Br710-A2 Engine Description

SECTION 2 - ENGINE DESCRIPTION

1. ENGINE INTRODUCTION

The engine has a two spool, axial flow, high bypass ratio configuration

featuring:

- Single stage wide chord fan - Ten stage HP compressor

- Annular combustor - Two stage HP turbine

- Two stage LP turbine - Full Authority Digital Engine Control (FADEC)

The basic engine is illustrated diagrammatically in Fig.1.

The low pressure shaft is supported by the No.1 thrust bearing, No.2 roller bearing and No.5 roller bearing. The high pressure shaft is

supported by the No.3 thrust bearing and No.4 roller bearing. All bearings are damped.

The lubrication system is a full flow recirculatory oil system. Primary

cooling of the oil is done by a fuel-oil heat exchanger.

2. ENGINE AIR FLOW

The output from the fan is divided into two separate flows.

One flow enters the HP compressor to be further compressed before

entering the combustion section, where fuel is added and the resultant mixture ignited. The hot gas, expands through the HP and LP turbines

before entering the multi-lobe mixer.

The other flow is directed through the bypass duct to mix with the

turbine exhaust in a common nozzle.

The mixer produces a low velocity gas stream which expands through the common nozzle to atmosphere.

A portion of the core compressor air flow is used to cool certain parts

of the engine and to pressurize oil seals, before being vented to atmosphere.

RR BR710-A2 OI RevDate: Nov 15/10


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14.07.2011https://www.aeromanager.com/pubs1/lifecdrom/xslt?nocache&xsl=https://www.aero...

Main Rotating Assemblies, Pressures and Temperature Locations.

Fig. 1

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3. INDICATED ENGINE PARAMETER INSTRUMENTATION

The following presents a summary description of the engine parameter instrumentation. The signals are processed by the

Electronic Engine Controller (EEC) and transmitted to the aircraft via an Aeronautical Radio Incorporated (ARINC) 429 digital databus for

display in the cockpit, unless otherwise stated.

A. Engine Pressure Ratio (EPR)

The Engine Pressure Ratio (EPR) is used for thrust setting and is

defined as:

Core Engine Exhaust Total Pressure (P50) = EPR

Engine Inlet Total Pressure (P20)

B. LP Compressor Speed Signal (N1)



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The LP compressor speed signal N1 is measured by pulse probes

(3 in total) mounted in the engine.

C. HP Compressor Speed Signal (N2)

The HP compressor speed signal N2 is measured by pulse

probes (3 in total) mounted in the engine.

D. Turbine Gas Temperature (TGT)

An alternative nomenclature for TGT is Inter Turbine Temperature (ITT).Thermocouples (7 dual element dual

immersion) are mounted in the LP Turbine inlet nozzle guide vanes measuring an average TGT.

E. Oil Pressure

Two strain gauge type transducers are provisioned on the

engine to measure the differential oil pressure between the oil feed pressure and the rear bearing chamber scavenge return

line.

F. Oil Temperature

Oil temperature is measured utilizing two single element Resistance Bulb Thermometers (RBTs), in the combined

scavenge return line.

G. Engine Oil Quantity

Engine oil quantity is measured by a reed switch/float type transmitter located in the engine oil tank.

H. Fuel Flow

Fuel flow is measured by a mass flow meter located on the

engine downstream of the Fuel Metering Unit (FMU).

I. Engine Fuel Temperature

Fuel temperature is measured utilizing two single element RBTs, downstream of the fuel flowmeter. An additional fuel

temperature sensor is hard wired directly to the aircraft and is



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not processed by the EEC. This single element RBT is located in

the fuel filter.

J. Engine Vibration

Engine vibration is measured using a single element natural crystal accelerometer located on the engine intermediate case.

The signal is not processed by the EEC.

4. FULL AUTHORITY DIGITAL ENGINE CONTROL (FADEC)

The engine Full Authority Digital Engine Control (FADEC) provides the necessary engine control functions and operates in association with

appropriate aircraft subsystems. An overview of the engine FADEC is shown in Fig.2.

