WO2014122568A1 - System for preventing icing on an aircraft surface operationally exposed to air - Google Patents

System for preventing icing on an aircraft surface operationally exposed to air Download PDF

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Publication number
WO2014122568A1
WO2014122568A1 PCT/IB2014/058751 IB2014058751W WO2014122568A1 WO 2014122568 A1 WO2014122568 A1 WO 2014122568A1 IB 2014058751 W IB2014058751 W IB 2014058751W WO 2014122568 A1 WO2014122568 A1 WO 2014122568A1
Authority
WO
WIPO (PCT)
Prior art keywords
exposed
air
aircraft
electrode portion
intermediate portion
Prior art date
Application number
PCT/IB2014/058751
Other languages
French (fr)
Inventor
Emanuele MERLO
Alessandro Gurioli
Ermanno MAGNOLI
Giordano MATTIUZZO
Roberto PERTILE
Original Assignee
Alenia Aermacchi S.Pa.
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 Alenia Aermacchi S.Pa. filed Critical Alenia Aermacchi S.Pa.
Priority to US14/765,744 priority Critical patent/US10227923B2/en
Priority to EP14710970.6A priority patent/EP2953852B1/en
Publication of WO2014122568A1 publication Critical patent/WO2014122568A1/en

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Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
    • F02C7/04—Air intakes for gas-turbine plants or jet-propulsion plants
    • F02C7/047—Heating to prevent icing
    • 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
    • B64D15/00—De-icing or preventing icing on exterior surfaces of aircraft
    • B64D15/12—De-icing or preventing icing on exterior surfaces of aircraft by electric heating
    • 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
    • B64D15/00—De-icing or preventing icing on exterior surfaces of aircraft
    • 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
    • B64D15/00—De-icing or preventing icing on exterior surfaces of aircraft
    • B64D15/12—De-icing or preventing icing on exterior surfaces of aircraft by electric heating
    • B64D15/14—De-icing or preventing icing on exterior surfaces of aircraft by electric heating controlled cyclically along length of surface
    • 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
    • B64D15/00—De-icing or preventing icing on exterior surfaces of aircraft
    • B64D15/20—Means for detecting icing or initiating de-icing
    • 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
    • B64D15/00—De-icing or preventing icing on exterior surfaces of aircraft
    • B64D15/20—Means for detecting icing or initiating de-icing
    • B64D15/22—Automatic initiation by icing detector
    • 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
    • F01D25/02—De-icing means for engines having icing phenomena
    • H—ELECTRICITY
    • H05—ELECTRIC TECHNIQUES NOT OTHERWISE PROVIDED FOR
    • H05H—PLASMA TECHNIQUE; PRODUCTION OF ACCELERATED ELECTRICALLY-CHARGED PARTICLES OR OF NEUTRONS; PRODUCTION OR ACCELERATION OF NEUTRAL MOLECULAR OR ATOMIC BEAMS
    • H05H1/00—Generating plasma; Handling plasma
    • H05H1/24—Generating plasma
    • H05H1/2406—Generating plasma using dielectric barrier discharges, i.e. with a dielectric interposed between the electrodes
