Mi-8MTV2 Magnificent Eight. PC flight simulator manual (2016)

all modes of helicopter operation. Service fuel cell Left saddle tank Right saddle tank

Auxiliary tank foam-filled

7.2.1. Fuel Storage Location

Onboard helicopter, fuel is located in three main fuel tanks lined with self-sealing

polyurethane (PU) foam: two saddle tanks are located on either side of fuselage and

one service fuel cell is located in container behind main gear, Fig. 7.5

Fig. 7.5. Location of Fuel Systems Units

1. Service fuel cell

3. Left auxiliary tank (not implemented in the

2. Right saddle tank

simulator)

4. Left saddle tank

For increase in flight range and duration, one or two auxiliary fuel tanks may be

installed inside fuselage (not implemented in the simulator).

Fuel tanks capacity in ltr/kg:

service fuel cell with foam-filled

415/322

right saddle tank with foam-filled

1040/832

left saddle tank with foam-filled

1130/904

auxiliary tank with foam-filled

895/694 (not modeled)

7.2.2. Fuel Distribution System

Reliable operation of fuel system is ensured by pumps, valves, pressure sensors,

solenoid and shut-off valves, Fig. 7.6

7. Breating line

16. Drain valves

8. Pump 463Б

17. Saddle tanks

9. Pressure detectors СД-29А

18. Float valve 766300А-1

A - Fuel feed line

B - Breathing lines

Helicopter fuel distribution system comprises the following: electric driven centrifugal

fuel pump 463Б (8); two centrifugal fuel pumps ЭЦН91С (15); electric driven fuel

pump 748Б (of heater КО-50, not shown); float valve 766300A-1 (18); five shut-off

cocks 768600MA (6, 13); one shut-off cock 633600A (14); two solenoid valves

610200А (2, 3) for feed lines of engine АИ-9 and heater КО-50 accordingly; bypass

cock 637000 (12); fuel filter 11 ТФ30 СТ (not shown); check valves block (not

shown); pipes and hoses.

From saddle tanks, fuel is fed to service fuel cell by two pumps ЭЦН-91С (15) via

pipelines; pump 463Б (8) distributes fuel from service fuel cell for feeding of engines

ТВ3-117ВМ. In the lines running from pump 463Б to engines, emergency shut-off

cocks 768600 МА (4) are fitted.

For feeding of engine АИ-9В and kerosene-combustion heater fuel is taken from the

line running to the right engine ТВ3-117ВМ, upstream the emergency shut-off cock.

This scheme (Fig. 7.6) does not contain the engines fuel shutoff levers, which are

opened / closed by crew for engines start / shut down.

Function of Fuel System Units

ELECTRIC DRIVEN CENTRIFUGAL FUEL PUMP 463Б (8) creates fuel overpressure at the inlet of main

fuel pumps of engines ТВ3-117ВМ and engine АИ-9В, and inlet of kerosene-combustion heater КО-50

pump 748Б as well.

Pump 463Б is located outside service fuel cell, in its lower part.

Pump 463Б failure does not lead to interruption of engines operation: fuel would be fed from service

fuel cells to engines by gravity.

Pump power supply circuit is connected to battery bus via circuit breaker ТОПЛИВНАЯ СИСТЕМА.

НАСОСЫ ТОПЛИВН. БАКОВ-РАСХОД (FUEL SYSTEM. FUEL TANKS PUMPS - FLOW) located on the

breakers’ right section of cockpit electric panel.

Pump is activated by switch НАСОСЫ БАКОВ-РАСХОД (TANKS PUMPS - FLOW) installed on the mid

section of electric panel.

ELECTRIC DRIVEN CENTRIFUGAL PUMPS ЭЦН-91С (15) located in the mounting units inside saddle

tanks (in the front part of left tank and in the rear part of right tank) pump fuel to service fuel cell.

Pumps motors power supply circuits are connected to battery bus via circuit breakers ТОПЛИВНАЯ

СИСТЕМА. НАСОСЫ ТОПЛИВН. БАКОВ-ЛЕВОГО (FUEL SYSTEM. FUEL TANKS PUMPS - LEFT) and

ТОПЛИВНАЯ СИСТЕМА. НАСОСЫ ТОПЛИВН. БАКОВ-ПРАВОГО (FUEL SYSTEM. FUEL TANKS PUMPS

- RIGHT) on the breakers’ right section of cockpit electric panel. Pumps are turned on by switches

НАСОСЫ БАКОВ-ЛЕВЫЙ (TANKS PUMPS - LEFT) and НАСОСЫ БАКОВ-ПРАВЫЙ (TANKS PUMPS -

RIGHT), on the mid section of cockpit electric panel.

ELECTRIC DRIVEN FUEL PUMP 748Б (11) of geared type for fuel feed to kerosene-combustion heater

КО-50 injectors, installed in kerosene-combustion heater bay.

Pump power supply and control circuit is connected to rectifier bus via circuit breaker of heater КО-50,

located on the breakers right section of cockpit electric panel. Pump activates after heater activation.

FLOAT VALVE 766300А-1 (18) prevents overfill of service fuel cell when fuel is pumped from saddle

tanks; it is installed in service fuel cell and fixed to its plate.

During pumping, when service fuel cell is not yet filled completely, valve goes off its seat under fuel

pressure thus opening passage area for fuel flow. Fuel goes to the tank via valve body openings and

fills the tank. After tank is filled the valve float takes top position and shuts fuel supply to the tank.

EMERGENCY SHUT-OFF COCKS 768600А (4) are operated remotely via electric circuits; they serve

for shutting and opening of fuel lines. The Emergency Shut-off Cocks inslalled in the main gear part

(see Fig. 7.7 1, 2). Designed to shutoff of fuel lines in case of fire.

