|Publication number||US3679156 A|
|Publication date||Jul 25, 1972|
|Filing date||Jul 20, 1970|
|Priority date||Jul 20, 1970|
|Also published as||DE2135366A1, DE2135366C2|
|Publication number||US 3679156 A, US 3679156A, US-A-3679156, US3679156 A, US3679156A|
|Inventors||Redmond William G Jr|
|Original Assignee||Ltv Electrosystems Inc|
|Export Citation||BiBTeX, EndNote, RefMan|
|Patent Citations (8), Referenced by (26), Classifications (7)|
|External Links: USPTO, USPTO Assignment, Espacenet|
United States Patent Redmond, Jr. July 25, 1972 s41 FLY-BY-WIRE 3,426,650 2/1969 Jenney ..9l/363 A x 3,505,929 4/1970 Coppola et al.. ...9l/4l1 R X  wiu'am kedmm" Dallas 2,898,889 8/1959 Foster ..244/85  Assignee: LTV Electrosystems, lnc., Greenville, Tex. 3,027,378 1962 K y 6 /73 X 3,422,767 1 196 M Al l 5 22 Filed: July 20, 1970 9 vay 9  Appl. No.: 56,551 Primary Examiner-Milton Buchler Assistant Examiner-F. K. Yee Attorney-Richards, Harris & Hubbard and James D. Will-  0.8. Cl. ..244/77 R, 91/411 R, 244/78, born 244/85, 318/564  Int. Cl. ..B64c 13/50 57 B TR  Field of Search ..244/ R, 75 A, 76 R, 76 A,
244/76 B, 76 C 77 R. 77 v, 78' 83 R 83 E 85; Three separately generated electr cal control signals are up- 91/363 A, 363 R, 411 A 411 B 411 R I 505407; plied to respective electromechanical transducers In a power 74/479; 235/61 A 61 C, 417/222 426; actuator control system. Each electromechanical transducer 318563465 generates a mechanical output motion related to the applied control signal and independent of the remaining transducers. A mechanical linkage arrangement coupled to the outputs of  References Cned each of the three electromechanical transducers combines the UNITED S ES PATENTS motions into a single mechanical movement. This single mechanical movement positions the control element for a 3,554,084 l/l97l Rasmussen et al. ..9l/4li R X servo pump supplying a power actuaton 3,469,162 9/1969 Goshn ..3l8/564 3,401,600 9/ 1968 Wood ..91/4ll R X 5 Claims, 4Drawing Figures a2 66 so 74 o 2o 78 a0 a 22 as e iir'e I a4 SUPPLY /0\ 40 54 62 7o 76 25 J 26 3o 42 45 72 5' a 7 a 46 56 36 I2 34 32 3e 92 44 SERVO SERVO a PUMP PUMP I42 4 SO-f 50 52 COMPARATOR CHANNEL MONITOR SELECTOR PATENTEDJULZS 1912 SHEET 8 [IF 2 INVENTOR: WILLIAM G. REDMOND, JR.
ATTORNEYS FLY-BY-WIRE This invention relates to redundant electrical control, and more particularly to redundant electrical control employing displacement summing techniques.
It was early realized that as aircraft increase in size and speed that conventional cable and mechanical linkage control mechanisms are inadequate and there is a need for electrical flight control systems. There has, however, been some reluctance to accept the electrical flight control system because it is thought that mechanical systems are more reliable. To improve the reliability of electrical control, a system of redundant parallel channels has been implemented.
I-Ieretofore, approaches for providing redundancy have typically resulted in double or triple control chains or channels in which a failure in one channel hopefully would permit the other channels to carry on the necessary command functions. Such a system, depending upon the particular failure suffered, generally experienced at least degradation of control when the failed channel must be dragged by the operating channel or channels. Because most electrical failures are not just loss of electrical power or other passive failures but are often hardover or other grossly discrepant commands, three or more redundant channels are frequently employed. In the usual redundant control, the complete system from the signal generating source to the control surface is repeated. As the number of redundant channels increases, synchronization of the control motion at the control surface becomes a problem.
