The present invention relates to actuation of aircraft components, and in particular to determining an actual deformation undergone by an actuated aircraft component.
Many aircraft components have essentially compromised designs which provide benefits in one or more flight phases, but which generate penalties (e.g. drag penalties) in other flight phases. In aircraft design it is necessary to balance the degree of compromise to ensure that the benefits outweigh the penalties. In practice this means that the design is often compromised in all flight phases.
The present invention is concerned with actuating such aircraft components to tailor their configuration to achieve improved performance in one or more flight phases.
In general terms, the invention provides a method and system for actuating an aircraft component comprising an actuating material in which the actuating material generates an output signal in response to an actual deformation it undergoes, this output signal being representative of that actual deformation.
A first aspect of the invention provides a method of actuating an aircraft component, at least a portion of the aircraft component comprising an actuating material which undergoes deformation in response to the application of an electrical signal thereto, and which generates an electrical signal in response to a deformation of the actuating material, the method comprising the steps of: applying an activation input signal to the actuating material of the aircraft component, the activation input signal corresponding to a desired deformation of the actuating material, the actuating material of the aircraft component undergoing an actual deformation in response to the activation input signal; and generating an output signal representative of the actual deformation of the actuating material.
A second aspect of the invention provides an aircraft component actuating system for actuating an aircraft component, the system comprising: an aircraft component comprising an actuating material which is configured to deform in response to the application of an electrical signal thereto, and which is configured to generate an electrical signal in response to a deformation of the actuating material; a controller configured to transmit an activation input signal to the actuating material of the aircraft component corresponding to a desired deformation of the actuating material, and further configured to receive from the actuating material a generated output signal representative of an actual deformation of the actuating material.
The arrangement of the first and second aspect enables the configuration (i.e. shape or internal mechanical properties such as stiffness) of an aircraft component to be changed by simple application of the activation input signal. Moreover, the output signal generated by the actuating material in response to the deformation provides an indication of the nature of the actual deformation undergone. In this way the actuating material provides both an actuating function and a sensing function. The performance of the aircraft component (e.g. throughout each flight, the life of the component, or a test phase of the aircraft) can be determined using the actuating material, without the need for separate sensing devices. This provides a significant weight saving advantage, in addition to the increased ability for performance analysis.
The portion of the aircraft component comprising the actuating material may be bonded to, co-bonded with, or otherwise fixed to a substrate. Alternatively, the aircraft component may comprise a fibre-reinforced composite material, and the actuating material may comprise particles dispersed throughout the matrix of the composite material and/or may comprise a plurality of filaments extending through the matrix.
The actuating material may comprise an electro-active polymer, or any material which has piezoelectric properties such that it undergoes deformation in response to an electrical signal and produces an electrical signal in response to a deformation.
The actuating material may deform by expanding, contracting or in any other way. In some embodiments the application of the activation input signal may result in a change in stiffness of the actuating material.
The activation input signal is preferably actively controlled based upon the generated output signal, e.g. via closed-loop control. Thus, if the generated output signal indicates that the actual deformation is not within acceptable margins of the desired deformation then the activation input signal can be modified accordingly. For example, external forces applied to the aircraft component, such as aerodynamic forces, may influence the aircraft component such that the actuating material is unable to achieve the desired deformation. By monitoring the actual deformation via the output signal, the characteristics of the input activation signal can be controlled to counteract those external forces and achieve closer correspondence between the desired and actual deformations.
In some embodiments the activation input signal is controlled based upon an instruction from a flight control computer of the aircraft. The instruction from the flight control computer may itself be based on a control input from the cockpit (pilot control) or from a fly-by-wire (auto pilot) control input, or a monitoring input from a movable control surface.
In some embodiments the aircraft component is formed from a fibre-reinforced composite material, wherein the actuating material is embedded in the matrix of the composite material. The actuating material may comprise particles dispersed through the matrix, and/or filaments embedded in the matrix. Such filaments may be interwoven with reinforcing fibres of the composite material. By embedding the actuating material in the matrix in this way the number of fabrication steps is reduced, and the risk of delamination or debonding of is minimised. Alternatively, the actuating material may be co-bonded or co-cured with a substrate to form a composite material.
The aircraft component may be located upon an aerodynamic surface of the aircraft. For example, the aircraft component may comprise one of the following components: a rain gutter; a vortex generator; a NACA duct; a fuel system access component. Such components provide benefits at one or more flight phases (or on the ground), and penalties at one or more other flight phases. Thus, controlling the configuration of such components to maximise the benefits and minimise the penalties can be very advantageous.