Emergency shut-off cocks power supply circuits are connected to battery bus via circuit breakers

ТОПЛИВНАЯ СИСТЕМА. ПЕРЕКРЫВ. КРАНЫ-ЛЕВЫЙ (FUEL SYSTEM. SHUT-OFF COCKS - LEFT) and

ТОПЛИВНАЯ СИСТЕМА. ПЕРЕКРЫВ. КРАНЫ-ПРАВЫЙ (FUEL SYSTEM. SHUT-OFF COCKS - RIGHT) on

the breakers’ right section of cockpit electric panel. The cocks are opened and shut by switches

ПЕРЕКРЫВ. КРАНЫ-ЛЕВЫЙ (SHUT-OFF COCKS - LEFT) and ПЕРЕКРЫВ. КРАНЫ-ПРАВЫЙ (SHUT-OFF

COCKS - RIGHT), which are located on the electric panel mid section and safeguarded. Shut position

of cocks is indicated by lamps ЛЕВЫЙ ЗАКРЫТ (LEFT SHUT) and ПРАВЫЙ ЗАКРЫТ (RIGHT SHUT),

located below those switches.

TWO FUEL SHUTOFF LEVERS are installed in the fuel lines of engines TV3-117VM, they cut off fuel

(handle) flow upstream engines inlet. It necessary open/close in case start/stop engines by crew

members.

ONE SHUT-OFF COCK (6) serves for fuel bypassing from saddle tanks to service fuel cell in case of

failure of float valve 766300А-1 in shut position. The cock is installed on the service fuel cell plate.

Shut-off cock power supply circuit is connected to battery bus via circuit breaker ТОПЛИВНАЯ

СИСТЕМА-КРАН ПЕРЕПУСК (FUEL SYSTEM - BYPASS COCK) located on the right section of cockpit

electric panel. The cock is opened and shut by switch ТОПЛИВНАЯ СИСТЕМА-ПЕРЕПУСК FUEL

SYSTEM - BYPASS) located on the mid section of electric panel.

TWO SHUT-OFF COCKS are installed in tanks cross-feeding lines that interconnect saddle tanks in

their front and rear parts. They serve for uniform fuel use from saddle tanks in case of failure of either

pump ЭЦН-91С.

Normal position of cocks is open.

The cocks shall be shut when helicopter enters hazardous zone, to save fuel in one saddle tank if

another gets damaged.

Cross-feeding cocks power supply circuits are connected to battery bus. The cocks are opened and

shut by safeguarded switch КОЛЬЦЕВ БАКОВ ЗАКР-ОТКР (TANKS CROSS-FEEDING OPEN-SHUT)

located on the mid section of electric panel. Shut position of cocks is indicated by yellow lamp

КОЛЬЦЕВ ОТКЛ. (CROSS-FEEDING OFF) below the switch.

SHUT-OFF COCK 633630 (not implemented in the simulator) shuts front cross-feeding line. The cock

is controlled manually and should be open. The cock is shut in case of right tank removal and fuel

drain from auxiliary tanks.

SOLENOID VALVES 610200А ensure control of fuel feed to engine АИ-9В and kerosene-combustion

heater КО-50. The solenoid valve in the fuel line of engine АИ-9В is installed in the main gear

compartment. It opens automatically after start button of engine АИ-9В is pressed. Cock closes when

engine АИ-9В shuts down.

The solenoid valve in the fuel feed line of kerosene-combustion heater КО-50 is installed on the cargo

compartment ceiling panel below emergency shut-off cock; it opens automatically when heater is

started and shuts when heater is turned off.

BYPASS COCK 637000 (not implemented in the simulator) serves for connection of one or two

auxiliary tanks to the front line connecting the saddle tanks, and for fuel drain from auxiliary tanks.

154

The cock is opened manually. It is installed under cargo compartment deck close to shut-off cock

633600А. The cock ensures fuel use from auxiliary tanks: either separate or simultaneous.

CHECK VALVES BLOCK includes two check valves installed in the lines of fuel transfer from saddle

tanks to service fuel cell. The valves pass fuel in one direction only - to service fuel cell. They are

installed on the service fuel cell plate upstream float valve and shut-off cock.

The system of pipelines and check valves ensures fuel feed of engines from either pump of saddle

tanks, if another fails.

PRESSURE DETECTORS СД-29А (SD-29A) turn off the lamps, which indicate running of transfer

pumps ЭЦН-91С or pump 463Б if pressure in relevant line drops below 0.5 kg/cm².

7.2.3. Monitoring and Control of Fuel System Operation

Fuel System Units Control

The units of fuel system are controlled from the fuel system control panel located on

the mid section of electric panel, Fig. 7.7. Moreover, there are manual shut-off cocks

that are not implemented in the simulator.

Fig. 7.7. Fuel System Control Panel, Center overhead console

1. Indicator lamp of left fuel shut-off cock and

6. Switch and indicator lamp of RIGHT tank

its closed position

ЭЦН-91С pump operating status

2. Indicator lamp of right fuel shut-off cock

7. Switch and indicator lamp of LEFT tank

and its closed position

ЭЦН-91С pump operating status

3. Indicator lamp of fuel cross-feeding

8. Switch and indicator lamp of FEED tank

disabling (interconnecting line is shut)

463Б pump operating status

4. Switch of saddle tanks cross-feed enabling

9. Switch for selecting operating mode of

solenoid valves (top position: interconnecting

lamps indicating 100% fill of tanks. The lamps

line is open - ОТКРЫТА)

are located close to fillers, not implemented in

5. Switch of solenoid valve for manual fuel

the simulator

bypassing from saddle tanks to service fuel

cell (in case of service fuel cell float valve

failure). Normal position is ЗАКРЫТО

(CLOSED) - down

Fuel System Operation Monitoring

Operation of fuel system may be monitored by pumps’ indicator lamps and fuel

gauge, which indicates remaining fuel quantity in litres, separately or in total (for

saddle tanks and service fuel cell), Fig. 7.8:

Fig. 7.8. Fuel gauge

1. Inner scale, for reading of fuel quantity in

2. Outer scale for reading total fuel quantity

separate tanks (highlighted red)

(highlighted blue)

For tanks switching over, fuel gauge switch, Fig. 7.9 is used:

156

Fig. 7.9. Fuel Gauge Switch

Fuel gauge switch positions

3. РАСХ - Service fuel cell

1. Пл - Left Saddle tank

4. Д - Auxiliary Tank (not implemented in the

2. Ппр - Right Saddle tank

simulator)

5. СУММА - total fuel quantity in saddle tanks

and service fuel cells

7.2.4. Normal Operation

Before Start

1. Set the following circuit breakers to on: ТОПЛИВНАЯ СИСТЕМА.

ТОПЛИВОМЕР (FUEL SYS FUEL GAUGE); НАСОСЫ ТОПЛИВН. БАКОВ-РАСХОД-

ЛЕВОГО-ПРАВОГО (FUEL TANKS PUMPS - FLOW - LEFT - RIGHT); ПЕРЕКРЫВ

КРАНЫ ЛЕВЫЙ-ПРАВЫЙ (SHUT-OFF COCKS LEFT - RIGHT), КРАН ПЕРЕПУС.

