Vortex generator control system, aircraft, and method for controlling an airflow changing device

Abstract

The present application provides a vortex generator control system (814) for controlling a vortex generator structure (115) of an aircraft, which includes: a controller (825) configured to receive one or more deployment or retraction command signals from a flight control unit (117), and further configured to send one or more command signals to a fluid control valve (819); a fluid pressure sensor (817) configured to sense one or more pressure values from actuators (511, 711, 803) of the vortex generator structure (115) and send the pressure values to the controller (825), and the fluid control valve (819) is configured to control the fluid transfer between the actuators (511, 711, 803) and a reservoir (823) in response to the command signal from the controller (825). The present application also relates to an aircraft and a method of using the vortex generator control system to control the actuation of an air flow changing device of the vortex generator structure from a first position to a second position.

Description

Technical Field

[0001] The present technology relates to a vortex generator structure for an aircraft and a vortex generator control system. Background Art

[0002] The vertical fin of an aircraft and a vertical rudder mounted to the vertical fin provide a method of controlling a yaw moment for causing the aircraft to rotate about its vertical axis or yaw axis. Thus, the vertical fin and the vertical rudder provide directional stability and directional control. The maximum yaw moment that the rudder can generate depends on the size of the rudder and the degree to which the rudder can be deflected about the rudder axis. In other words, a larger rudder or a rudder that can be deflected through a larger angle can be used to generate a larger yaw moment. Since the rudder can only be deflected through a limited angle to avoid stall flow at the rudder, the maximum yaw moment is mainly determined by the overall shape and surface area of the vertical fin and the rudder.

[0003] Under normal operating conditions, the maximum yaw moment is never required. Therefore, most of the yaw moment that the vertical rudder can generate remains unused, which results in unnecessary aerodynamic drag, and as the drag increases, the fuel consumption increases, thereby reducing the maximum range of the aircraft. Therefore, any unnecessary drag should be avoided.

[0004] In view of the above, it can be considered that the object of the present technology is to improve the flow generated by the vertical fin and the vertical rudder without having to increase the size of the vertical fin. Another object of the present technology is to provide the following vortex generator structure: the vortex generator structure is less likely to fail or has a lower maintenance cost due to having a small number of moving parts in its design. Summary of the Invention

[0005] Embodiments of the present technology provide a vortex generator control system for controlling a vortex generator structure of an aircraft, the vortex generator control system including: a controller configured to receive one or more deployment or retraction command signals from a flight control unit and further configured to send one or more command signals to a fluid control valve; a fluid pressure sensor configured to sense one or more pressure values of an actuator of the vortex generator structure and send the pressure values to the controller, wherein the fluid control valve is configured to control the fluid transfer between the actuator and a reservoir in response to a command signal from the controller.

[0006] The controller can also be configured to send command signals to a fluid pump, which is configured to drive fluid from a reservoir to an actuator in response to the command signals from the controller. The system can be configured to operate using hydraulic fluid or using pneumatic fluid. The vortex generator control system can include a plurality of pressure sensors or control valves. The control valves can be electromagnetically operated.

[0007] The vortex generator control system can be configured to control a linear fluid actuator of a vortex generator structure. In an alternative embodiment, the vortex generator control system can be configured to control a fluid muscle actuator of a vortex generator structure. The vortex generator control system can be configured to operate a fluid muscle actuator of a vortex generator structure within a pressure range of 0 Pa to +600,000 Pa relative to ambient pressure. The vortex generator control system can be configured to control the vortex generator structure in a deployed state corresponding to an actuator pressure of approximately +600,000 Pa relative to ambient pressure. The vortex generator control system can be configured to control a plurality of aircraft vortex generator structures. An aircraft including a vortex generator control system is also provided.

[0008] In another embodiment of the present technology, a method for controlling the actuation of an air flow changing device of a vortex generator structure from a first position to a second position using a vortex generator control system is provided. The method includes the following steps: a flight control unit signals a controller to move the air flow changing device of the vortex generator structure from the first position to the second position; the controller signals a control valve to open, thereby allowing fluid to be transferred between a reservoir and an actuator of the vortex generator structure, such that the air flow changing device is actuated from the first position to the second position.

[0009] The first position can be a contracted state of the air flow changing device, and the second position can be a deployed state of the air flow changing device. Alternatively, the first position can be a deployed state of the air flow changing device, and the second position can be a contracted state of the air flow changing device.

[0010] The advantages of the present technology will now become apparent from the following detailed description with appropriate reference to the accompanying drawings.

[0011] Description of the Drawings

[0012] Embodiments of the present technology will now be described by way of example only with reference to the following drawings, in which:

[0013] Figure 1 is a side view of an aircraft including a plurality of vortex generator structures and a system for operating the vortex generator structures;

[0014] Figure 2 is Figure 1A close-up view of the vertical tail of the aircraft, which shows in more detail the position of the vortex generator structure;

[0015] Figure 3A is Figure 1 and Figure 2 An isometric view of one of the vortex generator structures in the vortex generator structure of the aircraft airflow changing device in the extended first state;

[0016] Figure 3B is Figure 3A An isometric view of the aircraft airflow changing device in the contracted second state;

[0017] Figure 4A is Figure 1 and Figure 2 An isometric view of one of the vortex generator structures in the vortex generator structure, the vortex generator structure including the aircraft airflow changing device in the extended first state;

[0018] Figure 4B is Figure 3B An isometric view of the vortex generator structure in the contracted second state;

[0019] Figure 5A Provides a side cross-sectional view of the vortex generator structure 115 along the Figure 4A YZ plane indicated in;

[0020] Figure 5B Provides a side cross-sectional view of the vortex generator structure 115 along the Figure 4B YZ plane indicated in;

[0021] Figure 6A is Figure 4A The front cross-sectional view of the vortex generator structure of through the Figure 4A plane XZ given in;

[0022] Figure 6B is Figure 4B The front cross-sectional view of the vortex generator structure of through the Figure 4B plane XZ given in;

[0023] Figure 7A is according to another embodiment of the present technology Figure 4A The front cross-sectional view of the vortex generator structure of;

[0024] Figure 7B is according to another embodiment of the present technology Figure 4B The front cross-sectional view of the vortex generator structure of.

