CA2147183A1 - Active gas turbine (jet) engine noise suppression - Google Patents

Abstract

A method and system for reducing the acoustic levels of internal and external sound fields (23 and 24) generated by gas turbine en-gines (2) has several actuators to generate sound (50, 84, and 115), several sensors to measure the acoustic levels (49 and 69), and one or more controllers (113). The controllers are adaptive self-learning neural networks (112) that control the actuators to generate sound in order to effect the reduction of the in-ternal and external sound field as measured by the sensors.

Description

- W094/08~0 2 1 4 7 1 8 ~ PCT/US93/09739 1 ACTIVE GA8 TURBINB (JET) ~N'~lN~ NOIS~ ~u~rK~88ION

2 This invention relates to acoustical apparatus, and 3 to methods to actively and adaptively suppress the acoustic 4 noise, such as adaptive vibration control system. In particular it relates to such apparatus and methods for 6 suppressing acoustic noise produced internally and externally 7 by a jet (gas turbine) engine and perceived far from the 8 engine.

g BACRGROUND OF THE lNV~ ON
Commercial jet aircraft currently are required to 11 meet governmentally specified perceived noise criteria. The 12 prior systems used in engine silencing include both passive 13 and active methods and apparatus. Passive means include 14 mufflers, acoustic treatment in the inlets and exhaust ducts and tailpipes. Active means include engine surge control, 16 turbofans with high bypass ratios, bypass/exhaust mixers, 17 external suppressors, and acoustic wave interference in engine 18 inlet ducts, exhaust tailpipes, and mufflers.
19 The previous art, for example, Wanke U.S. patent 3,936,606, issued February 3, 1976, utilized acoustic wave 21 interference to achieve sound reduction in gas turbine 22 engines. A microphone, or equivalent, measured the existing 23 acoustic wave. Conventional adaptive control apparatus 24 created a time-delayed and phase reversed mirror symmetry signal to generate an anti-noise acoustic wave downstream from 26 the microphone via a "speaker" in a turbojet's inlet duct and 27 exhaust tailpipe. The Wanke patent, however required detailed 28 modelling of the acoustic wave and the counter wave that would 29 cancel it. Since the wave that effected cancellation was a simple time-delayed and phase- reversed wave front, Wanke 31 found it necessary to direct the waves through a wave guide 32 that converted all the wave energy of both the acoustic wave 33 and the counter wave into plane waves or other predictable 34 wave modes. Cancellation could only be accomplished within such wave guides, and therefore they had to be present within W094/08~0 2 1 ~ 7 1 8 3 PCT/US93/09739 1 the region where the acoustic energy would be canceled. The 2 use of such conventional noise control systems has had limited 3 results in turbojet engine inlet ducts.
4 Other prior systems of active noise suppression on gas turbines has not produced cancellation of non-linear, 6 random noise over the needed acoustic frequency range in real 7 time.
8 The prior methods and apparatus for active sound 9 control, noise cancellation, noise abatement, noise attenuation, and the like, involve conventional adaptive 11 controllers or adaptive filters. These systems require 12 extensive system modeling in order to operate successfully.
13 They have limited abilities in non-stationary and non-linear 14 acoustic applications.

BRIEF DESCRIPTION OF THE lNV~ lON
16 The invention realizes a reliable, adaptable, and 17 cost-effective means to upgrade aircraft jet engines to meet 18 current and future noise criteria. The methods of this 19 invention also has application to helicopter, industrial and military gas turbine engines. They also may be applicable to 21 steam turbines, reciprocating engines, and electrical motor 22 acoustic noise as well. The invention provides an active 23 apparatus and method for significantly reducing the acoustic 24 levels of sound generated by gas turbine and other engines.
The reductions of the sound level on aircraft gas turbine 26 (jet) engines are addressed specifically in this patent 27 description; however, the invention is not limited to jet 28 turbine engines.
29 In the present invention, sound sources that are installed at appropriate locations within the gas turbine 31 engine are the noise suppression means. The system that 32 controls these sound sources learns, self tunes, and adapts to 33 the in situ noise environments to produce acoustic waves. The 34 acoustic waves are equal to and opposite to the mixture of periodic, harmonic, and random noise acoustic waves produced 36 by the gas turbine's internal processes. The two sets of - W094/08540 21~7183 PCT/US93/09739 1 acoustic waves are nearly mirror images of each other.
2 Therefore, their mutual interference causes them to cancel 3 each other. This process reduces the sound levels produced 4 within the engine. Also, this process reduces the sound levels externally propagated from the engine. Measurements of 6 any residual acoustic waves assess the effectiveness of the 7 cancellation process. The residual acoustic signals become 8 the error signals fed back to the control system. The control 9 system learns to minimize the error signals under time-varying (non-stationary) conditions. Thus the final, residual 11 acoustic noise field is minimized.
12 The primary advantage of this invention over the 13 prior systems is its ability to tune itself in situ to 14 variations measured downstream from both the engines acoustic noise sources and the canceling acoustic noise generators.

16 The Structure Of Gas Turbine Engines 17 To appreciate the context in which the present 18 invention was made it is necessary to have some understanding 19 of the structure of gas turbine engines. A gas turbine has two main acoustic energy paths to the atmosphere, namely, its air 21 inlet and its exhaust outlet. Within the engine there are two 22 major energy paths. The fan bypass duct represents one path.
23 The air compressor/combustion and gas generator/gas 24 turbine/exhaust duct is the other path. Air entering the inlet is compressed by the bypass fan stage and divided 26 between the bypass duct and the low pressure compressor inlet 27 duct. The air in the bypass duct continues through to the 28 exhaust end of the engine. Successive compressor stages of 29 the low pressure compressor compress the remaining air until it enters the high pressure compressor. The high pressure 31 compressor further compresses this air in successive stages.
32 High pressure air enters the combustion chambers, mixes with 33 fuel and ignites to produce large amounts of high temperature, 34 high pressure gas. This gas drives successive power turbine sections. The successive power turbine stages provide power 36 to the compressor sections including the bypass fan. The flow 2t471~3 W094/08~0 ' ~ PCT/US93/09739 , . . .

1 of hot exhaust gas from the power turbines enters the exhaust 2 tailpipe section where it mixes with the cooler bypass air.
3 This total air/gas mass flow provides the engine thrust.

4 Sound Sources In Gas Turbine Engines An operating gas turbine engine presents three 6 significant acoustic noise sources to the surrounding air 7 medium. The primary noise source is the engine exhaust.
8 Another noise source is the air inlet. Mechanical engine 9 noises that radiate through the nacelle structures represent a third acoustic source of an operating engine.

