Patent Publication Number: US-2009222148-A1

Title: Autonomous Outer Loop Control of Man-Rated Fly-By-Wire Aircraft

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
     This is a continuation-in-part of U.S. patent application Ser. No. 11/425,600 filed on Jun. 21, 2006, the content of which is relied upon and incorporated herein by reference in its entirety, and the benefit of priority under 35 U.S.C. §120 is hereby claimed. 
    
    
     BACKGROUND OF THE INVENTION 
     1. Field of the Invention 
     The present invention relates generally to flight control systems, and particularly to fly-by-wire flight control systems for unmanned airborne vehicles (UAVs). 
     2. Technical Background 
     The market for UAVs is growing and is in the range of several billion dollars per year. UAVs may be used for many purposes including aerial surveillance, weapons delivery, and target training. Many UAVs are used as target drones by providing military pilots with realistic, high performance targets during airborne training. Irregardless of the use, one method for making a UAV is by converting a retired man-rated aircraft into an unmanned vehicle that is remote controlled or preprogrammed to follow a predetermined trajectory. The process of conversion typically involves modifying the retired aircraft&#39;s flight control system. A discussion of basic aircraft terminology may be useful before presenting some of the conventional approaches for converting retired aircraft into target drones. 
     Note that a typical aircraft includes a fuselage, wings, one or more engines, and a tail section that includes horizontal stabilizers and a vertical stabilizer. The engines generate the thrust that drives the aircraft forward and the wings provide the lift necessary for the aircraft to become airborne. Control surfaces are disposed on the wings, the horizontal stabilizers and the vertical stabilizer. The control surfaces enable the aircraft to respond to the flight control system command inputs provided by the pilot(s) by directing air flow in a controlled manner. The major control surfaces disposed on the typical aircraft are the ailerons, the elevators, and the rudder. 
     The ailerons are disposed on the trailing edges of the wings and are used to control the roll of the aircraft. Roll refers to the tendency of the aircraft to rotate about the aircraft&#39;s central longitudinal axis. If the pilot moves the control stick (or alternatively the control wheel) to the left, the left aileron will rise and the right aileron will fall and the aircraft will begin rolling to the port side. In like manner, if the control stick is moved to the right, the aircraft will roll to the starboard side. The elevators are disposed on the rear edges of the horizontal stabilizers or on the entire horizontal stabilizer and are used to control the aircraft pitch. Pitch refers to the tendency of the aircraft to rotate around the transverse axis of the aircraft. For example, if the pilot adjusts the control stick aft, the elevators will cause the nose to pitch upward and the aircraft will tend to lose airspeed. If the stick is moved foreword, the nose of the aircraft pitches downward. 
     The rudder is disposed on the vertical stabilizer and is usually employed to adjust the yaw of the aircraft. The yaw is the tendency of the aircraft to rotate around the vertical axis, i.e., the axis normal to the longitudinal axis and the transverse axis. The rudder is typically controlled by a pair of foot-operated pedals. 
     The aircraft may also include secondary control surfaces such as spoilers, flaps, and slats. The spoilers are also located on the wings and are employed for a variety of functions. The flaps and the slats are also disposed on the wing and are typically used to adjust the aircraft&#39;s lift and drag during landing and take off. As noted above, the means for transmitting the pilot&#39;s commands to the above described control surfaces is commonly referred to as the flight control system. 
     In the description provided above, the most common control surfaces were discussed. However, those of ordinary skill in the art will understand that aircraft may employ other such control surfaces such as flaperons, elevons, ruddervators, and thrust vectoring nozzles to name a few. A flaperon is a combination flap and aileron and is used, for example, on the F-16. An elevon is a combination elevator and aileron and is used on flying wing aircraft and delta-wing aircraft such as the B-2, F-106, B-58, etc. The ruddervator is a combination of the rudder and the elevator and is used, for example, on the F-117. The F-22 also employs a specialized control surface known as a thrust vectoring nozzle in addition to the horizontal stabilizer. 
     The flight control system is designed to actuate the control surfaces of the aircraft, allowing the pilot to fly the aircraft. The flight control system is, therefore, the control linkage disposed between the control input mechanisms, i.e., the control stick, pedals and the like, and the control surface actuator devices. One criteria of flight control system design relates to the aircraft&#39;s handling characteristics. The flight control system is also designed and implemented in accordance with certain specifications that ensure a very high level of reliability, redundancy and safety. These issues are especially important for man-rated aircraft, i.e., those that are to be flown by a pilot, and carry aircrew or passengers. The system&#39;s reliability and redundancy ensures that there is a very low probability of failure and the resulting loss of the aircraft and life due to a control system malfunction. All of these factors ensure that the airplane can be operated safety with a minimum risk to human life. 
     In older aircraft, the control stick and the pedals are coupled to the control surfaces by a direct mechanical linkage. The pilot&#39;s commands are mechanically or hydraulically transferred to the control surface. The pilot&#39;s control inputs are connected to hydraulic actuator systems that move the control surfaces by a system of cables and/or pushrods. In recent years, aircraft having flight control systems featuring direct mechanical linkages have been replaced by newer aircraft that are equipped with an electrical linkage system commonly referred to as a fly-by-wire system. 
     A fly-by-wire system translates the pilot&#39;s commands into electrical signals by transducers coupled to the control stick and the pedals. The electrical signals are interpreted by redundant flight control computers. Thus, the flight control system performs multiple digital or analog processes that combine the pilot&#39;s inputs with the measurements of the aircraft&#39;s movements (from its sensors) to determine how to direct the control surfaces. The commands are typically directed to redundant control surface actuators. The control surface actuators control the hydraulic systems that physically move the control surface of the aircraft. 
     After a man-rated aircraft is retired, it may be re-used for airborne missions that do not require a pilot or on-board crew. This type of aircraft, known as an Unmanned Air Vehicle (UAV) or Target Drone is modified to take advantage of the existing systems by replacing the functionality typically provided by a pilot. The flight control system may be changed in order to allow control by a ground controller. Alternatively, conversion is implemented by modifying flight control processor logic to merge external sensor signals and commands into the control surface commands that drive the UAV. 
     Currently, the primary aircraft employed for full-scale target missions is the F-4 Phantom fighter aircraft, which is a 1960&#39;s vintage aircraft. Retired F-4 Phantom aircraft have been used as target drones for several years. Approximately 5,000 F-4s were produced over the years. Unfortunately, the fleet of available F-4 aircraft is dwindling and the supply of F-4 aircraft will soon be depleted. This problem may be solved by pressing newer retired fly-by-wire aircraft (such as the F-16 or F-18) into service to meet the demand for target drones. However, it must be noted that the F-4 Phantom is not a fly-by-wire system. The F-4 is equipped with an older hydro-mechanical flight control system. Accordingly, different technological means are required to convert the newer fly-by-wire aircraft into target drones. 