[0025] Figure 8A and Figure 8Bis a front - view sectional view of a vortex generator structure 115 according to another embodiment of the present technology.

[0026] Figure 9A and Figure 9B respectively provide the deployment operation of the vortex generator control system 814 and the retraction operation of the vortex generator control system 814 according to an embodiment of the present technology.

[0027] Figure 10A and Figure 10B is a front - view sectional view of a vortex generator structure 115 according to an embodiment of the present technology. Detailed Embodiment

[0028] Referring to Figure 1 , the aircraft 101 includes a pair of lift - surface assemblies in the form of wings 103 joined to the fuselage 105. Each wing 103 includes hinged control surfaces such as ailerons and spoilers (not shown), and high - lift devices such as flaps and slats (not shown). Each wing 103 carries an engine 107 mounted to an engine pylon 108 using the wing. The aircraft 101 includes additional lift - surface assemblies in the form of a horizontal tail 109 and a vertical tail 111, where the horizontal tail 109 and the vertical tail 111 are each attached at the rear of the fuselage 105. The vertical tail 111 is provided with a hinged control surface in the form of a rudder 113. The aircraft 101 includes other surface assemblies such as a belly fairing 112 that smoothly connects the wing 103 to the fuselage 105, and high - lift device fairings 114 that smoothly connect the respective parts of the flaps (not shown) to the corresponding wings 103.

[0029] In this embodiment, the aircraft 101 further includes a plurality of vortex generator structures 115 mounted in the vertical tail 111. The deflection of each control surface and the deflection of each vortex generator structure 115 are controlled by at least one flight control unit 117. The flight control unit 117 receives input from at least one sensor unit 119 and commands the deflection according to a predetermined flight envelope stored on a medium (not shown) connected to the flight control unit 117.

[0030] Referring to Figure 2 , in this embodiment, each vortex generator structure 115 is separated from an adjacent vortex generator structure 115 along the span direction of the vertical tail 111 by a set distance. Each vortex generator structure 115 is mounted in the main body 203 of the vertical tail 111 and has a chordwise position set in front of the leading edge 201 of the rudder 113.

[0031] Referring to Figure 3A, in the present embodiment, an aircraft airflow changing device 301 is provided. The aircraft airflow changing device 301 includes an elastically deformable baffle member 303, and the elastically deformable baffle member 303 is laminated to an elastically deformable base member 305 to form a single integrally formed component, that is, the component does not have a hinge or a separate connector element. In the first state, the elastically deformable baffle member 303 is substantially flat and rectangular in shape, and extends in the YZ plane in a manner substantially perpendicular to the elastically deformable base member 305 that is also in the corresponding first state, such that the outermost edge 307 of the elastically deformable baffle member 303 substantially coincides with the YZ plane.

[0032] The elastically deformable baffle member 303 is formed of 8 layers of co-cured GFRP laminate, and the co-cured GFRP laminate tapers down from 8 layers to 4 layers at the edge 307 of the elastically deformable baffle member 303. Each layer can use an epoxy resin / E-glass composite material.

[0033] The elastically deformable baffle member 303 can have a dimension Hf of at least 10 mm in the Z direction, but preferably, it has a dimension Hf of 20 mm in the Z direction. The elastically deformable baffle member 303 can have a dimension Wf of at least 30 mm in the Y direction, but preferably, it has a dimension Wf of 60 mm in the Z direction. The elastically deformable baffle member 303 can have a dimension Tf of at least 0.5 mm in the X direction, but preferably, it has a dimension Tf of 1 mm in the X direction.

[0034] The elastically deformable base member 305 is formed of a rectangular GFRP laminate profile, such that the elastically deformable baffle member 303 has a relatively low lateral stiffness compared to the elastically deformable base member 305.

[0035] The elastically deformable base member 305 can have a dimension Hb of at least 2 mm in the Z direction, but preferably it has a dimension Hb of 5 mm in the Z direction.

[0036] The elastically deformable base member 305 has a dimension Wb of at least 60 mm in the Y direction (i.e., the same Y dimension as the elastically deformable baffle member 303), but preferably it has a dimension Wb of 80 mm in the Y direction.

[0037] The elastically deformable base member 305 preferably has a dimension Tb of 20 mm in the X direction.

[0038] Alternatively, the elastically deformable baffle member 303 or the elastically deformable base member 305 can be formed of any other suitable material and also by any suitable method, such materials as non-metallic materials, metallic materials or composite materials, and such methods including but not limited to laying and curing the material, manufacturing by adding layers, or machining, shaping, casting or molding the material. Additionally, the elastically deformable baffle member 303 and the elastically deformable base member 305 can be integrally formed.

[0039] In addition, the elastically deformable baffle member 303 can be not rectangular in shape, but can have a swept leading edge or trailing edge. Alternatively, the elastically deformable baffle member 303 can be triangular. Alternatively, the elastically deformable baffle member 303 can include curved edges. Alternatively, the elastically deformable baffle member 303 can be semi-circular or semi-elliptical.

[0040] shown in the Figure 3A first state of the aircraft airflow changing device 301 is shown in Figure 3B the second state. In response to a positive bending moment Mx about the X axis, the elastically deformable base member 305 undergoes uniaxial elastic bending to the second state, thereby causing an asymmetric elastic buckling of the elastically deformable baffle member 303 to the corresponding second state. This is caused by a lateral torsional buckling of the elastically deformable baffle member 303 having a lower total lateral stiffness compared to the total lateral stiffness of the elastically deformable base member 305. When the elastically deformable baffle member 303 is in the second state, its outermost side edge 307 forms an angle A that is substantially perpendicular to the YZ plane.

[0041] Removing the bending moment Mx from the elastically deformable base member 305 (i.e., causing the elastically deformable base member 305 to return from the second state to the first state) causes the elastically deformable baffle member 303 and the elastically deformable base member 305 to elastically deform back to their respective non-buckled first states, where the outermost side edge 307 of the elastically deformable baffle member 303 returns to a position substantially coincident with the YZ plane, and the elastically deformable base member 305 is substantially rectangular in shape.

