WO1996018538A1 - Montage pour moteur a turbine a gaz - Google Patents

Montage pour moteur a turbine a gaz Download PDF

Info

Publication number
WO1996018538A1
WO1996018538A1 PCT/US1995/012451 US9512451W WO9618538A1 WO 1996018538 A1 WO1996018538 A1 WO 1996018538A1 US 9512451 W US9512451 W US 9512451W WO 9618538 A1 WO9618538 A1 WO 9618538A1
Authority
WO
WIPO (PCT)
Prior art keywords
mount
gas turbine
turbine engine
thrust
onto
Prior art date
Application number
PCT/US1995/012451
Other languages
English (en)
Inventor
Paul W. Duesler
Constantino V. Loffredo
Stephen E. Potz
Stephen A. Sarcich
Jon A. Marx
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
Publication of WO1996018538A1 publication Critical patent/WO1996018538A1/fr

Links

Classifications

    • 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
    • B64D27/18—Aircraft characterised by the type or position of power plants of jet type within, or attached to, wings
    • 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
    • B64D27/40—Arrangements for mounting power plants in aircraft
    • B64D27/402—Arrangements for mounting power plants in aircraft comprising box like supporting frames, e.g. pylons or arrangements for embracing the power plant
    • 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
    • F02C7/20—Mounting or supporting of plant; Accommodating heat expansion or creep