     In one approach, fly-by-wire conversion methods requiring flight control computer re-programming are being considered. In another approach that is being considered, the flight control computer is removed altogether and replaced with a new computer. The new computer is programmed to perform the functions normally performed by the pilot, in addition to the traditional flight control system functions. However, both of these approaches have their drawbacks. Reprogramming or replacing the original man-rated flight control processor is a complex and costly proposition. The new flight control processor has to pass many, if not all, of the aircraft development tests originally required. The fact that most of the fly-by-wire aircraft expected to be used for this application are now more than 20 years old further complicates matters. The designers of the new replacement systems are faced with replicating the original system&#39;s functions and capabilities without having the necessary documentation. The system design and test definitions for these functions have been lost over time. 
     Accordingly, the effort required to replicate and prove a replacement system having identical fit/form/function and repeat the required development testing has been found to be prohibitively expensive. What is needed is an alternative, and less expensive, method for converting retired fly-by-wire aircraft into UAVs and/or target drones. 
     SUMMARY OF THE INVENTION 
     The present invention addresses the needs described above by providing a system and method for converting a fly-by-wire aircraft into a UAV. 
     One aspect of the present invention is directed to a system for converting a man-rated fly-by-wire (FBW) aircraft into a remote controlled unmanned airborne vehicle (UAV). The FBW aircraft includes a FBW flight control system (FBW-FCS) configured to control aircraft control surfaces disposed on the aircraft. The system includes a controller coupled to the FBW aircraft. The controller is configured to generate substantially real-time pilot control data from at least one aircraft maneuver command. The real-time pilot control data is generated in accordance with a predetermined control law. The at least one aircraft maneuver command is derived from at least one command telemetry signal received from a remote control system not disposed on the FBW aircraft or from a pre-programmed trajectory. An FBW-FCS interface system is coupled to the controller. The FBW-FCS interface system is configured to convert the substantially real-time pilot control data into substantially real-time simulated FBW-FCS pilot control signals. The substantially real-time simulated FBW-FCS pilot control signals are configured to direct the FBW-FCS such that the FBW aircraft performs in accordance with the at least one aircraft maneuver command. 
     In another aspect, the present invention is directed to a method for converting a man-rated fly-by-wire (FBW) aircraft into a remote controlled unmanned airborne vehicle (UAV). The FBW aircraft includes a FBW flight control system (FBW-FCS) configured to control aircraft control surfaces disposed on the aircraft. The method includes decoupling existing pilot controls from the FBW-FCS. An embedded control system is coupled to the FBW aircraft and the FBW-FCS. The embedded system includes a controller configured to generate substantially real-time pilot control data from at least one aircraft maneuver command. The real-time pilot control data is generated in accordance with a predetermined control law. The at least one aircraft maneuver command is derived from at least one command telemetry signal received from a remote control system not disposed on the FBW aircraft or from a pre-programmed trajectory. An FBW-FCS interface system is coupled to the controller. The FBW-FCS interface system is configured to convert the substantially real-time pilot control data into substantially real-time simulated FBW-FCS pilot control signals. The substantially real-time simulated FBW-FCS pilot control signals are configured to direct the FBW-FCS such that the FBW aircraft performs in accordance with the at least one aircraft maneuver command. 
     Additional features and advantages of the invention will be set forth in the detailed description which follows, and in part will be readily apparent to those skilled in the art from that description or recognized by practicing the invention as described herein, including the detailed description which follows, the claims, as well as the appended drawings. 
     It is to be understood that both the foregoing general description and the following detailed description are merely exemplary of the invention, and are intended to provide an overview or framework for understanding the nature and character of the invention as it is claimed. The accompanying drawings are included to provide a further understanding of the invention, and are incorporated in and constitute a part of this specification. The drawings illustrate various embodiments of the invention, and together with the description serve to explain the principles and operation of the invention. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS  
         FIG. 1  is a block diagram of an airborne control system in accordance with one embodiment of the present invention; 
         FIG. 2  is a schematic diagram illustrating the disposition of outer loop control processor (OLCP) within the UAV; 
         FIG. 3  is a perspective view of the OLCP enclosure in accordance with the present invention; 
         FIG. 4  is a hardware block diagram of the OLCP in accordance with an embodiment of the present invention; 
         FIG. 5  is a diagram illustrating the OLCP control system architecture in accordance with the present invention; 
         FIG. 6  is a flow chart illustrating the software control of the OLCP; 
         FIGS. 7A-7B  are diagrammatic depictions of the FBW interface circuit shown in  FIG. 3  in accordance with an embodiment of the present invention; 
         FIG. 8  is a detailed schematic of the quadrature multiplier circuit shown in  FIG. 7 ; 
         FIG. 9  is a detailed schematic of the power bus provided by the power supply depicted in  FIG. 3 ; 
         FIGS. 10A-10C  are voltage waveforms provided by the DACs shown in  FIG. 7  and  FIG. 15 ; 
         FIG. 11  is an example of a time varying voltage waveform in accordance with  FIGS. 10A-10C ; 
         FIG. 12  is an AC reference voltage signal in accordance with the embodiments depicted in  FIG. 7  and  FIG. 15 ; 
         FIGS. 13A-13C  are command voltage waveforms provided to the existing fly-by-wire aircraft in accordance with an embodiment of the invention; 
         FIG. 14  is an example of a time varying command voltage waveform in accordance with  FIGS. 13A-13C ; and 
         FIG. 15  is detailed block diagram of the FBW interface circuit depicted in  FIG. 3  in accordance with yet another embodiment of the present invention. 
     
    
    
     DETAILED DESCRIPTION  
     Reference will now be made in detail to the present exemplary embodiments of the invention, examples of which are illustrated in the accompanying drawings. Wherever possible, the same reference numbers will be used throughout the drawings to refer to the same or like parts. An exemplary embodiment of the system of the present invention is shown in  FIG. 1 , and is designated generally throughout by reference numeral  10 . 
     As embodied herein, and depicted in  FIG. 1 , a block diagram of a UAV control system  10  in accordance with one embodiment of the present invention is disclosed. The system  10  includes an outer loop control platform (OLCP)  20  disposed on an airborne platform, and a ground control system (GCS)  30 . Those of ordinary skill in the art will understand that GCS  30  may also be implemented on an airborne platform depending on mission requirements. 
     Although not shown in  FIG. 1 , GCS  30  typically includes communications and telemetry systems that are adapted to communicate with the communications and telemetry systems disposed aboard the aircraft. The GCS telemetry system is coupled to a processing system that is programmed to format GCS operator commands in accordance with both the telemetry system requirements and the aircraft requirements. The processing system is coupled to an operator I/O system and an operator display. 
     In one embodiment, the operator I/O provides the processor with input control signals that are substantially identical to the signals generated by cockpit control devices, such as the pitch/roll sticks, pedals, engine thrust control, etc., that are disposed in the aircraft. For example, if the UAV is a converted F-16 fighter aircraft, the processor in GCS  30  is programmed to provide GCS  30  telemetry/communication system with compatible signals. These commands are provided to the communication/telemetry systems  32  and transmitted to OLCP  20 . This is described herein as the “joystick” method. 
