Engine Control Panels Dornier 328Jet

Dornier 328Jet - Engines

Engine Control Panels

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MESSAGE (SYNOPTIC) WARN INHIBIT CONDITION Location (COLOR) TONE 1 2 3

L or R ENG FAIL Uncommanded engine shut down in flight (N2 2% below FI or Fuel Flow below 70 pph) or power llever set to CUT OFF position. i i CAS Field (AMBER)

L or R FADEC FAULT

CAS Field (AMBER) FADEC power interruption, sensor failure, or, software malfunction. No didispatch t h after ft llanding di or bbefore f ttake.off. k ff L or R FADEC FAULT ENGINE Page (AMBER)

L or R FADEC MAJOR Majjqor FADEC fault – No Pilot action required X X A fault is presented, no dispatch permitted. CAS Field (AMBER) L or R FADEC MAJOR ENGINE Page (AMBER)

L or R ENG EXCEEDED Normal ooperationperation pparameter(s)arameter(s) of the related enenginegine Is/are out of X X CAS Field (AMBER) range.

L or R FADEC MINOR Minor FADEC fault – No Pilot action required. X X CAS Field (BLUE) A fault is presented, no dispatch permitted. L or R FADEC MINOR X X ENGINE Page (BLUE)

L or R OIL PRS LOW Insufficient oil pressure available to allow further engine operation. X Red Warning Panel (RED)

Message inhibit logic: 1. WOW, Engines off and Electrical Bus Failure refer to section 12–31–17–04 2. Takeoff phase 3. Landing phase

CAS Field and System Messages

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Indications/Messages on EICAS Display

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NOTE: NOT ALL THE SYMBOLOGY SHOWN MAY SIMULTANEOUSLY OCCUR ON AN ACTUAL DISPLAY.

Indications/Messages on ENGINE Page

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Red Warning Panel

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RMU Engine Backup Pages

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ENGINE

GENERAL The Pratt and Whitney type PW306B Engine is a twin–spool, turbofan engine, that has a full–length annular bypass duct. The engine is built–up around a Low Pressure (LP) spool and a High Pressure (HP) spool that rotate on mechanically independent concentric running shafts. The single–stage LP compressor (fan) is mounted on the inner shaft and is driven by the three–stage LP turbine located in the hot section of the engine. The four stage and one centrifugal stage HP compressor is mounted on the outer shaft and is driven by the two–stage HP turbine which is located forward of the LP turbine. The inner LP shaft is supported on bearings No. 1 and 4 and the outer HP shaft is supported on bearings No. 2 and 3. The ENGINE SECTIONS provide information on the flow of primary air through the engine for propulsion purposes and describe the engine cold and hot sections of the engine. For information on air for non–propulsion purposes refer to SECONDARY AIR.

ENGINE SECTIONS

Primary Air, Inlet Section and Bypass Duct The primary air for propulsion purposes enters the engine through the fan case, is accelerated rearwards by the single–stage LP compressor and is split into bypass and core airflow streams through concentric dividing ducts. The bypass air passes through a single stage of stators and a faired bypass duct before exiting with the core flow through a common mixing nozzle.

Compressor Section The core airflow passes through variable inlet guide vanes and first–stage variable stator vanes, which provide optimum airflow into the HP compressor. These vanes are hydraulically actuated by fuel pressure from the hydro–mechanical metering unit and controlled by the electronic engine control. From the HP compressor, the core airflow is routed through diffuser tubes to the annulus surrounding the combustion chamber liner.

Combustion Section The air enters the annular combustion chamber liner through 24 secondary nozzles where it is mixed with the fuel. Two of these secondary nozzles are hybrids and have an additional primary nozzle which is required for engine starting. During the engine start phase, LP fuel is injected into the combustion chamber and atomized by 2 primary nozzles which are supplied by a separate primary fuel manifold. The resulting fuel/air mixture is then ignited by two spark igniters that protrude into the combustion chamber liner. When the engine speed begins to increase, the secondary fuel manifold and nozzles are pressurized to enable sustained normal engine operation.

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Turbine Section The resultant gases that expand from the combustion chamber liner flow through the first–stage HP turbine stator to the first–stage HP turbine. The first–stage HP vanes and rotor blades are cooled by air that is routed through internally cast passages. The expanding gases are then directed rearward through the engine to the second–stage HP vanes and turbine and the three–stage LP turbine and associated stator vanes.

Exhaust Section The hot gasses from the LP turbine are directed through the core exhaust to the exhaust mixer nozzle which are then mixed with the flow of cold air from the bypass duct. The controlled exit area of the nozzle accelerates the core gas velocity to produce the required thrust. The corrugated shape of the exhaust mixer nozzle optimizes the mixing of the core and fan streams to reduce engine noise levels.