Definitions

  • This invention relates to gas turbine engines and, more particularly, to mounting of the gas turbine engines onto an aircraft.
  • Conventional gas turbine engines include a core engine having a compressor section, a combustor, and a turbine section.
  • the sections of the gas turbine engine are enclosed in an engine case and are sequentially situated about a longitudinal axis.
  • a fan section is disposed forward of the compressor section.
  • the compressor section of many gas turbine engines frequently includes a low pressure compressor and a high pressure compressor. Both, the turbine section and the compressor section include alternating rows of rotating blades and stationary vanes. Air flows substantially axially through the sections of the gas turbine engine.
  • the air, compressed in the compressor is mixed with fuel which is burned in the combustor. The products of combustion are then expanded in the turbine, thereby rotating the turbine and driving the compressor.
  • a gas turbine engine is enclosed in a nacelle with the nacelle being disposed radially outward from the engine case and spaced apart therefrom.
  • a pylon is suspended from an aircraft wing to facilitate mounting of the gas turbine engine onto the aircraft.
  • the gas turbine engine attaches onto the pylon by means of a mounting arrangement that includes a front mount, a rear mount, and a thrust mount.
  • the front and rear mounts counteract vertical and lateral loads acting on the gas turbine engine, such as the engine's own weight and wind gusts acting on the nacelle. Also, either the front or rear mount counteracts the torque accompanying engine acceleration, deceleration or rotor seizure.
  • the thrust mount counteracts axial loads acting on the gas turbine engine.
  • the nacelle aerodynamic load includes an upwardly vertical force that becomes significant during takeoff and other maneuvers involving higher angles of attack.
  • Another force acting on the engine is the engine thrust.
  • the nacelle aerodynamic loading and the engine thrust result in axial flexing of the engine, also known in the art as backbone bending.
  • the backbone bending tends to distort the shape of the generally round engine case, a condition known as ovalization and translation.
  • the excessive backbone bending introduces interference between the inner surface of the engine case and the tips of the rotating blades. Such interference results in permanent erosion of the inner surface of the engine case and/or rotating blade tips, thereby permanently increasing a tip clearance therebetween.
  • the increased tip clearance adversely impacts the performance of the gas turbine engine. For example, in the compressor, increased tip clearance will result not only in a performance penalty, but it also may cause instability conditions leading to engine stall.
  • the newer generations of the gas turbine engines are more susceptible to the problem of backbone bending.
  • the newer engines tend to have longer fan blades and a longer and a more slender core engine to achieve a higher bypass ratio.
  • Each of these attributes of the newer engines contributes to additional backbone bending.
  • One feature of the newer engines that contributes to increased backbone bending is a larger diameter fan.
  • the larger diameter fan increases the aerodynamic load on the nacelle since the aerodynamic load on the nacelle is proportional to the square of the fan diameter.
  • Another contributing factor is that the longer front end of the core engine results in an increased distance between the aerodynamic load application location on the forward portion of the nacelle and the conventional front mount location. The increase in the distance results in a greater bending moment.
  • the smaller diameter of the core engine is the smaller diameter of the core engine.
  • the stiffness of the core is proportional to the cube of the diameter of the core.
  • reduction of the core diameter results in a significant reduction in the stiffness of the core.
  • the high pressure compressor is more sensitive to the backbone bending, ovalization and translation than the low pressure compressor.
  • the high pressure compressor typically has a smaller diameter than the low pressure compressor and therefore the high pressure compressor is significantly less stiff than the low pressure compressor since stiffness is proportional to the diameter cubed.
  • the sensitivity of the rotating blade to the tip clearance is inversely proportional to its length.
  • the high pressure compressor Since the high pressure compressor has shorter blades than the low pressure compressor, the high pressure compressor is more sensitive to large tip clearance.
  • the front mount attaches to the gas turbine engine between the low pressure compressor and high pressure compressor, thus incurring the largest bending moment at the high pressure compressor.
  • a gas turbine engine having a core engine enclosed in an engine case is secured onto a pylon by means of a front mount, a rear mount, and a thrust mount with the front mount and a forward end of the thrust mount being axially separated and with the thrust mount being oriented to point substantially at an intersection line between a vertical plane traversing the front mount and a horizontal plane passing through a longitudinal central axis of the gas turbine engine.
  • the front mount attaches onto the gas turbine engine at an inlet case disposed on the engine case aft of the fan section and forward of the low pressure compressor.
  • the forward end of the thrust mount link attaches onto an intermediate case disposed on the engine case between the low pressure compressor and the high pressure compressor.
  • the mounting system of the present invention significantly reduces the backbone bending in gas turbine engines.
  • the first feature of the present invention is that the axial separation of the front mount and the forward end of the thrust mount coupled with the specific orientation of the thrust mount has the capability of reducing the thrust induced bending moment from a prohibitive magnitude to zero.
  • the second feature of the present invention is that the positioning of the front mount forward of the low pressure compressor reduces backbone bending of the gas turbine engines due to the aerodynamic loading. The reduction of the backbone bending due to the aerodynamic loading is achieved since the distance between the load application forward of the fan and the front mount disposed aft the fan is reduced as compared to the distance in conventional engines.
  • the third feature of the present invention is that the positioning of the front mount forward of the low pressure compressor isolates the high pressure compressor from the maximum bending moment. The maximum bending moment occurs forward of the low pressure compressor, thereby reducing the effective bending moment acting on the high pressure compressor.
  • One advantage of the present invention is that the mounting arrangement utilizes the existing structure of the gas turbine engine, such as the inlet case and the intermediate case, without adding undesirable weight and cost to the gas turbine engine.
