US20150369065A1 - Nacelle air scoop assembly - Google Patents

Nacelle air scoop assembly Download PDF

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Publication number
US20150369065A1
US20150369065A1 US14/743,584 US201514743584A US2015369065A1 US 20150369065 A1 US20150369065 A1 US 20150369065A1 US 201514743584 A US201514743584 A US 201514743584A US 2015369065 A1 US2015369065 A1 US 2015369065A1
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US
United States
Prior art keywords
air
flowpath
hood
set forth
scoop assembly
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
US14/743,584
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English (en)
Inventor
John M. Feiereisen
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 US14/743,584 priority Critical patent/US20150369065A1/en
Assigned to UNITED TECHNOLOGIES CORPORATION reassignment UNITED TECHNOLOGIES CORPORATION ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: FEIEREISEN, John M.
Publication of US20150369065A1 publication Critical patent/US20150369065A1/en
Abandoned legal-status Critical Current

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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
    • F02K3/075—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 controlling flow ratio between flows
    • 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
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C7/00—Structures or fairings not otherwise provided for
    • B64C7/02—Nacelles
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
    • F01D25/24—Casings; Casing parts, e.g. diaphragms, casing fastenings
    • 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
    • F02C6/04—Gas-turbine plants providing heated or pressurised working fluid for other apparatus, e.g. without mechanical power output
    • F02C6/06—Gas-turbine plants providing heated or pressurised working fluid for other apparatus, e.g. without mechanical power output providing compressed gas
    • 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
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
    • B64D2241/00—NACA type air intakes
    • 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
    • F05D2260/00—Function
    • F05D2260/60—Fluid transfer