11 The Location Of Sound suPPression Sources 12 Therefore, a plurality of acoustic sources placed 13 within the bypass air ducts and exhaust ducts suppress the 14 rearward-radiated exhaust noise. Acoustic cancellation source locations are at or near the junction where the bypass air 16 meets and mixes with the exhaust gas. An additional plurality 17 of acoustic sources placed in the inlet duct suppress the 18 forward-radiated inlet noise. If necessary, an additional 19 plurality of vibration/acoustic sources placed on or near the engine structures suppress the mechanically radiated noise.

21 Neurocontroller Suppression 22 Microphones, or equivalents (such as dynamic 23 pressure sensors, and/or accelerometers) strategically placed 24 in and around the engines, on the aircraft fuselage, wings, nacelles, tail, and stabilizer structures measure the 26 resulting residual sound levels. A neural network is used to 27 control the acoustic sources that effect the noise 28 cancellation. This network is termed a neurocontroller.
29 The measurements of residual sound levels provide 'error' signals to the neurocontroller for adaptive 31 suppression of the engine noise. The neurocontroller will 32 continue to minimize the sound levels at the measurement 33 locations by controlling the plurality of acoustic sources at 34 all locations. The neuro-controller produces sound waves that 1 interfere destructively with the engine-produced sound waves.
2 This wave interference process cancels these latter waves.
3 The net result is suppression of the external acoustic field 4 produced by the operating engines.
The system of the present invention comprises a - 6 plurality of means to generate sound (termed the 'Actuators');
7 a plurality of means to measure sound (termed the 'Sensors');
8 one or more adaptive self-learning neural network-based 9 Controllers (termed the 'Controllers'.) The Controllers control the plurality of Actuators to affect the reduction of 11 the measured internal and external sound field. This system 12 is integrated, installed, and operated on one or more engines 13 (termed the 'Apparatus').
14 The preferred embodiments of the present invention have the following features:
16 The acoustic sources comprise high-intensity 17 Actuators that include high-intensity air-stream modulators, 18 high-intensity speakers, and high-intensity mechanical 19 actuators. These Actuators may be placed inside the engine to suppress the noise at or very near the noise sources. Air 21 modulator-based acoustic sources may be employed that use a 22 flow of compressed air. Adequate compressed air for these 23 Actuators is available from appropriate compression stages of 24 the operating turbofan engine to achieve a very energy-efficient operation.
26 The Sensors that control performance of the noise 27 suppression system are located both inside the engine and in 28 the far and near external radiation fields of the engine. A
29 plurality of such error sensors control a plurality of Actuators simultaneously to suppress acoustic noise from 31 multiple engines. That is, the controller is a Multiple-32 Input, Multiple-Output (MIMO) neuro-controller.
33 The neural network based controller is of a unique 34 form. It has the following features: It incorporates a MIMO
neurocontroller that includes one or more neural network based 36 embodiments of the filtered-x algorithm. It includes one or 37 more MIMO neuroemulators that automatically learn, ln situ, 2147183 ~ -1 the acoustic-coupled transfer functions between the 2 pluralities of both error Sensors and Actuators. It includes 3 one or more MIMO neurocontrollers that automatically learn, in 4 situ, to control the plurality of Actuators simultaneously to suppress the sound at the plurality of error Sensor locations.
6 As a result, the system of the present invention 7 learns and adapts in real time to changes in engine 8 performance, aircraft operation, and external aircraft 9 environment. It may be synchronized by engine keyphasors or tachometers. It learns and adapts in real time to changes in 11 residual acoustic fields. It is non-obtrusive to engine 12 performance without degrading fuel performance since the 13 system is lightweight, efficient, and has low-power use. The 14 system is adaptable to both gas and steam turbines. The system is adaptable to other rotating engines and machinery 16 types.

18 The above description, as well as further objects, 19 features and advantages of the present invention, will be more fully appreciated by reference to the following detailed 21 description of a presently preferred, but nonetheless 22 illustrative, embodiment in accordance with the present 23 invention when taken in conjunction with the accompanying 24 drawings wherein: Fig. lA illustrates a side view of the major sections of a turbofan engine and the zones of the 26 engine and nacelle reserved for the installations of Actuators 27 and Sen~-ors.
28 Fig. lB is a schematic diagram showing the primary 29 noise source locations relative to a turbofan engine.
Fig. lC illustrates the process of wave cancellation 31 by wave interference.
32 Fig. lD is a partly block, partly schematic diagram 33 of a novel method and Apparatus for suppressing acoustic noise 34 by use of wave interference.
Fig. 2A is one embodiment of the invention 36 illustrating a side cross-section of part of a turbofan 21~7183 1 engine.
2 Fig. 2B is a cross-section along lines A-A of Fig.
3 2A.
4 Fig. 2C is a cross-section along lines B-B of Fig.
2A.
6 Fig. 2D is a cross-section along lines C-C of Fig.
7 2A.
8 Fig. 2E is a cross-section along lines D-D of Fig.
9 2A.
Fig. 2F is a cross-section along lines E-E of Fig.
11 2A.
12 Fig. 2G is a cross-section along lines F-F of Fig.
13 2A.
14 Fig. 2H is a cross-section along lines G-G of Fig.
2A.
16 Fig. 2I is a cross-section along lines H-H of Fig.
17 2A.
18 Fig. 3 is a plan view of a typical aircraft showing 19 typical additional Sensor locations.
Fig. 4A is a cross-section view of a typical air 21 modulator and horn.
22 Fig. 4B is a cross-section along lines I-I ofFig.
23 4A.
24 Fig. 4C is a cross-section along lines J-J of Fig.
4A.
26 Fig. 4D is a cross-section along lines K-K of Fig.
27 4A.
28 Fig. 4B is a plan view of a typical impedance-29 matching acoustic horn.
Fig. 5A is a top and side plan view of a typical 31 high-intensity speaker.
32 Fig. 5B is a cross-section along lines L-L of Fig.
33 5A.
34 Fig. 6A is a plan view of a typical h gh-force piezoceramic Actuator 'patch'.
36 Fig. 6B is a cross-section along lines M-M of Fig.
37 6A.