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Sectional View of PW306 Turbofan Engine

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External View of PW306 Turbofan Engine

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SECONDARY AIR SYSTEM

General Secondary air for non–propulsion purposes is bled off from the HP compressor. During sustained engine operation, secondary engine generated compressed air (P2.5, P2.8 and P3) is supplied throughout the engine for various purposes such as sealing of the bearing compartments, hot section cooling, accessory cooling, anti–icing and engine control. Secondary air is also supplied in the form of bleed air to the respective aircraft services.

Sealing of Bearing Compartments To prevent oil loss from No. 1, 2, 3 and 4 bearing compartments, the seals of each compartment are externally pressurized with P2.8 compressed air. The P2.8 pressure counter–acts the oil pressure in the respective compartment to create an internal/external pressure balance and assists the seals to prevent leakage. The air pressure used to seal the bearing compartments is returned to the AGB along with the scavenge oil. The unwanted air is then vented overboard through the AGB breather.

Hot Section Cooling The hot section components that require cooling are the combustion chamber liner, HP turbine stators, HP turbine disks blades. Internal cooling is provided to 1st and 2nd stator and 1st stage blades. It is P3 compressed air which is used to cool high turbine section and P2.8 compressed air which is used to cool low turbine section.

Engine Nacelle and Engine Accessory Cooling The nacelle has various openings for air inlet and exhaust. They are in the inlet cowl, the LH and RH cowl doors and the afterbody assembly.

The inlet cowl has openings for the AC alternator air intake and for the anti–ice exhaust.

The LH cowl door has the opening fot the wing de–ici pre–cooler exhaust.

The RH cowl door has the opening for:

– The DC generator inlet and exhaust – The AC alternator exhaust – The Air Turbine Starter (ATS) exhaust – The Nacelle Cooling Air (NACA) inlet and the nacelle exhaust.

The afterbody assembly has the opening for the ECS pre–cooler exhaust.

Anti–icing For information on the internal engine anti–icing system refer to section 30–00–00 ICE AND RAIN PROTECTION.

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Engine Control P3 pressure is measured by the Engine Electronic Control (EEC) to calculate air/fuel ratio which prevents an engine surge or engine flame–out. The EEC together with the Hydro–mechanical Metering Unit (HMU) control the HP compressor variable vane actuator.

Aircraft Services Compressed air for the air conditioning system is taken from the 2.5 and 3.0 port, cabin pressurization and airfoil and nacelle anti–icing systems is provided from the P2.5 and P3.0 ports on either side of the engine. A limiting orifice is installed on the outlet of each of the P2.5 and P3.0 ports. They prevent excessive extraction of compressed air from the engine.

FADEC / EEC

General The FADEC with its integrated EEC is designed to ensure continued safe operation of the FADEC in the event of component failures. This is achieved by the configuration of dual channels where all the control and indication functions are duplicated in two independent channels and no single channel can affect the other. The two FADEC control channels are identified as Channel A and Channel B and have identical software. Only one channel can be in control of the primary output devices at any one given time. The channel that is not in control provides data for the primary engine indication displays in the flight compartment. The FADEC is configured such that either channel can fully control the engine. Should one channel become impaired, control is transferred automatically to the healthier of the two channels. Each channel receives inputs from dedicated sensors, and from the other channels sensor by means of a cross channel communication link. Therefore in the event of a dedicated sensor failure, the affected channel can continue to control using the other channels sensor input. Control of the engine is still achieved in the event that both channels have become impaired, although in a reduced capacity.

FADEC Health Status The software of each channel reads the health status of the other channel to determine which is the healthiest channel. The healthiest channel will remain in control. During each start, two serviceable fit channels alternate their controlling designation. This ensures that alternate channels control on alternate flight legs, thus minimizing fault dormancy. Based on the health status of each channel, the EICAS displays a status message which relates to the level of operability that can be provided. The health status is transferred by ARINC data between channel A and channel B so that each channel is aware of the other channels health. This ensures that information is still available on the EICAS even if one channel sustains a major failure and is unable to reliably transmit ARINC data. Based on the health status of either FADEC software channel the EICAS may display the blue status message ”L or R FADEC MINOR”. As long as the degraded system status exists, the status message will be displayed. It is allowed to fly for 150 hours after the first occurrence of the blue status message ”L or R FADEC MINOR”. Should an additional fault occur the FADEC logic assesses the new status of the electronic i t l (EEC) d if d t i d th f il t t t d d d

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FADEC Logic and CAS Field Indications For each channel, the fault data is transmitted as a 3 bit binary word Lane Failure Value. The 3 bits represent soft fail bits SF1 and SF2 and Hard Fail (HF). It is the SF1, SF2 and HF bits that provide the information for channel status message display on the EICAS. The meaning of these bits are as follows:

SF1...... = Channel minor fault SF2 or SF1 and SF2...... = Channel major fault HF...... = Channel has failed

Tables 1, 2 and 3 show the loss of operability of the control system that is displayed on the CAS field together with the implications for dispatch.

