GAS TURBINE WITH OPTIMIZED AIRFOIL ELEMENT ANGLES

• US 20130089415A1
• Filed: 08/20/2012
• Published: 04/11/2013
• Est. Priority Date: 10/06/2011
• Status: Active Grant

• Current Assignee: Siemens Energy Incorporated (Siemens AG)
• Original Assignee: Siemens Energy Incorporated (Siemens AG)
• Inventors: Lee, Ching-Pang, Malandra, Anthony J., Munoz, Eric, Brown, Barry J.
• Granted Patent: US 8,864,457 B2
• Time in Patent Office: Days
• US Class Current: 415/193
• CPC Class Codes: F01D 5/141 Shape, i.e. outer, aerodynamic form F01D5/148 - F01D5/20 take precedence; blade construction F01D5/147
• CPC Class Codes: F01D 9/041 using blades F01D5/148 takes precedence
• CPC Class Codes: F05D 2220/3213 an intermediate stage of the turbine
• CPC Class Codes: F05D 2250/74 given by a set or table of xyz-coordinates

Abstract

A turbine airfoil assembly for installation in a gas turbine engine. The airfoil assembly includes an endwall and an airfoil extending radially outwardly from the endwall. The airfoil includes pressure and suction sidewalls defining chordally spaced apart leading and trailing edges of the airfoil. An airfoil mean line is defined located centrally between the pressure and suction sidewalls. An angle between the mean line and a line parallel to the engine axis at the leading and trailing edges defines gas flow entry angles, α, and exit angles, β. Airfoil inlet and exit angles are substantially in accordance with pairs of inlet angle values, α, and exit angle values, β, set forth in one of Tables 1, 3, 5 and 7.

Claims

1. A turbine airfoil assembly for installation in a gas turbine engine having a longitudinal axis, the turbine airfoil assembly including an endwall for defining an inner boundary for an axially extending hot working gas path, and an airfoil extending radially outwardly from the endwall, said airfoil having an outer wall comprising a pressure sidewall and a suction sidewall joined together at chordally spaced apart leading and trailing edges of said airfoil, an airfoil mean line is defined extending chordally and located centrally between said pressure and suction sidewalls, airfoil inlet and exit angles are defined at said airfoil leading and trailing edges that are substantially in accordance with pairs of inlet angle values, α

, and exit angle values, β

, set forth in one of Tables 1, 3, 5 and 7, where said inlet and exit angle values are generally defined as angles between a line parallel to the longitudinal axis and the airfoil mean line lying in an X-Y plane of an X, Y, Z Cartesian coordinate system in which Z is a dimension perpendicular to the X-Y plane and extends radially relative to the longitudinal axis, and wherein each pair of inlet and exit angle values is defined with respect to a distance from said endwall corresponding to a Z value that is a percentage of the total span of said airfoil from said endwall, and wherein a predetermined difference between each pair of said airfoil inlet and exit angles is defined by a delta value, Δ

, in said Table, and a difference between any pair of said airfoil inlet and exit angles varies from the delta values, Δ

, in said Table by at most 5%.

2. The turbine airfoil assembly of claim 1, wherein said airfoil comprises an airfoil for a third stage vane in a turbine engine, and said Table defining said airfoil inlet and exit angles is Table 1.

3. The turbine airfoil assembly of claim 1, wherein said airfoil comprises an airfoil for a third stage blade in a turbine engine, and said Table defining said airfoil inlet and exit angles is Table 3.

4. The turbine airfoil assembly of claim 1, wherein said airfoil comprises an airfoil for a fourth stage vane in a turbine engine, and said Table defining said airfoil inlet and exit angles is Table 5.

5. The turbine airfoil assembly of claim 1, wherein said airfoil comprises an airfoil for a fourth stage blade in a turbine engine, and said Table defining said airfoil inlet and exit angles is Table 7.

6. The turbine airfoil assembly of claim 1, including four of said airfoil assemblies comprising, in succession, an airfoil for a third stage vane having said airfoil inlet and exit angles defined by Table 1, an airfoil for a third stage blade having said airfoil inlet and exit angles defined by Table 3, an airfoil for a fourth stage vane having said airfoil inlet and exit angles defined by Table 5 and an airfoil for a fourth stage blade having said airfoil inlet and exit angles defined by Table 7.

7. The turbine airfoil assembly of claim 6, wherein said difference between any pair of said airfoil inlet and exit angles varies from said delta values, Δ

, in a respective Table by at most 3%.

8. The turbine airfoil assembly of claim 6, wherein said difference between any pair of said airfoil inlet and exit angles varies from said delta values, Δ

, in a respective Table by at most 1%.