The Electronic Engine Controller (EEC) is the major part of the FADEC

interfacing between the aircraft and the engine, and providing a means of controlling the engine. The EEC architecture is of a dual

channel type with electrical isolation between channels but with a certain degree of interchannel data communications. Either channel is

able to control the operation of the engine.

The FADEC system includes the following items:

- Electronic Engine Controller (EEC)

- Fuel pumping system - low pressure and high pressure - Hydromechanical Fuel Metering Unit (FMU)

- Independent Overspeed Protection (IOP) for N1 and N2 - Dedicated generator supplying:

- a. Three phase supply to each channel of the EEC - b. Single phase supply to each channel of the IOP unit

- Engine handling bleed valves - dual wound solenoids controlling HP5 and HP8 stage bleed valves

- HP compressor Variable Inlet Guide Vane (VIGV) and Variable Stator Vane (VSV) control - the VIGVs and the three VSV stages are

ganged into a single drive mechanism. Control is achieved through a dual wound torque motor

- Igniters and ignition control units (dual) - Data entry plug

- Starter Air Valve (SAV) - Power lever position

- Heat management system for the engine - Thrust reverser control

- Associated transducers for engine control and monitoring



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The EEC has built-in protection against overspeed, burner over

pressure and is also capable of surge detection and surge recovery.

Schematic - FADEC System

Fig.2 (Sheet 1) Print gfedcb



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Schematic - FADEC System Fig.2 (Sheet 2).

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5. ELECTRONIC ENGINE CONTROLLER (EEC)

The EEC provides the computational and interfacing means for the

operation of the Engine. The EEC is a single unit, which is bypass duct mounted and which comprises all inputs, outputs and power

conditioning.

The EEC receives digital aircraft data available from three Air Data Computers (ADC) and from three Integrated Avionics Computers

(IAC). The EEC outputs data to the two Data Acquisition Units (DAUs) The EEC system controls among other functions the following:

- Automatic engine starting,

- Fuel flow scheduling - Variable stator vanes,

- Engine handling bleed valves, - Thrust reverser (partial control only, refer to Chapter 1 Section 3.)

The EEC system is a dual channel system designed with each channel



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capable of controlling the engine. Each channel has its own input/output and processing capability and its own power supply. Both

channels are powered and continuously processing but only one channel is in control at a given time.

The EEC will act to limit N1 and N2 to their respective red lines.

For EEC SW standards C7.0.2 (SB 73-101472) and above an

automatic N1/Fan Speed Keep-Out-Zone (KOZ) logic has been introduced in the EEC software to prevent stabilised engine on ground

maintenance running between 66 and 80 percent N1 with the aircraft static and the parking brake applied. This protects the engine from

prolonged exposure to crosswind induced vibrations in forward thrust only. This logic is equally active in primary (EPR) and alternate (N1)

control mode. The EEC will activate the Keep-Out-Zone function upon detection of all the following conditions:

- Aircraft is on ground (weigth on wheels is true)

- The parking brake is engaged - Indicated airspeed is below 31 knots

- N1 is between 66 and 80 percent

If the rate of change of N1 within the Keep-Out-Zone is less than 1.4

percent/sec, the N1 is commanded to either the 66 or 80 percent N1 boundary depending on which is closer to the N1 at the time of

activation. The EEC will transit from the Keep-Out-Zone function once the N1 is commanded outside the Keep-Out-Zone.

Independent Overspeed Protection (IOP) is provided to protect

mechanical integrity such that if either N1 or N2 exceed their IOP limits the IOP will command the fuel high pressure shut-off valve to

close.

The EEC's digital transmission to the airplane includes information that defines the status of the FADEC system. This information is described

in Appendix 1.

