US20110044796A1 - Fluidfoil tip vortex disruption - Google Patents

Fluidfoil tip vortex disruption Download PDF

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Publication number
US20110044796A1
US20110044796A1 US12/851,776 US85177610A US2011044796A1 US 20110044796 A1 US20110044796 A1 US 20110044796A1 US 85177610 A US85177610 A US 85177610A US 2011044796 A1 US2011044796 A1 US 2011044796A1
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Prior art keywords
fluidfoil
arrangement
fluid
tip
fluid outlet
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Abandoned
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US12/851,776
Inventor
Zahid M. Hussain
Christopher T.J. SHEAF
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Rolls Royce PLC
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Rolls Royce PLC
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Assigned to ROLLS-ROYCE PLC reassignment ROLLS-ROYCE PLC ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: SHEAF, CHRISTOPHER T. J., HUSSAIN, ZAHID M.
Publication of US20110044796A1 publication Critical patent/US20110044796A1/en
Abandoned legal-status Critical Current

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    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C11/00—Propellers, e.g. of ducted type; Features common to propellers and rotors for rotorcraft
    • B64C11/16—Blades
    • B64C11/18—Aerodynamic features
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C11/00—Propellers, e.g. of ducted type; Features common to propellers and rotors for rotorcraft
    • B64C11/46—Arrangements of, or constructional features peculiar to, multiple propellers
    • B64C11/48—Units of two or more coaxial propellers
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C23/00—Influencing air flow over aircraft surfaces, not otherwise provided for
    • B64C23/06—Influencing air flow over aircraft surfaces, not otherwise provided for by generating vortices
    • B64C23/065—Influencing air flow over aircraft surfaces, not otherwise provided for by generating vortices at the wing tips
    • 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
    • B64D27/00—Arrangement or mounting of power plants in aircraft; Aircraft characterised by the type or position of power plants
    • B64D27/02—Aircraft characterised by the type or position of power plants
    • B64D27/026—Aircraft characterised by the type or position of power plants comprising different types of power plants, e.g. combination of a piston engine and a gas-turbine
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/20—Specially-shaped blade tips to seal space between tips and stator
    • 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
    • F02K3/072—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 counter-rotating, e.g. fan rotors
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D29/00—Details, component parts, or accessories
    • F04D29/66—Combating cavitation, whirls, noise, vibration or the like; Balancing
    • F04D29/661—Combating cavitation, whirls, noise, vibration or the like; Balancing especially adapted for elastic fluid pumps
    • F04D29/667—Combating cavitation, whirls, noise, vibration or the like; Balancing especially adapted for elastic fluid pumps by influencing the flow pattern, e.g. suppression of turbulence
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D29/00—Details, component parts, or accessories
    • F04D29/66—Combating cavitation, whirls, noise, vibration or the like; Balancing
    • F04D29/68—Combating cavitation, whirls, noise, vibration or the like; Balancing by influencing boundary layers
    • F04D29/681—Combating cavitation, whirls, noise, vibration or the like; Balancing by influencing boundary layers especially adapted for elastic fluid pumps
    • F04D29/684—Combating cavitation, whirls, noise, vibration or the like; Balancing by influencing boundary layers especially adapted for elastic fluid pumps by fluid injection
    • 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
    • F05D2240/00—Components
    • F05D2240/20—Rotors
    • F05D2240/30—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor
    • F05D2240/307—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor related to the tip of a rotor blade
    • 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
    • Y02T50/10—Drag reduction
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/60—Efficient propulsion technologies, e.g. for aircraft

