WO2008132009A2 - Dispositif d'insonorisation pour moteur à réaction ou turbine - Google Patents

Dispositif d'insonorisation pour moteur à réaction ou turbine Download PDF

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Publication number
WO2008132009A2
WO2008132009A2 PCT/EP2008/053956 EP2008053956W WO2008132009A2 WO 2008132009 A2 WO2008132009 A2 WO 2008132009A2 EP 2008053956 W EP2008053956 W EP 2008053956W WO 2008132009 A2 WO2008132009 A2 WO 2008132009A2
Authority
WO
WIPO (PCT)
Prior art keywords
sound
engine
region
protection device
lining
Prior art date
Application number
PCT/EP2008/053956
Other languages
German (de)
English (en)
Other versions
WO2008132009A3 (fr
Inventor
Dominik Broszat
Reinhard Pongratz
Rudolf Maier
Original Assignee
Eads Deutschland Gmbh
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 Eads Deutschland Gmbh filed Critical Eads Deutschland Gmbh
Publication of WO2008132009A2 publication Critical patent/WO2008132009A2/fr
Publication of WO2008132009A3 publication Critical patent/WO2008132009A3/fr

Links

Classifications

    • 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
    • F04D29/663—Sound attenuation
    • F04D29/664—Sound attenuation by means of sound absorbing material
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
    • F02C7/04—Air intakes for gas-turbine plants or jet-propulsion plants
    • F02C7/045—Air intakes for gas-turbine plants or jet-propulsion plants having provisions for noise suppression
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02K—JET-PROPULSION PLANTS
    • F02K1/00—Plants characterised by the form or arrangement of the jet pipe or nozzle; Jet pipes or nozzles peculiar thereto
    • F02K1/44—Nozzles having means, e.g. a shield, reducing sound radiation in a specified direction
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02K—JET-PROPULSION PLANTS
    • F02K1/00—Plants characterised by the form or arrangement of the jet pipe or nozzle; Jet pipes or nozzles peculiar thereto
    • F02K1/78—Other construction of jet pipes
    • F02K1/82—Jet pipe walls, e.g. liners
    • F02K1/827—Sound absorbing structures or liners

