EP1013894A2 - Revêtement pour la virole d'une soufflante - Google Patents

Revêtement pour la virole d'une soufflante Download PDF

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Publication number
EP1013894A2
EP1013894A2 EP99310555A EP99310555A EP1013894A2 EP 1013894 A2 EP1013894 A2 EP 1013894A2 EP 99310555 A EP99310555 A EP 99310555A EP 99310555 A EP99310555 A EP 99310555A EP 1013894 A2 EP1013894 A2 EP 1013894A2
Authority
EP
European Patent Office
Prior art keywords
fan
fan case
rotor
engine
case
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.)
Granted
Application number
EP99310555A
Other languages
German (de)
English (en)
Other versions
EP1013894A3 (fr
EP1013894B1 (fr
Inventor
Keven G. Van Duyn
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
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 Corp filed Critical United Technologies Corp
Publication of EP1013894A2 publication Critical patent/EP1013894A2/fr
Publication of EP1013894A3 publication Critical patent/EP1013894A3/fr
Application granted granted Critical
Publication of EP1013894B1 publication Critical patent/EP1013894B1/fr
Anticipated expiration legal-status Critical
Expired - Lifetime legal-status Critical Current

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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
    • F01D21/00—Shutting-down of machines or engines, e.g. in emergency; Regulating, controlling, or safety means not otherwise provided for
    • F01D21/04—Shutting-down of machines or engines, e.g. in emergency; Regulating, controlling, or safety means not otherwise provided for responsive to undesired position of rotor relative to stator or to breaking-off of a part of the rotor, e.g. indicating such position
    • 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
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/08—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator
    • F01D11/12—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator using a rubstrip, e.g. erodible. deformable or resiliently-biased part
    • 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
    • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
    • F01D25/24—Casings; Casing parts, e.g. diaphragms, casing fastenings

