US20150285081A1 - Gas turbine engine component with flow separating rib - Google Patents

Gas turbine engine component with flow separating rib Download PDF

Info

Publication number
US20150285081A1
US20150285081A1 US14/665,170 US201514665170A US2015285081A1 US 20150285081 A1 US20150285081 A1 US 20150285081A1 US 201514665170 A US201514665170 A US 201514665170A US 2015285081 A1 US2015285081 A1 US 2015285081A1
Authority
US
United States
Prior art keywords
cooling
component
recited
gas turbine
cooling cavity
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.)
Abandoned
Application number
US14/665,170
Other languages
English (en)
Inventor
Brandon W. Spangler
Gina CAVALLO
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.)
Raytheon Technologies 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
Priority to US14/665,170 priority Critical patent/US20150285081A1/en
Publication of US20150285081A1 publication Critical patent/US20150285081A1/en
Priority to US15/919,253 priority patent/US10774655B2/en
Abandoned legal-status Critical Current

Links

Images

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
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • F01D5/187—Convection cooling
    • 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
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • F01D5/186—Film cooling
    • 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
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/147—Construction, i.e. structural features, e.g. of weight-saving hollow blades
    • 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
    • F01D9/00—Stators
    • F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/60—Efficient propulsion technologies, e.g. for aircraft

