Triplex cockpit control data acquisition electronics

A system and method are provided for controlling aircraft flight control surfaces. The system may include at least three pilot sensor channels, each pilot sensor channel including a set of pilot sensor data. The system may also include at least three aircraft sensor channels, each aircraft sensor channel including a set of aircraft sensor data. The system may further include an actuator control component configured to synchronously receive and vote on the pilot sensor data and the aircraft sensor data, such that a voted output of the at least three pilot sensor channels is transmitted to a flight control computer and augmented before being transmitted to remote electronics units. The voted output of the at least three pilot sensor channels providing for the control of the aircraft surfaces coupled to the remote electronics units.

BACKGROUND

In the past, flight control systems used in many aircraft relied on direct mechanical linkages between the pilot's control devices and the aircraft flight control surfaces. In such systems, pilot-manipulated flight control devices, such as pedals, levers, and the control column, would cause connected mechanical linkages to transmit movement of the pilot control devices to the appropriate aircraft flight control surface, such as the rudders, ailerons, and elevators. Such a rudimentary system of flight control allowed for direct control of the aircraft flight control surfaces and was typically a highly reliable system. However, the mechanical flight control system required frequent, intensive inspection due to the wear and tear on the mechanical linkage, added an undesirable amount of weight to the aircraft, and required a large amount of space to properly operate.

With the advent of analog and digital circuitry, a new flight control system, commonly referred to as “fly-by-wire” or FBW, was developed whereby electronic signals generated in response to the pilot-manipulated control devices could be sent to electronic control devices to process and control movement of the flight control surfaces of the aircraft. The implementation of the fly-by-wire aircraft control system reduced the amount of weight added to the aircraft and provided a smaller footprint in the aircraft as well.

Generally, in conventional fly-by-wire aircraft control systems, electronic signals are generated by the pilot-manipulated control devices, such as the control column, and sent to a flight control computer (FCC), which also receives electronic signals from aircraft sensors providing information related to the aircraft's speed, altitude, angle of attack, and the like. The FCC processes the received electronic signals and generates a correlating instruction signal and sends the instruction signal to one or more actuator control electronics (ACE), where the instruction signal is further processed and transmitted to the appropriate actuator, which may be mechanically coupled to the respective aircraft flight control surface, e.g., the rudder, aileron, or elevator. The instruction signal provided by the FCC causes the actuator to move the aircraft control surface to correspond to the input provided by the pilot, and in some cases, the input provided by the pilot and augmented by the input from the FCC. In such cases in which the FCC augments input provided by the pilot, the fly-by-wire aircraft control system is commonly referred to as operating in “normal mode.”

Although the fly-by-wire aircraft control systems are easier to maintain than mechanical systems, the signals generated by modern control fly-by-wire systems can be very complex, and certain failures of electronic subsystems may lead to loss of operational control. In addition, the data buses and/or wires interconnecting the control electronics, actuators, and sensors may become damaged or disconnected, thereby causing interference with, or loss of, the pilot's ability to control the aircraft.

To counter such safety concerns, conventional fly-by-wire aircraft control systems provide for a plurality of microprocessors in each of the FCCs and the ACEs to protect against “random failures.” In addition, the conventional fly-by-wire aircraft control systems typically also include a plurality of dissimilar microprocessors in each of the FCCs and the ACEs to protect against “common mode failures” and may also include a plurality of dissimilar FCCs and ACEs to protect against “generic failures.” Further yet, to provide for the safe operation of the aircraft, many conventional fly-by-wire aircraft control systems include redundant FCCs and ACEs and/or bypasses to ensure that a failure in part of the fly-by-wire aircraft control system does not cause failure of the entire fly-by-wire aircraft control system. For example, in a conventional fly-by-wire aircraft control system, the failure of one or more FCCs may provide for the bypass of the FCCs, in which the instruction signal corresponding to the input provided by the pilot may be sent directly to the corresponding actuator via one or more ACEs such that the fly-by-wire aircraft control system may be referred to as operating in “direct mode” or stick-to-surface mode.

However, the hardware, e.g., redundant FCCs, ACEs, sensors, and the associated buses created there between, added to provide a high level of fault tolerance in the fly-by-wire aircraft control system, has correspondingly added to the footprint, weight, and maintenance of the fly-by-wire aircraft control system. Such additional weight is highly undesirable in view of the extra fuel required by the aircraft. Further, the larger footprint provides for less room in the aircraft to be utilized for transporting cargo and/or people. Still yet, the additional hardware requires increased maintenance hours and an increased number of replaced components, which adds undesirable cost.

What is needed, then, is a fly-by-wire aircraft control system having a smaller footprint and providing for a reduction in weight and associated maintenance costs over conventional fly-by-wire aircraft control systems currently in use.

SUMMARY

Embodiments of the disclosure may provide a system for controlling aircraft flight control surfaces. The system may include at least three pilot sensor channels including a first channel containing a first set of pilot sensor data, a second channel containing a second set of pilot sensor data, and a third channel containing a third set of pilot sensor data. The system may also include at least three aircraft sensor channels including a primary channel containing a first set of aircraft sensor data, a secondary channel containing a second set of aircraft sensor data, and a tertiary channel containing a third set of aircraft sensor data. The system may further include a first actuator control component operatively connected to the at least three pilot sensor channels and the at least three aircraft sensor channels and configured to synchronously receive and vote on the pilot sensor data from each of the at least three pilot sensor channels and the aircraft sensor data from each of the at least three aircraft sensor channels, such that a voted output of the at least three pilot sensor channels is transmitted to a flight control computer and augmented before being transmitted to remote electronics units. The voted output of the at least three pilot sensor channels may provide for the control of the aircraft surfaces coupled to the remote electronics units.

Embodiments of the disclosure may further provide a method for controlling aircraft flight control surfaces. The method may include generating pilot control signals from each of a first, second, and third set of pilot control sensors, and generating aircraft sensor signals from each of a first, second, and third set of aircraft sensors. The method may also include synchronously transmitting the pilot control signals from the first, second, and third set of pilot control sensors to a first actuator control component via a respective first, second, and third channel operatively connected to the first actuator control component, and synchronously transmitting the aircraft sensor signals from the first, second, and third set of aircraft sensors to the first actuator control component via a respective primary, secondary, and tertiary channel operatively connected to the first actuator control component. The method may further include voting the first, second, and third channels such that a voted output of the pilot control sensors is generated, and voting the primary, secondary, and tertiary channels such that a voted output of the aircraft sensors is generated. The method may also include transmitting the voted output of the pilot control sensors or the voted output of the aircraft sensors to remote electronics units operatively connected to the aircraft flight control surfaces, such that the voted output of the pilot control sensors or the voted output of the aircraft sensors controls the aircraft flight control surfaces.

