US20160130949A1 - Low noise turbine for geared turbofan engine - Google Patents

Abstract

A turbine module according to an example of the present disclosure includes, among other things, a fan drive rotor having a plurality of blade rows each including a number of blades. A majority of the blade rows are capable of rotating at a rotational speed, so that when measuring the rotational speed in revolutions per minute: (number of blades×the rotational speed)/60≧about 5500 Hz.

Description

CROSS-REFERENCE TO RELATED APPLICATION
• This application is a continuation-in-part of U.S. patent application Ser. No. 14/795931, filed Jul. 10, 2015, which was a continuation-in-part of U.S. patent application Ser. No. 14/248,386, filed Apr. 4, 2014, which was a continuation-in-part of International Application No. PCT/U.S.2013/020724 filed Jan. 9, 2013 which claims priority to U.S. Provisional Application No. 61/592,643, filed Jan. 31, 2012. U.S. patent application Ser. No. 14/248,386 further claims priority to U.S. Provisional Application No. 61/884,660 filed Sep. 30, 2013.
BACKGROUND
• This application relates to the design of a turbine which can be operated to produce noise to which human hearing is less sensitive.
• Gas turbine engines are known, and typically include a fan delivering air into a compressor. The air is compressed in the compressor and delivered downstream into a combustor section where it was mixed with fuel and ignited. Products of this combustion pass downstream over turbine rotors, driving the turbine rotors to rotate.
• Typically, there is a high pressure turbine rotor, and a low pressure turbine rotor. Each of the turbine rotors includes a number of rows of turbine blades which rotate with the rotor. Typically interspersed between the rows of turbine blades are vanes.
• The low pressure turbine can be a significant noise source, as noise is produced by fluid dynamic interaction between the blade rows and the vane rows. These interactions produce tones at a blade passage frequency of each of the low pressure turbine stages, and their harmonics.
• The noise can often be in a frequency range to which humans are very sensitive. To mitigate this problem, in the past, a vane-to-blade ratio of the fan drive turbine has been controlled to be above a certain number. As an example, a vane-to-blade ratio may be selected to be 1.5 or greater, to prevent a fundamental blade passage tone from propagating to the far field. This is known as acoustic “cut-off.”
• However, acoustically cut-off designs may come at the expense of increased weight and reduced aerodynamic efficiency. Stated another way, if limited to a particular vane to blade ratio, the designer may be restricted from selecting such a ratio based upon other characteristics of the intended engine.
• Historically, the low pressure turbine has driven both a low pressure compressor section and a fan section. More recently, a gear reduction has been provided such that the fan and low pressure compressor can be driven at different speeds.
SUMMARY
• A gas turbine engine according to an example of the present disclosure includes a fan and a turbine having a fan drive rotor. There also is a second turbine rotor, and a gear reduction effecting a reduction in a speed of the fan relative to an input speed from the fan drive rotor. The fan drive rotor has a number of turbine blades in a majority of a plurality of blade rows of the fan drive rotor, and the turbine blades are configured to operate at least some of the time at a rotational speed. The number of turbine blades in the majority of the blade rows and the rotational speed is such that the following formula holds true for each row of the majority of the blade rows of the fan drive turbine: (number of blades×speed)/60≧about 5500 Hz. The rotational speed is in revolutions per minute.
• In a further embodiment of any of the forgoing embodiments, the formula results in a number greater than or equal to about 6000 Hz.
• In a further embodiment of any of the forgoing embodiments, the gas turbine engine is rated to produce 15,000 pounds of thrust or more.
• In a further embodiment of any of the forgoing embodiments, the formula holds true for all of the blade rows of the fan drive rotor.
• In a further embodiment of any of the forgoing embodiments, the formula does not hold true for all of the blade rows of the fan drive rotor.
• In a further embodiment of any of the forgoing embodiments, the formula results in a number less than or equal to about 7000 Hz, the rotational speed being an approach speed.