  • FIG. 1 is a schematic, cross-sectional elevation of a gas turbine engine secured onto a pylon by means of a front mount, thrust mount, and a rear mount, according to the present invention, with the pylon being secured onto an aircraft wing; and
  • FIG. 2 is an enlarged, partially broken away perspective view of the gas turbine engine mounted onto the pylon by means of the front mount, thrust mount, and rear mount of FIG. 1.
  • a gas turbine engine 10 includes a core engine 12 comprising a low pressure compressor 14, a high pressure compressor 15, a combustor 16, a high pressure turbine 17, and a low pressure turbine 18.
  • the sections 14-18 of the engine are disposed sequentially about a longitudinal axis 20 and are enclosed in an engine case 22 .
  • a fan section
  • the engine 24 is disposed forward of the low pressure compressor 1 .
  • the air 30 flows through the sections 24, 14-18 of the engine 10.
  • the engine case 22 provides an outer boundary for the air 30 flowing through the sections of the gas turbine engine 10.
  • the engine case 22 includes an inlet case 36, disposed forward of the low pressure compressor 14 and aft of the fan section 24 , an intermediate case 38 disposed between the low pressure compressor 14 and the high pressure compressor 15, and a turbine exhaust case 40 disposed aft the low pressure turbine 18.
  • the inlet case 36, intermediate case 38, and the turbine exhaust case 40 are generally stiff and fixedly support the core engine 12.
  • a torque box 41 surrounds the inlet case 36 and provides structural support to the gas turbine engine.
  • the torque box having a trailing edge 42, also serves as a base for fan frame vanes 43.
  • a nacelle 44 is disposed radially outward from the engine case 22 and spaced apart therefrom.
  • the gas turbine engine 10 mounts onto a pylon 45 to facilitate the attachment of the gas turbine engine 10 onto an aircraft 46.
  • the gas turbine engine 10 is secured onto the pylon 45 by means of a front mount
  • the front mount 48 includes three front mount links 48A, 48B, 48C, as shown in FIG. 2. One end of each front mount link 48A-C is secured to front ears 49A-C protruding from the trailing edge 42 of the torque box 41. The second end of each front mount link 48A-B is secured onto the pylon 45.
  • the rear mount 52 includes two rear mount links 52A, 52B that attach to rear ears 53A-B protruding from the turbine exhaust case 40 of the gas turbine engine 10.
  • a vertical plane Y traversing the gas turbine engine 10 passes through the front mount 48 so that it generally passes through a center point of all three front mount links 48A-C.
  • a horizontal plane X passes through the longitudinal center axis 20 and intersects the plane Y along an intersection line L.
  • the thrust mount 50 includes two thrust mount links 50A, 50B.
  • Each thrust mount link 50A, 50B has an elongated body 56 with a forward end 58 and an aft end 60.
  • the aft end 60 of the elongated body 56 of each thrust mount link 50A, 50B is secured onto the pylon 45 and the forward end 58 of each thrust mount link 50A, 50B attaches onto ears 61 A, B formed on the intermediate case 38 of the engine case 22.
  • the elongated body 56 of each thrust mount link 50A, 50B points onto the intersection line L, so that the line of action of each thrust mount link 50A, 50B, shown as a dotted line 62 on FIG. 1 , intersects the intersection line L of the aforementioned planes X, Y.
  • the aerodynamic load is designated by F-j and the engine thrust is designated by F2 on FIG. 1.
  • the front and rear mounts 48, 52 react to vertical and lateral loads as to well as to torque. Generally, either the front 48 or the rear 52 mount counteracts torque.
  • the front mount 48 includes the torque link 48C to react to torque.
  • the thrust mount links 50A, 50B react to axial loads, such as the thrust load and the longitudinal inertial loads, and to a lesser degree to vertical loads.
  • Axial separation of the front mount 48 and the forward end 58 of the thrust mount 50 coupled with the specific orientation of the thrust mount links 50A, B pointing onto the intersection line L reduces backbone bending in the gas turbine engine 10 caused by the aerodynamic loading and the engine thrust acting on the gas turbine engine.
  • a number of factors contribute to reduction of backbone bending in the gas turbine engine having the mounting arrangement of the present invention.
  • the specific orientation of the thrust mount links 50A, B eliminates or significantly reduces the bending moment due to the engine thrust. This reduction in bending moment due to the engine thrust is the result of the thrust mount links, directly or substantially, pointing onto the intersection line L.
  • is relatively small since the distance between the load application at the forward portion of the nacelle 44 and the front mount 48 is reduced as compared to the prior art. Since bending moment is equal to the product of the force and distance, as is well known in the art, the magnitude of the bending moment is decreased as the distance is reduced. Third, since the front mount 48 is placed forward of the low pressure compressor 14 , the high pressure compressor 15 is shielded from incurring maximum bending moment. This represents a very significant improvement over the prior art. Since the high pressure compressor is more sensitive to the bending moment than the low pressure compressor, a decrease in the bending moment at the high pressure compressor is crucial.
  • the high pressure compressor is much less stiff than the low pressure compressor.
  • Another reason for the high pressure compressor being more sensitive to the bending moment is that the high pressure compressor includes shorter blades than the low pressure compressor. It is well known in the art that shorter blades are more sensitive to tip leakage between the tips thereof and the inner surface of the engine case 22. Therefore, it is a great advantage of the present invention to isolate the high pressure compressor from the maximum bending moment.
  • One major advantage of the present invention is that existing structure of the engine case 22 is utilized for mounting the gas turbine engine onto the pylon while significantly reducing backbone bending in the gas turbine engine. Since the existing structure, such as the inlet case 36 and the intermediate case 38, was used, no significant weight was added to the gas turbine engine.
  • the front mount 48 may include two links 48A, 48B rather than three links 48A-C, and the rear mount 52 may include an additional link to counteract torque. Other configurations of links may be employed as well.
  • the preferred embodiment is described as having a low pressure compressor, high pressure compressor and a high pressure turbine and a low pressure turbine, the present invention can be applicable to other types of gas turbine engines as long as the front mount is axially separated from the forward end of the thrust mount and the thrust mount links point onto the intersection line L.
  • the action line 62 can be slightly offset from the intersection line L.
  • the offset will result in a slightly greater bending moment due to the engine thrust, it is still within the scope of this invention to minimize the offset.

