WO2013122713A2 - Low noise compressor rotor for geared turbofan engine - Google Patents
Low noise compressor rotor for geared turbofan engine Download PDFInfo
- Publication number
- WO2013122713A2 WO2013122713A2 PCT/US2013/022035 US2013022035W WO2013122713A2 WO 2013122713 A2 WO2013122713 A2 WO 2013122713A2 US 2013022035 W US2013022035 W US 2013022035W WO 2013122713 A2 WO2013122713 A2 WO 2013122713A2
- Authority
- WO
- WIPO (PCT)
- Prior art keywords
- low pressure
- set forth
- gas turbine
- compressor
- turbine engine
- Prior art date
Links
Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C9/00—Controlling gas-turbine plants; Controlling fuel supply in air- breathing jet-propulsion plants
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C9/00—Controlling gas-turbine plants; Controlling fuel supply in air- breathing jet-propulsion plants
- F02C9/26—Control of fuel supply
- F02C9/28—Regulating systems responsive to plant or ambient parameters, e.g. temperature, pressure, rotor speed
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K3/00—Plants including a gas turbine driving a compressor or a ducted fan
- F02K3/02—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber
- F02K3/04—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type
- F02K3/06—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type with front fan
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2260/00—Function
- F05D2260/40—Transmission of power
- F05D2260/403—Transmission of power through the shape of the drive components
- F05D2260/4031—Transmission of power through the shape of the drive components as in toothed gearing
- F05D2260/40311—Transmission of power through the shape of the drive components as in toothed gearing of the epicyclical, planetary or differential type
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2260/00—Function
- F05D2260/96—Preventing, counteracting or reducing vibration or noise
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2270/00—Control
- F05D2270/01—Purpose of the control system
- F05D2270/02—Purpose of the control system to control rotational speed (n)
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2270/00—Control
- F05D2270/01—Purpose of the control system
- F05D2270/02—Purpose of the control system to control rotational speed (n)
- F05D2270/023—Purpose of the control system to control rotational speed (n) of different spools or shafts
Definitions
- This application relates to the design of a gas turbine engine rotor which can be operated to produce noise that is less sensitive to human hearing.
- Gas turbine engines 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.
- each of the turbine rotors include a number of rows of turbine blades which rotate with the rotor. Interspersed between the rows of turbine blades are vanes.
- the high pressure turbine rotor has typically driven a high pressure compressor rotor
- the low pressure turbine rotor has typically driven a low pressure compressor rotor.
- Each of the compressor rotors also include a number of compressor blades which rotate with the rotors. There are also vanes interspersed between the rows of compressor blades.
- the low pressure turbine or compressor 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 rotors, the low pressure compressor rotors, and their harmonics.
- the noise can often be in a frequency range that is very sensitive to humans.
- a vane-to-blade ratio has been controlled to be above a certain number.
- 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 "cut-off.”
- 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 distinct speeds.
- a gas turbine engine has a fan, a compressor section having a low pressure portion and a high pressure portion, a combustor section, and a turbine having a low pressure portion.
- the low pressure turbine portion drives the low pressure compressor portion and the fan.
- a gear reduction effects a reduction in the speed of the fan relative to a speed of the low pressure turbine and the low pressure compressor portion.
- At least one of the low pressure turbine portion and low pressure compressor portion has a number of blades in each of a plurality of rows. The blades operate at least some of the time at a rotational speed.
- the number of blades and the rotational speed are such that the following formula holds true for at least one of the blade rows of the at least one of the low pressure turbine portion and/or the low pressure compressor sections: (number of blades x rotational speed)/60 > 5500.
- the rotational speed is an approach speed in revolutions per minute.
- the formula results in a number greater than or equal to 6000 Hz.
- the gas turbine engine is rated to produce 15,000 pounds of thrust or more.
- the at least one of the low pressure turbine portion and the low pressure compressor portion is the low pressure compressor portion.
- the formula results in a number greater than or equal to 6000.