(BYPASS COCK) on the breakers’ mid section

. At

that, lamps ЛЕВЫЙ ЗАКРЫТ (LEFT CLOSED), ПРАВЫЙ ЗАКРЫТ (RIGHT CLOSED)

on the mid section of electric panel come on (Fig. 7.7, 1, 2);

2. Check fuel quantity using fuel gauge indicator;

3. Check functioning of backing up and transfer pumps by alternate training of

them: make sure that lamps РАСХОД РАБОТАЕТ (FLOW ON), ЛЕВЫЙ (ПРАВЫЙ)

РАБОТАЕТ (LEFT (RIGHT) ON)

come on;

4. Prior to starting auxiliary power unit and main engine, turn on the fuel

gauge, backing up and transfer pumps to check functioning of them by relevant

indicator lamps coming on; open emergency shut-off cocks: indicator lamps ЛЕВЫЙ

ЗАКРЫТ (LEFT CLOSED) and ПРАВЫЙ ЗАКРЫТ (RIGHT CLOSED)

must go off.

Operation in Flight

Fuel pumped from outboard pumps goes to service fuel cells by two fuel lines via

check valves and float valve; from service fuel cell, fuel is fed to engines ТВЗ-117ВМ

via open emergency shut-off cocks.

In flight, it is required to monitor lamp indicating pumps operation and remaining

fuel in service fuel cell (once in 5 to 10 min).

7.2.5. Failures

Indications of failure and appropriate actions.

Failure of service fuel cell float valve (not implemented)

Failure of service fuel cell float valve has the following indication:

 with saddle tanks’ pumps running, fuel quantity in service fuel cell

decreases.

Crew Procedure

 Take decision about continuing flight.

 Set switch ПЕРЕПУСК (BYPASS)

on the panel ТОПЛИВНАЯ

СИСТЕМА (FUEL SYSTEM) (Fig. 7.7, 5) to ОТКР. (OPEN, up): fuel will be

delivered to service fuel cell through open shut-off cock, with float valve

bypassing. Then, further fuel use should be adjusted manually while

maintaining fuel quantity in service fuel cell at the level of 370...390 litres

and avoiding its overfill.

158

Failure of One/Both Transfer Pump

Failure of transfer pumps has the following indications:

 indcator lamp НАСОСЫ БАКОВ (TANKS PUMPS): ЛЕВЫЙ РАБОТАЕТ

(LEFT RUNNING) or ПРАВЫЙ РАБОТАЕТ (RIGHT RUNNING)

goes off, or both lamps go off;

 voice recorder message "Отказали насосы основных топливных баков"

(Main fuel tanks’ pumps failure);

 fuel quantity reduction in service fuel cell.

Crew Procedure

IF ONE TRANSFER PUMP FAILS:

 Turn off the failed pump.

 In case of either pump failure, another pump ensures service fuel cell fill.

 Make sure that fuel is fed to service tank and continue your mission.

IF BOTH TRANSFER PUMPS FAIL:

 Turn off НАСОСЫ БАКОВ ЛЕВЫЙ, ПРАВЫЙ (TANKS PUMPS LEFT, RIGHT)

 Abort the mission.

 Land on the nearest airfield or an appropriate site noticed in flight.

WARNING. Please note that fuel quantity in service fuel cell is sufficient for flight at altitude 500 m

and speed 220 km/h for 21 min to the range of 70 km.

Service fuel cell Backing Pump Failure

Failure of service fuel cell backing pump has the following indications:

 short-term drop of engines RPM by 2-5%, fuel pressure by 3-4 kgf/cm²

and main rotor RPM by 1-3%;

 green indicator lamp РАСХОД РАБОТАЕТ (FLOW ON)

on the

mid section of electric panel goes off;

 voice recorder message "Отказал насос расходного бака" (Service fuel

cell pump fails).

Crew Procedure

 if backing pump fails, continue the mission. At that, normal operation of

engines are ensured by pumps of engine ДЦН-70А.

 turn pump off

 act smoothly when changing engine operating parameters and helicopter

piloting.

7.3. Hydraulic System

7.3.1. Brief Description

Helicopter has two independent hydraulic systems:

 main system;

 backup system

Fig. 7.10. Location of Hydraulic System Units

1. Hydraulic units’ panel

5. Backup system hydraulic pump НШ-39М

2. Hydraulic boosters

6. Hydraulic lock cylinder

3. Main system hydraulic pump НШ-39М

7. Collective pitch control clutch dumping

4. Charging connections panel

cylinder

(1) HYDRAULIC UNITS CYLINDER serves for compact location of both hydraulic systems

units (solenoid valves, pressure accumulators, filters, check valves);

(2) HYDRAULIC BOOSTERS convert manual control rods motion autopilot signals into

power rod travel;

(3) HYDRAULIC PUMP НШ-39М of main system serves for creation of pressure in main

system line;

(4) CHARGING CONNECTIONS PANEL is used for charging of hydraulic systems with power

fluid and check of those system from ground unit (not implemented in the

simulator);

160

(5) HYDRAULIC PUMP НШ-39М of backup system serves for creation of pressure in

backup system line;

(6) HYDRAULIC LOCK CYLINDER serves for force creation on cyclic pitch control stick

when swash plate angle reaches 2°12'. Force makes 16 kg (not implemented in the

simulator);

(7) COLLECTIVE PITCH CONTROL CLUTCH DUMPING CYLINDER serves for release of collective

pitch control forces, when pressing on collective pitch control button (not

implemented in the simulator)

Main Hydraulic System

Main hydraulic system serves for feeding of combined control units (hydraulic

boosters) КАУ-30Б (KAU-30B, installed in longitudinal, lateral and collective pitch

control systems) and РА-60Б (RA-60B, lateral control system), collective pitch control

clutch dumping cylinder, variable lock cylinder (lateral control).

Hydraulic boosters КАУ-30Б and РА-60Б can operate in two modes:

 manual control (by pilot);

 combined control (autopilot on).