Definitions

  • the present invention relates to tip vortex disruption for a fluidfoil. It finds utility for aerofoils or hydrofoils in applications including, but not limited to, gas turbine engine rotors, propellers, wind turbines and helicopters.
  • fluidfoil is used throughout to encompass aerofoil and hydrofoil.
  • Rotor blades, propeller blades and aeroplane wings are known to shed vortices into their wakes, which reduces propulsive efficiency and increases the acoustic signature of the blades or wings. It is therefore desirable, particularly in relation to aircraft where there are increasingly stringent environmental requirements, to reduce or remove the vortices.
  • the vortices are predominantly shed from the tip of the fluidfoil, although other vortices are shed from the trailing edge.
  • ailerons such as ailerons, counter-rotating propellers and winglets.
  • Cordier U.S. Pat. No. 6,607,168 proposes fitting a wing with toothed platelets at its free edges to break up tip vortices.
  • Such mechanical solutions increase parasitic drag losses, add mass at the tip thereby greatly increasing bending moments and stress, and are prone to damage from small scale impacts such as ice.
  • the increased complexity of the fluidfoil increases cost for manufacture and repair.
  • Loth (US 2006/0006290) proposes active flow control to aerodynamically disrupt wing tip vortices. Perforations are provided through which air is supplied to change the boundary layer behaviour. However, this arrangement requires extra components to feed air to the perforations. Additionally, it is impractical and mechanically highly inefficient to apply Loth's teaching to rotating fluidfoils, such as propellers or rotors, since this would require fluid to be ducted into the rotating component, significantly increasing complexity and weight. A further problem with this arrangement is that the perforations are provided in a highly stressed component which therefore requires additional strengthening.
  • the present invention seeks to disrupt tip vortices of a fluidfoil and seeks to address the aforementioned problems with prior methods.
  • the present invention provides a fluidfoil tip vortex disruption arrangement comprising a fluid inlet on a pressure side of the fluidfoil; and a fluid outlet in a low pressure region of at or near the tip of the fluidfoil, whereby fluid exiting the fluid outlet inserts instability into a tip vortex to disrupt or destroy the vortex.
  • This arrangement has the advantages of increasing the propulsive efficiency of the fluidfoil and reducing its acoustic signature.
  • the fluid outlet may be located on a suction side of the fluidfoil.
  • the fluid inlet may be located towards the root of the fluidfoil or at substantially the same distance from the tip as the fluid outlet.
  • the former provides centrifugal boost whilst the latter has higher pressure recovery due to the higher tip speed relative to the root speed.
  • the fluid inlet may be angled substantially in a chordal direction. It may also be angled in a spanwise direction. This guides the fluid flow past the fluidfoil into the fluid inlet.
  • the fluid inlet may comprise a NAGA duct.
  • the fluid outlet may comprise an aperture including vanes arranged to pre-swirl the fluid prior to exiting the fluid outlet.
  • the fluid outlet may comprise a plurality of apertures arranged to pre-swirl the fluid prior to exiting the fluid outlet. The fluid exiting the fluid outlet thus forms a vortex that disrupts the tip vortex or otherwise disrupts or destroys the tip vortex.
  • the fluid outlet may comprise a plurality of apertures arranged to create an air curtain to increase the span of the fluidfoil. Beneficially this causes the over tip flow to travel further in the spanwise direction and therefore weakens the tip vortex.
  • the fluid outlet may comprise a plurality of apertures arranged to create an aerodynamic winglet at or near the tip of the fluidfoil. This causes relatively unstable fluid flows to be directed towards the tip vortex to insert instability therein, as well as increasing the aerodynamic span of the fluidfoil.
  • the fluidfoil may comprise an aerofoil or a hydrofoil.
  • the present invention provides a gas turbine engine, a propeller, a helicopter or a wind turbine comprising an arrangement as described above.
  • any application which comprises a fluidfoil may use the arrangement of the present invention with felicity.
  • FIG. 1 is a sectional side view of a gas turbine engine having contra-rotating propeller stages.
  • FIG. 2 is a schematic drawing of fluid flow over a fluidfoil tip.
  • FIG. 3 and FIG. 4 are perspective sectional views of a fluidfoil according to the present invention having first and second embodiments of a fluid inlet.
  • FIG. 5 is a sectional view of a fluidfoil according to the present invention showing fluid flow.
  • FIG. 6 is a partial perspective view of a fluidfoil according to the present invention having the fluid inlet aligned with the fluid outlet.
  • FIG. 7 shows four arrangements of the fluid outlet.
  • a twin-spooled, contra-rotating propeller gas turbine engine is generally indicated at 10 and has a principal rotational axis 9 .
  • the engine 10 comprises a core engine 11 having, in axial flow series, an air intake 12 , an intermediate pressure compressor 14 , a high-pressure compressor 15 , combustion equipment 16 , a high-pressure turbine 17 , an intermediate pressure turbine 18 , a free power (or low-pressure) turbine 19 and a core exhaust nozzle 20 .
  • a nacelle 21 generally surrounds the core engine 11 and defines the intake 12 and nozzle 20 and a core exhaust duct 22 .
  • the engine 10 also comprises two contra-rotating propeller stages 23 , 24 attached to and driven by the free power turbine 19 via shaft 26 .
  • the gas turbine engine 10 works in a conventional manner so that air entering the intake 12 is accelerated and compressed by the intermediate pressure compressor 14 and directed into the high-pressure compressor 15 where further compression takes place.
  • the compressed air exhausted from the high-pressure compressor 15 is directed into the combustion equipment 16 where it is mixed with fuel and the mixture combusted.
  • the resultant hot combustion products then expand through, and thereby drive the high-pressure, intermediate pressure and free power turbines 17 , 18 , 19 before being exhausted through the nozzle 20 to provide some propulsive thrust.
  • the high-pressure, intermediate pressure and free power turbines 17 , 18 , 19 respectively drive the high and intermediate pressure compressors 15 , 14 and the propellers 23 , 24 by suitable interconnecting shafts.
  • the propellers 23 , 24 normally provide the majority of the propulsive thrust. In the embodiments herein described the propellers 23 , 24 rotate in opposite senses so that one rotates clockwise and the other anti-clockwise around the engine's rotational axis 9 .
  • Each propeller 23 , 24 consists of a plurality of rotor blades 28 , one of which is illustrated in FIG. 2 .
  • the blade 28 has a pressure side 30 and a suction side 32 and extends spanwise from its root 34 to its tip 36 .
