CA2235309A1 - Active noise and vibration control, especially in aircraft - Google Patents

Abstract

A system for controlling noise and/or vibration in a structure, especially an aircraft fuselage, of the type having structural components which include stiffening members fixed to a wall or skin. A series of sensors attached to the structure is capable of registering noise or changes in accelerations or strains in the structure caused by noise or vibrations in the environment, and actuators of adaptive material, for example formed of piezoelectric elements, are connected to the sensors by control means for causing the actuators to counteract noises or vibrations sensed by the sensors. In accordance with the invention the actuators are fixed to the stiffening members.
Preferably, the actuators are fixed to flanges of the stiffening members which are spaced from the wall or skin by a web. In an aircraft fuselage, some actuators are fixed to stiffening members which extend across other stiffening members.

Description

Title: "Active noise and vibration control, especially in aircraft"
Background of the Invention.
l.Field of the Invention _5 The present invention relates to noise and vibration control, especially to reduce the noise and vibrations experienced by passengers in turbopropeller driven aircraft.
However, the invention may also have uses for reducing noise and vibration in other aircraft, including helicopters, and in motor vehicles, ships, submarines, and in static structures such as buildings and bridges.

2.Prior Art In turbopropeller driven aircraft undesirable amounts of noise and vibration are caused by excitation of the fuselage by _15 the unsteady aerodynamic pressure field of the propeller. Since the interior noise spectrum is dominated by tones occurring at integral multiples of the blade passage frequency (BPF), the reduction of interior noise and fuselage vibrations at these discrete frequencies provides a significant improvement.
_20 Various means have been tried in the past to reduce noise and vibration in propeller aircraft, including passive techniques, tuned vibration absorbers, and active noise control through the use of secondary acoustic sources. So-called active structural acoustic control has also been proposed.
_25 Passive techniques of noise and vibration reduction for this purpose are generally not effective for a number of reasons. Since the excitation is usually neither broadband nor resonant, the addition of damping material does not have a major effect on noise and vibration levels. Moreover, passive damping treatments are ineffective at the low frequencies where propeller noise is significant, i.e. 50 to 300 Hz, since for the sound absorption material to be effective the thickness of the material must be comparable to the wavelengths of vibration, which are long at this frequency range.
Tuned Vibration Absorbers (TVA), which are located on fuselage frames and designed with a resonant frequency equal to that of the BPF, thereby reducing the fuselage vibration and noise transmission, have also been considered for aircraft application. When subjected to a stationary harmonic disturbance, TVA's that are tuned accurately can be effective, but the variations in engine speed greatly limit their performance in propeller aircraft due to the required trade-off between bandwidth and peak achievable attenuation.
Systems of Active Noise Control (ANC) of the aircraft cabin through the use of secondary acoustic sources are presently offered by a number of aircraft manufacturers. These systems incorporate a large number of trim mounted speakers along with microphones and/or accelerometers to achieve noise reduction through destructive interference between the primary (disturbance) sources and the secondary (control) sources.
These systems have drawbacks. Although they are capable of reducing the peak noise level at discrete locations, the overall effect is not generally reduced. Also, the performance of these systems can be affected by changes in the passenger cabin, including the number and location of passengers, thereby reducing their effectiveness. In addition, they are not effective in reducing structural vibration.

Another previously proposed technique is known as Active Structural Acoustic Control (ASAC) . Here an array of structural actuators is distributed over an area of the aircraft fuselage, and is connected via a control circuit to sensors embedded in or adhered to the structure. The control circuit causes the actuators to adaptively alter the mechanical properties of the structure, and reduces the levels of noise and vibration. The sensors include accelerometers and/or microphones, and the actuators may be piezoelectric or inertia force actuators.
Such ASAC systems require fewer control sources, such as the actuators, and require less complex control, than required with the ANC approach.
Summarv of the Invention Hitherto, proposed systems for using ASAC have involved covering a substantial area of an aircraft fuselage skin with PZT actuators. These are a form of piezoceramic actuators available in sheet-like form which are fixed to the skin of the fuselage. The sheet-like form of these actuators limits the kind of surface to which they can be attached, and also limits their size, since the interior skin of a fuselage has only fairly small flat or substantially flat areas between stiffeners. Applicants have discovered that better results for reducing propeller noise and fuselage vibration can be achieved by using actuators fixed to the circumferential or longitudinal stiffening members which reinforce and are fixed to the fuselage skin or shell. A smaller number of actuators is required compared to the conventional approach. Also, the actuators used in accordance with this invention are not limited in their dimensions to the distances between adjacent stiffening members.