[0042] In the present embodiment, the elastically deformable baffle member 303 is in an extended configuration when in the first state and non-buckled, and in a contracted configuration when in the second state and buckled. Correspondingly, the elastically deformable base member 305 is non-buckled in the first state and buckled in the corresponding second state. However, it should be understood that those skilled in the art will understand that different variations can be achieved. For example, the aircraft airflow changing device can be constructed and manufactured such that the elastically deformable baffle member 303 is extended and non-buckled in a first state corresponding to the first buckled state of the elastically deformable base member 305.

[0043] Referring to Figure 4A , in the present embodiment, each vortex generator structure 115 includes a housing 401 that is generally substantially rectangular in shape; an outer cover 403 having a portion 405 that is substantially coplanar with the outermost surface (commonly also referred to as the outer mold line or OML) of the vertical fin 111 (not shown). It should be understood that the exact shape of the housing 401 can be determined by the available space in which it is to be installed, and thus the housing 401 may not be rectangular in shape. The housing 401 includes a foremost wall 407, a rearmost wall 409, a pair of side walls 411, 413, and an innermost wall 414 (not shown). The foremost wall 407 and the rearmost wall 409 are each arranged substantially parallel to the XZ plane, the pair of side walls 411, 413 are each arranged substantially parallel to the YZ plane, and the innermost wall 414 is arranged substantially parallel to the XY plane and offset from the outer cover 403. The housing 401 is a monolithic component formed of CFRP material. However, the housing 401 can be an assembly of two or more separate sub-components formed of the same material, the sub-components being attached to each other by any suitable means, such as by replaceable fasteners. Such a structure may be preferred for subsequent assembly and disassembly of the vortex generator structure 115. The housing 401 can also be formed of any other suitable non-metallic material, metallic material, or composite material and using any suitable manufacturing method, such as additive layer manufacturing, machining, forming, casting, molding, or laying and curing.

[0044] The outer peripheral edge 417 of the outer cover 403 is configured to extend substantially perpendicular to the walls 407, 409, 411, 413 of the housing 401 so as to define an overhang 419 that is configured to be attached to the corresponding outer peripheral edge 501 (not shown) provided by the vertical fin 111.

[0045] The overhang portion 419 extends generally along the entire outer periphery 417, i.e., along the front wall 407, the rear wall 409, and the two side walls 411, 413 of the vortex generator structure 115. However, this overlapping structure may be discontinuous around the outer periphery of the top cover 403. In the present embodiment, countersunk quick-release fasteners (not shown) are used in combination with corresponding nut plates (not shown) attached to portions of the surrounding structure to assemble the vortex generator structure 115 to the vertical fin 111. This provides an in-service replaceable system, which is important for in-service operations, where a failed unit can be quickly replaced to prevent unnecessary delays during normal operation. Alternatively, the vortex generator structure 115 can form an integral part of the vertical fin 111, i.e., a part of the cover of the vertical fin 111.

[0046] Portion 405 also defines an opening 421 in the outer cover 403 of the housing 401 leading to a corresponding chamber 423 within the housing 401, where the chamber 423 is also defined by the inner surfaces of the walls 407, 409, 411, and 413. The inner surfaces may be offset as shown, or may be very close to the edge of the elastically deformable baffle member 303. The edge of the opening 421 may include a large radius or chamfer configured to improve the airflow through the chamber 423.

[0047] An aircraft airflow altering device 301 is disposed within the chamber 423 and is supported by the foremost wall 407 and the rearmost wall 409 of the housing 401 and includes an elastically deformable baffle member 303 that extends through the opening 421 in the Figure 4A first state shown.

[0048] When the elastically deformable baffle member 303 extends in the first state during operation of the aircraft 101, the portion of the airflow 431 adjacent to the outer cover 403 is altered by the elastically deformable baffle member 303 to create one or more airflow vortices 433 that flow rearward from the location of the vortex generator structure 115 and over the vertical fin 111 and the rudder 113.

[0049] Referring to Figure 4B Figure 4A the vortex generator structure 115 of Figure 3B is shown in a second configuration in which the elastically deformable baffle member 303 of the aircraft airflow altering device 301 is retracted from the opening 421 when it is in its second buckled state and the elastically deformable base member is in a corresponding second buckled state, i.e., corresponding to the state shown in

[0050] A known effect of the vortex 433 near a lift surface subjected to flow separation is that even though such a vortex 433 also generates a relevant amount of form drag and vortex drag, it still delays the onset of flow separation from the lift surface. Despite this drag penalty, the vortex 433 is still desirable in certain situations. For example, the vertical fin 111 or the rudder 113 is sized considering parameters such as surface area, airfoil section, etc., such that when the vertical fin 111 and the rudder 113 are operated at the maximum angle of incidence with respect to the direction of the free stream airflow, they can provide sufficient yaw moment to control the aircraft 101. The sizing of the vertical fin 111 and the rudder 113 for these situations can account for a certain amount of lift loss due to possible flow separation. Thus, in the context of the present embodiment of the vertical fin 111 or the rudder 113, when the vortex 433 is introduced upstream (in front) of the region of the vertical fin 111 and the rudder 113, the flow separation losses can be reduced or completely eliminated, which would otherwise occur at a large angle of incidence. The vortex generator structure 115 of the present technology is a means to achieve this, and the vortex generator structure 115 in turn allows for a reduction in the overall size of the vertical fin 111 and the rudder 113 and thus a reduction in the total weight, which is advantageous in terms of the manufacturing and operating costs of the aircraft 101.

[0051] Referring Figure 4B , the contraction of the elastically deformable baffle member 303 restores the airflow near the vortex generator structure 115 to its nominal characteristics and avoids the form drag and vortex drag losses that would otherwise occur. This configuration is desirable, for example, when the vertical fin 111 or the rudder 113 does not need to be operated at a high angle of incidence with respect to the free stream airflow direction during cruise. In the present embodiment, the opening 423 can also be provided with a thin frangible membrane that is configured to extend over the opening 423 in the plane of the portion 405 and is fastened or bonded to the outer covering 403. The frangible membrane will provide aerodynamic coverage for the chamber 423 but is also configured to permanently rupture due to the elastically deformable baffle member 303 contacting the frangible membrane during the initial stage of its first extension, thus not inhibiting the full extension or performance of the elastically deformable baffle member 303.