In preferred embodiments the desired deformation serves to alter an air gap between the aircraft component and a movable control surface configured to be movable between a stowed configuration and a deployed configuration. Such gap control can be used to manage flow separation during different flight phases. For example, a convergent gap which narrows in the direction of airflow may be especially desirable in some embodiments. In particular, the movable control surface may comprise a trailing edge flap and the desired deformation may alter the air gap to provide a convergent gap between the aircraft component and the flap in the deployed configuration of the flap. The aircraft component may comprise an actuatable trailing edge portion extending from a trailing edge of a spoiler or fixed wing panel.
In other embodiments the aircraft component is a seal located between first and second surfaces, the second surface comprising a moveable control surface configured to move between a stowed configuration and a deployed configuration, wherein the desired deformation of the seal tends to urge the seal in a first direction towards the second surface.
Thus, the seal can be urged into contact with the movable control surface during cruise, in order to provide an aerodynamically beneficial profile and avoid undesirable flutter. Any such flutter, or other unfavourable aerodynamic profiles, can be diagnosed by analysis of the output signal. Moreover, the shape, contour and/or stiffness of the seal along its length can be varied in accordance with the aerodynamic requirements and/or the seal's performance.
Such embodiments preferably comprise the further steps of: applying a second activation input signal to the actuating material of the seal, the second activation input signal corresponding to a second desired deformation of the seal tending to urge the seal in a second direction opposite to the first direction, the actuating material of the seal undergoing a second actual deformation in response to the second activation input signal; and generating a second output signal representative of the second actual displacement of the actuating material of the seal.
Thus, the seal may be deflected away from the movable control surface during deployment thereof in order to avoid entrapment of the seal by the movable control surface.
In preferred seal embodiments the first or second activation input signal is applied to the actuating material of the seal in response to the movement of the second surface between the stowed configuration and deployed configuration.
The aircraft component may project from a moveable control surface, and the desired deformation of the actuating material may provide movement of the aircraft component relative to the movable control surface. Thus the aircraft component may be actuatable to provide an additional degree of control to the lift or drag influencing characteristics of the movable control surface. For example, the aircraft component may be actuatable to vary the curvature of the upper aerodynamic profile provided by the movable control surface. Such an arrangement can provide a continuously adaptable flight control surface. Moreover, by analysing the output signal it is possible to provide load alleviation and buffet reduction for reduced wing loading and/or passenger comfort.
In preferred embodiments the moveable control surface comprises a trailing edge flap configured to move between a stowed configuration and a deployed configuration, and the aircraft component projects from a trailing edge of the flap. This arrangement provides a tabbed flap which can provide a flap with a variable camber in high-lift flight phases, without adding a significant drag penalty during cruise.
In some embodiments the desired deformation and the actual deformation comprise a desired shape change and an actual shape change, respectively. That is, the deformation results in a change of shape (geometry) of the actuating material.
In other embodiments the desired deformation and the actual deformation comprise a desired generation of mechanical stress and an actual generation of mechanical stress, respectively. Thus, the deformation results in a change to the internal form/configuration of the actuating material in which internally generated forces result in a mechanical stress. This stress may result in a shape change of the actuating material (as a result of the induced strain) or a change in mechanical properties such as stiffness.
A third aspect of the invention provides an aircraft comprising the system of the second aspect.
Any of the features of the invention described herein as optional or desirable may be applied to any aspect of the invention, either individually or in any combination.
Embodiments of the invention will now be described with reference to the accompanying drawings, in which:
Electro-active polymers are polymers that undergo a deformation, i.e. change in shape and/or internal mechanical stresses/stiffness, when stimulated by an electrical field. Thus, application of an electrical signal to an electro-active polymer causes the electro-active polymer to undergo a deformation. The electrical signal is representative of the desired deformation to be achieved by the electro-active polymer. In this way, characteristics of the electrical signal can be controlled to achieve desired characteristics of the deformation.
Moreover, electro-active polymers can also provide an inverse response; that is, an electro-active polymer will generate an electrical signal in response to an actual deformation of the electro-active polymer. The generated electrical signal is representative of the actual deformation, such that characteristics of the signal can be interpreted to determine characteristics of the deformation.