The invention is not limited to aircraft fuselages, and in general the invention provides a system for controlling noise and/or vibration in a structure of the type having structural components which include stiffening members fixed to a wall, of the type comprising:
a series of sensors fixed to the structure and capable of registering noise or changes in accelerations or strains in the structure caused by noise or vibrations in the environment adjacent the structure;
a series of actuators of adaptive material fixed to the structural components and capable of actively varying the vibrational properties of the components to which they are attached, and control means connected to the sensors and to the actuators for causing the actuators to counteract noise or vibrations sensed by the sensors.
The invention is characterized in that the actuators are fixed to the stiffening members.
An advantage of the ASAC system is that it involves reducing the transmission of noise before it enters the fuselage; generally treating noise at or close to its source yields superior performance as compared to treating the problem further down the transmission path. The piezoelectric elements may be used to cover critical regions only of the structure thereby providing the actuator with the ability of selectively controlling modes of vibration. For vibroacoustic problems, these will be those structural modes which are efficient radiators of sound.
The actuators used with this invention may be fixed exclusively to the stiffening members, or a series of actuators may be fixed to the stiffening members, and others fixed to a fuselage skin. Where, as is usual, such members have an outer flange fixed to the skin of the fuselage and an inner flange spaced from the outer flange by a web, the actuators may be fixed along the inner or outer flanges or along the web.
In addition to its use in propeller driven aircraft, the system may be used for other structures which have a wall, such as a skin or metal sheet , f fixed to st i f fening members . Such structures include the fuselages of helicopters, and the bodies of motor vehicles, and also static structures such as buildings and bridges.
For aircraft fuselage use, piezoceramic elements are preferred. However, other types of piezoelectric elements may be used. Also, other forms of so-called "smart" materials which respond to electrical or other signals may be used; such materials are generically referred to as "adaptive" materials since they adapt to circumstances of use. Suitable materials also include electrostrictive or magnetostrictive materials.
Brief Description of the drawinQS.
Preferred embodiments of the invention will now be described by way of example with reference to the accompanying drawings, in which;
Fig.l is a diagrammatic representation of a fuselage section showing the area which is affected by acoustic disturbance from the propellers, and the propeller pressure distribution;
Fig.2a is a cross-sectional view through a portion of an aircraft fuselage, showing a known arrangement of several actuators and their relationship to the structure, and Fig.2b is a similar view showing an arrangement in accordance with the invention;
Figs. 3a, 3b, and 3c are each a perspective view of a portion of the inside of a structure having a metal sheet such as a fuselage skin reinforced by stiffeners, and showing different arrangements of actuators on the stiffeners, in accordance with the present invention;
Fig.4 is a diagrammatic cross-sectional view of a fuselage showing one arrangement of actuators in accordance with the invention and parts of a control circuit; and Figs.5(a), (b), (c) and (d) are diagrammatic views of piezoelectric actuators attached to structural members, illustrating different relationships between length of actuator and mode of vibration.
Detailed Description.
Fig.l illustrates propeller pressure distribution over the fuselage of a typical turbo-prop aircraft which was used in experiments on this invention which are described below. Fig.l also indicates the propeller plane P, and the positions of the fore and aft bulkheads B1 and B2 and typical ribs 26. The actuators used to control noise and vibration from the propellers are placed in or close to the area indicated at A in Fig. 1. Typical actuators in accordance with the invention, which are described more fully below, are indicated at 20.
Fig.2a shows diagrammatically the prior art arrangement of a typical piezoceramic PZT type actuator 18 as previously used on a section of fuselage. The actuator 18 comprises a sheet of PZT material typically 5 inches wide and about 20 inches in length. As shown, each actuator 18 is fixed by bonding material 21 to the interior surface of the metal fuselage skin 22.
Typical bonding materials which have been used are structural adhesives and epoxys, and typical thickness of the bonding material is 2 to 10 thousands of an inch. The actuators are arranged in the relatively flat, although usually slightly curved, areas beside longitudinal structural stiffeners or so-y called stringers 24, and between circumferential stiffeners or frames or ribs 26, described further below. The term "circumferential" in this context means that these ribs extend around the fuselage, and does not imply that all their parts are curved, since in some cases fuselages have flat sides.
As illustrated in Fig.2a, the actuator 18 transfers strain to the fuselage skin 22 largely by bending moments indicated at M to which the skin is subjected at each end of the actuator.
As the bonding layer is made thinner and/or stiffer, the transfer of loading occurs over progressively smaller regions located at the edges of the actuator. In the limit, the loading reduces to moments along the four edges of the actuator.
It has been realized that the effectiveness of the actuators in transferring energy to the fuselage also depends on the moment arm represented by the distance between the neutral axis of the structure, shown at NA in Fig.2a, and the interface of the actuator 18 and the fuselage skin 22, represented by the bonding layer. This distance is relatively small in the known Fig.2a arrangement.
