GB2372019A - Turbofan engine negatively scarfed nacelle for uniform flow to the fan - Google Patents

Turbofan engine negatively scarfed nacelle for uniform flow to the fan Download PDF

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Publication number
GB2372019A
GB2372019A GB0103367A GB0103367A GB2372019A GB 2372019 A GB2372019 A GB 2372019A GB 0103367 A GB0103367 A GB 0103367A GB 0103367 A GB0103367 A GB 0103367A GB 2372019 A GB2372019 A GB 2372019A
Authority
GB
United Kingdom
Prior art keywords
intake
nacelle
gas turbine
turbine engine
axis
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Withdrawn
Application number
GB0103367A
Other versions
GB0103367D0 (en
Inventor
Adam Macgregor Bagnall
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.)
Rolls Royce PLC
Original Assignee
Rolls Royce PLC
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 Rolls Royce PLC filed Critical Rolls Royce PLC
Priority to GB0103367A priority Critical patent/GB2372019A/en
Publication of GB0103367D0 publication Critical patent/GB0103367D0/en
Publication of GB2372019A publication Critical patent/GB2372019A/en
Withdrawn legal-status Critical Current

Links

Classifications

    • 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
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C7/00—Structures or fairings not otherwise provided for
    • B64C7/02—Nacelles
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENTS OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
    • B64D29/00—Power-plant nacelles, fairings, or cowlings
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENTS OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
    • B64D33/00—Arrangements in aircraft of power plant parts or auxiliaries not otherwise provided for
    • B64D33/02—Arrangements in aircraft of power plant parts or auxiliaries not otherwise provided for of combustion air intakes
    • 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
    • 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
    • F02C7/045—Air intakes for gas-turbine plants or jet-propulsion plants having provisions for noise suppression
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENTS OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
    • B64D33/00—Arrangements in aircraft of power plant parts or auxiliaries not otherwise provided for
    • B64D33/02—Arrangements in aircraft of power plant parts or auxiliaries not otherwise provided for of combustion air intakes
    • B64D2033/0266—Arrangements in aircraft of power plant parts or auxiliaries not otherwise provided for of combustion air intakes specially adapted for particular type of power plants
    • B64D2033/0286—Arrangements in aircraft of power plant parts or auxiliaries not otherwise provided for of combustion air intakes specially adapted for particular type of power plants for turbofan engines
    • 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
    • F05D2270/00—Control
    • F05D2270/01—Purpose of the control system
    • F05D2270/10—Purpose of the control system to cope with, or avoid, compressor flow instabilities
    • F05D2270/101—Compressor surge or stall
    • F05D2270/102—Compressor surge or stall caused by working fluid flow velocity profile distortion
    • 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

Abstract

A turbofan gas turbine engine (10, fig 1) includes a nacelle 34 having an intake 12 comprising a non-axis-symmetric flared intake 36, an intake throat 38 and an axis-symmetric diffuser 40. The intake throat 38 is circular in cross-section and is arranged substantially coaxially with, and substantially perpendicularly to the axis X of the gas turbine engine 10. The axis-symmetric diffuser 40 is arranged substantially coaxially with the axis X of the gas turbine engine 10. The non-axis-symmetric flared intake 36 has a longer flared lip 36 at a bottom region 36A of the nacelle than the flared lip 36 at other regions 36B,(36C,36D fig 3) of the nacelle 34 to form a negatively scarfed intake. The intake throat 38 is arranged a distance of 0.5 to 1.0 times the fan diameter T upstream of the fan 14. A more uniform airflow to the fan is thus provided.

