Patent Publication Number: US-2013253735-A1

Title: Distributed electronic engine control architecture

Description:
BACKGROUND OF THE INVENTION 
     Generally, the present invention is directed to aircraft, and more particularly, exemplary embodiments of the present invention are directed to distributed electronic engine control architectures. 
     More electric engines may include gas turbine engines with increasingly electrically powered means of activation, including electric actuators and pumps driven by electricity rather than mechanical power. Conventionally, gas turbine electronic engine controls (EEC) include a central controller configured to receive line-by-line electrical value (including voltage, current, frequency, and/or resistance) information from a plurality of individual EEC sensors. This results in significant harness weight between the central controller and an associated engine. Furthermore, if conventional mechanical actuators and pumps are eventually replaced with electrically actuated systems, harness weight increases further, providing a limiting factor in the design and implementation of more electric engines. 
     BRIEF DESCRIPTION OF THE INVENTION 
     According to an exemplary embodiment of the present invention, a distributed electronic control system for an aircraft includes an engine data controller arranged on an airframe of the aircraft and a plurality of engine data concentrators arranged proximate the engine of the aircraft in signal communication with the engine data controller. The engine data controller is configured to process information related to an engine of the aircraft. Also, the plurality of engine data concentrators are configured to receive engine sensor information from a plurality of engine sensors and to transmit the engine sensor information to the engine data controller. 
     According to another exemplary embodiment of the present invention, a distributed electronic control system for an aircraft includes a plurality of airframe-mounted engine control components arranged on an airframe of the aircraft and a plurality of engine-mounted engine control components arranged proximate the engine of the aircraft in signal communication with the airframe-mounted engine control components. The engine control components are configured to process information related to an engine of the aircraft. The plurality of engine-mounted engine control components are configured to receive engine sensor information from a plurality of engine sensors and to transmit the engine sensor information to the airframe-mounted engine control components. The plurality of engine-mounted engine control components include a centralized power conditioning system configured to condition and distribute power to the plurality of airframe-mounted engine control components and the plurality of engine-mounted engine control components. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The subject matter which is regarded as the invention is particularly pointed out and distinctly claimed in the claims at the conclusion of the specification. The foregoing and other features, and advantages of the invention are apparent from the following detailed description taken in conjunction with the accompanying drawings in which: 
         FIG. 1  is a schematic of a distributed engine control system, according to an exemplary embodiment of the present invention; and 
         FIG. 2  is a detailed schematic of a distributed engine control system with redundancy, according to an exemplary embodiment of the present invention. 
     
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
     More electric engines may include gas turbine engines with increasingly electrically powered means of activation such as, for example, electric actuators and pumps that are driven by electricity rather than mechanical power. According to exemplary embodiments of the present invention, a distributed electronic engine control (EEC) architecture has been provided which reduces harness weight and complexity as compared to conventional EEC systems. The EEC architecture also provides for more electric engines with increasingly electrically powered means of activation. 
     Turning to  FIG. 1 , a schematic of a distributed engine control system  100  of an aircraft is illustrated. The system  100  includes a plurality of airframe-mounted components  101  and a plurality of engine-mounted components  102 . The airframe-mounted components  101  may be control components mounted to an airframe of an aircraft, for example, within an avionics compartment of an airframe of an aircraft. An avionics compartment may be a compartmentalized portion of an airframe of an aircraft with a relaxed environment as compared to an environment proximate an engine. The engine-mounted components  102  may be control components mounted to and/or proximate an aircraft engine, for example, within or on the actual engine or within an engine pylon affixed to a wing section of an aircraft. 
     The airframe-mounted components may include an aircraft controller  112 . The aircraft controller  112  may be any controller suitable for processing information related to an aircraft, including a centralized processor configured to control aircraft operations. 