Definitions

  • the present disclosure relates to gas turbine engines, and more particularly to an air scoop assembly in a nacelle of the gas turbine engine and method of operating.
  • Air scoops are known to project into a primary air flowpath for redirecting a portion of the air flow to serve a particular purpose such as cooling of components in a remote region.
  • One example of such air scoops are those that project into a bypass air flowpath of a nacelle for a turbofan engine often called ram scoops.
  • the air scoops are known to disrupt efficient airstream flow. This disruption is further aggravated where air scoop assemblies include valves that intermittently limit or prevent diversion of air flow through the air scoop assembly.
  • An air scoop assembly configured to mount to a surface defining a primary flow path, the assembly according to one, non-limiting embodiment of the present disclosure includes a hood having an irregular upstream edge and configured to project into the primary flowpath.
  • the edge is at least one vortex generator.
  • the edge is scalloped.
  • the air scoop assembly includes an air duct having an inlet defined at least by the surface and the edge.
  • the air duct defines a secondary flowpath in fluid communication with the primary flowpath and for the intermittent flow of air from the primary flowpath.
  • the air scoop assembly includes a valve configured for the intermittent isolation of the secondary flowpath.
  • the valve includes a closed position and an open position and the at least one vortex generator forms at least one vortex that breaks up any coherent shedding off of air flow from the hood when the valve is in the closed position.
  • the edge has at least one apex projecting in an upstream direction and into the primary flowpath.
  • the edge has at least one apex projecting into an upstream direction and into the primary flowpath.
  • the hood is flush with the surface.
  • the hood projects transversely into the primary flowpath and outward from the surface.
  • the surface is carried by a nacelle.
  • the apex projects upstream by a downstream-most portion of the edge by a distance that is about twice a thickness of the hood.
  • the at least one apex is spaced from a next adjacent apex by a distance that is about one fifth to about one half a height of the hood.
  • a nacelle for a gas turbine engine includes a fan cowling concentrically disposed to an engine axis and surrounding a fan section of the gas turbine engine; a core cowling concentrically disposed to the engine axis and located downstream of the fan section with an annular air bypass flowpath defined between and by the fan and core cowlings; and a hood formed to the core cowling and projecting into the bypass flowpath, the hood including a leading edge that includes at least one apex projecting in an upstream direction and acting as at least one vortex generator.
  • the nacelle includes an air duct having an inlet defined at least by the core cowling and the edge; and wherein the air duct defines a secondary flowpath in fluid communication with the bypass flowpath and for the intermittent flow of air from the bypass flowpath.
  • the nacelle includes a valve in the air duct for the intermittent isolation of the secondary flowpath; and wherein the valve includes a closed position and an open position and the at least one vortex generator forms at least one vortex that breaks up any coherent shedding off of air flow from the hood when the valve is in the closed position.
  • the hood is flush with the core cowling.
  • the hood projects transversely into the bypass flowpath and radially outward from the core cowling.
  • a method of operating an air scoop assembly includes the steps of substantially closing the air scoop assembly having a hood that projects into a primary air flowpath; and forming at least one air vortex that stems from an irregular leading edge of a hood of the air scoop assembly and extends in a substantially downstream direction thereby minimizing disruption of airstreams in the primary air flowpath.
  • FIG. 1 is a schematic cross section of a gas turbine engine
  • FIG. 2 is a partial perspective view of the engine viewing in an upstream direction
  • FIG. 3 is a side view of an air scoop assembly
  • FIG. 4 is a perspective view of a hood of the air scoop assembly.
  • FIG. 5 is a cross section of a leading edge of the hood taken along line 5 - 5 of FIG. 4 .
  • FIG. 1 schematically illustrates a gas turbine engine 20 disclosed as a two-spool turbo fan that generally incorporates a fan section 22 , a compressor section 24 , a combustor section 26 and a turbine section 28 .
  • the fan section 22 drives air along a bypass or primary flowpath 29 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 .
  • a turbofan depicted as a turbofan 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 engine architecture such as turbojets, turboshafts, three-spool turbofans, land-based turbine engines, and others.
  • the engine 20 generally includes a low spool 30 and a high spool 32 mounted for rotation about an engine axis A via several bearing structures 38 and relative to a static engine case 36 .
  • the low spool 30 generally includes an inner shaft 40 that interconnects a fan 42 of the fan section 22 , a low pressure compressor 44 (“LPC”) of the compressor section 24 and a low pressure turbine 46 (“LPT”) of the turbine section 28 .
  • 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 may be 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”) of the compressor section 24 and a high pressure turbine 54 (“HPT”) of the turbine section 28 .
  • a combustor 56 of the combustor section 26 is arranged between the HPC 52 and the HPT 54 .
  • the inner shaft 40 and the outer shaft 50 are concentric and rotate about the engine axis A. Core airflow is compressed by the LPC 44 then the HPC 52 , mixed with the fuel and burned in the combustor 56 , then expanded over the HPT 54 and the LPT 46 .
  • the LPT 46 and HPT 54 rotationally drive the respective low spool 30 and high spool 32 in response to the expansion.
  • 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 2.3:1, and in another example is greater than about 2.5:1.
  • the geared turbofan enables operation of the low spool 30 at higher speeds that can increase the operational efficiency of the LPC 44 and LPT 46 and render increased pressure in a fewer number of stages.
  • a pressure ratio associated with the LPT 46 is pressure measured prior to the inlet of the LPT 46 as related to the pressure at the outlet of the LPT 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 the LPC 44 ; and the LPT 46 has a pressure ratio that is greater than about five (5:1). It should be understood; however, that the above parameters are only exemplary of one example 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 flowpath 29 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 (10,668 meters). This flight condition, with the gas turbine engine 20 at its best fuel consumption, is also known as Thrust Specific Fuel consumption (TSFC).
  • TSFC Thrust Specific Fuel consumption
  • Fan Pressure Ratio is the pressure ratio across a blade of the fan section 22 without consideration of the effect of a fan exit guide vane assembly 58 located downstream of the fan 42 (also see FIG. 2 ).
