US20160003146A1 - Rotating inlet cowl for a turbine engine, comprising an eccentric forward end - Google Patents

Rotating inlet cowl for a turbine engine, comprising an eccentric forward end Download PDF

Info

Publication number
US20160003146A1
US20160003146A1 US14/620,798 US201514620798A US2016003146A1 US 20160003146 A1 US20160003146 A1 US 20160003146A1 US 201514620798 A US201514620798 A US 201514620798A US 2016003146 A1 US2016003146 A1 US 2016003146A1
Authority
US
United States
Prior art keywords
inlet cowl
forward end
cone
rotation axis
cowl
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
US14/620,798
Other versions
US9243562B1 (en
Inventor
Mickael Delapierre
Herve Gignoux
Gael Loro
Sylvie Wintenberger
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.)
Safran Aircraft Engines SAS
Original Assignee
SNECMA SAS
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 SNECMA SAS filed Critical SNECMA SAS
Priority to US14/620,798 priority Critical patent/US9243562B1/en
Publication of US20160003146A1 publication Critical patent/US20160003146A1/en
Application granted granted Critical
Publication of US9243562B1 publication Critical patent/US9243562B1/en
Assigned to SAFRAN AIRCRAFT ENGINES reassignment SAFRAN AIRCRAFT ENGINES CHANGE OF NAME (SEE DOCUMENT FOR DETAILS). Assignors: SNECMA
Assigned to SAFRAN AIRCRAFT ENGINES reassignment SAFRAN AIRCRAFT ENGINES CORRECTIVE ASSIGNMENT TO CORRECT THE COVER SHEET TO REMOVE APPLICATION NOS. 10250419, 10786507, 10786409, 12416418, 12531115, 12996294, 12094637 12416422 PREVIOUSLY RECORDED ON REEL 046479 FRAME 0807. ASSIGNOR(S) HEREBY CONFIRMS THE CHANGE OF NAME. Assignors: SNECMA
Active legal-status Critical Current
Anticipated expiration legal-status Critical

Links

Images

Classifications

    • 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
    • F02C7/04—Air intakes for gas-turbine plants or jet-propulsion plants
    • F02C7/047—Heating to prevent icing
    • 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
    • F02C6/00—Plural gas-turbine plants; Combinations of gas-turbine plants with other apparatus; Adaptations of gas-turbine plants for special use
    • F02C6/20—Adaptations of gas-turbine plants for driving vehicles
    • F02C6/206—Adaptations of gas-turbine plants for driving vehicles the vehicles being airscrew driven
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02K—JET-PROPULSION PLANTS
    • F02K3/00—Plants including a gas turbine driving a compressor or a ducted fan
    • F02K3/02—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber
    • F02K3/04—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type
    • F02K3/06—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type with front fan
    • 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
    • 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
    • Y10T137/00—Fluid handling
    • Y10T137/0536—Highspeed fluid intake means [e.g., jet engine intake]
    • 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
    • Y10T137/00—Fluid handling
    • Y10T137/0536—Highspeed fluid intake means [e.g., jet engine intake]
    • Y10T137/0645—With condition responsive control means