6. THRUST MANAGEMENT

The engine thrust management system is provided by the EEC. The EEC uses three distinct control modes:

A. Primary (EPR) Control Mode

The Primary Control Mode utilizes EPR to calculate the



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powersetting for forward thrust control under steady state operating conditions. The EEC calculates an EPR command

corresponding to the actual power lever position from linear interpolation between maximum available EPR at maximum

forward Throttle RVDT Angle (TRA) and an idle EPR reference at idle TRA. Thrust is then set and held by closed loop control of

EPR. These EPR rated thrust settings are provided as a reference by the engine manufacturer.

EPR idle reference is a calculated datum. At idle TRA the engine

will be controlled to one of the minimum limiters of Shaft Speeds (N1 and N2) and HP Compressor Delivery Pressure and

Temperature (P30 and T30), not EPR. Minor variations in thrust can occur depending on environmental conditions and bleed air

demand. When the aircraft is in landing configuration the EEC will automatically raise the idle setting to high idle in order to

assure compliance with go-around requirements in the case of an aborted landing.

B. Alternate (N1) Control Mode

In Alternate (N1) Control Mode thrust is set and held by closed

loop control of low pressure shaft speed N1. N1 rated thrust setting charts are provided by the engine manufacturer. For

detailed information on N1 setting procedure refer to Chapter 4 paragraph 1. (Alternate Control Mode). N1 command is

calculated from interpolation between 99 per cent N1 at maximum forward TRA and an idle N1 reference. Similar to the

Primary Control Mode at idle TRA the engine will be controlled to an appropriate limiter. High idle will be set for aircraft in landing

configuration.

Normally the EEC operates in the Primary Control Mode. If any of the required resources for EPR control are not available the

EEC reverts to the Alternate Control Mode (soft reversion). Alternate Control Mode can also be selected by a switch in the

cockpit (hard reversion). Once reverted to Alternate Control Mode an engine will not automatically recover Primary Control

Mode even when the EEC has the required resources available for this mode. Return from Alternate Control Mode to Primary

Control Mode can only be achieved when hard reversion is deselected via the switch in the cockpit and the EEC has all the

required resources available. Therefore following a soft reversion an engine must first be hard reverted before a return to Primary

Control Mode is possible.



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When an engine soft reverts to Alternate Control Mode the EEC

ensures that the transfer to an N1 control schedule is bumpless. This is achieved within the EEC software by applying a delta to

the N1 command. N1 command is determined as a function of the power lever position and an N1 schedule. The N1 delta is

determined by the EEC as the difference between the N1 command and the measured N1 actual at the time of soft

reversion. To remove the delta the engine must be hard reverted, for which it is recommended that the pilot retard the

power lever to the idle position before hard reverting the engine. This ensures that the engine is controlled on one of the

limiters and not N1 therefore making the transition smoothly. The power lever can then be advanced to the desired setting.

C. Reverse Thrust Control Mode

Reverse Thrust Control Mode is entered upon selection of

reverse thrust. Similar to Alternate Control Mode, in Reverse Control Mode the EEC controls the engine to a N1 command as a

function of the TRA in the reverse thrust region. The reverse idle TRA corresponds to reverse N1 idle reference and the maximum

reverse TRA position equals maximum reverse N1 which is a function of inlet total temperature, altitude and calibrated

airspeed. N1 command is calculated from linear interpolation between these set points. Maximum reverse thrust is

automatically reduced from 70 to 50 percent N1 as a linear function of calibrated air speed between 65 and 50 knots

forward speed.

7. ENGINE START SYSTEM

The EEC utilizes certain command signals from the aircraft for the

control of the following functions, each of which is characterized by the degree of authority of the EEC:

- Normal autostart ground and air starts

- Alternate manual ground and air starts - Alternate manual windmilling air starts

- Cranking - Auto-Relight

- Quick-Relight

A. Normal Ground Start:

Engine starting mode on the ground in which the engine limits



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and functions are monitored and controlled by the EEC. Fuel

HPSOV control is automatic once the pilot has selected the Fuel Switch to ON.