Embodiments of the disclosure may further provide a system for controlling aircraft flight control surfaces of an aircraft. The system may include a plurality of pilot controlled devices, each operatively connected to at least one of the aircraft control surfaces and configured to be manipulated by an action provided by a pilot of the aircraft. The system may also include a first plurality of pilot control sensors operatively connected to each of the plurality of pilot controlled devices, such that at least one of the first plurality of pilot control sensors is operatively connected to a corresponding one of the plurality of pilot controlled devices, and the at least one of the first plurality of pilot control sensors is configured to generate a first signal containing information related to the respective one of the plurality of pilot controlled devices. The system may further include a second plurality of pilot control sensors operatively connected to each of the plurality of pilot controlled devices, such that at least one of the second plurality of pilot control sensors is operatively connected to the corresponding one of the plurality of pilot controlled devices, and the at least one of the second plurality of pilot control sensors is configured to generate a second signal containing information related to the respective one of the plurality of pilot controlled devices. The system may also include a third plurality of pilot control sensors operatively connected to each of the plurality of pilot controlled devices, such that at least one of the third plurality of pilot control sensors is operatively connected to the corresponding one of the plurality of pilot controlled devices, and the at least one of the third plurality of pilot control sensors is configured to generate a third signal containing information related to the respective one of the plurality of pilot controlled devices. The system may further include a first, second, and third plurality of aircraft sensors, each aircraft sensor configured to detect information related to the aircraft or surrounding environment. The system may also include a first actuator control component. The first actuator control component may include a first, second, and third programmable device, each configured to synchronously receive and vote on the first, second, and third plurality of pilot control sensors to produce a pilot control sensors voted output, and each programmable device further configured to synchronously receive and vote on the first, second, and third plurality of aircraft sensors to produce an aircraft sensors voted output. The first actuator control component may also include a first monitor configured to receive each of the pilot control sensors voted outputs and perform a bit by bit compare to produce a direct mode voted output, and a second monitor configured to receive each of the aircraft sensors voted outputs and perform a bit by bit compare to produce a normal mode voted output. The system may further include a plurality of remote electronics units configured to receive either the direct mode voted output or the normal mode voted output from the first actuator control component, either voted output providing information relative to the control of the aircraft flight control surfaces operatively connected to the plurality of remote electronics units, such that the aircraft flight control surfaces may be controlled by either the direct mode voted output or the normal mode voted output received.

DETAILED DESCRIPTION

FIG. 1illustrates an exemplary fly-by-wire aircraft control system10according to an embodiment of this disclosure. The fly-by-wire aircraft control system10may include a plurality of pilot control sensors12operatively connected to corresponding pilot controlled devices14and a plurality of copilot control12asensors operatively connected to corresponding copilot controlled devices14a. In an exemplary embodiment, the pilot controlled devices14and the copilot controlled devices14aeach include a wheel16,16a, a control column18,18a, and a set of pedals20,20a. The pilot's set of pedals20and copilot's set of pedals20amay be mechanically interconnected such that the pilot and copilot pedals20,20aoperate in unison. In an exemplary embodiment, the pilot's wheel16and control column18and the copilot's wheel16aand control column18amay be connected by an override device17and configured such that in normal operation, the pilot and copilot wheels16,16aand control columns18,18aalso move in unison; however, as needed, the override device17may be manipulated such that the pilot wheel16and control column18and the copilot wheel16aand control column18amay move independently of the other. In addition, the pilot controlled devices14may include a speed brake lever22and a set of elevator feel actuators (not shown). It will be understood by one of ordinary skill in the art that the speed break lever22and the set of elevator feel actuators may be disposed between the pilot (not shown) and copilot (not shown) and further operated by the copilot as needed.

As will be further explained in the detailed description, in an exemplary embodiment, each of the pilot controlled devices14and copilot controlled devices14amay be operatively connected to one or more aircraft flight control surfaces24coupled to an aircraft26having a fuselage28, a first and second horizontal stabilizer30,30a, a vertical stabilizer32, and first and second wings34,34aas shown inFIG. 2. It will be appreciated by one of ordinary skill in the art that the configuration of the aircraft may vary, and may include, for example, a plurality of vertical stabilizers or a lack of vertical stabilizers. The aircraft flight control surfaces24may include primary aircraft flight control surfaces36and secondary aircraft flight control surfaces38. The primary aircraft flight control surfaces36may include an elevator40disposed on each horizontal stabilizer30,30a, a rudder42disposed on the vertical stabilizer32, and an aileron44disposed on each wing34,34a. The primary aircraft flight control surfaces36may control movement of the aircraft26in one or more of the three directional aircraft axes (roll axis R, pitch axis P, and yaw axis Y). For example, manipulation of the pilot's set of pedals20may cause the rudder42to move, which may be used to control movement of the aircraft26about the yaw axis Y, whereas movement of the pilot's and/or copilot's control column18,18amay control movement of the aircraft26about the roll axis R or pitch axis P depending on the movement of the control column18,18a, i.e., port/starboard or forward/aft, respectively. Specifically, movement of the control column18,18ain the forward/aft direction may cause the movement of the elevators40, thereby affecting movement about the pitch axis P of the aircraft26, while movement of the control column18,18ain the port/starboard direction may cause the movement of the ailerons44, thereby affecting movement of the aircraft26about the roll axis R. It will be appreciated by one of ordinary skill in the art that the control column18,18amay be moved in a combined forward-starboard direction, in a combined forward-port direction, in a combined aft-starboard direction, and in a combined aft-port direction.

The secondary aircraft flight control surfaces38may include a plurality of flaps46, slats48, and spoilers50disposed on each wing34,34a. The secondary aircraft flight control surfaces38may influence the lift and drag of the aircraft26. For example, manipulation of the pilot's speed brake lever22may cause one or more spoilers50to extend from the aircraft26, such that drag is increased and the speed of the aircraft26is decreased.

As shown inFIG. 1, the plurality of pilot and copilot control sensors12,12amay include three pilot control pitch sensors52a,b,cand three copilot control pitch sensors54a,b,c, each configured to detect the position of the aircraft26about the pitch axis P as controlled by manipulation of the respective pilot and copilot control columns18,18ain the forward/aft direction. The plurality of pilot and copilot sensors12,12amay also include three pilot control roll sensors56a,b,cand three copilot control roll sensors58a,b,c, each configured to detect the position of the aircraft26about the roll axis R as controlled by manipulation of the respective pilot and copilot control columns18,18ain the port/starboard direction.

The plurality of pilot sensors12may also include three pilot control yaw sensors60a,b,c, each configured to detect the position of the aircraft26about the yaw axis Y as controlled by manipulation of the pilot's pedals20. As noted above, in an embodiment, the pilot and copilot pedals20,20aare mechanically interconnected such that the pedals20,20amove in unison. Thus, movement of the pedals20,20aby either the pilot or copilot results in the same positioning of the aircraft26. It will be appreciated by one of ordinary skill in the art that additional yaw sensors may be employed in an embodiment including independent pilot and copilot pedals20,20a.