In the typical push rod and cable controlled mechanical systems, dual-tandem hydromechanical power servos are used with each section thereof provided with a separate hydraulic supply. The signal or intelligence paths to the power servo is mechanical and single, except at the dual servo for controlling the hydraulic fluid to the dual hydromechanical power servo. In such systems, synchronization is feasible and easily accomplished. In the dual hydromechanical power servo systems, the failure of one of the hydraulic systems is passive and not fighting the remaining operating system. These last two features of mechanical system have, heretofore, not been found in the redundant electrical control signal.
An object of the present invention is to provide a redundant electrical control system terminating at a single power actuator. Another object of this invention is to provide a redundant cal motion equal to the sum of the individual movements. This single mechanical motion is connected to a power actuator which responds thereto and has an output controlling the movement of the aircraft control surface.
A more complete understanding of the invention and its advantages will be apparent from the specification and claims and from the accompanying drawings illustrative of the invention.
Referring to the drawings:
FIG. 1 is a schematic of the redundant control system including displacement summed electromechanical transducers for controlling a dual-tandem power actuator connected to a control surface;
FIG. 2 is a cross section of an electromechanical transducer for the system of FIG. 1;
FIG. 3 is a sectional view of a motor driven variable displacement servo-pump for controlling hydraulic fluid to the power actuator of the system of FIG. 1; and
FIG. 4 is a schematic of a three channel comparator system.
Referring to FIG. 1, there is shown a fly-by-wire redundant control system wherein the electrical portion is made triply redundant to give a low probability of all channels failing during a single operation. The three electrical channels are totally independent; the output of each being summed in the form of position or displacement" summation techniques of mechanical movements of electromechanical transducers l0, l2, and 14. Each of the electromechanical transducers is energized by a separately generated electrical signal. The electrical signals may be generated by a pilot control stick transducer that converts a mechanical motion into electrical signals. In addition to the pilot generated signals, the electrical signals to the transducers may be received from autopilot sensors, a stability augmentation system, or from other systems, such as a navigation control.
Although the system to be described has electromechanical transducers in each control channel, other electrically responsive transducers may be used. For example, electrohydraulic transducers responsive to an electrical signal and generating a electrical control system without channel synchronization. A
further object of this invention is to provide a redundant electrical control system wherein a failure of one channel does not light the remaining operating systems. Still another object of this invention is to provide a redundant electrical control signal having a low probability of all channels failing during a control operation.
In accordance with the present invention, redundant control channels are employed to improve the system reliability. The several electrical channels terminate at a single power actuator that positions the control surface of an aircraft. The several channels of the system are displacement summed that is, the output of each channel is combined by a mechanical linkage arrangement to generate a single motion related to the sum of all the channels. Displacement summing has the advantage that the remaining active channels do not have to drag a failed channel. When the several redundant channels are combined to produce the total desired movement of the aircraft control surface, a failure of one of these channels results in some loss of stroke of the power actuator, with the attendant loss in the range of movement of the control surface, but control is maintained. Further, with displacement summing, over-stroke capability in each of the three channels can be built into each actuator to provide full stroke upon the loss of a single channel.
In accordance with a more specific embodiment of the invention, an aircraft control system for positioning a control surface in response to electrical control signals includes at least three transducers each having an output motion related to an electrical control signal applied thereto. Each of the output motions of the individual transducers are displacement summed such that the combination results in a single mechanimechanical motion through hydraulic action may be used.
Each of the electromechanical transducers 10, 12 and 14 operates independent of each other. For example, a signal for energizing the transducers 10 is supplied from an amplifier 16 connected to a power supply 18 and receiving an input signal from a summing junction 20. The summing junction 20 differentially combines the electrical control signal from the source, as explained, with a feedback signal from a position transducer 22. Similarly, the transducer 12 is energized by a signal from an amplifier 24 coupled to a power supply 26 and receiving an input signal from a summing junction 28. The summing junction differentially combines the control signal for the transducer 12 with a feedback signal from a position transducer 30. The third channel, containing the electromechanical transducer 14, is similar to the previous two. The transducer 14 is energized by a signal from an amplifier 32 connected to a power supply 34 and receiving an input signal from a summing junction 36. The summing junction 36 differentially combines a control signal for the transducer 14 with a feedback signal from a position transducer 38.