Definitions

  • the internal cooling circuit may include a cooling cavity and a rib that separates the cooling cavity into separate portions.
  • gas turbine engine components Because they are commonly exposed to hot combustion gases, many gas turbine engine components employ internal cooling circuits that channel a dedicated cooling fluid for cooling regions of the component. Thermal energy is transferred from the component to the cooling fluid to cool the component.
  • a component according to an exemplary aspect of the present disclosure includes, among other things, a wall that extends about a cooling cavity.
  • the cooling cavity is a dual-fed cavity that is fed from at least two different locations.
  • a rib separates the cooling cavity into a first portion and a second portion that is fluidly isolated from the first portion.
  • the component is one of a vane, a blade, a blade outer air seal (BOAS), and a liner.
  • BOAS blade outer air seal
  • the first portion is an outer diameter portion and the second portion is an inner diameter portion of the cooling cavity.
  • the components include a plurality of openings through portions of the wall associated with both the first portion and the second portion.
  • the plurality of openings are film cooling holes.
  • a gas turbine engine includes, among other things, a component that defines a cooling circuit configured to cool the component with a cooling fluid.
  • the cooling circuit includes a cooling cavity disposed inside of the component and an axial rib that divides the cooling cavity into a first portion and a second portion that is separate from the first portion.
  • the rib fluidly isolates the first portion from the second portion.
  • the cooling circuit is disposed inside of a body of the component.
  • a method of cooling a gas turbine engine component includes, among other things, dividing a cooling cavity disposed inside the gas turbine engine component into a first portion and a second portion with a rib, communicating a first cooling fluid from a first location into the first portion, and communicating a second cooling fluid from a second location into the second portion.
  • the method includes expelling the first and second cooling fluids from the cooling cavity through a plurality of openings.
  • the second cooling fluid is separate from the first cooling fluid.
  • FIG. 1 illustrates a schematic, cross-sectional view of a gas turbine engine.
  • FIG. 2 illustrates a gas turbine engine component according to a first embodiment of this disclosure.
  • FIG. 3 illustrates a cross-sectional view through Section A-A of FIG. 2 .
  • FIG. 6 illustrates another flow separating rib.
  • This disclosure relates to a gas turbine engine component that includes an internal cooling circuit.
  • the cooling circuit employs one or more cooling cavities disposed inside of the component.
  • a flow separating rib is positioned to divide the cooling cavity into at least two portions.
  • the cooling cavity may be fed with separate cooling fluids at opposite sides of the cavity. These opposite fluid flows are fluidly isolated between the first portion and the second portion by the rib in order to maintain a constant fluid flow within each portion even where pressure differentials may exist between the opposite sides.
  • a more evenly cooled part is achieved by maintaining constant fluid flows within each portion of the cooling cavity.
  • FIG. 1 schematically illustrates a gas turbine engine 20 .
  • the gas turbine engine 20 is disclosed herein as a two-spool turbofan that generally incorporates a fan section 22 , a compressor section 24 , a combustor section 26 and a turbine section 28 .
  • Alternative engines might include an augmenter section (not shown) among other systems or features.
  • the fan section 22 drives air along a bypass flow path B in a bypass duct defined within a nacelle 15
  • the compressor section 24 drives air along a core flow path C for compression and communication into the combustor section 26 then expansion through the turbine section 28 .
  • the exemplary engine 20 generally includes a low speed spool 30 and a high speed spool 32 mounted for rotation about an engine central longitudinal axis A relative to an engine static structure 36 via several bearing systems 38 . It should be understood that various bearing systems 38 at various locations may alternatively or additionally be provided, and the location of the bearing systems 38 may be varied as appropriate to the application.
  • the low speed spool 30 generally includes an inner shaft 40 that interconnects a fan 42 , a first (or low) pressure compressor 44 and a first (or low) pressure turbine 46 .
  • the inner shaft 40 is connected to the fan 42 through a speed change mechanism, which in exemplary gas turbine engine 20 is illustrated as a geared architecture 48 to drive the fan 42 at a lower speed than the low speed spool 30 .
  • the high speed spool 32 includes an outer shaft 50 that interconnects a second (or high) pressure compressor 52 and a second (or high) pressure turbine 54 .
  • a combustor 56 is arranged in exemplary gas turbine 20 between the high pressure compressor 52 and the high pressure turbine 54 .
  • a mid-turbine frame 57 of the engine static structure 36 is arranged generally between the high pressure turbine 54 and the low pressure turbine 46 .
  • the mid-turbine frame 57 further supports bearing systems 38 in the turbine section 28 .
  • the inner shaft 40 and the outer shaft 50 are concentric and rotate via the bearing systems 38 about the engine central longitudinal axis A which is collinear with their longitudinal axes.
  • the core airflow is compressed by the low pressure compressor 44 then the high pressure compressor 52 , mixed and burned with fuel in the combustor 56 , then expanded over the high pressure turbine 54 and low pressure turbine 46 .
  • the mid-turbine frame 57 includes airfoils 59 which are in the core airflow path C.
  • the turbines 46 , 54 rotationally drive the respective low speed spool 30 and high speed spool 32 in response to the expansion.