[0052] However, it should be understood that preferably, when the elastically deformable baffle member 303 of the airflow changing device 301 is in the second buckled state and the elastically deformable base member is in the corresponding second buckled state, the elastically deformable baffle member 303 of the airflow changing device 301 does not fully contract into the chamber but is located in a position substantially parallel to the plane of the portion 405. In this configuration, the elastically deformable baffle member 303 will remain in the airflow in the second state of contraction but will not generate any significant vortices 433 and associated drag. However, it can generate a certain degree of acceptable form drag.

[0053] Another advantage of using the vortex generator structure 115 including such an aircraft airflow altering device 301 is that the aircraft airflow altering device 301 has no moving parts, simplifying its design. Thus, it can also have a lower probability of failure due to wear or due to jamming in the presence of contaminants such as ice, oil, or dust. This can result in a smaller number of vortex generator structures 115 being required and thus lower costs or reduced complexity due to longer maintenance inspection intervals.

[0054] It should also be understood that one or more vortex generator structures 115 can be arranged at other locations on the vertical fin 111 - for example, at the foremost "front" edge portion of the vertical fin 111 or the rudder 113 - in order to enhance aerodynamic performance. It should also be understood that the one or more vortex generator structures 115 can be arranged at other locations on the aircraft 101 proximate to any other lift surface where flow separation may desirably be inhibited. For example, the one or more vortex generator structures 115 can be mounted at a portion of the leading or trailing "rear" edge of the wing 103 or the horizontal stabilizer 109. Alternatively, the one or more vortex generator structures 115 can be mounted on another surface component such as a pylon 108 or an engine 107. Additionally, one or more vortex generator structures 115 can be used to improve the aircraft acoustic performance. For example, one or more vortex generator structures 115 can be positioned in front of (upstream) and adjacent to an opening 110 defined by a portion of the outer surface of the aircraft 101 in order to generate one or more vortices 433 that can reduce the audible Helm - Holtz induction frequency from such an opening 110.

[0055] Referring Figure 5A , the airflow altering device 301 is mounted within the chamber 423. The first end 502 and the second end 503 of the elastically deformable base member 305 are each positioned within corresponding slots 509 and 507 defined respectively by the rear wall 409 and the front wall 407 such that the airflow altering device 301 is supported at both ends. An actuator structure 510 including an electric linear actuator 511 and a coupler 513 are also mounted within the chamber 423. It should be understood that the actuator structure can include any suitable alternative actuator, such as a rotary type actuator or a piezoelectric type actuator. The actuator can alternatively be hydraulically actuated. The first end 515 of the linear actuator 511 has a lug and is pivotally mounted to a corresponding set of lugs provided by the innermost wall 414 of the chamber 423 using a U - shaped clip pin 519.

[0056] The second end 521 of the linear actuator 511 includes a lug pivotally connected via another U-bolt pin 523 to a set of lugs provided by the coupler 513. The coupler 525 is fixedly attached to a position at a substantially middle length in its Y-direction on the lower surface 527 of the elastically deformable base member 305.

[0057] Optionally, a universal ball joint type structure allowing for a greater degree of freedom can be alternatively used to replace either the pin structure for connecting the linear actuator 511 to the housing 401 or the coupler 525 as needed.

[0058] When the actuator structure 510 is in the first position as shown in Figure 5A , the elastically deformable base member 305 is in a first state of being non-buckled, and the elastically deformable baffle member 303 is thus in a corresponding first state that is also non-buckled. In this state, the elastically deformable baffle member 303 extends through the opening 421 and acts as a vortex generator in the presence of an air flow passing through the vertical fin 111.

[0059] Referring to Figure 5B , shows all the features of Figure 5A , however, in Figure 5B , the actuator structure 510 is actuated to a second position, thereby providing a force generally in the direction of the vertical axis of the elastically deformable base member 305, such that the elastically deformable base member 305 deforms to a second state of being buckled and the elastically deformable baffle member 303 thus also deforms to a corresponding second state of being buckled. In this state, the elastically deformable baffle member 303 contracts from the opening 421, such that the elastically deformable baffle member 303 extends substantially in the plane of the portion 405 (i.e., in the plane of the OML), so that the air flow near the opening 421 is not substantially changed. By removing the vertical force applied by the actuator structure 510, for example, by de-energizing the linear actuator 511, the elastic energy stored within the aircraft air flow changing device 301 is sufficient to reverse drive the linear actuator 511 to return the aircraft air flow changing device 301 from the buckled second state to the non-buckled first state. It is advantageous that no power or command is required to be supplied to the linear actuator 511 to extend the elastically deformable baffle member 303, as the design is inherently fail-safe and ensures that the maximum yaw moment capabilities of the vertical fin 111 and the rudder 113 can still be obtained as needed in the absence of power, although at the expense of a minimal drag penalty.

[0060] It should be understood that the aircraft airflow changing device 301 can be actuated and deformed between the first state and the second state alternatively by the deformation of the surrounding structure acting on the elastically deformable base member 305 or by the pressure difference acting on the elastically deformable baffle member 303 due to the airflow outside the chamber, rather than by the actuation of the specific type of power actuator structure described so far.

[0061] Referring Figure 6A and Figure 6B , a gap 601 is provided between the side wall 413 of the housing 401 and the airflow changing device 301. Such a gap 601 may be preferred for providing a path for installing or replacing the actuator structure 510 or other internal components within the housing 401 or for performing maintenance or repair inspections. It should be understood that such a gap 601 may not be required, but rather the side wall 413 of the housing 401 may be configured such that when the elastically deformable base member 305 is in the first state and the second state, the outermost edge 309 of the elastically deformable base member 305 substantially abuts the inner surface of the side wall 413, and the side wall 413 of the housing 401 may also be configured to substantially abut the outermost edge 307 of the elastically deformable baffle member 303 when the elastically deformable baffle member 303 is in the buckled second state. This may preferably be done to prevent liquids and contaminants from collecting within the chamber 423.