- the gas turbine engine is rated to produce 15,000 pounds of thrust or more.
- the gear reduction has a gear ratio of greater than about 2.3.
- the gear reduction has a gear ratio of greater than about 2.5.
- the fan delivers air into a bypass duct, and a portion of air into the compressor section.
- a bypass ratio is defined as the volume of air delivered into the bypass duct compared to the volume of air delivered into the compressor section. The bypass ratio is greater than about 6.
- the bypass ratio is greater than about 10.
- the gear reduction has a gear ratio of greater than about 2.3.
- the fan delivers air into a bypass duct, and a portion of air into the compressor section.
- a bypass ratio is defined as the volume of air delivered into the bypass duct compared to the volume of air delivered into the compressor section. The bypass ratio is greater than about 6.
- the bypass ratio is greater than about 10.
- a compressor module has a low pressure portion having a number of blades in each of a plurality of rows of the low pressure portion.
- the blades operate at least some of the time at a rotational speed.
- the number of blades and the rotational speed are such that the following formula holds true for at least one of the blade rows of the low pressure portion: (number of blades x rotational speed)/60 > 5500 Hz.
- the rotational speed is an approach speed in revolutions per minute.
- the formula results in a number greater than or equal to 6000 Hz.
- the gas turbine engine is rated to produce 15,000 pounds of thrust or more.
- Figure 1 shows a gas turbine engine.
- FIG. 1 schematically illustrates a gas turbine engine 20.
- the gas turbine engine 20 is disclosed herein as a two-spool turbofan that generally incorporates a fan section 22, a compressor section 24, a combustor section 26 and a turbine section 28.
- Alternative engines might include an augmentor section (not shown), or an intermediate spool, among other systems or features.
- the fan section 22 drives air along a bypass flowpath B in a bypass duct defined within a nacelle 15, while the compressor section 24 drives air along a core flowpath C for compression and communication into the combustor section 26 then expansion through the turbine section 28.
- FIG. 1 schematically illustrates a gas turbine engine 20.
- the gas turbine engine 20 is disclosed herein as a two-spool turbofan that generally incorporates a fan section 22, a compressor section 24, a combustor section 26 and a turbine section 28.
- Alternative engines might include an augmentor section (not shown), or an intermediate spool, among other systems or features.
- the fan section 22 drives air along
- 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 engine 20 in one example is a high-bypass geared aircraft engine.
- 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 planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3 and the low pressure turbine 46 has a pressure ratio that is greater than about 5.
- 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
- the low pressure turbine 46 has a pressure ratio that is greater than about 5: 1.
- Low pressure turbine 46 pressure ratio is pressure measured prior to inlet of low pressure turbine 46 as related to the pressure at the outlet of the low pressure turbine 46 prior to an exhaust nozzle.
- the geared architecture 48 may be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.5: 1. 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.
- 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 1.45.
- 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)] ⁇ ° 5 .
- the "Low corrected fan tip speed” as disclosed herein according to one non-limiting embodiment is less than about 1150 ft / second.
- the maximum fan speed is at its outer tip, and in larger engines, the fan diameter is much larger than it may be in smaller power engines.
- the use of the gear reduction has freed the designer from limitation on the low pressure turbine and compressor speeds caused by a desire to not have unduly high fan speeds.
- the number of rotating blades in any low pressure turbine stage, multiplied by the rotational speed of the low pressure turbine 46 (in revolutions per minute), divided by 60 should be greater than or equal to 5500. The same holds true for the low pressure compressor stages. More narrowly, the amounts should be above 6000. A worker of ordinary skill in the art would recognize that the 60 factor is to change revolutions per minute to Hertz, or revolutions per one second.
- the operational speed of the low pressure turbine 46 and low pressure compressor 44 as utilized in the formula should correspond to the engine operating conditions at each noise certification point defined in Part 36 or the Federal Airworthiness Regulations. More particularly, the rotational speed may be taken as an approach certification point as 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.