    • H05H1/2439—Surface discharges, e.g. air flow control
    • 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
    • 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
    • B64D2033/0233—Arrangements in aircraft of power plant parts or auxiliaries not otherwise provided for of combustion air intakes comprising de-icing means
    • 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
    • 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
    • F05D2270/17—Purpose of the control system to control boundary layer
    • F05D2270/172—Purpose of the control system to control boundary layer by a plasma generator, e.g. control of ignition
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F15—FLUID-PRESSURE ACTUATORS; HYDRAULICS OR PNEUMATICS IN GENERAL
    • F15D—FLUID DYNAMICS, i.e. METHODS OR MEANS FOR INFLUENCING THE FLOW OF GASES OR LIQUIDS
    • F15D1/00—Influencing flow of fluids
    • F15D1/002—Influencing flow of fluids by influencing the boundary layer
    • F15D1/0065—Influencing flow of fluids by influencing the boundary layer using active means, e.g. supplying external energy or injecting fluid
    • F15D1/0075—Influencing flow of fluids by influencing the boundary layer using active means, e.g. supplying external energy or injecting fluid comprising electromagnetic or electrostatic means for influencing the state of the fluid, e.g. for ionising the fluid or for generating a plasma
    • H—ELECTRICITY
    • H05—ELECTRIC TECHNIQUES NOT OTHERWISE PROVIDED FOR
    • H05H—PLASMA TECHNIQUE; PRODUCTION OF ACCELERATED ELECTRICALLY-CHARGED PARTICLES OR OF NEUTRONS; PRODUCTION OR ACCELERATION OF NEUTRAL MOLECULAR OR ATOMIC BEAMS
    • H05H1/00—Generating plasma; Handling plasma
    • H05H1/24—Generating plasma
    • H05H1/2406—Generating plasma using dielectric barrier discharges, i.e. with a dielectric interposed between the electrodes
    • H05H1/2418—Generating plasma using dielectric barrier discharges, i.e. with a dielectric interposed between the electrodes the electrodes being embedded in the dielectric
    • H—ELECTRICITY
    • H05—ELECTRIC TECHNIQUES NOT OTHERWISE PROVIDED FOR
    • H05H—PLASMA TECHNIQUE; PRODUCTION OF ACCELERATED ELECTRICALLY-CHARGED PARTICLES OR OF NEUTRONS; PRODUCTION OR ACCELERATION OF NEUTRAL MOLECULAR OR ATOMIC BEAMS
    • H05H1/00—Generating plasma; Handling plasma
    • H05H1/24—Generating plasma
    • H05H1/2406—Generating plasma using dielectric barrier discharges, i.e. with a dielectric interposed between the electrodes
    • H05H1/2425—Generating plasma using dielectric barrier discharges, i.e. with a dielectric interposed between the electrodes the electrodes being flush with the dielectric
    • H—ELECTRICITY
    • H05—ELECTRIC TECHNIQUES NOT OTHERWISE PROVIDED FOR
    • H05H—PLASMA TECHNIQUE; PRODUCTION OF ACCELERATED ELECTRICALLY-CHARGED PARTICLES OR OF NEUTRONS; PRODUCTION OR ACCELERATION OF NEUTRAL MOLECULAR OR ATOMIC BEAMS
    • H05H1/00—Generating plasma; Handling plasma
    • H05H1/24—Generating plasma
    • H05H1/2406—Generating plasma using dielectric barrier discharges, i.e. with a dielectric interposed between the electrodes
    • H05H1/2437—Multilayer systems