Description

A GAS TURBINE ENGINE NACELLE The present invention relates to a gas turbine engine nacelle and in particular to a turbofan gas turbine engine nacelle.
A turbofan gas turbine engine nacelle intake is required to supply the fan of the turbofan gas turbine engine with favourably conditioned air during all operational conditions of the turbofan gas turbine engine, irrespective of the aircraft environment and aircraft attitude, whether the aircraft is in flight or on the ground. The nacelle intake may also be required to absorb noise generated by the gas turbine engine.
Prior art nacelle intakes are designed to minimise the pressure loss of the intake at the maximum incidence the aircraft experiences in flight and at the maximum crosswind conditions on the ground. In prior art nacelle intakes good pressure recovery at high incidence is achieved by attempting to ensure that the air flow remains attached to the bottom lip of the nacelle intake by drooping down the nacelle throat relative to the axis of the fan of the turbofan gas turbine engine. The drooping down of the intake throat has been achieved either by providing an angled straight diffuser between the intake throat and the fan or by providing a curved diffuser between the intake throat and the fan. A contraction is provided upstream of the intake throat to ensure the airflow remains attached at crosswind and static conditions. The contraction varies around the circumference of the nacelle intake according to the worst case airflow expected at each circumferential location.
A problem with these nacelle intakes is that the flow path in the nacelle immediately upstream of the fan is not symmetric. This results in an asymmetric flow of air to the fan causing it to operate away from its optimum operating point and hence there is a loss of efficiency.
In addition the non-uniform aerodynamic operation of the
fan generates a non-uniform flow of air to the components downstream of the fan, causing a further loss of efficiency. Also the non-uniform aerodynamic operation of the fan generates extra fan noise. A further problem is that the non-uniform flow to the fan results in non-uniform loading of the fan blades, which may excite vibrations in the fan blades and thereby generates excessive mechanical stresses in the fan blades.
The asymmetry of the flow path in the nacelle immediately upstream of the fan makes it impossible to provide a uniform and axis-symmetric noise attenuation treatment in the intake diffuser upstream of the fan.
The drooping nacelle intake may direct additional noise preferentially downwards towards the ground.
Accordingly the present invention seeks to provide a novel gas turbine engine having a nacelle which reduces, preferably overcomes, the above mentioned problems.
Accordingly the present invention provides a gas turbine engine comprising a nacelle, the nacelle comprising an intake at a first end of the nacelle and an exhaust at a second end of the nacelle, the intake comprising a nonaxis-symmetric flared intake, an intake throat and an axissymmetric diffuser, the intake throat is circular in crosssection, the intake throat is arranged substantially coaxially with the axis of the gas turbine engine, the intake throat is arranged substantially perpendicularly to the axis of the gas turbine engine, the non-axis-symmetric flared intake has a longer flared lip at a first region of the nacelle than the flared lip at other regions of the nacelle to form a negatively scarfed intake, the axissymmetric diffuser is arranged substantially coaxially with the axis of the gas turbine engine.
Preferably the first region of the flared intake lip is at the bottom of the flared intake lip.
Preferably the gas turbine engine comprises a turbofan gas turbine engine having a fan at its upstream end.
Preferably the intake throat is arranged a distance of 0. 5 to 1. 0 times the fan diameter upstream of the fan.
The present invention also provides a gas turbine engine nacelle comprising an intake at a first end of the nacelle and an exhaust at a second end of the nacelle, the intake comprising a non-axis-symmetric flared intake, an intake throat and an axis-symmetric diffuser, the intake throat is circular in cross-section, the intake throat is arranged substantially coaxially with the axis of the axissymmetric diffuser, the intake throat is arranged substantially perpendicularly to the axis of the axissymmetric diffuser, the non-axis-symmetric flared intake has a longer flared lip at a first region of the nacelle than the flared lip at other regions of the nacelle to form a negatively scarfed intake.
The present invention will be more fully described by way of example with reference to the accompanying drawings in which: Figure 1 is a partially cut away view of a turbofan gas turbine engine having a nacelle according to the present invention.
Figure 2 is an enlarged vertical cross-section through the intake of the turbofan gas turbine engine nacelle shown in figure 1.
