EP2914819A2 - Flüssigkeitsgekühlte dichtungsanordnung für einen gasturbinenmotor - Google Patents
Flüssigkeitsgekühlte dichtungsanordnung für einen gasturbinenmotorInfo
- Publication number
- EP2914819A2 EP2914819A2 EP13874213.5A EP13874213A EP2914819A2 EP 2914819 A2 EP2914819 A2 EP 2914819A2 EP 13874213 A EP13874213 A EP 13874213A EP 2914819 A2 EP2914819 A2 EP 2914819A2
- Authority
- EP
- European Patent Office
- Prior art keywords
- passage
- gas turbine
- turbine engine
- recited
- rotational component
- 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- Withdrawn
Links
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
- F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
- F01D11/005—Sealing means between non relatively rotating elements
-
- 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
- F01D25/18—Lubricating arrangements
- F01D25/183—Sealing means
-
- 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
- F02C7/06—Arrangements of bearings; Lubricating
-
- 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
- F05D2250/00—Geometry
- F05D2250/20—Three-dimensional
- F05D2250/24—Three-dimensional ellipsoidal
- F05D2250/241—Three-dimensional ellipsoidal spherical
-
- 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
- Y10—TECHNICAL SUBJECTS COVERED BY FORMER USPC
- Y10T—TECHNICAL SUBJECTS COVERED BY FORMER US CLASSIFICATION
- Y10T29/00—Metal working
- Y10T29/49—Method of mechanical manufacture
- Y10T29/49316—Impeller making
- Y10T29/4932—Turbomachine making
Definitions
- FLUID-COOLED SEAL ARRANGEMENT FOR A GAS TURBINE ENGINE
- the present disclosure relates to a gas turbine engine and, more particularly, to fluid-cooled seal arrangements therefor.
- Certain sections of gas turbine engines may be subjected to high temperatures and pressures.
- Some engine components may be sensitive to the harsh environment thereof and are shielded therefrom. These components are typically vented by ambient or cooling bleed-off air or have cooling oil flowing therethrough. In order to maintain cool air in the cavities housing these components, the cavities must be shielded from engine pressure and temperature differentials.
- seal systems are positioned to prevent high temperature and pressure air from flowing downstream into the areas with lower temperature and pressure air.
- One such seal system includes arcuate carbon material segments arranged to form a stationary carbon ring that forms a rubbing interface with a rotating seal runner.
- the rubbing interface between the rotating seal runner and the carbon ring minimizes or prevents leakage, however if the heat generated by the rubbing interface is not adequately dissipated, the rotating seal runner may thermally distort. This may degrade performance manifested by excessive fluid leakage.
- One approach to minimize overheating of the seal interface includes the delivery of a cooling fluid onto the underside of the rotating seal runner sprayed from a stationary nozzle positioned proximate the rotating seal runner.
- the relative motion between the rotating seal runner and the stationary nozzle causes a uniform film of cooling fluid to be deposited on the seal runner to extract thermal energy.
- Stationary nozzles provide a consistently even film of cooling fluid on the underside of the rotating seal runner, however their applicability on many gas turbine engines is limited by physical constraints that prevent the nozzle from being located proximate the seal runner.
- Another approach utilizes a rotating distributor to deliver the cooling fluid onto the underside of the rotating seal runner.
- the rotating distributor is typically affixed to the seal runner, and a steady stream of cooling fluid is delivered through a central passageway in the rotating distributor to the underside of the seal runner.
- a series of openings in the rotating distributor dispense the cooling fluid onto the seal runner.
- Carbon seal systems with a rotating distributor include a significant quantity of openings to deliver an even film of cooling fluid. If this design parameter is not satisfied, an uneven film of cooling fluid is distributed across the seal runner, which may cause an uneven extraction of heat.
- carbon seal system designs with large quantities of cooling fluid dispensing openings may, under centrifugal loads, concentrate stresses which may lead to fatigue life shortfalls.
- expensive rotor grade material may be used to meet fatigue life requirements. In other cases even rotor grade material is insufficient to meet desired fatigue life requirements.
- a rotational component for a gas turbine engine includes at least one passage with a semi- spherical end.
- the at least one passage is an oil communication passage.
- the at least one passage is a blind hole.
- a gas turbine engine includes a rotational component having at least one passage with a semi-spherical end.
- the rotational component is a seal runner.
- the at least one passage is an oil communication passage.
- the at least one passage is a blind hole.
- the oil communication passage is an outlet passage.
- the rotational component includes a cantilevered section.
- the cantilevered section collects and directs a lubricant toward an internal inlet passage.
- the cantilevered section include a hook-shaped end section.
- the method comprises drilling the passage and inserting a ball-endmill into the passage to form the end.
