EP2809936B1 - Turbine à gaz dotée d'une efficience énergétique améliorée - Google Patents

Turbine à gaz dotée d'une efficience énergétique améliorée Download PDF

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Publication number
EP2809936B1
EP2809936B1 EP13775841.3A EP13775841A EP2809936B1 EP 2809936 B1 EP2809936 B1 EP 2809936B1 EP 13775841 A EP13775841 A EP 13775841A EP 2809936 B1 EP2809936 B1 EP 2809936B1
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EP
European Patent Office
Prior art keywords
fan
section
low pressure
guide vanes
exit guide
Prior art date
Legal status (The legal status 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 status listed.)
Active
Application number
EP13775841.3A
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German (de)
English (en)
Other versions
EP2809936A2 (fr
EP2809936A4 (fr
Inventor
Karl L. Hasel
Peter G. Smith
Stuart S. Ochs
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
RTX Corp
Original Assignee
United Technologies Corp
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Publication date
Priority claimed from US13/361,987 external-priority patent/US20120124964A1/en
Application filed by United Technologies Corp filed Critical United Technologies Corp
Publication of EP2809936A2 publication Critical patent/EP2809936A2/fr
Publication of EP2809936A4 publication Critical patent/EP2809936A4/fr
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Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D17/00—Regulating or controlling by varying flow
    • F01D17/10—Final actuators
    • F01D17/12—Final actuators arranged in stator parts
    • F01D17/14—Final actuators arranged in stator parts varying effective cross-sectional area of nozzles or guide conduits
    • F01D17/16—Final actuators arranged in stator parts varying effective cross-sectional area of nozzles or guide conduits by means of nozzle vanes
    • F01D17/162—Final actuators arranged in stator parts varying effective cross-sectional area of nozzles or guide conduits by means of nozzle vanes for axial flow, i.e. the vanes turning around axes which are essentially perpendicular to the rotor centre line
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D29/00—Details, component parts, or accessories
    • F04D29/40—Casings; Connections of working fluid
    • F04D29/52—Casings; Connections of working fluid for axial pumps
    • F04D29/54—Fluid-guiding means, e.g. diffusers
    • F04D29/56—Fluid-guiding means, e.g. diffusers adjustable
    • F04D29/563—Fluid-guiding means, e.g. diffusers adjustable specially adapted for elastic fluid pumps
    • 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
    • F05D2220/00—Application
    • F05D2220/30—Application in turbines
    • F05D2220/36—Application in turbines specially adapted for the fan of turbofan engines
    • 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
    • F05D2240/00—Components
    • F05D2240/10—Stators
    • F05D2240/12—Fluid guiding means, e.g. vanes
    • F05D2240/121—Fluid guiding means, e.g. vanes related to the leading edge of a stator vane