W094/08~0 PCT/US93/09739 1 Fig. 7 is a three-dimensional view of a section of 2 duct wall illustrating typical installations of piezoceramic 3 Actulator "patches".
4 Fig. 8 is a plan view of a typical internal sensor installation along the lines A-A, E-E and H-H of Fig. 2A.
6 Fig. 9, is a schematic of typical instrumentation 7 installed on a two-engine aircraft.
8 Fig. 10A, is a schematic block diagram illustrating 9 the processes of Controller synchronization.
Fig. 10B is a block diagram of a typical embodiment 11 of the neural network based filtered-x algorithm implemented 12 in the current noise suppression MIMO neurocontroller.
13 Fig. llA is a schematic illustrating the emulator 14 network of the present invention.
Fig. llB is a schematic diagram showing the emulator 16 input signal.
17 Fig. llC illustrates the timing during one sample 18 epoch of neuroemulator training.
19 Fig. 12A is a schematic block diagram illustrating the training of the neurocontroller of Fig. 10B.
21 Fig. 12B depicts the trained neuroemulator input 22 signal sequence.
23 Fig. 12C is a schematic depicting the training of 24 the neurocontroller.
Fig. 12D illustrates the timing involved with the 26 events during one sample epoch of neurocontroller operation.
27 Fig. 13A further depicts the emulator network of the 28 present invention.
29 Fig. 13B is a schematic depicting overlapped feedforward operations of the present invention.

31 Fig. 13C further depicts the timing in the present 32 invention.

34 Fig. lA is a schematic diagram of a typical jet engine nacelle 114. The nacelle consists of cowl 9 and inlet 36 duct 13, the engine and cowl panels 10 and bypass duct 14, and - W094/08540 2 1 4 7 1 ~ 3 PCT/US93/09739 1 the exhaust duct 15 that includes tailpipe extension 11 and 2 the thrust reverser and tailpipe 12. The nacelle and engine 3 schematic shows the zones for the installation of noise source 4 Sensors 7, internal error Sensors 5, external error Sensors 4 and 8, and canceling Actuators 6. Fig. lA also illustrates 6 the positions of the turbofan 1 and the engine 2 within the 7 Nacelle 114.
8 Fig. lB shows a schematic diagram of the primary 9 turbofan engine-produced acoustic noise sources. The rotating turbine blade-generated noise region 21 includes the noise 11 generated by the combustion chambers 19 and the noise 12 generated by the power turbines 20. The noise in region 21 13 combines with the turbofan blade-generated bypass noise in 14 bypass air 17. The acoustic noise sources 21 and 17 combine to become the rearward-radiated internal engine-generated 16 exhaust noise in exhaust noise zone 22. This noise 22 17 combines with the rearward-radiated external acoustic noise in 18 shear noise zone 23 generated in the shear mechanisms between 19 the exhaust gases and the ambient atmosphere. The inlet acoustic noise in source zone 24 consists of the forward-21 radiating acoustic noise generated by the rotating turbofan 22 blades 1 and the rotating compressor blades 18.
23 Fig. lC is a schematic diagram illustrating the 24 process of acoustic wave destructive interference. Fig. lC
shows the noise source waveforms as they would appear at a 26 plurality of measurement locations 26. The noise suppression 27 system uses these measurements to control a plurality of 28 'anti-noise' generators 50, 115, 84 (see Fig. 2A) located in 29 canceling Actuator zones 6 to produce canceling acoustic noise 'waves' 27 as shown in Fig. lC. Fig. lC shows that these 31 waves collectively interfere with the existing acoustic noise 32 'waves' 28 at the plurality of measurement locations in the 33 error sensor zones 5. The result of this action markedly 34 reduces the amplitudes of the internally-produced acoustic 'waves' 29 that continue to propagate out of the exhaust 15 or 36 out of the inlet 13. When this process is extended to an 37 additional plurality of measurement locations 8 and 4, ~147`1~3 W094/08~0 PCT/US93/09739 1 external to the engine, additional noise reduction results.
2 Fig. lD illustrates a method and apparatus for 3 providing canceling acoustic waves propagating through an 4 elastic medium such as air. Acoustic error Sensors 49 and 69, (see Fig. 2A) such as microphones, generate audio varying 6 output signals representative of the acoustic waves that exist 7 at the noise source Sensor locations 31 and 33 in zones 5 and 8 the error Sensor locations 30, 32, 34 and 38 in zones 7, 4 and 9 8. These signals provide inputs to a neurocontroller 113 that learns to generate audio varying output signals that in turn 11 cause acoustic generators 50, 115 and 84 (see Fig. 2A) at the 12 canceling Actuator locations 35, 36 and 37 in zone 6 to 13 produce acoustic waves that are mirror images of the engine-14 generated acoustic waves measured at error Sensor locations 30, 32, 34 and 38. The generated acoustic waves generated in 16 zones 6 interfere with the existing acoustic noise waves at 17 the error Sensor locations 30, 32, 34, and 38. The final 18 result is wave interference and cancellation. The 19 neurocontroller means 113 will continue to adapt its output signals 35, 36, 37 such that the audio signals measured by the 21 microphones 49 and 69 (see Fig. 2A) are always minimized. The 22 net result of the invention is the reduction in the measured 23 sound levels at the microphone locations in both the internal 24 error Sensor zones 49 and the external error Sensor zones 69.