Definitions

  • TITLE System for preventing icing on an aircraft surface operationally exposed to air
  • the present invention relates to a system for preventing icing on an aircraft surface operationally exposed to air.
  • Ice formation and deposition on exposed surfaces, in particular aerodynamic surfaces, of an aircraft is a problem of primary importance as regards the full operability, the safety and the efficiency of an aircraft during the use or operation thereof, whether the same is intended for civil or military applications.
  • pneumatic, electric or mechanic systems which utilize power to heat the surfaces affected by the problem, so as to avoid that the humidity or water lapping said exposed surface might enter the solid state, or to mechanically remove ice after it has formed.
  • FIGS. 1 to 3 are schematic sectional views of three exemplificative embodiments of a system for preventing icing on an aircraft surface operationally exposed to air, in accordance with the present invention
  • FIG. 4 is a longitudinal sectional view of an engine nacelle having a surface exposed to air, equipped with one of the systems shown in Figures 1 to 3;
  • FIG. 5 is a top view of an aircraft equipped with a plurality of surfaces exposed to air and fitted with one of the systems shown in the preceding Figures.
  • numeral 10 designates as a whole an exemplificative embodiment of a system according to the present invention.
  • System 10 comprises a plasma actuator, designated as a whole by reference numeral 12, intended for application onto a surface exposed to air of an aircraft (said surface being schematically identified by reference numeral 28) and arranged for generating at least one plasma discharge D for inducing a flow (schematically represented by dashed arrow F in Figure 1) of ionized hot-air particles towards said exposed surface.
  • the plasma discharge generated by using actuator 12 triggers ionization in the particles contained in the gas present in the surrounding air. In order to cause this ionization, heat is generated in the air which is then yielded to the gas particles.
  • the intense electric field generated by the plasma can interact with the latter's ionized particles, thereby producing mass forces that create said (oriented) flow of ionized air particles, accelerating them in a direction that converges towards the exposed surface.
  • the heat generated by ionization is directed towards said exposed surface 28 on which system 10 has been positioned, thus preventing icing from occurring on such surfaces.
  • said plasma actuator 12 is of the dielectric barrier discharge (DBD) type, which generates so-called cold plasma.
  • DBD dielectric barrier discharge
  • cold plasma comes from the fact that the temperature of the ionized particles thus generated is typically in the range of 40°C to 100°C, so that the presence of plasma has no destructive effects upon the surrounding material (also including, in particular, the dielectric barrier incorporated in actuator 12) .
  • actuator 12 comprises: - an intermediate portion 14 of dielectric material, at least one exposed electrode portion 16, 18 positioned on an outer side 20 of the intermediate portion 14, at least a part thereof being exposed to air, and
  • covered electrode portion 22 corresponds to the outer covering of the aircraft, which will necessarily be made of metal
  • electrode portions 16, 18, 22 are electrically connectable to a high-voltage electric power generator 24 and are adapted to be energized by said electric power to generate said plasma discharge D between them.
  • actuator 12 comprises at least one pair of exposed electrode portions 16, 18.
  • At least one exposed electrode portion 18 is positioned on top of intermediate portion 14.
  • said exposed electrode portion 18 is prominent, e.g. it protrudes like a relief or an "embossment", relative to the intermediate portion 14.
  • the top surface and the side surface of the exposed electrode portion 18 are operationally exposed to air, whereas the bottom surface of the exposed electrode portion 18 remains in contact with intermediate portion 14, and is thus electrically insulated from the air.
  • At least one exposed electrode portion 16 is partially embedded in intermediate portion 14.
  • the exposed electrode portion 16 is flush with outer side 20 of intermediate portion 14.
  • the top surface of the exposed electrode portion 16 is exposed to air, whereas its side surface and its bottom surface are electrically insulated from the air, in that said bottom surface is in contact with intermediate portion 14.
  • covered electrode portion 22 is at least partially positioned under an inner surface or inner side 24, opposite to the outer surface or outer side 20, of the portion of dielectric material 14. In this embodiment, covered electrode portion 22 is completely arranged under inner side 24.
  • covered electrode portion 22 is at least partly constituted by exposed surface 28 of the aircraft, which is generally electrically conductive, in particular being made of metallic material.
  • exposed surface 28 that at least partly constitutes covered electrode portion 22 may be a portion of the covering of the aircraft.
  • generator 24 is configured for generating an alternating-current electric voltage, e.g. having a frequency of the order of magnitude of kHz and/or an amplitude of the order of magnitude of kV.
  • generator 24 can originate a voltage which is modulated as concerns at least one of frequency and amplitude, for the purpose of optimally adapting the operation of system 10 to the desired specifications in terms of performance and energy consumption.
  • said voltage may be amplitude-modulated according to an ON/OFF pulsing, and possibly frequency-modulated as concerns at least one of :
  • system 10 may comprise a control module (not shown) which can automatically control generator 24 according to predetermined criteria, allowing for automatic intervention only when particular operating conditions occur (e.g. if the temperature of said surface 28 operationally exposed to air is lower than a predetermined threshold value) .
