ACTUATOR SYSTEM FOR AEROSPACE CONTROLS AND FUNCTIONS

Abstract: The present invention provides an actuator system (10, 100, 200) for use in an aircraft control or operating system, comprising control means (12) operable in response to an input for generating a control signal, and an electrical actuator (20) responsive to the control signal for operating an aircraft flight control surface or other aircraft apparatus. Means (24; 124) for aerodynamically assisting the electrical actuator are also provided in order to reduce the load on the electrical actuator in use.

Description: ACTUATOR SYSTEM FOR AEROSPACE CONTROLS AND FUNCTIONS The present invention concerns actuator systems tor aerospace controls and other aerospace functions In particular, the invention concerns an actuator system tor operating the aerodynamic control surfaces ot an aircraft such as the ailerons, the wing leading edge and trailing edge flaps and slats, i e the tlaperons, the elevators, the spoi lers and the rudder The actuator system of the present invention mav also be employed for controlling the airbrakes, arrester hook flight re-fuelling probe, the undercai πage or pallets, and the doors and locks generally Actuator systems for operating aircraft control surfaces and tor performing other aerospace functions are well known Early such svstems were in the form of mechanical l inkage and lever arrangements operated manually by means ot a control column or rudder bar by which the pilot could directly adj ust the deflection or the aerodynamic control surfaces, for example Then, as aircraft increased in size and speed, additional force was required to operate the aerodynamic control surfaces, and various developments ensued The first ot these involved the mounting ot devices known as control tabs on the aerodynamic control surfaces themselves Such tabs were in the toi m ot small auxiliary control surfaces operated by the pilot and arranged in use to generate automatical I v an aerodynamic rorce designed to assist in moving the main control surface in the mannei desired In this wav the control tabs served to amplify aerodvnamically the pilot's effort for application to the main control surface, or they served to reduce the resistance to movement ot the main control surface, in situations or aircraft where direct operation by the pilot would otherwise be impractical Tnere were, however, a number ot drawbacks to the use ot control tabs the most significant being potential risks ot control surface flutter at resonant frequencies and potential shock wave problems at aircraft velocities approaching sonic conditions Flutter arises from a structural interaction between the aircraft frame and the control surface resulting in a relative oscillatory movement between them, which becomes exacerbated as resonant frequencies are approached Consequently, alternative actuator systems were developed for enabling the pilot to move the control surfaces, such systems constituting either power-assisted control systems or power-operated control systems. In both cases, the power was provided by hydraulic actuators, comprising pumps driven from the mam aircraft engιne(s) and hydraulic jacks or piston and cylinder arrangements connected bv fluid lines to the relevant pumps In a power assisted control system, the pilot still employs a control column or rudder bar tor operating the control surface, and the hydraulic jacks are arranged in such a way as to assist his efforts. In a power-operated control system, the pilot simply operates switches and valves to actuate the hydraulic jacks which themselves actuate the control surfaces. The hvdrau c actuators employed in such powei-assisted and powei- operated control systems are capable of producing high forces, and the necessary hinge moments tor deflecting the control surfaces can readi ly be generated within relatively small space requirements However, hydraulic systems are prone to leakage and wear, and ensuring that they remain at all times in a satisfactory operating condition requires a high level ot maintenance This has led to attempts to produce electrical ly powered actuation systems A number of proposals have been put forward and are employed in limited applications, but hitherto no satisfactory electrical ly powered actuation system has yet been produced. One currently available electrical arrangement features an electro-hydraulic actuator, comprising an electrical power source driven from the mam aircraft propulsion engιne(s) and a hydraulic actuator arranged to be operated by the electrical power source tor deflecting the control surfaces This system suffers tiom al l ot the disadvantages ot the known hydraulic actuator systems, together with the additional problems that the power source has a high mass, is expensive to produce and requires a substantial amount of aircraft space. Another electrical arrangement currently available features a conventional electric motor driven from the mam aircraft propulsion engιne(s) together with a gearbox/ball screw and mechanical l inkage to provide the required force and stroke tor operating the control surfaces Such arrangements