Definitions

  • This invention relates to the field of turbine engines, and more particularly to aircraft turbine engines, preferably of the turbojet type. More specifically, the invention relates to the rotating inlet cowl fitted on these turbine engines.
  • Such a rotating inlet cowl is usually composed of two parts fixed to each other, the forward part in the form of a cone and the aft part in the form of a shroud.
  • the aft end of the aft shroud is flush with the fan blade platforms in a known manner, and is in aerodynamic continuity with them and in front of them.
  • the front cone has a forward end shaped like a cone tip centred on the axis of rotation of the inlet cowl, also corresponding to the longitudinal axis of the fan and the entire turbine engine.
  • This tip is known as being a point on the turbine engine at which ice can easily collect, because the fact that it is centred on the axis of rotation makes it impossible to apply large centrifugal forces. Consequently, ice forming on the cone tip can become large before it breaks off, introducing the risk of causing damage to the fan blades that it strikes when it finally breaks free from the tip.
  • the purpose of the invention is therefore to provide at least partially a solution to the disadvantages mentioned above, compared with the embodiments of the prior art.
  • the first purpose of the invention is a rotating inlet cowl for a turbine engine, said cowl having a rotation axis and comprising a forward cone defining a forward end of the cowl.
  • this forward end is arranged to be eccentric relative to this rotation axis of the inlet cowl, and said forward cone is truncated by a truncation surface defining said forward end of the inlet cowl.
  • the eccentric nature of the end advantageously means that ice will be subject to high centrifugal forces. These will facilitate its ejection so that it will separate from the rod before it reaches a critical size that could damage the downstream fan blades.
  • the invention has the advantage that it is based on a simple design, is extremely reliable and is not very penalising in terms of cost and size.
  • the fact that it results from a forward cone being truncated makes a large contribution to the simplicity of its design.
  • said forward cone is oblique with an axis inclined from said axis of rotation of the inlet cowl. Nevertheless, it is also possible that the forward cone could be straight, that its axis could be coincident with said inlet cowl rotation axis. The advantage is then that it is possible to start from a conventional forward cone according to prior art and to truncate it in order to achieve the required embodiment.
  • said truncation surface is approximately plane, inclined relative to a plane orthogonal to the axis of rotation of the inlet cowl.
  • the rotating inlet cowl preferably comprises said front cone and a rear shroud.
  • Another purpose of the invention is a turbine engine, preferably for an aircraft, comprising a rotating inlet cowl like that described above.
  • FIG. 1 shows a longitudinal half-sectional view of a forward part of an aircraft turbine engine according to a preferred embodiment of this invention
  • FIG. 2 schematically shows an enlargement of the rotating inlet cowl fitted on the turbine engine shown in FIG. 1 ;
  • FIG. 3 shows a perspective diagrammatic view of the forward cone fitted on the rotating inlet cowl shown in FIG. 2 .
  • FIG. 1 shows a forward part 1 of a turbine engine for a turbojet type aircraft, according to a preferred embodiment of this invention.
  • FIG. 1 only the low-pressure compressor 3 of the gas generator has been represented, which is, for example, a two-compressor generator.
  • the turbine engine is provided with an air intake 4 , a fan 6 , a splitter 14 from which an annular core engine flow 16 and an annular fan flow 18 arranged radially outside the core engine flow 16 , and an inner ring 10 supporting the fan outlet guide vane assemblies 12 .
  • each of these conventional elements known to those skilled in the art is annular in shape and centred on a longitudinal axis 22 of the turbine engine.
  • the air flow F passing through the fan 6 is divided into two separate flows after it comes into contact with the upstream end of the splitter 14 , namely into a primary flow F 1 entering channel 16 and a secondary flow F 2 entering channel 18 and passing through the fan outlet guide vane assemblies 12 .
  • the turbine engine comprises a rotating inlet cowl 30 at its forward end fixed in rotation to the fan 6 .
  • the cowl 30 is provided with a forward cone 32 with axis 33 , and an aft shroud 36 installed fixed on the cone 32 , preferably by bolts 38 . Its aft end is flush with the platforms 40 of the fan blades 42 , in aerodynamic continuity in front of these platforms.
  • the forward end 44 of the rotating inlet cowl 30 is eccentric from the rotation axis 34 of this cowl 30 , the axis 34 also corresponding to the axis of the fan 6 , and more generally to the longitudinal axis 22 of the turbine engine.
  • the eccentric nature of the forward end 44 is obtained using a straight forward cone 32 , the axis 33 of which is coincident with the rotation axis 34 of the cone and the longitudinal axis of the turbine engine 22 .
  • the forward part of this cone is truncated by an approximately plane truncation surface 70 inclined relative to a plane P orthogonal to the axes 22 , 34 , for example by an angle B between 1 and 15°.
  • the truncation can define the forward eccentric end 44 because it corresponds to the most forward part of the ellipse 72 formed by the intersection between the cone 32 and the approximately plane truncation surface 70 as can be seen in FIG. 3 .
  • a balancing bead 50 could be fitted on the forward cone 32 , on the inside close to the bolted connection to the aft shroud 36 . Therefore, the purpose of this bead 50 is to compensate for the unbalanced mass, and therefore it has a variable thickness along the circumferential direction as shown diagrammatically in FIG. 2 . It may be made by making a reaming 52 with axis 54 eccentric from the axes 22 , 34 .
  • Another balancing bead 62 is provided to complete the balancing bead 50 and to compensate for the unbalanced mass resulting essentially from the offset of the forward end 44 relative to the rotation axis 34 , arranged on the inside close to the forward end 44 .
  • this bead 62 has a variable thickness along the circumferential direction as shown diagrammatically in FIG. 2 , and it could also be made by making a reaming 64 with axis 66 eccentric from axes 22 , 34 . Alternatively, or simultaneously, the unbalanced mass could be compensated by varying the thickness of the skin from which the cone 32 is formed, in the circumferential direction.

