High lift transonic turbine blade

The present invention relates to turbine blades for turbomachinery such as gas turbine engines and, more particularly, to improved geometrical shapes for such blades and the platform from which they extend.

Turbine blades are mounted circumferentially on an inner diameter platform about a turbine shaft to allow rotation thereof in streams of working medium fluids to extract energy and work therefrom as those fluids flow around and past such blades through the flow passages between them. Such flows can be transonic in accelerating from below the speed of sound coming into the blade region to being above the speed of sound in the exit region. In these situations inefficiencies due to shock wave losses are important.

Blade airfoil surface profiles and the geometry of the airfoil and the platform on which the blade is mounted become very important in achieving high efficiencies. In addition, a blade design to be suitable must minimize the penalties imposed on the engine system by the weight and cost of the blades, and need for cooling air therein, a goal sought through reducing the number of blades provided about the platform circumference for each blade stage used. However, the remaining blades must then extract more work, the lift or loading of the blade, from the fluid streams passing thereby which tends to reduce blade efficiency. A high lift blade will have a Zweifel lift coefficient (the ratio of the actual load to the ideal load) that is greater than 1.1. Thus, there is a desire for a blade that achieves high efficiency during transonic operation while having a large lift loading.

The present invention provides a turbine blade system including a blade airfoil having an airfoil shape with the blade airfoil having a nominal profile substantially in accordance with normalized Cartesian coordinate values of X, Y and Z set forth in Table 1 below and which values are dimensionless values that are convertible to corresponding absolute distance values through manipulating them in accord with corresponding normalization equations. The X and Y absolute distance values, when connected by smooth continuing arcs, define nominal airfoil profile sections at each Z absolute distance value, and these nominal airfoil profile sections, when joined smoothly with adjacent ones thereof, form a complete nominal airfoil shape that is substantially matched by the airfoil shape of the blade airfoil. This blade airfoil can be supported on a ring platform having a support surface with a support surface shape in the vicinity of the location at which that blade airfoil is supported thereon such that the support surface smoothly joins with the airfoil shape of that blade airfoil, the support surface having a nominal profile substantially in accordance with normalized Cartesian coordinate values of X, Y and Z set forth in Table 4 below and which values are dimensionless values that are convertible to corresponding absolute distance values through manipulating them in accord with corresponding normalization equations, and wherein X and Y absolute distance values at various Z absolute distance values, when connected by smooth continuing arcs, form a nominal support surface shape substantially matched by the support surface shape of the support surface.

FIG. 1 provides a diagrammatic side view of a blade mounted on a rotatable platform adjacent a fluid guide vane,

FIG. 2 shows a pictorial perspective view of the turbine blade embodying the present invention,

FIG. 3 shows a side pictorial view of a platform ring stage having blades like that of FIG. 2 mounted thereon as an embodiment of the present invention,

FIG. 4 shows a diagrammatic top view of a portion of FIG. 3 marked with certain variables,

FIG. 5 shows a perspective diagram of the blade of FIG. 2 that is shown with corresponding three-dimensional coordinate axes,

FIG. 6 shows a diagrammatic side view of the blade of FIG. 2 marked with a plurality of sectioning lines arrayed along with a corresponding one of the coordinate dimensional axis from FIG. 5 ,

FIG. 7 shows a typical blade cross section view of the blade of FIG. 2 for a section of that blade designated in FIG. 6 that is shown with two of the corresponding two dimensional coordinates axes from FIG. 5 , and

FIG. 8 shows a plurality of blade cross section views of the blade of FIG. 2 for sections of that blade designated in FIG. 6 as the sections selected ascend from the blade mounting platform.

FIG. 1 shows the relationship of a blade, 10 , mounted on a rotatable platform extending about an unseen turbine shaft adjacent a fluid guide vane, 11 . Above them is shown the edge of the engine wall limiting the outer extent of the fluid flow path past blade 10 and vane 11 , that is, forming the outer diameter, 12 , of this fluid flow passage containing blade 10 and vane 11 . Rotatable blade 10 has a gap, 13 , between the tip of that blade and outer diameter 12 but vane 11 is fixed to the structure providing this fluid flow passageway outer diameter as well as to the structure supporting it from below. That vane supporting structure and the platform on which blade 10 is mounted together limit the inner extent of the fluid flow path past blade 10 and vane 11 , that is, forming the inner diameter, 12 , of this fluid flow passage containing blade 10 and vane 11 . Blade 10 is shown having an outward extent from the platform on which it is mounted to its tip designated as its span, 15 .