     In another embodiment, the GCS  30  operator I/O provides the operator with various maneuver options, such as turn, roll, etc. Of course, this GCS implementation is much easier to implement. In fact, the operator may transmit maneuver commands to the GCS command telemetry system via a personal computer or a laptop computer. The maneuver commands are transmitted to the UAV command telemetry unit, and OLCP  20  translates the maneuver commands appropriately. 
     In yet another embodiment, OLCP  20  maneuvers in accordance with a preprogrammed flight trajectory. For example, OLCP  20  programming may direct the FBW aircraft to follow and repeat a certain flight path at a predetermined airspeed and altitude. In this case, GCS  30  does not have to provide moment-to-moment control of the UAV. However, GCS  30  may reprogram OLCP  20  by way of the command telemetry uplink and direct OLCP  20  to follow a new trajectory. This feature of the present invention may be very beneficial during surveillance missions or weapons delivery missions. 
     Regardless of the type of GCS  30  employed to control the UAV, OLCP  20  processes these commands on a real-time basis to fly the aircraft, i.e., use the existing fly-by-wire flight control system, avionics, and other existing aircraft systems in accordance with operator commands. OLCP  20  provides the existing fly-by-wire flight control system (FBW-FCS) with pseudo pitch stick commands, roll stick commands, and rudder pedal commands in accordance with GCS  30  instructions. 
     The present invention also includes an electromechanical throttle actuator  22  that is electrically coupled to OLCP  20 . Throttle actuator  22  is disposed and mounted in the cockpit, and mechanically coupled to the existing aircraft throttle. Throttle actuator  22  receives scaled and calibrated servo control signals from OLCP  20  and physically manipulates the existing throttle mechanism in response thereto. 
     OLCP  20  may also be equipped, coupled to, or used in conjunction with, with one or more digital or analog cameras  24 . Digital cameras  24  may be disposed within the aircraft canopy to obtain a “cockpit view” of the UAV. OLCP  20  transmits aircraft navigational data, altitude, aircraft attitude data, and video (when so equipped) to GBCS  30 . This information may be displayed on a GCS  30  display for the benefit of the operator/pilot that is “flying” the UAV via GCS  30 . 
       FIG. 2  is a schematic diagram that illustrates the disposition of OLCP  20  within the UAV. Before the aircraft is converted into a UAV, the existing FBW-FCS is coupled to the existing pilot controls by way of redundant electrical interfaces. The present invention takes advantage of this arrangement by decoupling the cockpit pilot controls from the FBW-FCS, and replacing them with OLCP  20 . The present invention is also equipped with means for overriding the OLCP inputs. The overriding means are employed by an on-board safety pilot during developmental testing of the FBW aircraft or during other such manual operation of the FBW aircraft. OLCP  20  is also electrically coupled to existing aircraft landing gear interfaces, communications and telemetry interfaces, and existing avionics. OLCP  20  may also be coupled to a flight termination system and a scoring system developed for existing drone systems. OLCP  20  is configured to transmit and receive both analog and digital data in accordance with the existing electrical interfaces deployed in the aircraft. Once OLCP  20  is programmed and configured for deployment on a given fly-by-wire airborne platform, it is easily installed by connecting OLCP  20  to existing aircraft systems by way of signal cable interfaces  26 . OLCP  20  may be coupled to existing avionics by way of redundant high speed serial data bus interfaces  28 . As noted previously, OLCP  20  is coupled to the existing throttle via an electromechanical actuator  22 . 
     Although a single OLCP  20  is shown in  FIG. 2 , the present invention typically employs multiple-redundant systems for safety and reliability. Those skilled in the art will understand that redundant systems may be implemented by using a single OLCP that includes multiple processing channels or multiple OLCPs  20 , each having a single processing channel. When redundant systems are employed, the system includes a voting algorithm that selects an appropriate channel output. 
     As embodied herein and depicted in  FIG. 3 , a simplified hardware block diagram of the OLCP  20  in accordance with one embodiment of the present invention is disclosed. Again, OLCP  20  typically includes redundant processing channels for reliability and safety reasons.  FIG. 3  shows a single channel embodiment for clarity of illustration. 
     OLCS  20  is implemented as an embedded processor system  200  that includes I/O circuits  202 , embedded processor  204 , memory  206 , high speed serial data bus interface (I/F) circuits  210 , fly-by-wire interface (FBW I/F) circuits  212 , throttle interface circuit  214 , landing gear interface  216 , and OLCP sensor package  218  coupled to bus  220 . System  200  also includes power supply  222 . System  200  is also shown to include video processor circuit  208 . The video processor is configured to process the data provided by digital camera  24 . On the other hand, those of ordinary skill in the art will understand that the video system may be implemented using an existing video system and be deployed in the UAV as a separate stand-alone unit. 
     Further, any suitable communications/telemetry unit, scoring system, and flight termination equipments may be employed by the present invention. The command telemetry system may be implemented with off-the-shelf equipment developed for existing drone systems or custom designed equipment, depending on the UAV implementation. As those skilled in the relevant arts will understand, the communications and telemetry equipment employs a high speed radio link having the signal bandwidth to support OLCP  20  functionality. In any event, the design and implementation of I/O circuitry  202  is a function of the command telemetry system disposed on the aircraft and is considered to be within the abilities of one of ordinary skill in the art. 
     In one embodiment, processor  204  is implemented using a PowerPC. However, as those of ordinary skill in the art will appreciate, processor  204  may be of any suitable type depending on the timing and the sizing requirements of the present invention. Accordingly, processor  204  may be implementing using an X86 processor, for example, or by DSP devices manufactured by Freescale, Analog Devices, Texas Instruments, as well as other suitable DSP device manufacturers. The processor  204  may be implemented using application specific integrated circuits (ASIC) and/or field programmable gate array (FPGA) devices as well. Combinations of these devices may also be used to implement processor  204 . 
     Memory  206  may include any suitable type of computer-readable media such as random access memory (RAM), flash memory, and various types of read only memory (ROM). The term “computer-readable media” as used herein refers to any medium that may be used to store data and computer-executable instructions. Computer readable media may be implemented in many different forms, including but not limited to non-volatile media, volatile media, and/or transmission media. As those of ordinary skill in the art will understand, RAM or DRAM may be used as the “main memory,” and employed to store system data, digital audio, sensor data, status information, instructions for execution by the processor, and temporary variables or other intermediate data used by the processor  204  while executing instructions. 
     Memory  206  may employ non-volatile memory such as flash memory or ROM as system firmware. Flash memory is also advantageous for in-flight reprogramming operations. In this instance, GCS  30  may provide OLCP with programmed trajectory data that supersedes previously stored trajectory data. Static data, start-up code, the real-time operating system and system applications software are embedded in these memory chips. Of course, non-volatile memory does not require power to maintain data storage on the memory chip. Flash memory is physically rugged and is characterized by fast read access times. ROM may be implemented using PROM, EPROM, E 2 PROM, FLASH-EPROM and/or any other suitable static storage device. 