Landscapes

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

Abstract

A rotating inlet cowl for a turbine engine includes a rotation axis. The rotating inlet cowl includes a forward cone defining a forward end of the inlet cowl. The forward end is configured to be eccentric relative to the rotation axis of the inlet cowl. Furthermore, the forward cone is truncated by a truncation surface defining the forward end of the inlet cowl.

Description

    CROSS-REFERENCE TO RELATED APPLICATIONS
  • This application is a continuation of U.S. application Ser. No. 13/262,386, filed Sep. 30, 2011. U.S. application Ser. No. 13/262,386 is a U.S. National Stage Application under 35 U.S.C. §371 of International Application No. PCT/EP2010/054080 filed Mar. 29, 2010, which claims priority to French Application No. 09 52056 filed Mar. 31, 2009. The entire contents of each of the above applications are incorporated herein by reference.
    • TECHNICAL FIELD
  • This invention relates to the field of turbine engines, and more particularly to aircraft turbine engines, preferably of the turbojet type. More specifically, the invention relates to the rotating inlet cowl fitted on these turbine engines.
    • STATE OF THE PRIOR ART
  • Such a rotating inlet cowl is usually composed of two parts fixed to each other, the forward part in the form of a cone and the aft part in the form of a shroud. The aft end of the aft shroud is flush with the fan blade platforms in a known manner, and is in aerodynamic continuity with them and in front of them.
  • The front cone has a forward end shaped like a cone tip centred on the axis of rotation of the inlet cowl, also corresponding to the longitudinal axis of the fan and the entire turbine engine.
  • This tip is known as being a point on the turbine engine at which ice can easily collect, because the fact that it is centred on the axis of rotation makes it impossible to apply large centrifugal forces. Consequently, ice forming on the cone tip can become large before it breaks off, introducing the risk of causing damage to the fan blades that it strikes when it finally breaks free from the tip.
  • It is known that a de-icing system can be installed to reduce this risk, the purpose of which is to assure that ice collected on the cone tip is ejected before it reaches a critical size. However, this type of system is expensive in terms of mass and dimensions, and especially it is particularly difficult to install due to the rotation of the inlet cowl on which it is fitted.
    • SUMMARY OF THE INVENTION
  • The purpose of the invention is therefore to provide at least partially a solution to the disadvantages mentioned above, compared with the embodiments of the prior art.
  • To achieve this, the first purpose of the invention is a rotating inlet cowl for a turbine engine, said cowl having a rotation axis and comprising a forward cone defining a forward end of the cowl. According to the invention, this forward end is arranged to be eccentric relative to this rotation axis of the inlet cowl, and said forward cone is truncated by a truncation surface defining said forward end of the inlet cowl.
  • Thus, during operation, when ice has collected on the forward end of the inlet cowl, the eccentric nature of the end advantageously means that ice will be subject to high centrifugal forces. These will facilitate its ejection so that it will separate from the rod before it reaches a critical size that could damage the downstream fan blades.
  • Consequently, the invention has the advantage that it is based on a simple design, is extremely reliable and is not very penalising in terms of cost and size. The fact that it results from a forward cone being truncated makes a large contribution to the simplicity of its design.
  • According to one preferred embodiment of this invention, said forward cone is oblique with an axis inclined from said axis of rotation of the inlet cowl. Nevertheless, it is also possible that the forward cone could be straight, that its axis could be coincident with said inlet cowl rotation axis. The advantage is then that it is possible to start from a conventional forward cone according to prior art and to truncate it in order to achieve the required embodiment.
  • Preferably, said truncation surface is approximately plane, inclined relative to a plane orthogonal to the axis of rotation of the inlet cowl.
  • As mentioned above, regardless of what embodiment is envisaged, the rotating inlet cowl preferably comprises said front cone and a rear shroud.
  • Finally, another purpose of the invention is a turbine engine, preferably for an aircraft, comprising a rotating inlet cowl like that described above.
  • Other advantages and characteristics of the invention will appear in the non-restrictive detailed disclosure below.