  • gear system 48 may be located aft of combustor section 26 or even aft of turbine section 28
  • fan section 22 may be positioned forward or aft of the location of gear system 48 .
  • Low pressure turbine 46 pressure ratio is pressure measured prior to inlet of low pressure turbine 46 as related to the pressure at the outlet of the low pressure turbine 46 prior to an exhaust nozzle.
  • the gear system 48 may be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3:1. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention is applicable to other gas turbine engines including direct drive turbofans and turboshafts.
  • 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 (10,668 meters).
  • “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 1,150 ft/second (350.5 meters/second).
  • Each of the compressor section 24 and the turbine section 28 may include alternating rows of rotor assemblies and vane assemblies (shown schematically).
  • the rotor assemblies can carry a plurality of rotating blades 25
  • each vane assembly can carry a plurality of vanes 27 that extend into the core flow path C.
  • the blades 25 may either create or extract energy in the form of pressure from the core airflow as it is communicated along the core flow path C.
  • the vanes 27 direct the core airflow to the blades 25 to either add or extract energy.
  • a rib 82 may axially extend inside of the cooling cavity 72 to separate the cooling cavity 72 into a first portion 84 and a second portion 86 .
  • the first portion 84 is an outer diameter portion of the cooling cavity 72 and the second portion 86 is an inner diameter portion.
  • other configurations are also contemplated as being within the scope of this disclosure, including but not limited to circumferentially spaced portions (see, e.g., FIG. 5 ) and axially spaced portions (see, e.g., FIG. 6 ).
  • the rib 82 extends between the opposing sides 90 , 92 of the wall 80 to completely seal and separate the first portion 84 of the cooling cavity 72 from the second portion 86 of the cooling cavity 72 . Said another way, the rib 82 is a solid flow separator that fluidly isolates the first portion 84 from the second portion 86 of the cooling cavity 72 .
  • the rib 82 may be positioned at a mid-span M of the cooling cavity 72 .
  • the actual location of the rib 82 could vary part-by-part and may depend on pressure differentials that exist between the first portion 84 and the second portion 86 of the cooling cavity 72 , among other factors.
  • the cooling cavity 72 is a dual-fed cavity that is fed with a cooling fluid at both of its opposite sides (i.e., fed from two distinct locations).
  • the cooling cavity 72 is fed from distinct coolant sources that communicate coolant from different locations.
  • the first portion 84 of the cooling cavity 72 may be fed with a first cooling fluid F 1 , such as a first bleed airflow
  • the second portion 86 of the cooling cavity 72 may be fed with a second cooling fluid F 2 , such as a second bleed airflow.
  • the cooling fluids F 1 and F 2 may be separate from one another.
  • an inlet 85 of the first portion 84 of the cooling cavity 72 is positioned in a relatively high pressure area and an inlet 87 of the second portion 86 of the cooling cavity 72 is positioned at a relatively low pressure area.
  • an opposite configuration is also possible in which the inlet 85 of the first portion 84 is located at a relatively low pressure area and the inlet 87 of the second portion 86 is within a relatively high pressure area (see FIG. 4 ).
  • flow of the first and second cooling fluids Fl, F 2 remains constant within both the first portion 84 and the second portion 86 because these portions are sealed from one another by the rib 82 . Maintaining consistent flow in this manner results in relatively consistent Mach numbers, pressure losses, heat transfer and metal temperatures throughout the cooling cavity 72 . In other words, the component 60 is more evenly cooled by virtue of the flow separating rib 82 .
  • a plurality of openings 96 may extend through portions of the wall 80 associated with both the first portion 84 and the second portion 86 of the cooling cavity 72 .
  • the cooling fluids F 1 , F 2 that are circulated in the first and second portions 84 , 86 , respectively, may be expelled through the openings 96 .
  • the openings 96 are film cooling holes.
  • the openings 96 are slots. Any type of opening may extend through the wall 80 for expelling the cooling fluids F 1 , F 2 from the cooling cavity 72 .
  • FIG. 5 illustrates another component 160 that can be incorporated into a gas turbine engine.
  • like reference numbers designate like elements where appropriate and reference numerals with the addition of 100 or multiples thereof designate modified elements that are understood to incorporate the same features and benefits of the corresponding original elements.
  • the component 160 is represented as a BOAS.
  • the component 160 includes a body 161 having a radially inner face 163 and a radially outer face 165 .
  • the radially inner face 163 and the radially outer face 165 extend circumferentially between a first mate face 167 and a second mate face 169 and extend axially between a leading edge 171 and a trailing edge 173 .
  • a cooling cavity 172 may be disposed inside the body 161 .
  • the cooling cavity 172 of this embodiment circumferentially extends between the first mate face 167 and the second mate face 169 .
  • a wall 180 may extend about the cooling cavity 172 .
  • the cooling cavity 172 is divided into a first portion 184 and a second portion 186 by a rib 182 .
  • the rib 182 fluidly isolates the first portion 184 from the second portion 186 .
  • the first portion 184 of the cooling cavity 172 may be fed with a first cooling fluid F 1 at a location adjacent to the first mate face 167 and the second portion 186 may be fed with a second cooling fluid F 2 at a location adjacent to the second mate face 169 .
  • the rib 182 is adapted to maintain these split flows at relatively constant flow levels despite potential pressure differentials that may exist between the first mate face 167 and the second mate face 169 .