Fig.2b shows an arrangement in accordance with the invention, in which an actuator 20 comprising a series of piezoelectric elements 20' is located along the inner flange 24a of a longitudinal stiffener or stringer 24. The elements 20' have their ends close together, and although their ends are not touching the f act that the adj acent ends are all f firmly bonded to the stiffener surface means that the elements are effectively fixed end-to-end. Also, all the elements are provided with an identical electrical field, so that the series of elements behaves as one large element or actuator. The use of a large actuator formed from a series of elements is usually desirable since it allows the actuator to control long wavelength, and therefor low frequency, vibration.
As also shown more fully in Fig.3a, each stringer 24 is of generally Z form and has an outer flange 24c, parallel to the inner flange 24a, and fastened, a . g. rivetted or bonded, to the fuselage skin 22, the flanges being connected by a radial web 24b. The stringers pass through apertures 28 in the ribs or circumferential stiffeners or ribs 26, each of which is in the form of a channel, having an outer flange 26c fastened the skin 22, a web 26b projecting into the fuselage interior, and an inner flange 26a parallel with the outer flange.
Referring again to Fig.2b, the elements 20' are held to the inner flange 24a by a relatively thin layer of bonding material 21, so as to cause primarily bending strain in the stiffener 24. It will be seen that in the Fig.2b arrangement there is a relatively large off-set between the stress points represented by the bonding layer 21 and the neutral axis NA, and accordingly the actuator applies much more bending moment to the structure than in the prior art arrangement.
A further advantage of the Fig.2b and Fig.3a arrangement is that the effective length of an actuator is not limited to the distance between adjacent stiffeners, as in the prior art arrangement. It will be seen from Fig.3a that the actuator 20 in fact passes through the aperture 28 in the rib 26. Such apertures are conventional, allowing stiffeners to run a large part of the fuselage length, and in practice the apertures are usually large enough to accommodate the actuators which are perhaps 1/8 inch in thickness.
Fig.3a also shows actuators 20 which extend along portions of the inner flanges of ribs 26. These inner flanges are uninterrupted by the stringers, allowing long actuators to be used. Such actuators may extend between 70° and 90° around the circumference of the fuselage.
Figs. 3b and 3c show further arrangements of actuators also in accordance with the invention. In Fig.3b, an actuator 20 formed by a row of the elements 20' is secured to the inside of each outer flange 24c, 26c both of the stringers and of the ribs, being separated from the metal skin 22 by the thickness of the outer flange. Fig.3c shows a third arrangement, in which an actuator 20 is secured to each web both of the stringer 24s and of the ribs 36.
Although Figs.3a, 3b, and 3c show a flat metal sheet or skin 22, this may be somewhat curved, as when the invention is applied to a fuselage F. When used in a propeller driven aircraft the actuators are confined to the portion of the fuselage shown at A in Fig.l, i.e between the bulkheads B1 and B2, and around a large part of the two sides of a fuselage.
In each case the actuators 20 are formed by a series of piezoceramic elements 20' of nominally rectangular form each measuring 1" by 1" by 1/4" , with a distance of 1/8" maintained between neighbouring elements. The number of elements will depend on the size of the elements and the wavelength of vibration to be controlled, in accordance with principles described with reference to Figs.S(a) to 5(b) discussed below.
Both the stiffeners 24 and 26 are about 1 inch wide, and the stringer is about 3 inches deep. The ribs 26 are about 20 inches apart.
Sensors may be attached between each of the actuators, and connected to control systems of known type. A suitable control circuit is shown diagrammatically in Fig.4.
Fig.4 shows a turbo propeller driven aircraft having two wing mounted engines E on opposite sides of a generally cylindrical fuselage section F having floor F1. Several of the circumferential ribs 26 of the fuselage are each provided with two actuators 20, each comprising a row of piezoceramic elements, generally as described with reference to Fig.3a. The actuators each extend about 70-80° around the upper half of each side of the fuselage, terminating near the central horizontal plane. The controls for the actuators include both microphones and accelerometers, a typical microphone being indicated at 40 and a typical accelerometer being indicated at 42. A reference microphone 43 is also located near an engine E.
All microphones 40, 43 are connected via measuring amplifiers 44, 45 to a DSP (digital signal processor) controller 47, and the accelerometers are connected via charge amplifiers 48 to the same controller. This controller feeds signals to the actuators 20 via a power amplifier 50.
The one-sided or asymmetric actuation of a structure such as an aircraft fuselage by bonded piezoelectric actuators will produce both flexural and in-plane vibration. For the case of noise transmission into an aircraft cabin, flexural vibrations are of greater significance due to their ability to couple with acoustic waves. As indicated above, for flexural vibration, the predominant actuation mechanism due to a uniform rectangular piezoelectric actuator is the presence of bending moments at its ends. This is true of both monolithic and segmented actuators. For best results in accordance with the invention, a segmented actuator with all its elements connected in phase should be positioned such that its ends coincide with regions of high (and opposite) angular rotation in order to have high control authority over that particular mode or deflection pattern. If the ends are located in regions of high displacement (and hence low rotation), or if the rotations at the ends are nearly equal, low control authority over that mode or deflection shape results. Based on this approach, the control authority over a mode can also be enhanced by appropriately reversing the electrical field of elements over regions of opposite modal displacement. These principles are illustrated in Figs.5(a) to 5(d).