     Those of ordinary skill in the art will understand that the present invention may also be implemented using other forms of computer-readable media including floppy-disks, flexible disks, hard disks, magnetic tape or any other type of magnetic media, CD-ROM, CDRW, DVD, as well as other forms of optical media such as punch cards, paper tape, optical mark sheets, or any other physical medium with hole patterns or other optically recognizable media. The present invention also defines carrier waves or any other media from which a computer may access data and instructions, as computer-readable media. 
     Embedded system  200  also includes high speed serial data bus interface circuitry  210 . The high speed serial data bus interfaces are configured to transmit and receive information to and from the existing avionics systems disposed on the aircraft. These existing systems may include GPS Navigation systems, inertial navigation systems, and sensor systems that provide altimeter, airspeed, and aircraft attitude (i.e., pitch, roll, yaw, and etc.) data. Those of ordinary skill in the art will understand that high speed serial data bus defines the electrical, mechanical, and functional characteristics of the bus system. The present invention may employ any suitable high speed data bus interface such as MIL-STD-1553, IEEE-1394, ARINC-429, ARINC-629, RS-485, RS-422, and RS-232. Those of ordinary kill in the art will also understand that the present invention should not be construed as being limited by the foregoing examples. For example, the high speed serial data bus interface bus employs a differential interface that supports up to thirty-two interface devices on the bus. The bus is asynchronous and uses a half-duplex format. Data is transmitted using Manchester encoding. 
     Turning to the fly-by-wire interface (FBW I/F) circuit  212 , note that in a man-rated FBW aircraft, the pilot stick and rudder controls are coupled to control transducers that are configured to generate pilot control transducer signals. As the pilot actuates the cockpit control devices (control stick, wheel, pedals, etc.), transducer signals that are proportional to the position of the control device are generated. One common means for measuring such displacements is a linear variable differential transformer (LVDT) sensor. When rotational angles are measured, rotary variable differential transformer (RVDT) sensors may be employed. Accordingly, the FBW I/F circuit  212  of the present invention includes a bus  220  interface that receives digital commands from the processor circuit  204 . These digital signals are converted into analog signals that, in at least one embodiment of the present invention, may be combined with a reference signal provided by the FCS to simulate LVDT or RVDT sensor outputs. The LVDT and/or RDVT simulated output signals are directed to the existing FBW-FCS. The existing FBW-FCS cannot tell the difference between the pilot controls and the simulated signals, and functions as before, driving the various control surface actuators (CSA) disposed on the airplane to cause the elevators, ailerons, rudder, flaps, spoilers, stabilizers, slats, flaperons, elevons, ruddervators, thrust vectoring nozzles, and/or other such control surfaces to move in accordance with the digital commands from the processor circuit  204 . Of course, the digital commands generated by processor circuit  204  are ultimately provided by GCS  30  via the existing command telemetry system. Those of ordinary skill in the art will understand that the present invention should not be construed as being limited to any particular type of aircraft. Obviously, the number and type of control surfaces is a function of aircraft type (F-16, F-18, Airbus A380, B2, F-22, F-117, Boeing 777, etc.). Any FBW aircraft may be converted into a UAV in accordance with the principles of the present invention. 
     The existing aircraft throttle control must be physically manipulated. Thus, throttle interface circuit  214  is configured to provide electromechanical (E/M) actuator  22  with servo-control signals that correspond to the throttle commands provided by GCS  30 . Any suitable linear E/M actuator, such as a ball screw actuator, may be employed to implement E/M actuator  22 . Some aircraft include a servoed throttle (e.g., F-18), and in this instance, an electronic signal is provided directly to the actuator. 
     Embedded system  200  also includes a landing gear interface circuit  216 . The implementation of circuit  216  is largely dependent on the landing gear employed by the FBW aircraft. The details of implementing a landing gear interface circuit that provides appropriate signaling to an existing landing gear system is deemed to be within the skill of one of ordinary skill in the art. 
     System  200  may also include an optional sensor package  218  that is configured to augment the aircraft&#39;s existing sensor systems. Certain older FBW aircraft have analog sensors that are not accommodated by the high speed serial data bus. For example, older F-16 aircraft may be equipped with analog altimeter and airspeed sensors. OLCP  20  requires the aircraft&#39;s heading, roll, pitch, normal acceleration, pressure altitude, true velocity, roll rate, and other such sensor inputs to generate the stick, rudder pedal, and throttle commands that are used to fly the UAV. 
     Finally, embedded system  200  includes a power supply  222 . The power supply  222  includes various DC/DC converters that are configured to convert +28 VDC voltages into the voltages required by OLCP  20  and/or AC/DC converters that convert AC voltages into the voltages required by OLCP  20 . 
     Referring to  FIG. 4 , a perspective view of OLCP  20  in accordance with one embodiment of the present invention is disclosed. As described above, OLCP  20  may be implemented as an embedded electronic control system  200 . The embedded system is environmentally sealed and protected within a rugged enclosure  250 , engineered to withstand the environmental forces applied during flight. In the embodiment depicted in  FIG. 4 , enclosure  250  may be implemented using a ruggedized Airline Transport Rack (ATR) that supports a VME (Versa Modular European) bus format. The front side of enclosure  250  includes a plurality of connectors  252 . The connectors  252 , of course, mate with connectors disposed on the cables  26  that connect OLCP  20  with the existing aircraft systems. Connectors  252  are electrically coupled to I/O plane  254  and provides a means for coupling the multiple VME control channel boards ( 256 ,  258 ,  260 ) to connectors  252 . 
     As those of ordinary skill in the art will understand, the VME bus is a flexible, memory mapped bus system that recognizes each system device as an address, or a block of addresses. The VME bus supports a data transfer rate of approximately 20 Mbytes per second. The VME bus is a “TTL” based backplane that requires +5 VDC as well as ±12 DC. Accordingly, power supply  262  converts +28V DC from the aircraft power bus into +5 VDC and ±12 VDC power. 
     The size of the ATR rack  250  and/or the number of boxes depends on how system redundancy is achieved. In the embodiment depicted herein, each VME board ( 256 ,  258 ,  260 ) implements a single control channel and includes a special purpose processor, memory, various interface circuits, and a power supply. On the other hand, if each ATR rack accommodates one processing channel, several smaller ATR racks may be connected together to achieve redundancy. 
     As those of ordinary skill in the art appreciate, electrical and electronic components generate thermal energy that must be conducted away from the electronic components. As such, the thermal design, including various heat sinking devices and the like, directs the thermal energy to fan unit  266  disposed at the rear portion of the enclosure  250  or through forced air or liquid cooled from the aircraft&#39;s environmental control system (ECS). The fan unit  266  expels the heated air mass into the surrounding space where it dissipates without causing damage to the electronic components. 
     As embodied herein and depicted in  FIG. 5 , a diagram illustrating the OLCP software control system architecture  50  in accordance with the present invention is disclosed. The OLCP control system architecture includes a sensor module  52  and a maneuver module  54  coupled to control module  56 . The output of the control module  56  is coupled to the command module  58 . As described in the hardware description, software modules  52 - 58  are implemented in firmware and executed by processor  204 . 
