WO2015112231A2 - Réacteur à double flux à engrenages avec trois sections de turbine - Google Patents

Réacteur à double flux à engrenages avec trois sections de turbine Download PDF

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Publication number
WO2015112231A2
WO2015112231A2 PCT/US2014/064490 US2014064490W WO2015112231A2 WO 2015112231 A2 WO2015112231 A2 WO 2015112231A2 US 2014064490 W US2014064490 W US 2014064490W WO 2015112231 A2 WO2015112231 A2 WO 2015112231A2
Authority
WO
WIPO (PCT)
Prior art keywords
equal
rotor
ratio
gas turbine
set forth
Prior art date
Application number
PCT/US2014/064490
Other languages
English (en)
Other versions
WO2015112231A3 (fr
Inventor
Frederick M. Schwarz
Joseph Brent Staubach
Original Assignee
United Technologies Corporation
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by United Technologies Corporation filed Critical United Technologies Corporation
Priority to EP14880333.1A priority Critical patent/EP3084173A4/fr
Publication of WO2015112231A2 publication Critical patent/WO2015112231A2/fr
Publication of WO2015112231A3 publication Critical patent/WO2015112231A3/fr

Links

Classifications

    • 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
    • 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
    • F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
    • F02C7/36—Power transmission arrangements between the different shafts of the gas turbine plant, or between the gas-turbine plant and the power user
    • 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