Fig. lD illustrates the application of the invention 26 to one or more gas turbine (turbofan) engine assemblies 114.
27 A turbofan engine 2 creates thrust by moving air from the 28 surrounding air medium through the engine, compressing it 18 29 mixing fuel with it, and igniting the fuel air mixture to produce a hot exhaust gas 19. The hot gas rotates the power 31 turbines 20, that, in turn, rotate the turbofan 1 and 32 compressors 18. The hot exhaust gasses mix with the cooler 33 bypass air 21. Finally, the exh~ust mixture 22 moves out of 34 the engine 2 through the exhaust tailpipe 15. Thrust forces produced by these processes on the engine move the aircraft 36 forward. The engine generates acoustic noise due to the - W094/08540 2 ~ ~ 7 1 8 3 PCT/US93/09739 1 aforedescribed processes. In addition, the turbulent boundary 2 layer between the hot exhaust mixture and the relatively 3 'undisturbed' cooler atmosphere produces acoustic noise 23.
4 These two sources of acoustic noise combine to radiate into the far field (that is, far from the engine).
6 Acoustic noise generated by the compressing blades 18 radiates 7 forward out of the inlet duct 13. Similar acoustic noise 17 8 is radiated rearward into the bypass duct 16 from the actions 9 of the bypass fan 2. The inlet's forward-radiated noise 24 also combines with the exhaust's rearward-radiated noise 22 11 and 23 at locations far (far field) from the engine. Multiple 12 engines combine to increase the radiated far-field noise.
13 One or more acoustic sources (usually actuators 50, 14 air stream modulators 115 (see Fig. 4A), or piezoceramic actuators 84 (see Fig. 6A) are the means to create sound.
16 Fig. lA, lD and 2A show where the Actuators are located, 17 namely in the actuator zones 6 of the engine exhaust 15, inlet 18 13 and bypass ducts 16. Fig. lD depicts the electrical drive 19 signals 35, 36 and 37, that are furnished to the Actuators 50, 115 and 84 from the Controllers 113. The controllers 113 21 condition the drive signals to match the electrical 22 characteristics of the Actuators 50, 115 and 84. Fig. lC
23 shows that the Actuators 50, 115 and 84 at locations 35, 36 24 and 37 adaptively produce canceling acoustic waveforms 27 that interfere destructively with the engine generated acoustic 26 waveforms 26 at the internal and external engine locations 30, 27 32, 34, and 38 of the error Sensors 49 (see Fig. 2A) and 69.
28 This destructive waveform interference reduces the sound 29 levels measured by the error Sensors 49 and 69. This process reduces the sound levels internal to the engine and realizes a 31 reduction in the external perceived sound levels.
32 The means to measure sound are preferably of two 33 categories: source Sensors 49 (usually dynamic pressure 34 sensors or microphones) to measure the normal engine generated sound; and error Sensors 49 and 69 (usually dynamic pressure 36 sensors or microphones) to measure reduced sound.
37 The source Sensors 49 are 'upstream' from the W094/08540 ~1~ 7 1 8 3 PCT/US93/09739 1 Actuators 50, 115 and 84 and the error Sensors 49 and 69 are 2 'downstream' from the Actuators 50, 115 and 84. "Upstream"
3 refers to closer to the sound source 1 and 15 than the 4 Actuators and "downstream" refers to further from the sound source 1 and 15 than the Actuators. The term Sound source 6 refers to the inlet duct/compressor bypass fan end 13. Sound 7 source also refers to the turbine 8 exhaust/bypass/duct/tailpipe/thrust reverser end 15. Fig. lA, 9 lD and 2A show typical sound source locations.
As seen in Fig. lB and 2A, the turbofan engine 2 11 generates acoustic waves propagating forward from its air 12 inlet 13, propagating rearward from its exhaust outlet 15, and 13 propagating outward from the engine's vibrating cowling 14 (nacelle) structure 3. The processes of compressing the air entering the inlet by the rotating bypass fan blades 1 and the 16 successive stages of rotating compressor blades 18 generate 17 the acoustic waves propagating from the inlet 13.
18 Acoustic waves generated at the inlet propagate 19 rearward through the turbofan engine's air bypass duct 16.
These acoustic waves also propagate rearward through the air 21 compressor inlet duct 18. They are combined with the acoustic 22 waves generated by each successive compressor stage. This 23 complex acoustic field enters the gas generator ducts 19, 24 combines with the complex acoustic fields generated by the combustion processes and enters the power turbines 20 and 26 exhaust duct 15. These two sets of complex acoustic waves 27 recombine at the exhaust/tailpipe end of the engine and 28 propagate rearward out of the tailpipe 11 and 12. The 29 exhausted acoustic waves 21 combine with acoustic waves generated by the shear mechanisms created at the boundary 31 between the relatively cool ambient air medium and the hot 32 moving exhaust gases 23.
33 In a preferred embodiment, each engine has a 34 plurality of installed Actuators. Fig. 2C illustrates typical Actuator 50 locations 35 at the exhaust end of the engine 15 36 along lines B-B of Fig. 2A. Fig. 2H illustrates typical 37 actuator 50 locations 37 at the inlet end 13 of the engine - W094/08~0 ~1 4 7 ~ 8 3 PCT/US93/09739 1 along lines G-G of Fig. 2A. Fig. 2G illustrates typical 2 actuator 50 locations 36 on the bypass duct 16 portions of the 3 engine along lines F-F of Fig. 2A.
~ 4 Fig. 2D is a typical cross-section of an air modulator actuator horn along the lines C-C of Fig. 2A. The 6 horn section 53 of the air modulator serves to increase the 7 sound levels of the air modulation 51. The horn 116 provides 8 impedance loading at the throat 70 of the horn to enhance low-9 frequency performance. And, the horn section matches the impedance of the ambient bypass air flow at the mouth 71 of 11 the horn. Fig. 2E is a typical cross-section of an air 12 modulator actuator driver 51 and plenum 52 along the lines D-D
13 of Fig. 2A.
14 The horn mouth attaches to an entry hole in the bypass duct wall by a flange 72 (see Fig. 4E). The plenum 52 16 attaches by a pipe manifold 54. The pipe manifold 54 attaches 17 to a source of high pressure air. Appropriate bleed ports(s) 18 on the engine's compressor stages provide the sources for 19 high-pressure air. The only stringent requirements are that the compressed air be very clean and the plenum 52 pressure 21 must be at least twice as high as the ambient bypass duct 16 22 pressure. Thus, the throat 70 of the horn 116 serves as the 23 exit from the high pressure plenum 52 and air modulation.
24 And, the mouth of the horn 71 serves as the entrance back into the bypass duct 16. Each engine has a plurality of air 26 modulator actuators 115.
27 Each engine has a plurality of internal source 28 Sensors 49 and error Sensors 49 and 69. Typical internal-to-29 the-engine Sensor locations 30, 31, 32, 33 and 34 are shown in Fig. 2A. The Sensors 49 in the exhaust tailpipe 15 and inlet 31 13 provide the error signals to the neurocontroller 113. The 32 neurocontroller 113 optimizes the generation of canceling 33 acoustic waves 27 to improve the interference between the 34 internal, controlled, acoustic wave generation and the internal, engine generated acoustic waves 26. These signals 36 fuse with additional near-field error Sensor signals 69 at 37 locations 38 to assure far-field noise reductions. Fig. 2B