[0062] It is also shown that the coupler 513 extends substantially across the elastically deformable base member 305 in the X direction. This ensures that the load vertically applied by the actuator structure 510 in the Z direction is evenly distributed across the X - dimension span of the elastically deformable base member 305 via the coupler 513. Attaching the linear actuator 511 to the coupler 513 at the second end using a pivoting U - shaped clip pin 523 may allow for easier installation and replacement of the linear actuator 511 when needed. The use of pivoting connectors 519 and 523 in the actuator structure 510 ensures that no bending load is applied to the elastically deformable base member 305 when a vertical load is applied, which bending load may cause unexpected deformation of the aircraft airflow changing device 303.

[0063] Referring Figure 7A , an alternative embodiment is shown. The vortex generator structure 115 of this embodiment is substantially the same as Figure 4A the vortex generator structure 115 shown in

[0064] An air flow changing device 301 is installed within a chamber 423. The elastically deformable base member 305 of the device is formed such that in its first non-buckled state it is generally slightly curved in the positive Z direction, represented in the figure by dashed line 701 and upper and lower surfaces 702. The first end 709 of the elastically deformable base member 305 is of a circular shape and is positioned within a corresponding first slot 509 which also has a corresponding circular shape such that the first end 709 generally conforms to the inner surface of the first slot 509.

[0065] The second end 707 of the elastically deformable base member 305 is positioned within a corresponding second slot 507 of a similar corresponding shape. An actuator structure 710 is provided within the second slot 507. The actuator structure includes a linear actuator 711. The linear actuator 711 is fixed to the foremost wall 407 of the housing 401 at its first end 713.

[0066] The second end 715 of the linear actuator 711 is pivotally attached to the second end 707 of the elastically deformable base member 305 by a U-shaped clip pin 717. The actuator structure 710 enables the elastically deformable base member 305 to rotate relative to the second end 715 of the linear actuator 711 and to rotate generally parallel to the YZ plane in a sense.

[0067] When the actuator structure 710 is in the first position as shown in Figure 7A , the elastically deformable base member 305 is in a first state which is a non-buckled state, and the elastically deformable baffle member 303 is thus in a corresponding first state which is also non-buckled. In this state, the elastically deformable baffle member 303 extends through the opening 421 and serves as a vortex generator in the presence of an air flow passing over the vertical fin 111.

[0068] Referring to Figure 7B , all features of Figure 7A are shown. However, in Figure 7BIn [the situation], the actuator structure 710 is actuated to the second position, thereby providing a force generally along the transverse axis (Y direction) of the elastically deformable base member 305, such that the elastically deformable base member 305 deforms to a second state that is a buckled state and the elastically deformable baffle member 303 thus also deforms to a corresponding second state that is a buckled state. In this state, the elastically deformable baffle member 303 contracts from the opening 421, such that the elastically deformable baffle member 303 extends substantially in the plane of the portion 405 (i.e., in the plane of the OML), so that the airflow near the opening 421 is not substantially changed. By, for example, de-energizing the linear actuator 711 to remove the transverse force applied by the actuator structure 710, the elastic energy stored within the aircraft airflow changing device 301 is sufficient to reversely drive the linear actuator 711 to return the aircraft airflow changing device 301 from the buckled second state to the unbuckled first state.

[0069] When compared with Figure 5A 、 Figure 5B 、 Figure 6A and Figure 6B 's embodiments, it can be seen that the housing 401 in the present embodiment does not need to have such a large chamber 423 because the actuator structure 710 requires a much shorter stroke to achieve the same deformation amount of the aircraft airflow changing device 301. For such a structure, a piezoelectric type actuator can be preferred because their power density is relatively high for a given size. Therefore, the inner wall 414 of the housing 401 is closer to the outer wall 403 and thus shallower, making the vortex generator structure 115 of the present embodiment much more compact. Such a structure can be preferred when the available internal space of the vertical fin 111 is limited, and can be particularly required for the outer trailing edge region of the vertical fin 111 or other lift surfaces.

[0070] Referring to Figure 8A , an embodiment of the vortex generator structure 115 is shown. An airflow changing device 301 shown in the first position is installed within the chamber 423. The elastically deformable base member 305 of the device 301 is again formed such that in its unbuckled first state, it has a shape that is generally slightly curved in the positive Z direction, represented in the figure by the dashed line 701 and the upper surface 702 and the lower surface 702. The first end 709 of the elastically deformable base member 305 is positioned within a corresponding first slot 509 of a similar corresponding shape. The second end 707 of the elastically deformable base member 305 is positioned within a corresponding second slot 507 of a similar corresponding shape. Each end 707, 709 can have a shape corresponding to the housing 401, for example, each end can be circular to allow the airflow changing device 301 to rotate more easily.

[0071] A fluid muscle actuator structure 801 is provided, which includes a fluid muscle actuator 803 pivotally connected to a pair of connectors 805; there is a connector 805 at each end of the actuator 803. Each connector 805 extends in a generally vertical direction from the lower part of the airflow changing device 301 at the first end 709 and the second end 707 adjacent to the slots 509 and 507. Alternatively, the connector can extend from the base member 305 at an acute angle. However, this may increase the width of the chamber 423 and thus increase the overall size of the vortex generator structure 115. The connector has a substantially L-shaped body, which is formed of CFRP and bonded to the lower surface 702 of the base member 305 at the upper surface. Alternatively, any other suitable material, such as CRESS, can be selected. In addition, it may be preferable to integrally form one or more connectors from the same material as the airflow changing device 301.

[0072] The opposite ends of each connector form a U-shaped clamp 811, and the U-shaped clamp 811 is configured to receive the corresponding lug 813 and pin 809. When the lug 813 formed by the body of the fluid muscle actuator 803 is received in the corresponding U-shaped clamp 811 and fixed at each connector 805 using the pin 809, a pivotable connection between the fluid muscle actuator 803 and the connector 805 is achieved. The connector 805 is used to support the fluid muscle actuator 803 within the chamber 423 and offset the fluid muscle actuator 803 by an offset length D relative to the base member 305 or the baffle member 303 of the airflow changing device 301. The offset length D is between 5 mm and 15 mm in length. Preferably, 10 mm is used.