Blade 10 and the inner diameter 12 surface of the platform on which blade 10 is mounted of the present invention are shown in the pictorial perspective view of FIG. 2 , and a plurality of blades 10 so mounted are shown in FIG. 3 circumferentially positioned about the rotatable ring platform supporting them on which the inner diameter surface 14 of the fluid passageway is indicated. Sixty such blades are so provided on the ring platform example shown in FIG. 3 . In addition to what is shown in FIGS. 2 and 3 , blades 10 can be positioned through being selectively rotated about a radial axis extending outward from the engine centerline from what is shown by up to about ±10°.

FIG. 4 shows a diagrammatic top view of a portion of the platform ring in FIG. 3 with certain variables marked thereon used in calculating the Zweifel lift coefficient for blades 10 . The Zweifel lift coefficient is defined as the ratio of the actual load and the ideal load, and is given by:

Zweifel=[(1 /g )* {dot over (m)}*ΔCu ]/[0.5 *ρ*W 2 ^2 *bx*h]

where:

g=acceleration due to gravity of 32.2 ft/s 2

{dot over (m)}=mass flow (lbm/s)

ΔCu=Cu 2 −Cu 1 =change in tangential velocity across the blade (ft/s)

ρ=density (lbm-s 2 /ft 4 )

W 2 =exit relative velocity (ft/s)

bx=axial chord (ft)

h=span (ft)

A perspective diagram of the blade of FIG. 2 is shown in FIG. 5 with the three dimensional coordinates, X, Y and Z, of a Cartesian coordinate system shown that will be used in providing a description of the geometry of the blade, and surface of the platform supporting that blade, that is, a description of the shapes involved. The shape of blade 10 of the present invention can be specified using an outwardly stacked, or radially stacked, series of cross section portions of that blade, stacked along its span 15 extending outward from the ring center as indicated in FIG. 6 , with the radial axis extending perpendicular to the engine centerline (and so the platform ring centerline) being designated the Z axis shown in FIG. 5 . The remaining two axes in this coordinate system, the X and the Y axes, are used to specify the shape of a stack section in a cross section plane as indicated in FIG. 7 with the X axis being in the axial direction extending parallel to the engine centerline and with the Y axis being in the tangential direction that is tangential to the ring platform and parallel to a direction of blade rotation at a point. The radially stacked sections so specified are to be smoothly joined together in providing a complete surface for a blade 10 . The inner diameter 14 surface of the flow passageway between the blades is also contoured as specified here and smoothly joined with each corresponding one of blades 10 .

Some definitions of terms used in the description include the leading edge as blade or airfoil point farthest toward the forward direction of the engine and the trailing edge as blade or airfoil point farthest toward the rearward from the forward direction of the engine. The axial chord of a blade or airfoil is the axial distance between the leading and trailing edges thereof. The pitch is the tangential distance between adjacent blades or airfoils mounted on the platform ring. The root, or inner diameter (ID) section of the blade or airfoil is the section thereof closest to the engine centerline, and the tip, or outer diameter (OD) section is the section thereof farthest form the engine centerline. The root radius is the radial distance from the engine centerline to the root section.

In addition, there are some normalization parameters used in normalization equations given below for use in converting normalized blade or airfoil surface coordinates for sections of the blade given in the tables below to absolute blade or airfoil coordinates to selectively scale such blades to the various sizes suited for various sized corresponding turbine engines. The blade root axial chord normalization parameter represented as Bx root has a normalization parameter value of 1.155630 in. The blade span (form the ID section to the OD section) represented as h, as above, has a normalization parameter value of 1.905000 in. The blade root pitch represented as Pitch root has a normalization parameter value of 1.081550 in., as determined through the platform ring radius and the number of blades selected to be mounted thereon, and the blade root radius represented as R root has a normalization parameter value of 10.328000 in.