Definitions

  • the present application relates to a gas turbine engine having an improved fuel consumption based upon a combination of operational parameters.
  • Gas turbine engines typically include a fan which drives air into a bypass duct, and also into a compressor section. The air is compressed in the compressor section, and delivered into a combustor section where it is mixed with fuel and burned. Products of this combustion pass downstream over turbine rotors, driving the turbine rotors to rotate.
  • a low pressure turbine has rotated at a given speed, and driven a low pressure compressor, and the fan at the same rate of speed. More recently, gear reductions have been included such that the fan in a low pressure compressor can be driven at different speeds.
  • the present invention provides a gas turbine engine in accordance with claim 1.
  • the gear ratio is less than or equal to about 4.2.
  • the expansion ratio is greater than or equal to about 5.7.
  • the fan has an outer diameter that is greater than an outer diameter of the low pressure turbine section.
  • the present invention provides a method of operating a gas turbine engine in accordance with claim 5.
  • the gear reduction is less than or equal to 4.2.
  • the fan has an outer diameter that is greater than an outer diameter of the low pressure turbine section.
  • Figure 1 illustrates a general partial fragmentary schematic view of a gas turbofan engine 10 suspended from an engine pylon P within an engine nacelle assembly N as is typical of an aircraft designed for subsonic operation.
  • the turbofan engine 10 includes a core section within a core nacelle 12 that houses a low spool 14 and high spool 24.
  • the low spool 14 includes a low pressure compressor 16 and low pressure turbine 18.
  • the low spool 14 drives a fan section 20 directly or through a gear train 22.
  • the high spool 24 includes a high pressure compressor 26 and high pressure turbine 28.
  • a combustor 30 is arranged between the high pressure compressor 26 and high pressure turbine 28.
  • the low and high spools 14, 24 rotate about an engine axis of rotation A.
  • the engine 10 in the disclosed embodiment is a high-bypass geared turbofan aircraft engine in which the engine 10 bypass ratio is greater than ten (10), the turbofan diameter is significantly larger than that of the low pressure compressor 16, and the low pressure turbine 18 has a pressure, or expansion, ratio greater than five (5).
  • the gear train 22 may be an epicycle gear train such as a planetary gear system or other gear system with a gear reduction ratio of greater than 2.5. It should be understood, however, that the above parameters are exemplary of only one geared turbofan engine and that the present invention is likewise applicable to other gas turbine engines including direct drive turbofans.
  • the fan section 20 communicates airflow into the core nacelle 12 for compression by the low pressure compressor 16 and the high pressure compressor 26. Core airflow compressed by the low pressure compressor 16 and the high pressure compressor 26 is mixed with the fuel in the combustor 30 then expanded over the high pressure turbine 28 and low pressure turbine 18.
  • the turbines 28, 18 are coupled for rotation with respective spools 24, 14 to rotationally drive the compressors 26, 16 and, through the gear train 22, the fan section 20 in response to the expansion.
  • a core engine exhaust E exits the core nacelle 12 through a core nozzle 43 defined between the core nacelle 12 and a tail cone 32.
  • a bypass flow path 40 is defined between the core nacelle 12 and the fan nacelle 34.
  • the engine 10 generates a high bypass flow arrangement with a bypass ratio in which approximately 80 percent of the airflow entering the fan nacelle 34 becomes bypass flow B.
  • the bypass flow B communicates through the generally annular bypass flow path 40 and may be discharged from the engine 10 through a fan variable area nozzle (FVAN) 42 which defines a variable fan nozzle exit area 44 between the fan nacelle 34 and the core nacelle 12 at an aft segment 34S of the fan nacelle 34 downstream of the fan section 20.
  • FVAN fan variable area nozzle
  • the core nacelle 12 is generally supported upon a core engine case structure 46.
  • a fan case structure 48 is defined about the core engine case structure 46 to support the fan nacelle 34.
  • the core engine case structure 46 is secured to the fan case 48 through a multiple of circumferentially spaced radially extending fan exit guide vanes (FEGV) 50.
  • the fan case structure 48, the core engine case structure 46, and the multiple of circumferentially spaced radially extending fan exit guide vanes 50 which extend therebetween is typically a complete unit often referred to as an intermediate case. It should be understood that the fan exit guide vanes 50 may be of various forms.
  • the intermediate case structure in the disclosed embodiment includes a variable geometry fan exit guide vane (FEGV) system 36.
  • Thrust is a function of density, velocity, and area. One or more of these parameters can be manipulated to vary the amount and direction of thrust provided by the bypass flow B. A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio.
  • the fan section 20 of the engine 10 is nominally designed for a particular flight condition -- typically cruise at 0.8M and 35,000 feet (10,668 m).