  • the low Fan Pressure Ratio according to one, non-limiting, example of the gas turbine engine 20 is less than 1.45:1.
  • Low Corrected Fan Tip Speed is the actual fan tip speed divided by an industry standard temperature correction of (T/518.7 0.5 ), where “T” represents the ambient temperature in degrees Rankine.
  • the Low Corrected Fan Tip Speed according to one non-limiting example of the gas turbine engine 20 is less than about 1150 fps (351 m/s).
  • a nacelle assembly 60 of the turbine engine 20 has core cowling 62 , a fan cowling 64 and a pylon 66 .
  • the core and fan cowlings 62 , 64 are generally concentric to the engine axis A with the core cowling 62 generally supporting the core engine and surrounding the engine sections 24 , 26 , 28 .
  • the fan cowling 64 is spaced radially outward from the core cowling 62 and surrounds the fan 42 and guide vane assembly 58 .
  • the annular bypass flow path 29 is defined between and by the core and fan cowlings 62 , 64 .
  • the pylon 66 may be attached to both cowlings 62 , 64 and generally supports and attaches the entire engine 20 to, for example, an aircraft (not shown).
  • an annular engine cavity 68 is located radially between and may be defined by the core cowling 62 and the inner engine case 36 .
  • Various engine components may be generally located in the cavity 68 and may require cooling by a ventilation or air cooling system 70 through a series of tube and hoses directed to the component and/or otherwise ventilate the cavity 68 to prevent the accumulation of fumes.
  • Cooling system 70 may include an air scoop assembly 72 having a hood 74 mounted to the core cowling 62 (as one example) and projecting into the bypass flowpath 29 for receipt of a secondary airflow.
  • the amount of secondary airflow is dependent mainly upon the pressure difference between the bypass flowpath 29 and the ambient air, and other systems may include control logic with associated flow control valve(s).
  • an air cooling system 70 is an Active Clearance Control (ACC) cooling system 70 that assists in the control of a blade tip clearance between the engine case 36 and the tips of the rotating blades of the turbine section 28 . More specifically, the ACC cooling system 70 adjusts cooling flow to the engine case 36 thereby controlling thermal expansion of the case relative to centrifugal and thermal expansion of the turbine rotor thereby minimizing and/or controlling the blade tip clearance throughout varying engine operating conditions.
  • ACC Active Clearance Control
  • the air scoop assembly 72 of the ACC cooling system 70 includes the hood 74 , a duct 76 defining a secondary air flowpath 78 , and a flow control device or at least one valve 80 that intersects the duct 76 for controlling the amount of secondary air flow.
  • the hood 74 has a contoured or scalloped leading edge 82 that, with a radially outward facing surface 84 of the core cowling 62 , defines an inlet 86 of the secondary air flowpath 78 . It is further contemplated and understood that the air scoop assembly 72 may be generally mounted on any surface that defines at least in-part a primary air flowpath and where the retrieval of a secondary air flow is required.
  • bypass air flows generally in an axially downstream direction (see the bypass airstreams shown as arrows 90 in FIG. 3 ) along the bypass flowpath 29 .
  • a portion of this bypass air is scooped-up by the hood 74 of the air scoop assembly 72 and flows through or along the secondary air flowpath 78 .
  • no (or minimal) air vortices are created about the hood which could hinder efficient flow of the bypass airstreams 90 . Absent the features of the present disclosure, with the flow control device closed or substantially closed, no (or minimal) secondary air flows through the secondary air flowpath and the hood of the air scoop assembly may function more as an airstream obstruction in the bypass flowpath.
  • air vortices or disturbances may be created that undesirably disrupt the bypass airstreams 90 , reduce airflow efficiency, and lead to undesired resonance that may produce significant unsteady pressures resulting in elevated noise and elevated unsteady stresses in surrounding structures, if not properly controlled.
  • the irregular or scalloped shape of the leading edge 82 of the hood 74 generally functions as a plurality of vortex generators that create, controlled, air vortices 92 generally when the flow control device 80 is closed, and which stem from the leading edge 82 of the hood 74 and generally co-extend in the downstream direction with the bypass airstreams 90 thereby minimizing any disruption of the bypass air flow. That is, with the flow control device 80 generally closed, there is a continuous shedding of streamwise vorticity.
  • the “continuous shedding” is desirable over discontinuous or periodic shedding because it eliminates the production of undesired resonance that produce excessive noise in the duct and stress upon surrounding structure.
  • the irregular or non-uniform edge 82 avoids the tendancy to shed a full-width coherent vortex and avoids any coupling with the natural frequency of the cowlings 62 , 64 .
  • the non-uniform edge 82 acts as vortex generators in the flow spilling around the edge, shedding continuous, streamwise, vorticity and reducing any flow separation in the bypass air flowpath 29 downstream of the hood 74 .
  • the irregular shape of the leading edge may not be scalloped but may take the form of any variety of shapes that may produce air vortices as described.
  • the irregular shape of the leading edge 82 may be a plurality of scallops 94 (i.e. each scallop generally being one vortex generator).
  • Each scallop 94 meets the next adjacent scallop at an apex or convex portion 96 that generally projects and substantially faces in an upstream direction, and which contributes toward the shedding of the air vortices 92 .
  • Each apex 96 is spaced from the next adjacent apex by a distance 100 .
  • Each scallop 94 also has a concave portion 98 that substantially faces in the upstream direction and is generally spaced axially (with respect to the engine axis A) from the apex 96 by a distance 102 .
  • the hood 74 has a width 104 measured between the joinder(s) of the hood 74 to the surface 84 of the core cowling 62 , and a height 106 that is generally the maximum projection of the hood into the bypass air flowpath 29 (i.e. maximum radial distance from the surface 84 to the hood 74 ).
  • the width 104 is substantially greater than the height 106
  • the height 106 is generally two to five times greater than the distance 100 between apexes 96 .
  • the hood 74 has a general thickness 108 and the concave portion 98 may have a parabolic shaped cross section that generally begins where the hood has the thickness 108 and projects upstream to a vertex by a distance 110 that may be substantially equal to thickness 108 .
  • the apex or convex portion 96 may have a parabolic shaped cross section similar to the concave portion 98 but generally more pointed (i.e. more tapered).
  • the distance 102 between the apex 96 and the concave portion 98 may be about equal to twice the thickness 108 .