     The OLCP  20  inputs sensor measurements and maneuver type commands. The sensor measurements may be obtained by way of the high speed serial data bus interface  210  or OLCP sensor package  218  and are pre-conditioned with appropriate scaling. As noted previously, OLCP  20  provides the existing aircraft systems with the pitch stick commands, roll stick commands, and rudder pedal commands in a form that is identical to the LVDT and the RVDT sensors that generate the pilot control transducer signals in a man-rated aircraft. Again, the pitch and roll stick and rudder pedal command signals replace the normal pilot&#39;s stick and rudder pedal input signals. OLCP  20  also generates the throttle servo position commands in a form compatible with electromechanical actuator  22 . Linear E/M actuator  22  moves the throttle lever in accordance with the throttle servo position commands to control engine thrust. In another embodiment of the present invention, the aforementioned E/M actuator may be replaced with other types of actuation devices including electro-hydraulic actuators or other actuators configured to convert an electrical command into a mechanical movement or physical deflection whereby the throttle is displaced. These actuators may also be applied to modulate the fuel flow to the engine (or engines) to control the thrust produced by the engine (or engines) accordingly. 
     Sensor Module  52  mainly is used to convert discontinuous signals such as heading, pitch, and roll angle into continuous signals. The sensor inputs include pitch, roll, heading, normal acceleration, pressure altitude, true velocity, roll rate, etc. Those of ordinary skill in the art will understand that certain sensor measurements such as heading, for example, are provided as continuous analog or digital signals. Sensor module  52  formats the signal and provides the Control module  56  with measurements properly filtered and formatted for computation. The sensor module  52  also performs latching of appropriate sensors in accordance with Control Module  56  requirements, when a maneuver type is commanded. Of course, the sensor module also conditions the sensor data received from the high speed serial data bus interface. 
     GCS  30  may transmit maneuvers or commands to OLCP  20  via the “joystick” method or by way of the maneuver command method. OLCP  20  may also be preprogrammed to follow a predetermined trajectory. Maneuver module  54  is programmed to decipher each type of command and provide control module  56  with “discrete flag counts” and the appropriate reference signals for maneuver types. The discrete flag counts correspond to a maneuver type. Examples of the reference signals include velocity, heading, and altitude reference signals. 
     In the “joystick” method, GCS  30  input controls are substantially identical to the cockpit control devices disposed on a man-rated aircraft, such as the pitch/roll sticks, rudder pedals, engine thrust control, brakes, etc. As the ground based operator manipulates the pitch stick, roll stick, rudder pedals and brakes provided in the GCS simulator, GCS  30  generates the electrical signals corresponding to the operator/pilot commands. These commands are provided to the communication/telemetry systems  32  and transmitted to OLCP  20 . Maneuver module  54  processes these commands on a real-time basis. 
     When GCS  30  employs the maneuver command format, a suite of aircraft maneuvers are available to the ground based GCS operator for input. For example, the operator may select a “2 g turn to the right, hold altitude” command. GCS  30  may use this mode to provide simple autopilot commands, such as “fly at 300 knots at a heading of 270°, at an altitude of 20,000 feet.” The maneuver module  54  responds by generating the discrete flag count and the reference signals corresponding to the maneuver command. 
     In the embodiment wherein OLCP  20  is preprogrammed, processor  204  follows the trajectory instructions stored in firmware memory  206 . Thus, maneuver module receives the reference maneuver command internally, rather than from GCS  30 . 
     As those of ordinary skill in the art will appreciate, the discrete flag count may be stored in a look-up table as a function of the maneuver command. Discrete reference signals may also be stored therein. Maneuver module  54  may be configured to extrapolate between the discrete reference values stored in the table to limit the table size. However, the maneuver module  54  should not be construed as being limited to the table embodiment discussed above. In any event, the Maneuver Module  54  is configured to decipher numerical GCS commands and generate appropriate discrete flags for Control Module  56 . 
     Control Module  56  is programmed to convert the sensor module input and the maneuver module input into a “control law” for each maneuver type. Several types of control laws may be implemented within the Control Module  56  to perform each maneuver type. Each control law is determined by an error-loop type architecture implemented by a Proportional Integral Differential (PID) control law. PID control employs a continuous feedback loop that regulates the controlled system by taking corrective actions in response to any deviation from the desired values (i.e., the reference signals from the maneuver module—velocity, heading, altitude, and other such values). Deviations are generated when the GCS  30  operator changes the desired value or aircraft experiences an event or disturbance, such as wind or turbulence, that results in a change in measured aircraft parameters. The PID controller  56  receives signals from the sensors and computes the error signal (proportional/gain), the sum of all previous errors (integral) and the rate of change of the error (derivative). 
     The gains for the PID control laws are determined prior to the implementation of the code and are typically schedule-based static pressure and dynamic pressure measurements. For a FBW aircraft such as the F-16, with the landing gear retracted, the measurements and the predetermined gain values are related to the desired normal acceleration and roll rate commands. Accordingly, Control Module  56  provides the command module  58  with desired longitudinal acceleration (throttle control), normal acceleration, and roll rate reference signal to the Command Module  58 . 
     The Command Module  58  converts the output of the error-loop command control law to signals that replace the FBW aircraft&#39;s stick, rudder and throttle servo. Four commands are output: pitch stick, roll stick, rudder pedal commands and a throttle servo position command. The Command Module  58  consists of a reverse breakout routine to overcome the hardware/software breakout which is present on the pitch, roll and rudder command paths. The routine adds the breakout value if the Control Module control command signal is within the breakout limits of the breakout function. When the Control Module control command signal is above the pitch and roll breakout value the command is allowed to pass through directly to the pitch and roll stick summing point. The FBW aircraft&#39;s control law will also contain a stick gradient function converting stick measurements to normal acceleration command signals for the pitch flight control system and roll rate command signals for the lateral/directional flight control system. The Control Module  56  is designed to command normal acceleration and roll rate. Therefore, an additional algorithm within the Command Module  56  is required to provide a “reverse” stick gradient function for the Control Module  58  outputs. A table lookup routine may be used to interpolate between the discrete points determined from the optimization routine creating a continuous output signal. 
     Referring to  FIG. 6 , a flow chart illustrating the software control of the OLCP is disclosed. The control loop is implemented by scheduling events within a predetermined timing frame  60  that is continuously repeated. In one embodiment of the present invention, the frame rate is substantially equal to 64 Hz. Therefore, the software calls each scheduled event once every 15.625 milliseconds. For reliability and extensibility reasons, i.e., the ability to add new functionality as mission requirements change and grow, the frame rate includes a 50-100% execution margin depending on the implementation. Those of ordinary skill in the art will understand that the frame rate may be any suitable rate consistent with the aircraft&#39;s maneuvering and stability requirements. For example, the F-18 may require an 80 Hz frame rate. 