Definitions

  • the present invention relates to gas turbine engines, and more particularly, to providing a hardened liner in the fan case of the engine to reduce damage to the fan case and support structure in the event of a high rotor imbalance condition, such as a fan blade loss.
  • a gas turbine engine such as a turbofan engine for an aircraft, includes a fan section, a compression section, a combustion section, and a turbine section. An axis of the engine is centrally disposed within the engine, and extends longitudinally through these sections. A primary flow path for working medium gases extends axially through the engine. A secondary flow path for working medium gases extends parallel to and radially outward of the primary flow path.
  • the fan draws air into the engine.
  • the fan raises the pressure of the air drawn along the secondary flow path, thus producing useful thrust.
  • the air drawn along the primary flow path into the compressor section is compressed.
  • the compressed air is channeled to the combustor section, where fuel is added to the compressed air, and the air-fuel mixture is burned.
  • the products of combustion are discharged to the turbine section.
  • the turbine section extracts work from these products to power the fan and compressor. Any energy from the products of combustion not needed to drive the fan and compressor contributes to useful thrust.
  • the fan section includes a rotor assembly and a stator assembly.
  • the rotor assembly of the fan includes a rotor disk and a plurality of outwardly extending rotor blades.
  • Each rotor blade includes an airfoil portion, a root portion, and a tip portion.
  • the airfoil portion extends through the flow path and interacts with the working medium gases to transfer energy between the rotor blade and working medium gases.
  • the stator assembly includes a fan containment case assembly, which circumscribes the rotor assembly in close proximity to the tips of the rotor blades.
  • This invention is in part predicated on the recognition that by constraining the interaction of fan blade tips and the fan case to a predetermined radial zone in which is disposed hardened structure, there is a decrease of the loads transmitted to the interfaces of the engine by approximately the same percentage of the loads transmitted to the interfaces of the aircraft, and will allow an additional factor of safety during an abnormal imbalance condition of the rotor assembly.
  • a fan case in a gas turbine engine has a radial zone of interaction bounded outwardly by an inwardly facing surface, preferably a hardened metallic surface, of the fan case, the zone being a clearance which is less than one hundredth of the fan case diameter measured from the blade tips in a non-operative, zero speed engine condition with the rotor centered, a hardened structure disposed in the zone, such that during a high rotor imbalance condition, the blade tips skid on the hardened structure and reduce the destructive cutting away of the fan case, and reduce torque and imbalance loads transmitted to the interface of the engine and the aircraft.
  • the optimal radial zone of clearance is defined as a constant approximately five thousandths (0.005) of the fan case diameter.
  • the lower limit of the radial zone of clearance is defined as a constant approximately two and one half thousandths (0.0025) of the fan case diameter, below which fan blades would destroy themselves due to high interaction loads between the fan blades and the fan case.
  • the structural clearance lies in a range of 0.20 inches to 1.25 inches (5 mm to 32 mm) for corresponding jet engine fan case diameters which lie in a range of 20 inches to 120 inches (0.5 m to 3.05 m).
  • the hardened structure or material is preferably a liner which provides a skid-surface for the blades to circumferentially glide on and thus minimizes torque loading of the fan case. Further, the fan case structure limits the deflection of the rotor shaft during a fan blade loss event.
  • the liner of the present invention comprises shingles of hardened material.
  • a primary advantage of the present invention is the minimization of damage to the fan case thus, resulting in a durable fan case in the event of a fan blade loss.
  • the hardened fan case liner of the present invention reduces the destructive cutting away of the fan case by the fan blades.
  • a further advantage of the fan case of the present invention is its ability to provide an appropriate restraining structure to the deflection of the rotor shaft during a fan blade loss event.
  • the hardened liner reduces frictional forces and therefore, the torque transmitted from the rotor to the engine cases.
  • Another advantage is the ease and cost of manufacturing and incorporating into the fan case the liner of the present invention. The simplicity of the structure of the liner and the use of economic materials, allows for cost effective manufacturing processes. Further, fan cases of the prior art can be retrofitted to include the present invention in a cost-effective manner.
  • FIG. 1 is a perspective view of a typical axial flow, turbofan gas turbine engine.
  • FIG. 2 is a perspective view of the rotor assembly of the prior art showing a released fan blade.
  • FIG. 3 is a cross-sectional schematic representation of the fan containment case assembly a fan case of the present invention taken along the lines 3-3 of FIG. 2.
  • the fan case liner 42 is made from hardened material such as from alloys of stainless steel or nickel.
  • the nickel alloy Inconel 718, or stainless steel alloys, such as AISI 321 or AISI 347, are examples of alloys that can be used to manufacture the liner.
  • the liner is thus manufactured from material that is harder than the fan blade tip material which is typically titanium. For ease of installation, the liner could be manufactured as arced segments, which can then be bonded to the fan case.
  • the radial zone of interaction 60 of the present invention is a clearance bounded inwardly by the blade tips 38 in a non-operative, zero speed engine condition with the rotor centered about the engine centerline and the blades in their engaged position with the rotor.
  • the radial zone of interaction 60 is bounded outwardly by the hardened inner surface 53 of the fan case 48.
  • the radial zone of interaction is referred to hereinafter as the structural clearance.
  • the hardened liner 42 is disposed in the radial zone of interaction.
  • the structural clearance 60 is less than one hundredth of the fan case diameter.
  • the optimal structural clearance measured from the fan blade tips is about five thousandths (0.005) of the fan case diameter.
  • the normalized engine interface loads are plotted versus a ratio of the structural clearance to the fan case diameter for a typical modern gas turbine engine.
  • the normalization of the engine interface loads is based on a typical structural clearance of one inch (1") (25 mm).
  • the curve shown in FIG. 7 is representative of loads at different engine to aircraft interfaces and is dependent on several factors some of which are the weight of the fan case and related hardware attached to the fan case such as a nacelle, the fan case stiffness relative to the engine, the ratio of the weight of the combination of the fan and blades to the weight of the fan case, and the dynamics of the rotor such as the frequency of the rotor.
  • the interface loads cannot be reduced beyond the normalized value of about 0.5 due to the structural characteristics of the fan case, i.e., a heavier fan case would be required to increase the transmission of loads to the fan case thereby reducing rotor deflections.
  • the fan case interacts more closely with the fan blade tips and as such the fan case constrains the deflection of the imbalanced rotor by inertial resistance.
  • there is a decrease in the amplitude of the rotor deflections which results in the decrease of the forces or loads transmitted through the bearing support structure.
  • the kinetic energy associated with the imbalance of the rotor is transmitted through the fan blade tips into the fan case and is largely dissipated by the translational (radial) movement of the fan case.
  • a portion of the kinetic energy associated with the imbalance of the rotor is dissipated by the movement of the fan blades relative to the fan case.
  • the associated heat generated due to the frictional forces between the fan blade tips and the fan case is dissipated in the materials of the fan case and blade structure.
  • the blade loss produces an imbalance in the rotor and causes the rotor to deflect radially outwardly in close proximity to the fan case.
  • the separation between the fan blades and the inner surface of the fan case is minimized in modern engines to decrease the radial deflection of the rotor assembly. Due to the rotor deflection and the reduced clearance between the fan blades and the fan case, the fan blade tips rapidly cut away the compliant rub strip 44 in the innermost surface of the fan containment case assembly.
  • the thin, fan case liner made from hardened materials such as steel or nickel alloys, provides a skid surface for the relatively softer blades. The fan blades move circumferentially along on the skid-surface of the liner.
  • the shingled embodiment also provides a skid-surface for the fan blades to circumferentially rotate upon.
  • the damage to the liner after a fan blade loss event is limited to the loss of one or more adjacent shingles 52.
  • the remaining shingles continue to provide an effective skid-surface for the fan blades to glide on.
  • the diameter of the fan case can be an inner, outer or mean diameter of the fan case in the region of the fan blades.
  • a primary advantage of the present invention fan case liner is the minimization of damage to the fan case thus, resulting in a durable fan case in the event of a fan blade loss.
  • the liner reduces the destructive cutting away of the fan case by the fan blades.
  • a further advantage of the present invention fan case is its ability to provide an appropriate restraining structure to the deflection of the rotor shaft during a fan blade loss event.
  • the liner reduces frictional forces, and as a result, reduces torque loads transmitted from the fan rotor to the case.
  • Another advantage is the ease and cost of manufacturing and incorporating the hardened fan case liner of the present invention.
  • the simplicity of the structure of the liner and the use of economical materials allows for cost effective manufacturing processes.
  • current, prior art fan cases can be retrofitted to include the fan case liner in a cost effective manner. By incorporating the present invention liner, current engines limit damage to the fan containment case assembly and to the rotor shaft.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
EP99310555A 1998-12-23 1999-12-23 Moteur à turbine à gaz Expired - Lifetime EP1013894B1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
US22054498A 1998-12-23 1998-12-23
US220544 1998-12-23