[0073] As shown, it is advantageous to use the connector 805 that provides the offset of length D, because a moment arm with substantially the same length is introduced between the base member 305 and the load application point of the actuator 803, so that a higher positive bending moment or negative bending moment Mx about the X axis can be achieved for a fluid muscle actuator 803 of a given size. This may allow the use of a smaller fluid muscle actuator 803 in some cases, which reduces the cost, power consumption, and weight of the entire vortex generator structure 115. The pivotable connection of the connector 805 optimizes the direction of the load from the actuator 803 and ensures that the fluid muscle actuator 803 itself is not subjected to bending or torsional loads caused by its own actuation. However, it should be understood that alternatively, the fluid muscle actuator 803 can be directly attached to the elastically deformable base member 305 or the elastically deformable baffle member 303 without the need for the connector 805. In an alternative embodiment, the actuator 803 can also be non-pivotally connected to the connector 805. When the airflow changing device 301 is in the first position, another function of the connector 805 is to orient the actuator 803 substantially parallel to the line 701.

[0074] The fluid muscle actuator 803 is similar to commercially available artificial muscle type actuators such as those from Festo Vertrieb GmbH. The actuator 803 in this embodiment is configured to be pneumatically driven by a gas pressure source supplied within an operating pressure range of 0 Pascals to +600,000 Pascals relative to the ambient atmospheric pressure value. In this embodiment, considering the variations in air pressure and temperature according to the altitude and weather conditions of the aircraft, the ambient atmospheric pressure value can be within the range of values determined according to the operating envelope of the aircraft. The pressure source can be provided by an independent pump, a pressurized or non-pressurized reservoir, or any combination of both, such as an aircraft bleed air system or a central hydraulic system. In the case of powering the aircraft vortex generator structure 115, a pneumatic type fluid muscle actuator 803 is desirable as pneumatic power systems tend to be lighter in weight. However, depending on the required operating characteristics, the fluid muscle actuator 803 can alternatively be powered by any other suitable fluid or hydraulically driven. A hydraulically driven fluid muscle actuator 803 may be preferred as hydraulic systems have a higher available power density and less variation in power characteristics due to ambient air temperature.

[0075] The fluid muscle actuator 803 includes an elongated body provided with a closed fluid bladder 804 having a substantially constant cylindrical cross-section along its length. At each end, the bladder 804 forms a lug 813 which, as previously described, can be attached to a U-bolt 811 at each end. The bladder 804 is formed of an elastic impermeable material reinforced with a woven material, such as a rubber bladder 804 reinforced by woven glass material. Those skilled in the art will understand that the configuration, material selection, and tightness of the reinforcing weave within the bladder 804 can be selected to determine the specific elastic strain energy, deflection characteristics, and / or force requirements of the bladder 804.

[0076] The fluid muscle actuator 803 is connected via a fluid conduit 815 to a vortex generator control system 814 which includes at least one fluid pressure sensor 817, at least one solenoid-operated control valve 819, at least one pump 821, at least one reservoir 823, at least one controller 825 including a combined signal processor and memory unit 827, and a flight control unit 117, all of which components are positioned and connected downstream of the fluid muscle actuator 803. The advantage of the vortex generator control system 814 as described above is that it is relatively simple and formed of reliable components. Additional pressure sensors or other components may be installed in the control system 814 in order to achieve a higher level of reliability and thus redundancy of the control system 814. For example, an additional control valve 819 may be installed in parallel with the first pressure control valve 819 to ensure that the failure of either valve 819 does not cause the system 814 to become inoperable. It should be understood that such a vortex generator control system 814 may be applied to the vortex generator structure 115 of any of the foregoing embodiments, which vortex generator structure 115 may be provided with a hydraulically driven actuator, such as a linear hydraulic actuator and may not necessarily require the fluid muscle actuator 803.

[0077] When the fluid muscle actuator structure 801 is in the first position as shown in Figure 8A the elastically deformable base member 305 is in a first state which is a non-buckled state, and thus the elastically deformable baffle member 303 is in a corresponding first state which is also non-buckled. In this state, the elastically deformable baffle member 303 extends through the opening 421 and serves as a vortex generator in the presence of an air flow passing over the vertical fin 111. Further, in this current state, the fluid muscle actuator 803 is without a source of energy, that is, the bladder 804 of the actuator 803 has no elastic strain, and the fluid pressure within the actuator 803 is at or below a threshold P1 which is 0 bar relative to ambient atmospheric pressure in the present embodiment. At the pressure P1, that is, in the state without a source of energy, the actuator 803 is designed to have a length L1 which is substantially equal to 60 mm, but depending on the dimensions of the air flow changing device 301, the length L1 may be less than 60 mm or up to 150 mm.

[0078] If the fluid pressure in the bladder 804 increases to be higher than the threshold, the deformation (and thus the elastic strain) of the actuator 803 increases in a manner substantially proportional to the increase in pressure. In this case, the outer diameter of the cylindrical bladder 804 begins to increase as the strain energy increases, and the elastic strain energy stored by the bladder 804 increases, and overall, the length of the actuator 803 decreases. At the peak supply pressure P2 - which is about 600,000 Pascals higher than the threshold in this embodiment - the actuator 803 is designed to have a length L2 equal to about 45 mm, and the length L2 is about 25% shorter than L1. In this state, it can be said that the fluid muscle actuator 803 is shortened, and the air flow changing device 301 will be in the second buckled state, and the air flow changing device 301 contracts from the opening 421, as will be described in further detail later with the support of Figure 8B as follows.

[0079] When the actuator 803 loses energy from the peak pressure P2 to P1, there is a situation contrary to this behavior. This is advantageous in the context of the aircraft vortex generator structure 115 because the entire system has an enhanced fail-safe design. If an accidental leak in the bladder 804 of the actuator 803 or the vortex generator control system 814 causes the pressure to continuously drop, the fluid muscle actuator 803 returns to its original length L1, thereby causing the deployment state (generating vortices) of the vortex generator structure 115.

[0080] It should be understood that one or more vortex generator structures 115 can be connected to a single vortex generator control system 814. This is advantageous because in the case of the failure just described, since the pressure of the entire system will be affected, the risk of asymmetric deployment of multiple vortex generator structures 115 will be avoided, which might otherwise introduce undesirable aircraft handling qualities. It should also be understood that multiple vortex generator control systems 814 can be connected to multiple vortex generator structures 115 to obtain a higher level of redundancy.