     In step  600 , processor  204  performs initialization and built-in testing. As those of ordinary skill in the art will appreciate, each processing channel in OLCP  20  must perform a self-test to ensure system reliability. The processor, RAM, and firmware are tested to ensure that these circuits are operating properly. The processor may be required to perform certain predetermined computations to ensure computational reliability. Memory may be checked by determining whether various memory locations may be accessed. The BIT tests may test each of the interface circuits to determine whether these circuits are able to read and write to the existing aircraft systems. The self-tests also test the power supply  222  to ensure that aircraft input power (+28 VDC), and measure the output of the various power rails (+5 VDC, ±12 VDC, etc.). The self-tests may also perform communication tests to ensure that OLCP  20  is able to communicate to GCS  30  via the aircraft command telemetry unit. After step  600  is completed, embedded processor  204  begins continuous execution of the control loop. 
     In step  602 , processor  204  obtains the various avionics signals from the high speed serial data bus interface. These signals typically include navigation and aircraft status inputs. In step  604 , discrete signals and various analog signals are also obtained. An example of a discrete signal is the landing gear status. In older FBW aircraft, certain parameters such as dynamic pressure (airspeed) and static pressure (altitude) may not be available on the high speed serial data bus. These parameters may be provided by analog sensors. Both of these steps are performed by calling the sensor module  52 . 
     At this point in the frame (step  606 ), the maneuver module  54  determines the state of the OLCP  20 . As noted previously, GCS  30  commands may be provided by GCS  30  in either the “joystick” mode or the “maneuver command” mode, or the state of OLCP  20  may be provided by a preprogrammed trajectory stored in firmware. For example, GCS  30  may order the UAV to proceed on a straight and level path, perform a barrel roll, perform a turn, or any other such maneuver. As described above, maneuver module  54  responds by generating the appropriate discrete flag count and reference signals corresponding to the maneuver command. Those of ordinary skill in the art will also understand that the desired state of OLCP  20  may include actuation of weapons delivery systems when the UAV is configured as a combat air vehicle (CAV). 
     In step  608 , processor  204  calls the control module  56  to compute the OLCP  20  control law. Again, the control law is determined by an error-loop type architecture implemented by a Proportional Integral Differential (PID) control law. 
     Subsequently, in step  610 , Command Module  58  converts the output of the error-loop command control law into pitch stick, roll stick, rudder pedal, and throttle servo position commands. 
     At this point in the discussion it is important to recall that OLCP  20  is implemented with redundant processing channels. If OLCP employs three redundant channels, the activities of the sensor module, the maneuver module, the control module, and the command module are performed in parallel by three machines. In step  612 , the channel commands for the frame are exchanged and a voting algorithm is performed. In one embodiment of the present invention, all of the channel outputs are compared to a failure threshold. If a given channel exceeds the threshold, its result is thrown out. Thus, the remaining two channels are averaged. In another embodiment, the high and low value may be disregarded and the middle value selected. Alternatively, in a two channel system, both values may be averaged. In a four channel system, the voting algorithm may be configured to throw out the high and low values for each parameter and average the middle values. Those of ordinary skill in the art will understand that the present invention may be implemented using any reasonable voting algorithm. 
     In step  614 , processor  204  writes the pitch stick, roll stick, rudder pedal output commands to FBW I/F circuit  212  (See  FIG. 3 ) which converts these values into simulated LVDT/RVDT signals for use by the existing FBW-FCS on board the aircraft. Similarly, processor  204  provides a throttle position command to the throttle I/F circuit  214 . Throttle I/F circuit  214  transmits a throttle servo position command to the E/M actuator  230  in response thereto. 
     At this point in frame  60 , continuous BIT testing is performed. Continuous BIT (step  616 ) may be implemented as sub-set of the tests performed in step  600 . This testing provides in flight failure detection and isolation and tests each processing channel on a frame-by-frame basis. 
     Finally, processor  204  enters an idle state and waits for the remainder of the 15.625 millisecond frame to complete. As noted above, frame  60  may include a margin of 50%-100%. In the latter case, processor  204  may be idle for 7.8125 milliseconds before repeating steps  602 - 618  in the next frame sequence. 
     As embodied herein and depicted in  FIG. 7A , a high-level block diagram of the FBW interface circuit  212  depicted in  FIG. 3  in accordance with another embodiment of the present invention is disclosed. This block diagram of  FIG. 7A  illustrates an “analog solution” for the OLCP interface. As shown, pitch, roll, and rudder commands are provided by the OLCP  20  to the interface circuit  212 . In one embodiment, this data is provided by a 16 bit data bus. The digital data is converted into an analog signal by DAC  2120  and multiplied with a analog legacy aircraft reference signal by multiplier  2124 . The output of multiplier  2124  yields an analog OLCP input command to the FBW-FCS of the legacy aircraft. 
     Referring to  FIG. 7B , a detailed block diagram of the FBW interface circuit  212  depicted in  FIG. 7A  is provided. Interface circuit  212  includes four digital-to-analog converters (DAC)  2120 ,  2126 ,  2132  and  2138  coupled to the microprocessor  204  by way of bus  220 . By way of example, DAC  2120  may be employed in the data channel corresponding to digital pitch stick commands, DAC  2126  may be employed in the data channel corresponding to roll stick commands, DAC  2132  may be employed in the data channel corresponding to rudder commands, and DAC  2138  may be employed in the data channel corresponding to brake commands. In one embodiment of the present invention the DACs include 16 bit data registers that latch data present on the data bus in response to a control signal provided by microprocessor  204 . 
     DAC  2120  converts the 16 bit digital data into an analog command signal directed into a multiplication circuitry  2124 . The multiplication circuitry  2124  multiplies the analog command signal an AC reference signal, amplifies the product and performs further analog signal formatting before providing the channel output signal to the aircraft fly-by-wire (FBW) system. 
     In the example provided above, the channel 0 output signal (CH 0 OUT) provided by multiplier circuitry  2124  is the exact representation of a pilot pitch stick command. In other words, the fly-by-wire system cannot tell the difference between an actual pilot pitch stick command and the CH 0 OUT signal. In similar fashion, DACs ( 2126 ,  2132   213 ) provide their corresponding analog command signals to their respective multiplier circuits ( 2130 ,  2136 ,  2142 ). Accordingly, the FBW interface circuit  212  may be configured to provide FBW pitch stick commands via channel 0 output, FBW roll pitch commands via channel 1 output, FBW rudder commands via the channel 2 output, and FBW brake commands via the channel 3 output. As noted previously, throttle commands are directed to the aircraft by way of a mechanical actuator. This may be implemented using a servo-throttle mechanism of the type employed in both commercial airliners and military aircraft autopilot systems. 