B. Normal Air Start:

Engine starting mode in-flight which is controlled by the EEC.

Fuel HPSOV control is automatic once the pilot has selected the Fuel Switch to ON. The EEC will determine if starter assist is

required. The EEC does not limit TGT in this mode.

C. Alternate Manual Starter Assisted Start:

Engine starting mode on the ground or in-flight within the starter assisted envelope in which the pilot controls the start

sequence. Ignition and fuel HPSOV are controlled by the EEC upon commands from the cockpit. The EEC does not limit TGT in

this mode.

D. Alternate Manual Windmilling Air Start:

Engine starting within the windmill start envelope is controlled by the pilot. The EEC does not limit TGT in this mode.

E. Cranking:

Engine operating mode in which the engine is motored but

without igniters ON.

F. Auto-Relight:

Function within the EEC which upon detection of a flameout of

the engine will automatically initiate an engine relight by turning

on both igniters and controlling fuel flow. This function is not pilot selectable but is always enabled within the EEC software.

G. Quick-Relight

The EEC provides a quick relight function which provides a means of automatically relighting the engine if the Engine Run

Switch has been deselected to the OFF position and then reselected to the ON position.

8. FUEL SYSTEM



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A. General

The Engine fuel system consists of a low pressure system and a high pressure system. Fuel is supplied from the aircraft fuel

system to the LP stage of the combined LP and HP pump which is mounted on the engine accessory gearbox. The LP pump

delivers fuel to the HP pump inlet through a Fuel Cooled Oil Cooler (FCOC) and the main fuel filter. The FCOC extracts heat

from the engine oil providing oil cooling and fuel heating.

The HP pump raises the pressure further to supply the Fuel Metering Unit (FMU) which controls the fuel flow to the engine in

response to inputs from the EEC. The FMU incorporates the fuel HP Shut-Off Valve (HPSOV). Fuel is passed to the burners via

the fuel flow meter and the LP turbine Overspeed and Splitter Unit (OSU).

The FMU supplies controlled servo fuel flow to the VSV actuation

system.

Refer to Fig.3 for a schematic illustration of the fuel system.

B. LP Pump

The interface for the LP fuel inlet is on the LP fuel pump which is

mounted on the rear face of the engine accessory gearbox at the bottom of the fancase. The pump is part of the Main Engine

Pump (MEP). The LP pump is a centrifugal design with the purpose of ensuring that there is sufficient pressure available for

the HP pump.

A fuel low pressure switch is provided to monitor for low fuel pressure at the LP pump inlet. The switch is electrically

connected to the EEC and is set to give an indication of low fuel pressure at 55(+/-2) psig. Low fuel pressure is provided to the

aircraft via ARINC status words.

C. Oil Coolers

The Fuel Cooled Oil Cooler (FCOC) is situated between the LP pump delivery and the LP filter and enables the engine oil heat

to be dissipated to the fuel. Oil from the engine is used to heat the fuel to reduce ice build up in the filter and FMU.

D. Fuel Filtration



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The fuel system is protected by a 10 micron nominal (40 micron

absolute) filter in the LP system.

The filter has a bypass valve which opens at 25 psid. A pop-up

indicator on the LP filter also indicates the bypass has been operated.

The EEC monitors the pressure drop across the filter element by

means of a single differential pressure switch. The indication of impending filter blockage is given at 5(+/-2) psid. The status of

the pressure switch is transmitted via ARINC to the aircraft

A fuel filter Differential Pressure Indicator (DPI) is used to provide a visual indication to the maintenance crew that the fuel

filter by-pass valve has opened and that unfiltered fuel has entered the system. The DPI is a conventional mechanical pop-

up type indicator and will operate at 21.5 (+/-1.5) psid on increasing pressure. After the DPI has activated it can only be

reset by opening the LP filter housing. This ensures that the indication and associated maintenance action are not easily

overridden.