- the method includes drilling the passage with a drill tool having spherical shaped flutes.
- Figure 1 is a schematic cross-section of a gas turbine engine
- Figure 2 is a partial longitudinal schematic sectional view of a bearing compartment that may be used with the gas turbine engine shown in Figure 1 ;
- Figure 3 is a sectional view of a seal runner with a passage with a semi- spherical end according to one disclosed non-limiting embodiment.
- Figure 4 is an expanded view of a RELATED ART passage. DETAILED DESCRIPTION
- 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 augmentor section (not shown) among other systems or features.
- the fan section 22 drives air along a bypass flowpath while the compressor section 24 drives air along a core flowpath for compression and communication into the combustor section 26 then expansion through the turbine section 28.
- turbofan gas turbine engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to use with turbofans as the teachings may be applied to other types of turbine engines such as a three-spool (plus fan) engine wherein an intermediate spool includes an intermediate pressure compressor (IPC) between the LPC and HPC and an intermediate pressure turbine (IPT) between the HPT and LPT.
- IPC intermediate pressure compressor
- IPT intermediate pressure turbine
- the engine 20 generally includes a low spool 30 and a high spool 32 mounted for rotation about an engine central longitudinal axis A relative to an engine static structure 36 via several bearing structures 38.
- the low spool 30 generally includes an inner shaft 40 that interconnects a fan 42, a low pressure compressor 44 ("LPC") and a low pressure turbine 46 ("LPT").
- the inner shaft 40 drives the fan 42 directly or through a geared architecture 48 to drive the fan 42 at a lower speed than the low spool 30.
- An exemplary reduction transmission is an epicyclic transmission, namely a planetary or star gear system.
- the high spool 32 includes an outer shaft 50 that interconnects a high pressure compressor 52 (“HPC”) and high pressure turbine 54 (“HPT").
- a combustor 56 is arranged between the high pressure compressor 52 and the high pressure turbine 54.
- the inner shaft 40 and the outer shaft 50 are concentric and rotate about the engine central longitudinal axis A which is collinear with their longitudinal axes.
- Core airflow is compressed by the low pressure compressor 44 then the high pressure compressor 52, mixed with the fuel and burned in the combustor 56, then expanded over the high pressure turbine 54 and low pressure turbine 46.
- the turbines 54, 46 rotationally drive the respective low spool 30 and high spool 32 in response to the expansion.
- the main engine shafts 40, 50 are supported at a plurality of points by bearing structures 38 within the static structure 36. It should be understood that various bearing structures 38 at various locations may alternatively or additionally be provided.
- the gas turbine engine 20 is a high-bypass geared aircraft engine.
- the gas turbine engine 20 bypass ratio is greater than about six (6:1).
- the geared architecture 48 can include an epicyclic gear train, such as a planetary gear system or other gear system.
- the example epicyclic gear train has a gear reduction ratio of greater than about 2.3, and in another example is greater than about 2.5:1.
- the geared turbofan enables operation of the low spool 30 at higher speeds which can increase the operational efficiency of the low pressure compressor 44 and low pressure turbine 46 and render increased pressure in a fewer number of stages.
- a pressure ratio associated with the low pressure turbine 46 is pressure measured prior to the inlet of the low pressure turbine 46 as related to the pressure at the outlet of the low pressure turbine 46 prior to an exhaust nozzle of the gas turbine engine 20.
- the bypass ratio of the gas turbine engine 20 is greater than about ten (10:1)
- the fan diameter is significantly larger than that of the low pressure compressor 44
- the low pressure turbine 46 has a pressure ratio that is greater than about 5 (5:1). It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present disclosure is applicable to other gas turbine engines including direct drive turbofans.
- a significant amount of thrust is provided by the bypass flow path B due to the high bypass ratio.
- the fan section 22 of the gas turbine engine 20 is designed for a particular flight condition - typically cruise at about 0.8 Mach and about 35,000 feet. This flight condition, with the gas turbine engine 20 at its best fuel consumption, is also known as bucket cruise Thrust Specific Fuel Consumption (TSFC). TSFC is an industry standard parameter of fuel consumption per unit of thrust.
- Fan Pressure Ratio is the pressure ratio across a blade of the fan section 22 without the use of a Fan Exit Guide Vane system.
- the low Fan Pressure Ratio according to one non-limiting embodiment of the example gas turbine engine 20 is less than 1.45.
- Low Corrected Fan Tip Speed is the actual fan tip speed divided by an industry standard temperature correction of "T" / 518.7 0'5 in which "T" represents the ambient temperature in degrees Rankine.
- the Low Corrected Fan Tip Speed according to one non-limiting embodiment of the example gas turbine engine 20 is less than about 1150 fps (351 m/s).