• In a further embodiment of any of the forgoing embodiments, the formula results in a number less than or equal to about 10000 Hz, the rotational speed being a takeoff speed.
• In a further embodiment of any of the forgoing embodiments, the turbine section has a higher pressure turbine rotor and a lower pressure turbine rotor, with the fan drive rotor being the lower pressure turbine rotor and the second turbine rotor being the higher pressure turbine rotor.
• In a further embodiment of any of the forgoing embodiments, there is a third turbine rotor, with the fan drive turbine being a most downstream of the three turbine rotors.
• A method of designing a gas turbine engine according to an example of the present disclosure includes the steps of including a gear reduction between a fan drive turbine rotor and a fan, and selecting a number of blades in a majority of blade rows of the fan drive turbine rotor, in combination with a rotational speed of the fan drive turbine rotor, such that the following formula holds true for each row of the majority of blade rows of the fan drive turbine rotor: (number of blades×speed)/60≧about 5500 Hz. The rotational speed is in revolutions per minute and includes a second turbine rotor.
• In a further embodiment of any of the forgoing embodiments, the formula results in a number greater than or equal to about 6000.
• In a further embodiment of any of the forgoing embodiments, the gas turbine engine is rated to produce 15,000 pounds of thrust or more.
• In a further embodiment of any of the forgoing embodiments, the formula holds true for all of the blade rows of the fan drive turbine.
• In a further embodiment of any of the forgoing embodiments, the formula does not hold true for all of the blade rows of the fan drive turbine.
• In a further embodiment of any of the forgoing embodiments, the formula results in a number less than or equal to about 7000 Hz, and the rotational speed is an approach speed.
• In a further embodiment of any of the forgoing embodiments, the formula results in a number less than or equal to about 10000 Hz, and the rotational speed is a takeoff speed.
• In a further embodiment of any of the forgoing embodiments, a turbine section including a higher pressure turbine rotor and a lower pressure turbine rotor, and the fan drive turbine rotor being the lower pressure turbine rotor.
• A turbine module according to an example of the present disclosure includes a fan drive rotor having a plurality of blade rows each including a number of blades. A majority of the blade rows are capable of rotating at a rotational speed, so that when measuring the rotational speed in revolutions per minute: (number of blades×the rotational speed)/60≧about 5500 Hz.
• In a further embodiment of any of the forgoing embodiments, the formula results in a number greater than or equal to about 6000.
• In a further embodiment of any of the forgoing embodiments, the formula holds true for all of the blade rows of the fan drive rotor.
• In a further embodiment of any of the forgoing embodiments, the formula does not hold true for all of the blade rows of the fan drive rotor.
• In a further embodiment of any of the forgoing embodiments, a pressure ratio across the fan drive rotor is greater than about 5:1.
• In a further embodiment of any of the forgoing embodiments, the formula results in a number less than or equal to about 7000 Hz, and the rotational speed is an approach speed.
• In a further embodiment of any of the forgoing embodiments, the formula results in a number less than or equal to about 10000 Hz, and the rotational speed is a takeoff speed.
• In a further embodiment of any of the forgoing embodiments, there is a higher pressure turbine rotor and a lower pressure turbine rotor, and the fan drive rotor is the lower pressure turbine rotor.
• Although the different examples have the specific components shown in the illustrations, embodiments of this disclosure are not limited to those particular combinations. It is possible to use some of the components or features from one of the examples in combination with features or components from another one of the examples.
• These and other features disclosed herein can be best understood from the following specification and drawings, the following of which is a brief description.
BRIEF DESCRIPTION OF THE DRAWINGS
• FIG. 1 shows a gas turbine engine.
• FIG. 2 shows another embodiment.
• FIG. 3 shows yet another embodiment.
DETAILED DESCRIPTION