Further, the plurality of pilot sensors may include three pilot control speed brake sensors62a,b,c, each configured to detect the lift and/or drag of the aircraft26as controlled by the manipulation of the pilot's speed brake lever22. AlthoughFIG. 1illustrates the foregoing pilot and/or copilot control sensors12,12a, embodiments including other pilot and/or copilot control sensors known in the art are contemplated herein.

In an exemplary embodiment, the plurality of pilot and copilot sensors12,12amay include at least three sensors utilized to detect a certain aircraft information, characteristic, parameter, or status identifier, such as the aircraft's movement about the yaw axis Y, roll axis R, or pitch axis P. Utilizing at least three sensors to detect the aircraft characteristic, parameter, or status identifier allows for the data provided by the sensors to be included in a voting procedure, discussed in further detail below. The use of three pilot and/or copilot sensors12,12afor a particular aircraft characteristic allows for a high integrity command to be determined; however, embodiments in which two, or four or more pilot and copilot sensors12,12aare utilized to detect a particular aircraft characteristic or status component are contemplated herein, such that the data from the pilot and copilot sensors12,12amay be used to compare, average, vote, or select a mid-value to derive a high integrity output.

One or more of the plurality of pilot control sensors12may include pilot control transducers64. In an exemplary embodiment, each of the pilot control sensors12includes a corresponding pilot control transducer64. Correspondingly, one or more of the copilot controlled sensors12amay include copilot control transducers64a. In an exemplary embodiment, each of the copilot control sensors12aincludes a corresponding copilot control transducer64a. At least one of the pilot control transducers64and the copilot control transducers64amay be configured to generate one or more signals that are proportional to the movement and/or position of the respective pilot controlled device14and/or copilot controlled device14a. In at least one embodiment, the signals generated by the respective pilot and copilot control transducers64,64amay be analog signals. In an exemplary embodiment, the pilot control transducer64and/or the copilot control transducer64amay include a linear variable differential transformer (LVDT), a rotary variable differential transformer (RVDT), or a resolver; however, other electrical components capable of generating an electrical signal correlating to the change in position of the pilot controlled device and/or copilot controlled device are also contemplated herein. Further, in another embodiment, one or more of the pilot and copilot sensors12a,12amay be a respective pilot and copilot transducer64,64a.

In an exemplary embodiment, each of the three pilot control pitch sensors52a,b,cmay include a pilot control pitch transducer66a,b,csuch that the fly-by-wire aircraft control system may include a first pilot control pitch transducer66a, a second pilot control pitch transducer66b, and a third pilot control pitch transducer66c, each operatively connected to the pilot control column18and configured to generate a corresponding transducer signal related to the pitch axis P of the aircraft26depending on the movement of the pilot control column18in the forward/aft direction. Correspondingly, each of the three copilot control pitch sensors54a,b,cmay include a copilot control pitch transducer68a,b,csuch that the fly-by-wire aircraft control system10may include a first copilot control pitch transducer68a, a second copilot control pitch transducer68b, and a third copilot control pitch transducer68c, each operatively connected to the copilot control column18aand configured to generate a corresponding transducer signal related to the pitch axis P of the aircraft26depending on the movement of the copilot control column18ain the forward/aft direction.

In an exemplary embodiment, each of the three pilot control roll sensors56a,b,cmay include a pilot control roll transducer70a,b,csuch that the fly-by-wire aircraft control system10may include a first pilot control roll transducer70a, a second pilot control roll transducer70b, and a third pilot control roll transducer70c, each operatively connected to the pilot control column18and configured to generate a corresponding transducer signal related to the roll axis R of the aircraft26depending on the movement of the pilot control column18in the port/starboard direction. Correspondingly, each of the three copilot control roll sensors58a,b,cmay include a copilot control roll transducer72a,b,csuch that the fly-by-wire aircraft control system10may include a first copilot control roll transducer72a, a second copilot control roll transducer72b, and a third copilot control roll transducer72c, each operatively connected to the copilot control column18aand configured to generate a corresponding transducer signal related to the roll axis R of the aircraft26depending on the movement of the copilot control column18ain the port/starboard direction.

In an exemplary embodiment, each of the three pilot control yaw sensors60a,b,cmay include a pilot control yaw transducer74a,b,csuch that the fly-by-wire aircraft control system10may include a first pilot control yaw transducer74a, a second pilot control yaw transducer74b, and a third pilot control yaw transducer74c, each operatively connected to the pilot pedals20and configured to generate a corresponding transducer signal related to the yaw axis Y of the aircraft26depending on the movement of the pilot pedal20manipulated. It will be appreciated by one of ordinary skill in the art that the aircraft movement about the yaw axis Y may also be achieved by varying the thrust levels from the engines76a,76bon opposing sides of the aircraft26. Embodiments in which the yaw sensors60a,b,cand corresponding yaw transducers74a,b,care operatively connected to the throttle (not shown) are contemplated herein.

In an exemplary embodiment, each of the three pilot control speed brake sensors62a,b,cmay include a pilot control speed brake transducer78a,b,csuch that the fly-by-wire aircraft control system may include a first pilot control speed brake transducer78a, a second pilot control speed brake transducer78b, and a third pilot control speed brake transducer78c, each operatively connected to the pilot's speed brake lever22and configured to generate a corresponding transducer signal related to the lift and/or drag of the aircraft26depending on the amount of manipulation of the pilot's speed brake lever22. As noted above, embodiments in which the copilot controls the speed brake lever22and pedals20a, thereby affecting the lift and/or drag of the aircraft26and the yaw axis Y of the aircraft26, respectively, are contemplated herein.

The fly-by-wire aircraft control system10may include a plurality of actuator control components, referred to as actuator control electronics (ACEs)80a,80b. In an exemplary embodiment, the plurality of ACEs includes a first ACE80a, referred to as a first cockpit control data acquisition electronics component (CCDAE), and a second ACE80b, referred to as a second CCDAE. The first and second CCDAE80a,80bmay each be operatively coupled to the plurality of pilot and copilot control sensors12,12avia a plurality of lines or channels82a-f. Each channel may be configured to transmit the one or more transducer signals generated by the respective pilot and/or copilot transducer64,64a.

In an exemplary embodiment, each CCDAE80a,80bmay be operatively connected to the plurality of pilot and copilot sensors12,12avia three channels82a-f. The first CCDAE80amay be operatively connected to the plurality of pilot and copilot sensors12,12avia a first CCDAE first channel82a, a first CCDAE second channel82b, and a first CCDAE third channel82c. The first CCDAE first channel82amay be configured such that the one or more transducer signals generated by one or more of the first pilot and copilot control pitch transducers66a,68a, the first pilot and copilot control roll transducers70a,72a, the first pilot control yaw transducer74a, and the first pilot control speed brake transducer78amay be transmitted to the first CCDAE80a. The first CCDAE second channel82bmay be configured such that the one or more transducer signals generated by one or more of the second pilot and copilot control pitch transducers66b,68b, the second pilot and copilot control roll transducers70b,72b, the second pilot control yaw transducer74b, and the second pilot control speed brake transducer78bmay be transmitted to the first CCDAE80a. The first CCDAE third channel82cmay be configured such that the one or more transducer signals generated by one or more of the third pilot and copilot control pitch transducers66c,68c, the third pilot and copilot control roll transducers70c,72c, the third pilot control yaw transducer74c, and the third pilot control speed brake transducer78cmay be transmitted to the first CCDAE80a.