  • the roots 34 of a plurality of blades 28 are connected, by any suitable means known in the art, to a rotor hub to form the propeller 23 , 24 .
  • fluid flows over the tip 36 from the pressure side 30 to the suction side 32 of the blade 28 , as indicated by arrow 38 .
  • This over tip fluid flow 38 causes tip vortices 40 to form, in a clockwise sense as illustrated, and be shed behind the blade 28 as the blade 28 travels through the fluid. Since the propellers 23 , 24 rotate, the tip vortices 40 from each blade 28 actually describe a helical path.
  • FIG. 3 A first embodiment of a fluid inlet 42 of the present invention is shown in FIG. 3 .
  • the fluid inlet 42 takes the form of a slot that is inclined into the fluidfoil 28 and generally aligned with the direction of fluid flow past the fluidfoil 28 such that fluid is diverted to flow into the fluid inlet 42 , as shown by arrows 44 .
  • the fluid inlet 42 is chordally angled.
  • the fluid inlet 42 may also be angled in the spanwise direction so as to better capture the fluid flow migration from root 34 to tip 36 of the fluidfoil, blade 28 .
  • This arrangement of fluid inlet 42 is short and light.
  • An alternative, second embodiment of a fluid inlet 42 is shown in FIG. 4 and takes the form of a NACA duct or scoop. This has the conventional form wherein the inlet becomes progressively deeper and wider in the direction of inlet flow 44 .
  • This arrangement of fluid inlet 42 is longer and heavier than the inclined slot fluid inlet 42 , but has a higher pressure recovery.
  • FIG. 5 shows a blade 28 having a NACA duct fluid inlet 42 , although it is to be understood that an inclined slot could be substituted without prejudice to the function of the invention.
  • the fluid inlet 42 is located towards the root 34 of the blade 28 . In the case of a propeller 23 , 24 of a gas turbine engine 10 for aircraft propulsion, this position reduces the likelihood of ice forming around and blocking the fluid inlet 42 during flight.
  • a duct 46 extends from the fluid inlet 42 through the interior of the blade 28 . Preferably the duct 46 forms part of the structure of the blade 28 rather than being an additional component.
  • a fluid outlet 48 is provided at the distal end of the duct 46 to the fluid inlet 42 .
  • the fluid outlet 48 is formed in a low pressure region of the suction side 32 of the blade 28 , close to the tip 36 .
  • the fluid outlet 48 is arranged, as will be described in more detail with respect to FIG. 7 , to expel the fluid as an exit vortex 50 .
  • the exit vortex 50 may rotate in the opposite sense to the tip vortex 40 , so as to disrupt that flow. Alternatively it may rotate in the same sense as the tip vortex 40 provided that it destabilizes the tip vortex 40 and causes it to be disrupted or destroyed.
  • the arrangement of the present invention is passive since the rate of fluid egress from the fluid outlet 48 is determined by the rate of fluid flow over the blade 28 itself. This has advantages over an active system since it requires fewer components and does not require controlling. Furthermore, there are no moving parts in a passive system which reduces failure modes.
  • FIG. 6 shows an alternative arrangement wherein the fluid inlet 42 , shown as an inclined slot but which could be a NACA duct with equal felicity, is located towards the tip 36 of the fluidfoil, blade 28 .
  • the duct 46 is short so there are minimal ducting losses.
  • this is balanced by the higher speed of the tip 36 relative to the root 34 providing a higher pressure recovery at the inlet 42 than in the inclined slot arrangement.
  • the skilled reader will make a judgement of the optimal location of the fluid inlet 42 , and consequent length of the duct 46 , dependent on the precise application of the present invention.
  • FIG. 7 gives four embodiments of the fluid outlet 48 .
  • the fluid outlet 48 comprises a vaned aperture 52 , that is an aperture having vanes extending therein. The vanes are arranged to pre-swirl the fluid exiting the fluid outlet 48 to promote formation of the exit vortex 50 .
  • the fluid outlet 48 comprises a plurality of apertures 54 that are grouped and arranged to pre-swirl the fluid exiting the fluid outlet 48 to promote formation of the exit vortex 50 in a similar manner to the vaned aperture 52 of the first embodiment.
  • the apertures 54 are arranged in a ring and each aperture 54 is angled to promote the pre-swirling of the fluid.
  • other suitable arrangements of the apertures 54 may be contemplated and fall within the scope of this embodiment of the fluid outlet 48 of the present invention.
  • the fluid outlet 48 is in the form of a plurality of radially, spanwise directed apertures 56 .
  • the apertures 56 are arranged to direct the exiting fluid to form a curtain, which aerodynamically increases the span of the blade 28 . This causes the over tip flow 38 to travel further in the spanwise direction and therefore weakens the tip vortex 40 subsequently formed.
  • the apertures 56 are angled so as to further weaken the tip vortex 40 by directing relatively unstable fluid flows at it thus acting as an aerodynamic winglet, whilst also increasing the aerodynamic span of the fluidfoil 28 somewhat.
  • the arrangement is passive so there are no moving parts and no need for a high pressure fluid feed.
  • the duct 46 can be embedded within the fluidfoil 28 , or be integrally formed therewith, providing increased reliability and mechanical robustness.
  • the amount of fluid exiting the fluid outlet 48 increases proportionately with the increase in fluid flow over the fluidfoil 28 , which causes an increase in over tip flow 38 and thus in the intensity of tip vortex 40 formation.
  • the arrangement is self-compensating since it is more effective at disrupting or destroying tip 40 vortices when the vortices 40 are more prone to form.
  • the structural efficiency of the blade 28 is not degraded by the introduction of stress raisers, apertures and the like, or by introducing significant mass moment at the tip 36 . This is particularly the case when the fluid inlet 42 is located towards the root 34 of the blade 28 rather than towards the tip 36 . This location of the fluid inlet 42 also reduces susceptibility of the arrangement to icing since this area of an aircraft aerofoil is less prone to icing. Additionally, it is easier to provide de-icing heat to the root 34 of a blade 28 than to duct it all the way to its tip 36 .
  • the arrangement of the present invention finds utility for a variety of fluidfoils 28 in a variety of applications and technical fields.
  • fluidfoils 28 For example, aircraft wings, gas turbine engine aerofoils such as fan and compressor blades, propeller blades including for contra-rotating propellers, helicopter propeller blades including for contra-rotating propellers, wind turbine aerofoils, marine turbine hydrofoils and other hydrofoils such as stabilising fins for ships.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Aviation & Aerospace Engineering (AREA)
  • Physics & Mathematics (AREA)
  • Fluid Mechanics (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
  • Cyclones (AREA)
  • Nozzles (AREA)
  • Fertilizing (AREA)