Publications (3)

Publication Number Publication Date
EP1013894A2 true EP1013894A2 (fr) 2000-06-28
EP1013894A3 EP1013894A3 (fr) 2002-01-09
EP1013894B1 EP1013894B1 (fr) 2003-06-04

Family

ID=22823963

Family Applications (1)

Application Number Title Priority Date Filing Date
EP99310555A Expired - Lifetime EP1013894B1 (fr) 1998-12-23 1999-12-23 Moteur à turbine à gaz

Country Status (3)

Country Link
EP (1) EP1013894B1 (fr)
JP (1) JP2000220472A (fr)
DE (1) DE69908540T2 (fr)

Cited By (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP1475516A1 (fr) * 2003-05-02 2004-11-10 General Electric Company Système et procédé de réglage élastique du jeu d'extrémité d'une turbine à haute pression
EP1589195A1 (fr) * 2004-04-20 2005-10-26 ROLLS-ROYCE plc Dispositif de rétention des aubes pour turbine à gaz
GB2494137A (en) * 2011-08-31 2013-03-06 Rolls Royce Plc Rotor casing lining comprising multiple sections
WO2013154642A1 (fr) * 2012-01-31 2013-10-17 United Technologies Corporation Ensemble support d'architecture à engrenages de turbomachine
WO2015076882A3 (fr) * 2013-09-09 2015-07-30 United Technologies Corporation Carters de ventilateur et procédés de fabrication

Families Citing this family (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR2935017B1 (fr) * 2008-08-13 2012-11-02 Snecma Paroi interne d'une nacelle de turbomachine
DE102015204893B3 (de) * 2015-03-18 2016-06-09 MTU Aero Engines AG Schutzeinrichtung für eine Strömungsmaschine

Family Cites Families (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4547122A (en) * 1983-10-14 1985-10-15 Aeronautical Research Associates Of Princeton, Inc. Method of containing fractured turbine blade fragments
US5403148A (en) * 1993-09-07 1995-04-04 General Electric Company Ballistic barrier for turbomachinery blade containment
JPH11200813A (ja) * 1997-11-11 1999-07-27 United Technol Corp <Utc> ガスタービンエンジン

Non-Patent Citations (1)

* Cited by examiner, † Cited by third party
Title
None

Cited By (10)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP1475516A1 (fr) * 2003-05-02 2004-11-10 General Electric Company Système et procédé de réglage élastique du jeu d'extrémité d'une turbine à haute pression
EP1589195A1 (fr) * 2004-04-20 2005-10-26 ROLLS-ROYCE plc Dispositif de rétention des aubes pour turbine à gaz
US7402022B2 (en) 2004-04-20 2008-07-22 Rolls-Royce Plc Rotor blade containment assembly for a gas turbine engine
GB2494137A (en) * 2011-08-31 2013-03-06 Rolls Royce Plc Rotor casing lining comprising multiple sections
US9097114B2 (en) 2011-08-31 2015-08-04 Rolls-Royce Plc Rotor casing liner
GB2494137B (en) * 2011-08-31 2016-02-17 Rolls Royce Plc A rotor casing liner comprising multiple sections
WO2013154642A1 (fr) * 2012-01-31 2013-10-17 United Technologies Corporation Ensemble support d'architecture à engrenages de turbomachine
US8961113B2 (en) 2012-01-31 2015-02-24 United Technologies Corporation Turbomachine geared architecture support assembly
WO2015076882A3 (fr) * 2013-09-09 2015-07-30 United Technologies Corporation Carters de ventilateur et procédés de fabrication
US10221718B2 (en) 2013-09-09 2019-03-05 United Technologies Corporation Fan cases and manufacture methods

Also Published As

Publication number Publication date
DE69908540D1 (de) 2003-07-10
JP2000220472A (ja) 2000-08-08
EP1013894A3 (fr) 2002-01-09
EP1013894B1 (fr) 2003-06-04
DE69908540T2 (de) 2003-12-18

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