Correspondingly, the second CCDAE80bmay be operatively connected to the plurality of pilot and copilot sensors12,12avia a second CCDAE first channel82d, a second CCDAE second channel82e, and a second CCDAE third channel82f. The second CCDAE first channel82dmay be configured such that the one or more transducer signals generated by one or more of the first pilot and copilot control pitch transducers66a,68a, the first pilot and copilot control roll transducers70a,72a, the first pilot control yaw transducer74a, and the first pilot control speed brake transducer78amay be transmitted to the second CCDAE80b. The second CCDAE second channel82emay be configured such that the one or more transducer signals generated by one or more of the second pilot and copilot control pitch transducers66b,68b, the second pilot and copilot control roll transducers70b,72b, the second pilot control yaw transducer74b, and the second pilot control speed brake transducer78bmay be transmitted to the second CCDAE80b. The second CCDAE third channel82fmay be configured such that the one or more transducer signals generated by one or more of the third pilot and copilot control pitch transducers66c,68c, the third pilot and copilot control roll transducers70c,72c, the third pilot control yaw transducer74c, and the third pilot control speed brake transducer78cmay be transmitted to the second CCDAE80b.

In addition to receiving the transducer signals generated and transmitted by the pilot and copilot transducers64,64a, the first and second CCDAE80a,80beach receive transducer signals generated by a plurality of aircraft sensors84operatively connected to each CCDAE80a,80b. In an exemplary embodiment, the plurality of aircraft sensors84,86of the fly-by-wire aircraft control system10includes direct mode aircraft sensors84and normal mode aircraft sensors86. As illustrated inFIG. 1, each CCDAE80a,80bmay be operatively coupled to the direct mode aircraft sensors84via lines or channels88a-f. In an exemplary embodiment, channels88a-fmay form a digital communication link.

The direct mode aircraft sensors84may include roll rate, yaw rate, pitch rate sensors and the like. In an exemplary embodiment, the direct mode aircraft sensors84may be configured to supply discrete or digital inputs to each CCDAE80a,80b. Each CCDAE80a,80bmay include one or more circuits (not shown) configured to acquire the plurality of discrete inputs. In an exemplary embodiment, the discrete inputs may be acquired from trim switches, command switches, and the like. Further, one or more of the discrete inputs may be latched via a latch (not shown) upon initially supplying power to the CCDAEs80a,80b. In an embodiment in which the discrete input may be utilized for a dynamic response, a filtering component (not shown) may be applied to the discrete input. The time constant of the circuits configured to acquire the plurality of discrete inputs may be configured to have a time constant of five to ten milliseconds.

In an exemplary embodiment, the fly-by-wire aircraft control system10includes a plurality of direct mode aircraft sensors84including a first set of three direct mode aircraft sensors84a, a second set of three direct mode aircraft sensors84b, and a third set of three direct mode aircraft sensors84c. In an exemplary embodiment, the three direct mode aircraft sensors in each set84a,b,cdetect an identical property or characteristic of the aircraft26. For example, the three sensors of the first set84amay detect the roll rate, yaw rate, and pitch rate, respectively. Accordingly, the second and third set84b,84cof sensors may also detect the roll rate, yaw rate, and pitch rate, respectively. By doing so, the direct mode aircraft sensors84may undergo a voting process in each CCDAE80a,80bto determine a direct mode command, as will be discussed below in greater detail. Each set84a,b,cof direct mode aircraft sensors84may be operatively connected to each CCDAE80a,80bvia the respective three respective channels88a-c,88d-f, such that the first CCDAE80amay be configured to receive the signals generated by the first set84aof direct mode aircraft sensor via a first CCDAE direct mode aircraft sensor first channel88a, the signals generated by the second set84bof direct mode aircraft sensors via a first CCDAE direct mode aircraft sensor second channel88b, and the signals generated by the third set84cof direct mode aircraft sensors via a first CCDAE direct mode aircraft sensor third channel88c. Correspondingly, the second CCDAE80bmay be configured to receive the signals generated by the first set84aof direct mode aircraft sensor via a second CCDAE direct mode aircraft sensor first channel88d, the signals generated by the second set84bof direct mode aircraft sensors via a second CCDAE direct mode aircraft sensor second channel88e, and the signals generated by the third set84cof direct mode aircraft sensors via a second CCDAE direct mode aircraft sensor third channel88f.

Each CCDAE80a,80bmay be configured to be a line replaceable unit including at least one printed circuit board (PCB)90a,b,c. Each PCB90a,b,cmay include a plurality of programmable devices92a,b,c. In an exemplary embodiment, the programmable devices may be field programmable gate arrays (FPGAs)92a,b,coperatively connected to and disposed on each of the PCBs90a,b,c. In another embodiment, the programmable devices may be, for example, programmable logic devices (PLDs), application-specific integrated circuits (ASICs), or central processing units (CPUs). In an exemplary embodiment, each CCDAE80a,80bincludes three PCBs90a,b,c, each having a first FPGA92a, a second FPGA92b, and a third FPGA92c. The first, second, and third FPGA92a,b,cof the first CCDAE80amay be configured to be operatively connected to a respective one of the first, second, and third, first CCDAE channels82a,b,cand a respective one of the first, second, and third first CCDAE direct mode aircraft sensor channels88a,b,c.

In an exemplary embodiment, each CCDAE80a,80bmay include a plurality of demodulators (not shown) configured to recover the information content from each transducer signal generated by the pilot and copilot transducers64,64aand transmitted to the respective CCDAE80a,80bvia the respective CCDAE channel82a-f. Each CCDAE80a,80bmay also include a plurality of anti-alias filters (not shown), such that the demodulator outputs in each channel82a-fmay be fed through an anti-alias filter to reduce noise susceptibility and alias effects of sampling. In an exemplary embodiment, each demodulator output may have rigging electronic capability; however, the authority of rigging may be limited to a percentage of full scale based on the aircraft requirements.