An additional fuel temperature sensor is hardwired directly to the aircraft and is not processed by the EEC.

The HP filter is the final protection for the burners. The filter is a

cleanable strainer type adjacent to the fuel manifold on the combustion chamber.

The servo supply for the Variable Stator Vane (VSV) unit is

taken directly from the HP pump/FMU unit via an internal flow washed filter.

E. HP Pump

This is a gear pump in the MEP and receives its drive direct from

the engine accessory gearbox. The incoming fuel from the LP pump is fed via the FCOC and fuel filter to the HP pump inlet.

The high pressure fuel is passed to the FMU which is mounted on the MEP.

F. Fuel Metering Unit



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The FMU has eight functions:

1. Fuel metering

2. HP Shut-Off Valve (HPSOV)

3. Pressure raising

4. Overspeed shut-off

5. Pump unloading

6. VSV control valve

7. Manifold drains switching

8. Spill Divereter Valve (SDV)

The Engine fuel schedule is controlled by this unit as commanded by the EEC via the position of a metering valve. A

Spill Diverter Valve (SDV) is incorporated, for heat management purposes, controlling the return of excess fuel delivered by the

HP pump to either the FCOC inlet or HP pump inlet. At high engine power the spill diverter closes off the return flow to the

FCOC inlet.

The HPSOV, as part of the FMU, is used for starting and shut-down. During a start when the HPSOV is commanded open, the

torque motor also cancels the pump unloading function in the FMU. The HPSOV maintains a minimum HP pump pressure rise.

This ensures that over the engine operating range there is sufficient pressure to power the fuel servo systems. With the

HPSOV commanded closed the HP pump pressure rise is maintained sufficiently high to allow servo system control.

Closure of the HPSOV also opens a passageway from the splitter valve to the drains tank. The declining air pressure within the

engine is sufficient to purge any remaining fuel in the nozzle manifolds back to the drains tank.

G. Fuel Drains Tank

The tank is mounted at the bottom of the structural bypass duct to provide a collection point for fuel from the fuel manifold on

HPSOV closure. An ejector pump, operated by low pressure fuel,



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returns the contents of the drains tank to the LP pump inlet. A

float within the tank prevents the ingress of air into the system when the level falls. The tank is sized to accommodate the

amount of fuel collected as a result of three engine shut-downs. If the tank become full for any reason an overflow pipe carries

surplus fuel overboard.

H. LP Turbine Overspeed Fuel Shut-Off/Splitter Unit (OSU)

The emergency shut-off valve in the OSU is mechanically actuated in the event of a LP shaft failure and closes the fuel

line to the burners. It is located close to the fuel manifold adjacent to the combustion chamber.

It also contains the splitter function which directs fuel flow to

the upper and lower manifold segments. This is to overcome the head effect in the manifold at low fuel flows.

I. Servo Systems

HP fuel is used as servo power to the following systems:

- HP compressor variable stator vanes

- Fuel metering valve



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Schematic - Fuel System Fig. 3

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9. IGNITION SYSTEM

The ignition system comprises two ignition units, two igniter leads and

two surface discharge semi conductor igniters.

The ignition system is controlled by the EEC, each channel of the EEC has the capability of energizing either or both ignition units.

Continuous ignition is available by pilot selection.

Although selection of continuous ignition is not time limited it will

reduce overall igniter life, as such it is not recommended.

The inclement weather protection system will activate continuous ignition automatically.

10. ENGINE BLEED AIR SYSTEM

Engine bleed air is used for the following functions:



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- Aircraft systems - Inlet cowl anti-icing

- Cross bleed starting - Pre-cooling of air for aircraft systems

- Handling bleed valve actuation

A. Aircraft Service Bleeds (Fig.4)

Air is supplied for aircraft systems, from tappings on the HP

compressor 5th and 8th stages. HP5 compressor air is normally used for aircraft systems. However, at low engine power

settings or bleed air temperatures HP5 air pressure is insufficient to maintain the aircraft services and HP8 compressor

air is used.