suffer from problems ot wear and a tendency to jam, w hich is a significant drawback considering the satetv requirements in aircraft fl ight A further difficulty in employing electrical actuation systems tor aircraft control surfaces concerns the dissipation or heat, particularly tor high load high duty cvcle devices especially in the kinds of environment in which manv actuators are required to function Bv wav of example the tlaperon actuatois ot a fast jet wil l tvpicallv be located in an unconditioned wing bav having an ambient temperatuie of the ordei of 90°C In hvdrau c actuators the oi l employed for driving them can also be used for cooling purposes and the same pipework can serve tor both functions The majority ot the heat generated is thus removed bv the circulating hvdrauhc fluid and is simply cooled by an additional fuel cooled oil cooler Bv contrast an electric actuator requires a separate cooling system including separate pipi ng and cooling fl uid as well as a cooler and this adds to the bulk ot the actuator system Current efforts to introduce electrical actuation systems tor aircraft control surfaces have to a large extent been based simply on replacing the existing hydraulic actuators with electric actuators However as indicated above, this leads to a new set of problems For a similar power and dut\ cvcle a change from hvdrauhc actuation to electrical actuation wi ll result in Significantly increased actuatoi cooli ng problems Increased space lequirements Increased actuator mass Increased actuator cost Another problem concerns the power requirements tor the currently proposed and available electrical actuators Such an actuator when employed to operate the tlaperons ot a modern fast jet typical ly requires a powei supply ot the Older or 270 volt DC and draws peak currents between 100 and 1 50 amperes This poses a significant safety risk in passenger applications It is an aim of the present invention to provide an actuator system which overcomes the above problems Its is also an aim ot the present invention to proviσe an actuator svstem employing an electrical actuator which is compact and rel iable and in which cooling requirements and power consumption are reduced According to the piesent invention there is provided an actuator system for use in an aircraft control or operating system, comprising control means operable in response to an input for generating a control signal, an electrical actuatoi responsive to the control signal tor opeiatmg an aircraft flight contiol surface oi other ancratt apparatus and means foi aeioctvnamicailv assisting the electrical actuatoi in oider to reduce the load on the electrical actuator in u^e In a preferred form ot the invention described below tne actuator svstem is employed tor activating an aircraft flight control surface, and tne electrical actuator constitutes a linear electric motor while the means tor aerodynamicallv assisting the electrical actuator comprise a control tab incorporated in or mounted on the control surface Λs described the linear electric motor may be mounted within the aircraft frame that is within the fuselage oi a wing, or it mav be mounted in tne control tab itself The linear electric motor mav be ai ranged to deflect the control tab αirectly or bv means ot a l inkage mechanism or it mav be aπanged to deflect servo tabs mounted on the control tab, again eitner directly or thiough a l inkage mechanism Advantageously the control means are arranged to receive feedback signals tor example relating to control surface position and acceleration for generating an appropriate control signal to control tab acceleration Such feedback signals also represent the stroke position of the linear electric motor The present invention combines the advantages ot a relativelv low powered electrical actuator with aerodynamic amplification ot or an aerodynamic reduction in resistance to, the force required tor operating the aircrart control surface or other apparatus Potentially, this can enable very significant reductions in actuator mass and size, and in the cooling requirements and power consumption Linear electric motor technology otters simplicity, low wear, low maintenance and high accuracy In particular, linear electric motors are capable ot long strokes fast response times, and a high degree ot positional accuracy compared with existing aircraft actuators Although the achievable output forces compaied with known actuation techniques are relatively low, this is compensated by the use ot aerodynamic amplification or other aerodynamic assistance, tor example in the form ot control tabs The present invention is described further, by way of example, with reference to the accompanying drawings, in which Figure 1 is a diagrammatic view ot a first embodiment of actuator system according to the present invention, Figure 2 is a diagrammatic view ot a second embodiment ot actuator svstem and Figure 3 is a diagrammatic view ot a third embodiment ot actuator system Referring initial ly to Figure 1 , this shows an actuator system 10 according to the present invention tor operating an aircraft control surface, such