Detuning trailing edge compound lean contour

Patent metadata

Field	Value
Publication number	US-10233758-B2
Application number	US-201415023912-A
Country	US
Kind code	B2
Filing date	Oct 2, 2014
Priority date	Oct 8, 2013
Publication date	Mar 19, 2019
Grant date	Mar 19, 2019

Abstract

A rotor blade comprises a blade platform and an airfoil. The airfoil comprises a blade tip, a leading edge, and a trailing edge with a stiffening compound lean contour. The blade tip is radially opposite the blade platform, and defines a blade span between the blade itself and the blade platform. The leading and trailing edges extend from the blade platform to the blade tip to define blade chords between the leading edge and the trailing edge. The compound lean contour comprises a positive lean section located at the lean tip, and extending along a lean axis to a lean end. The negative lean section is located radially inward of the positive lean section, and extends along the lean axis to the lean end.

First claim

The invention claimed is:

1. A rotor blade comprising:
a blade platform and
an airfoil comprising:
a blade tip radially opposite the blade platform, and defining a blade span therebetween; and
leading and trailing edges extending from the blade platform to the blade tip to define blade chords between the leading edge and the trailing edge;
wherein the trailing edge has a compound lean contour comprising:
a positive lean section located at the blade tip, and extending along a lean axis to a lean end;
a negative lean section located radially inward of the positive lean section, and extending along the lean axis to the lean end; and
wherein the lean axis is angled radially outward from the trailing edge to the lean end and oriented at an angle between 0° and 30° with respect to the blade chords.

2. The rotor blade claim 1 , wherein the lean axis is oriented at an angle between 0° and 15° with respect to the blade chords.

3. The rotor blade claim 1 , wherein the lean axis is oriented at an angle between 15° and 30° with respect to the blade chords.

4. The rotor blade of claim 1 , wherein the compound lean contour extends in the spanwise direction from the blade tip across less than 40% of the blade span.

5. The rotor blade of claim 4 , wherein the compound lean contour extends in a spanwise direction from the blade tip across less than 20% of the blade span.

6. The rotor blade of claim 1 , wherein the compound lean contour extends in the chordwise direction from the trailing edge across less than 75% of the blade chord, to the lean end.

7. The rotor blade of claim 6 , wherein the compound lean contour extends in a chordwise direction from the trailing edge across less than 50% of the blade chord, to the lean end.

8. The rotor blade of claim 1 , wherein the compound lean contour increases the local stiffness of the rotor blade near a tip of the rotor blade trailing edge.

9. The rotor blade of claim 8 , wherein the rotor blade is a turbine blade of a turbine section of a gas turbine engine.

10. The rotor blade of claim 9 , wherein the increased local stiffness of the rotor blade near a tip of the rotor blade trailing edge blade serves to detune a natural frequency of the rotor blade away from an excitation engine order.

11. The rotor blade of claim 9 , wherein the turbine section is a low pressure turbine power head module of an aircraft auxiliary power unit.

12. The rotor blade of claim 1 , wherein lean angle within the positive lean section is between 90° and 135°.

13. The rotor blade of claim 1 , wherein lean angle within the negative lean section is between 90° and 45°.

14. The rotor blade of claim 1 , wherein the rotor blade is a turbine blade or compressor blade.

15. A gas turbine engine including a turbine section that includes a plurality of alternating stages of blades and vanes, each blade comprising:
a blade platform;
an airfoil extending from the blade platform to a blade tip radially opposite the blade platform, thereby defining a blade span therebetween; and
leading and trailing edges extending from the blade platform to the blade tip to define blade chords between the leading edge and the trailing edge;
wherein the trailing edge has a stiffening compound lean contour with a positive lean section at the blade tip, and a negative lean section located radially inward of the positive lean section;
wherein the lean axis is angled radially outward from the trailing edge to the lean end and oriented at an angle between 0° and 30° with respect to the blade chords.

16. The gas turbine engine of claim 15 , wherein the compound lean contour extends along a lean axis that is angled radially outwards from the trailing edge to a lean end, and that extends across less than 75% of one of the blade chords.

17. The gas turbine engine of claim 16 , wherein the lean axis extends across less than 50% of one of the blade chords.

18. The gas turbine engine of claim 15 , wherein the compound lean contour extends across less than 40% of the blade span.

19. The gas turbine engine of claim 18 , wherein the compound lean contour extends across less than 20% of the blade span.

20. The gas turbine engine of claim 15 , wherein gas turbine engine is an aircraft auxiliary power unit, and the turbine section is a low pressure turbine of a power heat module.

21. An airfoil comprising:
an external surface formed in in substantial conformance with a plurality of cross-sectional profiles of the airfoil, each of the cross-sectional profiles being defined by a plurality of points including normalized Cartesian coordinates as set forth in Table 1.

22. The airfoil of claim 21 , wherein substantial conformance includes manufacturing tolerances of about ±0.005 inches (±0.127 mm).

23. The airfoil of claim 21 , wherein the external surface is defined such that the points and the adjacent cross-sectional profiles are connected via smooth arcs to define a smooth curve.

Assignees

• United Technologies Corp

Inventors

• Duong Loc Quang
• Hu Xiaolan
• Yang Gao
• Garcia Jesus A
• Sapiro Dan G

Classifications

• F05D2240/307

  related to the tip of a rotor blade · CPC title

• F01D5/26

  Antivibration means not restricted to blade form or construction or to blade-to-blade connections {or to the use of particular materials} · CPC title

• F01D5/145

  Means for influencing boundary layers or secondary circulations (for compressors F04D29/68) · CPC title

• F01D5/141 Primary

  Shape, i.e. outer, aerodynamic form (F01D5/148 - F01D5/20 take precedence; blade construction F01D5/147) · CPC title

• F01D5/16 Primary

  for counteracting blade vibration · CPC title