Abstract

A fluidfoil tip vortex disruption arrangement. There is a fluid inlet on a pressure side of the fluidfoil; and a fluid outlet in a low pressure region at or near the tip of the fluidfoil. Fluid exiting the fluid outlet inserts instability into a tip vortex to disrupt or destroy the vortex.

Description

  • The present invention relates to tip vortex disruption for a fluidfoil. It finds utility for aerofoils or hydrofoils in applications including, but not limited to, gas turbine engine rotors, propellers, wind turbines and helicopters.
  • The term fluidfoil is used throughout to encompass aerofoil and hydrofoil.
  • Rotor blades, propeller blades and aeroplane wings are known to shed vortices into their wakes, which reduces propulsive efficiency and increases the acoustic signature of the blades or wings. It is therefore desirable, particularly in relation to aircraft where there are increasingly stringent environmental requirements, to reduce or remove the vortices. The vortices are predominantly shed from the tip of the fluidfoil, although other vortices are shed from the trailing edge.
  • Known methods of disrupting the tip vortices comprise mechanical tip modifications such as ailerons, counter-rotating propellers and winglets. For example, Cordier (U.S. Pat. No. 6,607,168) proposes fitting a wing with toothed platelets at its free edges to break up tip vortices. However, such mechanical solutions increase parasitic drag losses, add mass at the tip thereby greatly increasing bending moments and stress, and are prone to damage from small scale impacts such as ice. Furthermore, the increased complexity of the fluidfoil increases cost for manufacture and repair.
  • Loth (US 2006/0006290) proposes active flow control to aerodynamically disrupt wing tip vortices. Perforations are provided through which air is supplied to change the boundary layer behaviour. However, this arrangement requires extra components to feed air to the perforations. Additionally, it is impractical and mechanically highly inefficient to apply Loth's teaching to rotating fluidfoils, such as propellers or rotors, since this would require fluid to be ducted into the rotating component, significantly increasing complexity and weight. A further problem with this arrangement is that the perforations are provided in a highly stressed component which therefore requires additional strengthening.
  • The present invention seeks to disrupt tip vortices of a fluidfoil and seeks to address the aforementioned problems with prior methods.
  • Accordingly the present invention provides a fluidfoil tip vortex disruption arrangement comprising a fluid inlet on a pressure side of the fluidfoil; and a fluid outlet in a low pressure region of at or near the tip of the fluidfoil, whereby fluid exiting the fluid outlet inserts instability into a tip vortex to disrupt or destroy the vortex.
  • This arrangement has the advantages of increasing the propulsive efficiency of the fluidfoil and reducing its acoustic signature.
  • The fluid outlet may be located on a suction side of the fluidfoil. The fluid inlet may be located towards the root of the fluidfoil or at substantially the same distance from the tip as the fluid outlet. The former provides centrifugal boost whilst the latter has higher pressure recovery due to the higher tip speed relative to the root speed.
  • The fluid inlet may be angled substantially in a chordal direction. It may also be angled in a spanwise direction. This guides the fluid flow past the fluidfoil into the fluid inlet. The fluid inlet may comprise a NAGA duct.
  • The fluid outlet may comprise an aperture including vanes arranged to pre-swirl the fluid prior to exiting the fluid outlet. Alternatively the fluid outlet may comprise a plurality of apertures arranged to pre-swirl the fluid prior to exiting the fluid outlet. The fluid exiting the fluid outlet thus forms a vortex that disrupts the tip vortex or otherwise disrupts or destroys the tip vortex.
  • The fluid outlet may comprise a plurality of apertures arranged to create an air curtain to increase the span of the fluidfoil. Beneficially this causes the over tip flow to travel further in the spanwise direction and therefore weakens the tip vortex. Alternatively, the fluid outlet may comprise a plurality of apertures arranged to create an aerodynamic winglet at or near the tip of the fluidfoil. This causes relatively unstable fluid flows to be directed towards the tip vortex to insert instability therein, as well as increasing the aerodynamic span of the fluidfoil.
  • The fluidfoil may comprise an aerofoil or a hydrofoil.
  • The present invention provides a gas turbine engine, a propeller, a helicopter or a wind turbine comprising an arrangement as described above. Thus any application which comprises a fluidfoil may use the arrangement of the present invention with felicity.
  • The present invention will be more fully described by way of example with reference to the accompanying drawings, in which:
  • FIG. 1 is a sectional side view of a gas turbine engine having contra-rotating propeller stages.
  • FIG. 2 is a schematic drawing of fluid flow over a fluidfoil tip.
