EP3387244A1 - Turbine à gaz d'aéronef à tuyère de sortie variable d'un canal de flux secondaire - Google Patents

Turbine à gaz d'aéronef à tuyère de sortie variable d'un canal de flux secondaire

Info

Publication number
EP3387244A1
EP3387244A1 EP16805124.1A EP16805124A EP3387244A1 EP 3387244 A1 EP3387244 A1 EP 3387244A1 EP 16805124 A EP16805124 A EP 16805124A EP 3387244 A1 EP3387244 A1 EP 3387244A1
Authority
EP
European Patent Office
Prior art keywords
gas turbine
aircraft gas
outlet nozzle
ring element
core engine
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.)
Withdrawn
Application number
EP16805124.1A
Other languages
German (de)
English (en)
Inventor
Frank Uwe KÖPF
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.)
Rolls Royce Deutschland Ltd and Co KG
Original Assignee
Rolls Royce Deutschland Ltd and Co KG
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 Rolls Royce Deutschland Ltd and Co KG filed Critical Rolls Royce Deutschland Ltd and Co KG
Publication of EP3387244A1 publication Critical patent/EP3387244A1/fr
Withdrawn legal-status Critical Current

Links

Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02K—JET-PROPULSION PLANTS
    • F02K1/00—Plants characterised by the form or arrangement of the jet pipe or nozzle; Jet pipes or nozzles peculiar thereto
    • F02K1/06—Varying effective area of jet pipe or nozzle
    • F02K1/08—Varying effective area of jet pipe or nozzle by axially moving or transversely deforming an internal member, e.g. the exhaust cone
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
    • B64D27/00—Arrangement or mounting of power plants in aircraft; Aircraft characterised by the type or position of power plants
    • B64D27/02—Aircraft characterised by the type or position of power plants
    • B64D27/16—Aircraft characterised by the type or position of power plants of jet type
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
    • B64D31/00—Power plant control systems; Arrangement of power plant control systems in aircraft
    • B64D31/02—Initiating means
    • B64D31/06—Initiating means actuated automatically
    • 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
    • F05D2220/32—Application in turbines in gas turbines
    • F05D2220/323—Application in turbines in gas turbines for aircraft propulsion, e.g. jet engines