as the ailerons the wing leading edge and trailing edge flaps and slats, or tlaperons, the elevators, the spoilers oi the rudder In the present instance, the control surface is an aileron or flaperon As shown, the actuator system 10 comprises a control ler 12, which may be a part ot the flight control computer or which may be a part of another aircraft system The controller 10 is located within the frame ot the aircraft tor operating a control surface 14 attached by a main pivot 1 6 to a wing 18 of the aircraft The pivot 1 6 is appropriately placed relative to the control surface 14 to provide a balanced mounting for the surface 14 The controller 12 is connected to an electric actuator 20 in the form of a linear electric motor, wherein the shaft and armature comprise a moving slide or carriage running on an air gap above a flat stator The electric actuator 20 is connected by way ot a linkage mechanism 22 to a control tab 24, the control tab 24 being pivotal ly mounted by way ot a hinge 26 on the control surface 14 The control ler 1 2 is arranged to receive an input signal from the pi lot, which signal is in the form or a position demand providing an instruction tor deflecting the control surface 14 to a new position In response to such input, the controller 12 is ai ranged to generate a coπesponding position control signal tor operating the electr ic actuator 20 The controller 12 is also arranged to receive as input signals feedback signals indicative ot the movement ot the control surface 14 and the control tab 24, tor generating a flutter control signal as descπoed below More particularly a position sensor 28 mounted between the aircraft wing 18 and the control surface 14 is ananged to supply signals to the control ler 1 2 lepresenting the current position or the control suitace 14 and accelerometers 30, 32 mounted on the control surface 14 and control tab 24 lespectivelv are arranged to supply to the control lei 12 signals representing the current acceleiation ot the control surface 14 and the control tab 24 Finally, an additional position sensor (not shown) within the electric actuator 20 is arranged to supply to the controller 12 a signal representing the linear stroke position of the electric actuator 20. The position sensors and accelerometers employed here are al l of a tvpe known generally as rotary or linear variable differential transformers However, potentiometers or Hal l Effect devices may be used instead Operation or the actuatoi svstem 10 is as fol lows On leceipt or an input signal from the pilot, the controller 12 monitors the position and acceleration signals from the position sensor 28, the accelerometers 30, 32 and the sensor within the electric actuator 20 and generates a position control signal representing a new stroke position for the electric actuator 20 The electric actuator 20 responds causing the l inkage mechanism 22 to adjust the position ot the control tab 24 The control tab 24 thus moves to a new position in which the aerodynamic forces acting on the control tab 24 generate a corresponding aerodynamic adjustment force acting on the control surface 14 It will be appreciated that the combined effect ot the force applied by the electric actuator 20 and the aerodynamic force applied by the controlled tab 24 serve together to generate a sufficient force to deflect the control surface 14 even in passenger aircraft or high speed jet aircraft In other words, the control tab 24 serves in use as an aerodynamic amplifier of the force supplied bv the electric actuator 20, in order to ensure that a sufficient moment is generated on the control surface 14 to deflect it on its hinge 1 6 bv the required amount Consequently, an electric actuator 20 having a relatively low output can be employed, thereby reducing the mass ot the actuatoi , its space requirements, its cool ing needs and its power consumption. As aircraft speeds increase, and in particular as they approach sonic velocities, the aircraft wing 18, the control surface 14 and the control tab 24 mav be subject to shock waves At some speeds where the resonant frequency of the control surface approaches the aerodynamic excitation frequency, control surface flutter may also arise. In order to counteract these effects, two steps have been taken in the present actuator system 10: Firstly, the stiffness ot the control surface 14 including also tne stiffness ot the attachment ot the control surface 14 to the wing 1 8, are made as high as possible. Likewise, the stiffnesses ot the control tab 24 and its attachment to the control suitace 14, as well as the actuator impedance, are made as high as possible This has the effect ot ensuring that the conditions at which resonance occur arise only at speeds above the normal speed range ot the aircraft, so that control surface flutter is minimised. Secondly, the control surface flutter is detected bv means or the accelerometers 30, 32 and actively compensated bv means of the control ler 1 2 The controller 12 is arranged continually to monitor the outputs from the accelerometers 30, 32 and to generate a flutter control output accordingly to the