Aircraft Fuel System - Report Example

Response to Questions: Air Craft Fuel System TASK Engine management system The two types of electronic engine management system found on the modern aircraft are Full Authority Digital Electronic control system and the hydro mechanically controls system, but of which FADEC is the most common. The FADEC is a system in which the engine is fully under the control of a combination of electronic microprocessor and the sensors that measure oil pressure, RPM, atmospheric pressure, CHT, MAP and EGT. The FADEC system controls the injection of fuel and timing of fuel, thereby optimizing engine power. The system adds considerable wiring to the system since it needs the separate power supply for backup that comes from the alternator. Due to the extra weight of parts of the full authority digital engine control, the balance and the weight of the aircraft has to be checked and recalculated. There are additional cockpit controls and switches to control the fuel pump with FADEC power supply protecting these systems (BTEC Unit 81 handout). The ECU uses a 3D memory map in controlling the injection of appropriate fuel amount taking into consideration varied ambient circumstances, such as air pressure density and the air temperature, in relation to throttle settings and RPM. The ECU can sense barometric pressure and in response compensates the appropriate fuel amount to be injected. The spark plug ignition timing is regulated depending on throttle settings for each load. The variations in ignition timing yields faster engine start-up, as well as smooth operations with the variable loads. The FADEC engine does not require chocking during engine start since the ECU regulates fall fuel cylinders and helps in retarding the ignition. FADEC system replaces the carburetor. In this case, the ignition does not wholly depend on aircraft electrical system. It combines well with the system of fuel injection (Federal Aviation Administration). The hydro mechanical control system, on the other hand, makes the driver control the power plant of the aircraft by utilizing an internal combustion engine. The sets of sensors and control are the alternator master and the battery master. The battery master activates battery contractor that links the battery to the electrical bus aircraft while the master alternator supplies power to alternators field circuit to activate it (BTEC Unit 81 handout). The two switches are the source of power to the aircraft systems. The throttle sets up the power level required and controls the massive airflow rate in the carbureted engines that is to be delivered to the cylinder. Pitch control makes adjustments to the speed unit which, in effect, adjusts the pitch of the propeller and works to control the load required by the engine for maintaining the RPM. The mixture control, in turn, sets the fuel amount required to be added the airflow intake. At high altitudes, the oxygen level reduces and, therefore, the volume of fuel must be readjusted. In this process, the ignition opens the P-lead circuit by activating the magnetos, which send the output voltage to spark plugs. The magnetrons connect the engine through the gearing. Whenever there is movement in the crankshaft, the magnetrons are turned which generate sparking voltage. This keeps the engine working even during the electric failure. TASK 2 FADEC SYSTEM Fig 02 INPUT AND OUTPUTS OF A FADEC FADEC system controls the engine. FADEC computers have the ability to process more data than the hydro mechanical system. This feature enables the FADEC to optimize the operation of the engine. Other than fuel control and power management functions, the system can perform the following functions: Sourcing data for ECM, shutting down, starting up and controlling the ignition, controlling of thrust reverser, detecting of faults for the system and monitoring all components connected to the engine and sourcing of data to be used for engine indication. For the FADEC to achieve its multiple tasks, it has various components with distinctive functions. The FADEC components are centered on an Electronic Engine Control (EEC), which comprise the FADEC computers and the auxiliary components. These components include: the sensors, ignition systems, stator valves, actuator controller, alternator, reverser systems and data buses (Federal Aviation Administration) THE ELECTRONIC ENGINE CONTROL (EEC) The EEC consists of digital computers called channels A and B that constitute the “brains” of FADEC and are located at the fan case area of the aircraft engine. All the computers have their own power supply with one connection to the aircraft alternator and the other for connection to FADEC. Furthermore, the EEC is divided into two: the upper and the lower. The upper part contains two microprocessor controllers that are independent from each other, both of which serve as control channels. On the other hand, the lower portion has an electronic circuit