  • the FEGV system 36 and/or the FVAN 42 is operated to adjust fan bypass air flow such that the angle of attack or incidence of the fan blades is maintained close to the design incidence for efficient engine operation at other flight conditions, such as landing and takeoff.
  • the FEGV system 36 and/or the FVAN 42 may be adjusted to selectively adjust the pressure ratio of the bypass flow B in response to a controller C. For example, increased mass flow during windmill or engine-out, and spoiling thrust at landing.
  • the FEGV system 36 will facilitate and in some instances replace the FVAN 42, such as, for example, variable flow area is utilized to manage and optimize the fan operating lines which provides operability margin and allows the fan to be operated near peak efficiency which enables a low fan pressure-ratio and low fan tip speed design; and the variable area reduces noise by improving fan blade aerodynamics by varying blade incidence.
  • the FEGV system 36 thereby provides optimized engine operation over a range of flight conditions with respect to performance and other operational parameters such as noise levels.
  • each fan exit guide vane 50 includes a respective airfoil portion 52 defined by an outer airfoil wall surface 54 between the leading edge 56 and a trailing edge 58.
  • the outer airfoil wall 54 typically has a generally concave shaped portion forming a pressure side and a generally convex shaped portion forming a suction side.
  • respective airfoil portion 52 defined by the outer airfoil wall surface 54 may be generally equivalent or separately tailored to optimize flow characteristics.
  • Each fan exit guide vane 50 is mounted about a vane longitudinal axis of rotation 60.
  • the vane axis of rotation 60 is typically transverse to the engine axis A, or at an angle to engine axis A.
  • various support struts 61 or other such members may be located through the airfoil portion 52 to provide fixed support structure between the core engine case structure 46 and the fan case structure 48.
  • the axis of rotation 60 may be located about the geometric center of gravity (CG) of the airfoil cross section.
  • An actuator system 62 illustrated schematically; Figure 1A ), for example only, a unison ring operates to rotate each fan exit guide vane 50 to selectively vary the fan nozzle throat area ( Figure 2B ).
  • the unison ring may be located, for example, in the intermediate case structure such as within either or both of the core engine case structure 46 or the fan case 48 ( Figure 1A ).
  • the FEGV system 36 communicates with the controller C to rotate the fan exit guide vanes 50 and effectively vary the fan nozzle exit area 44.
  • Other control systems including an engine controller or an aircraft flight control system may also be usable with the present invention.
  • Rotation of the fan exit guide vanes 50 between a nominal position and a rotated position selectively changes the fan bypass flow path 40. That is, both the throat area ( Figure 2B ) and the projected area ( Figure 2C ) are varied through adjustment of the fan exit guide vanes 50.
  • bypass flow B is increased for particular flight conditions such as during an engine-out condition.
  • engine bypass flow may be selectively vectored to provide, for example only, trim balance, thrust controlled maneuvering, enhanced ground operations and short field performance.
  • the FEGV system 36' includes a multiple of fan exit guide vane 50' which each includes a fixed airfoil portion 66F and pivoting airfoil portion 66P which pivots relative to the fixed airfoil portion 66F.
  • the pivoting airfoil portion 66P includes a leading edge flap which is actuatable by an actuator system 62' as described above to vary both the throat area ( Figure 3B ) and the projected area ( Figure 3C ).
  • the FEGV system 36" includes a multiple of slotted fan exit guide vane 50" which each includes a fixed airfoil portion 68F and pivoting and sliding airfoil portion 68P which pivots and slides relative to the fixed airfoil portion 68F to create a slot 70 vary both the throat area ( Figure 4B ) and the projected area ( Figure 4C ) as generally described above.
  • This slatted vane method not only increases the flow area but also provides the additional benefit that when there is a negative incidence on the fan exit guide vane 50" allows air flow from the high-pressure, convex side of the fan exit guide vane 50" to the lower-pressure, concave side of the fan exit guide vane 50" which delays flow separation.
  • the use of the gear reduction 22 allows control of a number of operational features in combination to achieve improved fuel efficiency.
  • the expansion ratio (or pressure ratio) across the low pressure turbine, which is the pressure entering the low pressure turbine section divided by the pressure leaving the low pressure turbine section is greater than or equal to about 5.0.
  • the bypass ratio is greater than 10.0.
  • the gear reduction ratio is greater than 2.5. In an embodiment, the gear reduction ratio is less than or equal to about 4.2.
  • This combination provides a low pressure turbine section that can be very compact, and sized for very high aerodynamic efficiency with a small number of stages (3 to 5, in accordance with the present invention). Further, the maximum diameter of these stages can be minimized to improve installation clearance under the wings of an aircraft.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
  • Control Of Turbines (AREA)