  • FIG. 3 and FIG. 4 are perspective sectional views of a fluidfoil according to the present invention having first and second embodiments of a fluid inlet.
  • FIG. 5 is a sectional view of a fluidfoil according to the present invention showing fluid flow.
  • FIG. 6 is a partial perspective view of a fluidfoil according to the present invention having the fluid inlet aligned with the fluid outlet.
  • FIG. 7 shows four arrangements of the fluid outlet.
    •
  • Referring to FIG. 1, a twin-spooled, contra-rotating propeller gas turbine engine is generally indicated at 10 and has a principal rotational axis 9. The engine 10 comprises a core engine 11 having, in axial flow series, an air intake 12, an intermediate pressure compressor 14, a high-pressure compressor 15, combustion equipment 16, a high-pressure turbine 17, an intermediate pressure turbine 18, a free power (or low-pressure) turbine 19 and a core exhaust nozzle 20. A nacelle 21 generally surrounds the core engine 11 and defines the intake 12 and nozzle 20 and a core exhaust duct 22. The engine 10 also comprises two contra-rotating propeller stages 23, 24 attached to and driven by the free power turbine 19 via shaft 26.
  • The gas turbine engine 10 works in a conventional manner so that air entering the intake 12 is accelerated and compressed by the intermediate pressure compressor 14 and directed into the high-pressure compressor 15 where further compression takes place. The compressed air exhausted from the high-pressure compressor 15 is directed into the combustion equipment 16 where it is mixed with fuel and the mixture combusted. The resultant hot combustion products then expand through, and thereby drive the high-pressure, intermediate pressure and free power turbines 17, 18, 19 before being exhausted through the nozzle 20 to provide some propulsive thrust. The high-pressure, intermediate pressure and free power turbines 17, 18, 19 respectively drive the high and intermediate pressure compressors 15, 14 and the propellers 23, 24 by suitable interconnecting shafts. The propellers 23, 24 normally provide the majority of the propulsive thrust. In the embodiments herein described the propellers 23, 24 rotate in opposite senses so that one rotates clockwise and the other anti-clockwise around the engine's rotational axis 9.
  • Each propeller 23, 24 consists of a plurality of rotor blades 28, one of which is illustrated in FIG. 2. The blade 28 has a pressure side 30 and a suction side 32 and extends spanwise from its root 34 to its tip 36. The roots 34 of a plurality of blades 28 are connected, by any suitable means known in the art, to a rotor hub to form the propeller 23, 24. In use, fluid flows over the tip 36 from the pressure side 30 to the suction side 32 of the blade 28, as indicated by arrow 38. This over tip fluid flow 38 causes tip vortices 40 to form, in a clockwise sense as illustrated, and be shed behind the blade 28 as the blade 28 travels through the fluid. Since the propellers 23, 24 rotate, the tip vortices 40 from each blade 28 actually describe a helical path.
  • A first embodiment of a fluid inlet 42 of the present invention is shown in FIG. 3. The fluid inlet 42 takes the form of a slot that is inclined into the fluidfoil 28 and generally aligned with the direction of fluid flow past the fluidfoil 28 such that fluid is diverted to flow into the fluid inlet 42, as shown by arrows 44. Thus the fluid inlet 42 is chordally angled. The fluid inlet 42 may also be angled in the spanwise direction so as to better capture the fluid flow migration from root 34 to tip 36 of the fluidfoil, blade 28. This arrangement of fluid inlet 42 is short and light. An alternative, second embodiment of a fluid inlet 42 is shown in FIG. 4 and takes the form of a NACA duct or scoop. This has the conventional form wherein the inlet becomes progressively deeper and wider in the direction of inlet flow 44. This arrangement of fluid inlet 42 is longer and heavier than the inclined slot fluid inlet 42, but has a higher pressure recovery.
  • FIG. 5 shows a blade 28 having a NACA duct fluid inlet 42, although it is to be understood that an inclined slot could be substituted without prejudice to the function of the invention. The fluid inlet 42 is located towards the root 34 of the blade 28. In the case of a propeller 23, 24 of a gas turbine engine 10 for aircraft propulsion, this position reduces the likelihood of ice forming around and blocking the fluid inlet 42 during flight. A duct 46 extends from the fluid inlet 42 through the interior of the blade 28. Preferably the duct 46 forms part of the structure of the blade 28 rather than being an additional component. A fluid outlet 48 is provided at the distal end of the duct 46 to the fluid inlet 42. The fluid outlet 48 is formed in a low pressure region of the suction side 32 of the blade 28, close to the tip 36. The fluid outlet 48 is arranged, as will be described in more detail with respect to FIG. 7, to expel the fluid as an exit vortex 50. The exit vortex 50 may rotate in the opposite sense to the tip vortex 40, so as to disrupt that flow. Alternatively it may rotate in the same sense as the tip vortex 40 provided that it destabilizes the tip vortex 40 and causes it to be disrupted or destroyed.