Backup Hydraulic System

Backup hydraulic system duplicates main hydraulic system; it performs main system

functions if that fails. Backup system is activated automatically, if main system

pressure drops to 30±5 kgf/cm².

In case of main system failure and backup system activation, the following units are

deactivated automatically: autopilot АП-34Б (AP-34B), collective pitch control clutch

dumping system and hydraulic lock. At that, hydraulic boosters are operated in

manual mode only; to create optimum forces for collective pitch control movement,

clutch tightening is to be adjusted (not implemented in the simulator).

Control of Hydraulic Systems

Hydraulic systems are controlled from hydraulic system panel, Fig. 7.11

Fig. 7.11. Hydraulic System Control Panel

1. Main system switch

5. Backup system pressure gauge

2. Main system pressure gauge

6. Backup system switch

3. Lamp indicating activation of main system

7. Backup system deactivation button

4. Lamp indicating activation of backup system

Button ОТКЛ.ДУБЛИР. (DEACTIVATE BACKUP SYS) (Fig.

7.11,

7) for stable

switching from backup system to main one during engine start and hydraulic system

ground check (ground check is not implemented in the simulator).

Technical details of hydraulic systems are given in Table 7.1.

Table 7.1

Power fluid

масло АМГ-10

Operating pressure in main and backup hydraulic systems



(4500± 300...650

) kPa





[(45±3…6

) kgf/cm²]

2

Ambient air operation range for normal operation of hydraulic

от -50 до +60 °С

systems

Power fluid permissible temperature

до 70°С

162

АМГ-10 oil quantity in hydraulic system

22 l (по 11)

System pressure for pump switching to operating mode

(4500±300) kРа [(45±3)

(hydraulic system feeding)

kgf/cm²]

System pressure of pump switching to idle mode (fluid pumping



(650

) kРа [(6

)

200

2

to tank)

kgf/cm²]

Minimum pressure in main system for hydraulic boosters feed

(3000±500) kPa [(30±5)

switching to backup system

kgf/cm²]

7.3.2. Normal Operation

Check before starting the engine

Set circuit breakers ГИДРОСИСТ. ОСНОВН/ДУБЛИР. (HYDRO SYS MAIN/BACKUP)

to on. Make sure that hydraulic systems switches are set to ВКЛ (ON)

(main

and backup

hydraulic systems are activated), and backup

system switch ГИДРОСИСТЕМА ДУБЛИР. (HYDRO SYS BACKUP) is safeguarded.

During Engine Start

Warning! During engine start with main and backup systems switches set to on,

backup system may activate (not main one). At backup system pressure of

(2500±160) kPa [(25±1.6) kgf/cm²] indicator lamp ДУБЛИР ВКЛЮЧЕНА (BACKUP

SYS ON)

must come on. (At that, electromagnetic relay deactivates main

hydraulic system). Pressure read by pressure gauges of both systems will be

stabilized within (6500) kPa [(65) kgf/cm²]. To switch to operation from main

hydraulic system, press button ОТКЛ ДУБЛИР (DEACTIVATE BACKUP SYS)

and hold it for 1…1.5 s. Indicator lamp ОСНОВНАЯ ВКЛЮЧЕНА (MAIN SYS ON)

must come on and indicator lamp ДУБЛИР ВКЛЮЧЕНА (BACKUP SYS ON) must go

off. Pressure in backup system shall drop to (0-500) kPa [(0-5) kgf/cm²].

Prior to hydraulic system serviceability check, it is required to switch to main system;

for that, press button ОТКЛ. ДУБЛИР. (DEACTIVATE BACKUP SYS) on the mid

section of cockpit electric panel and hold it until green indicator lamp ОСНОВН.

ВКЛЮЧ. (MAIN SYS ON)

comes on and red indicator lamp ДУБЛИР.

ВКЛЮЧ. (BACKUP SYS ON) goes off.

Make sure that main system pressure increases to 42... 73 kgf/cm²

. At

that, one should pay attention to frequency of main system pressure gauge indicator

movement.

Hydraulic System Check During Power Plant Warm-up

Check operation of hydraulic system controls; for that:



deflect control stick, pedals and collective pitch control lever alternately,

by no more than 1/3 of their full travel, to make sure that controls operate

smoothly (no jerks or sticking);



make sure that operation of controls causes main system pressure

variation within 42...73 kgf/cm², no pressure in backup system, green

indicator lamp ОСНОВН. ВКЛЮЧЕНА (MAIN SYS ON) is on;



Set switch (MAIN HYDRO SYS) to ВЫКЛ (OFF). Pressure in main system

(downstream valve) will drop to zero; at the moment it passes the value

(3000±160) kPa [(30±1.6) kgf/cm²] main system pressure detector will

operate, indicator lamp ОСНОВНАЯ ВКЛЮЧЕНА (MAIN SYS ON) goes off;

when pressure becomes (3000±500) kPa [(30±5) kgf/cm²] emergency

feeding valve will operate. Pressure in backup system must go up; when

pressure reaches (2500±1.6) kPa [(25±1.6) kgf/cm²] the backup system

pressure detector will operate and indicator lamp ДУБЛИР ВКЛЮЧЕНА

(BACKUP SYS ON) must come on. Pressure in backup system grows

800



rapidly and changes within (4500±300...6500

) kPa [(45±3...6

)

200

2

kgf/cm²], when controls are moved;



set main system switch to on and make sure that green indicator lamp

ОСНОВН. ВКЛЮЧЕНА (MAIN SYS ON) comes on, pressure in main system

is maintained within

42…73 kgf/cm², red indicator lamp ДУБЛИР.

ВКЛЮЧЕНА (BACKUP SYS ON) goes off and pressure in backup system

drops to zero. For the helicopters with updated electric circuit of hydraulic

system: to switch to main system, after setting ГИДРОСИСТЕМА ОСНОВН

(HYDRO SYS MAIN) to on, it is required to press button ОТКЛ. ДУБЛИР.

(DEACTIVATE BACKUP SYS) on the mid section of cockpit electric panel

and hold it until green indicator lamp ОСНОВН. ВКЛЮЧЕНА (MAIN SYS

ON) comes on and red indicator lamp ДУБЛИР. ВКЛЮЧЕНА. (BACKUP

SYS ON) goes off.