Each of the transducers 10, 112, and 14 thus produces a unique output motion for the electrical control signal applied thereto. The separate mechanical motions are summed by mechanical summing linkages 40 and 42. Summing link 40 combines the output motion of the transducer 10 with the output motion of the transducer 12. The summing link 42 combines the movement of the summing link 40 with the output motion of the transducer 14. With the arrangement illustrated, each of the transducers has a gain factor (a fraction of the total motion) equal to one-third of the total output motion produced by the links 40 and 42, as evidenced by the mechanical motion of a connecting rod 44.
The combined movement of the transducers l0, l2, and 14 is imparted to the connecting rod 44 coupled to lever arms 46 and 48 of variable displacement servo-pumps 50 and 52,
respectively. As an alternative, instead of servo-pumps, the connecting rod 44 may be coupled to a power servo using a flow-controlled servo valve. An over limit stop 54 is provided for the connecting rod 44 to limit the travel thereof to the maximum allowable displacement of the lever arms 46 and 48. This over limit stop thus absorbs the excess movement at the output of each of the transducers 10, 12, and 14 when provided with over-stroke capability. The lever arms 46 and 48 are interconnected by a synchronizing arm 56 having a synchronizing adjustment nut 58. In the system shown, this is the only synchronization required. This adjustment is made when the system is initially put into operation.
The servo-pumps 50 and 52 are part of separate hydraulic sources connected to a dual-tandem power actuator 60. Conduits 62 and 64 connect the servo-pump 50 to the first stage of the actuator 60 on opposite sides of a piston 66. A pressure relief valve 68 is connected between the conduits 62 and 64 in accordance with standard procedures. Conduits 70 and 72 similarly interconnect the servo-pump 52 to the second stage of the power actuator 60 on opposite sides of a piston 74. Again, a pressure relief '76 interconnects between the conduits 70 and 72.
Pistons 66 and 74 of the power actuator 60 are interconnected on a piston rod 78 that has an external coupling to a link 80. Link 80 is intended to represent the mechanical linkage between the power actuator 60 and one of the control surfaces 82 of an aircraft. Also connected to the piston rod 78 is a follow-up (feedback) rod 84 for repositioning the lever arms 46 and 48 in response to movement of the control surface 82.
As illustrated, each electrical control channel culminates in a mechanical link displacement with the three displacements of the transducers 10, 12 and 14 added to give a single mechanical input to the dual hydromechanical system including the actuator 60. Each channel gain can have as much percentage tolerance as would a single, non-redundant control system; the resultant percentage tolerance from the three summed channels will be no greater than the percentage tolerance of each channel. Also, the phase angle of each of the three channels does not need to be more precise than for a single non-redundant system, thereby alleviating the synchronization problem.
Referring to FIG. 2, there is shown in cross section the electromechanical transducer for the actuator system of FIG. 1. The other transducers of the system are similarly constructed. As illustrated, the rotary-to-linear motion transducer has a permanent magnet stator 120 in a housing 122. Pilot generated signals are applied to the transducer through a connector 124 to a brush ring 125 and then to an armature 126 in the form of a DC motor, an AC motor may also be used or a brushless DC motor. The armature 126 rotates by the interaction between its electrical field and the magnetic field of the stator 120. It is mounted to rotate in the housing 122 by means of bearings 128 and 130. A ball nut 131 is fixed to the armature 126 and engages a lead screw 132 that extends through the housing 122 as part of the output shaft 133. Linear motion results from the operation of the lead screw 132 and the ball nut 131; this motion is imparted to the shaft 133 which in turn is connected to the mechanical summing link 40. To produce this linear motion, the lead screw 132 is restrained from rotating by means of a dowel pin 134.
Also, included as part of the transducer is a linear voltage differential transformer 136 threaded into the housing 122 and engaging the lead screw 132 for generating two identical position feedback signals. These feedback signals are transmitted through the connector 124 to the summing junction 20 and to a comparator monitor, as will be explained.
An output motion from the transducer 10, as evidenced by movement of the lead screw 132 is transmitted to the at- I tachment point of the mechanical link 40 through a neutralizer 138. A neutralizer is a mechanical means by which a majority vote" of the respective actuator outputs can be transmitted to the connecting rod 44. Similar neutralizers are provided for each of the remaining channels. The transducer 12 includes a neutralizer 140 and the transducer 14 includes a neutralizer 142.