    • BRIEF DESCRIPTION OF THE DRAWINGS
  • This description will be made with reference to the attached illustrations, among which;
  • FIG. 1 shows a longitudinal half-sectional view of a forward part of an aircraft turbine engine according to a preferred embodiment of this invention;
  • FIG. 2 schematically shows an enlargement of the rotating inlet cowl fitted on the turbine engine shown in FIG. 1; and
  • FIG. 3 shows a perspective diagrammatic view of the forward cone fitted on the rotating inlet cowl shown in FIG. 2.
    •  DETAILED DESCRIPTION OF PREFERRED EMBODIMENTS
  • FIG. 1 shows a forward part 1 of a turbine engine for a turbojet type aircraft, according to a preferred embodiment of this invention.
  • In FIG. 1, only the low-pressure compressor 3 of the gas generator has been represented, which is, for example, a two-compressor generator.
  • Starting from the forward end and following the general direction of fluid flow through the turbine engine towards the aft end as shown diagrammatically by the arrow 9, the turbine engine is provided with an air intake 4, a fan 6, a splitter 14 from which an annular core engine flow 16 and an annular fan flow 18 arranged radially outside the core engine flow 16, and an inner ring 10 supporting the fan outlet guide vane assemblies 12. Obviously, each of these conventional elements known to those skilled in the art is annular in shape and centred on a longitudinal axis 22 of the turbine engine.
  • Thus, the air flow F passing through the fan 6 is divided into two separate flows after it comes into contact with the upstream end of the splitter 14, namely into a primary flow F1 entering channel 16 and a secondary flow F2 entering channel 18 and passing through the fan outlet guide vane assemblies 12.
  • Furthermore, the turbine engine comprises a rotating inlet cowl 30 at its forward end fixed in rotation to the fan 6. In a known manner, the cowl 30 is provided with a forward cone 32 with axis 33, and an aft shroud 36 installed fixed on the cone 32, preferably by bolts 38. Its aft end is flush with the platforms 40 of the fan blades 42, in aerodynamic continuity in front of these platforms.
  • One of the special features of this invention is that the forward end 44 of the rotating inlet cowl 30 is eccentric from the rotation axis 34 of this cowl 30, the axis 34 also corresponding to the axis of the fan 6, and more generally to the longitudinal axis 22 of the turbine engine.
  • In the preferred embodiment shown in FIGS. 1 to 3, the eccentric nature of the forward end 44 is obtained using a straight forward cone 32, the axis 33 of which is coincident with the rotation axis 34 of the cone and the longitudinal axis of the turbine engine 22. Furthermore, the forward part of this cone is truncated by an approximately plane truncation surface 70 inclined relative to a plane P orthogonal to the axes 22, 34, for example by an angle B between 1 and 15°. Thus, the truncation can define the forward eccentric end 44 because it corresponds to the most forward part of the ellipse 72 formed by the intersection between the cone 32 and the approximately plane truncation surface 70 as can be seen in FIG. 3.
  • It is noted that a similar embodiment could be envisaged with an oblique forward cone 32, namely with an axis 33 inclined relative to the rotation axis 34.
  • In a known manner, a balancing bead 50 could be fitted on the forward cone 32, on the inside close to the bolted connection to the aft shroud 36. Therefore, the purpose of this bead 50 is to compensate for the unbalanced mass, and therefore it has a variable thickness along the circumferential direction as shown diagrammatically in FIG. 2. It may be made by making a reaming 52 with axis 54 eccentric from the axes 22, 34. Another balancing bead 62 is provided to complete the balancing bead 50 and to compensate for the unbalanced mass resulting essentially from the offset of the forward end 44 relative to the rotation axis 34, arranged on the inside close to the forward end 44. Therefore this bead 62 has a variable thickness along the circumferential direction as shown diagrammatically in FIG. 2, and it could also be made by making a reaming 64 with axis 66 eccentric from axes 22, 34. Alternatively, or simultaneously, the unbalanced mass could be compensated by varying the thickness of the skin from which the cone 32 is formed, in the circumferential direction.
  • When the fan and the inlet cowl 30 rotate with the eccentric forward end 44 on which ice 60 has collected, significant centrifugal forces are applied to the ice facilitating its ejection from the cowl.
  • Naturally, various modifications can be made by the skilled man in the art to the invention which has just been described, solely as non-restrictive examples.