Fig. 5 (a) shows a piezoelectric actuator 20 having its ends at adjacent nodes which are regions of high rotation; this gives good control.
Fig.5 (b) shows a piezoelectric actuator 20 having its ends at points of high displacement and low rotation; this gives poor control.
Fig. 5 (c) shows a piezoelectric actuator 20 having its ends at nodes which are not adjacent, and hence which are points of high rotation but not opposite rotation; this also gives poor control.
Fig.S(d) shows, in effect, the piezoelectric actuator of Fig.5(c) split so that into two parts, 20a and 20b, which each have ends at adjacent nodes, being points of high and opposite rotation. This gives better control than the single element of Fig. 5 (a) .
While in theory it may be desirable to have an actuator of a length equal to 1/2 the wavelength of the vibration, in practice it is found that the fundamental frequency of a vibration is accompanied by higher harmonics, making it usually desirable to have an actuator somewhat shorter than the fundamental frequency.
Experimental Demonstration of Invention.
The feasibility of using piezoelectric actuators and structural sensors has been experimentally investigated and established on a full-scale fuselage. This test article was a full-sized de Havilland Dash-8 Series 100/200 fuselage without horizontal lifting surfaces. The floor was fitted in the interior, which did not include any interior trim or seats, but the cargo and pilot's cabin bulkheads were present. The fuselage rested on a cradle fitted trailer except during testing when an overhead crane was used to raise the fuselage using steel cables attached to the fuselage at the wing-fuselage interface. This arrangement subjected the fuselage to support conditions that resemble those of steady level flight.
The fuselage was tested in a high bay laboratory with the temperature maintained at approximately 20°C at ambient atmospheric pressure.
Multiple piezoelectric elements were connected together to form multiple piezoelectric actuators located near the plane of the propeller. Each actuator was a series of elements bonded closely together and subjected to the same electrical field to behave as a single long actuator. The vibration patterns or operating deflection shapes of the fuselage were used to determine the specific size and location of the actuators. In the case of flexural vibration, the predominant actuation mechanism for a rectangular piezoelectric actuator is the application of bending moments at its ends on the host _5 structure. Accordingly, each segmented actuator was positioned such that its ends coincided with regions of high (and opposite) angular rotation in order to have high control authority over that particular mode.
Structural response was measured by a number of discrete _10 sensors. The weighted average of the sensor signals was combined to ensure that the average vibration over the critical region was accurately measured and reduced by the control system. In doing so, the residual vibration pattern was reduced, resulting in lower interior noise levels.
15 A sophisticated sound source was provided consisting of a speaker-ring. The port side exterior propeller pressure field was simulated using the sound field generated by this speaker-ring with four 12" loudspeakers units. Each unit was a sealed plywood enclosure that forms the speaker box and a horn section _20 with an opening that measures 30 in. by 24 in. The speaker-ring covered an arc of approximately 100° on the fuselage and was maintained at distance of approximately 2" from the fuselage surface to allow for the blending of the sound from the individual speakers. The four individual waveforms were 25 synthesized by a National Instruments AT-AO-10 analog output card and amplified using a pair of Ashly MFA-8000 dual channel power amplifiers. This system simulated the exterior propeller pressure field at the surface of the fuselage provided by the propeller manufacturer, Hamilton Standard. Using this system, 30 the in-flight magnitude and phase distribution of noise on the exterior of the fuselage was simulated in the neighbourhood of the propeller plane. The objective of using the present speaker-ring was not to recreate the entire propeller field, but rather to represent its phase and magnitude characteristics over the key region close to the propeller plane. It can be extended in the future to provide a more representative simulation of the propeller field.
Two control strategies were implemented experimentally in this study. The first was a MIMO (multiple input multiple output) version of the filtered-x LMS (least mean square) adaptive feedforward a:Lgorithm. This control method required that the transfer functions between the control actuators and the error sensors (i.e. the secondary path) be estimated. Both on-line and off-line approaches to the estimation of these transfer functions were: used. The second control approach was _15 a MIMO extension to classical feedback design where high loop gain is introduced at the disturbance frequencies to render the sensitivity transfer functions small. Second order compensators were used in this case..
These controllers were implemented on a DSP (digital signal processor) system consisting of a Spectrum QPC40 ISA
processor board equipped with a single Texas Instruments TMS320C40 processor. I/O (Input/output) was provided on a separate ISA (Industry Standard Architecture) board with 16 channels of 12-bit analog inputs and 8 channels of 12-bit analog outputs. Communication between the processor and I/O
board was achieved via a high-speed 16-bit interface. Other features of the I/O board include programmable input gain, a maximum sample rate of 25 kHz (with all channels in use), and on-board third-order Butterworth anti-aliasing and _30 re-construction low-pass filters for all input and output channels (with -18 dB/octave roll-off).