  • system 10 may also include at least one sensor (not shown), e.g. a temperature sensor, for providing the control module with a signal indicating the occurrence of the above-mentioned operating conditions, e.g. that the temperature of surface 28 operationally exposed to air has gone below said threshold value.
  • the control module can then control generator 24 on the basis of said indicative signal, according to predetermined criteria (e.g. by activating the generator or by appropriately adjusting the voltage it generates, in particular through amplitude and/or frequency modulation, as a function of temperature or other parameters detected by the sensor) .
  • system 110 comprises an insulating portion 26 of dielectric material, positioned under covered electrode portion 22, to be applied onto the surface operationally exposed to air of the aircraft.
  • said surface is designated by reference numeral 28.
  • covered electrode portion 22 is interposed, in particular in a "sandwich" fashion, between intermediate portion 14, which is located in the outermost position, and insulating portion 26, which is located in the innermost position.
  • covered electrode portion 22 coincides with surface 28 operationally exposed to air of the aircraft, with no electric insulation between them.
  • system 210 may be manufactured by rolling.
  • system 10, 110, 210 can be structured as a rolled section processed into a continuous strip in which intermediate portion 14 is obtained, which includes a layer of dielectric material onto which the following elements are applied:
  • covered electrode portion 22 particularly in the form of a layer of electrically conductive material.
  • Said rolled section can then be easily applied onto the desired exposed surfaces of the aircraft.
  • system 10 may also be fabricated as a "multilayer" element, which can also be obtained by means of different stratification processes, wherein said intermediate portion 14 is initially deposited (with electrode portion 22) onto the exposed surface of the aircraft by using deposition techniques which are per se known to a man skilled in the art.
  • said at least one exposed electrode portion 16, 18 may be made as a longitudinal band of electrically conductive material, which is subsequently applied onto intermediate portion 14.
  • said band may either be applied over said intermediate portion 14 (which preferably includes a layer of dielectric material) or be partially embedded in the latter.
  • a plurality of exposed electrode portions 16, 18 may be conceived in the form of bands of electrically conductive material, which may be arranged, by way of example, in longitudinal rows, preferably parallel to one another, on/in intermediate portion 14.
  • Said at least one exposed electrode portion 16, 18 can be connected to said intermediate portion 14 by glueing or by using other deposition techniques which are per se known to a man skilled in the art.
  • FIG. 4 there is shown an engine nacelle N on which one or more of systems 10, 110, 210 are mounted, each one of them being designed in accordance with one of the previously described embodiments .
  • engine nacelle N comprises a "cowling" C having an aerodynamic shape.
  • cowling C has a substantially tubular shape, e.g. a barrellike shape, with an annular profile preferably having a prevalently airfoil longitudinal section.
  • cowling C has an air "inlet" I from which a through cavity develops, which extends in a substantially axial direction.
  • cowling C includes a radially external side wall and a radially internal side wall, which are referred to in the art as “outer barrel” and “inner barrel”, respectively.
  • devices for acoustic attenuation are mounted on the radially internal side wall of the cowling for reducing the noise which is typically generated during the operation of the aircraft on which said engine nacelle N is mounted.
  • a jet-type propulsion assembly is accommodated, which is adapted to receive air from the air inlet and accelerate it in order to generate a thrust.
  • the propulsion assembly is designed as a turbo-fan structure, i.e. it comprises:
  • cowling C and engine E there is an annular region or by-pass duct BD, which is structured for conveying that air fraction which flows through fan F without however entering engine E.
  • engine nacelle N includes system 10, 110, 210, which is positioned at air inlet I, in particular covering lip L and preferably extending over exposed surfaces of at least one radially external side wall (“outer barrel”) and one radially internal side wall (“inner barrel”) adjacent to lip L.
  • an aircraft e.g. a fixed-wing aircraft, on which one or more of systems 10, 110, 210 are mounted, each one of them being designed in accordance with one of the previously described embodiments.
  • the aircraft comprises a pair of fixed wings W, wherein at least one of them has at least one surface operationally exposed to air, onto which at least one of systems 10, 110, 210 has been applied.
  • system 10, 110, 210 is positioned in the proximity of the leading edge of wing W.
  • system 10, 110, 210 can be positioned on the exposed surface of the flaps of wing W, in particular of at least one of front flap WS (commonly referred to in the industry as "nose droop or slat”) and the rear flap WF (commonly referred to in the industry as "flap”) .
  • front flap WS commonly referred to in the industry as "nose droop or slat”
  • rear flap WF commonly referred to in the industry as "flap”
  • both front flap WS and the rear flap are equipped with a system 10, 110, 210 designed in compliance with the present invention .
  • At least one system 10, 110, 210 is arranged at the tail of the aircraft, i.e. on the horizontal empennages and on the vertical empennage, particularly on the leading edges of the tail itself.