Landscapes

  • Engineering & Computer Science (AREA)
  • Aviation & Aerospace Engineering (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)

Abstract

L'invention porte sur un montage, destiné à fixer un moteur à turbine à gaz (10) sur un pylône (45), comportant un bâti frontal (48), un bâti de poussée (50) et un bâti arrière (52). Le bâti frontal (48) est situé en avant d'une extrémité avant du bâti de poussée (50) qui décrit un angle tel qu'il se trouve orienté vers une ligne l'intersection (L) entre un plan vertical (Y), traversant le moteur (10) de la turbine et passant à travers le bâti frontal (48), et un plan horizontal (X) passant à travers un axe longitudinal central (20). Le montage réalisé au titre de la présente invention permet de réduire notablement le moment du couple de flexion résultant de la charge aérodynamique et de la poussée du moteur.
PCT/US1995/012451 1994-12-12 1995-09-28 Montage pour moteur a turbine a gaz WO1996018538A1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
US35439794A 1994-12-12 1994-12-12
US354,397 1994-12-12

Publications (1)

Publication Number Publication Date
WO1996018538A1 true WO1996018538A1 (fr) 1996-06-20

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ID=23393161

Family Applications (1)

Application Number Title Priority Date Filing Date
PCT/US1995/012451 WO1996018538A1 (fr) 1994-12-12 1995-09-28 Montage pour moteur a turbine a gaz

Country Status (1)

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WO (1) WO1996018538A1 (fr)

Cited By (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US6494403B2 (en) * 2000-03-22 2002-12-17 Eads Aribus Sa Device for aircraft thrust recovery capable of linking a turboshaft engine and an engine strut
EP1571082A1 (fr) * 2004-03-04 2005-09-07 Airbus France Système de montage interposé entre un moteur d'aéronef et une structure rigide d'un mât d'accrochage fixé sous une voilure de cet aéronef
FR2925016A1 (fr) * 2007-12-12 2009-06-19 Snecma Sa Suspension d'un turboreacteur a un aeronef
US20140326842A1 (en) * 2013-03-21 2014-11-06 United Technologies Corporation Oil tank mount with stiffeners
EP3561277A3 (fr) * 2018-04-06 2020-01-01 Rolls-Royce plc Boîtier

Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB1504290A (en) * 1975-08-01 1978-03-15 Rolls Royce Gas turbine engine mounting systems
GB2010969A (en) * 1977-12-22 1979-07-04 Rolls Royce Mounting for Gas Turbine Jet Propulsion Engine
GB2044358A (en) * 1979-03-10 1980-10-15 Rolls Royce Gas turbine jet engine mounting
EP0115914A1 (fr) * 1983-01-12 1984-08-15 British Aerospace Public Limited Company Disposition du montage de groupes moteurs pour ailes d'avion

Patent Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB1504290A (en) * 1975-08-01 1978-03-15 Rolls Royce Gas turbine engine mounting systems
GB2010969A (en) * 1977-12-22 1979-07-04 Rolls Royce Mounting for Gas Turbine Jet Propulsion Engine
GB2044358A (en) * 1979-03-10 1980-10-15 Rolls Royce Gas turbine jet engine mounting
EP0115914A1 (fr) * 1983-01-12 1984-08-15 British Aerospace Public Limited Company Disposition du montage de groupes moteurs pour ailes d'avion

Cited By (10)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US6494403B2 (en) * 2000-03-22 2002-12-17 Eads Aribus Sa Device for aircraft thrust recovery capable of linking a turboshaft engine and an engine strut
EP1571082A1 (fr) * 2004-03-04 2005-09-07 Airbus France Système de montage interposé entre un moteur d'aéronef et une structure rigide d'un mât d'accrochage fixé sous une voilure de cet aéronef
FR2867157A1 (fr) * 2004-03-04 2005-09-09 Airbus France Systeme de montage interpose entre un moteur d'aeronef et une structure rigide d'un mat d'accrochage fixe sous une voilure de cet aeronef.
US7156343B2 (en) 2004-03-04 2007-01-02 Airbus France Mounting system inserted between an aircraft engine and a rigid structure of an attachment strut fixed under a wing of this aircraft
FR2925016A1 (fr) * 2007-12-12 2009-06-19 Snecma Sa Suspension d'un turboreacteur a un aeronef
US20140326842A1 (en) * 2013-03-21 2014-11-06 United Technologies Corporation Oil tank mount with stiffeners
US9863324B2 (en) * 2013-03-21 2018-01-09 United Technologies Corporation Oil tank mount with stiffeners
EP3561277A3 (fr) * 2018-04-06 2020-01-01 Rolls-Royce plc Boîtier
US11156167B2 (en) 2018-04-06 2021-10-26 Rolls-Royce Plc Casing
US11525407B2 (en) 2018-04-06 2022-12-13 Rolls-Royce Plc Casing

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