Definitions

  • the invention relates to an aircraft gas turbine according to the features of the preamble of claim 1.
  • the invention relates to an aircraft gas turbine with a variable outlet nozzle of a bypass channel.
  • the bypass duct surrounds the core engine as known in the art.
  • Variable outlet nozzles of bypass ducts are particularly required in aircraft gas turbines with high bypass rates in order to optimize the efficiency of the fan.
  • By changing the effective exit area of the outlet nozzle of the operating point of the fan can be adjusted so that there are favorable pressure conditions, which take into account the surge limit of the fan.
  • adjustable outlet nozzles are previously known.
  • the US 2009/0208328 A1 and US 8,850,824 B2 show constructions in which in the region of the outlet nozzle elements are arranged on the lining of the core engine, which can be curved.
  • US 4,043,508 A shows a solution in which a multi-membered flap mechanism is used.
  • three flaps are connected in series pivotally to each other, which can be pivoted to achieve different exit surfaces in different positions.
  • Around the circumference of the outlet nozzle a plurality of such flap arrangements are provided.
  • a further measure for changing the exit face of the outlet nozzle is shown in US Pat. Nos. US 2010/0043394 A1 and US Pat. No. 3,598,318 A. Individual flaps distributed around the circumference are provided, which are pivoted into the bypass channel at different flight conditions.
  • the cross-section of the outlet nozzle can also be influenced according to US 2009/0067993 A1 by displacing an outer end region of the lining of the bypass channel in the axial direction.
  • the invention has for its object to provide an aircraft gas turbine of the type mentioned above, which allows a simple design and simple, cost manufacturability effective and flow-optimized adjustment of the outlet nozzle of the bypass channel.
  • a ring element is arranged, which is displaceable in the axial direction, wherein between the lining of the core engine and the ring member is formed by the displacement of the ring member variable annular channel.
  • a preferably aerodynamically formed ring is used, which is axially displaceable as a function of the respective flight conditions or operating conditions of the aircraft gas turbine.
  • the ring element is formed so that opens between the ring member and the outer lining of the core engine, an annular channel through which a portion of the flow of the bypass channel is passed. It is provided in a preferred embodiment of the invention that the annular channel opens when the ring member is moved in the axial direction to the rear.
  • the term "axial direction" refers to the engine axis
  • the annular channel by complete displacement of the ring element completely closed at the front.
  • the effective outlet cross section of the outlet nozzle can be adapted to the operating conditions of the aircraft gas turbine in a simple manner and be adapted to the respective operating point of the fan, especially in aircraft gas turbines with a high bypass ratio.
  • the displacement of the ring element is not limited to certain displacement positions, but rather it is possible to steplessly bring the ring element into any displacement position.
  • the ring element according to the invention opens up the possibility of optimizing the flow conditions in the region of the outlet nozzle of the bypass channel in comparison with the constructions known from the prior art. Since the ring element extends around the entire circumference of the outlet nozzle, uniform flow conditions result around the entire circumference. This is not possible with the flap solutions known from the prior art, in which individual flaps are distributed separately around the circumference.
  • a further, essential advantage of the invention is also that the mechanism for displacing the ring element can preferably be arranged and integrated on the core engine or the radially outer lining of the core engine such that the flow of the bypass channel itself is not disturbed. It is particularly advantageous if the ring element is displaceable by means of electrical or hydraulic actuators. There are thus no lever designs or the like required, as shown in the prior art.
  • the ring element is aerodynamically designed and optimized in its cross section, so that there is a minimal pressure loss in the bypass channel. This also leads to an increase in the efficiency in the respective positions of the ring element.
  • the ring element can be designed such that it merely opens or closes the additional annular channel during its axial displacement, while the outflow surface of the original outlet nozzle remains unchanged.
  • the cross-sectional area of the original outlet nozzle is understood to be the cross-section which results radially outside the ring element between the ring element and the outer housing wall.
  • the above-mentioned change or enlargement of effective exit area of the outlet nozzle thus comprises the effective area of the additionally provided annular channel plus the exit area of the actual, original outlet nozzle.
  • the effective cross-sectional area thus results from an addition of the cross-sectional area of the additionally openable annular channel.
  • the control of the displacement of the ring element can be done automatically by the electronic engine control, so that the respective engine conditions, such as maximum thrust during takeoff, end of the climb and cruise, are automatically taken into account.