Thus, a malfunctioning channel may be automatically neutralized by means of the respective neutralizer. The neutralizers primary function is to remove a transducer in a hardover" failure condition. A hardover failure is where a channel operates in opposition to the remaining channels. In the hardover" failure condition, it is desirable to automatically de-energize and neutralize a malfunctioning channel.
Referring to FIG. 3, there is shown in section a simplified diagram of the variable displacement servo-pump 50 with the lever arm 46 coupled to a rotatable swash plate 144. The servo-pump 52 of FIG. 1 may be similar to that illustrated in FIG. 3 with the lever arm 48 interconnected to the lever arm 46 through the synchronizing arm 56.
Conduits 62 and 64 are connected to the pump housing 146 through inlet/outlet connectors 148 and 150. These connectors have openings to a piston block 152 that contains a plurality of pistons, such as pistons 154 and 156, arranged in a circular pattern. The piston block 152 is coupled to the drive shaft of a motor 158, FIG. 1, through a pump shaft 160.
The operation of the pump is conventional with the pump capacity and fluid direction established by an angular position of the swash plate 144. By coupling the swash plate 144 directly to the connecting rod 44 through the mixing link 45 as also connected to the follow-up rod 84, there is eliminated the need for the usual hydraulic actuator heretofore used to position the swash plate of a servo-controlled variable displacement pump. Further, with the electromechanical transducers coupled directly to the servo-pump swash plate, the hydraulics of the system are separated from the electrical fly-by-wire portion.
With the system illustrated, the failure combination situation is improved by not jeopardizing the hydraulic system by tying the electrical-mechanical transducers directly thereto. Any one or two of the electromechanical transducers may fail and at the same time either of the hydraulic systems can fail and control of the surface 82 will be maintained.
To sense the failure of any of the electromechanical transducers 10, 12, or 14, each is provided with a second position transducer connected to a comparator monitor 86. Thus, the transducer 10 includes a position transducer 88 generating a feedback signal to the comparator monitor 86. Similarly, the transducer 12 includes a position transducer generating a feedback signal to the monitor, and the transducer 14 includes a position transducer 92 generating a feedback signal to the comparator monitor. The comparator monitor 86 determines when one channel disagrees with the other two by more than a predetermined limit and generates a signal to a channel failure indicator 94. The comparator monitor 86 also automatically de-energizes and neutralizes (centers) the disagreeing channel. This indicator may be located at the pilots position for information or possible override of a faulty monitor cutting off a good channel.
Referring to FIG. 4, there is shown one embodiment of a comparator monitor for a three channel redundant system. Position signals from the transducers 88 and 90 are applied to input terminals of differential amplifiers 96 and 98. Similarly, position signals from the transducers 90 and 92 are applied to the inputs of differential amplifiers and 102. Position signals from the transducers 88 and 92 are also applied to inputs of differential amplifiers 104 and 106.
Each of the differential amplifiers has a saturated output stage such that the output signal is either at a cut-off voltage level or a saturated voltage level. The input section of each of the amplifiers includes circuitry such that if the upper terminal is at a higher voltage level than the lower terminal, the output stage will be cut off and if the upper terminal is at the lower voltage level, the output stage will be operating at saturation. The input voltage differential required to switch the output stage may be set by an amplifier adjacent. Such amplifiers are conventional in the art.
The output voltage from each of these amplifiers is applied to inputs of an arrangement of NAND-gates 108 through 113 in the configuration as illustrated. Output terminals of the NANDS 108 and 109 are applied to NOR-gate 114. Output terminals of the NANDS lland 111 are similarly connected to a NOR gate-116 and outputs of the NANDS 112 and 113 are connected to a NOR-gate 118.
With all three channels operating properly, the output of each of the NORS 114, 116, and 118 will be at a logic ZERO level. Should any one of the three channels differ by a preset amount, (a differential amplifier setting) from the other two, one of the NOR gates will switch to a logic ONE level. As connected, when the transducer 10 fails the output of the NOR 114 will switch to a logic ONE level. If the transducer 12 fails, the output of the NOR 116 will switch to a logic ONE level, and if the transducer 14 fails, the output of the NOR 118 will change to the logic ONE level. it should be understood that the comparator logic of FIG. 4 is only one example of a comparator monitor 86 that may be used with the system of FIG. 1.