- the low pressure turbine 46 may meet the above formula. However, this application may also extend to low pressure turbines wherein only one of the blade rows in the low pressure turbine meet the above formula. In other embodiments, plural rows meet the formula and in other embodiments, the majority of the rows meet the formula. The same is true for low pressure compressors, wherein all of the rows in the low pressure compressor 44 may meet the above formula. However, the application may extend to low pressure compressors wherein only one of the blade rows in the low pressure compressor meet the above formula. In other embodiments, plural rows meet the formula and in other embodiments, the majority of the rows meet the formula. [0045] This will result in operational noise that would be less sensitive to human hearing.
- the formula can result in a range of greater than or equal to 5500, and moving higher.
- the number of blades and controlling the operational speed of the low pressure turbine 46 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.
- it may be only the low pressure turbine rotor 46, or the low pressure compressor rotor 44 which is designed to meet the meet the above formula. On the other hand, it is also possible to ensure that both the low pressure turbine 46 and low pressure compressor 44 meet the above formula.
- This invention is most applicable to jet engines rated to produce 15,000 pounds of thrust or more. In this thrust range, prior art jet engines have typically had frequency ranges of about 4000 hertz. Thus, the noise problems as mentioned above have existed.
- Lower thrust engines ( ⁇ 15,000 pounds) may have operated under conditions that sometimes passed above the 4000 number, and even approached 6000, however, this has not been in combination with the geared architecture, nor in the higher powered engines which have the larger fans, and thus the greater limitations on low pressure turbine or low pressure compressor speed.
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- Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Structures Of Non-Positive Displacement Pumps (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
Description
Claims
Priority Applications (4)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
CA2863620A CA2863620C (en) | 2012-01-31 | 2013-01-18 | Low noise compressor rotor for geared turbofan engine |
CN201380007480.8A CN104246132B (en) | 2012-01-31 | 2013-01-18 | For the low-noise compressor rotor of gear transmission turbofan engine |
JP2014549009A JP5898337B2 (en) | 2012-01-31 | 2013-01-18 | Low noise compressor rotor of geared turbofan engine |
EP13749721.0A EP2809881B1 (en) | 2012-01-31 | 2013-01-18 | Low noise compressor rotor for geared turbofan engine |
Applications Claiming Priority (7)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
US201261592643P | 2012-01-31 | 2012-01-31 | |
US61/592,643 | 2012-01-31 | ||
US13/363,154 | 2012-01-31 | ||
US13/403,005 US8246292B1 (en) | 2012-01-31 | 2012-02-23 | Low noise turbine for geared turbofan engine |
US61/619,124 | 2012-04-02 | ||
US13/446,510 | 2012-04-13 | ||
US13/590,328 US8517668B1 (en) | 2012-01-31 | 2012-08-21 | Low noise turbine for geared turbofan engine |
Publications (3)
Publication Number | Publication Date |
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WO2013122713A2 true WO2013122713A2 (en) | 2013-08-22 |
WO2013122713A3 WO2013122713A3 (en) | 2014-05-30 |
WO2013122713A8 WO2013122713A8 (en) | 2014-06-19 |
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Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
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PCT/US2013/022035 WO2013122713A2 (en) | 2012-01-31 | 2013-01-18 | Low noise compressor rotor for geared turbofan engine |
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Cited By (7)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
WO2015048214A1 (en) * | 2013-09-30 | 2015-04-02 | United Technologies Corporation | Low noise turbine for geared turbofan engine |