[0081] At least a portion of the fluid conduit 815 within the chamber 423 is elastically deformable such that it will not be damaged due to the movement of the fluid muscle actuator 803 within the chamber 423. The conduit 815 passes through a hole formed by the housing 401. The fluid muscle actuator controller 825 controls the actuation of the fluid muscle actuator 803 (and causes the actuation of the aircraft air flow changing device 301 from the buckled state to the unbuckled state). The controller 825 includes a combined signal processor and memory unit 827, which is configured to (i) receive a deployment or contraction command signal from the flight control unit 117, (ii) receive a pressure reading signal from the pressure sensor 817, and (iii) receive a status signal from the valve 819 (open / closed) and the pump (supply energy / lose energy) 821.

[0082] The controller 825 is also configured to (i) determine the deployed or contracted state of the fluid muscle actuator 803, (ii) send such state to the flight control unit 117, (iii) send commands to the control valve 819 to open or close it, and (iv) send commands to the pump 821 to start it and pump fluid from the reservoir 823 into the actuator 803, or to stop the pump 821 and allow the fluid to return from the actuator 803 to the reservoir 823.

[0083] Referring to Figure 8B which shows all the features of Figure 8A However, in Figure 8B , the fluid muscle actuator device 801 is actuated to a second position such that a force is provided substantially along the transverse axis (Y direction) of the elastically deformable base member 305 which is deformed to the second state. The elastically deformable base member 305 in this second state is in a buckled state, and the elastically deformable baffle member 303 is thus also deformed into a corresponding second state which is the buckled state. As shown in this second state, the elastically deformable baffle member 303 contracts from the opening 421 such that the elastically deformable baffle member 303 extends substantially in the plane of the part 405 (i.e., in the plane of the OML), so that the airflow near the opening 421 is not substantially changed.

[0084] Referring to Figure 9A which shows the contraction operation 901 for actuating the airflow changing device 301 from the first position shown in Figure 8A to the second position shown in Figure 8B . The operation 901 includes the following steps: 903 - the flight control unit 117 signals the controller 825 to move the airflow changing device 301, 905 - the controller 825 signals the valve 819 to open; signals the pump 821 to supply it with energy and once the signal is received, pump fluid through the conduit 815 into the bladder 804 of the actuator 803, causing the elastically deformable baffle member 303 of the airflow changing device 301 to move towards the contracted position; and 907 - when the controller 825 receives the peak P2 pressure value detected by the pressure sensor 817, the controller 825 signals the control valve 819 to close and signals the pump 821 to lose energy. As described above, once the peak pressure P2 is reached, the actuator 803 has a length L2, the elastically deformable baffle member 303 of the airflow changing device 301 also contracts from the opening 421, and the maximum strain energy is stored in the now deformed wall of the bladder 804 and the elastically deformable base member 305.

[0085] Referring to Figure 9B which shows the actuation for moving the airflow changing device 301 from the second position shown in Figure 8B toFigure 8A Deployment operation 909 at the first position in Figure 8A . Operation 909 includes the following steps: 911 - The flight control unit 117 signals the controller 825 to cause the airflow changing device 301 to contract; 913 - When receiving the signal, the controller 825 signals the valve 819 to open, causing the pressure in the bladder 804 to drop from the peak supply pressure P2, thereby causing the elastically deformable baffle member 303 of the airflow changing device 301 to move towards the deployed position; and 915 - When the controller 825 receives the threshold P1 pressure value detected by the pressure sensor 817, the controller 825 signals the control valve 819 to close. As previously described, once P1 is reached, the actuator 803 has a length L1, and the elastically deformable baffle member 303 of the airflow changing device 301 extends through the opening 421 and serves as a vortex generator in the presence of airflow flowing through the vertical fin 111.

[0086] In the present embodiment, the elastic strain energy stored in the wall of the bladder 804 of the actuator 803 and the elastically deformable base member 305 of the airflow changing device 301 is sufficient to deploy the airflow changing device 301 and force the fluid to flow back to the reservoir 823 through the conduit 815 without the need to energize the pump 819 in the opposite direction. However, this principle can be used when needed. Once the threshold pressure P1 is reached at the sensor 817, the controller 827 signals the valve 819 to close. The controller 825 can also send a signal indicating that the deployed state has been achieved to the flight control unit 117. The controller 827 can signal the valve 819 at a value slightly higher than the threshold pressure P1 to account for any processing or signal lag in the controller 825.

[0087] The vortex generator structure 115 is configured such that the elastic strain energy stored in the actuator 803 and the elastically deformable base member 305 works together to provide the elastic strain energy required to deploy the airflow changing device 301, where the pressure within the actuator 803 is released, either intentionally or as a result of the vortex generator structure 115 becoming inoperable due to accidental leakage of fluid from the system 814 or the actuator 803. Alternatively, the elastic strain energy stored in the fluid muscle actuator 803 can be configured to provide the energy required to deploy the airflow changing device 301.

[0088] Although Figure 8A and Figure 8BIn the embodiment, the fluid muscle actuator structure 801 is coupled to the airflow changing device 301 of the deformable type. However, the fluid muscle actuator structure 801 can also be applied to actuate the vortex generator structure 115 using the airflow changing device 901 of the rigid type. The airflow changing device 901 of the rigid type only includes a rigid baffle member 303, and the rigid baffle member 303 can be carried or hingedly installed within the chamber 423. By way of example, Figure 10A Such a structure substantially according to the previous embodiment is shown, but wherein the airflow changing device 301 is a rigid carbon fiber baffle member 301, which is hingedly installed and pivoted about an axis 1003 formed by a pair of concentrically aligned corrosion-resistant stainless steel shafts 1001. Each of the pair of concentrically aligned corrosion-resistant stainless steel shafts 1001 engages the corresponding sides of the member 301 and the housing 401. In the illustrated embodiment, the fluid muscle actuator structure 801 is configured to extend and contract substantially in the Z-axis direction along the ZX plane. The fluid muscle actuator 803 is coupled to the housing 401 at a first end by a first pivotable coupler 805 provided by a lug, a pin, and a U-bolt, and is coupled to the airflow changing device 301 at a second end by a second pivotable coupler 805 also provided by a lug, a pin, and a U-bolt. The pivot axis provided by the pin of the second coupler 805 is offset from the axis 1003 by a certain distance such that a lever arm is provided, and the extension and contraction of the fluid muscle actuator 803 enable the actuating rotation of the airflow changing device 301 about the axis 1003. In response to a deployment command from the flight control unit 117, the vortex generator control system 814 has commanded the vortex generator structure 115 to move to a first state, whereby the fluid muscle actuator 803 loses power to a threshold pressure P1 and extends to a length L1, and the airflow changing device 301 rotates about the axis 1003 to deploy into the airflow outside the opening 421, as shown.