     For example, in the “joystick” method, previously described above, the operator I/O in GCS  30  includes a joystick, peddles, and other such pilot control devices. The remote pilot is provided with aircraft sensor data via the telemetry link and has a “pilot&#39;s view” by way of video camera  24 . In one embodiment, the remote pilot wears head gear that provides a tracking signal to the on-board video camera such that the video camera moves within the canopy to provide the remote pilot with the desired vantage point. As described previously, the GCS  30  converts the signals received from the GCS  30  pitch/roll sticks, pedals, engine thrust control, etc., into data more suitable for RF transmission. A given stick command may be formatted as a digital block of data having an identification header and a block data representing the command. The data may be transmitted using spread spectrum techniques, frequency hopping techniques or by way of a satellite data link. The data is provided to the UAV computer in the manner previously described or in any suitable comparable manner via the telemetry unit. The processor  204  reads the header, processes the data accordingly and provides each DAC ( 2120 ,  2126 ,  2132  and  2138 ) with a digital representation of the pilot command in the manner described above. 
     As noted above, the GCS  30  may be configured to provide the remote pilot/operator with various maneuver commands, such as turn, roll, etc. In this case, the OLCP computer  204  is programmed to derive the digital stick, pedal, thrust commands, etc. from the maneuver command while taking account of the avionics systems data provided by the high speed serial data bus interface circuitry  210 . The OLCP computer  204  is will also derive the digital stick, pedal, thrust commands, etc. when it is programmed to perform maneuvers in accordance with a preprogrammed flight trajectory. 
     Referring to  FIG. 8 , a detailed schematic of the multiplier circuit  2124  shown in  FIG. 7  is disclosed. Because multiplier circuitry  2124  is substantially identical to the other multiplier circuits ( 2130 ,  2136 ,  2142 ) only multiplier circuit  2124  is shown in the interests of brevity. In one embodiment of the present invention, multiplier circuit  2124  includes a quadrature multiplier device  2133  which receives the analog command from DAC  2120  and an AC reference signal. The quadrature multiplier is a four-quadrant analog multiplier that is a purely analog circuit that creates an output that is proportional to the multiplication of the two input values (X, Y), i.e., Z=(X)(Y). The Four Quadrant term refers to the ability of the circuit to handle positive and negative values of input, so it can compute: Z=(+X)(+Y); Z=(−X)(+Y); Z=(+X)(−Y); or Z=(−X)(−Y). The two signals (X and Y) are multiplied and the product (Z) is provided to amplifier  2125 . The amplified signal is directed to output transformer  2121 ′. 
     Referring to  FIG. 8  and  FIG. 12 , the quadrature multiplier device  2133  receives an analog command signal that is a time varying +/− VDC signal centered around 0 volts and is proportional to the OLCP command. The AC reference signal received from the aircraft is a differential peak-to-peak AC signal, i.e., that it is centered around 0 volts and varies from +VAC to −VAC. One differential signal input is provided to one input of transformer  2121  and the other differential signal input is provided to its corresponding input of transformer  2121 . Because one end of the transformer output is grounded, the signal provided to quadrature multiplier  2123  at pin  3  varies from 0 volts to +VAC. One aircraft type is known to provide a 26 VAC peak-to-peak reference signal having a frequency of 800 Hz. This is shown in  FIG. 8  merely as an illustrative example. The output of quadrature multiplier device  2133  is a time varying AC voltage signal with a magnitude proportional to the OLCP command. The phase of the signal provides directional information. In the channel 0 example, the directional information relates to whether the stick is being moved forward or aft. The output of quadrature multiplier device  2133  is directed into operational amplifier  2125 . The gain of the amplifier is set by the RC circuit  2129 . Output transformer  2121 ′ provides a differential out signal  21240  that mimics an LVDT or RVDT signal. Thus, the output of the multiplier circuit is directed into the FBW system via a signal input previously occupied by an LVDT output. 
     It will be apparent to those of ordinary skill in the pertinent art that modifications and variations can be made to the DACs, quadrature multiplier device  2133  and the operational amplifier employed by the present invention depending on the application, type of aircraft being modified, various performance issues, etc. For example, the DACs ( 2120 ,  2126 ,  2132  and  2138 ) may be implemented by any suitable 16 bit monolithic D/A converter such as the AD669 manufactured by Analog devices. The quadrature multiplier device  2133  may be implemented by any suitable Four-Quadrant Analog Multiplier such as AD 633 which is also manufactured by Analog Devices. The amplifier  2125  may implemented using any suitable operational amplifier such as OP 727 which is manufactured by Analog Devices. 
     Referring to  FIG. 9 , a detailed schematic of the power bus provided by the power supply depicted in  FIG. 3  is shown. The power bus provides +5 V, +/−12V and ground as needed in the circuit depicted in  FIG. 8 . The various capacitors shown in  FIG. 9  provide noise immunity. 
     Referring to  FIGS. 10A-10C , voltage waveforms provided by the DACs ( 2120 ,  2126 ,  2132  and  2138 ) shown in  FIG. 7  and  FIG. 15  are disclosed.  FIG. 10A  is a representative example of DAC  2120  and shows the output when the stick is forward. The “+V” is a voltage level that is proportional to the displacement of the stick.  FIG. 10B  shows the output of DAC  2120  when the stick is in the neutral position.  FIG. 10C  depicts the output of the DAC  2120  when the stick is displaced in the aft direction. Again, the “−V” is a voltage level that is proportional to the displacement of the stick. Referring to  FIG. 11 , an example of a time varying voltage waveform  1100  in accordance with  FIGS. 10A-10C  is disclosed. Waveform  1100  follows directly from the explanation of  FIGS. 10A-10C . The various voltage levels represent DC voltages produced by the DAC  2120  over time. Each DC voltage represents a stick displacement. If the DC voltage is positive, the stick is displaced forwardly. Conversely, if the DC voltage is negative, the stick is displaced in the aft direction. 
     Referring to  FIG. 12 , an AC reference voltage signal in accordance with the embodiments depicted in  FIG. 7  and  FIG. 15  is disclosed. As explained above, the AC reference signal may be a sinusoidal peak-to-peak signal. In the example provided above, the AC reference signal may be 26 VAC having a frequency of 800 Hz (i.e., a period of 1/800 seconds or 5026 radians/sec). As those of ordinary skill in the art will appreciate, the frequency could be 1 KHz, 1.6 KHz, 4 KHz or any other frequency provided by the aircraft&#39;s electrical system. 
     Referring to  FIGS. 13A-13C , command voltage waveforms provided to the existing fly-by-wire aircraft in accordance with an embodiment of the invention are disclosed.  FIGS. 13A-13C  represent the peak-to-peak output of transformer  2121 ′ in  FIG. 8 . The command voltage waveforms, of course, are produced by multiplying the DAC output voltage by the AC reference signal. In  FIG. 13A , the stick is displaced forward by a distance proportional to the peak-to-peak voltage. In  FIG. 13C , the stick is displaced aft. Note that the signal depicted in  FIG. 13C  is 180° out of phase with the one shown in  FIG. 13A . The phase of the signal is indicative of the displacement direction.  FIG. 13B  shows the stick in the neutral position and the magnitude of the signal is equal to about 0 (zero) volts. 