- a bearing structure 38A includes a bearing 58 and fluid-cooled seal arrangements 60.
- the fluid-cooled seal arrangements 60A, 60B may each be, in the disclosed non-limiting embodiment a carbon seal system to seal a "wet" zone from a "dry" zone.
- regions or volumes containing lubricant will be referred to as a "wet” zone and a lubricant free region will be referred to as a "dry" zone.
- the fluid-cooled seal arrangement 60 generally includes a stationary component 62 and a rotational component 64.
- the stationary component 62 is coupled to a rotationally fixed structure such as the static structure 36 while the rotational component 64 is mechanically connected to a rotating structure such as the outer shaft 50. It should be appreciated, however, that any rotating structure such as a rotor hub may alternatively mount the rotational component 64.
- the rotational component 64 is referred to herein as a seal runner.
- the stationary component 62 is arranged with respect to the rotational component 64 to form a rubbing interface 66 therebetween which in the disclosed non-limiting embodiment, is axially oriented with respect to the engine axis A.
- the rubbing interface 66 may be radially oriented.
- a first annular surface 68 is defined by the stationary component 62 and a second annular surface 70 is defined by the rotational component 64 which are maintained in rubbing contact to form a fluid tight seal at the rubbing interface 66.
- the lubrication system (illustrated schematically) provides cooling fluid under pressure to lubricate and cool the moving parts of the engine 20, such as the bearing 58 and fluid-cooled seal arrangement 60 through a nozzle 72.
- the lubrication system discharges the fluid from the nozzle 72 with sufficient kinetic energy to spray an underside of a cantilevered section 74 of the rotational component 64.
- the cantilevered section 74 may include a hook- shaped end section 76 to collect and direct the lubricant toward an internal passage 78 ( Figure 3).
- the internal inlet passage 78 may be formed from one or a multiple of drill holes.
- the internal inlet passage 78 communicates with an internal outlet passage 82.
- the internal outlet passage 82 may also be formed from one or a multiple of drill holes.
- the internal outlet passage 82 includes a semi-spherical end 90 as compared to a conventional conical end E ( Figure 7; RELATED ART). It should be understood that although only the internal outlet passage 82 includes the semi-spherical end 90, any passage may benefit herefrom such as the internal inlet passage 78.
- the semi-spherical end 90 is applicable to any blind hole such as a drill hole and may be manufactured with, for example only, a ball- endmill type tool which is chased down a conventional drill hole or other special drill tool with spherical shaped flutes.
- the semi-spherical end 90 reduces concentrated stress in the passages and increases fatigue life of the rotational component 64. That is, the semi-spherical end 90 may be particularly applicable to any rotating structure such as the disclosed seal runner which is subject to high centrifugal forces and is potentially a less expensive alternative to usage of higher grade materials to meet fatigue life requirements.
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Applications Claiming Priority (2)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
US13/666,591 US20140119887A1 (en) | 2012-11-01 | 2012-11-01 | Fluid-cooled seal arrangement for a gas turbine engine |
PCT/US2013/068053 WO2014120309A2 (en) | 2012-11-01 | 2013-11-01 | Fluid-cooled seal arrangement for a gas turbine engine |
Publications (2)
Publication Number | Publication Date |
---|---|
EP2914819A2 true EP2914819A2 (de) | 2015-09-09 |
EP2914819A4 EP2914819A4 (de) | 2016-01-20 |
Family
ID=50547383
Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
EP13874213.5A Withdrawn EP2914819A4 (de) | 2012-11-01 | 2013-11-01 | Flüssigkeitsgekühlte dichtungsanordnung für einen gasturbinenmotor |
Country Status (3)
Country | Link |
---|---|
US (1) | US20140119887A1 (de) |
EP (1) | EP2914819A4 (de) |
WO (1) | WO2014120309A2 (de) |
Families Citing this family (12)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
EP2964935B1 (de) | 2013-03-08 | 2018-12-12 | United Technologies Corporation | Flüssigkeitsgekühlte dichtungsanordnung für einen gasturbinenmotor |