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  • Engineering & Computer Science (AREA)
  • Aviation & Aerospace Engineering (AREA)
  • Physics & Mathematics (AREA)
  • Plasma & Fusion (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Spectroscopy & Molecular Physics (AREA)
  • Plasma Technology (AREA)
  • Vehicle Cleaning, Maintenance, Repair, Refitting, And Outriggers (AREA)

Abstract

A system (10, 110, 210) for preventing icing on an aircraft surface comprises a plasma actuator (12) which is to be applied onto an aircraft surface (28) operationally exposed to air and which is arranged for generating at least one plasma discharge (D) for inducing a flow (F) of ionized hot-air particles towards said surface (28).

Description

TITLE: "System for preventing icing on an aircraft surface operationally exposed to air"
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DESCRIPTION
Technical field
The present invention relates to a system for preventing icing on an aircraft surface operationally exposed to air.
Background art
Ice formation and deposition on exposed surfaces, in particular aerodynamic surfaces, of an aircraft is a problem of primary importance as regards the full operability, the safety and the efficiency of an aircraft during the use or operation thereof, whether the same is intended for civil or military applications.
In order to prevent this phenomenon from arising, pneumatic, electric or mechanic systems are used which utilize power to heat the surfaces affected by the problem, so as to avoid that the humidity or water lapping said exposed surface might enter the solid state, or to mechanically remove ice after it has formed.
Summary of the invention
It is one object of the present invention to provide a system for preventing icing on an aircraft surface operationally exposed to air, said system being improved over those in accordance with the prior art.
According to the present invention, this and other objects are achieved through a system having the technical features set out in the appended independent claim.
It is to be understood that the appended claims are an integral part of the technical teachings provided in the following detailed description of the invention. In particular, the appended dependent claims define some preferred embodiments of the present invention which include some optional technical features.
Brief description of the drawings
Further features and advantages of the present invention will become apparent from the following detailed description, which is supplied by way of non-limiting example with reference to the annexed drawings, wherein:
- Figures 1 to 3 are schematic sectional views of three exemplificative embodiments of a system for preventing icing on an aircraft surface operationally exposed to air, in accordance with the present invention;
- Figure 4 is a longitudinal sectional view of an engine nacelle having a surface exposed to air, equipped with one of the systems shown in Figures 1 to 3; and
- Figure 5 is a top view of an aircraft equipped with a plurality of surfaces exposed to air and fitted with one of the systems shown in the preceding Figures.
Detailed description of the invention
With particular reference to Figures 1 to 3, there are shown different exemplificative embodiments of a system for preventing icing on an aircraft surface operationally exposed to air, in accordance with the present invention;
With particular reference to Figure 1, numeral 10 designates as a whole an exemplificative embodiment of a system according to the present invention.
System 10 comprises a plasma actuator, designated as a whole by reference numeral 12, intended for application onto a surface exposed to air of an aircraft (said surface being schematically identified by reference numeral 28) and arranged for generating at least one plasma discharge D for inducing a flow (schematically represented by dashed arrow F in Figure 1) of ionized hot-air particles towards said exposed surface.
More in detail, the plasma discharge generated by using actuator 12 triggers ionization in the particles contained in the gas present in the surrounding air. In order to cause this ionization, heat is generated in the air which is then yielded to the gas particles.
Furthermore, the intense electric field generated by the plasma can interact with the latter's ionized particles, thereby producing mass forces that create said (oriented) flow of ionized air particles, accelerating them in a direction that converges towards the exposed surface.
Therefore, the heat generated by ionization is directed towards said exposed surface 28 on which system 10 has been positioned, thus preventing icing from occurring on such surfaces.
Of course, by changing the geometries and operating parameters of plasma actuator 12, it is possible to obtain different force fields and associated mass flows, so as to ensure that the particle flow is controlled within the boundary layer on surface 28 exposed to air which is affected by plasma generation.
In this embodiment, said plasma actuator 12 is of the dielectric barrier discharge (DBD) type, which generates so-called cold plasma. As is known in the art, the term "cold plasma" comes from the fact that the temperature of the ionized particles thus generated is typically in the range of 40°C to 100°C, so that the presence of plasma has no destructive effects upon the surrounding material (also including, in particular, the dielectric barrier incorporated in actuator 12) .
Preferably, actuator 12 comprises: - an intermediate portion 14 of dielectric material, at least one exposed electrode portion 16, 18 positioned on an outer side 20 of the intermediate portion 14, at least a part thereof being exposed to air, and
- a covered electrode portion 22 not operationally exposed to air because of intermediate portion 14; in this embodiment, covered electrode portion 22 corresponds to the outer covering of the aircraft, which will necessarily be made of metal
wherein electrode portions 16, 18, 22 are electrically connectable to a high-voltage electric power generator 24 and are adapted to be energized by said electric power to generate said plasma discharge D between them.