• FIG. 1 schematically illustrates agas turbine engine20. Thegas turbine engine20 is disclosed herein as a two-spool turbofan that generally incorporates afan section22, acompressor section24, acombustor section26 and aturbine section28. Alternative engines might include an augmentor section (not shown), or an intermediate spool, among other systems or features. Thefan section22 drives air along a bypass flowpath B while thecompressor section24 drives air along a core flowpath C for compression and communication into thecombustor section26 then expansion through theturbine section28. Although depicted as a two-spool turbofan gas turbine engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to use with two-spool turbofans as the teachings may be applied to other types of turbine engines including three-spool architectures.
• Theengine20 generally includes alow speed spool30 and ahigh speed spool32 mounted for rotation about an engine central longitudinal axis A relative to an enginestatic structure36 viaseveral bearing systems38. It should be understood that various bearingsystems38 at various locations may alternatively or additionally be provided.
• Thelow speed spool30 generally includes aninner shaft40 that interconnects afan42, alow pressure compressor44 and alow pressure turbine46. Theinner shaft40 is connected to thefan42 through a gearedarchitecture48 to drive thefan42 at a lower speed than thelow speed spool30. Thehigh speed spool32 includes anouter shaft50 that interconnects ahigh pressure compressor52 andhigh pressure turbine54. Acombustor56 is arranged between thehigh pressure compressor52 and thehigh pressure turbine54. Amid-turbine frame57 of the enginestatic structure36 is arranged generally between thehigh pressure turbine54 and thelow pressure turbine46. Themid-turbine frame57 furthersupports bearing systems38 in theturbine section28. Theinner shaft40 and theouter shaft50 are concentric and rotate via bearingsystems38 about the engine central longitudinal axis A which is collinear with their longitudinal axes.
• The core airflow is compressed by thelow pressure compressor44 then thehigh pressure compressor52, mixed and burned with fuel in thecombustor56, then expanded over thehigh pressure turbine54 andlow pressure turbine46. Themid-turbine frame57 includesairfoils59 which are in the core airflow path. Theturbines46,54 rotationally drive the respectivelow speed spool30 andhigh speed spool32 in response to the expansion.
• The terms “low” and “high” as applied to speed or pressure for the spools, compressors and turbines are of course relative to each other. That is, the low speed spool operates at a lower speed than the high speed spool, and the low pressure sections operate at lower pressure than the high pressures sections.
• Theengine20 in one example is a high-bypass geared aircraft engine. In a further example, theengine20 bypass ratio is greater than about six (6), with an example embodiment being greater than ten (10), the gearedarchitecture48 is an epicyclic gear train, such as a star system, a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3:1 or greater than about 2.5:1. In some embodiments, the bypass ratio is less than about thirty (30), or more narrowly less than about twenty (20). In embodiments, the gear reduction ratio is less than about 5.0, or less than about 4.0. In one disclosed embodiment, theengine20 bypass ratio is greater than about ten (10:1), the fan diameter is significantly larger than that of thelow pressure compressor44, and thelow pressure turbine46 has a pressure ratio that is greater than about 5:1. Thelow pressure turbine46 pressure ratio is a ratio of the pressure measured at inlet oflow pressure turbine46 to the pressure at the outlet of the low pressure turbine46 (prior to an exhaust nozzle). It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention is applicable to other gas turbine engines including direct drive turbofans.
• A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. Thefan section22 of theengine20 is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft, with the engine at its best fuel consumption—also known as “bucket cruise Thrust Specific Fuel Consumption (‘TSFC’)”—is the industry standard parameter of lbm of fuel being burned divided by lbf of thrust the engine produces at that minimum point. “Low fan pressure ratio” is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.50 and, in some embodiments, is less than about 1.45. “Low corrected fan tip speed” is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tambient deg R)/518.7)̂0.5]. The “Low corrected fan tip speed” as disclosed herein according to one non-limiting embodiment is less than about 1150 ft/second.