  • The arrangement of the present invention is passive since the rate of fluid egress from the fluid outlet 48 is determined by the rate of fluid flow over the blade 28 itself. This has advantages over an active system since it requires fewer components and does not require controlling. Furthermore, there are no moving parts in a passive system which reduces failure modes.
  • FIG. 6 shows an alternative arrangement wherein the fluid inlet 42, shown as an inclined slot but which could be a NACA duct with equal felicity, is located towards the tip 36 of the fluidfoil, blade 28. This means that the duct 46 is short so there are minimal ducting losses. There is no boost effect from the centrifugal force inherent in having the fluid inlet 42 near to the root 34 and a duct 46 extending spanwise through the blade 28. However, this is balanced by the higher speed of the tip 36 relative to the root 34 providing a higher pressure recovery at the inlet 42 than in the inclined slot arrangement. The skilled reader will make a judgement of the optimal location of the fluid inlet 42, and consequent length of the duct 46, dependent on the precise application of the present invention.
  • FIG. 7 gives four embodiments of the fluid outlet 48. In the first embodiment the fluid outlet 48 comprises a vaned aperture 52, that is an aperture having vanes extending therein. The vanes are arranged to pre-swirl the fluid exiting the fluid outlet 48 to promote formation of the exit vortex 50. In the second embodiment the fluid outlet 48 comprises a plurality of apertures 54 that are grouped and arranged to pre-swirl the fluid exiting the fluid outlet 48 to promote formation of the exit vortex 50 in a similar manner to the vaned aperture 52 of the first embodiment. As illustrated, the apertures 54 are arranged in a ring and each aperture 54 is angled to promote the pre-swirling of the fluid. However, other suitable arrangements of the apertures 54 may be contemplated and fall within the scope of this embodiment of the fluid outlet 48 of the present invention.
  • In the third and fourth embodiments the fluid outlet 48 is in the form of a plurality of radially, spanwise directed apertures 56. In the third embodiment the apertures 56 are arranged to direct the exiting fluid to form a curtain, which aerodynamically increases the span of the blade 28. This causes the over tip flow 38 to travel further in the spanwise direction and therefore weakens the tip vortex 40 subsequently formed. In the fourth embodiment the apertures 56 are angled so as to further weaken the tip vortex 40 by directing relatively unstable fluid flows at it thus acting as an aerodynamic winglet, whilst also increasing the aerodynamic span of the fluidfoil 28 somewhat.
  • There are many benefits and advantages to the arrangement of the present invention. The arrangement is passive so there are no moving parts and no need for a high pressure fluid feed. The duct 46 can be embedded within the fluidfoil 28, or be integrally formed therewith, providing increased reliability and mechanical robustness. The amount of fluid exiting the fluid outlet 48 increases proportionately with the increase in fluid flow over the fluidfoil 28, which causes an increase in over tip flow 38 and thus in the intensity of tip vortex 40 formation. Thus the arrangement is self-compensating since it is more effective at disrupting or destroying tip 40 vortices when the vortices 40 are more prone to form.
  • The structural efficiency of the blade 28 is not degraded by the introduction of stress raisers, apertures and the like, or by introducing significant mass moment at the tip 36. This is particularly the case when the fluid inlet 42 is located towards the root 34 of the blade 28 rather than towards the tip 36. This location of the fluid inlet 42 also reduces susceptibility of the arrangement to icing since this area of an aircraft aerofoil is less prone to icing. Additionally, it is easier to provide de-icing heat to the root 34 of a blade 28 than to duct it all the way to its tip 36.
  • By disrupting or destroying tip vortices 40 the acoustic signature of the blades 28 is reduced and the aero- or hydro-dynamic loads, through interaction with downstream blades 28, are reduced. This is particularly pertinent in the contra-rotating propeller gas turbine engine 10 application since the acoustic and aerodynamic load interaction between the front and rear propellers 23, 24 can be reduced.
  • By obviating the requirement for mechanical modifications to the fluidfoils 28, particularly to the tip 36, the parasitic losses associated with prior proposals are reduced or eliminated. Blade drag is reduced by the arrangement of the present invention so that system propulsive efficiency is enhanced.
  • The arrangement of the present invention finds utility for a variety of fluidfoils 28 in a variety of applications and technical fields. For example, aircraft wings, gas turbine engine aerofoils such as fan and compressor blades, propeller blades including for contra-rotating propellers, helicopter propeller blades including for contra-rotating propellers, wind turbine aerofoils, marine turbine hydrofoils and other hydrofoils such as stabilising fins for ships.