  • an additional oil cooler is arranged in the annular channel. This is installed, for example, on the lining of the core engine. By opening or closing the additional annular channel, the amount of air is determined, which is passed through the oil cooler.
  • FIG. 1 is a schematic representation of a gas turbine engine according to the present invention
  • Fig. 2 is an enlarged detail view of an embodiment in a first
  • Fig. 3 is a representation, analogous to FIG. 2, in an operating position to the end of
  • Fig. 4 is a representation in an operating position in cruising flight.
  • the gas turbine engine 10 of FIG. 1 is a generally illustrated example of a turbomachine to which the invention may find application.
  • the engine 10 is formed in a conventional manner and comprises one behind the other in the flow direction Air inlet 1 1, a circulating in a housing fan 12, a medium-pressure compressor 13, a high-pressure compressor 14, a combustion chamber 15, a high-pressure turbine 16, a medium-pressure turbine 17 and a low-pressure turbine 18 and an exhaust nozzle 19, all of which are arranged around a central engine axis 1.
  • the intermediate pressure compressor 13 and the high pressure compressor 14 each include a plurality of stages, each of which includes a circumferentially extending array of fixed stationary vanes 20, commonly referred to as stator vanes, that radially inwardly from the core engine housing 21 into an annular flow passage through the compressors 13, 14 protrude.
  • the compressors further include an array of compressor blades 22 projecting radially outwardly from a rotatable drum or disc 26 coupled to hubs 27 of high pressure turbine 16 and mid pressure turbine 17, respectively.
  • the turbine sections 16, 17, 18 have similar stages, comprising an array of fixed vanes 23 projecting radially inward from the housing 21 into the annular flow passage through the turbines 16, 17, 18, and a downstream array of turbine rotor blades 24 projecting outwardly from a rotatable hub 27.
  • the compressor drum or compressor disk 26 and the blades 22 disposed thereon and the turbine rotor hub 27 and the turbine rotor blades 24 disposed thereon rotate about the engine axis 1 during operation.
  • FIG. 1 shows, in the case of an aircraft gas turbine only shown schematically, that a bypass duct 25 is formed between an outer housing wall 30 and a panel 29 of the core engine 10.
  • the air flow conveyed by the fan 12 flows through the auxiliary passage 25 and exits through an exhaust nozzle 31, which is also referred to as a cold exhaust nozzle, as opposed to a hot exhaust nozzle 35 of the core engine.
  • FIG. 1 shows in a highly simplified, schematic representation the arrangement and positioning of a ring element 32 according to the invention.
  • FIGS. 2 to 4 each show enlarged and more precise detail views of the ring element 32 according to the invention.
  • This ring is designed as an aerodynamically designed and flow-optimized ring which preferably extends around the entire circumference of the aircraft gas turbine.
  • FIGS. 2 to 4 each show an end region of the outer housing wall 30 and an end region of the lining 29 of the core engine.
  • a portion of the outlet cone 28 is shown.
  • the outlet nozzle 35 of the core engine results.
  • the arrows each show the direction of flow.
  • the reference numeral 36 shows the cross-section of the outlet nozzle 31 in a simplified form.
  • This exit surface of the cross section 36 forms the actual outlet nozzle 31, which can remain unchanged during a displacement of the ring element 32 according to the invention.
  • the arrow shows the flow through the bypass duct 25.
  • FIG 2 shows an operating state in which the ring element 32 according to the invention is displaced maximally to the rear, relative to the throughflow direction of the aircraft gas turbine.
  • an annular channel 33 opens between the surface of the lining 29 of the core engine 10 and the ring element 32.
  • An oil cooler 34 can be arranged in the annular channel 33.
  • FIG. 2 shows an operating position in which the effective total area of the outlet nozzle 31 is increased in addition to the cross-section 36 by the cross-sectional area of the annular channel 33. This can be done to increase the total area of 10%.
  • This position is provided at maximum take-off power (max take-off). Due to the higher overall effective cross-sectional area of the operating point of the fan 12 can be lowered, so that there is a greater overall performance of the aircraft gas turbine.
  • FIG. 4 shows an operating state in cruise, in which the effective total area of the outlet nozzle 31 is determined by a partial opening of the annular channel 33 in such a way that a desired state is reached at which no change takes place. It should be pointed out again at this point that the effective total area of the outlet nozzle 31 results from the respective effective outflow surface of the annular channel 33 and the cross-sectional area 36 of the outlet nozzle 31 in the region of the bypass channel 25.
  • the invention is not limited to the embodiment shown, but rather arise within the scope of the invention varied modification and modification options. These may relate both to the drive of the ring element, not shown in detail, as well as the cross-sectional design and aerodynamic design of the ring member 32 and the associated wall of the lining 29 of the core engine. ?