7.3.3. Failures

For indications of failures and appropriate actions refer to 10.7

164

7.4. Anti-Ice System

7.4.1. General Description

Helicopter Anti-Ice System serves for icing prevention and ice/water removal from

helicopter units (Fig. 7.12) and indication of icing as well. The protected elements

are as follows:

 two front glasses of cockpit;

 air intakes including air inlet Particle Separator System (PSS) and engines’

inlets;

 main and tail rotor blades.

Fig. 7.12. Anti-Ice System

1. Heated glasses

3. Heated parts of main rotor blades

2. Heated parts of air intakes including Particle

4. Heated parts of tail rotor blades

Separator System and engines’ inlets

For icing early warning, helicopter is furnished with ice detector РИО-3 and visual ice

detector.

7.4.2. Heated Glasses

Glasses are electrically heated.

Anti-ice system of glasses provides for automatic energizing of glasses heating by

СО-121ВМ ice detector signal and manual switching to glasses heating and wipers

also. Glasses heating temperature (30°С) is maintained by temperature regulators

ТЭР-1М.

Technical details of heated glasses:

Supply voltage, V

190; 208; 230 or 250

Power demand, W, max.

1930

Current consumption, A, max.

9.65

For adjustment of currents consumed by glasses heating elements, rotary switch

ТОКИ ПОТРЕБИТЕЛЕЙ (CONSUMERS CURRENTS) (Fig. 7.18) is to be set to

ОБОГРЕВ СТЕКОЛ (GLASSES HEATING), along with current check by dashboard

amperemeter. Dashboard amperemeter readings should be within 40-120 A. Current

value depends upon simultaneous energizing of glasses heating. To know true value

of current, amperemeter readings should be divided by 6.

Wipers

Heated glasses are equipped with wipers. Wipers serve for removal of water splashes

and snow off cockpit heated glasses. Wiper brushes are actuated directly with

electric actuators ЭПК-2Т-60.

Electric actuator ЭПК-2Т-60 has four modes of operation: starting mode; the first

speed; the second speed; brush return to initial position.

Wipers supply and control circuit is connected to battery bus via circuit breakers (CB)

СТЕКЛООЧИСТ. - ЛЕВЫЙ (WIPER - RIGHT) and СТЕКЛООЧИСТ. - ПРАВЫЙ

(WIPER - RIGHT)

. For wiper control, СТЕКЛООЧИСТИТЕЛЬ (WIPER)

switches on the left and right section of the electric panel are used, Fig. 7.13

Fig. 7.13. Position of switch СТЕКЛООЧИСТИТЕЛЬ (WIPER)

166

1. ПУСК (START) - starting operation, the

3. 1 СКОР (SPEED 1) - switch position for first

switch should be set to this position, for a

speed operation

short time

4. 2 СКОР (SPEED 2) - switch position for

2. СБРОС (RESET) - position to stop operation

second speed operation

The switches have four fixed positions: ПУСК (START), 1 СКОР. (SPEED 1 - 60..90

oscillations per minute), 2 СКОР. (SPEED 2 - 30..60 oscillations per minute) and

neutral position corresponding to off position of electric actuator. The fifth position,

СБРОС (RESET) is not fixed; it provides automatic return to off position.

To start wiper operation, circuit breakers СТЕКЛООЧИСТИТЕЛЬ - ЛЕВЫЙ (WIPER -

RIGHT), СТЕКЛООЧИСТИТЕЛЬ - ПРАВЫЙ (WIPER - RIGHT) should be set to on.

Then set switch СТЕКЛООЧИСТИТЕЛЬ (WIPER) to ПУСК (START) position; then,

depending upon external conditions (rain or snow rate), the switches should be set

to position 1 СКОР (SPEED 1) or 2 СКОР (SPEED 2) (rain and snow on helicopter

glasses is not implemented in simulator yet).

In the game, this switch is controlled by the mouse: player must click on one of the

five areas, corresponding to operation modes of windshield wipers:

After wiper is switched off, brush may stop at any position and obstruct viewing. To

set it to initial limit position, switch СТЕКЛООЧИСТИТЕЛЬ (WIPER) should be set to

СБРОС (RESET) and held in that position until brush takes limit position; at that,

electric actuator will stop automatically. After wiper is stopped in limit position,

switch is to be released; when released, the switch takes neutral (off) position.

7.4.3. Anti-Ice System of air inlet Particle Separator System

Air inlet Particle Separator System (PSS) anti-icing system is of combined type (bleed

air plus electrical heating); engines’ inlets are heated by bleed air only.

Heating hot air is bled from combustor cooling loop (5), Fig. 7.14.

Fig. 7.14. Diagram of Hot Air Bleed for PSS and Engines’ Inlets anti-ice system

and for PSS Needs

1. Anti-ice system of air intake (intake lip)

5. Fitting for air bleed from ТВ3-117ВМ engine

2. Electric shutter 1919Т, which opens hot air

combustor cooling loop

flow for anti-ice system needs

6. Heating of НР-ЗВМ governor pump thermal

3. Temperature regulator

compensator air receiver (for correct operation

4. Electric shutter 1919Т, which opens hot air

of governor's systems)

flow to PSS ejector (for vacuum creation)

7. Heated parts of inlet guide vane (vertical

and horizontal supports)

8. Air bleed for PSS trap heating

9. PSS ejector

Either engine has shutter 1919Т (2), which opens hot air flow from combustor

cooling loop (5).

NOTE. The second shutter (4) serves for air supply to PSS ejector for creation of vacuum, which

promotes suction and ejection of dust from engine inlet duct.

168

Fig. 7.16. PSS parts heated by heating elements

1. Fairing front

3. Dust removal pipeline casing

2. Nose of fairing support

4. Fairing back

5. Dust exhaust throat

Heating pads are pasted on the entire area of the said surfaces (inside or outside).

Between casing and heating pad, thermal sensors ТД-2 are pasted; those sensors

ensure stable temperature of heater under various temperatures of outside air, along

with temperature regulators ТЭР-1М.

Electric heating system is powered by 200 V AC, 400 Hz.

Heating of left engine, its air intake and PSS are energized and de-energized

manually, heating of right engine and its inlet components are energized both

manually and automatically from ice detector (see below). Heating is de-energized

manually.