Claims (10)

1. A rotating inlet cowl of a gas turbine engine, the inlet cowl including a rotation axis and comprising:
a forward cone defining a forward end of the inlet cowl,
wherein the forward end is configured to be eccentric relative to the rotation axis of the inlet cowl,
wherein the forward cone is truncated by a truncation surface defining the forward end of the inlet cowl, and
wherein the forward cone includes an axis parallel to and coincident with the rotation axis of the inlet cowl.
2. The rotating inlet cowl according to claim 1, wherein the forward cone includes at least one balancing bead that includes a variable thickness along a circumferential direction, to compensate for an unbalanced mass.
3. A turbine or aircraft engine, comprising the rotating inlet cowl according to claim 1.
4. A rotating inlet cowl of a gas turbine engine, the inlet cowl including a rotation axis and comprising:
a forward cone defining a forward end of the inlet cowl,
wherein the forward end is configured to be eccentric relative to the rotation axis of the inlet cowl,
wherein the forward cone is truncated by a truncation surface defining the forward end of the inlet cowl, and
wherein the truncation surface is approximately a plane that is inclined relative to a plane orthogonal to the rotation axis of the inlet cowl.
5. The rotating inlet cowl according to claim 4, wherein the forward cone includes at least one balancing bead that includes a variable thickness along a circumferential direction, to compensate for an unbalanced mass.
6. The rotating inlet cowl according to claim 4, wherein the truncation surface is inclined relative to the plane orthogonal to the rotation axis of the inlet cowl by an angle between 1° and 15°.
7. A turbine or aircraft engine, comprising the rotating inlet cowl according to claim 4.
8. A rotating inlet cowl of a gas turbine engine, the inlet cowl including a rotation axis and comprising:
a forward cone defining a forward end of the inlet cowl,
wherein the forward end is configured to be eccentric relative to the rotation axis of the inlet cowl,
wherein the forward cone is truncated by a truncation surface defining the forward end of the inlet cowl, and
wherein the forward cone is forward from a rear shroud.
9. The rotating inlet cowl according to claim 8, wherein the forward cone includes at least one balancing bead that includes a variable thickness along a circumferential direction, to compensate for an unbalanced mass.
10. A turbine or aircraft engine, comprising the rotating inlet cowl according to claim 8.
US14/620,798 2009-03-31 2015-02-12 Rotating inlet cowl for a turbine engine, comprising an eccentric forward end Active US9243562B1 (en)

Priority Applications (1)

Application Number Priority Date Filing Date Title
US14/620,798 US9243562B1 (en) 2009-03-31 2015-02-12 Rotating inlet cowl for a turbine engine, comprising an eccentric forward end

Applications Claiming Priority (5)

Application Number Priority Date Filing Date Title
FR0952056 2009-03-31
FR0952056A FR2943726B1 (en) 2009-03-31 2009-03-31 ROTATING INPUT COVER FOR TURBOMACHINE, COMPRISING AN EXTREMITY BEFORE EXCENTREE
PCT/EP2010/054080 WO2010112453A1 (en) 2009-03-31 2010-03-29 Rotating inlet cowl for a turbine engine, including an eccentric front end
US201113262386A 2011-09-30 2011-09-30
US14/620,798 US9243562B1 (en) 2009-03-31 2015-02-12 Rotating inlet cowl for a turbine engine, comprising an eccentric forward end

Related Parent Applications (2)

Application Number Title Priority Date Filing Date
PCT/EP2010/054080 Continuation WO2010112453A1 (en) 2009-03-31 2010-03-29 Rotating inlet cowl for a turbine engine, including an eccentric front end
US13/262,386 Continuation US8984855B2 (en) 2009-03-31 2010-03-31 Rotating inlet cowl for a turbine engine, comprising an eccentric forward end