Definitions

  • the present invention relates to a soundproofing device for a jet engine or a turbine, in particular for a turbofan engine of an aircraft, for attenuating a sound field in a directional range, in particular for shielding emitted sound in the direction of the ground. Moreover, the invention relates to a provided with such a soundproofing Thebtechniksummantelung and provided with such a soundproofing engine.
  • liners also referred to as liners.
  • Such liners are designed to absorb sound. It is possible to obtain certain properties of a soundproofing device by design and arrangement of the liner. Various approaches for improving the properties can be distinguished here.
  • zero-splice liners aim to reduce or avoid the hard transitions between the production-related necessary liner sections in the circumferential direction. This is described by way of example in EP 1 621 752 A2. On the one hand, such liners increase the effective absorptive area and, on the other hand, avoid the occurrence of so-called "mode scattering", a redistribution of the modal acoustic energy, which can adversely affect the achievable noise reduction.
  • Such liners are particularly advantageous when starting an aircraft, since the rotor harmonics predominate among the generated frequencies, to which this type liner is well tunable and works well, since the rotor harmonics have a large propagation angle.
  • the leaf foliage dominant in the approach however, frequencies are more capable of propagation and can not be so effectively dampened by the Zero Splice Liner.
  • EP 1 411 225 B1 shows an arrangement of liner segments in which two annular liner segments, which have a different impedance, are embedded in the engine duct in front of the fan. Sound stimulated by the fan, which moves in the axial direction, first encounters a first liner segment that excites certain modes that can be more effectively absorbed by the subsequent second segment. For the effect used, the respective impedances and the relative impedance jump at the interface of the two segments are decisive.
  • EP 1 701 016 A1 teaches to achieve a flowing or stepped transition between different impedances through different hole sizes of the perforated cover layer of the liner and / or the depth of the honeycomb core behind this cover layer. This aims to influence the angle at which the radiated sound is reflected by the liner. This will, ever after position in the engine intake, a different effect causes. Near the sound source (fan) the sound is deflected at a greater angle to the engine axis, as this results in a longer path of sound through the engine inlet and thus a higher absorption. In contrast, the propagation angle near the entry lip should be reduced as much as possible in order to radiate more in the axial direction, thereby increasing the run length of the sound to the bottom.
  • the object is to provide a soundproofing device, which makes it possible to achieve a reduction in the sound radiated towards the ground with the same aerodynamic properties of a conventional liner.
  • sound-influencing surfaces on a turbine housing are distributed unevenly such that sound is preferably directed away from the directional area to be shielded by an impedance jump between different surface structures and / or by reflection.
  • the soundproofing device according to the invention can be used in turbine housings of conventional design, so that it is aerodynamically more favorable as well as lighter than constructions with an advancing inlet lip. Furthermore, due to the influencing of the directional characteristic of the sound field, at least in a certain sector a higher noise reduction than in the prior art can be achieved.
  • the surface structures may be unevenly distributed circumferentially in a radial direction region on a surface facing the axis with respect to a central or rotational axis of the jet engine or the turbine.
  • Such arranged surface structures have the advantage that sound coming directly from the direction of the axis is reliably deflected.
  • the surface structures are distributed unevenly in the circumferential direction in a surface facing away from the axis A surface. This is particularly advantageous when sound from air-conducting housing elements which are located off the axis to be deflected.
  • the sound-influencing surface structures are provided on or in an inlet of the turbine. As a result, they are arranged between the sound source and the sound outlet, which allows a particularly efficient sound influencing.
  • the sound-influencing surface structures can advantageously be provided on or in a bypass duct of the jet engine, in particular in the region of an engine outlet and in particular in the form of an engine lining.
  • the surface structures have a sound-absorbing lining or are formed by such. This additionally reduces the radiated sound.
  • the surface structures advantageously have at least one reflection segment for targeted deflection of sound radiated into the directional area to be shielded.
  • the reflection segment has an impedance which is significantly different from the impedance of the sound-absorbing lining. differs. This makes it possible to achieve a deflection of the sound without geometric changes.
  • the lining is arranged on a surface which is directed towards the axis of rotation of the turbine and extends in the circumferential direction with respect to the axis of rotation of the turbine.
  • the reflection segment is preferably inserted in the lining in a peripheral area which is determined by the directional area in which sound radiation is to be reduced.
  • the reflection segment is arranged in the axial direction on a side of the lining facing away from the turbine.
  • the sound generated by the turbine must first pass through the sound-absorbing lining, which already dampens certain frequencies.
  • the residual sound remaining after that can be favorably influenced in his direction.
  • the impedance of the reflection segment is preferably lower than the impedance of the surrounding cladding.
  • the impedance jump deflects the sound away from the reflection segment.
  • the surface structure is preferably changeable.
  • the soundproofing device can be adapted to different installation situations or operating conditions and environmental conditions.