W094/08~0 2 1 4 7 1 8 3 PCT/US93/09739 ~~

1 illustrates typical sensor locations 34 at the exhaust end 15 2 of the engine along lines A-A of Fig. 2A. Fig. 2I illustrates 3 typical sensor locations 30 at the inlet end 13 of the engine 4 6 along line H-H of Fig. 2A. Fig. 2F illustrates typical sensor locations 31, 32 and 33 on the bypass duct 16 portions 6 of the engine along lines E-E of Fig. 2A.
7 Fig. 3 is a plan view of a typical turbofan 8 aircraft. Fig. 3 shows typical external-to-the-engine Sensor 9 69 locations (near field) 38 on the fuselage 65, wings 67, tail 68, and stabilizers 66. These error Sensor signals 38 11 and the internal error Sensor signals 30, 32, and 34 combine 12 to improve the noise reduction performance in the near field 13 external to the engines. A plurality of external error 14 Sensors 69 are installed on each aircraft. The error Sensors 69 may be common to one or more engines.
16 Fig. 4A is a cross section view of typical air 17 stream modulator 115 and a horn 116 installed on and within 18 the bypass ducts 16. This device is one type of acoustic 19 Actuator that provides cancellation acoustic waves 27. An air stream modulator 115 requires a flow of compressed air. To 21 operate efficiently, pressures in the plenum 52 are at least 22 twice as high as the ambient bypass duct 16 air pressure. The 23 higher the mass flow the greater the sound output. Therefore, 24 it is an efficient device for operation on a turbofan engine 2 where enough air flow is available. Since most of the needed 26 acoustic energy is provided by the compressor stages the 27 electrical input power required to drive the modulator's 28 armature 117 is a minimum. Fig. 4A is a cross-section view of 29 a typical air modulator 115 and horn 116. Fig. 4A illustrates the relationships of high pressure supply 54, high pressure 31 plenum 52, armature 117, stator 118, driver 51, acoustic horn 32 116 and attachment mechanisms 72.
33 The acoustic exit horn 116 couples the acoustic wave 34 energy to the air stream at the horn mouth 71. The air supply plenum 52 sustains a pressure head on the modulator 115. Fig.
36 4C is a cross-section view along the lines E-E of Fig. 4A that 37 illustrates the concentric relationships of the air - W094/08540 2 ~ ~ 7 1 8 3 PCT/US93/09739 1 modulator's cylindrical armature 117 and stator 118. The 2 driver 51 is a stiffness-controlled high-force, large-3 displacement electrodynamic actuator, or the equivalent. That 4 is, it operates in the stiffness-controlled region of its response spectrum. Fig. 4D is a top view of an embodiment of 6 an air modulator 115 and acoustic horn 116.
7 The plenum 52 provides a supply of air at sustained 8 pressure. The driver 51 vibrates the armature 117 such that 9 the moving slots in the armature 117 move in relation to the stationary slots in the stator 118. This motion modulates the 11 openings between the two opposing sets of slots. Changes in 12 opening size modulates the air flow, which in turn, modulates 13 the pressure at the horn throat 70. Controlled modulation of 14 pressure generated controlled acoustic waves that propagate from the throat 70 of the horn 116 to its mouth 71. The horn 16 geometry is designed to provide an impedance load at its 17 throat 70 that improves its low-frequency performance. This 18 extends the useful bandwidth of the air modulator 115. The 19 exit horn 116 provides an impedance match between the horn mouth 71 and the air stream in the bypass duct 16. As a 21 result, the generated sound level increases and the acoustic 22 transfer efficiency to the ambient air stream improves.
23 Cancellation occurs through the mechanisms of wave 24 interference. That is, the generated acoustic waves carried by the bypass air stream mix with the acoustic waves produced 26 by the bypass air flow 17 and exhaust gas flow 21.
27 Fig. SA is a plan view of typical installations of 28 high-intensity acoustic loudspeaker Actuators (speakers) 50 in 29 the inlet duct 13, the bypass ducts 16, and the forward end of the tailpipe 15. This device 50 represents another acoustic 31 Actuator type. This Actuator provides a source of 32 cancellation acoustic waves. Each separately controlled 33 speaker requires a cooling air flow 16 around its driver 34 mechanism 51 and heat exchanger 76. A speaker requires electrical energy in order to provide acoustic energy.
36 Although speakers are not very efficient, they provide a broad 37 range of high frequency responses. Fig. 5B is a cross-section W094/08540 2 1 4 7 1 8 3 PCT/US93/09739 ~~