Definitions

  • Gas turbine engines are known and typically include a fan delivering air as bypass flow for propulsion within a nacelle.
  • the fan often delivers core air flow into a core engine where it is delivered into an upstream compressor rotor.
  • the air may be compressed at the upstream compressor rotor and then delivered into a downstream compressor rotor where it may be further compressed to higher pressures.
  • the air from the downstream compressor rotor may be delivered into a combustion section where may be is mixed with fuel and ignited. Products of this combustion can pass downstream over turbine rotors driving them to rotate.
  • the turbine rotors in turn, can rotate the compressor rotors and the fan rotor.
  • a first turbine rotor drives the downstream or higher pressure compressor rotor.
  • An intermediate turbine rotor drives an upstream or lower pressure compressor rotor.
  • a third turbine rotor is a fan drive turbine and drives the fan rotor. All of these turbine rotors can have one stage or more than one stage, except for the fan drive turbine which for reasons of efficiency and in order to provide a reasonable fan tip speed have conventionally had many stages.
  • the fan drive turbine has driven the fan rotor at a common speed.
  • a gas turbine engine comprises a fan rotor configured to be driven by a fan drive turbine through a first shaft and a gear reduction.
  • the fan rotor is configured to deliver air into a bypass duct as bypass air and to deliver core air flow into a core engine where it reaches an upstream compressor rotor.
  • the upstream compressor rotor is configured to be driven through a second shaft by an intermediate turbine rotor.
  • a downstream compressor rotor is configured to be driven by an upstream turbine rotor through a third shaft.
  • An overall pressure ratio across the upstream and downstream compressor rotors is greater than or equal to about 35.0 and less than or equal to about 75.0.
  • a fan pressure ratio is greater than or equal to about 1.1 and less than or equal to about 1.6.
  • a pressure ratio across the upstream compressor rotor is greater than or equal to about 4.0 and less than or equal to about 12.0.
  • a pressure ratio across the downstream compressor rotor is greater than or equal to about 4.0 and less than or equal to about 8.0.
  • a gear ratio for the gear reduction is greater than or equal to about 2.6.
  • the intermediate turbine rotor has more than one stage.
  • the upstream compressor rotor has a first number of stages and the downstream compressor rotor has a second number of stages, and a ratio of the first number to the second number is greater than or equal to about 1.20.
  • a bypass ratio is defined as the volume of air delivered as bypass flow into the bypass duct compared to the volume of air delivered as core flow into the core engine.
  • the bypass ratio is greater than or equal to about 6.0.
  • a fan pressure ratio is greater than or equal to about 1.1 and less than or equal to about 1.6.
  • a pressure ratio across the downstream compressor rotor is greater than or equal to about 4.0 and less than or equal to about 8.0.
  • a gear ratio for the gear reduction is greater than or equal to about 2.6.
  • the intermediate turbine rotor has more than one stage.
  • the upstream compressor rotor has a first number of stages and the downstream compressor rotor has a second number of stages.
  • a ratio of the first number to the second number is greater than or equal to about 1.2.
  • a bypass ratio is defined as the volume of air delivered as bypass flow into the bypass duct compared to the volume of air delivered as core flow into the core engine.
  • the bypass ratio is greater than or equal to about 6.0.
  • a bypass ratio is defined as the volume of air delivered as bypass flow into the bypass duct compared to the volume of air delivered as core flow into the core engine.
  • the bypass ratio is greater than or equal to about 6.0.
  • the upstream compressor rotor has a first number of stages and the downstream compressor rotor has a second number of stages.
  • a ratio of the first number to the second number is greater than or equal to about 1.2.
  • the upstream compressor rotor has a first number of stages and the downstream compressor rotor has a second number of stages.
  • a ratio of the first number to the second number is greater than or equal to about 1.2.
  • the intermediate turbine rotor has more than one stage.
  • a gear ratio for the gear reduction is greater than or equal to about 2.6.
  • the upstream compressor rotor has a first number of stages and the downstream compressor rotor has a second number of stages. A ratio of the first number to the second number is greater than or equal to about 1.2.
  • Figure 1 schematically shows a gas turbine engine.
  • Gas turbine engine 120 is illustrated in Figure 1 having a fan rotor 122 driven through a gear reduction 124 by a fan drive turbine 128.
  • the fan drive turbine 128 drives a shaft 126 that drives the gear reduction 124 to, in turn, drive the fan rotor 122 at a reduced speed.
  • a gear ratio of the gear reduction is greater than or equal to about 2.6.
  • Fan rotor 122 delivers bypass air B within a nacelle 130 and core air C within a core housing 132.
  • the air in the core housing 132 enters into an upstream or lower pressure compressor rotor 134.
  • Air compressed by first or upstream compressor rotor 134 is then delivered into a second (higher pressure) or downstream compressor rotor 140.
  • the compressed air downstream of the compressor rotor 140 is delivered into a combustion section 146 where it is mixed with fuel and ignited.
  • Products of this combustion pass downstream over an upstream or high pressure turbine rotor 142.
  • a shaft 144 is driven by turbine rotor 142 to, in turn, drive the downstream compressor rotor 140.
  • Downstream of the turbine rotor 142 the products of combustion pass over a second (intermediate pressure) turbine rotor 136.
  • Turbine rotor 136 is shown driving a shaft 138 that is, in turn, connected to drive the upstream compressor rotor 134.
  • the products of combustion downstream of turbine rotor 136 pass over the fan drive turbine 128.
  • An overall pressure ratio is defined across the compressor rotors 134 and
  • the average fan pressure ratio may be lower than about 1.6 and, in certain embodiments, be between about 1.2 and about 1.45. Applicant has discovered that by shifting more of the work burden to the lower pressure compressor rotor 134, the speed of the downstream compressor rotor 140 may be reduced even though it is generally seen as highly desirable from an aerodynamic efficiency standpoint, and a compressor stability standpoint and in order to reduce the number of compressor stages, for a compressor to be designed with as much speed capability as possible. This is because even though there are benefits to increased compressor speed, there are also countering trends that can sharply reduce the gains from endlessly increasing the speed of the downstream compressor.
  • a slower speed in the last rotor of the downstream compressor allows an achievement of higher overall pressure ratio, by shifting speed (and stages) to the upstream compressor.
  • the speed of the downstream compressor rotor 140 has historically been a limit on overall achievable pressure ratio of about 50 for long range, twin aisle aircraft where the use of takeoff power is infrequent in the overall duty cycle of the engine.
  • regional jets single aisle aircraft
  • the use of takeoff power is more frequent so the OPR might be limited to 40 or even less owing to durability concerns for the upstream turbine stages.
  • an overall pressure ratio across the compressor rotors 134 and 140 may be greater than or equal to about 35.0 and less than or equal to about 75.
  • the lower pressure compressor rotor 134 may have a pressure ratio greater than or equal to about 4.0 and less than or equal to about 12.0.
  • the higher or downstream compressor rotor 140 may have a pressure ratio of greater than or equal to about 4.0 and less than or equal to about 8.0.
  • the intermediate turbine rotor 136 may have more than one stage if the pressure rise across the first compressor is selected to be high in relation to the second compressor in the ranges mentioned earlier.
  • the fan drive turbine should have at least three stages.
  • the upstream compressor rotor 134 may have at least about 1.20 times as many stages as the downstream compressor rotor 140.
  • the fan may have a fan pressure ratio greater than or equal to about 1.1 and less than or equal to about 1.6.
  • a bypass ratio may be defined as the ratio of the volume of air delivered as bypass flow B compared to the volume of air delivered as core air flow C.
  • the bypass ratio may be greater than or equal to about 6.0.
  • 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 about 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.
  • 'TSFC' Thrust Specific Fuel Consumption
  • 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 [(Tram °R) / (518.7 °R)] 0'5 .
  • the "Low corrected fan tip speed" as disclosed herein according to one non- limiting embodiment is less than about 1150 ft / second. In one example, the bypass ratio was greater than or equal to about 10.0.