In the absence of an electrical signal applied to it, the unstimulated actuating material of the aircraft component 100 remains in a neutral position as shown in
The second deformation position (
Thus, the electro-active polymer system of
In other embodiments the aircraft component 100 may not undergo a shape change, but may instead undergo another type of deformation, such as internal configuration change caused by a change in internal mechanical stresses leading to a change in mechanical stiffness.
The control system comprises an aircraft component which includes an actuating material element 200 in electrical communication with a controller 204. The element 200 is formed from an actuating material, which in this embodiment is an electro-active polymer. The controller is configured to apply an activation input signal 202 to the actuating material element 200. The activation input signal 202 corresponds to the desired deformation that the actuating material element 200 is intended to undergo.
In the embodiment of
In the embodiment of
The controller 204 is also configured to receive an output signal 206 generated by the actuating material of the element 200 in response to the actual deformation undergone by the actuating material. The actual deformation may differ to the desired deformation of the actuating material due to external forces acting upon the actuating material element 200 in addition to the actuating force. That is, the magnitude and/or direction of the actual deformation may be different from the intended desired deformation such that the actual deformation of the actuating material element 200 is not exactly as intended. Such external forces may be caused, for example, by the effects of resonant frequencies and component flutter during differing flight phases of the aircraft. The output signal 206 thus provides a signal that is representative of the actual deformation experienced by the actuating material of the element 200.
In open-loop control embodiments the output signal is not fed back to the controller 204. However, in the closed-loop control embodiment of
By feeding the output signal 206 back to the controller 204, the activation input signal 202—which corresponds to the desired deformation—may be actively controlled based upon the output signal 206, i.e. via closed-loop control. That is, if the output signal 206 indicates that the actual deformation of the actuating material element 200 is not within given tolerance boundaries of the intended desired deformation, then the activation input signal 202 may be adjusted accordingly.
This control feedback loop ensures that the actual deformation of the actuating material element 200 is within acceptable margins of the desired deformation, and allows external forces acting upon the actuating material element 200 to be compensated for. Such active control may be carried out in real time during flight of the aircraft, or may be carried out at discrete intervals during routine maintenance of the aircraft or during a flight test programme.
Signals and/or data from the controller 204 may also be fed back to the flight control computer 210 via controller feedback signal 207, and signals and/or data from the flight control computer may also be fed back to the pilot or fly-by-wire system by computer feedback signal 217. Thus, the flight control computer 210 may determine whether or not to adjust the activation input signal 202 based on the output signal 206.
The arrangement described above in relation to
This control can be directly integrated into a fly-by-wire aircraft, with the possibility of a continually variable profile which can be tunable as flight test data becomes available during early flights, or even at later stages of the aircrafts life where modifications to the aerodynamic performance of the wings, e.g. by changes to the movable control surfaces, can be accommodated by changes to the software controlling the aircraft component rather than by replacing the aircraft component.
Moreover, by taking advantage of the actuating materials ability to generate an electrical signal in response to a change in deformation, it is possible to provide positive feedback to identify and counteract undesirable behaviour of the aircraft component, such as resonant frequencies or other transient behaviours experienced by the aircraft component during a particular flight phase. Such unwanted behaviour can be identified by analysis of the generated output signal, and counteracted by modification of the activation input signal.
The arrangement illustrated in
The aircraft also comprises vortex generators 312 which serve to delay flow separation over the wing 304. The illustrated vortex generators 312 are shown on the wing lower surface, upstream of fuel tank air vents. Vortex generators may alternatively be located elsewhere on the aircraft, including on the upper wing surface. The aircraft further comprises NACA ducts 314, which provide air flow inlets, and rain gutters 316 which extend over the top of each aircraft door to divert rain flow over the fuselage 302 from passengers using the doors. Such aircraft components have in common the fact that they perform a useful function in a particular flight phase or on the ground, while providing a drag penalty in other flight phases. The invention can be embodied in improvements to such aircraft components to address this issue, as discussed further below.
When the aircraft is at cruise (
Referring to
By actuating the blade seal 400 away from the second surface 402 whilst the second surface 402 is in motion, the possibility of the seal 400 becoming entrapped between the first and second surfaces is minimised. This is beneficial since an entrapped seal has an adverse effect upon the aerodynamic profile of the component and also would require a maintenance stop to be scheduled to rectify the problem. After the second surface 402 has been fully actuated into its high lift configuration, the seal may be further actuated to achieve an aerodynamically favourable profile in the high lift configuration (not shown). This may be achieved by creating a lip, a convergent gap, a divergent gap or any other favourable configuration that is capable of influencing flow separation and laminar flow.