1 view along lines L-L of Fig. 5A that illustrates typical 2 relationships between driver 79, 'former' and 'cone' assembly 3 77, frame 75, and cooling heat exchanger 76. The motion of 4 the driver voice coil 79 drives the former and cone assembly 77. The 'cone' 77 couples the mechanical motion to the air 6 medium 13, 15 and 16 and produces acoustic waves 27.
7 Typically, an audio signal 35, 36 or 37 provided to the voice 8 coil 79 causes it to vibrate within the magnetic field that 9 exists between the permanent magnets 80 and the driver core and pole piece 81. This vibration, in turn, vibrates the 11 'cone' 77 and couples the audio signal 35, 36 or 37 to the air 12 medium 13, 15 or 16 in the form of an acoustic wave 27.
13 Fig. 6A is a cross-section view of a typical high-14 force piezoceramic Actuator assembly 84. A piezoceramic Actuator assembly 84 is made up of a mosaic of piezoceramic 16 Actuator blocks 82. These blocks 82 produce forces and 17 motions through piezoelectric activity when a voltage is 18 applied across opposite faces of the blocks 82. Typical 19 installations include: the inlet duct 13, the bypass ducts 16, the struts between the bypass ducts and the engine cases 64, 21 the stator blades of the bypass fan 56, and the stator blades 22 of successive compressor stages 18. Piezoceramic Actuator 23 assemblies 84 can bend and stretch panels 83 to produce 24 vibratory motion that also generates acoustic waves 27. Fig.
6B is a cross-section view that illustrates the piezoceramic 26 Actuator assembly 84 installed to impart motion to a panel 83.
27 Fig. 7 is a cross-section view of piezoceramic 28 Actuators assemblies 84 mounted on both sides of a cylindrical 29 engine duct 85. Applying the same in-phase audio voltage signal to both sides of the Actuator assembly 84 produces 31 stretching of the panel 83 or 85. Applying the same out-of-32 phase audio voltage signal to both sides of the Actuator 33 assembly 84 produces bending of the panel 83 or 85.
34 Piezoceramic Actuators 82 applied to only one side of a panel 83 or 85 produce combined stretching and bending. The 36 flexing panels act like speakers hence they will produce 37 sound. Piezoceramic Actuator assemblies 84 produce large - W094/08540 2 1 ~ 7 1 8 ~ PCT/US93/09739 1 forces, however the motions are small. Piezoceramic Actuators 2 produce high acoustic levels in the medium to high frequency 3 ranges.
4 Fig. 2A includes typical Sensor installations. Fig.
8 is a plan view of one of these installations along the lines - 6 I-I of Fig. 2A. The Sensors measure the analog fluctuation of 7 acoustic pressures and vibration versus time acting on its 8 sensitive face 86. The frequency sensitivity of the Sensors 9 is at least as wide as the bandwidth of engine noise. Their amplitude sensitivity has enough dynamic range to measure the 11 full variation of ranges of flight dynamic and environmental 12 conditions encountered over the engine's frequency response 13 bandwidth. The Sensor 49 consists of the active element 86, 14 mounting block 90, gas seal washer 89, jam nut 87 and electrical connector 88. Fig. 8 shows the Sensor assembly 49 16 attached to a panel 83.
17 Fig. 9 is a simplified block diagram of the typical 18 acoustic noise reduction system instrumentation 119 installed 19 on an aircraft. Sensors 91 (49 and 69) generate source and error signals. Sensor Signal Conditioning amplifiers 93 21 provide impedance matching, charge-to-voltage conversion, and 22 other conditioning of these signals. Limiting the frequency 23 range and adjusting the overall gain is the function of the 24 input programmable gains and filters 95. An analog-to-digital conversion subsystem 39 provides for conversion of these 26 analog input signals to digital samples. Synchronizing the 27 sampling rate to the engines rotational cycle reference signal 28 45 is the job of the rotational speed synchronizer 101. This 29 unit 101 maintains a constant number of samples per rotational cycle regardless of engine speed. Another way of stating this 31 process is that the digital samples now represent the signals 32 in the 'revolution' domain synonymous to the 'time' domain.
33 Fig. lOA is a series of diagrams that illustrate how 34 the Controller is synchronized to a reference signal 45, such as a once-per-revolution 'keyphasor' reference on a rotating 36 machinery shaft. In Fig. lOA, view 'A' shows a signal that 37 varies from 1 Hz to 2 Hz to 1.5 Hz in the time domain on the W O 94/08540 '~ 1 ~ 7 1 8 3 PC~r/US93/09739 ,;,, 1 left side of the figure. The right side of view A shows how 2 these waveforms correspond to the first order in the order 3 domain when the sampling rate is eight samples per cycle for 4 each example. In the frequency domain these components would appear at 1 Hz, 2 Hz, and 1.5 Hz, respectively, when the 6 sampling rate was a constant eight samples per second for all 7 three examples.
8 View B of Fig. 10A illustrates the correspondence of 9 time and revolutions for a single rotating frequency component when the speed of rotation is varying in time. Views C and D
11 show order-tracking in the order domain using synchronous 12 sampling at 10 samples per revolution. Note that the first 13 five harmonics of the rotational speed are stationary in the 14 order axis. Also, note that their amplitudes vary in the amplitude axis as speed varies. This feature is illustrated 16 by the second order in views C and D.
17 On a typical turbofan engine, the rotational speed 18 is available as Nl and N2. Nl and N2 are tachometers that 19 measure the speeds of the two shafts of the turbofan engine.
Synchronizing the neurocontroller's operations to either Nl or 21 N2 means that the inputs 31 and 33 errors 30, 32, 34 and 38 22 and outputs 35, 36 and 37 to/from the neurocontroller 112 will 23 be sampled at an integer multiple of either N1 or N2. Since 24 the sampled data rate controls the neurocontroller operation, the neurocontroller process synchronizes to this reference 26 function 45. Tachometer-controlled sampling of the input 27 source 31 and 33 and error 30, 32, 34 and 38 signals converts 28 these signals from the analog time domain to the digital 29 revolution domain. A tachometer-controlled operation means that the Controller operation is proportional to a multiple of 31 the machinery rotational speed. Synchronization transforms 32 the control topology from the traditional time and frequency 33 domains to the corresponding rotational domains of 34 'revolutions' and harmonic 'orders'. The neurocontroller places emphasis on control of rotation-related 36 acoustics/vibration and becomes more responsive to their 37 variations. The conversion of digital outputs 35, 36, and 37 -- W094/08~0 ~ 8 3 PCT/US93/09739 1 back to analog are tachometer-controlled. This digital-to-2 analog conversion process returns the digitally-sampled output 3 data from the revolution domain back to the time domain.
4 The current filtered-x, multiple neural network controller architecture eliminates shortcomings of earlier 6 work, (Bozich, D.J. and MacKay, H.B., Neurocontrollers Applied 7 To Real-Time Vibration Cancellation At Multi~le Locations, 8 Conference on Recent Advances In Active Control Of Sound And 9 Vibration, V.P.I., April 15-17 1991) using a single neural network with a feedforward architecture and the back 11 propagation (BP) learning algorithm. The architecture used 12 here is derived from the "filtered-x" adaptive controllers. A
13 significant difference between adaptive Neurocontrollers and 14 conventional adaptive controllers is the utilization of BP
neural networks in place of the more conventional LMS
16 networks.
17 A neural network is an arrangement of interconnected 18 units modelled after similar structures in the nervous system 19 of living organisms. The connections between the units are each governed by a modifiable weight. Each neuron analogue is 21 associated with a number termed its activity. Each unit 22 converts patterns of incoming activities into outgoing 23 activities that are sent to other units. This is accomplished 24 by multiplying each incoming activity by the weight of the connection and then summing the weighted inputs to get a total 26 input. An input-output function transforms the total input 27 into an outgoing activity. Thus the performance of the neural 28 network depends on both the weights and the input-output 29 function specified for the units. Typically there are three layers of units termed the input, hidden, and output layers.
31 The activity of the units in the hidden layer are determined 32 by the activities of the input units and the weights on the 33 connections between them and the hidden units. The activity 34 of the output units depends on the activity of the hidden units and the weights between the hidden and output units.
36 The advantage of the neural network is that the hidden units 37 representation of the inputs is not pre-programmed but is W094/08540 2~ ~7 ~83 PCT/US93/09739 1 adjusted during a process called training. Training consists 2 of exposing the neural network to a pattern of activities and 3 adjusting the weight of each connection so that the neural 4 network produces a more acceptable output response.
Fig. lOB is a block diagram of the Neurocontroller 6 112. The Controller 113 incorporates a neural network based 7 version of the filtered-x adaptive controller. The neuro-8 controller uses multiple networks, namely, an emulator network 9 ('Neuroemulator') 107 and a controller network ('Neurocontroller') 108. The Neuroemulator 107 develops an 11 on-line model of the dynamics of the physical system. The 12 physical system includes actuators 92 and sensors 91. This 13 emulation of the physical system provides the system 14 identification necessary to enable the neurocontroller 112 to span the required operating frequency and amplitude ranges.