Claims (16)

1. A fluidfoil tip vortex disruption arrangement for a fluidfoil having a pressure side, a suction side, a tip and a root; the disruption arrangement comprising a fluid inlet on a pressure side of the fluidfoil; and a fluid outlet in a low pressure region at or near the tip of the fluidfoil, the fluid inlet coupled to the fluid outlet, whereby fluid exiting the fluid outlet during use of the fluidfoil inserts instability into a tip vortex to disrupt or destroy the vortex.
2. A disruption arrangement as claimed in claim 1 wherein the fluid outlet is located on a suction side of the fluidfoil.
3. A disruption arrangement as claimed in claim 1 wherein the fluid inlet is located towards the root of the fluidfoil.
4. A disruption arrangement as claimed in claim 1 wherein the fluid inlet is located at substantially the same distance from the tip as the fluid outlet.
5. A disruption arrangement as claimed in claim 1 wherein the fluid inlet is angled substantially in a chordal direction of the fluidfoil.
6. A disruption arrangement as claimed in claim 1 wherein the fluid inlet comprises a NACA duct.
7. A disruption arrangement as claimed in claim 1 wherein the fluid outlet comprises an aperture including vanes arranged to pre-swirl the fluid prior to exiting the fluid outlet.
8. A disruption arrangement as claimed in claim 1 wherein the fluid outlet comprises a plurality of apertures arranged to pre-swirl the fluid prior to exiting the fluid outlet.
9. A disruption arrangement as claimed in claim 1 wherein the fluid outlet comprises a plurality of apertures arranged to create an air curtain to increase the span of the fluidfoil.
10. A disruption arrangement as claimed in claim 1 wherein the fluid outlet comprises a plurality of apertures arranged to create an aerodynamic winglet at or near the tip of the fluidfoil.
11. A disruption arrangement as claimed in claim 1 wherein the fluidfoil comprises an aerofoil.
12. A disruption arrangement as claimed in claim 1 wherein the fluidfoil comprises a hydrofoil.
13. A gas turbine engine comprising an arrangement as claimed in claim 1.
14. A propeller comprising an arrangement as claimed in claim 1.
15. A helicopter comprising an arrangement as claimed in claim 1.
16. A wind turbine comprising an arrangement as claimed in claim 1.
US12/851,776 2009-08-21 2010-08-06 Fluidfoil tip vortex disruption Abandoned US20110044796A1 (en)

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Cited By (10)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20160003149A1 (en) * 2013-02-28 2016-01-07 United Technologies Corporation Method and apparatus for handling pre-diffuser airflow for cooling high pressure turbine components
US9481449B2 (en) 2011-06-20 2016-11-01 Snecma Blade, particularly variable-pitch blade, propellor comprising such blades and corresponding turbomachine
US9593582B2 (en) 2011-04-15 2017-03-14 Snecma Propulsion device having unducted counter-rotating and coaxial rotors
JP2018514678A (en) * 2015-03-10 2018-06-07 ミドル・リバー・エアクラフト・システムズ Acoustic liner for use in turbine engines
US10202865B2 (en) * 2012-10-23 2019-02-12 General Electric Company Unducted thrust producing system
US20190170008A1 (en) * 2017-12-04 2019-06-06 Rolls-Royce Plc Gas turbine engine offtake
US20190211841A1 (en) * 2018-01-05 2019-07-11 Aurora Flight Sciences Corporation Composite Fan Blades with Integral Attachment Mechanism
US11485472B2 (en) 2017-10-31 2022-11-01 Coflow Jet, LLC Fluid systems that include a co-flow jet
US11725526B1 (en) 2022-03-08 2023-08-15 General Electric Company Turbofan engine having nacelle with non-annular inlet
US11920617B2 (en) 2019-07-23 2024-03-05 Coflow Jet, LLC Fluid systems and methods that address flow separation

Families Citing this family (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR2999151B1 (en) * 2012-12-07 2017-01-27 Snecma PROPELLER BLADE FOR TURBOMACHINE