Each FPGA92a,b,cin the respective CCDAE80a,80bis operatively coupled to one or more multiplexers94a,94b. In an exemplary embodiment, each transducer signal, e.g., pilot transducer signal or copilot transducer signal, having passed through the demodulator and anti-alias filter is fed into a first multiplexer94a, such that the transducer signals form a single line input. The digital or discrete inputs provided from the signals transmitted in the respective direct mode aircraft sensor channels88a-fmay be fed into a respective second multiplexer94bsuch that the discrete inputs form a separate single line input from the first multiplexer94a. The single line input formed from the transducer signals are fed into an analog to digital (A/D) converter96controlled by the respective FPGA92a,b,c. The resulting data output from the A/D converter96may be fed to the FPGA92a,b,cin addition to the single line input generated from the second multiplexer94band the discrete inputs.

In an exemplary embodiment, each FPGA92a,b,cof the first CCDAE80amay be configured to receive both the single line input generated from the transducer signals and the single line input generated from the discrete inputs for each of the other FPGAs92a,b,c. For example, the first FPGA92aof the first CCDAE80areceives the single line input generated from the transducer signals of the first CCDAE first channel82aand the single line input generated from the discrete inputs of the first CCDAE direct mode aircraft sensor first channel88a. In addition, the first FPGA92aof the first CCDAE80areceives the single line input generated from the transducer signals of the first CCDAE second channel82band the single line input generated from the discrete inputs of the first CCDAE direct mode aircraft sensor second channel88b. Further, the first FPGA92aof the first CCDAE80areceives the single line input generated from the transducer signals of the first CCDAE third channel82cand the single line input generated from the discrete inputs of the first CCADE direct mode aircraft sensor third channel88c. By providing each FPGA92a,b,cof the first CCDAE80awith the information from each channel82a-c,88a-c, a voting procedure may be implemented in each FPGA92a,b,c, which will be discussed in more detail below.

In an exemplary embodiment, each FPGA92a,b,coperatively connected to the respective channels82a-fprovides for the voting of the transducer sensors input into the CCDAE80a,80b. Each channel output is voted against the other channel outputs. Thus, in the three channel configuration of each CCDAE80a,80b, the loss of a channel does not cause the fly-by-wire aircraft control system10to enter direct mode control. The CCDAE80a,80bremains fail-operative. Such a voting procedure allows for a high integrity command, i.e., 10−9or greater, to be output. In an exemplary embodiment, the voting of the pilot and copilot sensors12,12aand the digital or discrete inputs is carried out synchronously. The ability to provide a voting procedure with synchronized inputs provides for the CCDAE output to be fail-operative, and also provides for the elimination of Byzantine failures. In order for the inputs of the respective FPGAs92a,b,cof each CCDAE80a,80bto be synchronized, the fly-by-wire aircraft control system may include a plurality of control systems or loops configured to synchronize the clock98of each FPGA92a,b,cand the inputs received via the respective channels82a-f,88a-f.

In an exemplary embodiment, the plurality of control system or loops includes a first, or inner, control system97or loop and a second, or outer, control system99or loop. The inner control loop97may be provided by a minor count generator100provided in each of the FPGAs92a,b,c. The minor count generator100of each FPGA92a,b,cmay receive minor counts generated for example at20MHz by each of the other FPGAs92a,b,c, and a first and second monitor102,104further discussed below. Each minor count generator100processes the minor counts received and generates a selected frequency, e.g. 1 MHz, pulse substantially synchronized, i.e. within ten or fewer nanoseconds, with the clocks98of the other FPGAs92a,b,c.

The outer control loop99may be provided by a task generator106provided in each of the FPGAs92a,b,c. The task generator106of each FPGA92a,b,cmay be configured to operate as an incremental counter according to a programmed logic. In an exemplary embodiment, the programmed logic may include the instructions such that if the counts of each channel are equal to the counts of one other channel, the counter is incremented. For example, if the first CCDAE first channel88ais equal to the count provide by first CCDAE second channel88b, the counter is incremented. Otherwise, the counter returns to zero. By doing so, the counts of each FPGA92a,b,cprovide for the synchronization of the three FPGAs92a,b,c. The counter in each task generator106may count up to 32 before recycling back to 1. Accordingly, each FPGA92a,b,cmay provide for a 32 bit cyclic redundancy check in an exemplary embodiment.

Each PCB90a,b,cmay include the first monitor102and the second monitor104introduced above. In an exemplary embodiment, each monitor102,104may be a logic device providing a bit by bit compare. Each monitor102,104may be configured to receive the voted output of each FPGA92a,b,c. The voted output of the each FPGA92a,b,cincludes the voted direct mode command and the voted normal mode command. The voted direct mode command may be sent to the first monitor via line108and the voted output of the normal mode command may be sent to the second monitor via line110.

In an exemplary embodiment, the voted output of the normal mode command may be sent to the FCCs112,114. An augmented normal mode command may be provide by the each FCC's flight control computer control system (not shown) based on the data received from the normal mode aircraft sensors86and flap/slat controller data provided by the flap/slat controller116. The augmented normal mode command may be passed back to the respective CCDAE80a,80bvia a bidirectional digital communication link118and sent to the second monitor104. The voted normal mode command or augmented normal mode command is monitored in the second monitor104to determine if the command is valid or corrupted. This output is a high integrity output with a bit by bit compare. In an exemplary embodiment, a switch120,122may be manipulated in the respective monitor102,104corresponding to the validity of the normal mode command. For example, the switch122may be oriented in the second monitor104in such a manner as to allow the normal mode command to be outputted to one or more remote electronics units (REUs)124a-hwhen the normal mode command is valid, and the corresponding switch120in the first monitor102may be oriented to prohibit the direct mode command from being outputted to the respective REUs124a-h. Correspondingly, if the normal mode command is determined to be corrupt or invalid, the orientation of the switches120,122of the respective monitors102,104may be reversed such that the normal mode command is prohibited from being outputted, and the direct mode command is allowed to be outputted to the respective REUs124a-h.

Each CCDAE80a,80bmay include a plurality of excitation components126a,b,c, each configured to provide excitation via an electrical current to one or more pilot and/or copilot transducers64,64asuch that data provided by the respective transducer64,64amay be transmitted to each CCDAE80a,80b. In an embodiment, the excitation component126a,b,cmay include a current generator. The excitation component126a,b,cmay include an independent buffer (not shown) and may be operatively connected to the transducers64,64asuch that a single fault does not affect the performance of the other transducers64,64a. The excitation component126a,b,cmay generate an electrical current having a frequency of about 1800 Hz to 2400 Hz; however, higher or lower frequencies are contemplated herein. In an exemplary embodiment, each CCDAE80a,80bmay include three excitation components126a,b,c, each excitation component operatively connected to the respective CCDAE first, second, and third channels82a,b,c. Each excitation component126a,b,cin the respective channel82a,b,cmay produce two synchronous electrical currents being 180 degrees out of phase with each other. In an embodiment, the amplitude of each electrical current may be controlled by a closed loop circuit as will be appreciated by one of ordinary skill in the art.

Conventionally, a fly-by-wire aircraft control system may require up to six ACEs to provide the appropriate availability as required by certain government regulations; however, in an aspect of the present disclosure, the fly-by-wire aircraft control system10may only employ two ACEs80a,80bto achieve the desired availability required. Further, the failure of a single ACE, or a single channel of the ACE, of the present disclosure does not prohibit the fly-by-wire aircraft control system10from operating in normal mode.