electric actuator 20 for suppressing the flutter. By employing an electric actuator 20 in the form of a l ineai electric motor, the actuator 20 wil l have a sufficiently high/wide bandwidth and high performance to be able to respond to such flutter control signals from the controller 12 in oider to achieve an active adjustment and suppression of control surface flutter. If necessary, in order to increase the bandwidth ot the electric actuator 20, a plurality ot l inear electric motors mav be connected in parallel to the control tab 24. This has the additional advantage ot providing failure redundancy and enhanced safety regulation Alter natively, a multichannel linear electric motor may be employed instead to increase the bandwidth Turning now to Figure 2 this snows a modified actuator svstem 100 operating on a similai principle to the actuator svstem 10 shown in Figure 1 Similar paits are designated bv the same reference numerals Only the differences in design and operation will be described The actuator svstem 100 comprises a controller 12 as before for operating a control surface 14 mounted on a wing 18 of an aircraft The control suiface 14 and its mounting on the wing 18 aie identical to the arrangement shown in Figuie 1 altnough some ot the details are omitted in Figure 2 In the present instance the control tab 24 is replaced a control tab 124 mounted on the control surTace 14 by way or a hmse 126 The control tab 124 is a compound tab comprising a mam tab 128 and a pair or servo tabs 130 nmgedlv connected to the uppei and lower surfaces of the main tab 128 Another modification is that the electric actuator 20 and the linkage mechanism 22 are now housed within the mam tab 128 ot the control tab 124 rather than within the frame ot the aircraft, and the linkage mechanism 22 is connected to the servo tabs 130 The operation or the actuatoi svstem 100 is Minilai to the actuatoi svstem 10 except in that the actuatoi 20 and linkage mechanism 22 aie anangeα to deflect tne servo tabs 130 rather than the contiol tab 124 as a whole an aerodynamic force toi deflecting the main control tab 128 and consequently the contiol surface 14 Active flutter oppression is achieved in the same wav as in the actuator system 10, by employing the signals from the accelerometers 30 32 and generating an appropriate compensatory output from the controller 12 A fuither modification of the invention is ^nown in Figure 3 which illustrates an actuator system 200 operating on a similar principle to the actuator system 10 shown in Figure 1 and the actuator system 100 shown in Figure 2 The same parts are designated by similar reference numerals, and only the differences will be described The actuator svstem 200 again comprises a control ler 1 2 for operating a control surface 14, and a compound tab 124 is mounted on the control surface 14 by way of a hinge 126 As in the case ot the Figure 2 embodiment the compound tab 124 comprises a mam control tab 128 and a pair a servo tabs 1 30 hinged on the upper and lower surfaces of tne main control tab 128 In the Figure 3 embodiment, the electric actuator 20 is again mounted within the compound tab 124, but in this instance the linkage mechanism 22 is omitted altogether. Instead, the electric actuator 20 is arranged to actuate the servo tabs 1 30 directly tor the purpose ot generating an aerodynamic force tor deflecting firstly the main control tab 128 and subsequently the control surface 14 In other respects i ncl uding in the pi ovision or activ e fl utter suppression the actuator system 200 operates identical ly with the actuatoi svstem 100 in Figure 2 Various modifications are possible within the scope ot the invention. For example, the described actuator systems may be employed for operating other control surfaces ot the aircraft than those described, or even tor other aircraft functions, including the operation ot aircraft apparatus such as the undercai πage When employed for operating the control surfaces of the ancratt, the actuator system mav comprise alternative forms or control tab, for example a control tab which is an integral part ot the control surface and which is effective bv virtue of the control surface having an offset mam pivot Alternative means other than a control tab may even be employed tor aerodynamical lv amplifying the output of the lineai electi ic actuatoi The described embodiments all feature a l inear electric actuator arranged to operate the control or servo tabs However, tor reasons of redundancy oi safetv, the same or an additional l inear electric actuatoi mav also be arranged to operate the control surface itself. Another possibil ity is toi the described linear electric actuatoi to operate the control surface directly whi le a mechanism between the control surrace and the contiol tab operates the control tab. The present invention offers a significant advance in aircraft control in terms of flexibility, reliability and performance. The need for maintenance is likely to be significantly reduced by comparison with known hydraulic actuators, and the costs oi production and operation are expected to be highly advantageous.