Claims (7)

  1. Turbine à gaz (10) comprenant :
    une section de noyau définie autour d'un axe (A) ;
    une section de ventilateur (20) montée au moins partiellement autour de ladite section de noyau pour définir un trajet d'écoulement de dérivation de ventilateur (40), dans laquelle un rapport de dérivation pour la turbine à gaz (10) qui compare l'air délivré par la section de ventilateur (20) dans un conduit de dérivation à la quantité d'air délivrée dans la section de noyau est supérieur à 10, un rapport de détente sur une section de turbine à basse pression (18) est supérieur à 5, et la section de turbine à basse pression (18) entraîne la section de ventilateur (20) par un réducteur à engrenages (22), le réducteur à engrenages (22) ayant un rapport supérieur à 2,5 ; et
    une pluralité d'aubes de guidage de sortie de ventilateur (50') en communication avec ledit trajet d'écoulement de dérivation de ventilateur (40),
    caractérisée en ce que :
    ladite pluralité d'aubes de guidage de sortie de ventilateur (50') peut tourner autour d'un axe de rotation (A) pour faire varier une zone de sortie de buse de ventilateur efficace pour ledit trajet d'écoulement de dérivation de ventilateur (40) ;
    ladite pluralité d'aubes de guidage de sortie de ventilateur (50') peut tourner indépendamment ;
    ladite pluralité d'aubes de guidage de sortie de ventilateur (50') peut tourner simultanément ;
    ladite pluralité d'aubes de guidage de sortie de ventilateur (50') est montée à l'intérieur d'une structure de carter de moteur intermédiaire (46, 48, 50) ; et
    chacune de ladite pluralité d'aubes de guidage de sortie de ventilateur (50') comprend une partie pivotante (66P) pouvant tourner autour dudit axe de rotation (A) par rapport à une partie fixe (66F), ladite partie pivotante (66P) comprenant un volet de bord d'attaque ;
    dans laquelle la section de turbine à basse pression (18) a 3 à 5 étages.
  2. Turbine à gaz selon la revendication 1, dans laquelle ledit rapport d'engrenage est inférieur ou égal à environ 4,2.
  3. Turbine à gaz selon la revendication 1 ou 2, dans laquelle ledit rapport de détente est supérieur ou égal à environ 5,7.
  4. Turbine à gaz selon une quelconque revendication précédente, dans laquelle ledit ventilateur (20) a un diamètre extérieur qui est supérieur à un diamètre extérieur de la section de turbine à basse pression (18).
  5. Procédé de fonctionnement d'une turbine à gaz (10) comprenant les étapes :
    d'entraînement d'un ventilateur (20) pour délivrer une première partie d'air dans un conduit de dérivation, et une seconde partie d'air dans un compresseur à basse pression (16), un rapport de dérivation de la première partie à la seconde partie étant supérieur ou égal à 8,0 ;
    la première partie de l'air étant délivrée dans le compresseur à basse pression (16), dans un compresseur à haute pression (26), puis dans une section de combustion (30), l'air étant mélangé avec du carburant et enflammé, et les produits de la combustion passant en aval sur une turbine à haute pression (28), puis une turbine à basse pression (18), la section de turbine à basse pression (18) fonctionnant avec un rapport de détente supérieur ou égal à 5,0 ; et
    ladite section de turbine à basse pression (18) étant entraînée en rotation et à son tour faisant tourner ledit compresseur à basse pression (16) et faisant tourner ledit ventilateur (20) par l'intermédiaire d'un réducteur à engrenages (22), ledit réducteur à engrenages (22) ayant un rapport supérieur ou égal à 2,5, dans lequel la turbine à gaz comprend une section de ventilateur (20) montée au moins partiellement autour d'une section de noyau pour définir un trajet d'écoulement de dérivation de ventilateur (40),
    caractérisé en ce que :
    ladite section de ventilateur (20) comprend une pluralité d'aubes de guidage de sortie de ventilateur (50') pouvant tourner autour d'un axe de rotation (A) pour faire varier une zone de sortie de buse de ventilateur efficace pour ledit trajet d'écoulement de dérivation de ventilateur (40) ;
    ladite pluralité d'aubes de guidage de sortie de ventilateur (50') peut tourner indépendamment ;
    ladite pluralité d'aubes de guidage de sortie de ventilateur (50') peut tourner simultanément ;
    ladite pluralité d'aubes de guidage de sortie de ventilateur (50') est montée à l'intérieur d'une structure de carter de moteur intermédiaire (46, 48, 50) ; et
    chacune de ladite pluralité d'aubes de guidage de sortie de ventilateur (50') comprend une partie pivotante (66P) pouvant tourner autour dudit axe de rotation (A) par rapport à une partie fixe (66F), ladite partie pivotante (66P) comprenant un volet de bord d'attaque ;
    dans lequel la section de turbine à basse pression (18) a 3 à 5 étages.
  6. Procédé selon la revendication 5, dans lequel ledit réducteur à engrenages (22) est inférieure ou égale à 4,2.
  7. Procédé selon la revendication 5 ou 6, dans lequel ledit ventilateur (20) a un diamètre extérieur qui est supérieur à un diamètre extérieur de la section de turbine à basse pression (18).
EP13775841.3A 2012-01-31 2013-01-17 Turbine à gaz dotée d'une efficience énergétique améliorée Active EP2809936B1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
US13/361,987 US20120124964A1 (en) 2007-07-27 2012-01-31 Gas turbine engine with improved fuel efficiency
PCT/US2013/021831 WO2013154639A2 (fr) 2012-01-31 2013-01-17 Turbine à gaz dotée d'une efficience énergétique améliorée