Readings of dashboard amperemeter (Fig. 7.18) shall be within 65-120 A. Current

value depends upon simultaneous energizing of PSS front and back heating. To know

true value of current, amperemeter readings should be divided by 3.

Technical details of heating elements:

Supply voltage, V

200-208V

Power demand, W, max.

8000

Current consumption, A, max.

170

7.4.4. Rotors Anti-Ice System

Blades of main and tail rotors are heated by heating elements powered by AC

voltage.

Main rotor heating element comprises four sections, that of tail rotor is divided in two

sections.

Fig. 7.17. Main and Tail Rotors Blades Anti-Ice System. Diagram of Heating Elements

Location Relative to Blade Profile

1. Diagram of Main Rotor Heating Elements

2. Diagram of Tail Rotor Heating Elements

(4 sections)

(2 sections)

Sections are energized in cycles by cyclic timer ПМК-21. In one cycle, the cyclic timer

activates heating of every section of main and tail rotors for 38.5 sec and cooling of

them for 115.5 sec (main rotor section) and 38.5 sec (tail rotor section). Tail rotor

heating elements sections are energized in the following sequence: the first sections

of tail rotor blades’ heating elements are energized along with the sections I and III

of main rotor blades’ heating elements; the second sections are energized along with

the sections II and IV of main rotor blades’ heating elements.

For check of current value in every section of every main rotor blade and in tail rotor

sections, set rotary switch ТОКИ ПОТРЕБИТЕЛЕЙ (CONSUMERS CURRENTS) to

positions ЛОПАСТИ НЕСУЩ. ВИНТА (MAIN ROTOR BLADES) 1-2-3-4-5 and ХВОСТ

ВИНТА (TAIL ROTOR) in series when lamp of appropriate section comes on and

check current in relevant sections of blades by dashboard amperemeter, Fig. 7.18.

Fig. 7.18. Check of Anti-Ice System Heating Elements Electric Circuits Operation

1. Current consumers rotary switch

2. Lamp indicating energizing of separate

sections of main rotor and tail rotor

3. AC amperemeter

Dashboard amperemeter should be within:



60-72 A for main rotor blade;



110-150 A for tail rotor blades.

To know true value of current, amperemeter readings are to be divided by 3 for main

rotor blade and by 6 for tail rotor blades.

7.4.5. Ice Detectors

For timely detection of icing, warning about helicopter structures icing and automatic

energizing of anti-icing system, helicopter is equipped with radioisotopic ice detector

РИО-3 (RIO-3). Detecting unit of ice detector РИО-3 is installed in fan air intake

duct. Operation is based on variation in conductivity of electric circuit section, which

is energized by radioisotope beta-ray emission.

Fig. 7.19. Radioisotopic Ice Detector

Moreover, visual ice detector is fitted outside, on left land blister. Visual ice detector

is a rod on the left sliding blister. The rod has red and black vertical stripes, 5 mm

wide each. Pilot may use the rod to evaluate rate of icing (not implemented in the

stimulator).

Fig. 7.20. Visual Ice Detector

172

Radioisotopic ice detector serves for sending signal about initiation of icing to the

crew (continuous warning signal when helicopter is in icing zone) and automatic

energizing of anti-icing system (of both rotors, right engine and glasses). After

helicopter leaves icing zone, ice detector РИО-3 stops generation of signal; ant-icing

system is to be de-energized manually.

Signal from ice detector РИО-3 may be generated with certain delay if significant

amount of ice is already accumulated on PSS and engine units. After heating is

energized, accumulated ice may come off and get into engines, thus causing failure

of engines.

To avoid this, heating of PSS and engines is to be energized manually before flight.

7.4.6. Control of Anti-Ice System

1. For anti-ice system functioning in manual and automatic modes, anti-ice system

should be on.

2. The control of anti-ice system is carried out using the anti-ice system Control

Panel, Fig. 7.21

Fig. 7.21. Anti-Ice System Control Panel

1. Switch ОБЩЕЕ РУЧН-АВТОМ (GENERAL

7. Switch (manual - automatic activation)

MANUAL-AUTO)

ДВИГ ПЗУ ПРАВ (ENG PSS RIGHT) for PSS

2. Anti-ice system disabling button

and right engine inlet heating

3. Switch ДВИГ ПЗУ ЛЕВ (ENG PSS LEFT) for

8. Monitoring of detector РИО-3 heating

PSS and left engine inlet heating

9. Switch of detector РИО-3 heating

4. Icing indicator lamp (red) and anti-ice

10. Glasses heating switch

system activation indicator lamp (green)

11. Detector РИО-3 heating serviceability

5. Indicator lamp of PSS and left (left side)

indicator lamp

and right (right side) engines’ inlets heating

12. AC amperemeter

activated

13. Lamp indicating successive activation of

6. Rotary switch, for consumers current

main and tail rotors’ separate sections

monitoring

174

3. For anti-icing system operation in automatic mode, all switches on the control

panel should be in lower position:

Operation in automatic mode is based on receiving a signal from ice detector РИО-3

by anti-ice system control unit. At that, under icing conditions, anti-ice system and

warning system generates the following signals:

 for ОБЛЕДЕН (ICING) lamp

coming on;

 for automatic energizing of anti-icing system of rotors blades, right

engine, its air intake, PSS and glass heating - the same may be verified

by appropriate lamps coming on

;

 for replay of РИ-65 (RI-65) voice recorder message: ОБЛЕДЕНЕНИЕ

(ICING);

 to the САРПП-12ДМ (SARPP-12DM) flight data recorder, for producing

signal ОБЛЕДЕН (ICING).

If anti-ice system was energized automatically, energize anti-ice system of PSS and

LEFT engine inlet manually (set switch ДВИГ ПЗУ ЛЕВ (LEFT ENG PSS) to top

position), after you make sure in stable operation of right engine.

4. For anti-icing system operation in manual mode, all switches on the control

panel

should

upper

position

(exept

РИО-3

swith):

Moreover, manual energizing of anti-ice system for separate units is possible: main

rotor and tail rotor; left engine (PSS and engine inlet); right engine (PSS and engine

inlet); glasses.

5. To de-energize anti-ice system, set all anti-ice system switches to off (down) on

the control panel and press button ВЫК (OFF)

7.4.7. The effect of Anti-Ice System system on the flight characteristics

Maximum takeoff weight

With PZU (PSS) turned on, reduce maximum takeoff weight indicated in charts Fig.