Table 1, Airfoil Surface (Normalized), tabulates the normalized section cold, uncoated airfoil surface coordinates for the 16 airfoil sections indicated in FIG. 6 that describe the three dimensional shape of blade or airfoil 10 with a surface profile tolerance of ±0.0500 in. including coating and manufacturing variability. Each of these sections is normalized by its own axial chord (Bx). The coordinates are arranged such that they start at the leading edge, traverse clockwise around the airfoil on the convex surface to the trailing edge, and then continue on the concave surface of the section back to the leading edge. The span fraction z/h ranges from the root section of 0.0 to the tip (or outer diameter) section of 1.0. These coordinates can be translated to absolute coordinates using the root axial chord (BX root ) along with the values in Table 2, Axial Chord Distribution, and Table 3, Radial Stacking Distribution, set out below. FIG. 8 shows a plurality of blade cross section views from the Table 1 defined sections 1, 4, 6, 9 and 16.

Table 2, Axial Chord Distribution, tabulates the variation in the radial direction of the axial chord of each of the cold, uncoated blade sections shown in FIG. 6 . The airfoil sections each have a unique axial chord, and this chord is used in normalizing the airfoil section coordinates. For simplicity the local axial chord has been normalized by the root axial chord. So, for example, the tip section has an axial chord that is only 65.9% of the length of the root section (typical of a turbine blade).

Table 3, Radial Stacking Distribution, tabulates the offsets for cold, uncoated blades required to accurately stack the blade airfoil sections of FIG. 6 relative to one another in accord with the sections as dimensioned in the foregoing. For example, the leading edge points are not all lined up along a radial line, there is some axial sweep and tangential lean to the three dimensional blade shape. A radial distribution of axial and tangential offsets is supplied, all relative to the root section. The offsets are normalized by the root axial chord in the axial direction and by the root pitch in the tangential direction.

Table 4, Inner Diameter Flowpath, tabulates the normalized coordinates of the cold, uncoated ring platform surface forming the inner diameter passageway limit described above with a surface profile tolerance of ±0.0500 in. including coating and manufacturing variability. The ring platform surface, or flowpath, is a three dimensional shape that begins upstream of the root section leading edge and concludes downstream of the root section trailing edge. The flowpath varies in radius from the engine centerline with both axial and tangential position, and so the coordinates describe a surface rather than a line. The axial coordinate is normalized by the root axial chord, the tangential coordinate is normalized by the root pitch, and the radial coordinate is normalized by the airfoil span. The origin (0,0,0) of the surface is aligned with the root section leading edge. The surface platform surface dips below the root section radius in some places which leads to negative radial values in the table for those locations.

The normalized coordinates for any of the sections of blade 10 shown in FIG. 6 , and for the ring platform surface 14 , set out in the foregoing tables can be converted to the coordinates for a selected blade size and platform surface in absolute dimension space using the normalization equations given below. The conversion for use in converting normalized blade or airfoil surface coordinates for sections of the blade to absolute blade or airfoil coordinates are provided by the following equations beginning with the axial coordinates equation

x absolute = ( x local Bx local ) ⁢ ( Bx local Bx root ) ⁢ ( Bx root ) + ( x ⁢ Offset local Bx root ) ⁢ ( Bx root )

where

x local /Bx local is the normalized airfoil coordinate from Table 1, Bx local /Bx root is obtained from the Axial Chord Distribution from Table 2,

and

XOffset local /Bx root is obtained from the Radial Stacking Distribution from Table 3.

Similarly, the tangential coordinates can be found using

y absolute = ( y local Bx local ) ⁢ ( Bx local Bx root ) ⁢ ( Bx root ) + ( y ⁢ Offse ⁢ t local Pitch root ) ⁢ ( Pitch root )

where

yOffset local /Pitch root is obtained from the Radial Stacking Distribution from Table 3.

The absolute radius of any section can be found through

R absolute =(SpanFraction)( h )+ R root

where

h is the airfoil span, and

R root is the reference root section radius relative to the engine centerline.

The ring platform surface, or inner diameter flowpath, can also be converted from normalized to absolute coordinates using the following normalization equations

Once determined in absolute space relative to the root section leading edge, the entire airfoil can then be shifted to any other location in space.

The airfoil geometry includes tolerances due to manufacturing, surface finish and coating variability of about ±0.0500 in. In addition, the airfoil as described can be rotated about its radial axis ±20° depending on the particular turbine application.

Although the present invention has been described with reference to preferred embodiments, workers skilled in the art will recognize that changes may be made in form and detail without departing from the spirit and scope of the invention.