16 Next, a copy of the trained Neuroemulator 110 placed in front 17 of the controller network 109 provides adaptive training of 18 the Neurocontroller 112. A current copy of the trained 19 controller network 108 placed in front of the physical system effectively preprocesses (a feedforward operation) the input 21 signal 31 and 33 such that the measured errors 30, 32, 34 and 22 38 are minimized.
23 This approach minimizes the instabilities due to the 24 phase delays of the error signals returned to the controller 113. The outstanding feature of the filtered-x approach is 26 that the process of adapting (training) the Neurocontroller 27 112 becomes the last operation instead of the first operation.
28 Therefore the delay between error signals 30, 32, 34 and 38 29 and the neuro-controller output signals 35, 36, 37 is minimal.
For a given Neurocontroller 112 output response, the resultant 31 error signal naturally delays through the actions of the 32 physical system's delay between the transfer functions. The 33 myriad transfer functions include the cross-transfer functions 34 between the cancellation Actuators 92, the acoustic and structural responses, and the error Sensors 91. The cross-36 transfer functions between the acoustic source Sensor(s) 31 37 and 33 and the acoustic error Sensors 30, 32, 34, and 38 are W094/08540 ~1 4 7 1 8 3 PCT/US93/09739 l products of the Controller transfer function and the 2 actuator/acoustic/structure/sensor transfer functions.
3 Therefore, measures of the effects of the Controller 113 on 4 canceling the effects of the system are delayed. For a moderately linear system, it would be possible to switch the 6 order of the Controller and the 7 actuator/acoustic/structure/sensor operations and obtain the 8 same overall transfer function.
9 To achieve this switch in the order of Controller 113 and system operations, the 11 actuator/acoustic/structure/sensor transfer functions are 12 first learned by the emulator network as shown in Fig. llA.
13 During this emulator training mode, the controller network 108 14 is disabled and bypassed in order to feed the acoustic source input 111 directly to the emulator 107 and the cancellation 16 actuators 92. Fig. llB shows the emulator input signal 17 consisting of the digitized 39 and summed 120 (fused) actuator 18 input signals (source sensor inputs). The summed input signal 19 120 is placed into a ring buffer or shift register 121 that contains a shifted sequence of past input signal samples.
21 This sequence of samples serves as the parallel input layer 22 for the neuroemulator 107. The emulator's outputs are 23 compared 122 to the digitized 39 error sensor responses 91 to 24 produce difference signals, that is, errors for training the emulator. The emulator training continues until all 26 differences are reduced to zero, or at least minimized.
27 Fig. llC illustrates the timing involved with the 28 events during one sample epoch of neuroemulator training. The 29 input signals are digitized (ADC) 39 and summed 120 to form the nth sample of the input signal 121 or the neuroemulator 31 and the summed input signal is converted to parallel analog 32 signals (DAC) 40 to drive the actuators 92. The neuroemulator 33 performs a feed forward operation of the nth set of the 34 shifted input signal sequence through the network and obtains a set of outputs. The nth set of error sensor response signal 36 samples that were digitized 39 at the same time as the input 37 signals are compared to these network outputs to obtain 21~7183 W O 94/08540 ~ ` PC~r/US93/09739 1 differences that serve as errors to be backpropagated through 2 the network to adjust the network weights for the next n+lth 3 sample epoch. This training process continues until the 4 differences (errors) reduce to zero or minimize.
When switching from the emulator training mode to 6 the control mode, the controller bypass and the emulator 7 training mode are disabled and the controller output signals 8 108 are enabled.
g As indicated in Fig. 12A shows that the trained emulator weights are the emulator image network 110 weights.
11 The emulator image applies the 12 actuator/acoustic/structure/sensor system transfer functions 13 between the acoustic source input signal and the controller 14 image network 109. Fig. 12B shows the trained neuroemulator input signal sequence consisting of the digitized 39 and 16 summed 120 (fused) source sensor inputs. The summed input 17 signal 120 is placed into a ring buffer or shift register 121 18 that contains a shifted sequence of past input signal samples.
19 This sequence of samples serves as the parallel input layer for the neuroemulator 107. The emulator performs a 21 feedforward operation. The outputs of the emulator image 22 ne~work 110 are summed (fused) and provide the input to the 23 controller image network's ring buffer or shift register 124.
24 The ring buffer or shift register 121 also contains the shifted sequence of the past n-lth input signal samples. This 26 n-lth sequence of samples serves as the parallel input layer 27 for the neurocontroller 108. The controller performs a 28 feedforward operation. The outputs of the controller network 29 108 are converted to analog signals 40 and drive the actuators 92. The emulator output sequence contained in the ring buffer 31 or shift register 124 is used to train the controller image 32 network 109. The controller image network trains the weights 33 of the controller network 108 as if the physical System 34 preceded the controller 113, thereby decreasing the phase delays of the measured error signals 30, 32, 34, and 38.
36 During this process, the controller network 108 provides the 37 actual signals to the physical System using the trained - W094/08540 2 1 ~ 7 1 ~ 3 - ` PCT/US93/09739 1 weights. The controller 112 requires constant-pressure (DC) 2 compensation loops. Also, high-pass filtering of the output 3 signals of the controller network extract the DC components of 4 these signals.
Fig. 12D illustrates the timing involved with the 6 events during one sample epoch of neurocontroller operation 7 consisting of feedforward operations through both the nth 8 instance of the trained neuroemulator and the n-lth instance 9 of the trained neurocontroller, and the training of the nth instance of neurocontroller weights through a process of 11 feedforward and error backpropagation. In advance of step 1 12 of Fig. 12D, as shown in Fig. 12B, the input signals are 13 digitized (ADC) 39 and summed 120 to provide the nth sample 14 sequence of the input signal 121 for the neuroemulator image 110 and the n-lth sample sequence for the neurocontroller 16 image 108. In step 1, as indicated in Figs. 12B and 12D, the 17 trained neuroemulator performs a feed forward operation of the 18 nth set of the shifted input signal sequence through the 19 network and obtains a set of outputs that are summed (fused) 123 and placed in the shifted ring buffer or shift register 21 124. The trained neurocontroller performs a feedforward 22 operation of the n-lth set of the shifted input signal 23 sequence through the network and obtains a set of outputs that 24 drive the actuators. In step 2, as indicated in Figs. 12C and 12D, the nth emulator output sequence from the ring buffer 124 26 are the parallel inputs for the feedforward operation of the 27 controller image network 109. The nth set of error sensor 28 response signal samples that were digitized 39 at the same 29 time as the nth set of input signal samples are the errors to be backpropagated through the controller image network 109 to 31 adjust the network weights for the next n+lth sample epoch.
32 This training process continues until the response sensor 33 outputs (errors) reduce to zero or minimize.
34 Once the errors minimize, the neurocontroller training is essentially complete. The training can continue 36 to track non-stationary changes in the excitation environment, 37 as shown in Fig. 12A. If the excitation is steady-state W094/08540 ~ 3 PCT/US93/09739 1 (stationary), the training period can end and the trained 2 controller 113 continues to control as shown in Fig. 13A.
3 Fig. 13B and 13C are the same operations as explained for Fig.
4 12B and step 1 of 12D, respectively.
Once the errors minimize, the neurocontroller 6 training is essentially complete. The training can continue 7 to track non-stationary changes in the excitation environment, 8 as shown in Fig. 12A-D. If the excitation is steady-state 9 (stationary), the training period can end and the trained controller 113 continues to control as shown in Fig. 13A-C.
11 The Controller 113 includes the adaptive 12 neurocontroller hardware and software system 112. The 13 controller 113 comprises: a very-high-speed parallel processor 14 42; program management computer 43; I/O signals management computer 41 (with analog-to-digital conversion sub-systems 16 39); digital-to-analog conversion sub-systems 40 (with 17 interfaces to source sensors 31 and 33, error Sensors 49 and 18 69, and cancellation Actuators 50, 115 and 84); and support 19 software and the adaptive control and neurocontrol software.
While there have been shown and described and 21 pointed out the fundamental novel features of the invention as 22 applied to preferred embodiments thereof, it will be 23 understood that various omissions and substitutions and 24 changes in the form and details of the device illustrated and in its operation may be made by those skilled in the art 26 without departing from the spirit of the invention. It is the 27 intention, therefore, to be limited only as indicated by the 28 scope of the claims appended hereto.