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Architecture (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
US14/665,170 2014-04-04 2015-03-23 Gas turbine engine component with flow separating rib Abandoned US20150285081A1 (en)

Priority Applications (2)

Application Number Priority Date Filing Date Title
US14/665,170 US20150285081A1 (en) 2014-04-04 2015-03-23 Gas turbine engine component with flow separating rib
US15/919,253 US10774655B2 (en) 2014-04-04 2018-03-13 Gas turbine engine component with flow separating rib

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
US201461975198P 2014-04-04 2014-04-04
US14/665,170 US20150285081A1 (en) 2014-04-04 2015-03-23 Gas turbine engine component with flow separating rib

Related Child Applications (1)

Application Number Title Priority Date Filing Date
US15/919,253 Continuation-In-Part US10774655B2 (en) 2014-04-04 2018-03-13 Gas turbine engine component with flow separating rib

Publications (1)

Publication Number Publication Date
US20150285081A1 true US20150285081A1 (en) 2015-10-08

Family

ID=52780972

Family Applications (1)

Application Number Title Priority Date Filing Date
US14/665,170 Abandoned US20150285081A1 (en) 2014-04-04 2015-03-23 Gas turbine engine component with flow separating rib

Country Status (2)

Country Link
US (1) US20150285081A1 (fr)
EP (1) EP2927429B1 (fr)

Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
USD946528S1 (en) * 2020-09-04 2022-03-22 Siemens Energy Global GmbH & Co. KG Turbine vane
USD947126S1 (en) * 2020-09-04 2022-03-29 Siemens Energy Global GmbH & Co. KG Turbine vane
USD947127S1 (en) * 2020-09-04 2022-03-29 Siemens Energy Global GmbH & Co. KG Turbine vane

Citations (13)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4312624A (en) * 1980-11-10 1982-01-26 United Technologies Corporation Air cooled hollow vane construction
US4461612A (en) * 1982-04-27 1984-07-24 Rolls-Royce Limited Aerofoil for a gas turbine engine
US5711650A (en) * 1996-10-04 1998-01-27 Pratt & Whitney Canada, Inc. Gas turbine airfoil cooling
US5993150A (en) * 1998-01-16 1999-11-30 General Electric Company Dual cooled shroud
US7104756B2 (en) * 2004-08-11 2006-09-12 United Technologies Corporation Temperature tolerant vane assembly
US20090067994A1 (en) * 2007-03-01 2009-03-12 United Technologies Corporation Blade outer air seal
US7665962B1 (en) * 2007-01-26 2010-02-23 Florida Turbine Technologies, Inc. Segmented ring for an industrial gas turbine
US20100183427A1 (en) * 2009-01-19 2010-07-22 George Liang Turbine blade with micro channel cooling system
US7921654B1 (en) * 2007-09-07 2011-04-12 Florida Turbine Technologies, Inc. Cooled turbine stator vane
US20120014808A1 (en) * 2010-07-14 2012-01-19 Ching-Pang Lee Near-wall serpentine cooled turbine airfoil
US8398371B1 (en) * 2010-07-12 2013-03-19 Florida Turbine Technologies, Inc. Turbine blade with multiple near wall serpentine flow cooling
US20130315725A1 (en) * 2011-05-13 2013-11-28 Mitsubishi Heavy Industries, Ltd. Turbine vane
US20140271153A1 (en) * 2013-03-12 2014-09-18 Rolls-Royce Corporation Cooled ceramic matrix composite airfoil

Family Cites Families (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4798515A (en) * 1986-05-19 1989-01-17 The United States Of America As Represented By The Secretary Of The Air Force Variable nozzle area turbine vane cooling
GB2405451B (en) * 2003-08-23 2008-03-19 Rolls Royce Plc Vane apparatus for a gas turbine engine
US8628298B1 (en) * 2011-07-22 2014-01-14 Florida Turbine Technologies, Inc. Turbine rotor blade with serpentine cooling