Figure 3 is an enlarged horizontal cross-section through the intake of the turbofan gas turbine engine nacelle shown in figure 1.
Figure 4 is an enlarged alternative vertical crosssection through the intake of the turbofan gas turbine engine nacelle shown in figure 1.
Figure 5 is an enlarged alternative horizontal crosssection through the intake of the turbofan gas turbine engine nacelle shown in figure 1.
A turbofan gas turbine engine 10, as shown in figure 1, comprises in axial flow series an intake 12, a fan section 14, a compressor section 16, a combustion section
18, a turbine section 20 and an exhaust 22. The fan section 14 comprises a fan disc 24 carrying a plurality of circumferentially spaced radially extending fan blades 26. The fan disc 24 and fan blades 26 are surrounded by a fan casing 28. The fan casing 28 is mounted from the core casing 30 by a plurality of radially extending fan outlet guide vanes 32. The fan section 14 is driven by a turbine in the turbine section 20 via a shaft (not shown). The compressor section 16 is driven by a turbine in the turbine section 20 by a shaft (not shown). The whole of the turbofan gas turbine engine 10 is placed within a nacelle 34.
The nacelle 34, as shown more clearly in figure 2 and 3, has an intake 12 at its upstream end and an exhaust 25 at its downstream end. The nacelle 34 intake 12 comprises in flow series a flared intake lip 36, an intake throat 38 and a diffuser 40 upstream of the fan section 14 of the turbofan gas turbine engine 10. The flared intake lip 36 forms a contraction for the supply of air to the intake throat 38. The diffuser 40 is arranged to diffuse the air from the intake throat 38 to the fan section 14. The nacelle 34 intake 12 as shown in figures 2 and 3 is designed for high incidence operation of the turbofan gas turbine engine 10.
The flared intake lip 36 is designed at a number of circumferential locations to guide the air smoothly at the most adverse conditions at each circumferential location into the intake throat 38. The flared intake lip 36 has a bottom portion 36A and a top portion 36B. The bottom portion 36A is axially longer and larger than the top portion 36B, as shown in figure 2, so that the bottom portion 36A of the flared intake lip 36 extends further upstream than the top portion 36B to cater for high positive incidence, or high angles of attack, of the oncoming air in flight. This is commonly called negative scarfing of the intake. The bottom portion 36A of the
flared intake lip 36 has an axial length A, from the most upstream point of the lip 36 to the throat 38, and radial dimension B, at the intake throat 38, sufficient to ensure that the airflow does not separate from its internal surface 37A.
The plane Y of the intake throat 38 is arranged perpendicularly to the axis X of the fan section 14 of the turbofan gas turbine engine 10. The intake throat 38 is circular in cross-section and is arranged coaxially with the axis X of the fan section 14 of the turbofan gas turbine engine 10. The intake throat 38 has a sufficient area to supply the maximum airflow requirement of the turbofan gas turbine engine 10. The intake throat 38 is positioned a sufficient distance Z axially upstream of the fan section 14 to produce a uniform air flow field to the fan section 14. The axial distance Z is preferably between 0.5 and 1.0 times the tip diameter T of the fan blades 26 of the fan section 14.
The shape of the bottom portion 36A of the flared lip 36 is integrated with the shape of the intake between the intake throat 38 and the fan section 14 to form the diffuser 40. The diffuser 40 guides the airflow through the diffuser 40 to the fan section 14 without separation of the airflow. The diffuser 40 is defined as a surface of revolution of the line 40A between the bottom portion 36A of the flared intake lip 36 and the fan section 14. Thus the diffuser 40 is circular in cross-section and is coaxial with the axis X of the fan section 14 of the turbofan gas turbine engine 10.
The exterior surface 39 of intake lip 36 is designed to satisfy the external aerodynamic requirements, according to conventional practice.
The side portions 36C and 36D of the flared lip 36, on the horizontal plane through the turbofan gas turbine engine 10, are arranged to have an axial length C and a radial dimension D, at the intake throat 38, sufficient to
ensure the airflow does not separate from the internal surfaces 37C and 37D.
The nacelle 34B, as shown more clearly in figure 4 and 5, has an intake 12 at its upstream end and an exhaust 25 at its downstream end. The nacelle 34B intake 12 comprises in flow series a flared intake lip 36, an intake throat 38 and a diffuser 40 upstream of the fan section 14 of the turbofan gas turbine engine 10. The flared intake lip 36 forms a contraction for the supply of air to the intake throat 38. The diffuser 40 is arranged to diffuse the air from the intake throat 38 to the fan section 14. The nacelle 34B intake 12 as shown is designed for crosswind operation of the turbofan gas turbine engine 10.