Claims (17)

- 25What is claimed is:

1. A system for reducing the acoustic levels of internal and external sound fields generated by gas turbine engines of a jet aircraft having an aircraft fuselage, wings, nacelles, tail, and stabilizer structures, said system comprising a plurality of actuator means to generate sound, a plurality of sensor means to measure said acoustic levels, said sensor means comprising microphone means located in and around the engines, and on a plurality of aircraft structures selected from among the aircraft fuselage, wings, nacelles, tail, and stabilizer structures, one or more controller means to control said actuators to generate sound in order to effect the reduction of the internal and external sound field as measured by the plurality of sensors.

2. The system for reducing the acoustic levels of internal and external sound fields generated by gas turbine engines of a jet aircraft of claim 1 wherein said actuator means comprises an air modulator-based acoustic source that receives a flow of compressed air derived from compression stages of one or more of said engines.

3. The system for reducing the acoustic levels of internal and external sound fields generated by gas turbine engines of a jet aircraft of claim 1 wherein said engines have external far and near acoustic radiation fields and said sensor means comprise first error sensors located inside the engine, and second error sensors located in the far and near external acoustic radiation fields of the engine.

4. The system for reducing the acoustic levels of internal and external sound fields generated by gas turbine engines of a jet aircraft of claim 1 wherein said controller means receives signals from a plurality of said sensors and comprises means to control a plurality of actuator means simultaneously to suppress acoustic noise from multiple engines.

5. The system for reducing the acoustic levels of internal and external sound fields generated by gas turbine engines of a jet aircraft of claim 1 wherein said engines provide signals to keyphasor or tachometer means and said controller receives rotational cycle reference signals from the keyphasor or tachometer means, wherein said system learns and adapts in real time to changes in residual acoustic fields.

6. The system for reducing the acoustic levels of internal and external sound fields generated by gas turbine engines of a jet aircraft of claim 5, wherein said engines provide a rotational cycle reference signal and said sensor means provide signals to said controller and said system further comprises means to convert signals from the sensors to digital samples, said means to convert operating at a sampling rate, and rotational speed synchronization means to synchronize the sampling rate to the engines rotational cycle reference signal.

7. A jet aircraft having an active system for reducing the acoustic levels of internal and external sound fields generated by gas turbine engines of the jet aircraft comprising a plurality of actuator means to generate sound, a plurality of sensor means to measure said acoustic levels, said sensor means comprising one or more controller means for controlling said actuators to radiate acoustic energy to cancel acoustic levels measured by said sensor means, said jet engine having a nacelle comprising a cowl, an inlet duct, engine and cowl panels, a bypass duct, an exhaust duct that comprises a tailpipe extension, thrust reverser, and tailpipe, said nacelle having zones containing noise source sensors, internal error sensors, external error sensors, and canceling actuators.

8. The jet aircraft of claim 7 having forward and rearward ends, wherein said nacelle has an inlet end and comprises a nose cowl having an inlet duct at the inlet end of said nacelle, an engine and cowl panel having a bypass duct immediately behind said nose cowl, a tailpipe extension behind said engine and cowl panel, said tailpipe extension having an exhaust duct and a thrust reverser and tail pipe behind said tailpipe extension, wherein said zones comprising noise source sensors comprise a first noise source sensor zone in the forward end of said engine and cowl panel, a second noise source sensor zone in the forward end of said tailpipe extension, and wherein said zones comprising error sensors comprise a first error sensor zone in the forward end of said nose cowl, a second error sensor zone in the rearward end of said engine and cowl panel, and a third error sensor zone in said thrust reverser and tailpipe, and actuator zones comprising a first actuator zone in the inlet duct of said nose cowl aft of said first error sensor zone, a second actuator zone in the bypass duct of said engine and cowl panel between said first noise source sensor zone and said second error sensor zone, and a third actuator zone located in the exhaust duct of said tailpipe extension located aft of said second noise source sensor zone.

9. The jet aircraft of claim 8 wherein the actuators located in one or more of said first, second and third actuator zones comprise air stream modulators.

10. The jet aircraft of claim 8 wherein the actuators located in one or more of said first, second and third actuator zones comprise piezoceramic actuators.

11. The jet aircraft of claim 9 wherein one or more of said air stream modulators comprises an air modulator actuator horn having a mouth opening into an ambient bypass air flow, said horn comprising throat means to provide impedance loading to enhance performance, and horn section means to match the impedance of the ambient bypass air flow at the mouth of the horn.

12. The jet aircraft of claim 11 wherein said horn comprises a mouth having a flange attached to an entry hole in an engine bypass duct wall, a plenum, and a pipe manifold attached to a source of air.

13. The jet aircraft of claim 12 wherein said engine comprises compressor stages and bleed ports on the engine's compressor stages, and wherein said pipe manifold is adapted to receive compressed air from said bleed ports.

14. The jet aircraft of claim 13 wherein said throat means of said horn provides an exit from the plenum, and said mouth of said horn provides an entrance back into a bypass duct, each engine having a plurality of air modulator actuators.

15. The jet aircraft of claim 14 wherein said horn comprises driver means.

16. The jet aircraft of claim 15 wherein said driver means comprises a stiffness-controlled, electrodynamic actuator.

17. The jet aircraft of claim 7 wherein said actuator means comprise acoustic loudspeakers located in one or more of the inlet duct, a bypass duct, and the forward end of the tailpipe.