Dornier 328Jet - Engines · 2012. 6. 27. · Dornier 328Jet - Engines Page 12. FADEC Logic and CAS Field Indications For each channel, the fault data is transmitted as a 3 bit binary

Engine Control Panels

Dornier 328Jet - Engines

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MESSAGE (SYNOPTIC) WARN CONDITIONINHIBIT

Location (COLOR) TONE CONDITION1 2 3

L or R ENG FAIL Uncommanded engine shut down in flight (N2 2% below FI or Fuel Flowbelow 70 pph)

l CUT OFF i iCAS Field (AMBER)

or power lever set to CUT OFF position.

L or R FADEC FAULT

CAS Field (AMBER) FADEC power interruption, sensor failure, or, software malfunction. Nodi t h ft l di b f t k ffL or R FADEC FAULT dispatch after landing or before take.off.

ENGINE Page (AMBER)

L or R FADEC MAJOR Major FADEC fault – No Pilot action requiredX X

CAS Field (AMBER)

j qA fault is presented, no dispatch permitted. X X

L or R FADEC MAJOR

ENGINE Page (AMBER)

L or R ENG EXCEEDEDNormal operation parameter(s) of the related engine Is/are out of X X

CAS Field (AMBER)

Normal operation parameter(s) of the related engine Is/are out ofrange.

X X

L or R FADEC MINOR Minor FADEC fault – No Pilot action required.X X

CAS Field (BLUE)

qA fault is presented, no dispatch permitted. X X

L or R FADEC MINORX X

ENGINE Page (BLUE)X X

L or R OIL PRS LOWInsufficient oil pressure available to allow further engine operation X

Red Warning Panel (RED)Insufficient oil pressure available to allow further engine operation. X

Message inhibit logic: 1. WOW, Engines off and Electrical Bus Failure refer to section 12–31–17–042. Takeoff phase3. 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

GENERALThe Pratt and Whitney type PW306B Engine is a twin–spool, turbofan engine, that has afull–length annular bypass duct. The engine is built–up around a Low Pressure (LP) spool and aHigh 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 thethree–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 isdriven 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 onbearings No. 2 and 3.The ENGINE SECTIONS provide information on the flow of primary air through the engine forpropulsion purposes and describe the engine cold and hot sections of the engine. Forinformation on air for non–propulsion purposes refer to SECONDARY AIR.

ENGINE SECTIONS

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

Compressor SectionThe 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 actuatedby fuel pressure from the hydro–mechanical metering unit and controlled by the electronicengine control.From the HP compressor, the core airflow is routed through diffuser tubes to the annulussurrounding the combustion chamber liner.

Combustion SectionThe air enters the annular combustion chamber liner through 24 secondary nozzles where it ismixed with the fuel. Two of these secondary nozzles are hybrids and have an additional primarynozzle which is required for engine starting. During the engine start phase, LP fuel is injected into the combustion chamber and atomized by2 primary nozzles which are supplied by a separate primary fuel manifold. The resulting fuel/airmixture 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 arepressurized to enable sustained normal engine operation.


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Turbine SectionThe resultant gases that expand from the combustion chamber liner flow through the first–stageHP turbine stator to the first–stage HP turbine. The first–stage HP vanes and rotor blades arecooled by air that is routed through internally cast passages. The expanding gases are thendirected rearward through the engine to the second–stage HP vanes and turbine and thethree–stage LP turbine and associated stator vanes.

Exhaust SectionThe hot gasses from the LP turbine are directed through the core exhaust to the exhaust mixernozzle which are then mixed with the flow of cold air from the bypass duct. The controlled exitarea of the nozzle accelerates the core gas velocity to produce the required thrust. Thecorrugated shape of the exhaust mixer nozzle optimizes the mixing of the core and fan streamsto 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

GeneralSecondary air for non–propulsion purposes is bled off from the HP compressor. Duringsustained engine operation, secondary engine generated compressed air (P2.5, P2.8 and P3) issupplied throughout the engine for various purposes such as sealing of the bearingcompartments, hot section cooling, accessory cooling, anti–icing and engine control. Secondaryair is also supplied in the form of bleed air to the respective aircraft services. Sealing of Bearing CompartmentsTo prevent oil loss from No. 1, 2, 3 and 4 bearing compartments, the seals of each compartmentare externally pressurized with P2.8 compressed air. The P2.8 pressure counter–acts the oilpressure in the respective compartment to create an internal/external pressure balance andassists the seals to prevent leakage. The air pressure used to seal the bearing compartments isreturned to the AGB along with the scavenge oil. The unwanted air is then vented overboardthrough the AGB breather.

Hot Section CoolingThe hot section components that require cooling are the combustion chamber liner, HP turbinestators, HP turbine disks blades. Internal cooling is provided to 1st and 2nd stator and 1st stageblades. It is P3 compressed air which is used to cool high turbine section and P2.8 compressedair which is used to cool low turbine section. Engine Nacelle and Engine Accessory CoolingThe nacelle has various openings for air inlet and exhaust. They are in the inlet cowl, the LHand 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–icingFor information on the internal engine anti–icing system refer to section 30–00–00 ICE ANDRAIN PROTECTION.