Landscapes

  • Engineering & Computer Science (AREA)
  • Aviation & Aerospace Engineering (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Control Of Turbines (AREA)

Abstract

L'invention concerne une turbine à gaz d'aéronef comprenant un groupe propulseur central (10) et un canal de flux secondaire (25) entourant celui-ci, lequel forme, avec un habillage (29) du groupe propulseur central (10) et une paroi de boîtier (30) extérieure dans le sens radial, une tuyère de sortie (3). L'invention est caractérisée en ce qu'un élément en anneau (32) est disposé dans la zone de la tuyère de sortie (31), lequel peut coulisser dans le sens axial. Un canal annulaire (33) variable par coulissement de l'élément en anneau (32) est formé entre l'habillage (29) du groupe propulseur central (10) et l'élément en anneau (32).
EP16805124.1A 2015-12-09 2016-12-01 Turbine à gaz d'aéronef à tuyère de sortie variable d'un canal de flux secondaire Withdrawn EP3387244A1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
DE102015224701.5A DE102015224701A1 (de) 2015-12-09 2015-12-09 Fluggasturbine mit variabler Austrittsdüse eines Nebenstromkanals
PCT/EP2016/079484 WO2017097665A1 (fr) 2015-12-09 2016-12-01 Turbine à gaz d'aéronef à tuyère de sortie variable d'un canal de flux secondaire

Publications (1)

Publication Number Publication Date
EP3387244A1 true EP3387244A1 (fr) 2018-10-17

Family

ID=57460525

Family Applications (1)

Application Number Title Priority Date Filing Date
EP16805124.1A Withdrawn EP3387244A1 (fr) 2015-12-09 2016-12-01 Turbine à gaz d'aéronef à tuyère de sortie variable d'un canal de flux secondaire

Country Status (4)

Country Link
US (1) US20180266361A1 (fr)
EP (1) EP3387244A1 (fr)
DE (1) DE102015224701A1 (fr)
WO (1) WO2017097665A1 (fr)

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US8877851B2 (en) 2012-05-02 2014-11-04 E I Du Pont De Nemours And Company Graphite filled polyester compositions

Families Citing this family (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE102017104036A1 (de) * 2017-02-27 2018-08-30 Rolls-Royce Deutschland Ltd & Co Kg Konvergent-divergente Schubdüse für ein Turbofan-Triebwerk eines Überschallflugzeugs und Verfahren zur Einstellung der Düsenhalsfläche in einer Schubdüse eines Turbofan-Triebwerks

Family Cites Families (11)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3598318A (en) 1970-04-10 1971-08-10 Boeing Co Movable acoustic splitter for nozzle area control and thrust reversal
US4043508A (en) 1975-12-01 1977-08-23 General Electric Company Articulated plug nozzle
EP2069629B1 (fr) 2006-10-12 2014-09-24 United Technologies Corporation Tuyère à section variable d'une soufflante de turboréacteur double flux équipée d'un système à pièces rapportées pivotantes
US7681399B2 (en) * 2006-11-14 2010-03-23 General Electric Company Turbofan engine cowl assembly and method of operating the same
US7966828B2 (en) 2007-01-08 2011-06-28 United Technologies Corporation Variable area nozzle with woven sleeve extension
EP1944475B1 (fr) * 2007-01-08 2015-08-12 United Technologies Corporation Système d'échange de chaleur
US20090067993A1 (en) 2007-03-22 2009-03-12 Roberge Gary D Coated variable area fan nozzle
US9010126B2 (en) 2008-02-20 2015-04-21 United Technologies Corporation Gas turbine engine with variable area fan nozzle bladder system
US8961114B2 (en) * 2010-11-22 2015-02-24 General Electric Company Integrated variable geometry flow restrictor and heat exchanger
DE102011106959A1 (de) 2011-07-08 2013-01-10 Rolls-Royce Deutschland Ltd & Co Kg Fluggasturbine mit variabler Nebenstromdüse
US9828943B2 (en) * 2014-06-09 2017-11-28 United Technologies Corporation Variable area nozzle for gas turbine engine

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US8877851B2 (en) 2012-05-02 2014-11-04 E I Du Pont De Nemours And Company Graphite filled polyester compositions

Also Published As

Publication number Publication date
US20180266361A1 (en) 2018-09-20
DE102015224701A1 (de) 2017-06-14
WO2017097665A1 (fr) 2017-06-15

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