The MIMO classical feedback controller has relatively low computational requirements. The real-time code was written in the C-language and implemented using the Texas Instruments C40 code development system. The resulting code for a 3-input, 3-output controller could be implemented at a sample rate in excess of 15 kHz. The outputs of the controller were amplified by a pair of dual channel Trek model 50/750 high voltage amplifiers designed specifically for driving active loads, which were connected to the actuators.
_10 In order to monitor the noise reduction performance, microphones were positioned at the seated height for the two port side seats in the propeller plane, and at the aisle center standing height for seat rows l, 2 and 3. The noise attenuation data for the two control strategy cases (i.e. feedback and _15 feedforward control with vibration error sensing, configuration I) showed interior noise attenuation as high as 14.8 dB, and a peak vibration reduction of 20 dB at the BPF. Moreover, in both feedback and feedforward control cases, the reduction in noise was global in nature. Noise reductions for other flight 20 conditions were also achieved.

Claims (20)

1. A system for controlling noise and/or vibration in a structure of the type having structural components which include stiffening members fixed to a wall, comprising:
a series of sensors fixed to the structure and capable of registering noise or changes in accelerations or strains in the structure caused by noise or vibrations in the environment adjacent the structure;
a series of actuators of adaptive material fixed to the structural components and capable of actively varying the vibrational properties of the components to which they are attached;
control means connected to the sensors and to the actuators for causing the actuators to counteract noise or vibrations sensed by the sensors;
characterized in that the actuators are fixed to the stiffening members.