• The use of the gear reduction between the low pressure turbine spool and the fan allows an increase of speed to the low pressure compressor. In the past, the speed of the low pressure turbine has been somewhat limited in that the fan speed cannot be unduly high. The maximum fan speed is at its outer tip, and in larger engines, the fan diameter is much larger than it may be in lower power engines. However, a gear reduction may be used to free the designer from compromising low pressure turbine speed in order not to have unduly high fan speeds.
• It has been discovered that a careful design between the number of rotating blades, and the rotational speed of the low pressure turbine can be selected to result in noise frequencies that are less sensitive to human hearing.
• A formula has been developed as follows:
• 
  (blade count×rotational speed)/(60 seconds/minute)≧4000 Hz.
• That is, the number of rotating blades in any low pressure turbine stage, multiplied by the rotational speed of the low pressure turbine (in revolutions per minute), divided by 60 seconds per minute (to put the amount per second, or Hertz) should be greater than or equal to about 4000 Hz. In one embodiment, the amount is greater than or equal to about 5500 Hz. And, in another embodiment, the amount is greater than or equal to about 6000 Hz. In embodiments, the amount is less than or equal to about 10000 Hz, or more narrowly less than or equal to about 7000 Hz. A worker of ordinary skill in the art would recognize that the 60 s factor is to change revolutions per minute to Hertz, or revolutions per one second. For the purposes of this disclosure, the term “about” means ±3% of the respective quantity unless otherwise disclosed.
• The operational speed of the low pressure turbine as utilized in the formula should correspond to the engine operating conditions at each noise certification point currently defined inPart 36 or the Federal Airworthiness Regulations. More particularly, the rotational speed may be taken as an approach certification point as currently defined inPart 36 of the Federal Airworthiness Regulations. For purposes of this application and its claims, the term “approach speed” equates to this certification point. In other embodiments, the rotational speed is taken as a takeoff or cruise certification point, with the terms “takeoff speed” and “cruise speed” equating to these certification points. In some embodiments, the above formula results in a number that is less than or equal to about 10000 Hz at takeoff speed. In other embodiments, the above formula results in a number that is less than or equal to about 7000 Hz at approach speed.
• Although the above formula only needs to apply to one row of blades in thelow pressure turbine26, in one embodiment, all of the rows in the low pressure turbine meet the above formula. In some embodiments, the majority of the blade rows in the low pressure turbine meet the above formula, but some perhaps may not.
• This will result in operational noise to which human hearing will be less sensitive.
• In embodiments, it may be that the formula can result in a range of greater than or equal to 4000 Hz, and moving higher. Thus, by carefully designing the number of blades and controlling the operational speed of the low pressure turbine (and a worker of ordinary skill in the art would recognize how to control this speed) one can assure that the noise frequencies produced by the low pressure turbine are of less concern to humans.
• This invention is most applicable to jet engines rated to produce 15,000 pounds of thrust or more and with bypass ratios greater than about 8.0.
• FIG. 2 shows anembodiment200, wherein there is afan drive turbine208 driving ashaft206 to in turn drive afan rotor202. Agear reduction204 may be positioned between thefan drive turbine208 and thefan rotor202. Thisgear reduction204 may be structured and operate like the gear reduction disclosed above. Acompressor rotor210 is driven by anintermediate pressure turbine212, and a second stage compressor rotor214 is driven by aturbine rotor216. Acombustion section218 is positioned intermediate the compressor rotor214 and theturbine section216.
• FIG. 3 shows yet anotherembodiment300 wherein afan rotor302 and a first stage compressor304 rotate at a common speed. The gear reduction306 (which may be structured as disclosed above) is intermediate the compressor rotor304 and ashaft308 which is driven by a low pressure turbine section.
• Although an embodiment of this invention has been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of this invention. For that reason, the following claims should be studied to determine the true scope and content of this invention.