Selection of HP5 or HP8 compressor air is controlled by the Bleed Management Computer via a pressure shut-off valve in

the HP8 supply pipe. A non return valve in the HP5 supply prevents recirculation of HP8 air. The total air supply pressure to

the aircraft systems is controlled by a pressure regulating and shut-off valve in the combined HP5, HP8 air supply duct.

B. Engine Handling Bleed

The HP compressor is equipped with three 5th stage and one

8th stage handling bleeds for engine starting, stability, surge recovery and flame out protection in inclement weather. Bleed

valves have two positions (open or closed) only. They are actuated by internal springs and 10th stage HP compressor air.

The fail-safe and engine not running position is open for handling bleed valves.

The EEC schedules each bleed valve independently as a function

of HP shaft speed N2 and HP compressor inlet total temperature T26. The bleed valves are open during start and close

successively with increased powersetting above idle. Different bleed schedules are provided for steady state and transient

operation of the engine.

For surge recovery bleed valves are commanded open when the onset of a surge is detected by the EEC.



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If inclement weather is detected bleed valves are commanded

open to cater for engine stability and to allow water to pass into the bypass flow.

Schematic - Bleed Air System. Fig. 4

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11. ICE PROTECTION

The power plant ice protection comprises engine de-icing and inlet anti-icing.

Engine de-icing is accomplished by the design of the fan spinner

ensuring that ice accretions are shed centrifugally before they build to potentially damaging levels.

Inlet anti-icing relies on hot air supplied by the engine's high pressure

compressor, which is ducted to the inside of the inlet cowl inlet lip. The system is activated by cockpit switches.

12. ENGINE OIL SYSTEM

The engine oil system comprises the following main components:



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- Oil tank and de-aerator

- Pressure pump

- Pressure filter

- Fuel Cooled Oil Cooler (FCOC)

- Scavenge pumps - Breather

Fig.5 shows a schematic of the engine oil system.

The oil tank is an integral part of the accessory gearbox and is located on the lower left-hand side of the engine. Oil filling can be achieved

either through a conventional gravity filling cap or through a remote replenishing system. Oil quantity indication is achieved by an

internally mounted transmitter and also by a sight glass. Oil quantity information will be continuously available to the aircraft. To determine

the amount of oil required to fill the system, the oil quantity indication is not valid until 5 minutes after shut-down on ground. Replenish the

engine oil system between 5 and 30 minutes after engine shut-down.

From the oil tank the oil passes to the pressure pump element of the oil pump unit. From the pressure pump the oil flows through the oil

filter which is mounted on the pump unit and then to the FCOC which cools the oil. The oil then passes to the front and rear bearing

chambers and the accessory gearbox.

The oil pressure is measured as a differential between the main supply line pressure and the pressure in the rear bearing chamber scavenge

line. Two transducers measure this pressure difference (one transducer for each channel of the EEC) and at a set limit a low oil

pressure warning is transmitted to the aircraft.

Magnetic chip detectors are mounted at the gearbox in the front and rear bearing chamber and gearbox scavenge lines.

The oil is removed from the front and rear bearing chambers,

accessory gearbox and breather by four scavenge pump elements which are part of the oil pump unit. The scavenge pumps return the



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oil to the oil tank through a de-aerator.

The temperature of the oil is measured in the combined scavenge line by two transducers (one transducer for each channel of the EEC).

The bearing chambers, accessory gearbox and oil tank are all vented

to the breather in the accessory gearbox.

13. ENGINE INDICATING AND CREW ALERTING SYSTEM (EICAS)

The aircraft EICAS system is part of an integrated electronic display system which is used for providing propulsion system indication and

alerts to the crew.

A list of engine parameters, indications and recommended crew actions for safe operation of the engine is provided in Appendix 1.

Schematic - Oil System

Fig. 5 Print gfedcb



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Br710-A2 Engine Description

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