured along the stacking line of the airfoil, the X and Y are coordinate values defining the profile at each distance Z, and wherein the X and Y values are scalable as a function of the same constant or number.
In another aspect, the present invention provides a turbine rotor for a gas turbine engine comprising a plurality of blades extending from a rotor disc, each blade including an airfoil having an intermediate portion defined by a nominal profile substantially in accordance with Cartesian coordinate values of X, Y, and Z of Sections 2 to 12 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 blades, 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 accordance with a still further general aspect of the present invention, there is provided a high pressure blade located at a stacking line in a gaspath, comprising an airfoil having a surface lying substantially on the points of Table 2, the airfoil extending between a platform and a tip, the platform being generally defined by the inner gaspath of Table 1, and wherein the tip is defined to provide a blade tip clearance with the outer gaspath defined in Table 1 at said stacking line.
The profile shape of the present invention provides maximum work for a two-stage high pressure turbine, while minimizing flow separation disadvantages in such an environment. It is also necessary to give consideration to the downstream component (in this case, the LP turbine), to ensure that it can accept the flow conditions as they leave the HP turbine, without any adverse effect on LPT performance. The exit conditions of this HPT must be optimized such that the flow can negotiate the flow path in the infer-turbine duct, and enter the LPT fully attached. To accomplish this, advanced 3D optimization techniques are used to ensure that the radial distribution of flow leaving the HPT lends itself to being able to negotiate the inter turbine duct shape without any flow separation. The tip airfoil section is optimized to minimize airfoil pulse while also minimizing any detrimental aerodynamic effects. The present invention also minimizes static pressure gradients in the spanwise and pitch-wise directions, to minimize secondary flow losses.
Further details of these and other aspects of the present invention 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 gas turbine engine 10 further includes a turbine exhaust duct 20 which is exemplified as including an annular core portion 22 and an annular outer portion 24 and a plurality of struts 26 circumferentially spaced apart, and radially extending between the inner and outer portions 22, 24.
The turbine section 18 has a first high pressure turbine (HPT) stage located downstream of the combustor 16 and a second HPT stage located further downstream in the gaspath 27. Referring to
More specifically, the first and second stage rotor assemblies 36, 38 each include a disc drivingly mounted to shaft 39 (see
The novel airfoil shape of each second stage blade 42b is 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 single-stage LPT design. 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 HPT blade stacking line 50 of each respective blade 42b in a generally radial direction and intersects the X axis at the center of rotation of the rotor assembly 36. The positive Z direction is radially outwardly toward the blade tip 62. The Y axis extends tangentially with the positive Y direction being in the direction of rotation of the rotor assembly 36. 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, the set of points which define the HPT stage blade airfoil profile relative to the axis of rotation of the turbine engine 10 and the stacking line 50 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 70 at various locations along its height, the locations represented by Z values. It should be understood that the Z values do not represent an actual radial height along the airfoil 56 but are defined with respect to the engine center line. For example, if the blades 42b are mounted about the rotor assembly 36 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 70 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 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 70 to one another to form the airfoil profile.
The coordinate values listed in Table 2 below represent the desired airfoil profiles in a “cold” (i.e. non-operating) condition. However, the manufactured airfoil surface profile will be slightly different as a result of manufacturing and applied coating tolerances. The coordinate values listed in Table 2 below are for an uncoated airfoil. According to an embodiment of the present invention, the finished HPT blades are coated for thermal protection.
The Table 2 values are generated and shown to three decimal places for determining the profile of the second HPT stage blade airfoil. However, as mentioned above, there are manufacturing tolerance issues, as well as coating thicknesses, which must be accounted for and, accordingly, the values for the profile given in Table 2 are for a theoretical airfoil, to which a ±0.003 inch manufacturing tolerance is additive to the X and Y values given in Table 2 below. A coating having a thickness of 0.001 inch to 0.002 inch is typically applied to the uncoated blade airfoil defined in Table 2. The HPT stage blade airfoil design functions well within these ranges. 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 blade airfoil profile.
It should be understood that the finished HPT blade 42b does not necessarily include all the sections defined in Table 2. The tip 62 and the airfoil portion proximal the platform 64 may not be defined by a profile section 70. For example, in a particular embodiment in which the tip 62 is angled, multiple tip 62 cross-sections would not be defined by a profile section 70. Notably, it should be considered that the airfoil profile proximal to the platform 64 may vary due to several imposed constraints. However, the HPT blade 42b has an intermediate airfoil portion 68 defined between the platform 64 and the tip 62 thereof and which has a profile defined on the basis of at least the intermediate sections of the various blade profile sections 70 defined in Table 2.
It should be appreciated that the intermediate airfoil portion 68 of the HPT stage blade 42b is defined between the inner and outer gaspath walls 28 and 30, and that the wall 28 is partially defined by the blade platform. More specifically, the Z values defining the gaspath in the region of the stacking line 46 fall within the range of Z=5.763 and Z=7.201 which are the z values defining the inner and outer gaspath walls 28 and 30 at the stacking line 50 (see Table 1). Therefore, the physical airfoil profile of HPT blade 42b includes Sections 2 to 11 of Table 2. Sections 1 and 13 are located completely outside of the boundaries set by the inner and annular outer gaspath walls 28 and 30 at the HPT blade stacking line, but are provided, in part, to fully define the airfoil surface and, 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 wall 28 (i.e. blade platform) and the airfoil portion 54 of the blade 42b, and that a suitable blade tip clearance is to be provided between tip 62 and outer wall 30.
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. For example, the airfoil and/or gaspath definitions of Tables 1 and 2 may be scaled geometrically, while maintaining the same proportional relationship and airfoil shape, for application to gas turbine engine of other sizes. Still other 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.
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