Claims (25)

1. A gas turbine engine comprising:

a fan and a turbine having a fan drive rotor, there also being a second turbine rotor;

a gear reduction effecting a reduction in a speed of said fan relative to an input speed from said fan drive rotor;

said fan drive rotor having a number of turbine blades in a majority of a plurality of blade rows of said fan drive rotor, and said turbine blades configured to operate at least some of the time at a rotational speed, and said number of turbine blades in said majority of said blade rows and said rotational speed being such that the following formula holds true for each row of said majority of said blade rows of the fan drive turbine:

(number of blades×speed)/60≧about 5500 Hz;

and

said rotational speed being in revolutions per minute.

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

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

4. The gas turbine engine as set forth inclaim 1, wherein the formula holds true for all of said blade rows of the fan drive rotor.

5. The gas turbine engine as set forth inclaim 1, wherein the formula does not hold true for all of said blade rows of the fan drive rotor.

6. The gas turbine engine as set forth inclaim 1, wherein the formula results in a number less than or equal to about 7000 Hz, said rotational speed being an approach speed.

7. The gas turbine engine as set forth inclaim 1, wherein the formula results in a number less than or equal to about 10000 Hz, said rotational speed being a takeoff speed.

8. The gas turbine engine as set forth inclaim 1, wherein said turbine section having a higher pressure turbine rotor and a lower pressure turbine rotor, with said fan drive rotor being said lower pressure turbine rotor and said second turbine rotor is said higher pressure turbine rotor.

9. The gas turbine engine as set forth inclaim 1, wherein there is a third turbine rotor, with said fan drive turbine being a most downstream of said three turbine rotors.

10. A method of designing a gas turbine engine comprising the steps of:

including a gear reduction between a fan drive turbine rotor and a fan, and selecting a number of blades in a majority of blade rows of the fan drive turbine rotor, in combination with a rotational speed of the fan drive turbine rotor, such that the following formula holds true for each row of said majority of blade rows of the fan drive turbine rotor:

(number of blades×speed)/60≧about 5500 Hz;

said rotational speed being in revolutions per minute; and

including a second turbine rotor.

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

12. The method of designing a gas turbine engine as set forth inclaim 11, wherein said gas turbine engine is rated to produce 15,000 pounds of thrust or more.

13. The method as set forth inclaim 10, wherein the formula holds true for all of the blade rows of the fan drive turbine.

14. The method as set forth inclaim 10, wherein the formula does not hold true for all of the blade rows of the fan drive turbine.

15. The method as set forth inclaim 10, wherein the formula results in a number less than or equal to about 7000 Hz, and said rotational speed is an approach speed.

16. The method as set forth inclaim 10, wherein the formula results in a number less than or equal to about 10000 Hz, and said rotational speed is a takeoff speed.

17. The method as set forth inclaim 10, wherein a 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.

18. A turbine module comprising:

a fan drive rotor having a plurality of blade rows each including a number of blades, a majority of the blade rows being capable of rotating at a rotational speed, so that when measuring said rotational speed in revolutions per minute:

(number of blades×said rotational speed)/60≧about 5500 Hz.

19. The turbine module as set forth inclaim 18, wherein the formula results in a number greater than or equal to about 6000.

20. The turbine module as set forth inclaim 18, wherein the formula does not hold true for all of the blade rows of the fan drive rotor.

21. The turbine module as set forth inclaim 18, wherein the formula holds true for all of the blade rows of the fan drive rotor.

22. The turbine module as set forth inclaim 21, wherein a pressure ratio across said fan drive rotor is greater than about 5:1.

23. The turbine module as set forth inclaim 21, wherein the formula results in a number less than or equal to about 7000 Hz, and the rotational speed is an approach speed.

24. The turbine module as set forth inclaim 21, wherein the formula results in a number less than or equal to about 10000 Hz, and the rotational speed is a takeoff speed.

25. The turbine module as set forth inclaim 18, wherein there being a higher pressure turbine rotor and a lower pressure turbine rotor, and said fan drive rotor being said lower pressure turbine rotor.

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