Throat distribution for a rotor and rotor blade having camber and location of local maximum thickness distribution

Patent metadata

Field	Value
Publication number	US-11280199-B2
Application number	US-201816198260-A
Country	US
Kind code	B2
Filing date	Nov 21, 2018
Priority date	Nov 21, 2018
Publication date	Mar 22, 2022
Grant date	Mar 22, 2022

Abstract

A rotor blade for a compressor of a gas turbine engine includes an airfoil. The airfoil has a span that extends from 0% at the root to 100% at the tip and a mean camber line that extends from a leading edge to a trailing edge. The airfoil has a location of local maximum thickness defined as a ratio of a first arc distance along the mean camber line between the leading edge and a position of the local maximum thickness to a total arc distance along the mean camber line from the leading edge to the trailing edge. A value of the ratio increases from the root to a first position value, decreases from the first position value to a second position value and increases from the second position value to the tip. The first position value is at a spanwise location within 20% to 50% of the span.

First claim

What is claimed is:

1. A rotor blade for a compressor of a gas turbine engine, comprising:
an airfoil extending from a root to a tip and having a leading edge and a trailing edge, with a span that extends from 0% at the root to 100% at the tip and a mean camber line that extends from the leading edge to the trailing edge, the airfoil having a location of local maximum thickness defined as a ratio of a first arc distance along the mean camber line between the leading edge and a position of the local maximum thickness to a total arc distance along the mean camber line from the leading edge to the trailing edge, and a value of the ratio increases from the root to a first position value, decreases from the first position value to a second position value and increases from the second position value to the tip and the first position value is at a spanwise location within 20% to 50% of the span.

2. The rotor blade of claim 1 , wherein the second position value is at a spanwise location within 60% to 90% of the span.

3. The rotor blade of claim 1 , wherein at the tip, the value of the ratio is a third position value that is greater than the first position value, and the third position value is an absolute maximum value of the ratio over the span of the airfoil.

4. The rotor blade of claim 1 , wherein at the tip, the value of the ratio is a third position value, the third position value is less than the first position value, and the first position value is an absolute maximum value of the ratio over the span of the airfoil.

5. The rotor blade of claim 1 , wherein the value of the ratio has a root position value at the root, and the root position value is an absolute minimum value of the ratio over the span of the airfoil.

6. The rotor blade of claim 1 , wherein the value of the ratio has a fourth position value at 10% of the span, and the fourth position value is less than the first position value.

7. The rotor blade of claim 1 , wherein the airfoil further comprises a total camber distribution that increases from the root to a maximum value of total camber between 5% of the span and 20% of the span.

8. The rotor blade of claim 1 , wherein between at least one of the root and the first position value, the first position value and the second position value, and the second position value and the tip, the value of the ratio has a local increase or a local decrease.

9. A rotor blade for a compressor of a gas turbine engine, comprising:
an airfoil extending from a root to a tip and having a leading edge and a trailing edge, with a span that extends from 0% at the root to 100% at the tip and a mean camber line that extends from the leading edge to the trailing edge, the airfoil having a location of local maximum thickness defined as a ratio of a first arc distance along the mean camber line between the leading edge and a position of the local maximum thickness to a total arc distance along the mean camber line from the leading edge to the trailing edge, a value of the ratio increases from a root position value at the root to a first position value, decreases from the first position value to a second position value and increases from the second position value to a third position value at the tip, and the second position value is within 60% to 90% of the span.

10. The rotor blade of claim 9 , wherein the first position value is at a spanwise location within 20% to 50% of the span.

11. The rotor blade of claim 9 , wherein the third position value is less than the first position value, and the first position value is an absolute maximum value of the ratio over the span of the airfoil.

12. The rotor blade of claim 9 , wherein the third position value is greater than the first position value, and the third position value is an absolute maximum value of the ratio over the span of the airfoil.

13. The rotor blade of claim 9 , wherein the root position value is an absolute minimum value of the ratio over the span of the airfoil.

14. The rotor blade of claim 9 , wherein the value of the ratio has a fourth position value at 10% of the span, and the fourth position value is less than the first position value.

15. A rotor for a compressor of a gas turbine engine, comprising:
a hub; and
an airfoil extending from a root to a tip and having a leading edge and a trailing edge, with a span that extends from 0% at the root to 100% at the tip and a mean camber line that extends from the leading edge to the trailing edge, the airfoil having a location of local maximum thickness defined as a ratio of a first arc distance along the mean camber line between the leading edge and a position of the local maximum thickness to a total arc distance along the mean camber line from the leading edge to the trailing edge, a value of the ratio increases from a root position value at the root to a first position value, decreases from the first position value to a second position value and increases from the second position value to a third position value at the tip, the root position value is an absolute minimum value of the ratio over the span of the airfoil and the second position value is at a spanwise location within 60% to 90% of the span.

16. The rotor of claim 15 , wherein the first position value is at a spanwise location within 20% to 50% of the span.

17. The rotor of claim 15 , wherein the third position value is less than the first position value and is greater than the second position value, and the first position value is an absolute maximum value of the ratio over the span of the airfoil.

18. The rotor of claim 15 , wherein the third position value is greater than the first position value, and the third position value is an absolute maximum value of the ratio over the span of the airfoil.

19. The rotor of claim 15 , further comprising a plurality of the airfoils, each of the plurality of the airfoils coupled to the hub at the root and spaced apart from adjacent ones of the plurality of the airfoils over the span by a throat dimension defined between the adjacent ones of the plurality of the airfoils, and the throat dimension has a maximum value at a spanwise location between 60% of the span and 90% of the span of the adjacent ones of the plurality of the airfoils.

20. The rotor of claim 15 , wherein the airfoil further comprises a total camber distribution that increases from the root to a maximum value of total camber between 5% of the span and 20% of the span.

Assignees

• Honeywell Int Inc

Inventors

• Mcnamee Michael R
• Nolcheff Nick
• Gebre-Giorgis Yoseph
• Gunaraj John A
• Hayes Jeffrey
• Vogiatzis Constantinos

Classifications

• F05D2250/70

  Shape · CPC title

• F01D5/141 Primary

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

• Y02T50/60

  Efficient propulsion technologies, e.g. for aircraft · CPC title

• F05D2240/301

  Cross-sectional characteristics · CPC title

• F04D29/324

  Blades · CPC title