The flared intake lip 36 is designed at a number of circumferential locations to guide the air smoothly at the most adverse conditions at each circumferential location into the intake throat 38. The flared intake lip 36 is designed specifically for crosswind operation, when the air is approaching from one side of the intake 12. The flared intake lip 36 has side portions 36C and 36D which are designed with sufficient axial length C, from the most upstream point of the lip 36 in this horizontal plane, and radial dimension D, at the throat 38, sufficient to ensure that the airflow does not separate from their internal surfaces 37C and 37D. The side portions 36C and 36D are identical in shape in this example.
The plane Y of the intake throat 38 is arranged perpendicularly to the axis X of the fan section 14 of the turbofan gas turbine engine 10. The intake throat 38 is circular in cross-section and is arranged coaxially with the axis X of the fan section 14 of the turbofan gas turbine engine 10. The intake throat 38 has a sufficient area to supply the maximum airflow requirement of the turbofan gas turbine engine 10. The intake throat 38 is positioned a sufficient distance Z axially upstream of the fan section 14 to produce a uniform air flow field to the
fan section 14. The axial distance Z is preferably between 0.5 and 1.0 times the tip diameter T of the fan blades 26 of the fan section 14.
The shape of the side portions 36C and 36D of the flared lip 36 are integrated with the shape of the intake between the intake throat 38 and the fan section 14 to form the diffuser 40. The diffuser 40 guides the airflow through the diffuser 40 to the fan section 14 without separation of the airflow. The diffuser 40 is defined as a surface of revolution of the line 40C, or 40D, between the side portions 36C and 36D of the flared intake lip 36 and the fan section 14. Thus the diffuser 40 is circular in cross-section and is coaxial with the axis X of the fan section 14 of the turbofan gas turbine engine 10.
The exterior surface 39 of intake lip 36 is designed to satisfy the external aerodynamic requirements, according to conventional practice.
The flared intake lip 36 has a bottom portion 36A and a top portion 36B and the bottom portion 36A is axially longer and larger than the top portion 36B, as shown in figure 4, so that the bottom portion 36A of the flared intake lip 36 extends axially further upstream than the top portion 36B. This is commonly called negative scarfing of the intake. The bottom portion 36A of the flared intake lip 36 has an axial length A, from the most upstream point of the lip 36 to the throat 38, and radial dimension B, at the intake throat 38, sufficient to ensure that the airflow does not separate from its internal surface 37A.
The advantages of the present invention are due to the fan throat being arranged perpendicularly and coaxially with the fan section of the turbofan gas turbine engine and the axis-symmetric diffuser being arranged coaxially with the fan section of the turbofan gas turbine engine. This arrangement results in a more uniform flow of air to the fan section with more uniform airflow angles onto the fan at all circumferential locations. This ensures that the
fan operates at its optimum operating point and efficiency at all circumferential locations, so increasing the overall efficiency of the fan. The fan therefore will generate a uniform airflow to components downstream of the fan, allowing them to operate with reduced aerodynamic performance losses. The circumferentially uniform operation of the fan removes extra noise generated by a non-uniform aerodynamic operation of the fan. The circumferentially uniform operation of the fan removes additional stresses, which would be generated by a nonuniform aerodynamic operation of the fan.
The perpendicular throat and the axis-symmetric diffuser ensure that the noise generated by the fan is not scattered by any asymmetry in the walls upstream of the fan and hence reduces the overall engine noise. An axissymmetric acoustic liner may be utilised to maximise sound absorption without scattering of the fan noise and hence reduce overall engine noise.
The absence of the conventional drooping and the use of the negative scarfing also have advantages. The longer bottom flared lip, in both embodiments, provides a degree of shielding of the fan noise to people on the ground and preferentially deflects noise upwards away from people on the ground. The flared side lip may be designed to give optimum crosswind operation. The flared bottom lip may be designed to give optimum high incidence operation. The diffuser may be designed either to give optimum crosswind operation or optimum high incidence operation. The shorter top flared lip reduces weight for a given high incidence performance and noise reduction performance.