board. In addition, the EEC has pressure transducers used for air pressure measurement and electric sensors, which function is to detect thrust lever angle and temperatures. EEC receives data from the airframe systems, calculates values from its own data and compares it with the data from the air frame system. EEC compares the calculated N1 command and the actual command. If the difference exists, the N1 control sends a deceleration or acceleration command to the metering valve control section for the necessary action to be taken (BTEC Unit 81 handout). The EEC has dual outputs connected to the torque motors for controlling the servo valves and switching off air pressure. With the help of software parts, the EEC processes the input data and uses one of its channels to send output signal for closed loop control of actuators. THE SENSORS The FADEC sensors and transducers convert input variables into output utilizable quantities. They include: gas value sensor, engine sensors, probes, thermocouples, control sensors, signal monitoring sensors, position signals for feedback, analog signals and pressure transducers, as well as temperature transducers. The EEC can sense several position actuators and thrust lever. Engine sensors are used to sense other peripherals. The thermocouple and the RTD are installed for the purpose of sensing temperature. The pressure transducers are used to measure air pressure, which are then picked up by pressure probes at airflow stations of engine gas and through a sensing tube transfers it to the transducer in the EEC. FADEC has temperature and thrust lever angle electrical sensors that serve to measure temperature. The feedback sensors found in valves and actuators are connected to the EEC with separate cables and connectors to achieve hardware redundancy. The monitoring sensors connected to one of the EEC computers sense data for ECM. The fuel flow sensor connected to one channel is used for fuel transmission. CONTROL VALVES The FADEC has various control valves. They include: HPT Clearance control, LPT clearance Control and Servo valve, starter valves among others. The servo valves transform the analog signals into hydraulic outputs. They also allow fuel pressure that move the actuator position which, in effect, moves the position. THE ALTERNATOR The alternator, located at the accessory gearbox, is the main electrical power supply for the FADEC. This makes it independent of the power supply generated by aircraft at normal operations. It is, thus, unaffected whenever aircraft system fails. The power from the aircraft main power supply is used for engine in case of FADEC alternator fails. CONTROL UNIT FADEC contains control channels within its settings. The control channels receive sensor inputs and monitor changes. The control channels use the inputs to adjust fuel to air ratio into the cylinder, allowing every cylinder to be enriched individually. In case one of the controls channels in ECU fails, other control channels operate assigned cylinders and the opposing cylinder is used as back up for fuel ignition and injection. HEALTH STATUS ANNUCIATOR It consists of five lights on WOT and panel. It functions to provide information regarding the FADEC system state FUEL PUMP It responds by illuminating whenever fuel boost pump switch mode is on or off. Illumination of light sends a signal to show that fuel pressure is out of 10-40 range. There is the need for coordination between the FADEC system and the aircraft system. This is achieved through interfacing the aircraft system and the EEC system. Data buses are used as data channels between the aircraft system and the EEC. The EEC data buses are connected to the interfacing microprocessor in the airframe. In order to cut down on the aircraft weight, the data buses between the interface computer and the EEC are reduced due to the long distance that exists between the two channels. For the safety of the engine, in the event that EIU fails, the air data computers’ data buses connect the computers to the EIU. Data received by the EEC from air data computers includes: the temperature values, mach no, air pressure totals, the altitude values, flap lever position, aircraft air condition, aircraft bleed air demands as well as auto-thrust values. The aircraft system computers transport these data items using the Engine Interface Unit. Interface computers digitalize discrete inputs from the engine controls. This input value is switch position for engine controls, which includes: the start switch, anti ice engine switch as well as the master switch. Digitalized inputs from aircraft systems are transmitted through the interface computer into the EEC’s data buses. The power supply emanating from the ignition system and the EEC is also transferred through the EIU (Federal Aviation Administration). TASK 3 FADEC SYSTEM OF THE V2500-A5 ON THE AIRBUS A320 SYSTEM LAYEROUT DIAGRAM FIG1: The AIRBUS A320 FADEC system layout for the V2500 