With the system as illustrated and described, each electrical channel is provided with displacement capability sufficient to satisfactorily control an aircraft (perhaps with somewhat less than full surface throw but possibly with full-stroke capability) should one or more of the other channels be neutralized. The gain of each channel will be varied in accordance with the number of channels operating, that is, with three operating channels, each will contribute about one-third the total movement of the connecting rod 44. With more than one channel operating (normal condition) it is possible that the mechanical stop 54, which limits the stroke command to the servo-pumps 50 and 52, will be reached, and the electrical actuators will be neutralized. This, however, is a satisfactory condition. The system described has the further advantage in that all channels are active when in the normal operating mode.
While only preferred embodiments of the invention, together with modifications thereof, have been described in detail herein and shown in the accompanying drawings, it will be evident that various further modifications are possible without departure from the scope of the invention.
What is claimed is:
1. A system for positioning a control element in response to electrical control signals, comprising:
at least three electro-mechanical transducers each having an output motion related to an electrical signal applied thereto,
mechanical summing linkage coupled to the output of each of said transducers such that each transducer contributes to a single mechanical motion by an amount inversely proportional to a number of said transducers,
a position transducer at each of said electro-mechanical transducers generating a first feedback signal and a second feedback signal,
means for summing said first feedback signal with one of the electrical control signals to selectively control the respective electromechanical transducer,
comparing means monitoring each of the second feedback signals and generating an indication when one differs from any combination of at least two of the other second feedback signals, and
actuator means responsive to the single mechanical motion of said mechanical summing linkage and having an output controlling the movement of the control elements.
2. A system as set forth in claim 1 including neutralizing means at each of said transducers to effectively disconnect the associated transducer from said mechanical summing linkage.
3. A control system as set forth in claim 2 wherein said actuator means includes a duo-tandem hydraulic actuator responsive to the single mechanical motion of said summing linkage and having an output controlling the movement of the control element.
20 4. An aircraft control system for positioning a control surface in res onse to electrical control signals comlprisjng:
at least ee electro-mechamcal transducers avrng an output motion related to a separate electrical signal applied thereto,
mechanical summing linkage coupled to the output of each of said transducers such that each transducer contributes to a single mechanical motion by an amount inversely proportional to the number of said transducers,
a position transducer at each of said electro-mechanica] transducers generating a first feedback signal and a second feedback signal,
means for summing said first feedback signal with one of the electrical control signals to selectively control the respective electromechanical transducer,
comparing means monitoring each of the second feedback signals and generating an indication when one differs from any combination of at least two of the other second feedback signals,
a duo-tandem hydraulic actuator having an output controlling the movement of the control surface, each section of said actuator supplied operating fluid from a separate source, and
pump means for each of said supply sources for controlling the fluid therefrom to the respective actuator stage, said pump means having a positionable swash plate coupled to the single mechanical motion of said displacement summing means.
5. An aircraft control system as set forth in claim 4 including neutralizing means at each electromechanical transducer to effectively disconnect the respective transducer from said mechanical summing linkage.