Channel A SF1 only SF2 only SF1/SF2 &/or HF Failure Value 0 1 2 > = 3 Lane Channel B 0 NO FAULTS DISPATCH DISPATCH NO DISPATCH SF1 only 1 DISPATCH DISPATCH DISPATCH NO DISPATCH SF2 only 2 DISPATCH DISPATCH NO DIS- NO DISPATCH PATCH other > = 3 NO DISPATCH NO DISPATCH NO DIS- NO DISPATCH PATCH Table 1 Operability on the Ground

Channel A SF1 only SF2 only SF1/SF2 &/or HF Failure Value 0 1 2 > = 3 Lane Channel B 0 NO FAULTS DISPATCH DISPATCH NO DISPATCH after Ldg. SF1 only 1 DISPATCH DISPATCH DISPATCH NO DISPATCH after Ldg. SF2 only 2 DISPATCH DISPATCH NO DIS- NO DISPATCH PATCH after Ldg. after Ldg. other > = 3 NO DISPATCH NO DISPATCH NO DIS- NO DISPATCH PATCH after Ldg. after Ldg. after Ldg. after Ldg. Table 2 Operability in the Air

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Channel A SF1 only SF2 only SF1/SF2 &/or HF Failure Value 0 1 2 > = 3 Lane Channel B 0 no indication MINOR MINOR MAJOR SF1 only 1 MINOR MINOR MINOR MAJOR SF2 only 2 MINOR MINOR MAJOR MAJOR other > = 3 MAJOR MAJOR MAJOR FAULT Table 3 CAS Messages (Same on ground and in the air)

The resulting blue status messages and amber cautions are subsequently shown on the CAS field in priority order:

– L – R FADEC FAULT (AMBER) – L – R FADEC MAJOR (AMBER) – L – R FADEC MINOR (BLUE)

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FADEC / EEC Engine Control and Indicating Systems Block Diagram

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ACCESSORY DRIVE The accessory gearbox is located on the underside of the intermediate case. Most of the engine–drive accessories are mounted on drive pads on the accessory gearbox. The exceptions are the N1 and N2 speed sensors and the LP/HP fuel pump. The N1 speed sensors are installed in the intermediate casing behind the LP fan. The N2 speed sensors are radially located in the RH side of the gearbox. The fuel pump is installed in the HMU, which is installed to a mounting pad on the PMA. The other engine–driven accessories are:

– The Permanent Magnet Alternator (MPA) – The DC generator – The Air Turbine Starter (ATS) – AC alternator – The oil pumps (one pressure and three scavenge) – The oil breather.

The accessory gearbox is driven by a tower drive shaft which is mated to the HP rotor shaft (N2) by a bevel gear. The tower drive shaft passes down through the intermediate casing and meshes with a bevel gear in the accessory gearbox.

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ENGINE SYNCHRONIZATION The engine synchronization system adjusts the N1 speed of one engine against the other. Depending on the performance criteria obtained from each engine at engine pass–off test, maintenance personnel select which engine will be master and which will be the slave. Selection is controlled by a switch on the maintenance panel, pilot intervention is prohibited.

The synchronization system operates in a master/slave configuration such that the master can alter the N1 speed of the slave engine by up to +/– 5 % N1. Synchronization is available between the above IDLE and MAX CLIMB TLA positions. In the air, a ’master/slave’ sync logic maintains the slave engine’s speed within a band of +/– 0.1% with respect to the master engines speed at a steady state. On the ground the slave engine speed is maintained within a band of +/– 0.4%.

The system has a capture range of +/– 5% N1, therefore if the thrust levers are set to slightly different positions, the master engine will alter N1 speed of the slave engine to achieve synchronization. If one or both thrust levers are moved such that they are out of the +/– 5% NI capture range, the synchronization system will de–activate and each engine will assume its TLA commanded N1 speed.

ENGINE VIBRATION Engine vibration is monitored by a sensor mounted on each engine and is indicated by an amber boxed ”VIB” for each engine below the N2 dials on the main EICAS display. Indication is also provided on the ENGINE page and on the RMU. On the ENGINE page, the LH and RH engine analog scales range from 0 to 1.38 in/s (inches per second). Under normal operating conditions, they each have a white pointer and range field which changes to amber when the vibration level exceeds 0.9 in/s. A digital readout is positioned above the analog scale and a white ”VIB” label is positioned between the left and right vibration scales. The digital readout ranges from 0 to 5 in/s and is amber dashed when invalid. The digital readout, the pointer and the range field of the respective engine simultaneously changes from white to amber or vice–versa in accordance with the vibration levels.

ENGINE SHUTDOWN Engine shutdown is accomplished by moving the thrust lever from IDLE to the CUT–OFF position. This will reset the HMU and the EEC monitors the shutdown phase until the engine has become stationary.

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PW 306B OVER TEMPERATURE LIMITS (EXCEPT STARTING)

OVER TEMPERATURE LIMITS (EXCEPT STARTING)

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PW 306B ROTOR OVERSPEED LIMITS

ROTOR OVERSPEED LIMITS

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