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Aviation & Aerospace Engineering (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
US14/743,584 2014-06-18 2015-06-18 Nacelle air scoop assembly Abandoned US20150369065A1 (en)

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Application Number Priority Date Filing Date Title
US14/743,584 US20150369065A1 (en) 2014-06-18 2015-06-18 Nacelle air scoop assembly

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US201462013880P 2014-06-18 2014-06-18
US14/743,584 US20150369065A1 (en) 2014-06-18 2015-06-18 Nacelle air scoop assembly

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Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP3388648A1 (de) 2017-04-11 2018-10-17 Rolls-Royce plc Einlasskanal
US20210040897A1 (en) * 2019-08-07 2021-02-11 United Technologies Corporation External turning vane for ifs-mounted secondary flow systems
US10934937B2 (en) 2016-07-19 2021-03-02 Raytheon Technologies Corporation Method and apparatus for variable supplemental airflow to cool aircraft components

Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4070827A (en) * 1976-05-03 1978-01-31 General Electric Company Method and apparatus for limiting ingestion of debris into the inlet of a gas turbine engine
US20050163963A1 (en) * 2004-01-12 2005-07-28 Munro Alexander S. Method and apparatus for reducing drag and noise for a vehicle
US20060060721A1 (en) * 2004-03-30 2006-03-23 Phillip Watts Scalloped leading edge advancements
US20070130912A1 (en) * 2005-12-08 2007-06-14 General Electric Company Shrouded turbofan bleed duct
US7600963B2 (en) * 2005-08-22 2009-10-13 Viryd Technologies Inc. Fluid energy converter
US20100223905A1 (en) * 2009-03-04 2010-09-09 Rolls-Royce Deutschland Ltd & Co Kg Scoop of a running-gap control system of an aircraft gas turbine

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Publication number Priority date Publication date Assignee Title
US9194330B2 (en) * 2012-07-31 2015-11-24 United Technologies Corporation Retrofitable auxiliary inlet scoop
US9108737B2 (en) * 2012-08-24 2015-08-18 United Technologies Corporation Nacelle scoop inlet

Patent Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4070827A (en) * 1976-05-03 1978-01-31 General Electric Company Method and apparatus for limiting ingestion of debris into the inlet of a gas turbine engine
US20050163963A1 (en) * 2004-01-12 2005-07-28 Munro Alexander S. Method and apparatus for reducing drag and noise for a vehicle
US20060060721A1 (en) * 2004-03-30 2006-03-23 Phillip Watts Scalloped leading edge advancements
US7600963B2 (en) * 2005-08-22 2009-10-13 Viryd Technologies Inc. Fluid energy converter
US20070130912A1 (en) * 2005-12-08 2007-06-14 General Electric Company Shrouded turbofan bleed duct
US20100223905A1 (en) * 2009-03-04 2010-09-09 Rolls-Royce Deutschland Ltd & Co Kg Scoop of a running-gap control system of an aircraft gas turbine

Non-Patent Citations (1)

* Cited by examiner, † Cited by third party
Title
GB 1144804 A; United Aircraft Corporation; March 12, 1969; Original Document *

Cited By (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US10934937B2 (en) 2016-07-19 2021-03-02 Raytheon Technologies Corporation Method and apparatus for variable supplemental airflow to cool aircraft components
EP3388648A1 (de) 2017-04-11 2018-10-17 Rolls-Royce plc Einlasskanal
US20210040897A1 (en) * 2019-08-07 2021-02-11 United Technologies Corporation External turning vane for ifs-mounted secondary flow systems
US11022047B2 (en) * 2019-08-07 2021-06-01 Raytheon Technologies Corporation External turning vane for IFS-mounted secondary flow systems

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Effective date: 20150617

STCB Information on status: application discontinuation

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