  • the or at least one of a plurality of reflection segments, in particular its arrangement, orientation or, more preferably, its impedance can be set. This allows you to react to different flight situations, such as takeoff and landing.
  • the soundproofing device is advantageously arranged on an inner surface of an inlet section of the turbine or of the engine.
  • the turbine housing may be, for example, an engine nacelle of an aircraft or other aircraft. In other embodiments, the turbine housing is embedded in the fuselage of an aircraft.
  • the inlet region advantageously has an inlet lip, in the region of which a sound-deflecting element, in particular the reflection segment, is arranged.
  • the reflection segment is advantageously arranged in a lower region of the inlet lip. This will deflect the sound upwards.
  • a further advantageous embodiment or a further field of application of the described soundproofing device with reflection segment is the engine outlet, for example a turbofan jet engine.
  • sound-absorbing linings are preferably on the inner wall and on the outer wall of a bypass duct (bypass duct) in order to achieve the greatest possible absorption of the exiting sound. Due to the ambient conditions, conventional liners are currently not usable in the core stream, which contains the hot combustion gases. Therefore, an arrangement in the bypass duct is preferred.
  • reflection segments are respectively arranged at the lower boundary of the bypass duct directly in front of the outlet of the engine. This means in the case of the upper sector of the bypass channel on the wall between bypass channel and core current, in the case of lower sector on the outer wall of the engine nacelle.
  • the expansion of the reflection segments in the circumferential direction can be favorably selected according to the application or the conditions of use.
  • a change of the surface structure or impedance can be provided.
  • the soundproofing device is particularly suitable for turbofan engines, in particular for arrangement on an engine casing, for example a (inner region of) a Thebtechniksgephaseuse (s).
  • FIG. 1 shows a radial longitudinal section of an inlet region of an engine nacelle with a soundproofing device according to a first embodiment of the present invention
  • Figure 2 is a section along Il - Il through the engine nacelle of Fig. 1.
  • Fig. 3 is a section as in Fig. 1 by an engine nacelle with a soundproofing device according to a second embodiment of the present invention and 4 shows a radial longitudinal section of an outlet of an engine with a soundproofing device according to a third embodiment of the present invention.
  • Exemplary embodiments of soundproofing devices 1 for jet engines or for turbines, such as gas turbines or turbofan engines, are illustrated in the attached figures using the example of a turbofan engine 2 of an aircraft.
  • the soundproofing device 1 has unevenly distributed surface structures 3 for targeted sound influencing.
  • the surface structures 3 are formed by an engine lining 4 on an inwardly facing inner surface 5 of a turbine casing element or of a turbine casing element 6.
  • the turbine shroud 6 is an engine nacelle 12 (also called a nacelle) in which the turbofan engine 2 is housed.
  • the turbofan engine 2 includes a fan 18 having blades 20 and a hub portion 22.
  • the hub portion 22 with the blades 20 is part of a rotor of the engine and rotates in operation about its axis A, which forms the axis of rotation of the turbofan engine 2.
  • the fan 18 promotes air from an inlet region 14 in the direction of a flow 24 to a turbine, not shown, or in the bypass channel.
  • Fig. 1 shows a first embodiment of the inlet region 14 having engine inlet 10 of a here as a turbofan engine 2 aircraft engine.
  • the turbofan engine 2 is enclosed by the engine nacelle 12 and has, on the inner surface 5 directed towards the axis A and extending in the circumferential direction, the soundproofing device 1 with the engine factory lining 4 on.
  • the inlet area 14 of the turbofan engine 2 having this soundproofing device 1 extends from a circumferential plane 16 of the sound forming circumferential front edge 17 of the engine nacelle 12 to the fan 18th
  • FIG. 1 shows a side view of a first embodiment of the soundproofing device 1
  • FIG. 2 shows a section along the line II-II of FIG. 1 in this first embodiment.
  • FIG. 3 shows a comparative illustration of the engine intake 10 with a second embodiment of the sound protection device 1.
  • the engine cowl 4 in both embodiments has a sound absorbing liner 26 and a reflective liner segment, hereinafter referred to as a reflective segment 30.
  • the nacelle or engine nacelle 2 is aerodynamically contoured in both embodiments and surrounds the engine to direct the free flow of air into the engine.
  • the fan 18 plays a more and more important role due to a steadily increasing bypass ratio during takeoff and landing over the years.
  • the sound field in these two flight conditions fundamentally different, since each modal tonal sound fields of different cause represent the dominant parts.
  • the so-called "Buzz Saw Noise” This occurs at the harmonics of the rotor speed and corresponds in its circumferential order to the number of rotor blades.
  • the engine is adjusted to a much lower speed on approach.
  • the rotor harmonics are "cut off", ie they are not capable of propagation, but the blade follower frequencies (abbreviated BPF for short) dominate, which consist of an interaction of the rotor blades of the fan 18 and the following (not shown) stator blades ("stators"). Their frequencies correspond to the product of the rotor speed and the number of rotor blades.
  • the sound absorbing liner 26 is mounted in the engine intake 10 to reduce the sound emitted by the engine.
  • the sound absorbing liner 26 (also called absorptive liner) is typically optimized for flight conditions to reduce the fan blade follow-up frequencies BPF 18 (as dominant tonal components).
  • BPF 18 the fan blade follow-up frequencies
  • the impedance of the materials of the sound-absorbing lining 26 is selected accordingly.
  • the technique, how one can tune impedances of linings to certain dominant tonal components for absorption thereof, is known in the prior art described above and will not be described here.