     The airframe-mounted components further include an engine data controller  103  in signal communication with the aircraft controller  112 . The engine data controller  103  may be a single or dual-channel controller configured to receive electrical power from power bus  108  and data from data bus  109 . The engine data controller  103  may process data received to determine actuation adjustments for control of an aircraft engine, for example, pump pressure, actuation position, or any other suitable adjustments. The power bus  108  and the data bus  109  may extend from the airframe-mounted components  101  to the engine-mounted components  102 . 
     The engine mounted components  102  may include a plurality of engine data concentrators  104 ,  105 , and  106  in signal communication with the engine data controller  103  over the data bus  109 . Each engine data concentrator  104 ,  105 , and  106  may include an Input/Output portion (I/O), an analog sensor interface associated with the I/O, an analog-to-digital converter (ADC) in communication with the analog sensor interface, and a processor in communication with the ADC. It shall be understood, however, that in some cases, the sensor interface may receive “on” or “off” values from switch and, in such cases, the sensor need not be an analog sensor and the ADC may be omitted. Power may be received at each engine data concentrator over power bus  108 . Thus, associated processors may be powered from the power bus  108 , and may perform a plurality of functions related to information received over the analog sensor interface. The received information may be relayed to the engine data controller  103  over data bus  109 . The information may be received from associated engine sensors  141 ,  151 , and  161 . Each of the engine sensors  141 ,  151 , and  161  may be engine-mounted sensors (e.g., analog temperature sensors, position sensors, pressure sensors etc.) configured to produce an analog signal in response to a change in engine operation (e.g., temperature, position information, pressure, etc.). Therefore, each data concentrator  104 ,  105 , and  106  may “concentrate” engine sensor information, assemble the concentration engine sensor information into at least one data packet, and transmit the data packet to the engine data controller  103  in a controlled manner dictated by any associated communications protocol implemented for the data bus  109 . The communications protocol may be designed to be a fast transmission protocol, controller area network protocol, or any other suitable protocol. 
     Turning back to  FIG. 1 , the engine-mounted components further include power conditioner/actuator controls  107 . The power conditioner/actuator controls  107  may be in signal communication with the data bus  109 , and may transmit and receive information from the engine data controller  103 . The power conditioner/actuator controls may receive engine power over power bus  110 , for example, from a permanent magnet generator generating electricity from an aircraft engine. The power conditioner/actuator controls  107  may also receive aircraft power over power bus  111 , for example, from an aircraft battery bank or other power supply. The power conditioner/actuator controls  107  may condition the received power from buses  110  and  111  into power for transmission across power bus  108 . Therefore, the power conditioner/actuator controls  107  allow for a centralized power conditioning system mounted on or proximate the aircraft engine which provides conditioned power to a plurality of engine data concentrators. This may reduce overall wire-weight allowing for more efficient aircraft operations. Furthermore, the power conditioner/actuator controls  107  may provide conditioned power to a plurality of electric engine actuators  171  based on control information received from engine data controller  103  over data bus  109 . As such, overall wire weight is further reduced while allowing for more electric engine controls distributed across airframe-mounted and engine-mounted components. 
     Although particularly illustrated and described as having singular power and data buses, it should be understood that the same may be varied in many ways to allow for increased redundancy while still realizing reduced wire weight and increased engine efficiency. For example, a distributed engine control system  200  with both power and data bus redundancy is illustrated in  FIG. 2 , according to an exemplary embodiment of the present invention. 
     The system  200  includes a plurality of airframe-mounted components  201  and a plurality of engine-mounted components  202 . The airframe-mounted components  201  may be control components mounted to an airframe of an aircraft, for example, within an avionics compartment of an airframe of an aircraft. The engine-mounted components  202  may be control components mounted to and/or proximate an aircraft engine, for example, within or on the actual engine or within an engine pylon affixed to a wing section of an aircraft. 
     The airframe-mounted components may include the aircraft controller  112  described above. 