Citations (14)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2156133A (en) * 1936-06-16 1939-04-25 Theodore H Troller Propeller
US2163655A (en) * 1938-04-02 1939-06-27 Charles H Zimmerman Slotted airplane wing tip
US3596854A (en) * 1969-06-09 1971-08-03 William R Haney Jr Vortex generator for airfoil structures
US4478380A (en) * 1982-12-03 1984-10-23 Frakes James F Wing tip vortices suppressor
US5806807A (en) * 1995-10-04 1998-09-15 Haney; William R. Airfoil vortex attenuation apparatus and method
US6203269B1 (en) * 1999-02-25 2001-03-20 United Technologies Corporation Centrifugal air flow control
US6607168B1 (en) * 1998-09-25 2003-08-19 Geco As Bearing structure with reduced tip vortex
US20040197194A1 (en) * 2003-04-02 2004-10-07 Leishman John G. Rotor blade system with reduced blade-vortex interaction noise
US20050281676A1 (en) * 2004-06-16 2005-12-22 Egolf Thomas A Multi-hedral rotary wing
US20060006290A1 (en) * 2004-06-10 2006-01-12 Loth John L Vorticity cancellation at trailing edge for induced drag elimination
US20060067821A1 (en) * 2004-09-28 2006-03-30 Wadia Aspi R Methods and apparatus for aerodynamically self-enhancing rotor blades
US7354247B2 (en) * 2005-10-27 2008-04-08 General Electric Company Blade for a rotor of a wind energy turbine
US20100030634A1 (en) * 2007-04-25 2010-02-04 Pe Systems Altering Card-Issuer Interchange Categories
US20110309202A1 (en) * 2008-02-04 2011-12-22 John Jaycott Smith Wingtec Holding Limited

Patent Citations (15)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2156133A (en) * 1936-06-16 1939-04-25 Theodore H Troller Propeller
US2163655A (en) * 1938-04-02 1939-06-27 Charles H Zimmerman Slotted airplane wing tip
US3596854A (en) * 1969-06-09 1971-08-03 William R Haney Jr Vortex generator for airfoil structures
US4478380A (en) * 1982-12-03 1984-10-23 Frakes James F Wing tip vortices suppressor
US5806807A (en) * 1995-10-04 1998-09-15 Haney; William R. Airfoil vortex attenuation apparatus and method
US6607168B1 (en) * 1998-09-25 2003-08-19 Geco As Bearing structure with reduced tip vortex
US6203269B1 (en) * 1999-02-25 2001-03-20 United Technologies Corporation Centrifugal air flow control
US20040197194A1 (en) * 2003-04-02 2004-10-07 Leishman John G. Rotor blade system with reduced blade-vortex interaction noise
US6948906B2 (en) * 2003-04-02 2005-09-27 University Of Maryland Rotor blade system with reduced blade-vortex interaction noise
US20060006290A1 (en) * 2004-06-10 2006-01-12 Loth John L Vorticity cancellation at trailing edge for induced drag elimination
US20050281676A1 (en) * 2004-06-16 2005-12-22 Egolf Thomas A Multi-hedral rotary wing
US20060067821A1 (en) * 2004-09-28 2006-03-30 Wadia Aspi R Methods and apparatus for aerodynamically self-enhancing rotor blades
US7354247B2 (en) * 2005-10-27 2008-04-08 General Electric Company Blade for a rotor of a wind energy turbine
US20100030634A1 (en) * 2007-04-25 2010-02-04 Pe Systems Altering Card-Issuer Interchange Categories
US20110309202A1 (en) * 2008-02-04 2011-12-22 John Jaycott Smith Wingtec Holding Limited

Cited By (16)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US9593582B2 (en) 2011-04-15 2017-03-14 Snecma Propulsion device having unducted counter-rotating and coaxial rotors
US9481449B2 (en) 2011-06-20 2016-11-01 Snecma Blade, particularly variable-pitch blade, propellor comprising such blades and corresponding turbomachine
US10202865B2 (en) * 2012-10-23 2019-02-12 General Electric Company Unducted thrust producing system
US11988099B2 (en) 2012-10-23 2024-05-21 General Electric Company Unducted thrust producing system architecture
US10704468B2 (en) * 2013-02-28 2020-07-07 Raytheon Technologies Corporation Method and apparatus for handling pre-diffuser airflow for cooling high pressure turbine components
US20160003149A1 (en) * 2013-02-28 2016-01-07 United Technologies Corporation Method and apparatus for handling pre-diffuser airflow for cooling high pressure turbine components
JP2018514678A (en) * 2015-03-10 2018-06-07 ミドル・リバー・エアクラフト・システムズ Acoustic liner for use in turbine engines
US10961913B2 (en) 2015-03-10 2021-03-30 Mra Systems, Llc Acoustic liners for use in a turbine engine
US11485472B2 (en) 2017-10-31 2022-11-01 Coflow Jet, LLC Fluid systems that include a co-flow jet
US11987352B2 (en) 2017-10-31 2024-05-21 Coflow Jet, LLC Fluid systems that include a co-flow jet
US10900370B2 (en) * 2017-12-04 2021-01-26 Rolls-Royce Plc Gas turbine engine offtake
US20190170008A1 (en) * 2017-12-04 2019-06-06 Rolls-Royce Plc Gas turbine engine offtake
US11644046B2 (en) * 2018-01-05 2023-05-09 Aurora Flight Sciences Corporation Composite fan blades with integral attachment mechanism
US20190211841A1 (en) * 2018-01-05 2019-07-11 Aurora Flight Sciences Corporation Composite Fan Blades with Integral Attachment Mechanism
US11920617B2 (en) 2019-07-23 2024-03-05 Coflow Jet, LLC Fluid systems and methods that address flow separation
US11725526B1 (en) 2022-03-08 2023-08-15 General Electric Company Turbofan engine having nacelle with non-annular inlet

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