Therefore, when generator 24 is activated, plasma discharges D are generated in the proximity of (at least one) exposed electrode portion 16, 18, but not on covered electrode portion 22, since the latter is insulated from the air by intermediate portion 14.
In this embodiment, actuator 12 comprises at least one pair of exposed electrode portions 16, 18.
Advantageously but not necessarily, at least one exposed electrode portion 18 is positioned on top of intermediate portion 14. In other words, said exposed electrode portion 18 is prominent, e.g. it protrudes like a relief or an "embossment", relative to the intermediate portion 14. In particular, the top surface and the side surface of the exposed electrode portion 18 are operationally exposed to air, whereas the bottom surface of the exposed electrode portion 18 remains in contact with intermediate portion 14, and is thus electrically insulated from the air.
Advantageously but not necessarily, at least one exposed electrode portion 16 is partially embedded in intermediate portion 14. In the illustrated embodiment, the exposed electrode portion 16 is flush with outer side 20 of intermediate portion 14. Particularly, the top surface of the exposed electrode portion 16 is exposed to air, whereas its side surface and its bottom surface are electrically insulated from the air, in that said bottom surface is in contact with intermediate portion 14.
In particular, covered electrode portion 22 is at least partially positioned under an inner surface or inner side 24, opposite to the outer surface or outer side 20, of the portion of dielectric material 14. In this embodiment, covered electrode portion 22 is completely arranged under inner side 24.
Optionally, covered electrode portion 22 is at least partly constituted by exposed surface 28 of the aircraft, which is generally electrically conductive, in particular being made of metallic material. In the embodiment shown in Figure 1, exposed surface 28 that at least partly constitutes covered electrode portion 22 may be a portion of the covering of the aircraft.
In this illustrated embodiment, generator 24 is configured for generating an alternating-current electric voltage, e.g. having a frequency of the order of magnitude of kHz and/or an amplitude of the order of magnitude of kV. Particularly, generator 24 can originate a voltage which is modulated as concerns at least one of frequency and amplitude, for the purpose of optimally adapting the operation of system 10 to the desired specifications in terms of performance and energy consumption. For example, said voltage may be amplitude-modulated according to an ON/OFF pulsing, and possibly frequency-modulated as concerns at least one of :
- relative duration between the ON pulse and the OFF pulse, and
- total duration of one complete pulsing period (i.e. the sum of the duration of the ON pulse and the duration of the OFF pulse) .
In the illustrated embodiment, the system may comprise a control module (not shown) which can automatically control generator 24 according to predetermined criteria, allowing for automatic intervention only when particular operating conditions occur (e.g. if the temperature of said surface 28 operationally exposed to air is lower than a predetermined threshold value) . For example, system 10 may also include at least one sensor (not shown), e.g. a temperature sensor, for providing the control module with a signal indicating the occurrence of the above-mentioned operating conditions, e.g. that the temperature of surface 28 operationally exposed to air has gone below said threshold value. In such a case, the control module can then control generator 24 on the basis of said indicative signal, according to predetermined criteria (e.g. by activating the generator or by appropriately adjusting the voltage it generates, in particular through amplitude and/or frequency modulation, as a function of temperature or other parameters detected by the sensor) .
With particular reference to Figure 2, there is shown a further exemplificative embodiment of a system 110 designed in accordance with the present invention.
Those parts or elements which are similar to - or which perform the same function as - those of the above- described embodiment have been assigned the same reference numerals. For simplicity, the description of such parts or elements will not be repeated below, and reference will be made to the above description of the previous embodiment.
One of the differences that can be noticed in comparison with the previously described embodiment is that system 110 comprises an insulating portion 26 of dielectric material, positioned under covered electrode portion 22, to be applied onto the surface operationally exposed to air of the aircraft. In Figure 2 said surface is designated by reference numeral 28. In other systems, covered electrode portion 22 is interposed, in particular in a "sandwich" fashion, between intermediate portion 14, which is located in the outermost position, and insulating portion 26, which is located in the innermost position.
Therefore, in the embodiment previously shown in Figure 1, covered electrode portion 22 coincides with surface 28 operationally exposed to air of the aircraft, with no electric insulation between them. Instead, in the embodiment shown in Figure 2 there is a separation between covered electrode portion 22 and said surface 28, in that they are two distinct elements which are electrically insulated from each other by insulating portion 26.
With particular reference to Figure 3, there is shown a further exemplificative embodiment of a system 210 designed in accordance with the present invention.
Those parts or elements which are similar to - or which perform the same function as - those of the above- described embodiments have been assigned the same alphanumeric references. For simplicity, the description of such parts or elements will not be repeated below, and reference will be made to the above description of the previous embodiments.
