Patent Application
20130253735
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.
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.
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:
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.
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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.
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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
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.
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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.