US9631508B2 (en) | 2013-06-13 | 2017-04-25 | Pratt & Whitney Canada Corp. | Internally cooled seal runner |
US20180045316A1 (en) * | 2016-08-09 | 2018-02-15 | United Technologies Corporation | Hydrodynamic seal seat cooling features |
US20180087404A1 (en) * | 2016-09-23 | 2018-03-29 | Rolls-Royce Corporation | Oil distributor |
US11028717B2 (en) | 2017-06-26 | 2021-06-08 | Raytheon Technologies Corporation | Bearing assembly for gas turbine engines |
US10746051B2 (en) * | 2018-09-21 | 2020-08-18 | United Technologies Corporation | Hybrid wet-dry face seal seat |
US10774684B2 (en) * | 2018-10-24 | 2020-09-15 | Raytheon Technologies Corporation | Gas turbine engine seal assemblies |
US10975723B2 (en) * | 2019-02-26 | 2021-04-13 | Raytheon Technologies Corporation | Gas turbine engine including seal plate providing increased cooling adjacent contact area |
US11248492B2 (en) | 2019-03-18 | 2022-02-15 | Raytheon Technologies Corporation | Seal assembly for a gas turbine engine |
US11203948B2 (en) | 2019-09-06 | 2021-12-21 | Pratt & Whitney Canada Corp. | Seal runner and method |
US11193389B2 (en) * | 2019-10-18 | 2021-12-07 | Raytheon Technologies Corporation | Fluid cooled seal land for rotational equipment seal assembly |
US11441448B2 (en) | 2020-02-13 | 2022-09-13 | Raytheon Technologies Corporation | Impingement cooled rotating seal |
Family Cites Families (20)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US1269951A (en) | 1916-08-28 | 1918-06-18 | Frederick Livingstone Morris | Trolley-wheel and bearing therefor. |
US2857182A (en) * | 1955-05-31 | 1958-10-21 | Gen Motors Corp | Shaft seal |
US2874803A (en) | 1955-07-26 | 1959-02-24 | United Aircraft Corp | Lubrication means |
US2990202A (en) * | 1958-09-18 | 1961-06-27 | United Aircraft Corp | Labyrinth face seal plate |
NL130801C (de) * | 1961-02-10 | |||
US4928978A (en) * | 1988-04-07 | 1990-05-29 | United Technologies Corporation | Rotating shaft seal |
US5284347A (en) * | 1991-03-25 | 1994-02-08 | General Electric Company | Gas bearing sealing means |
US5685671A (en) * | 1993-11-01 | 1997-11-11 | Smith International, Inc. | Diamond or CBN fluted center cutting end mill |
FR2713527B1 (fr) | 1993-12-08 | 1996-01-12 | Snecma | Foret de perçage à tête hémisphérique à coupe évolutive. |
US5626347A (en) * | 1994-09-29 | 1997-05-06 | Alliedsignal Inc. | Coning resistant face seal having a "U" shape |
US6196790B1 (en) * | 1998-12-17 | 2001-03-06 | United Technologies Corporation | Seal assembly for an intershaft seal in a gas turbine engine |
EP1496197B1 (de) * | 2000-09-26 | 2008-11-26 | Mitsubishi Heavy Industries, Ltd. | Wellenanordnung für eine dampfgekühlte Gasturbine |
US8657297B2 (en) * | 2004-05-28 | 2014-02-25 | Stein Seal Company | Air riding seal |
JP2007078072A (ja) | 2005-09-14 | 2007-03-29 | Toyota Industries Corp | 内燃機関のコンロッド |
US20090051120A1 (en) * | 2007-08-23 | 2009-02-26 | United Technologies Corp. | Gas Turbine Engine Systems Involving Hydrostatic Face Seals |
US8215895B2 (en) * | 2008-03-03 | 2012-07-10 | Rolls-Royce Corporation | Vapor phase lubrication system |
US8490982B2 (en) * | 2008-05-22 | 2013-07-23 | Stein Seal Company | Windback device |
US7931277B2 (en) * | 2009-08-27 | 2011-04-26 | Stein Seal Company | Hydrodynamic circumferential seal system for large translations |
GB2473828B (en) * | 2009-09-24 | 2011-08-10 | Rolls Royce Plc | Oil seal assembly |
US8845282B2 (en) * | 2011-09-28 | 2014-09-30 | United Technologies Corporation | Seal plate with cooling passage |
-
2012
- 2012-11-01 US US13/666,591 patent/US20140119887A1/en not_active Abandoned
-
2013
- 2013-11-01 EP EP13874213.5A patent/EP2914819A4/de not_active Withdrawn
- 2013-11-01 WO PCT/US2013/068053 patent/WO2014120309A2/en active Application Filing
Also Published As
Publication number | Publication date |
---|---|
WO2014120309A3 (en) | 2014-10-16 |
US20140119887A1 (en) | 2014-05-01 |
WO2014120309A2 (en) | 2014-08-07 |
EP2914819A4 (de) | 2016-01-20 |
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Inventor name: RANGANATH, SANTOSH Inventor name: LARSON, CHRISTOPHER J. Inventor name: BORATGIS, ERNEST Inventor name: MUTSENGI, FUNGAYI Inventor name: ULLAH, M RIFAT Inventor name: LYLE, STEPHEN J. |
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