The blade seal 400 of
In the embodiment of
The actuatable trailing edge device 504 comprises an actuating material such as an electro-active polymer. In response to deployment of the flap 500, the actuatable trailing edge device 504 is actuated by application of a first activation input signal to induce a first deformation to a first configuration (shape) which provides a downward curvature as shown in
It is also desirable that the actuatable trailing edge device 504 have a high stiffness when it is in the first configuration, to ensure that the convergent gap is maintained within acceptable tolerances. Thus, the first activation input signal induces an increase in stiffness of the activating material, in tandem with the shape change.
During non high-lift flight phases, especially cruise, it is desirable for the actuatable trailing edge device 504 to adopt a shape in which it provides the aerodynamic profile with the lowest drag penalty. Thus, in response to retraction of the flap 500, the actuatable trailing edge device 504 is actuated to induce a second deformation to a second shape which provides such an aerodynamic shape, as illustrated in
As described above in relation to other embodiments, the actual deformation achieved by application of the first and/or second input activation signals will typically not exactly correspond to the desired deformation. The actual deformation achieved is determined by analysis of an output signal generated by the actuating material in response to the actual deformation, via the methods described above (and in particular as illustrated in
In some embodiments the actuatable trailing edge device will be monolithically formed from the actuating material, and in others it will comprise one or more portions of actuating material. For example, as illustrated in
Although in the embodiment of
In yet further embodiments the actuatable trailing edge portion 504 may alternatively extend from the aft edge of another movable control surface, such as a trailing edge flap. In such embodiments the actuatable trailing edge portion provides a flap tab that extends along the aft edge of the flap and is actuatable relative to the flap. The tab can be fixed at one end to the flap so that actuation of the actuating material of the tab causes its free end to move relative to its fixed end. In this way, the tab can be actuated so that its free end moves downwardly relative to the flap, to increase the curvature of the flap and therefore increase lift in the deployed high lift configuration of the flap.
Although the invention has been described above with reference to one or more preferred embodiments, it will be appreciated that various changes or modifications may be made without departing from the scope of the invention as defined in the appended claims.
In particular, the embodiments described above utilise electro-active polymers to achieve the required displacement of the aircraft component, but in other embodiments other suitable materials, such as materials having piezoelectric properties, may be used instead.
In all of the embodiments described herein the aircraft component may be formed monolithically from the actuating material, such as an electro-active polymer, or may comprise a composite component in which one or more portions comprise the actuating material.
Alternatively, the aircraft component may comprise a fibre-reinforced composite material in which reinforcing fibres are embedded in a matrix. In such embodiments the actuating material may be dispersed as particles throughout the matrix, providing the benefit of reducing additional fabrication steps and reducing delamination/debonding risks. The actuating material/matrix fraction would need to be tailored to ensure the actuating material content is sufficiently high to deliver the required mechanical force without compromising the load carrying ability of the composite matrix, or increasing its mass/dimensions beyond acceptable limits. This is considered most likely to be an attractive option for structures that take advantage of the expansion of the actuating material to drive the deformation.
As another alternative for embodiments in which the aircraft component comprises a fibre-reinforced composite material, the actuating material may be incorporated in filament form into the fibre weave of the composite material. By varying the location within the matrix (e.g. above or below the neutral axis, or parallel to the ±45° weave) it is possible to design a composite structure that has the ability to be deformed in both the x, y, and a axes (i.e. in plane and out of plane) by inducing strain in the appropriate plane. Careful choice of materials is necessary in order to limit the charge dissipation of the actuating material filaments into the composite matrix when the resistivity of the matrix is such that it ‘bleeds’ charge away from the actuating material. This may be achieved by resistive coatings applied to the actuating material, analogous to those used in electric transformer and motor windings to prevent short circuits. Moreover, tailoring the bulk of the actuating material fibres in the desired direction enables the material to apply a greater force in the desired plane, or exhibit varying degrees of deflection capability; it also enables the tailoring of the stiffness/strength of the composite structure by altering the fibre/actuating material ratio in the desired plane.
Number | Date | Country | Kind |
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1618393.1 | Nov 2016 | GB | national |
Filing Document | Filing Date | Country | Kind |
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PCT/EP2017/077853 | 10/31/2017 | WO | 00 |