     The airframe-mounted components further include a dual-channel engine data controller  203  in signal communication with the aircraft controller  112 . The engine data controller  203  may be controller configured to receive electrical power from power buses  208  and  209 , and data from data buses  210  and  211 . The engine data controller  203  may process data received to determine actuation adjustments for control of an aircraft engine, for example, pump pressure, actuation position, or any other suitable adjustments. The power buses  208 - 209  and the data buses  210 - 211  may extend from the airframe-mounted components  201  to the engine-mounted components  202 . 
     The engine mounted components  202  may include a plurality of engine data concentrators  204  and  205  in signal communication with the engine data controller  203  over the data buses  210 - 211 . Each engine data concentrator  204 - 205  may include an Input/Output portion (I/O), an analog sensor interface associated with the I/O, an analog-to-digital converter (ADC) in communication with the analog sensor interface, and a processor in communication with the ADC. Power may be received at each engine data concentrator over power buses  208 - 209 . Thus, associated processors may be powered from the power buses  208 - 209 , and may perform a plurality of functions related to information received over the analog sensor interface. The received information may be relayed to the engine data controller over data buses  208 - 209 . The information may be received from associated engine sensors  241 ,  242 ,  251 , and  252 . Each of the engine sensors  241 ,  242 ,  251 , and  252  may be engine-mounted sensors (e.g., analog temperature sensors, position sensors, pressure sensors etc.) configured to produce an analog signal in response to a change in engine operation (e.g., temperature, position information, pressure, etc.). Therefore, each data concentrator  204 - 205  may “concentrate” engine sensor information, assemble the concentration engine sensor information into at least one data packet, and transmit the data packet to the engine data controller  203  in a controlled manner dictated by any associated communications protocol implemented for the data buses  210 - 211 . The communications protocol may be designed to be a fast transmission protocol, controller area network protocol, or any other suitable protocol. 
     Turning back to  FIG. 2 , the engine-mounted components further include optional airframe interface controllers  206  in communication with the engine data controller  203  over data buses  210 - 211 , and configured to receive power from power buses  208 - 209 . The airframe interface controllers  206  may include air supply controllers, bleed controllers, or any other controllers which may be integrated with data buses  210 - 211 , which may therefore further reduce wire weight. The airframe interface controllers  206  may receive associated control information from the engine data controller  203  which may be relayed from aircraft controller  112  (e.g., associated bleed air flow measurements, etc.). The airframe interface controllers  206  may also transmit information to the engine data controller  203  over data buses  210 - 211 . 
     Turning back to  FIG. 2 , the engine-mounted components further include power conditioner system  207 . The power conditioner system  207  may be in signal communication with the data buses  210 - 211 , and may transmit and receive information from the engine data controller  203 . 
     The power conditioner system  207  may include at least two power conditioners  271  and  272  disposed to condition power received from engine power over power bus  110 , for example, from a permanent magnet generator generating electricity from an aircraft engine. The power conditioners  271  and  272  may also receive aircraft power over power bus  111 , for example, from an aircraft battery bank or other power supply. The power conditioners  271  and  272  may condition the received power from buses  110  and  111  into power for transmission across power buses  208  and  209 . Furthermore, the power conditioners  271  and  272  may provide conditioned power to a plurality of electric engine actuator controls  273 ,  274 , and  275  integrated in the power conditioner system  207 . Each electric engine actuator control  273 ,  274 , and  275  may control an associated actuator  276 ,  277 , and  278  based on control information received from engine data controller  203  over data buses  210 - 211 . 
     While the invention has been described in detail in connection with only a limited number of embodiments, it should be readily understood that the invention is not limited to such disclosed embodiments. Rather, the invention can be modified to incorporate any number of variations, alterations, substitutions or equivalent arrangements not heretofore described, but which are commensurate with the spirit and scope of the invention. Additionally, while various embodiments of the invention have been described, it is to be understood that aspects of the invention may include only some of the described embodiments. Accordingly, the invention is not to be seen as limited by the foregoing description, but is only limited by the scope of the appended claims.