engine The FADEC system for V2500 is designed with EEC, the HMU and the sensors as the components that form the core of the engine system. The other components are assigned to systems such as ignition system. Forces necessary for movement of stator vanes are brought about by the use of air pressure. Control valves from the hydraulic parts are arranged centrally in HMU or decentralized around the actuators with FMU fixed in fuel distribution system. The alternator located at the accessory gear box is the main power supply. The aircraft electrical system provides alternative power supply and is used in case the alternator fails. FADEC system requires relatively low power since this power is only used for EEC, the servo valves and sensor operation. The EEC is sited at the fan case engine area where the temperature has values that are favorable during the flight. To create redundancy for the hardware, there are channels A and B with each computer having its own power supply. One channel is linked to the FADEC Alternator and the other is connected to the aircraft network. The pressure to be sensed within the EEC is picked by pressure probes at air flow station of engine gas path and then transmitted by sensing tube to the transducer in the EEC. The pressure, temperature and TLA transducers are dual so that redundancy sensors can function effectively (Massachusetts Institute of Technology, 2009). For maximum redundancy hardware, sensors are linked to EEC with separate cables. FMU is used in hydro mechanical part, and the fuel metering unit inlet supplies the actuators with fuel. The complex fuel return valve and diverter is used for heat management on V2500 engine. For this reason, an air cooled oil cooler is placed within installed fuel pressure system and air modulating valve. The fan air is used in the operation of air-cooled oil cooler. Engine pressure Ratio is used in thrust control as the main control parameter. The Engine Pressure Ratio can be found from the value of P2 and p49. The Electronic Engine Control receives EPR command during the auto-thrust operation from the system of auto flight. In case the EEC does not sense the pressure for the calculation of EPR or N1, the switch has to be pressed. While operating in N1 mode, N1 speed will be the controlling parameter so that the thrust control and engine are operated at a full thrust for all flight phases. The auto flight system sends EPR and no N1cmd values. The Electronic Engine control unit makes use of FADEC alternator frequency, enabling it to sense the N2 since there is no installed sensor for N2. Sensors found at airflow station and N1 sensors are not designed as LRUs and can be replaced in case the engine is dissembled. Installed in front of the bearing compartment is N1, which has a spare probe to connect to Electronic Engine control, should one of the probes fail. The rear face of intermediate case consists of terminal for probe connection. The ignition system is operated by an installed relay box located in front of the Electronic Engine Control. An electrically heated equipment located close to the fan is activated by the Electronic Engine control. The EEC uses the third relay and the electric power as ignition system. ENGINE SUBSYSTEMS CONTROLLED FADEC Other than power management, fuel control and EPR control, FADEC controls the subsystems of an aircraft using the EEC control software. The EEC in a closed loop manages the actuators. Position sensor is installed in every actuator which is then connected to the Electronic Engine Control so that feedback information is sent to the EEC. The required command is derived from the respective function law, and it is sent to the control section closed loop section, where the real position actuator is compared to the position commanded. In case of an error, the output current is applied to keep torque motor controlling the servo valve out of the neutrality position. Pressure generated by fuel from the servo valves causes the actuator position to move, which, in turn, moves the sensor position. Once the actuator is at the commanded position, the servo valves move to neutral position by the control section of the closed loop (Hispano-Suiza 2001). THE POWER SOURCE FOR FADEC V2500 ENGINE The V2500 engine derives its power from small alternator that is located at the accessory gear box. The alternative power is derived from the electrical system of an aircraft. The FADEC has, therefore, its own power supply generated by the alternator and it is designed for the aircraft electrical power. This is pivotal in the cases of aircraft failure, ensuring that the engine is not affected. For the hydro-mechanical part, an FMU is used and the servo valves are sited within the actuators. The two computers of the EEC’s hardware are each subdivided into 3 groups, that is, the control computer hardware, over-speed protection section and the power converter system. The power converter system provides electric