The first and second CCDAEs80a,80bmay each be supplied by a respective primary power supply128a,128b. In an exemplary embodiment, the primary power supply128a,128bmay have a voltage of 28 VDCand may be rated for approximately 7.5 Watts per channel; however, embodiments in which the voltage and power rating may be higher or lower are contemplated herein. The primary power supply128a,128bmay include a single filter (not shown) and inrush limiter (not shown). The primary power supply source128a,128bmay be permanent magnet generators coupled to the respective aircraft engine76a,76bconfigured to generate AC power. The AC may be converted into the 28 VDC. Secondary power supply sources (not shown) may include the aircraft's ram air turbine, the 28 VDCmain aircraft buses, and the aircraft hot battery bus. In an exemplary embodiment, transistor-transistor logic (TTL) may be added to select between the primary power source128a,128band the secondary power source. In such an embodiment, the current limited primary power source128a,128bmay be distributed to generate power. Each of the three channels82a-fin the first and second CCDAE80a,80bmay include switching regulators (not shown) to generate the voltages used by the components, i.e., digital and analog, in each of the CCDAEs80a,80b. In an exemplary embodiment, a secondary voltage may be generated off of the filtered and current limited primary power supply128a,128b. The secondary voltage may be utilized to monitor the power supply generated voltages.

As indicated above, the fly-by-wire aircraft control system may include a plurality of flight control computers (FCCs)112,114, each operatively coupled to the first CCDAE80aand the second CCDAE80bvia a digital data bus. In an exemplary embodiment, the digital data bus may be a bidirectional digital communication link118. Multiple FCCs112,114may be utilized in a redundant manner to increase the availability of the fly-by-wire aircraft control system10and to ensure safe operation of the aircraft26in case of failure of a single FCC112,114. In an exemplary embodiment, the fly-by-wire aircraft control system10may include a first FCC112and a second FCC114. As noted above, failure of one or more FCCs112,114may cause the aircraft26to switch from a normal mode operation to a direct mode operation in which the normal mode operation is bypassed. In an exemplary embodiment, each FCC112,114may include a plurality of processors (not shown), including a plurality of dissimilar processors to substantially reduce the occurrence of common mode and random failures.

In a conventional FCC, the FCC is operative to receive an operational instruction via a plurality of channels from the pilot via the pilot control sensors or an operation instruction from the copilot via the copilot control sensors. The FCC, in the known art, is configured to undertake a voting process to determine the validity of the signals received, wherein a label or word is scanned and monitored from the signals received from each channel or lane. Conventionally, a single computing lane may be declared the “master” lane, and that lane is responsible for transmitting all data onto the data buses for use by the ACEs and other airplane systems; however, each of the channels are simultaneously computing the same control law. The outputs of the channels are compared against each other. Correspondingly, each FCC output may be compared against the outputs of the other FCCs to determine the value output to the ACE and the related actuator. As disclosed above, in an exemplary embodiment of this disclosure, the monitoring and voting of the pilot and/or copilot control sensors12,12amay be carried out in each CCDAE80a,80bupstream of the plurality of FCCs112,114. Such architecture of an exemplary embodiment of the present disclosure provides for the reduction in the amount of data that must be transmitted on the associated data buses, which reduces the amount of wire required for data buses and provides for a reduction in weight of the fly-by-wire aircraft control system10.

As noted above, the fly-by-wire aircraft control system10may include a plurality of aircraft sensors84,86including direct mode aircraft sensors84and normal mode aircraft sensors86. The normal mode aircraft sensors86may form at least a portion of the inertial reference system (IRS) and may include accelerometers, compasses, magnetometers, clinometers, and the like. In an exemplary embodiment, the normal mode aircrafts sensors86may be operatively coupled to each FCC112,114. The normal mode aircraft sensors86may include pressure sensors, such as pitot tubes, positioning sensors, strain gauges, and heat gauges and sensors, such as total air temperature (TAT) probes. The normal mode aircraft sensors86may be configured to provide information related to the performance and/or status of the aircraft26and the surrounding environment to each FCC112,114. For example, the normal mode aircraft sensors86may include information related to the pressure, altitude, speed, heading, temperature, G-forces, and the like of the aircraft. In an exemplary embodiment, multiple normal mode aircraft sensors86may be employed to detect a defined performance or status characteristic of the aircraft26in a redundant manner to ensure safe operation of the aircraft26in case of failure of one or more of the normal mode aircraft sensors86.

Each FCC112,114may be operatively connected to the flap/slat controller116. Conventionally, the flap/slat controller may have a direct interface with a flat/slat lever at the pilot controls and excitation of additional sensors may be required to receive the signals carrying data regarding the flap/slat controller. In an exemplary embodiment, the flap/slat controller116may receive the required data from each of the CCDAEs80a,80bdue to the high availability, i.e., 10−9or greater, provided by each CCDAE80a,80b. Thus, additional excitation components may be reduced by providing the required data from each CCDAE80,80bto the flap/slat controller116via each FCC112,114as illustrated inFIG. 1.

In an exemplary embodiment, each FCC112,114may be configured to receive the output from each CCDAE80a,80band one or more information signals generated by and transmitted from each of the normal mode aircraft sensors86operatively coupled to each FCC112,114. Each FCC112,114may be further configured to compare the output received from each CCDAE80a,80bwith the information signals received from the normal mode aircraft sensors86to determine the appropriate output to return to each CCDAE80a,80b.

The fly-by-wire aircraft control system10may include a plurality of remote electronics units (REUs)124a-h. Each REU124a-hmay be operatively coupled to each of the CCDAEs80a,80bvia a plurality of digital buses130a,130b. In an exemplary embodiment, each REU124a-his operatively coupled to each of the CCDAEs80a,80bvia a first digital bus130aand a second digital bus130b. Each REU124a-hmay be operatively connected to the each of the FCCs112,114via a digital bus132and configured to allow a signal to be output to each FCC112,114via the digital bus132. The output signal provided by one or more REUs124a-hto the FCCs112,114may provide status information regarding the one or more REUs124a-h. In an exemplary embodiment, each CCDAE80a,80bsends each REU124a-hthe voted normal mode command if no corruption is found in the voted normal command. In an embodiment in which the voted normal mode command is found to be deficient, the direct mode command is transmitted to each REU124a-hfrom the respective CCDAE80a,80b. In an embodiment in which the direct mode command may be sent to each REU124a-h, one or more REUs124a-hmay output a signal to each CCDAE80a,80bin relation to the direct mode command. The outputted signal may include data corresponding to an acknowledgement of receipt of the direct mode command. In another embodiment, one or more REUs124a-hmay output one or more signals produced by the corresponding operatively coupled actuator(s)134a-hto the first and second monitors102,104of each CCDAE80a,80b. Such outputted signal(s) may be routed to each CCDAE80a,80bvia the digital bus132and the corresponding FCCs112,114.

