Patent Application
20160153279
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.
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.
The engine 20 generally includes a low speed spool 30 and a high speed spool 32 mounted for rotation about an engine central longitudinal axis A relative to an engine static structure 36 via several bearing systems 38. It should be understood that various bearing systems 38 at various locations may alternatively or additionally be provided.
The low speed spool 30 generally includes an inner shaft 40 that interconnects a fan 42, a low pressure compressor 44 and a low pressure turbine 46. The inner shaft 40 is connected to the fan 42 through a geared architecture 48 to drive the fan 42 at a lower speed than the low speed spool 30. The high speed spool 32 includes an outer shaft 50 that interconnects a high pressure compressor 52 and high pressure turbine 54. A combustor 56 is arranged between the high pressure compressor 52 and the high pressure turbine 54. A mid-turbine frame 57 of the engine static structure 36 is arranged generally between the high pressure turbine 54 and the low pressure turbine 46. The mid-turbine frame 57 further supports bearing systems 38 in the turbine section 28. The inner shaft 40 and the outer shaft 50 are concentric and rotate via bearing systems 38 about the engine central longitudinal axis A which is collinear with their longitudinal axes.
The core airflow is compressed by the low pressure compressor 44 then the high pressure compressor 52, mixed and burned with fuel in the combustor 56, then expanded over the high pressure turbine 54 and low pressure turbine 46. The mid-turbine frame 57 includes airfoils 59 which are in the core airflow path. The turbines 46, 54 rotationally drive the respective low speed spool 30 and high speed spool 32 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.
The engine 20 in one example is a high-bypass geared aircraft engine. In a further example, the engine 20 bypass ratio is greater than about six (6), with an example embodiment being greater than ten (10), the geared architecture 48 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, the engine 20 bypass ratio is greater than about ten (10:1), the fan diameter is significantly larger than that of the low pressure compressor 44, and the low pressure turbine 46 has a pressure ratio that is greater than about 5:1. The low pressure turbine 46 pressure ratio is a ratio of the pressure measured at inlet of low pressure turbine 46 to the pressure at the outlet of the low pressure turbine 46 (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. The fan section 22 of the engine 20 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 in Part 36 or the Federal Airworthiness Regulations. More particularly, the rotational speed may be taken as an approach certification point as currently defined in Part 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 the low pressure turbine 26, 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.
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.
This application is a continuation of U.S. patent application Ser. No. 14/996,544 filed Jan. 15, 2016, which is a continuation-in-part of U.S. patent application Ser. No. 14/795,931, 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/US2013/020724 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.
| Number | Date | Country | |
|---|---|---|---|
| 61592643 | Jan 2012 | US | |
| 61884660 | Sep 2013 | US |
| Number | Date | Country | |
|---|---|---|---|
| Parent | 14996544 | Jan 2016 | US |
| Child | 15007784 | US |
| Number | Date | Country | |
|---|---|---|---|
| Parent | 14795931 | Jul 2015 | US |
| Child | 14996544 | US | |
| Parent | 14248386 | Apr 2014 | US |
| Child | 14795931 | US | |
| Parent | PCT/US2013/020724 | Jan 2013 | US |
| Child | 14248386 | US |