  • such a sound-absorbing lining 26 with a liner segment or lining segment-the reflection segment 30- is combined with impedance deviating significantly from the impedance of the sound-absorbing lining 26.
  • the impedance of the reflection segment 30 is set very low here.
  • the reflection segment 30 is formed here as part of the engine casing 4. It is in the first embodiment shown in FIGS. 1 and 2 in the front region of the engine intake 10 at a side remote from the fan 18, here front side of the sound-absorbing lining 26 mounted (see Fig. 1). In the second embodiment shown in FIG. 3, the reflection segment is arranged in the region of the inlet lip, so to speak as a lip-liner, ie as a lip lining.
  • This Refklexionssegment 30 aims to achieve by the resulting impedance jump at the interface, here at the interface 27 between the sound-absorbing liner 26 and the reflection segment 30, a directional influence of the incident sound field.
  • the noise radiated towards the ground is thereby to be reduced and the noise pollution of an area surrounding an airport is to be reduced.
  • the mode of action here corresponds to that of a "negatively scarfed inlets", ie an engine intake with an advanced lower lip, but dispenses with the additional structure of the extended lower lip and thus on weight penalties.
  • the acoustic effect is achieved by the targeted positioning of the reflective-acting liner segment - reflection segment 30.
  • Other embodiments not shown here allow a variation of the impedance of the sound-absorbing lining 26 (absorptive Liner- share) to adjust the flight state according to most favorable impedance and thus to achieve the highest possible absorption before the forward of the fan 18 propagating sound on the reflection segment 30 hits.
  • This adjustment of the impedance can be made either by adjusting the entire sound-absorbing lining 26 or even the entire engine lining 4 (sound-absorbing lining 26 with reflection segment 30). Additionally or alternatively, adjusting means are provided with which an adjustment of discrete, distributed in the engine lining 4, adaptive elements takes place, resulting in averaged over the area ratio impedance.
  • the sound-absorbing lining 26 is arranged on the inside of the engine nacelle 12 in the circumferential direction relative to the axis A.
  • the reflection segment 30 is inserted on the side of the sound-absorbing lining 26, which faces away from the fan 18 along the axis A in the axial direction.
  • the reflection Xionssegment 30 has compared to the sound-absorbing lining 26 to a much lower impedance.
  • FIG. 2 shows a view from the exit plane 16 in the direction of the axis A of the engine intake 10 and the fan 18.
  • the blades 20 and the hub portion 22 are clearly visible. Further, the circumferentially arranged sound absorbing liner 26 is visible. In the lower region 28, the reflection segment 30 is inserted into the sound-absorbing lining 26.
  • the entire engine lining 4 and thus also the sound-absorbing lining 26 are preferred up to the inlet lip 25.
  • the reflection segment 30 is now arranged on an inlet lip 25.
  • the function of the reflection segment 30 can be well explained with reference to FIG. 1.
  • the sound generated by the fan 18 propagates in the direction of the axis A in the direction of the flow 24. In doing so, it passes through the sound-absorbing lining 26, which absorbs frequencies in a certain frequency range.
  • the sound strikes the reflection segment 30, which has a much lower impedance than the sound absorbing lining 26. As this hard impedance jump passes, the sound is deflected upwards.
  • the second embodiment shown in FIG. 3 goes one step further and additionally arranges the sound-absorbing lining 26 in the area of the inlet lip 25. This increases the effective absorptive distance at which the sound is attenuated by the sound absorbing liner 26. Furthermore, the reflection segment 30 is arranged on the inlet lip 25, which makes it possible to deflect an even larger portion of the downward sound upward because there is no longer the possibility that the deflected upward sound from the upper region of the engine nacelle 12th is thrown back down.
  • the reflection segment 30 extends in the circumferential direction to an angle of about ⁇ 40 ° from the vertical V (see Fig. 2).
  • a third embodiment of the invention can also be used on an engine outlet 34.
  • the engine exhaust 34 of the turbofan engine 2 includes end portions of the engine nacelle 12, a circumferentially circumferential bypass duct 36 disposed within the engine nacelle 12, and a core stream 38 separated from the bypass duct 36 by a substantially cylindrical wall.
  • the core stream 38 carries hot combustion gases 40 of the engine 2 while the bypass channel 36 carries ambient air 42.
  • a sound-absorbing lining 26 is provided as an engine lining 4, which acts as a soundproofing device 1.
  • reflection segments 30 inserted into the sound-absorbing lining 26 are provided on the end-side edges 44 of the wall 37 and the engine nacelle 12.
  • the reflection segment 30 By extending the reflection segment 30 in the direction of the circumference, it is possible to set the directional region 32 in which a sound radiation is to be reduced.
  • This directional range 32 is indicated in FIG. 2 by arrows.
  • the directional region 32 is here a radial region extending over a certain angle in the circumferential direction.
  • the reflective segment 30 and / or the liner 26 may be constructed so that their impedance is adjustable. This could for example be achieved in that the surface of the lining 26 consists of two mutually displaceable, abutting, perforated plates. If the perforated plates are displaced relative to one another, the hole size of the surface of the lining 26 or of the reflection segment 30 changes. With such a variable surface structure, the radiated sound field can be adapted to a wide variety of flight situations.
  • a suitable non-uniform distribution of sound-absorbing lining 26 and reflection segment 30 in the inlet region 14 of the engine nacelle 12 results in a directional influence of the radiated sound field, which causes a greater noise reduction on the ground than with conventional measures.