[0089] Figure 10B Shows the same as Figure 10A the same structure, wherein, in response to a contraction command from the flight control unit 117, the vortex generator control system 814 has commanded the vortex generator structure 115 to move to a second state, whereby the fluid muscle actuator 803 is supplied with power to a peak pressure P2 and shortens to a length L2, and the airflow changing device 301 rotates about the axis 1003 to contract to a position inside the cavity 423 away from the airflow. This embodiment is also a fail-safe design, wherein the elastic strain energy stored in the fluid muscle actuator 803 is configured to be sufficient to deploy the airflow changing device 301, whether it is a commanded operation or in response to an accidental pressure loss in the vortex generator control system 814 or the actuator 803.

[0090] Using the fluid muscle actuator 803Figures 8A to 10B The implementation manner can be preferred because the use of the fluid muscle actuator 803 enables elastic strain energy to be stored by the actuator 803, and the stored elastic strain energy can then be additionally or fully used for the fail-safe deployment of the air flow changing device 301. In some embodiments, this can also allow reducing the elastic energy required to be stored by the base member 305, such that a less rigid and thus smaller or lighter base member 305 can be used. This allows more space for the actuation structure 801 to exist within the chamber 423, potentially allowing the use of a higher power actuator, or alternatively a more compact vortex generator structure 115, which is advantageous for the reasons described previously. Additionally, the use of fluid muscle actuators 803 is advantageous because they have a higher strength-to-weight ratio than conventional electric motors and hydraulic linear actuators, such that the weight of the vortex generator structure 115 is reduced. This is advantageous in aerospace where the weight of flying components is an important design requirement. Finally, the vortex generator structure 115 including the fluid muscle actuator structure 801 includes fewer moving parts and a substantially non-metallic configuration, which makes the design less susceptible to wear, corrosion, and blockage by contaminants such as water, ice, hydraulic oil, and / or dust.

[0091] In the case where known, obvious, or foreseeable equivalents of a whole or an element are mentioned in the foregoing description, then these equivalents are incorporated herein as if individually set forth. Reference should be made to the claims to determine the true scope of the present invention, and the claims should be construed to include any such equivalents. The reader will also understand that the overall components or features of the present invention described as preferred, advantageous, convenient, etc. are optional and do not limit the scope of the independent claims. Furthermore, it should be understood that such optional overall or features, while they may be beneficial in some embodiments of the present invention, may not be desirable in other embodiments and may thus be absent.

Claims

1. A vortex generator control system (814) for controlling a vortex generator structure (115) of an aircraft, the vortex generator control system (814) comprising: A controller (825) configured to receive one or more deployment or retraction command signals from a flight control unit (117) and further configured to send one or more command signals to a fluid control valve (819); A fluid pressure sensor (817) configured to sense one or more pressure values of a fluid muscle actuator (803) of the vortex generator structure (115) and send the pressure values to the controller (825); wherein the fluid control valve (819) is configured to control fluid transfer between the fluid muscle actuator (803) and a reservoir (823) in response to a command signal from the controller (825), and wherein the vortex generator control system (814) is configured to operate the fluid muscle actuator (803) of the vortex generator structure (115) in a pressure range of 0 Pa to 600000 Pa relative to ambient pressure.

2. The vortex generator control system (814) according to claim 1, wherein, The controller (825) is further configured to send a command signal to a fluid pump (821), the fluid pump (821) being configured to drive fluid from the reservoir (823) to the fluid muscle actuator (803) in response to a command signal from the controller (825).

3. The vortex generator control system (814) according to claim 1, wherein, The system is configured to operate using hydraulic fluid.

4. The vortex generator control system (814) according to claim 1, wherein, The system is configured to operate using pneumatic fluid.

5. The vortex generator control system (814) according to claim 1, comprising a plurality of pressure sensors (817) or fluid control valves (819).

6. The vortex generator control system (814) according to claim 1, wherein, The fluid control valve (819) is solenoid-operated.

7. The vortex generator control system (814) according to claim 1, wherein, The vortex generator control system (814) is configured to control the vortex generator structure (115) in a deployed state at an actuator pressure corresponding to approximately 600000 Pa relative to ambient pressure.

8. The vortex generator control system (814) according to any one of claims 1-7, wherein, The vortex generator control system (814) is configured to control a plurality of aircraft vortex generator structures (115).

9. An aircraft (101), wherein, The aircraft (101) comprises the vortex generator control system (814) according to any one of the preceding claims.

10. A method of controlling actuation of an air flow changing device (301) of a vortex generator structure (115) from a first position to a second position using a vortex generator control system (814), the method comprising the steps of: A flight control unit (117) signals a controller (825) to move the air flow changing device (301) of the vortex generator structure (115) from the first position to the second position, the vortex generator structure (115) comprising a fluid muscle actuator (803) operating in a pressure range of 0 Pa to 600000 Pa relative to ambient pressure; The controller (825) signals the fluid control valve (819) to open, allowing fluid to be transferred between the reservoir (823) and the fluidic muscle actuator (803) of the vortex generator structure (115), such that the airflow changing device (301) is actuated from the first position to the second position.

11. The method according to claim 10, wherein, The first position is the contracted state of the airflow changing device (301), and the second position is the deployed state of the airflow changing device (301).

12. The method according to claim 10, wherein, The first position is the deployed state of the airflow changing device (301), and the second position is the contracted state of the airflow changing device (301).

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