2. A system according to claim 1, wherein the actuators are provided along several of said stiffening members.

3. A system according to claim 1, wherein at least one of said actuators is fixed to a first stiffening member and extends across or through another stiffening member.

4. A system according to claim 1, wherein said stiffening members include an outer flange fixed to the wall, and a web projecting inwardly from the wall, and wherein at least some of said actuators are fixed to said outer flange.

5. A system according to claim 1, wherein said stiffening members include an outer flange fixed to the wall, and a web projecting inwardly from the wall, and wherein at least some of said actuators are fixed to said web.

6. A system according to claim 1, wherein said stiffening members include an outer flange fixed to the wall, a web projecting inwardly from the wall, and an inner flange projecting from an inner edge of the web and parallel to the outer flange, and wherein at least some of said actuators are fixed to said inner flange.

7. A system according to claim 1, wherein said actuators each comprise a series of piezoelectric elements effectively connected together in end-to-end relationship.

8. A system according to claim 7, wherein said piezoelectric elements are piezoceramic elements.

9. A system according to claim 1 or claim 7, wherein ends of said actuators are fixed to regions of high and opposite flexural rotation during vibration of the structure.

10. A system for controlling noise and/or vibration in an aircraft fuselage of the type having structural components which include stiffening members fixed to an outer skin, comprising:
a series of sensors fixed to the fuselage and capable of registering noise or changes in accelerations or strains in the fuselage caused by noise or vibrations in the environment adjacent to said fuselage;
a series of actuators of adaptive material fixed to the structural components of the fuselage and capable of actively varying the vibrational properties of the components to which they are attached;

control means connected to the sensors and to the actuators for causing the actuators to counteract noise or vibrations sensed by the sensors;
characterized in that the actuators are fixed to the stiffening members.

11. A system according to claim 10, wherein the actuators are provided along several of said stiffening members.

12. A system according to claim 10, wherein the ends of the actuators are fixed to regions of high and opposite flexural rotation during vibration of the structure.

13. A system according to claim 10, wherein at least one of said actuators is fixed to a first stiffening member and extends across or through another stiffening member.

14. A system according to claim 10, wherein said stiffening members include an outer flange fixed to the fuselage skin, and a web projecting inwardly from the skin, and wherein at least some of said actuators are fixed to said outer flange.

15. A system according to claim 10, wherein said stiffening members include an outer flange fixed to the fuselage skin, and a web projecting inwardly from the wall, and wherein at least some of said actuators are fixed to said web.

16. A system according to claim 10, wherein said stiffening members include an outer flange fixed to the fuselage skin, a web projecting inwardly from the wall, and an inner flange projecting from an inner edge of the web and parallel to the outer flange, and wherein at least some of said actuators are fixed to said inner flange.

17. A system according to claim 1, wherein said actuators each comprise a series of piezoelectric elements effectively connected together in end-to-end relationship.

18. A system according to claim 17, wherein said piezoelectric elements are piezoceramic elements.

19. A system according to claim 10, wherein at least two actuators are situated along portions of the an inner flange of circumferential stiffening members at opposite sides of the fuselage.

20. A system according to claim 19, wherein said actuators occupy between 70 and 90° of the circumference of the fuselage.