Claims: CLAIMS An actuator system for use in an aircraft control or operating system comprising control means (1 2) operable in response to an input tor generating a control signal, and characterised by the combination of an electrical actuator (20) responsive to the control signal tor operating an aircraft flight control surface or other aircraft apparatus, and means (24, 124) tor aerodynamical ly assisting the electrical actuator in order to reduce the load on the electrical actuator in use An actuator system according to claim 1 tor operating an aircraft control surface, characterised in that the aerodynamic assistance means comprise a control tab forming a part of the control surface and obtained bv mounting the control surface on the associated section ot the aircraft by means ot an offset hinge An actuator system according to claim 1 for operating an aircraft control surface characterised in that the aerodynamic assistance means comprise a control tab (24, 124) mounted on the control surface An actuator system according to claim 3, characterised in that the control tab comprises a compound control tab (124) having a main tab (128) connected bv means of a hinge (126) to the control surface, and one or more servo tabs (1 30) pivotablv connected to the mam tab An actuator system according to any ot claims 2 to 4 characterised in that the electrical actuator is connected to the control tab tor operating the control surface An actuator system according to claim 5, characterised in that the or a further electrical actuator is connected to the control surface tor operating the same An actuator system according to anv of claims 2 to 6 characterised in that the electrical actuator comprises a l inear electric motor An actuator system according to claim 7, characterised in that the linear electric motor is mounted within the control tab 9 An actuator system according to any of claims 2 to 8, further characterised by feedback means (28, 30, 32) for generating signals representing the present operating position or condition of the aircraft control surface, or apparatus, for providing an input to the control means 5 10 An actuator system according to claim 9, characterised in that the feedback means comprise a position sensor (28) associated with the control surface, and/or an accelerometer (30, 32) associated with at least one ot the control surface and the control tab 101 1 An actuator system according to any one of claims 2 to 10, further characterised by means (12, 30, 32) for suppressing flutter ot the control surface, the flutter suppression means comprising means foi generating signals lepresentmg the movement ot the control surface 1512 An aircraft including an actuator system according to any preceding claim

Inventors: BLACKWELL NIGEL, CHAMBERS PETER WILLIAM, UTTLEY ANDREW EDWARD, WELLER BRIAN, BLACKWELL, NIGEL, CHAMBERS, PETER, WILLIAM, UTTLEY, ANDREW, EDWARD, WELLER, BRIAN
Applicants: BAE SYSTEMS PLC, BLACKWELL NIGEL, CHAMBERS PETER WILLIAM, UTTLEY ANDREW EDWARD, WELLER BRIAN, BLACKWELL, NIGEL, CHAMBERS, PETER, WILLIAM, UTTLEY, ANDREW, EDWARD, WELLER, BRIAN
Application year: 2000
Main classification: B64C13
IPC codes: B64C9, B64C13