Publications (2)

Publication Number Publication Date
US20160003146A1 true US20160003146A1 (en) 2016-01-07
US9243562B1 US9243562B1 (en) 2016-01-26

Family

ID=41092060

Family Applications (2)

Application Number Title Priority Date Filing Date
US13/262,386 Active 2032-07-21 US8984855B2 (en) 2009-03-31 2010-03-31 Rotating inlet cowl for a turbine engine, comprising an eccentric forward end
US14/620,798 Active US9243562B1 (en) 2009-03-31 2015-02-12 Rotating inlet cowl for a turbine engine, comprising an eccentric forward end

Family Applications Before (1)

Application Number Title Priority Date Filing Date
US13/262,386 Active 2032-07-21 US8984855B2 (en) 2009-03-31 2010-03-31 Rotating inlet cowl for a turbine engine, comprising an eccentric forward end

Country Status (9)

Country Link
US (2) US8984855B2 (en)
EP (1) EP2414655B1 (en)
JP (1) JP5466291B2 (en)
CN (1) CN102378855B (en)
BR (1) BRPI1013370B1 (en)
CA (1) CA2756845C (en)
FR (1) FR2943726B1 (en)
RU (1) RU2529766C2 (en)
WO (1) WO2010112453A1 (en)

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2022200733A1 (en) * 2021-03-25 2022-09-29 Safran Aircraft Engines Inlet cone for an aircraft turbomachine

Families Citing this family (11)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR2989733B1 (en) * 2012-04-19 2014-05-02 Snecma TURBOMACHINE BLOWING INPUT CONE
US9969489B2 (en) * 2013-02-08 2018-05-15 General Electric Company Hybrid spinner support
US9644497B2 (en) * 2013-11-22 2017-05-09 Siemens Energy, Inc. Industrial gas turbine exhaust system with splined profile tail cone
CN103630362B (en) * 2013-11-29 2016-05-18 中国航天科技集团公司第六研究院第十一研究所 Blanking cover acting device and method for punching engine separation test
US10099772B2 (en) * 2014-10-31 2018-10-16 Hamilton Sundstrand Corporation Ice-shedding spinner for ram air turbine
US10190539B2 (en) * 2015-07-01 2019-01-29 The Boeing Company Inlet flow restrictor
US10794398B2 (en) 2015-12-18 2020-10-06 Raytheon Technologies Corporation Gas turbine engine with one piece acoustic treatment
FR3084696B1 (en) * 2018-07-31 2021-06-04 Safran Aircraft Engines IMPROVED BALANCING SYSTEM FOR AIRCRAFT TURBOMACHINE
US10975720B2 (en) 2018-07-31 2021-04-13 Safran Aircraft Engines Balancing system for an aircraft turbomachine
CN109630273B (en) * 2018-11-23 2021-04-16 中国航发沈阳黎明航空发动机有限责任公司 Aeroengine fairing based on Magnus effect
FR3097256B1 (en) 2019-06-14 2021-05-21 Safran Aircraft Engines INPUT CONE FOR AN AIRCRAFT TURBOMACHINE