Landscapes

  • 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)
  • Control Of Turbines (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Abstract

Moteur à turbine à gaz comprenant un rotor de soufflante conçu pour être entraîné par une turbine d'entraînement de soufflante par l'intermédiaire d'un premier arbre et d'une démultiplication. Le rotor de soufflante est conçu pour apporter de l'air dans un conduit de dérivation en tant qu'air de dérivation et pour apporter un écoulement d'air central dans un moteur central où il atteint un rotor de compresseur amont. Le rotor de compresseur amont est conçu pour être entraîné par l'intermédiaire d'un deuxième arbre par un rotor de turbine intermédiaire. Un rotor de compresseur en aval est conçu pour être entraîné par un rotor de turbine amont par l'intermédiaire d'un troisième arbre. Un rapport de pression globale sur les rotors de compresseur amont et aval est supérieur ou égal à environ 35,0 et inférieur ou égal à environ 75,0.
PCT/US2014/064490 2013-12-16 2014-11-07 Réacteur à double flux à engrenages avec trois sections de turbine WO2015112231A2 (fr)

Priority Applications (1)

Application Number Priority Date Filing Date Title
EP14880333.1A EP3084173A4 (fr) 2013-12-16 2014-11-07 Réacteur à double flux à engrenages avec trois sections de turbine

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
US201314107169A 2013-12-16 2013-12-16
US14/107,169 2013-12-16

Publications (2)

Publication Number Publication Date
WO2015112231A2 true WO2015112231A2 (fr) 2015-07-30
WO2015112231A3 WO2015112231A3 (fr) 2015-11-12

Family

ID=53682101

Family Applications (1)

Application Number Title Priority Date Filing Date
PCT/US2014/064490 WO2015112231A2 (fr) 2013-12-16 2014-11-07 Réacteur à double flux à engrenages avec trois sections de turbine

Country Status (2)

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EP (1) EP3084173A4 (fr)
WO (1) WO2015112231A2 (fr)

Family Cites Families (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP3128164B1 (fr) * 2006-08-22 2019-07-10 Rolls-Royce North American Technologies, Inc. Moteur à turbine à gaz avec accélérateur de vitesse intermédiaire
US8277174B2 (en) * 2007-09-21 2012-10-02 United Technologies Corporation Gas turbine engine compressor arrangement
US8337147B2 (en) * 2007-09-21 2012-12-25 United Technologies Corporation Gas turbine engine compressor arrangement
EP2592235A3 (fr) * 2011-11-11 2014-10-15 United Technologies Corporation Agencement de compresseur de turbine à gaz
US20130192258A1 (en) * 2012-01-31 2013-08-01 United Technologies Corporation Geared turbofan gas turbine engine architecture
US8424313B1 (en) * 2012-01-31 2013-04-23 United Technologies Corporation Gas turbine engine mid turbine frame with flow turning features
US20130318998A1 (en) * 2012-05-31 2013-12-05 Frederick M. Schwarz Geared turbofan with three turbines with high speed fan drive turbine

Non-Patent Citations (1)

* Cited by examiner, † Cited by third party
Title
See references of EP3084173A4 *

Also Published As

Publication number Publication date
EP3084173A2 (fr) 2016-10-26
WO2015112231A3 (fr) 2015-11-12
EP3084173A4 (fr) 2016-12-28

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