One of the differences that can be noticed in comparison with the previously described embodiments is that covered electrode portion 22 is embedded in intermediate portion 14.
By way of example, system 210 may be manufactured by rolling.
In the embodiments shown in Figures 1 to 3, system 10, 110, 210 can be structured as a rolled section processed into a continuous strip in which intermediate portion 14 is obtained, which includes a layer of dielectric material onto which the following elements are applied:
on one side, the exposed electrode portion(s) 16, 18, and
- on the other side, covered electrode portion 22, particularly in the form of a layer of electrically conductive material.
Said rolled section can then be easily applied onto the desired exposed surfaces of the aircraft.
As an alternative, system 10 may also be fabricated as a "multilayer" element, which can also be obtained by means of different stratification processes, wherein said intermediate portion 14 is initially deposited (with electrode portion 22) onto the exposed surface of the aircraft by using deposition techniques which are per se known to a man skilled in the art.
Advantageously but not necessarily, said at least one exposed electrode portion 16, 18 may be made as a longitudinal band of electrically conductive material, which is subsequently applied onto intermediate portion 14. For example, said band may either be applied over said intermediate portion 14 (which preferably includes a layer of dielectric material) or be partially embedded in the latter. In particular, a plurality of exposed electrode portions 16, 18 may be conceived in the form of bands of electrically conductive material, which may be arranged, by way of example, in longitudinal rows, preferably parallel to one another, on/in intermediate portion 14.
Said at least one exposed electrode portion 16, 18 can be connected to said intermediate portion 14 by glueing or by using other deposition techniques which are per se known to a man skilled in the art.
With particular reference to Figure 4, there is shown an engine nacelle N on which one or more of systems 10, 110, 210 are mounted, each one of them being designed in accordance with one of the previously described embodiments .
In a per se known manner, engine nacelle N comprises a "cowling" C having an aerodynamic shape. In particular, cowling C has a substantially tubular shape, e.g. a barrellike shape, with an annular profile preferably having a prevalently airfoil longitudinal section.
More in detail, cowling C has an air "inlet" I from which a through cavity develops, which extends in a substantially axial direction.
The structure of cowling C includes a radially external side wall and a radially internal side wall, which are referred to in the art as "outer barrel" and "inner barrel", respectively.
In particular, devices for acoustic attenuation, also referred to in the art as soundproof panels, are mounted on the radially internal side wall of the cowling for reducing the noise which is typically generated during the operation of the aircraft on which said engine nacelle N is mounted.
In cowling C, within the through cavity and downstream of inlet I, a jet-type propulsion assembly is accommodated, which is adapted to receive air from the air inlet and accelerate it in order to generate a thrust. In particular, the propulsion assembly is designed as a turbo-fan structure, i.e. it comprises:
- a fan F for accelerating the air flow coming in through inlet I, and
- an engine E positioned downstream of fan F.
In the illustrated embodiment, it can be seen that between cowling C and engine E there is an annular region or by-pass duct BD, which is structured for conveying that air fraction which flows through fan F without however entering engine E.
Of course, the structure and operation of said propulsion assembly, as diagrammatically shown in Figure 1, are per se known in the industry, and for simplicity they will not be described in detail below. In said drawing, the air flows that go through the engine nacelle are shown, by way of example, by means of a series of arrows.
In this embodiment, engine nacelle N includes system 10, 110, 210, which is positioned at air inlet I, in particular covering lip L and preferably extending over exposed surfaces of at least one radially external side wall ("outer barrel") and one radially internal side wall ("inner barrel") adjacent to lip L.
With particular reference to Figure 5, there is shown an aircraft, e.g. a fixed-wing aircraft, on which one or more of systems 10, 110, 210 are mounted, each one of them being designed in accordance with one of the previously described embodiments.
In particular, the aircraft comprises a pair of fixed wings W, wherein at least one of them has at least one surface operationally exposed to air, onto which at least one of systems 10, 110, 210 has been applied. Preferably but not necessarily, system 10, 110, 210 is positioned in the proximity of the leading edge of wing W.
Advantageously but not necessarily, system 10, 110, 210 can be positioned on the exposed surface of the flaps of wing W, in particular of at least one of front flap WS (commonly referred to in the industry as "nose droop or slat") and the rear flap WF (commonly referred to in the industry as "flap") . In the embodiment shown in Figure 5, both front flap WS and the rear flap are equipped with a system 10, 110, 210 designed in compliance with the present invention .
In the embodiment shown in Figure 5, at least one system 10, 110, 210 is arranged at the tail of the aircraft, i.e. on the horizontal empennages and on the vertical empennage, particularly on the leading edges of the tail itself.
The technical characteristics that differentiate the various embodiments and variants thereof described and illustrated herein are freely interchangeable, whenever compatible .
Of course, without prejudice to the principle of the invention, the forms of embodiment and the implementation details may be extensively varied from those described and illustrated herein by way of non-limiting example, without however departing from the scope of the invention as set out in the appended claims.
/GV