Claims (10)

  1. Claims :1. A gas turbine engine including a nacelle, the nacelle comprising an intake at a first end of the nacelle and an exhaust at a second end of the nacelle, the intake comprising a non-axis-symmetric flared intake, an intake throat and an axis-symmetric diffuser, the intake throat is circular in cross-section, the intake throat is arranged substantially coaxially with the axis of the gas turbine engine, the intake throat is arranged substantially perpendicularly to the axis of the gas turbine engine, the axis-symmetric diffuser is arranged substantially coaxially with the axis of the gas turbine engine and the non-axissymmetric flared intake has a longer flared lip at a first region of the nacelle than the flared lip at other regions of the nacelle to form a negatively scarfed intake.
  2. 2. A gas turbine engine as claimed in claim 1 wherein the first region of the flared intake lip is at the bottom of the flared intake lip.
  3. 3. A gas turbine engine as claimed in claim 1 or clam 2 wherein the gas turbine engine comprises a turbofan gas turbine engine having a fan at its upstream end.
  4. 4. A gas turbine engine as claimed in claim 3 wherein the intake throat is arranged a distance of 0.5 to 1.0 times the fan diameter upstream of the fan.
  5. 5. A gas turbine engine substantially as hereinbefore described with reference to figures 1,2 and 3 of the accompanying drawings.
  6. 6. A gas turbine engine substantially as hereinbefore described with reference to figures 1,4 and 5 of the accompanying drawings.
  7. 7. A gas turbine engine nacelle comprising an intake at a first end of the nacelle and an exhaust at a second end of the nacelle, the intake comprising a non-axis-symmetric flared intake, an intake throat and an axis-symmetric diffuser, the intake throat is circular in cross-section, the intake throat is arranged substantially coaxially with
    the axis of the axis-symmetric diffuser, the intake throat is arranged substantially perpendicularly to the axis of the axis-symmetric diffuser, the non-axis-symmetric flared intake has a longer flared lip at a first region of the nacelle than the flared lip at other regions of the nacelle to form a negatively scarfed intake.
  8. 8. A gas turbine engine nacelle as claimed in claim 7 wherein the first region of the flared intake lip is at the bottom of the flared intake lip.
  9. 9. A gas turbine engine nacelle substantially as hereinbefore described with reference to figures 2 and 3 of the accompanying drawings.
  10. 10. A gas turbine engine nacelle substantially as hereinbefore described with reference to figures 4 and 5 of the accompanying drawings.
GB0103367A 2001-02-10 2001-02-10 Turbofan engine negatively scarfed nacelle for uniform flow to the fan Withdrawn GB2372019A (en)

Priority Applications (1)

Application Number Priority Date Filing Date Title
GB0103367A GB2372019A (en) 2001-02-10 2001-02-10 Turbofan engine negatively scarfed nacelle for uniform flow to the fan

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
GB0103367A GB2372019A (en) 2001-02-10 2001-02-10 Turbofan engine negatively scarfed nacelle for uniform flow to the fan

Publications (2)