Family Cites Families (39)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB155082A (en) * 1919-10-22 1920-12-16 Thomas Watson Paterson Improvements in supplementary liquid fuel tanks of the vacuum gravity type
US1867809A (en) * 1930-09-09 1932-07-19 Herbert L Chase Propeller assembly for airships
GB540711A (en) * 1940-11-05 1941-10-27 William Creighton Clay Improvements in and relating to de-icing apparatus for aeroplanes
US2401247A (en) * 1941-09-20 1946-05-28 Goodrich Co B F Spinner assembly
US2612227A (en) * 1951-01-30 1952-09-30 Curtiss Wright Corp Rotatable seal for cowlings
SU108948A1 (en) * 1956-11-17 1956-11-30 Б.Д. Попов Propeller de-icer
US3234866A (en) * 1963-11-26 1966-02-15 Northrop Corp Aircraft camera mount
DE6801232U (en) * 1968-10-08 1969-01-16 Siemens Ag SEMI-AXIAL FAN IMPELLER
BE795867A (en) * 1972-03-01 1973-06-18 Gen Electric DEVICE FOR UNIFORMISING THE FLOW OF AIR IN A GAS TURBINE
US3990814A (en) * 1975-06-25 1976-11-09 United Technologies Corporation Spinner
GB1524908A (en) * 1976-06-01 1978-09-13 Rolls Royce Gas turbine engine with anti-icing facility
GB1557856A (en) * 1977-04-20 1979-12-12 Rolls Royce Spinner or nose bullet
US4699568A (en) * 1984-06-25 1987-10-13 Hartzell Propeller Inc. Aircraft propeller with improved spinner assembly
FR2621554B1 (en) * 1987-10-07 1990-01-05 Snecma NON-ROTATING INPUT COVER OF CENTRALLY FIXED TURBOREACTOR AND TURBOREACTOR THUS EQUIPPED
US5088277A (en) * 1988-10-03 1992-02-18 General Electric Company Aircraft engine inlet cowl anti-icing system
US5214914A (en) * 1990-04-30 1993-06-01 The Johns Hopkins University Translating cowl inlet with retractable propellant injection struts
US5149251A (en) * 1990-11-15 1992-09-22 Auto Air Composites, Inc. Metal/composite spinner cone
US6167829B1 (en) * 1997-10-09 2001-01-02 Thomas G. Lang Low-drag, high-speed ship
GB9828812D0 (en) * 1998-12-29 1999-02-17 Rolls Royce Plc Gas turbine nose cone assembly
US6354538B1 (en) * 1999-10-25 2002-03-12 Rohr, Inc. Passive control of hot air injection for swirling rotational type anti-icing system
EP1237641B1 (en) * 1999-11-23 2015-09-09 Marina Ellen Marinella Pavlatos Engine with upstream and rotationally attached guard
FR2813581B1 (en) * 2000-09-06 2002-11-29 Aerospatiale Matra Airbus AIR INTAKE COVER FOR REACTION ENGINE PROVIDED WITH DEFROSTING MEANS
US6447250B1 (en) * 2000-11-27 2002-09-10 General Electric Company Non-integral fan platform
US6520742B1 (en) * 2000-11-27 2003-02-18 General Electric Company Circular arc multi-bore fan disk
US6439840B1 (en) * 2000-11-30 2002-08-27 Pratt & Whitney Canada Corp. Bypass duct fan noise reduction assembly
RU2243392C2 (en) * 2003-01-04 2004-12-27 Открытое акционерное общество "Авиадвигатель" Gas-turbine engine compressor fairing
US6887043B2 (en) * 2003-03-28 2005-05-03 General Electric Company Methods and apparatus for assembling gas turbine engines
US7063291B2 (en) * 2004-05-25 2006-06-20 Rado Kenneth S Amphibian delta wing jet aircraft
US20050274103A1 (en) * 2004-06-10 2005-12-15 United Technologies Corporation Gas turbine engine inlet with noise reduction features
FR2873751B1 (en) * 2004-07-28 2006-09-29 Snecma Moteurs Sa INPUT CONE OF A TURBOMACHINE
FR2898939B1 (en) * 2006-03-22 2008-05-09 Snecma Sa SYSTEM FOR DEFROSTING A TURBOMOTEUR INPUT CONE FOR AIRCRAFT
US7650678B2 (en) * 2006-03-30 2010-01-26 United Technologies Corporation Fabric bushing installation to repair a hole
US7730715B2 (en) * 2006-05-15 2010-06-08 United Technologies Corporation Fan frame
FR2912467B1 (en) * 2007-02-14 2009-05-15 Snecma Sa CONE OIL DEFROSTING SYSTEM BEFORE AN AIRCRAFT TURBOJET.
US20090260341A1 (en) * 2008-04-16 2009-10-22 United Technologies Corporation Distributed zoning for engine inlet ice protection
US8616854B2 (en) * 2009-03-05 2013-12-31 Rolls-Royce Corporation Nose cone assembly
US8708642B2 (en) * 2009-04-17 2014-04-29 Romeo Prasad Stable wind power turbine
US9156561B2 (en) * 2010-03-24 2015-10-13 Thomas Lucian Hurlburt System and method for preventing objects from entering the intake of a jet engine
GB201005053D0 (en) * 2010-03-26 2010-05-12 Rolls Royce Plc A gas turbine engine nose cone

Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2022200733A1 (en) * 2021-03-25 2022-09-29 Safran Aircraft Engines Inlet cone for an aircraft turbomachine
FR3121169A1 (en) * 2021-03-25 2022-09-30 Safran Aircraft Engines INLET CONE FOR AN AIRCRAFT TURBOMACHINE
US12104498B2 (en) 2021-03-25 2024-10-01 Safran Aircraft Engines Inlet cone for an aircraft turbomachine

Also Published As

Publication number Publication date
FR2943726A1 (en) 2010-10-01
US8984855B2 (en) 2015-03-24
JP2012522170A (en) 2012-09-20
US20120036827A1 (en) 2012-02-16
FR2943726B1 (en) 2014-04-25
RU2011143863A (en) 2013-05-10
CN102378855A (en) 2012-03-14
EP2414655B1 (en) 2013-06-05
EP2414655A1 (en) 2012-02-08
CA2756845C (en) 2017-01-24
BRPI1013370B1 (en) 2020-06-23
BRPI1013370A2 (en) 2016-03-29
CA2756845A1 (en) 2010-10-07
JP5466291B2 (en) 2014-04-09
US9243562B1 (en) 2016-01-26
CN102378855B (en) 2014-03-12
WO2010112453A1 (en) 2010-10-07
RU2529766C2 (en) 2014-09-27

Similar Documents

Publication Publication Date Title
US9243562B1 (en) Rotating inlet cowl for a turbine engine, comprising an eccentric forward end
CN109415940B (en) Metal leading edge for composite fan blade
EP2290207B1 (en) Cambered aero-engine inlet
US8945255B2 (en) Jet engine with deflector
CN107013268B (en) Compression fairing for jet engine exhaust
CA2833986A1 (en) Low hub-to-tip ratio fan for a turbofan gas turbine engine
EP3985240A1 (en) Nacelle short inlet
EP3159524B1 (en) Compact nacelle with contoured fan nozzle
EP3159512B1 (en) Improved crosswind performance aircraft engine spinner
CN106988889B (en) Turbine rear frame for a turbine engine
EP3590846B1 (en) Engine enclosure air inlet section
CN115675836A (en) Ductless thrust producing system
US10393019B2 (en) Assembly and method for influencing flow through a fan of a gas turbine engine
CN110015432B (en) Nacelle for an aircraft
EP3536902B1 (en) Gas turbine engine component
EP3795802B1 (en) A gas turbine engine
EP3473841A1 (en) Turbofan engine
US20240133303A1 (en) Turbine engine having a compressor with an inducer
US9915227B2 (en) Discharge system of a separated twin-flow turbojet for an aircraft, corresponding turbojet and associated design method

Legal Events

Date Code Title Description
STCF Information on status: patent grant

Free format text: PATENTED CASE

AS Assignment

Owner name: SAFRAN AIRCRAFT ENGINES, FRANCE

Free format text: CHANGE OF NAME;ASSIGNOR:SNECMA;REEL/FRAME:046479/0807

Effective date: 20160803

AS Assignment

Owner name: SAFRAN AIRCRAFT ENGINES, FRANCE

Free format text: CORRECTIVE ASSIGNMENT TO CORRECT THE COVER SHEET TO REMOVE APPLICATION NOS. 10250419, 10786507, 10786409, 12416418, 12531115, 12996294, 12094637 12416422 PREVIOUSLY RECORDED ON REEL 046479 FRAME 0807. ASSIGNOR(S) HEREBY CONFIRMS THE CHANGE OF NAME;ASSIGNOR:SNECMA;REEL/FRAME:046939/0336

Effective date: 20160803

MAFP Maintenance fee payment

Free format text: PAYMENT OF MAINTENANCE FEE, 4TH YEAR, LARGE ENTITY (ORIGINAL EVENT CODE: M1551); ENTITY STATUS OF PATENT OWNER: LARGE ENTITY

Year of fee payment: 4

MAFP Maintenance fee payment

Free format text: PAYMENT OF MAINTENANCE FEE, 8TH YEAR, LARGE ENTITY (ORIGINAL EVENT CODE: M1552); ENTITY STATUS OF PATENT OWNER: LARGE ENTITY

Year of fee payment: 8