Low noise compressor rotor for geared turbofan engine

Patent metadata

Field	Value
Publication number	US-2016138474-A1
Application number	US-201514967478-A
Country	US
Kind code	A1
Filing date	Dec 14, 2015
Priority date	Sep 28, 2012
Publication date	May 19, 2016
Grant date	—

Abstract

A gas turbine engine according to an example of the present disclosure includes, among other things, a fan and a turbine section having a fan drive turbine rotor, and a compressor rotor. A gear reduction effects a reduction in a speed of the fan relative to an input speed from the fan drive turbine rotor. The compressor rotor has a number of compressor blades in at least half of a plurality of blade rows of the compressor rotor. The blades are configured to operate at least some of the time at a rotational speed. The number of compressor blades in the at least half of the blade rows and the rotational speed is such that the following formula holds true for each row of the at least half of the blade rows of the compressor rotor: (the number of blades×the rotational speed)/60 s≧about 5500 Hz.

First claim

1 . A gas turbine engine comprising:
a fan and a turbine section having a fan drive turbine rotor, and a compressor rotor; a gear reduction effecting a reduction in a speed of said fan relative to an input speed from said fan drive turbine rotor; said compressor rotor having a number of compressor blades in at least half of a plurality of blade rows of said compressor rotor, and said blades configured to operate at least some of the time at a rotational speed, and said number of compressor blades in said at least half of said blade rows and said rotational speed being such that the following formula holds true for each row of said at least half of said blade rows of the compressor rotor: (said number of blades×said rotational speed)/60 sec≧about 5500 Hz; and said rotational speed being an approach speed in revolutions per minute.

2 . The gas turbine engine as set forth in claim 1 , wherein the formula results in a number greater than or equal to about 6000 Hz.

3 . The gas turbine engine as set forth in claim 2 , wherein said gas turbine engine is rated to produce about 15,000 pounds of thrust or more.

4 . The gas turbine engine as set forth in claim 1 , wherein the formula holds true for a majority of the blade rows of the compressor rotor.

5 . The gas turbine engine as set forth in claim 4 , wherein the formula holds true for all of the blade rows of the compressor rotor.

6 . The gas turbine engine as set forth in claim 1 , wherein said gas turbine engine is rated to produce about 15,000 pounds of thrust or more.

7 . The gas turbine engine as set forth in claim 1 , wherein said gear reduction has a gear ratio of greater than about 2.3.

8 . The gas turbine engine as set forth in claim 7 , wherein said gear reduction has a gear ratio of greater than about 2.5.

9 . The gas turbine engine as set forth in claim 1 , wherein said fan delivers air into a bypass duct, and a portion of air into said compressor rotor, with a bypass ratio defined as the volume of air delivered into the bypass duct compared to the volume of air delivered into the compressor rotor, and said bypass ratio being greater than about 6.

10 . The gas turbine engine as set forth in claim 9 , wherein said bypass ratio is greater than about 10.

11 . The gas turbine engine as set forth in claim 10 , wherein the formula results in a number greater than or equal to about 6000 Hz.

12 . The gas turbine engine as set forth in claim 1 , wherein the formula results in a number less than or equal to about 7000 Hz.

13 . The gas turbine engine as set forth in claim 1 , wherein the formula results in a number less than or equal to about 10000 Hz.

14 . The gas turbine engine as set forth in claim 1 , wherein said turbine section including a higher pressure turbine rotor and a lower pressure turbine rotor, and said fan drive turbine rotor being said lower pressure turbine rotor.

15 . The gas turbine engine as set forth in claim 14 , wherein said compressor rotor is a lower pressure compressor rotor, and said higher pressure turbine rotor driving a higher pressure compressor rotor.

16 . The gas turbine engine as set forth in claim 1 , wherein there are three turbine rotors, the fan drive rotor turbine driving the fan, and a second and third turbine rotor each driving respective compressor rotors of the compressor section.

17 . The gas turbine engine as set forth in claim 1 , wherein the gear reduction is positioned intermediate the fan and a compressor rotor driven by the fan drive turbine rotor.

18 . The gas turbine engine as set forth in claim 1 , wherein the gear reduction is positioned intermediate the fan drive turbine rotor and a compressor rotor driven by the fan drive turbine rotor.

19 . A method of designing a gas turbine engine comprising the steps of:
including a first turbine rotor to drive a compressor rotor and a fan turbine rotor for driving a fan through a gear reduction, and selecting a number of blades in at least half of blade rows of the compressor rotor, in combination with a rotational speed of the compressor rotor, such that the following formula holds true for each row of said at least half of said blade rows of the compressor rotor: (said number of blades x said rotational speed)/60 sec≧about 5500 Hz; and said rotational speed being an approach speed in revolutions per minute.

20 . The method as set forth in claim 19 , wherein the formula results in a number greater than or equal to about 6000 Hz.

21 . The method as set forth in claim 20 , wherein said gas turbine engine is rated to produce about 15,000 pounds of thrust or more.

22 . The method as set forth in claim 20 , wherein the formula holds true for a majority of the blade rows of the compressor rotor.

23 . The method as set forth in claim 19 , wherein the formula holds true for all of the blade rows of the compressor rotor.

24 . The method as set forth in claim 19 , wherein the formula results in a number less than or equal to about 7000 Hz.

25 . The method as set forth in claim 19 , wherein the formula results in a number less than or equal to about 10000 Hz.

26 . The method as set forth in claim 19 , wherein said fan drive turbine rotor is a lower pressure turbine rotor and said first turbine rotor is a higher pressure turbine rotor.

27 . The method as set forth in claim 26 , wherein said compressor rotor is a lower pressure compressor rotor, and said higher pressure turbine rotor driving a higher pressure compressor rotor.

28 . The method as set forth in claim 19 , wherein said first turbine rotor and said fan turbine rotor are provided by a single rotor.

29 . The gas turbine engine as set forth in claim 1 , wherein the formula does not hold true for all of the blade rows of the compressor rotor.

30 . The gas turbine engine as set forth in claim 1 , wherein the formula results in a number less than or equal to about 6000 Hz.

Assignees

• United Technologies Corp

Inventors

• Topol David A
• Morin Bruce L

Classifications

• F02C7/24 Primary

  Heat or noise insulation (air intakes having provisions for noise suppression F02C7/045; turbine exhaust heads, chambers, or the like F01D25/30; silencing nozzles of jet-propulsion plants F02K1/00) · CPC title

• F02K3/06 Primary

  with front fan · CPC title

• F02C7/32

  Arrangement, mounting, or driving, of auxiliaries · CPC title

• F02C3/10

  with another turbine driving an output shaft but not driving the compressor · CPC title

• Y02T50/60

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