Each REU124a-hmay include a digital to analog converter (D/A) (not shown) configured to receive the digital signal transmitted from the respective CCDAE80a,80band convert the digital signal to an analog signal. The converted analog signal may be transmitted via one or more relays (not shown) and amplified via one or more amplifiers (not shown) before being transmitted to a corresponding actuator134a-h. In an exemplary embodiment, each REU124a-his installed integrally with the corresponding actuator134a-h; however, embodiments in which the REU124a-his installed proximal to the corresponding actuator134a-hare contemplated herein as well.

As introduced above, the fly-by-wire aircraft control system10may include a plurality of actuators134a-h, each operatively coupled to the corresponding remote electronics unit124a-h. Each actuator134a-hmay be a hydraulic actuator controlled by an electric signal. In an exemplary embodiment, the electrical signal may be the analog signal received from the corresponding REU124a-h. The actuator134a-hmay transform the analog signal received from the REU124a-hinto motion, which may be used to manipulate a corresponding aircraft flight control surface24, which will be discussed further below. The actuator134a-hmay be a servo valve-controlled linear cylinder, a high speed rotary motor driving a reduction gear, or the like. In another embodiment, the actuator134a-hmay be an electromechanical actuator. The actuator134a-hmay further include a sensor, such as a potentiometer, configured to transmit a signal to the REU124a-hreporting the position of the actuator134a-h. The transmitted signal provides the REU124a-hwith a reference to determine when the actuator134a-hhas reached the desired position.

As stated above, one or more actuators134a-hmay be operatively coupled to the corresponding aircraft flight control surface24of the aircraft26. In an exemplary embodiment, one or more of the aircraft flight control surfaces24may be driven by two or three actuators134a-h, each powered by a separate REU124a-h. In an exemplary embodiment, the actuator134a-hmay be a servo valve-controlled linear cylinder configured to manipulate a piston, or ram, to move forward or aft to move a hinge arm connected to the aircraft flight control surface24.

Turning to the operation of the fly-by-wire aircraft control system10, exemplary operation of the fly-by-wire aircraft control system10embodied inFIGS. 1 and 2is presented now. In such an exemplary operation, an excitation is provided by the respective excitation components126a,b,cin the first CCDAE80aand the second CCDAE80bto the plurality of pilot and copilot sensors12,12a. Each CCDAE80a,80bmay be substantially similar in function to the other CCDAE; thus, like numerals correspond to like elements and the operation of only one CCDAE80awill be detailed for the sake of brevity. It will be appreciated by one of ordinary skill in the art that the following description of the operation of the first CCDAE80ais applicable to the second CCDAE80b. Such excitation may be provided to the pilot and copilot control sensors12,12aincluding pilot and copilot control transducers64,64avia an electrical current. The pilot and copilot control transducers64,64amay be transformers such that the electrical current provides excitation and the generation of a transducer signal from each of the plurality of pilot and copilot control transducers64,64a.

As shown inFIG. 1, the first CCDAE80amay receive, via the first CCDAE first channel82a, the transducer signals generated by the first pilot and copilot control roll transducers70a,72a, the transducer signals generated by the first pilot and copilot control pitch transducers66a,68a, the transducer signal generated by the first pilot control yaw transducer74a, and the transducer signal generated by the first pilot control speed brake transducer78a. Correspondingly, the first CCDAE80amay receive, via the first CCDAE second channel82b, the transducer signals generated by the second pilot and copilot control roll transducers70b,72b, the transducer signals generated by the second pilot and copilot control pitch transducers66b,68b, the transducer signal generated by the second pilot control yaw transducer74b, and the transducer signal generated by the second pilot control speed brake transducer78b. Further, the first CCDAE80amay receive, via the first CCDAE third channel82c, the transducer signals generated by the third pilot and copilot control roll transducers70c,72c, the transducer signals generated by the third pilot and copilot control pitch transducers66c,68c, the transducer signal generated by the third pilot control yaw transducer signal74c, and the transducer signal generated by the third pilot control speed brake transducer78c.

The first CCDAE80amay further receive, via the direct mode aircraft sensor first channel88a, the respective signals generated from each direct mode aircraft sensor84of the first set84aof three direct mode aircraft sensors. Correspondingly, the first CCDAE80amay receive, via the direct mode aircraft sensor second channel88b, the respective signals generated from each direct mode aircraft sensor84of the second set84bof three direct mode aircraft sensors. Further, the first CCDAE80amay receive, via the direct mode aircraft sensor third channel88c, the respective signals generated from each direct mode aircraft sensor84of the third set84cof three direct mode aircraft sensors. As noted above, each direct mode aircraft sensor84may be a digital signal or may form a discrete input after being fed through circuitry provided by the first CCDAE80a.

In an exemplary embodiment, the first CCDAE80aincludes a first FPGA92a, a second FPGA92b, and a third FPGA92c, and each of the three FPGAs92a,b,cmay be coupled to a first and second multiplexer94a,94b. The transducer signals transmitted via the first CCDAE first channel82amay be demodulated and filtered before being fed into the first multiplexer94acoupled to the first FPGA92ato form a primary first channel single line input. The primary first channel single line input may be fed into the A/D converter96coupled to the first multiplexer94a. The digital or discrete inputs transmitted via the direct mode aircraft sensor first channel88amay be fed into the second multiplexer94bcoupled to the first FPGA92ato form a secondary first channel single line input. The primary first channel single line input and the secondary first channel single line input may be fed into the first FPGA92a. The primary first channel single line input and the secondary first channel single line input may then be split such that a portion of the primary first channel single line input and the secondary first channel single line input may be fed to the second FPGA92band a portion of the primary first channel single line input and the secondary first channel single line input may be fed to the third FPGA92c.

The transducer signals transmitted via the first CCDAE second channel82bmay be demodulated and filtered before being fed into the first multiplexer94acoupled to the second FPGA92bto form a primary second channel single line input. The primary second channel single line input may be fed into the A/D converter96coupled to the first multiplexer94a. The digital or discrete inputs transmitted via the direct mode aircraft sensor second channel88bmay be fed into the second multiplexer94bcoupled to the second FPGA92bto form a secondary second channel single line input. The primary second channel single line input and the secondary second channel single line input may be fed into the second FPGA92b. The primary second channel single line input and the secondary second channel single line input may then be split such that a portion of the primary second channel single line input and the secondary second channel single line input may be fed to the first FPGA92aand a portion of the primary second channel single line input and the secondary second channel single line input may be fed to the third FPGA92c.