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Engine ControlP3 pressure is measured by the Engine Electronic Control (EEC) to calculate air/fuel ratio whichprevents an engine surge or engine flame–out. The EEC together with the Hydro–mechanicalMetering Unit (HMU) control the HP compressor variable vane actuator.

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

FADEC / EEC

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

FADEC Health StatusThe software of each channel reads the health status of the other channel to determine which isthe healthiest channel. The healthiest channel will remain in control. During each start, twoserviceable fit channels alternate their controlling designation. This ensures that alternatechannels control on alternate flight legs, thus minimizing fault dormancy.Based on the health status of each channel, the EICAS displays a status message whichrelates to the level of operability that can be provided. The health status is transferred by ARINCdata between channel A and channel B so that each channel is aware of the other channelshealth. This ensures that information is still available on the EICAS even if one channel sustainsa 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 bluestatus message ”L or R FADEC MINOR”. As long as the degraded system status exists, thestatus message will be displayed. It is allowed to fly for 150 hours after the first occurrence ofthe 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 IndicationsFor each channel, the fault data is transmitted as a 3 bit binary word Lane Failure Value. The 3bits represent soft fail bits SF1 and SF2 and Hard Fail (HF). It is the SF1, SF2 and HF bits thatprovide the information for channel status message display on the EICAS. The meaning ofthese bits are as follows:

SF1..........................................= Channel minor faultSF2 or SF1 and SF2...............= Channel major faultHF............................................= Channel has failed

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

Lane

Channel A

Failure Value 0

SF1 only

1

SF2 only

2

SF1/SF2 &/orHF

> = 3

Channel B 0 NO FAULTS DISPATCH DISPATCH NO DISPATCHSF1 only 1 DISPATCH DISPATCH DISPATCH NO DISPATCHSF2 only 2 DISPATCH DISPATCH NO DIS-

PATCHNO DISPATCH

other > = 3 NO DISPATCH NO DISPATCH NO DIS-PATCH

NO DISPATCH

Table 1 Operability on the Ground

Lane

Channel A

Failure Value 0

SF1 only

1

SF2 only

2

SF1/SF2 &/orHF

> = 3

Channel B 0 NO FAULTS DISPATCH DISPATCH NO DISPATCHafter Ldg.

SF1 only 1 DISPATCH DISPATCH DISPATCH NO DISPATCHafter Ldg.

SF2 only 2 DISPATCH DISPATCH NO DIS-PATCHafter Ldg.

NO DISPATCH

after Ldg.

other > = 3 NO DISPATCH

after Ldg.

NO DISPATCH

after Ldg.

NO DIS-PATCHafter Ldg.

NO DISPATCH

after Ldg.

Table 2 Operability in the Air


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Lane

Channel A

Failure Value 0

SF1 only

1

SF2 only

2

SF1/SF2 &/orHF

> = 3

Channel B 0 no indication MINOR MINOR MAJORSF1 only 1 MINOR MINOR MINOR MAJORSF2 only 2 MINOR MINOR MAJOR MAJORother > = 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 CASfield 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 DRIVEThe accessory gearbox is located on the underside of the intermediate case. Most of theengine–drive accessories are mounted on drive pads on the accessory gearbox. Theexceptions are the N1 and N2 speed sensors and the LP/HP fuel pump. The N1 speed sensorsare installed in the intermediate casing behind the LP fan. The N2 speed sensors are radiallylocated in the RH side of the gearbox. The fuel pump is installed in the HMU, which is installedto 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 andmeshes with a bevel gear in the accessory gearbox.


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ENGINE SYNCHRONIZATIONThe 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 canalter the N1 speed of the slave engine by up to +/– 5 % N1. Synchronization is availablebetween the above IDLE and MAX CLIMB TLA positions. In the air, a ’master/slave’ sync logicmaintains the slave engine’s speed within a band of +/– 0.1% with respect to the masterengines speed at a steady state. On the ground the slave engine speed is maintained within aband of +/– 0.4%.

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

ENGINE VIBRATIONEngine vibration is monitored by a sensor mounted on each engine and is indicated by anamber boxed ”VIB” for each engine below the N2 dials on the main EICAS display. Indication isalso 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 (inchesper second).Under normal operating conditions, they each have a white pointer and range field whichchanges 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 positionedbetween the left and right vibration scales. The digital readout ranges from 0 to 5 in/s and isamber dashed when invalid.The digital readout, the pointer and the range field of the respective engine simultaneouslychanges from white to amber or vice–versa in accordance with the vibration levels.

ENGINE SHUTDOWNEngine shutdown is accomplished by moving the thrust lever from IDLE to the CUT–OFFposition. This will reset the HMU and the EEC monitors the shutdown phase until the enginehas 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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Documents

Dornier 328Jet - Engines · 2012. 6. 27. · Dornier 328Jet - Engines Page 12. FADEC Logic and CAS Field Indications For each channel, the fault data is transmitted as a 3 bit binary