Publication Number Publication Date
GB0103367D0 GB0103367D0 (en) 2001-03-28
GB2372019A true GB2372019A (en) 2002-08-14

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GB0103367A Withdrawn GB2372019A (en) 2001-02-10 2001-02-10 Turbofan engine negatively scarfed nacelle for uniform flow to the fan

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Cited By (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP1813529A1 (en) * 2006-01-27 2007-08-01 Snecma Air intake of a turbofan engine
GB2461718A (en) * 2008-07-10 2010-01-13 Rolls Royce Plc An aircraft propulsion arrangement having a single fan located within a curved guide path
EP2151381A2 (en) 2008-08-06 2010-02-10 Honeywell International Inc. Ducted fan lip shaping for an unmanned aerial vehicle
EP3181863A1 (en) * 2015-12-18 2017-06-21 United Technologies Corporation Gas turbine engine with minimized inlet distortion
EP3225789A1 (en) * 2016-04-01 2017-10-04 Rolls-Royce Deutschland Ltd & Co KG Engine assembly with fan casing and intake diffuser
US11015550B2 (en) 2012-12-20 2021-05-25 Raytheon Technologies Corporation Low pressure ratio fan engine having a dimensional relationship between inlet and fan size
US11286811B2 (en) 2012-12-20 2022-03-29 Raytheon Technologies Corporation Low pressure ratio fan engine having a dimensional relationship between inlet and fan size

Citations (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO1999061316A2 (en) * 1998-04-14 1999-12-02 The Boeing Company Biplanar scarfed nacelle inlet

Patent Citations (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO1999061316A2 (en) * 1998-04-14 1999-12-02 The Boeing Company Biplanar scarfed nacelle inlet

Cited By (15)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US7735776B2 (en) 2006-01-27 2010-06-15 Snecma Air inlet for a turbofan engine
FR2896771A1 (en) * 2006-01-27 2007-08-03 Snecma Sa DOUBLE FLOW TURBOJET AIR INTAKE
EP1813529A1 (en) * 2006-01-27 2007-08-01 Snecma Air intake of a turbofan engine
GB2461718A (en) * 2008-07-10 2010-01-13 Rolls Royce Plc An aircraft propulsion arrangement having a single fan located within a curved guide path
US8240597B2 (en) 2008-08-06 2012-08-14 Honeywell International Inc. UAV ducted fan lip shaping
EP2151381A3 (en) * 2008-08-06 2012-05-09 Honeywell International Inc. Ducted fan lip shaping for an unmanned aerial vehicle
EP2151381A2 (en) 2008-08-06 2010-02-10 Honeywell International Inc. Ducted fan lip shaping for an unmanned aerial vehicle
US11015550B2 (en) 2012-12-20 2021-05-25 Raytheon Technologies Corporation Low pressure ratio fan engine having a dimensional relationship between inlet and fan size
US11286811B2 (en) 2012-12-20 2022-03-29 Raytheon Technologies Corporation Low pressure ratio fan engine having a dimensional relationship between inlet and fan size
US11781447B2 (en) 2012-12-20 2023-10-10 Rtx Corporation Low pressure ratio fan engine having a dimensional relationship between inlet and fan size
US11781505B2 (en) 2012-12-20 2023-10-10 Rtx Corporation Low pressure ratio fan engine having a dimensional relationship between inlet and fan size
EP3181863A1 (en) * 2015-12-18 2017-06-21 United Technologies Corporation Gas turbine engine with minimized inlet distortion
US20170175626A1 (en) * 2015-12-18 2017-06-22 United Technologies Corporation Gas turbine engine with minimized inlet distortion
EP3181863B1 (en) 2015-12-18 2019-04-10 United Technologies Corporation Gas turbine engine with minimized inlet distortion
EP3225789A1 (en) * 2016-04-01 2017-10-04 Rolls-Royce Deutschland Ltd & Co KG Engine assembly with fan casing and intake diffuser

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