|Cited Patent||Filing date||Publication date||Applicant||Title|
|US2898889 *||Oct 3, 1958||Aug 11, 1959||John V Foster||Mechanically-limited, electrically operated hydraulic valve system for aircraft controls|
|US3027878 *||Jun 10, 1959||Apr 3, 1962||Honeywell Regulator Co||Fluid motor control apparatus|
|US3401600 *||Dec 23, 1965||Sep 17, 1968||Bell Aerospace Corp||Control system having a plurality of control chains each of which may be disabled in event of failure thereof|
|US3422767 *||Dec 5, 1966||Jan 21, 1969||Webster Electric Co Inc||Variable displacement swashplate pumps|
|US3426650 *||Dec 23, 1965||Feb 11, 1969||Bell Aerospace Corp||Triple channel redundant hydraeric control system|
|US3469162 *||Apr 1, 1966||Sep 23, 1969||Hawker Siddeley Dynamics Ltd||Multiplex-type control apparatus|
|US3505929 *||Apr 16, 1968||Apr 14, 1970||Gen Electric||Redundant flight control servoactuator|
|US3554084 *||Nov 17, 1967||Jan 12, 1971||Honeywell Inc||Redundant force summing servo unit|
|Citing Patent||Filing date||Publication date||Applicant||Title|
|US3955783 *||Dec 4, 1974||May 11, 1976||Joseph Lucas (Industries) Limited||Hydraulic actuating arrangement for aircraft control surfaces|
|US4466597 *||Dec 2, 1981||Aug 21, 1984||Pneumo Corporation||Electro-mechanical direct drive valve servo system with rotary to linear valve drive mechanism|
|US4596177 *||Nov 12, 1982||Jun 24, 1986||Rockwell International Corporation||Actuator system|
|US5033694 *||Sep 8, 1989||Jul 23, 1991||Daiichi Electric Kabushiki Kaisha||Attitude control device for air or sea transportation craft|
|US5074495 *||Mar 16, 1990||Dec 24, 1991||The Boeing Company||Load-adaptive hybrid actuator system and method for actuating control surfaces|
|US5082208 *||Sep 29, 1989||Jan 21, 1992||The Boeing Company||System and method for controlling an aircraft flight control member|
|US5182505 *||Jun 19, 1991||Jan 26, 1993||Honeywell Inc.||Aircraft control surface position transducer|
|US5493497 *||Jun 3, 1992||Feb 20, 1996||The Boeing Company||Multiaxis redundant fly-by-wire primary flight control system|
|US5670856 *||Nov 7, 1994||Sep 23, 1997||Alliedsignal Inc.||Fault tolerant controller arrangement for electric motor driven apparatus|
|US5678786 *||Dec 6, 1995||Oct 21, 1997||Mcdonnell Douglas Helicopter Co.||Reconfigurable helicopter flight control system|
|US5806805 *||Aug 7, 1996||Sep 15, 1998||The Boeing Company||Fault tolerant actuation system for flight control actuators|
|US5934614 *||Jul 11, 1997||Aug 10, 1999||Daimlerchrysler Aerospace Airbus Gmbh||Closed loop control system for controlling an air discharge out of an aircraft body|
|US6209825||Feb 27, 1998||Apr 3, 2001||Lockheed Martin Corporation||Low power loss electro hydraulic actuator|
|US6439512 *||Aug 24, 2000||Aug 27, 2002||Hr Textron, Inc.||All-hydraulic powered horizontal stabilizer trim control surface position control system|
|US6443399 *||Jul 14, 2000||Sep 3, 2002||Honeywell International Inc.||Flight control module merged into the integrated modular avionics|
|US6481203||Jun 5, 2000||Nov 19, 2002||Tecumseh Products Company||Electric shifting of a variable speed transmission|
|US6923405 *||Nov 17, 2003||Aug 2, 2005||The Boeing Company||Enhanced rudder control system|
|US7786684||Oct 22, 2007||Aug 31, 2010||Honeywell International Inc.||Electromechanical flight control system and method for rotorcraft|
|US8172174||Nov 13, 2008||May 8, 2012||Honeywell International Inc.||Hybrid electromechanical/hydromechanical actuator and actuation control system|
|US9327825 *||Mar 22, 2012||May 3, 2016||Bae Systems Plc||Actuator control system|
|US20050116095 *||Nov 17, 2003||Jun 2, 2005||Cline Paul J.||Enhanced rudder control system|
|US20090102413 *||Oct 22, 2007||Apr 23, 2009||Honeywell International, Inc.||Electromechanical flight control system and method for rotorcraft|
|US20100116929 *||Nov 13, 2008||May 13, 2010||Honeywell International Inc.||Hybrid electromechanical/hydromechanical actuator and actuation control system|
|US20140172203 *||Mar 22, 2012||Jun 19, 2014||Bae Systems Plc||Actuator control system|
|EP0152714A1 *||Dec 27, 1984||Aug 28, 1985||AEROSPATIALE Société Nationale Industrielle||Aircraft flight control system|
|WO1996014610A1 *||Nov 7, 1995||May 17, 1996||Alliedsignal Inc.||Fault tolerant controller arrangement for electric motor driven apparatus|
|U.S. Classification||244/194, 318/564, 244/227, 91/510|