Repeating airfoil tip strong pressure profile

Patent metadata

Field	Value
Publication number	US-11248622-B2
Application number	US-201615255663-A
Country	US
Kind code	B2
Filing date	Sep 2, 2016
Priority date	Sep 2, 2016
Publication date	Feb 15, 2022
Grant date	Feb 15, 2022

Abstract

A compressor section for a gas turbine engine includes a blade including a platform, a tip and an airfoil extending between the platform and tip. The airfoil includes a root portion adjacent to the platform, a midspan portion and a tip portion. Each of the root portion, midspan portion and tip portion define a meridional velocity at stage exit with the tip portion including a first meridional velocity greater than a second meridional velocity of the midspan portion. A blade for an axial compressor of a gas turbine engine and a method of operating a compressor section of a gas turbine engine are also disclosed.

First claim

What is claimed is:

1. A compressor section for a gas turbine engine comprising:
a plurality of blade stages including a first blade stage and a second blade stage immediately downstream of the first blade stage, wherein the first blade stage includes a blade including:
a first platform;
a first tip;
a first leading edge and a first trailing edge extending between the first platform and the first tip; and
a first airfoil extending between the first platform and the first tip, the first airfoil including a first root portion adjacent to the first platform, a first midspan portion and a first tip portion, wherein the first airfoil includes a first airfoil shape defined as a series of coordinates that provide a meridional velocity profile across a second airfoil within the second blade stage within each of a second root portion, a second midspan portion and a second tip portion of airflow flowing across the second airfoil of the second blade stage from a second leading edge to a second trailing edge during operation of the compressor section,

wherein each of the first tip portion and the second tip portion comprises between 65% and 100% of a radial span of the corresponding one of the first airfoil and the second airfoil and each of the first midspan portion and the second midspan portion comprises between 20% and 65% of the radial span of the airfoil and the first airfoil shape within the first tip portion is configured to provide a first meridional velocity in the second blade stage outside of the second midspan portion that increases relative to a second meridional velocity at 50% of the radial span within the second midspan portion in a direction away from the second midspan portion in the second blade stage; and wherein the second blade stage includes a second airfoil shape different than the first airfoil shape and the first meridional velocity is between 6% and 15% greater than the second meridional velocity through the second midspan portion.

2. The compressor section as recited in claim 1 , wherein a first total pressure within the second tip portion is greater than a second total pressure within the second midspan portion.

3. The compressor section as recited in claim 2 , wherein the first total pressure is between 1.2% and 3% greater than the second total pressure within the second midspan portion.

4. The compressor section as recited in claim 1 , wherein the compressor section includes a plurality of stages and the first blade and the second blade comprise portions of at least two of the plurality of stages of the compressor section.

5. The compressor section as recited in claim 4 , wherein the at least two of the plurality of stages of the compressor section comprises at least two of the last 5 of the plurality of stages of the compressor section before an outlet of the compressor section.

6. The compressor section as recited in claim 4 , wherein the compressor comprises between 2 and 12 stages.

7. The compressor section as recited in claim 1 , wherein a clearance between each of the first tip and the second tip and a fixed structure of the compressor section is greater than 1.5% of the radial span of the airfoil.

8. The compressor section as recited in claim 1 , wherein the clearance between the first tip and the second tip and a fixed structure of the compressor section is between 1.5% and 3% of the radial span of the corresponding first airfoil and the second airfoil.

9. A method of operating a compressor section of a gas turbine engine, the method comprising:
configuring a first blade to include a first platform, a first tip, a first leading edge, a first trailing edge and a first airfoil that includes a first root portion adjacent the first platform, a first midspan portion and a first tip portion;
configuring a second blade downstream of the first blade to include a second platform, a second tip, a second leading edge, a second trailing edge and a second airfoil between the second platform and the second tip, the second airfoil including a second root portion adjacent the second platform, a second midspan portion and a second tip portion, wherein each of the first tip portion and the second tip portion comprises between 65% and 100% of a radial span of the corresponding one of the first airfoil and the second airfoil and each of the first midspan portion of the first airfoil and the second midspan portion of the second airfoil comprises between 20% and 65% of the radial span of the corresponding first airfoil and the second airfoil;
configuring a surface of the first airfoil for generating a first meridional velocity across the tip portion of the second blade that increases within the tip portion of the second blade in a direction toward the tip of the second blade relative to a second meridional velocity at 50% of the radial span within the midspan portion of the second blade portion for a common airflow from the first blade across the airfoil of the second blade flowing from the leading edge toward the trailing edge of the second blade; and
generating the first meridional velocity within the tip portion of the second blade that is between 6% and 15% greater than the second meridional velocity through the midspan portion of the second blade.

10. The method as recited in claim 9 , including generating a first total pressure within the tip portion of the second blade that is greater than a second total pressure within the midspan portion of the second blade.

11. The method as recited in claim 10 , wherein the generated first total pressure is between 1.2% and 3% greater than the generated second total pressure within the midspan portion of the second blade.

12. The compressor section as recited in claim 1 , wherein the first meridional velocity is between 6% and 10% greater than the second meridional velocity through the second midspan portion.

13. The compressor section as recited in claim 1 , wherein the plurality of blade stages comprise nine blade stages and the first blade stage and the second blade stage comprise the eighth blade stage and the ninth blade stage respectively.

14. The compressor section as recited in claim 1 , wherein the first airfoil shape is defined as series of x, y and z coordinates along the radial span between the first platform and the first tip.

15. The compressor section as recited in claim 1 , wherein the compressor section comprises a high pressure compressor section disposed aft of a low pressure compressor section.

16. The compressor section as recited in claim 15 , wherein the high pressure compressor is coupled to a high pressure turbine and the low pressure compressor section is coupled to a low pressure turbine disposed aft of the high pressure turbine.

17. The compressor section as recited in claim 2 , wherein the first total pressure is between 1.2% and 1.8% greater than the second total pressure within the second midspan portion.

18. The method as recited in claim 10 , wherein the generated first total pressure is between 1.2% and 1.8% greater than the generated second total pressure within the midspan portion of the second blade.

19. The method as recited in claim 11 , further comprising configuring the compressor to comprise nine blade stages with the first blade and the second blade disposed within the eighth blade stage and the ninth blade stage respectively.

20. The method as recited in claim 9 , further comprising setting a clearance between each of the first tip and the second tip and a fixed structure of the compressor section to be greater than 1.5% of the

Assignees

• United Technologies Corp
• Raytheon Tech Corp

Inventors

• Evans Simon W
• Dong Yuan
• Nolan Sean

Classifications

• F01D5/145 Primary

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

• F04D29/324

  Blades · CPC title

• F05D2270/3013

  Outlet pressure · CPC title

• F05D2220/3216

  for a special compressor stage · CPC title

• F05D2240/307

  related to the tip of a rotor blade · CPC title