8.1 , Fig. 8.2Fig. 8.3 by 200 kg. With engine and rotor anti-ice systems turned on,

reduce maximum takeoff weight indicated in this charts by 1000 kg.

Fuel consumption

With anti-ice system engaged, fuel consumption rates indicated in the tables Table

8.4 increase as follows:

 engine anti-ice system: 3%

 main and tail rotor anti-ice system: 2%

With the PZU (PSS) engaged, fuel consumption rates per hour provided in Table 8.4

increase by 3%.

7.5. Pneumatic System

Helicopter pneumatic system serves for breaking of main landing gear wheels and

charging of wheels tubes from on-board bottles when helicopter is not on an airfield

(Fig. 7.22).

176

Fig. 7.24. Wheel Break Lever on Left Cyclic Pitch Control Stick

1. Breaks control lever

[W]

2. Parking break trigger

[LShift +W]

(7) REDUCING ACCELERATOR УПО3/2М (UPO3/2M) accelerates compressed air feed to

main gears wheels breaks; also, it ensures air bleed to atmosphere during breaks

release. Reducing accelerator operates from pilot pressure fed from reducing valve

УП25/2; the accelerator creates pressure of (3300+300) kPa [(33+3) kgf/cm²] in

break line.

(11) PRESSURE CONTROL UNIT АД-50 (AD-50) switches over modes of АК-50Т

compressor operation (operating mode to idle mode and vice versa), automatically.

Compressor is switched from operating to idle mode under air bottles pressure

(5000+400) kPa [(50+4) kgf/cm²], and from idle to operating mode under pressure

not less than 4000 kPa (40 kgf/cm²). Pressure control unit is installed on the

pneumatic units’ panel.

NOTE. Pneumatic units’ panel facilitates units installation and tightness check; also, it allows to reduce

number of pipes. It accommodates some pneumatic system units. The panel is located in fuselage

cargo compartment, left side.

Primary technical details of pneumatic system:

Parameter Description

Minimum

Nominal Value

Maximum

Value

Bottles air pressure, kgf/cm²

Wheel breaks air pressure, kgf/cm²

Normal Operation

Check charging of air system (system pressure read by pressure gauge ВОЗДУХ

(AIR) should be 40...50 kg/cm²) and operation of landing gear wheels break system

(when break lever is pressed, wheel breaks pressure read by pressure gauge

ТОРМОЗ (BREAK) should be 30...34 kgf/cm²; after brake release, no residual

pressure in breaks is allowed).

7.6. Fire Protection System

7.6.1. Brief Description

Fire protection equipment serves for fire detection, indication and extinguishing in

the protected units. The helicopter fire protection equipment comprises fire detection

and alarm system and fire extinguishing system that includes two ballons with fire-

extinguishing liquid (one ballon is the first order operation; another is the second

order operation). After СИГНАЛИЗАЦИЯ (INDICATION), 1 ОЧЕРЕДЬ (THE FIRST

ORDER), 2 ОЧЕРЕДЬ (THE SECOND ORDER) breakers are switched on, electric

circuit is powered by battery bus, Fig. 7.25:

Fig. 7.25. Fire Protection System circuit breakers

7.6.2. Fire detection and alarm system

ССП-ФК (SSP-FK) fire indication system serves for fire indication and crew warning.

The ССП-ФК (SSP-FK) fire indication system serves for:

 fire indication in the helicopter protected units

 crew warning by light alarm

 producing for additional signals on the voice data and flight variables

auto-recording equipment

 automatic energizing of the first order ballon discharging in the unit where

fire signal was detected

 indication of fire-extinguishing equipment operation

 monitoring of system serviceability and alertness.

The ССП-ФК (SSP-FK) equipment operation is based on receiving and processing electrical signal from

the special detecting units. This unit is a thermopile. It assembles of three alternately connected

chromel-alumel thermocouples. Electrodes hot junctions (fast-response) are much lesser size than

non-operating (slow-response) ones. When the unit enviroment heating fast, its fast-response

junctions heat much faster than slow-response ones as fast-response junctions are lesser mass. Due

to the difference of hot and non-operating junctions heating temperatures, thermal electromotive

force appears on the unit output rod.

Fire detection system

There is three sets of the ССП-ФК (SSP-FK) equipment on the helicopter. They

serves for fire detection in four units that are more dangerous in relation of fire

hazard. These units are as follows:

 the left engine

 the right engine

 КО-50 kerosene-combustion heater

 the main rotor transmission, the fuel consumed tank and the АИ-9В (AI-

9V) engine (these three sections are included into one protected unit).

Fire detection system operates on a multiple-circuit electric device. It is equipped with some

standalone sensor groups, which initiate signals on actuation devices, in each monitored unit.

In the actuating unit, each sensor group is connected to its signal conditioner (combined unit) forming

a fully independent sensitive circuit. Thereby, it is ensured high system reliability, as in case of any

180

sensor group failure, normal operation of the other sensor groups stays constant. The system is able

to detect a fire by one circuit of sensitive elements.

The ССП-ФК (SSP-FK) system comprises the following:



42 FAS (fire-alarm sensor) signal devices type integrated into 14 groups.

Fig. 7.26. Location of the ССП-ФК (SSP-FK) system sensors

 three ССП-ФК-БИ (SSP-FK-BI) actuating units, which receive electric

signals initiating in sensors circuit in event of fire and control indication

system and automatic equipment. The actuating units are installed in the

cockpit on the right rack.

Fire alarm system

Light alarm. Fire light alarm system comprises five indicating lamps with red filters

installed on the fire protection system switchboard, Fig. 7.27, 1:

 Two ПОЖАР ЛЕВ. ДВ. (LEFT ENGINE FIRE) and ПОЖАР ПРАВ. ДВ.

(RIGHT ENGINE FIRE) lamps indicate about a fire in the helicopter left

and right engine units

 ПОЖАР КО-50 (КО-50 FIRE) lamp indicates about a fire in the КО-50

kerosene-combustion heater unit

 ПОЖАР РЕДУК. АИ-9 (ROTOR TRANSMISSION FIRE АI-9) lamp indicates

about a fire in the units of the main rotor transmission, the fuel consumed

tank and the АИ-9В (AI-9V) engine

 ПОЖАР (FIRE) lamp is an additional lamp on the left dashboard which

duplicates any of four main lamps coming on, Fig. 7.28.