Patent Citations (17)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4312624A (en) * 1980-11-10 1982-01-26 United Technologies Corporation Air cooled hollow vane construction
US4461612A (en) * 1982-04-27 1984-07-24 Rolls-Royce Limited Aerofoil for a gas turbine engine
US5711650A (en) * 1996-10-04 1998-01-27 Pratt & Whitney Canada, Inc. Gas turbine airfoil cooling
US5993150A (en) * 1998-01-16 1999-11-30 General Electric Company Dual cooled shroud
US7104756B2 (en) * 2004-08-11 2006-09-12 United Technologies Corporation Temperature tolerant vane assembly
US7665962B1 (en) * 2007-01-26 2010-02-23 Florida Turbine Technologies, Inc. Segmented ring for an industrial gas turbine
US20090067994A1 (en) * 2007-03-01 2009-03-12 United Technologies Corporation Blade outer air seal
US7921654B1 (en) * 2007-09-07 2011-04-12 Florida Turbine Technologies, Inc. Cooled turbine stator vane
US20100183427A1 (en) * 2009-01-19 2010-07-22 George Liang Turbine blade with micro channel cooling system
US8398371B1 (en) * 2010-07-12 2013-03-19 Florida Turbine Technologies, Inc. Turbine blade with multiple near wall serpentine flow cooling
US20120014808A1 (en) * 2010-07-14 2012-01-19 Ching-Pang Lee Near-wall serpentine cooled turbine airfoil
US8535006B2 (en) * 2010-07-14 2013-09-17 Siemens Energy, Inc. Near-wall serpentine cooled turbine airfoil
US20130302167A1 (en) * 2010-07-14 2013-11-14 Mikro Systems, Inc. Near-Wall Serpentine Cooled Turbine Airfoil
US8870537B2 (en) * 2010-07-14 2014-10-28 Mikro Systems, Inc. Near-wall serpentine cooled turbine airfoil
US20130315725A1 (en) * 2011-05-13 2013-11-28 Mitsubishi Heavy Industries, Ltd. Turbine vane
US9523283B2 (en) * 2011-05-13 2016-12-20 Mitsubishi Heavy Industries, Ltd. Turbine vane
US20140271153A1 (en) * 2013-03-12 2014-09-18 Rolls-Royce Corporation Cooled ceramic matrix composite airfoil

Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
USD946528S1 (en) * 2020-09-04 2022-03-22 Siemens Energy Global GmbH & Co. KG Turbine vane
USD947126S1 (en) * 2020-09-04 2022-03-29 Siemens Energy Global GmbH & Co. KG Turbine vane
USD947127S1 (en) * 2020-09-04 2022-03-29 Siemens Energy Global GmbH & Co. KG Turbine vane

Also Published As

Publication number Publication date
EP2927429B1 (fr) 2018-07-04
EP2927429A1 (fr) 2015-10-07

Similar Documents

Publication Publication Date Title
US10907481B2 (en) Platform cooling core for a gas turbine engine rotor blade
US10774655B2 (en) Gas turbine engine component with flow separating rib
US10406596B2 (en) Core arrangement for turbine engine component
EP2944761B1 (fr) Composant de moteur à turbine à gaz avec circuit de refroidissement de plate-forme
US10041358B2 (en) Gas turbine engine blade squealer pockets
EP3084136B1 (fr) Aube rotorique et procédé associé de refroidissement d'une plateforme d'une aube rotorique
US20140075947A1 (en) Gas turbine engine component cooling circuit
EP2977555B1 (fr) Plateforme d'aube avec canaux de refroidissement
US10215051B2 (en) Gas turbine engine component providing prioritized cooling
EP3056680B1 (fr) Systèmes d'air de fuite pour turbomachines
US20170002662A1 (en) Gas turbine engine airfoil with bi-axial skin core
EP2885520B1 (fr) Composant pour un moteur à turbine à gaz et procédé de refroidissement associé
EP2927429B1 (fr) Composant de moteur à turbine à gaz avec nervure de séparation de flux
US20150354372A1 (en) Gas turbine engine component with angled aperture impingement
EP3101236B1 (fr) Joints de bord de fuite de plate-forme
EP3401509B1 (fr) Ensemble de joints externes d'aubes de turbine
EP2977557B1 (fr) Structure d'aube refroidie et procédé de refroidissement associé
EP3550109B1 (fr) Composant de moteur à turbine à gaz avec nervure de séparation de flux
US9790801B2 (en) Gas turbine engine component having suction side cutback opening

Legal Events

Date Code Title Description
STCB Information on status: application discontinuation

Free format text: ABANDONED -- FAILURE TO RESPOND TO AN OFFICE ACTION