Pressure recovery axial-compressor blading
In accordance with some embodiments of the present disclosure, a pressure recovery axial compressor blade is provided. The blade may comprise a high pressure surface and a low pressure surface connected at a leading edge and a trailing edge of the blade. Both the high and low pressure surfaces extend span wise from a first end to a second end. At least one of the high and low pressure surfaces has a finite discontinuity in curvature at an intermediate position along the chord of the blade.
1. A blade comprising:
a high pressure surface, and a low pressure surface connected at a leading edge and trailing edge of the blade;
wherein the high pressure and low pressure surfaces extend span wise from a first end to a second end; the low pressure surface having a discontinuity in a chord-wise curvature of an intermediate portion of the low pressure surface between the blade leading edge and blade trailing edge;
wherein the high pressure and low pressure surfaces form an uninterrupted surface extending from the first end to the second end and from the leading edge to the trailing edge wherein the blade thickness increases monotonically from the leading edge to a point of maximum blade thickness and the discontinuity in curvature occurs at, or immediately downstream of the point of maximum blade thickness, and
wherein the discontinuity in curvature causes a reduction in Mach number and increasing the static pressure at an edge of a boundary layer formed on the low pressure surface, the increase in static pressure approximating a shock.
2. The blade according to claim 1 , wherein the discontinuity in curvature is aft of the mid chord.
3. The blade according to claim 2 , wherein the discontinuity in curvature is proximate to the ⅔ chord.
4. The blade according to claim 2 , wherein the discontinuity in curvature is proximate to the ¾ chord.
5. The blade according to claim 1 , wherein the first end comprises an end wall substantially perpendicular to the span of the blade.
6. The blade according to claim 5 , wherein the second end comprises an end wall substantially perpendicular to the span of the blade.
7. The blade according to claim 1 , further comprising a solid interior between the high pressure and low pressure surfaces along the entire span.
8. The blade according to claim 1 , wherein the discontinuity in curvature is caused by a discontinuity in one or more of the first and second derivatives of the low pressure surface profile.
9. The blade according to claim 1 , wherein the blade thickness decreases monotonically from the point of maximum blade thickness to the trailing edge.
10. The blade according to claim 1 , wherein the blade is a compressor rotor blade.
11. The blade according to claim 1 , wherein the blade is a compressor stator blade.
12. The blade according to claim 1 , further comprising a second discontinuity in curvature in the chord-wise curvature of the high pressure surface.
13. The blade according to claim 1 , wherein the low pressure surface turning towards the high pressure surface at the discontinuity in curvature when moving downstream from the discontinuity.
14. A turbine compressor blade comprising:
a solid interior defined by a high pressure surface, a low pressure surface, a leading edge, a trailing edge, a first end, and a second end, wherein the high and low pressure surfaces are connected to each other at the leading and trailing edges and both the high and low pressure surfaces extend from the first end to the second end,
wherein at least one of the high pressure surface and the low pressure surface has a discontinuity in curvature at an intermediate position along a chord of the blade;
wherein the solid interior extends between the high pressure and low pressure surfaces along the entire span from the first end to the second end and from the leading edge to the trailing edge;
wherein the blade thickness increases monotonically from the leading edge to a point of maximum blade thickness and the discontinuity in curvature occurs at, or immediately downstream of the point of maximum blade thickness and
wherein the discontinuity in curvature causes a reduction in Mach number and increasing the static pressure at an edge of a boundary layer formed on the low pressure surface, the increase in static pressure approximating a shock.
15. The turbine compressor blade according to claim 14 , wherein the low pressure surface turning towards the high pressure surface at the discontinuity in curvature when moving downstream from the discontinuity.