9. Third and fourth stage vane and blade airfoil assemblies in a gas turbine engine having a longitudinal axis, each airfoil assembly including:

an endwall for defining an inner boundary for an axially extending hot working gas path, and an airfoil extending radially outwardly from the endwall, said airfoil having an outer wall comprising a pressure sidewall and a suction sidewall joined together at chordally spaced apart leading and trailing edges of said airfoil, an airfoil mean line is defined extending chordally and located centrally between said pressure and suction sidewalls, airfoil inlet and exit angles are defined at said airfoil leading and trailing edges that are substantially in accordance with pairs of inlet angle values, α

, and exit angle values, β

, where said inlet and exit angle values are generally defined as angles between a line parallel to the longitudinal axis and the airfoil mean line lying in an X-Y plane of an X, Y, Z Cartesian coordinate system in which Z is a dimension perpendicular to the X-Y plane and extends radially relative to the longitudinal axis, and wherein each pair of inlet and exit angle values is defined with respect to a distance from said endwall corresponding to a Z value that is a percentage of the total span of said airfoil from said endwall, wherein;

a) said pairs of inlet angle values, α

, and exit angle values, β

, for said third stage vane are as set forth in Table 1;

b) said pairs of inlet angle values, α

, and exit angle values, β

, for said third stage blade are as set forth in Table 3;

c) said pairs of inlet angle values, α

, and exit angle values, β

, for said fourth stage vane are as set forth in Table 5;

d) said pairs of inlet angle values, α

, and exit angle values, β

, for said fourth stage blade are as set forth in Table 7; and

wherein a predetermined difference between each pair of said airfoil inlet and exit angles is defined by a delta value, Δ

, in said Table, and a difference between any pair of said airfoil inlet and exit angles varies from the delta values, Δ

, in a respective Table by at most 5%.

10. The turbine airfoil assembly of claim 9, wherein said difference between any pair of said airfoil inlet and exit angles varies from said delta values, Δ

, in a respective Table by at most 3%.

11. The turbine airfoil assembly of claim 9, wherein said difference between any pair of said airfoil inlet and exit angles varies from said delta values, Δ

, in a respective Table by at most 1%.

12. A turbine airfoil assembly for installation in a gas turbine engine having a longitudinal axis, the turbine airfoil assembly including an endwall for defining an inner boundary for an axially extending hot working gas path, and an airfoil extending radially outwardly from the endwall, said airfoil having an outer wall comprising a pressure sidewall and a suction sidewall joined together at chordally spaced apart leading and trailing edges of said airfoil, an airfoil mean line is defined extending chordally and located centrally between said pressure and suction sidewalls, airfoil exit angles are defined at said airfoil trailing edge that are substantially in accordance with exit angle values, β

, set forth in one of Tables 1, 3, 5 and 7, where said exit angle values are generally defined as angles between a line parallel to the longitudinal axis and the airfoil mean line lying in an X-Y plane of an X, Y, Z Cartesian coordinate system in which Z is a dimension perpendicular to the X-Y plane and extends radially relative to the longitudinal axis, wherein each said exit angle value is defined with respect to a distance from said endwall corresponding to a Z value that is a percentage of the total span of said airfoil from said endwall, and wherein each said airfoil exit angle is within about 1% of a respective value set forth in said Table.

13. The turbine airfoil assembly of claim 12, wherein said airfoil comprises an airfoil for a third stage vane in a turbine engine, and said Table defining said airfoil exit angles is Table 1.

14. The turbine airfoil assembly of claim 12, wherein said airfoil comprises an airfoil for a third stage blade in a turbine engine, and said Table defining said airfoil exit angles is Table 3.

15. The turbine airfoil assembly of claim 12, wherein said airfoil comprises an airfoil for a fourth stage vane in a turbine engine, and said Table defining said airfoil exit angles is Table 5.

16. The turbine airfoil assembly of claim 12, wherein said airfoil comprises an airfoil for a fourth stage blade in a turbine engine, and said Table defining said airfoil exit angles is Table 7.

17. The turbine airfoil assembly of claim 12, including four of said airfoil assemblies comprising, in succession, an airfoil for a third stage vane having airfoil exit angles defined by Table 1, an airfoil for a third stage blade having airfoil exit angles defined by Table 3, an airfoil for a fourth stage vane having airfoil exit angles defined by Table 5 and an airfoil for a fourth stage blade having airfoil exit angles defined by Table 7.

18. The turbine airfoil assembly of claim 12, including at least two of said airfoil assemblies comprising, in succession, an airfoil for a third stage blade having airfoil exit angles defined by Table 3, and an airfoil for a fourth stage vane having airfoil exit angles defined by Table 5.