Claims

1. System (10, 110, 210) for preventing icing on an aircraft surface (28) operationally exposed to air;
said system being characterized in that it comprises a plasma actuator (12) to be applied onto said surface (28) operationally exposed to air, and arranged for generating at least one plasma discharge (D) for inducing a flow (F) of ionized hot-air particles towards said surface (28) operationally exposed to air.
2. System according to claim 1, wherein said plasma actuator (12) is of the dielectric barrier discharge type.
3. System according to claim 2, wherein said actuator ( 12 ) comprises :
- an intermediate portion (14) of dielectric material, - at least one exposed electrode portion (16, 18) positioned on an outer side (20), exposed to air, of said intermediate portion (14), and
- at least one covered electrode portion (22), which is not operationally exposed to air on said outer side (20) because of said intermediate portion (14);
said electrode portions (16, 18, 20) being electrically connectable to a high-voltage electric power generator (24) and being adapted to be energized by said electric power to generate said plasma discharge (D) between them (16, 18, 20) .
4. System according to claim 3, wherein said exposed electrode portion (18) is positioned on top of said intermediate portion (14) .
5. System according to claim 3 or 4, wherein said exposed electrode portion (16) is partially embedded in said intermediate portion (14) .
6. System according to any one of claims 3 to 5, wherein said covered electrode portion (22) is at least partially positioned under an inner side (24), opposite to said outer side (20), of said intermediate portion (14).
7. System according to claim 6, further comprising an insulating portion (26) of dielectric material positioned under said covered electrode portion (22), to be applied onto said exposed surface (28) of said aircraft.
8. System according to claim 6, wherein said covered electrode portion (22) is at least partly constituted by said exposed surface (28) of said aircraft.
9. System according to any one of claims 3 to 5, wherein said covered electrode portion (22) is embedded in said intermediate portion (14) .
10. Operational surface exposed to air of an aircraft, equipped with a system (10, 110, 210) according to any one of the preceding claims.
11. Engine nacelle (N) including at least one exposed surface according to claim 10, preferably positioned in the proximity of an air inlet (I) of said engine nacelle.
12. Wing (W) of an aircraft including at least one exposed surface according to claim 10, preferably positioned in the proximity of a leading edge.
13. Wing according to claim 12, wherein said exposed surface is positioned in the proximity of a leading edge of at least one flap or slat (RW, FW) of said wing (W) .
14. Aircraft including at least one exposed surface according to claim 10.
/GV
PCT/IB2014/058751 2013-02-06 2014-02-03 System for preventing icing on an aircraft surface operationally exposed to air WO2014122568A1 (en)

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US14/765,744 US10227923B2 (en) 2013-02-06 2014-02-03 System for preventing icing on an aircraft surface using a plasma actuator
EP14710970.6A EP2953852B1 (en) 2013-02-06 2014-02-03 Engine nacelle including a system for preventing icing on an aircraft surface operationally exposed to air

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ITTO2013A000095 2013-02-06
IT000095A ITTO20130095A1 (en) 2013-02-06 2013-02-06 SYSTEM TO COUNTER THE FORMATION OF ICE ON AN SURFACE OPERATIVELY EXPOSED TO THE AIR OF AN AIRCRAFT.

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US20160003147A1 (en) 2016-01-07
US10227923B2 (en) 2019-03-12
EP2953852B1 (en) 2017-09-20

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