Publications (3)

Publication Number Publication Date
EP2809936A2 EP2809936A2 (fr) 2014-12-10
EP2809936A4 EP2809936A4 (fr) 2015-09-02
EP2809936B1 true EP2809936B1 (fr) 2019-08-14

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EP13775841.3A Active EP2809936B1 (fr) 2012-01-31 2013-01-17 Turbine à gaz dotée d'une efficience énergétique améliorée

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WO (1) WO2013154639A2 (fr)

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* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20230027726A1 (en) * 2021-07-19 2023-01-26 Raytheon Technologies Corporation High and low spool configuration for a gas turbine engine

Family Cites Families (4)

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Publication number Priority date Publication date Assignee Title
US5259187A (en) 1993-02-05 1993-11-09 General Electric Company Method of operating an aircraft bypass turbofan engine having variable fan outlet guide vanes
JP3912989B2 (ja) * 2001-01-25 2007-05-09 三菱重工業株式会社 ガスタービン
US20110120078A1 (en) * 2009-11-24 2011-05-26 Schwark Jr Fred W Variable area fan nozzle track
WO2013141930A1 (fr) * 2011-12-30 2013-09-26 United Technologies Corporation Turbomoteur à faible rapport de pression de soufflante

Non-Patent Citations (1)

* Cited by examiner, † Cited by third party
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Also Published As

Publication number Publication date
EP2809936A2 (fr) 2014-12-10
WO2013154639A3 (fr) 2014-01-03
EP2809936A4 (fr) 2015-09-02
WO2013154639A2 (fr) 2013-10-17

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