The transducer signals transmitted via the first CCDAE third channel82cmay be demodulated and filtered before being fed into the first multiplexer94acoupled to the third FPGA92cto form a primary third channel single line input. The primary third channel single line input may be fed into the A/D converter96coupled to the first multiplexer94a. The digital or discrete inputs transmitted via the direct mode aircraft sensor third channel88cmay be fed into the second multiplexer coupled to the third FPGA92cto form a secondary third channel single line input. The primary third channel single line input and the secondary third channel single line input may be fed into the third FPGA92c. The primary third channel single line input and the secondary third channel single line input may then be split such that a portion of the primary third channel single line input and the secondary third channel single line input may be fed to the first FPGA92aand a portion of the primary third channel single line input and the secondary third channel single line input may be fed to the second FPGA92b.

The three CCDAE channels82a-cand the three direct mode aircraft sensor channels88a-cmay be fed to the respective first, second, and third FPGAs92a,b,cin a synchronized fashion as a result of the inner and outer control loops97,99provided. The inner clock loop97may be provided by a minor count generator100provided in each of the FPGAs92a,b,c. The minor count generator100of each FPGA92a,b,cmay receive minor counts generated for example at20MHz by the other FPGAs92a,b,c, and the first and second monitors102,104. Each minor count generator100processes the minor counts received and generates a selected frequency pulse, e.g. 1 MHz, substantially synchronized, i.e. within ten or fewer nanoseconds, with the clocks98of the other FPGAs92a,b,c.

The outer control loop99may be provided by a task generator106provided in each of the FPGAs92a,b,c. The task generator106of each FPGA92a,b,cmay be configured to operate as an incremental counter according to a programmed logic. In an exemplary embodiment, the programmed logic may include the instructions such that if the counts of each channel are equal to the counts of one other channel, the counter is incremented. For example, if the first CCDAE first channel88ais equal to the count provide by first CCDAE second channel88b, the counter is incremented. Otherwise, the counter returns to zero. The counter in each task generator may count up to 32 before recycling back to 1. Accordingly, the FPGA92a,b,cmay provide for a 32 bit cyclic redundancy check in an exemplary embodiment.

The synchronized three CCDAE channels82a-cfed into each FPGA92a,b,cmay undergo a voting procedure in the corresponding FPGA92a,b,cto determine a high integrity direct mode command output. The synchronized three direct mode aircraft sensor channels88a-cfed into each FPGA92a,b,cmay also undergo a voting procedure in the corresponding FPGA92a,b,cto determine a high integrity normal mode command output. Each direct mode command output of the respective FPGA92a,b,cmay be fed to the first monitor via line108, wherein a bit by bit compare is performed on the three direct mode outputs to determine a high integrity direct mode command output.

The high integrity normal mode command of each FPGA92a,b,cmay be sent to each FCC112,114, wherein the normal mode sensor data and the flap/slat controller data provided by the normal mode aircraft sensors86and flap/slat controller116operatively connected to each FCC112,114may provide the FCC control system with data correlating to the production of an augmented normal mode command from the normal mode command initially received from the first CCDAE80a. The augmented normal mode command may be sent back to the first CCDAE80avia the bidirectional data bus118where the augmented normal mode command may undergo a bit by bit compare in the second monitor104.

In an exemplary embodiment, a switch120,122may be manipulated in the respective monitor102,104corresponding to the validity of the normal mode command or augmented normal mode command. For example, a switch122may be oriented in the second monitor104in such a manner as to allow the normal mode command or augmented normal mode command to be outputted to the respective REUs124a-hwhen the normal mode command or augmented normal mode command is valid, and a corresponding switch120in the first monitor102may be oriented to prohibit the direct mode command from being outputted to the respective REUs124a-h. Correspondingly, if the normal mode command or augmented normal mode command is determined to be corrupt or invalid, the orientation of the switches120,122of the respective monitors102,104may be reversed such that the normal mode command or augmented normal mode command is prohibited from being outputted, and the direct mode command is allowed to be outputted to the respective REUs124a-h.

As stated above, if determined to be valid by the first CCDAE80a, the normal mode command or augmented normal mode command may be transmitted to the respective REUs124a-h; otherwise, the direct mode command may be transmitted to the corresponding REUs124a-h. The direct or normal mode command provided instructs the REUs124a-hon the positioning of the appropriate aircraft flight control surfaces24. Each REU124a-hmay include a digital to analog converter (D/A) configured to receive the digital signal transmitted from the first CCDAE80aand convert the digital signal to an analog signal. The converted analog signal may be transmitted via one or more relays and amplified via one or more amplifiers before being transmitted to the corresponding actuator134a-h.

Each actuator134a-hmay be a hydraulic actuator controlled by an electric signal. In an exemplary embodiment, the electrical signal may be the analog signal received from the corresponding REU124a-h. The actuator134a-hmay transform the analog signal received from the REU124a-hinto motion, which may be used to manipulate a corresponding aircraft flight control surface24. The actuator134a-hmay further include a sensor, such as a potentiometer, configured to transmit a signal to the REU124a-hreporting the position of the actuator134a-h. The transmitted signal provides the REU124a-hwith a reference to determine when the actuator134a-hhas reached the desired position. In an exemplary embodiment, the actuator134a-hmay be a servo valve-controlled linear cylinder configured to manipulate a piston, or ram, to move forward or aft to move a hinge arm connected to the aircraft flight control surface24.

In another embodiment illustrated inFIG. 3, an exemplary method200for controlling one or more aircraft flight control surfaces of an aircraft is provided. The method200may include generating pilot control signals from each of a first, second, and third set of pilot control sensors, as at202. The method may also include generating aircraft sensor signals from each of a first, second, and third set of aircraft sensors, as at204.

The method may further include synchronously transmitting the pilot control signals from the first, second, and third set of pilot control sensors to a first actuator control component via a respective first, second, and third channel operatively connected to the first actuator control component, as at206, and synchronously transmitting the aircraft sensor signals from the first, second, and third set of aircraft sensors to the first actuator control component via a respective primary, secondary, and tertiary channel operatively connected to the first actuator control component, as at208.

The first actuator control component may include a first programmable device, a second programmable device, and a third programmable device, each of the first, second, and third programmable devices being operatively connected to the first, second, and third channels and the primary, secondary, and tertiary channels, and each of the first, second, and third programmable devices being configured to vote on the first, second, and third channels and the primary, secondary, and tertiary channels, such that the first, second, and third programmable devices produce a respective first, second, and third pilot control sensor voted output and a respective first, second, and third aircraft sensor voted output. A clock of each programmable device may be synchronized with a respective clock of the other programmable devices.

The method may also include voting the first, second, and third channels such that a voted output of the pilot control sensors is generated, as at210, and voting the primary, secondary, and tertiary channels such that a voted output of the aircraft sensors is generated, as at212. The method may further include transmitting the voted output of the pilot control sensors or the voted output of the aircraft sensors to remote electronics units operatively connected to the aircraft flight control surfaces, such that the voted output of the pilot control sensors or the voted output of the aircraft sensors controls the aircraft flight control surfaces, as at214.