JP2015137649A (en) * | 2014-01-21 | 2015-07-30 | ユナイテッド テクノロジーズ コーポレイションUnited Technologies Corporation | Gas turbine engine and gas turbine engine design method |
JP2016125500A (en) * | 2015-01-08 | 2016-07-11 | ユナイテッド テクノロジーズ コーポレイションUnited Technologies Corporation | Low noise compressor rotor for geared turbofan engine |
US9624834B2 (en) | 2012-09-28 | 2017-04-18 | United Technologies Corporation | Low noise compressor rotor for geared turbofan engine |
US9650965B2 (en) | 2012-09-28 | 2017-05-16 | United Technologies Corporation | Low noise compressor and turbine for geared turbofan engine |
US11143109B2 (en) | 2013-03-14 | 2021-10-12 | Raytheon Technologies Corporation | Low noise turbine for geared gas turbine engine |
US11719161B2 (en) | 2013-03-14 | 2023-08-08 | Raytheon Technologies Corporation | Low noise turbine for geared gas turbine engine |
Family Cites Families (6)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US2957655A (en) * | 1950-06-01 | 1960-10-25 | Curtiss Wright Corp | Turbine propeller control system |
US2850226A (en) * | 1954-09-28 | 1958-09-02 | Curtiss Wright Corp | Gas turbine engine with air flow modulating means |
WO2000019082A2 (en) * | 1998-08-17 | 2000-04-06 | Ramgen Power Systems, Inc. | Ramjet engine with axial air supply fan |
WO2008063154A2 (en) * | 2006-10-12 | 2008-05-29 | United Technologies Corporation | Fan variable area nozzle with adaptive structure |
US8167540B2 (en) * | 2008-01-30 | 2012-05-01 | Hamilton Sundstrand Corporation | System for reducing compressor noise |
US20100192595A1 (en) * | 2009-01-30 | 2010-08-05 | Robert Joseph Orlando | Gas turbine engine assembly and methods of assembling same |
-
2013
- 2013-01-18 WO PCT/US2013/022035 patent/WO2013122713A2/en active Application Filing
Non-Patent Citations (2)
Title |
---|
None |
See also references of EP2809881A2 |
Cited By (13)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US9733266B2 (en) | 2012-09-28 | 2017-08-15 | United Technologies Corporation | Low noise compressor and turbine for geared turbofan engine |
US9624834B2 (en) | 2012-09-28 | 2017-04-18 | United Technologies Corporation | Low noise compressor rotor for geared turbofan engine |
US9650965B2 (en) | 2012-09-28 | 2017-05-16 | United Technologies Corporation | Low noise compressor and turbine for geared turbofan engine |
US9726019B2 (en) | 2012-09-28 | 2017-08-08 | United Technologies Corporation | Low noise compressor rotor for geared turbofan engine |
US11143109B2 (en) | 2013-03-14 | 2021-10-12 | Raytheon Technologies Corporation | Low noise turbine for geared gas turbine engine |
US11168614B2 (en) | 2013-03-14 | 2021-11-09 | Raytheon Technologies Corporation | Low noise turbine for geared gas turbine engine |
US11560849B2 (en) | 2013-03-14 | 2023-01-24 | Raytheon Technologies Corporation | Low noise turbine for geared gas turbine engine |
US11719161B2 (en) | 2013-03-14 | 2023-08-08 | Raytheon Technologies Corporation | Low noise turbine for geared gas turbine engine |
WO2015048214A1 (en) * | 2013-09-30 | 2015-04-02 | United Technologies Corporation | Low noise turbine for geared turbofan engine |
JP2015137649A (en) * | 2014-01-21 | 2015-07-30 | ユナイテッド テクノロジーズ コーポレイションUnited Technologies Corporation | Gas turbine engine and gas turbine engine design method |
JP2017198218A (en) * | 2014-01-21 | 2017-11-02 | ユナイテッド テクノロジーズ コーポレイションUnited Technologies Corporation | Gas-turbine engine, and its design method |
RU2656171C2 (en) * | 2014-01-21 | 2018-05-31 | Юнайтед Текнолоджиз Корпорейшн | Low-noise compressor rotor for reducer turbo-fan engine |
JP2016125500A (en) * | 2015-01-08 | 2016-07-11 | ユナイテッド テクノロジーズ コーポレイションUnited Technologies Corporation | Low noise compressor rotor for geared turbofan engine |
Also Published As
Publication number | Publication date |
---|---|
WO2013122713A8 (en) | 2014-06-19 |
WO2013122713A3 (en) | 2014-05-30 |
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