power to the speed protection hardware as well as to the control computer system. The speed compressor shafts probes, the N1 and the N1 turbine probes supply power to the over-speed protection system. Usually, the subsystem does not depend on the hardware computer operations. The N1 probe added to the turbine detects the shaft fracture of the N1 which is detected whenever the compressor shaft values are compared to the turbine probe values. Other supplementary sources of powers are the pressure, temperature transducers as well as motors (Gunston, 1990). WHAT FADEC CONTROLS The EEC forms the brains of the FADEC for the V2500 engine. All other components are connected to this unit. The control software is installed and is used for engine control. The system controls the servo valves for switching off air pressure. In doing so, the EEC output side is connected to the torque mortars. Each solenoid and the torque motor have the two electric circuits, for channel A and for channel B. Only one of the two channels sent the signal for actuator control. The two control software parts operate concurrently. In both channel A and channel B, the same software is used. The two basic software parts are: the control software and the maintenance software. The software section includes: the fuel metering valve control, the EPR control, the power management control and the subsystem control. The EPR serves as the main control parameter for thrust control where the actual value for EPR can be calculated from P49 and P2. When carrying out auto thrust operations, the Electronic Engine Control receives EPR value from auto flight. Should the EEC fail to sense pressures for the calculation of EPR, the pilot has to switch N1 mode with the system operating in alternate N1 mode. In such a mode, N1 speed is the controlling parameter for the thrust control and this enables the engine to operate at full thrust range in all flight phases (Gunston, 1990). REDUNDANCY IN THE SYSTEM For maximum hardware redundancy, the sensors are all connected to the Electronic Engine Control with separate cables and connectors. Temperature sensors, pressure transducers and the TLA sensor all are dual to make up redundant sensors. This ensures that reliability of the system is increased, especially for the case of any back up for safety purposes. THE ENGINE INTERFACE UNIT (EIU) AND ITS OPERATION To help in the exchanging of data between the aircraft and the Electronic Engine Control system and the aircraft systems, it is necessary to interface the aircraft system and the engine. In this case, data are exchanged through data buses. The EEC data buses get routed to interfaced computers in the airframe (Lombardo, 2008). In V2500 engine of theA320 Aircraft, there is an installed unit for the engine. To save on weight, the data buses between the interface computer and EEC are few in number due to the relatively long distance between the two computers. In an A320 engine, the EEC is connected through five data buses to the engine interface unit. In ensuring safety in engine operations – should it happen that the EIU fails – data buses that link the data computer and the EEC are directly connected. The EEC receives Mach number, total temperature, altitude and total air pressure from the air data computer. Other exchanged data between the EEC and the air data computer are the thrust parameter, flap lever position, aircraft ground condition and the bleed air demand. All this data is sent by the aircraft system computer through the EIU. Interface computers process digitalized engine controls for the aircraft system, and the inputs are switch positions that include anti ice engine switch, master switch, mode selector and start switch. These digital inputs transferred through the EIU and via data buses into the EEC are from various other aircraft systems (Linville, 1999). The EIU is also used as a channel for power supply to the ignition system and the Electronic Engine Control from the entire aircraft system. The engine interface of AIRBUS A320 FADEC SYSTEM is shown below. FIG2: The engine interface of AIRBUS A320 FADEC SYSTEM References BTEC Unit 81 handout. Aircraft control system. Federal Avaition Administration. (2008). Pilots handbook of aeronautical knowledge. Retrieved on 3rd March, 2012 from http://www.flightlearnings.com Retrieved 2012/03/05 Gunston k. (1990) Avionics: The story and technology of aviation electronics. Welling borough: Patrick Stephens Ltd. Hispano-Suiza. (2001). Digital engine control. Retrieved on 28th February, 2012 from http://www.hispano-siza-sa.com/spip.php?article62&lang=en. Linville, R. (1999). Embry-Riddle offers aviation and aerospace education. Logistics Spectrum 33 (1): 32 -33.  Lombardo, D. (2008). Aircraft systems. New York: Donnel and Sons Company. Massachusetts Institute of Technology. (2009). Aircraft systems. Retrieved on 3rd March, 2012 from http://web.mit.edu/aeroastro/academics/grad/aircraftsystems.pdf Read More