     Referring to  FIG. 14 , an example of a time varying command voltage waveform in accordance with  FIGS. 13A-13C  is disclosed.  FIG. 1400  is an example of the stick being displaced in the forward direction by an increasing amount ( 1402 ), then to the neutral position ( 1404 ) and the aft ( 1406 ). In the example embodiments depicted in  FIG. 9-14 , the method is directly applicable to FBW aircraft that use either LVDT or RVDT type of stick and rudder pedal sensors. the output of the four-quadrant analog multiplier circuits appears as an AC signal, whose frequency is identical to the reference input and whose magnitude is proportional to the magnitude of the DC signal (which was proportional to the command from the OLCP). The phase of the output (with respect to the reference oscillation) is dependent upon the sign of the OLCP command, this phase would represent the movement of the LVDT/RVDT measurement of pilot&#39;s stick input to be forward stick (nose down) or aft stick (nose up) for example. 
     As embodied herein and depicted in  FIG. 15 , a detailed block diagram of the FBW interface circuit  212  depicted in  FIG. 3  in accordance with yet another embodiment of the present invention is disclosed. This embodiment may be referred to as the digital solution because it replaces the analog AC reference signal with a digital timing circuit. Like all of the previous embodiments, the LVDT/RVDT elements are electrically removed from inputs to the FBW flight control system and the FBW interface circuit  212  is inserted in their place. The AC reference signal is directed into analog-to-digital converter  1502 . The A/D converts the AC signal into a time varying digital signal which, in the embodiment depicted in  FIG. 15 , is a 16 bit signal. The 16 bit timing signal is directed into a field programmable gate array (FPGA) circuit  205 . At the same time, a 16 bit digital input signal that represents the pitch, roll or rudder pedal input (depending on the channel) is also directed to FPGA  205 . As those skilled in the art will appreciate, the digital command data may be any suitable number of bits (10, 12, 16 or 18 bits), depending on the resolution required by the application. The digital command, while shown herein as being a parallel digital signal, may also be provided to FPGA  205  by way of a serial interface. 
     The FPGA is programmed to combine the digital command signal and the digital timing signal in a way that is analogous to the embodiment described previously. In other words, the gate circuits are programmed to represent the multiplication of the digital command signal X by the digital timing signal Y. In the previous embodiment, the multiplication of the command signal and the AC reference was done in the analog domain. In this embodiment the product (X*Y) is generated digitally. Like the previously described analog embodiment, the logic gates compute all combinations of positive and negative signals: Z=(+X)*(+Y); Z=(−X)*(+Y); Z=(+X)*(−Y); or Z=(−X)*(−Y). The output (Z) is directed to the DACS ( 2120 ,  2126 ,  2132 , and  2138 ) depending on the channel. Each DAC converts the digital data to an AC analog output signal. FPGA  205  is also configured to provide two clock signals. One clock signal is employed by the A/D  1502  to sample and convert the analog reference input into a digital value for use by the FPGA  205 . The other clock signal is employed by the DACS to generate the analog output signal from the digital FPGA  205  output. The circuit depicted in  FIG. 8  is modified accordingly, such that the analog command output signal mimics an LVDT signal as before. As those of ordinary skill in the art will appreciate, a Field Programmable Gate Array (FPGA) may be replaced by an application specific integrated circuit (ASIC). 
     Note that each DAC output is an AC signal, whose frequency is identical to the reference input and whose magnitude is proportional to the magnitude of the DC signal. The phase of the output (with respect to the reference oscillation) is dependent upon the sign of the command signal. As before, the phase of the AC output signal represents, e.g., the direction of the stick or rudder displacement. 
     Referring back to  FIG. 4 , the FBW interface circuit  212 , which may be thought of as a “stick interface circuit,” may be disposed the OLCP “box” enclosure  250 . Each interface circuit  212  (e.g., pitch stick, roll stick, rudder, brake, etc.) may be disposed on one or more of the circuit cards. Clearly, each simulated LVDT measurement signal is generated by one interface circuit  212 . For example, if the legacy FBW aircraft requires a pitch stick, roll stick, and rudder input, these inputs may be provided by a pitch LVDT/RVDT sensor, a roll LVDT/RVDT sensor and a rudder pedal LVDT/RVDT sensor. In a system that provides “quad-redundancy,” the interface circuitry  212  is configured to provide 12 individual interface circuits. As noted above in reference to  FIG. 4 , the interface  212  circuit card communicates with the main processor  204  via the backplane (for example a VME bus). 
     As noted above, processor  204  may be configured to perform autonomous control computations or use the remote control commands embedded in the uplinked signals. The RF signals from the uplink are demodulated, decoded and provided to interface circuits  212  to the appropriate address via the VME bus  212 . As noted above, certain legacy aircraft employ quad-redundancy. To insure the redundant FCS obtained the same signals for each of the 4 commands (e.g., pitch), the processor  204  is programmed to provide the same digital signal to each of the pitch stick interface circuits. In other embodiments of the present invention, instead of providing redundancy with one computer providing four outputs, two computers may be programmed to generate two inputs (four total) or four computers may be configured to generate one for each circuit. The benefit of using multiple computers is that the computing device itself does not become a single point of failure. 
     Certain aircraft use LVDT/RVDT sensors as a means for commanding Brakes (Brake by wire). As described above, the present invention is well suited for providing the legacy FBW system with brake commands to control the speed, deceleration and ability to stop of an aircraft under remote or autonomous control. 
     All references, including publications, patent applications, and patents, cited herein are hereby incorporated by reference to the same extent as if each reference were individually and specifically indicated to be incorporated by reference and were set forth in its entirety herein. 
     The use of the terms “a” and “an” and “the” and similar referents in the context of describing the invention (especially in the context of the following claims) are to be construed to cover both the singular and the plural, unless otherwise indicated herein or clearly contradicted by context. The terms “comprising,” “having,” “including,” and “containing” are to be construed as open-ended terms (i.e., meaning “including, but not limited to,”) unless otherwise noted. The term “connected” is to be construed as partly or wholly contained within, attached to, or joined together, even if there is something intervening. 
     The recitation of ranges of values herein are merely intended to serve as a shorthand method of referring individually to each separate value falling within the range, unless otherwise indicated herein, and each separate value is incorporated into the specification as if it were individually recited herein. 
     All methods described herein can be performed in any suitable order unless otherwise indicated herein or otherwise clearly contradicted by context. The use of any and all examples, or exemplary language (e.g., “such as”) provided herein, is intended merely to better illuminate embodiments of the invention and does not impose a limitation on the scope of the invention unless otherwise claimed. 
     No language in the specification should be construed as indicating any non-claimed element as essential to the practice of the invention. 
     It will be apparent to those skilled in the art that various modifications and variations can be made to the present invention without departing from the spirit and scope of the invention. There is no intention to limit the invention to the specific form or forms disclosed, but on the contrary, the intention is to cover all modifications, alternative constructions, and equivalents falling within the spirit and scope of the invention, as defined in the appended claims. Thus, it is intended that the present invention cover the modifications and variations of this invention provided they come within the scope of the appended claims and their equivalents.