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

Abstract

L'invention concerne un dispositif d'insonorisation destiné à un moteur à réaction ou à une turbine, en particulier à un turboréacteur à double flux (2) d'un aéronef, ce dispositif servant à atténuer un champ sonore dans une région directionnelle (32), en particulier à atténuer les sons émis en direction du sol. À cet effet, un revêtement (4) du moteur présente des structures superficielles (26, 30) réparties irrégulièrement de sorte que les sons soient déviés de la région directionnelle (32) de façon ciblée par un saut d'impédance et/ou par réflexion.
PCT/EP2008/053956 2007-04-25 2008-04-02 Dispositif d'insonorisation pour moteur à réaction ou turbine WO2008132009A2 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
DE102007019762.6 2007-04-25
DE200710019762 DE102007019762A1 (de) 2007-04-25 2007-04-25 Schallschutzvorrichtung für ein Strahltriebwerk oder eine Turbine

Publications (2)

Publication Number Publication Date
WO2008132009A2 true WO2008132009A2 (fr) 2008-11-06
WO2008132009A3 WO2008132009A3 (fr) 2009-02-26

Family

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Family Applications (1)

Application Number Title Priority Date Filing Date
PCT/EP2008/053956 WO2008132009A2 (fr) 2007-04-25 2008-04-02 Dispositif d'insonorisation pour moteur à réaction ou turbine

Country Status (2)

Country Link
DE (1) DE102007019762A1 (fr)
WO (1) WO2008132009A2 (fr)

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN111741899A (zh) * 2018-02-19 2020-10-02 赛峰飞机发动机公司 具有声音可透过壁的涡轮机短舱

Families Citing this family (5)

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Publication number Priority date Publication date Assignee Title
DE102011056826B4 (de) 2011-12-21 2014-06-26 Eads Deutschland Gmbh Schallschutzvorrichtung und damit versehenes Triebwerk und Verfahren zur Bereitstellung
WO2016190753A1 (fr) 2015-05-25 2016-12-01 Dotterel Technologies Limited Carénage pour aéronef
DE102016123096B4 (de) 2016-11-30 2023-06-22 Airbus Defence and Space GmbH Steuerflächenbauteil für eine Auftriebshilfevorrichtung eines Luftfahrzeugs sowie Herstellungsverfahren hierfür
AU2018306554A1 (en) 2017-07-24 2020-02-20 Dotterel Technologies Limited Shroud
US11721352B2 (en) 2018-05-16 2023-08-08 Dotterel Technologies Limited Systems and methods for audio capture

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US3890060A (en) * 1974-02-15 1975-06-17 Gen Electric Acoustic duct with asymmetric acoustical treatment
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GB0223697D0 (en) 2002-10-14 2002-11-20 Rolls Royce Plc Acoustic liner for gas turbine engineers
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US20060169532A1 (en) 2005-02-03 2006-08-03 Patrick William P Acoustic liner with nonuniform impedance

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GB1236854A (en) * 1968-09-16 1971-06-23 Gen Electric Improvements in sound suppression of compressors used in gas turbine engines
US3481427A (en) * 1968-11-29 1969-12-02 Mc Donnell Douglas Corp Acoustical panel structure
US3890060A (en) * 1974-02-15 1975-06-17 Gen Electric Acoustic duct with asymmetric acoustical treatment
US3964569A (en) * 1974-09-06 1976-06-22 General Electric Company Gas turbine engine noise shield
US5702231A (en) * 1996-08-09 1997-12-30 The Boeing Company Apparatus and method for reducing noise emissions from a gas turbine engine inlet

Cited By (2)

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Publication number Priority date Publication date Assignee Title
CN111741899A (zh) * 2018-02-19 2020-10-02 赛峰飞机发动机公司 具有声音可透过壁的涡轮机短舱
CN111741899B (zh) * 2018-02-19 2023-12-22 赛峰飞机发动机公司 具有声音可透过壁的涡轮机短舱

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Publication number Publication date
DE102007019762A1 (de) 2008-10-30
WO2008132009A3 (fr) 2009-02-26

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