Fig. 7.27. Fire protection system switchboard on Center overhead console

1. The lamp indicating fire in the protected

Letters mark lamps indicating fire, operation of

units (ПОЖАР ЛЕВ ДВ, ПОЖАР ПРАВ ДВ

fire extinguishing system ballons and push

ПОЖАР КО-50, ПОЖАР РЕД.АИ-9) (LEFT

buttons of units fire protection system as

ENGINE FIRE, RIGHT ENGINE FIRE, КО-50

follows:

FIRE, MAIN ROTOR АИ-9 FIRE)

a. The LEFT engine

2. The lamp indicating operation of the first

b. The RIGHT engine

order fire extinguishing system

c. The КО-50 kerosene-combustion heater

3. The lamp indicating operation of the second

d. The main rotor transmission, the fuel

order fire extinguishing system

consumed tank and APU

4. The first order fire extinguishing system

push button

5. The second order fire extinguishing system

push button

6. Alarm silence push button

Fig. 7.28. Lamp ПОЖАР (FIRE) on the left dashboard

182

To capture the crew attention to fire alert in any unit as soon as possible, it is

provided for operation these five indication lamps in flashing mode. For this purpose,

its supply circuits are connected into the helicopter FLASHER (МИГАЛКА) system

scheme.

Auxiliary alarm signal. Simultaneously with energizing the lamp, fire protection

system serves for sending auxiliary alarm signals by parallel circuit in the unit input in

the РИ-65 (RI-65) voice data equipment kit. Depending on the site of the fire, a

signal comes on one of the four channels, which herewith energizing and serves for

replay of the РИ-65 (RI-65) voice recorder message to the left pilot’s phones:

ПОЖАР ЛЕВЫЙ ДВИГАТЕЛЬ, ПОЖАР ПРАВЫЙ ДВИГАТЕЛЬ, ПОЖАР РЕДУКТОР

ВСУ, ПОЖАР КО-50 (LEFT ENGINE FIRE, RIGHT ENGINE FIRE, ROTOR

TRANSMISSION APS FIRE, КО-50 FIRE).

The voice recorder message comes to the left pilot’s phones repeatedly and replays

on a periodic basis of two messages per 12 s. Simultaneously, the same message is

sent to the flight ground control station and the unit input in the flight variables

auto-recording equipment (САРПП-12ДМ (SARPP-12DM)) via the helicopter

command communication VHF radio.

7.6.3. Indication and alarm system check

Indication and alarm system serviceability are monitored by checking of the

indicating lamps and monitoring of the sensors serviceability.

LAMPS SERVICEABILITY and its power supply circuits are checked by the helicopter lamps

check system with the ПРОВЕРКА ЛАМП. МИГАЛКА (LAMP CHECK. FLASHER) circuit

breaker and the МИГАЛКА (FLASHER) switch set to on.

When setting the ПРОВЕРКА МИГАЛКИ-СИГНАЛ. ЛАМП switch on the pilot’s central

console to the ПРОВЕРКА МИГАЛКИ (FLASHER CHECK) position, five fire indication

lamps should come on in flashing mode (Fig. 7.27, 1 и Fig. 7.28). When setting the

ПРОВЕРКА СИГНАЛ. ЛАМП switch on, eight 1 ОЧЕРЕДЬ (1 ORDER) and 2 ОЧЕРЕДЬ

(2 ORDER) lamps should come on (Fig. 7.27, 2, 3).

FAS SIGNAL DEVICES SERVICEABILITY and its power supply circuits are monitored by the

КОНТРОЛЬ ДАТЧИКОВ (MONITORING OF SIGNAL DEVICES) panel (Fig. 7.29) on

the mid section of electric panel after the ОГНЕТУШЕНИЕ-КОНТРОЛЬ ДАТЧИКОВ

(FIRE EXTINGUISHING - MONITORING OF SIGNAL DEVICES) switch set to the

КОНТРОЛЬ ДАТЧИКОВ (MONITORING OF SIGNAL DEVICES) position. Herewith,

КОНТРОЛЬ ДАТЧИКОВ (MONITORING OF SIGNAL DEVICES) red indicator lamp (Fig.

7.29, 3) comes on, indicating monitoring circuit readiness

Fig. 7.29. КОНТРОЛЬ ДАТЧИКОВ (MONITORING OF SIGNAL DEVICES)

Center overhead console

1. ОГНЕТУШЕНИЕ - КОНТРОЛЬ ДАТЧИКОВ

4. ВЫК - КОНТРОЛЬ - КАНАЛЫ (OFF -

(FIRE EXTINGUISHING - MONITORING OF

MONITORING - CHANNELS)1-2-3-4-5-6

SIGNAL DEVICES) switch

switch

2. Switch I - II checking of pyros on fire-

5. Positions for connection of signal devices

extinguishing liquid ballons

channels 1-2-3-4-5-6 or monitoring shutting

3. Lamp indicating position of switch (1)

down (ВЫК (OFF))

КОНТРОЛЬ ДАТЧИКОВ (MONITORING OF

SIGNAL DEVICES)

Monitoring system is operated by battery bus via ПРОТИВОПОЖАРНАЯ СИСТЕМА.

СИГНАЛИЗАЦ. (FIRE PROTECTION SYSTEM. INDICATION) circuit breaker and

КОНТРОЛЬ ДАТЧИКОВ (MONITORING OF SIGNAL DEVICES) rotary switch. Fourteen

groups of FAS signal devices are integrated into six monitoring channels (Fig. 7.29,

4). Each of them connected to its switch contact. For monitoring, each switch is

required to set to checking channels position alternately. In case of serviceability of

signal devices in the groups and its power supply circuits, on the mid section of the

electric panel and the left dashboard, appropriate lamps, indicating fire, should come

on (Fig. 7.27, 1 и Fig. 7.28).

Division of the groups of the different units signal devices for monitoring channels is

given in the table:

Monitored units

Monitoring channels

Lamps come on

III

Left engine

ПОЖАР ЛЕВ. ДВ. (LEFT ENGINE FIRE)

Right engine

ПОЖАР ПРАВ ДВ. (RIGHT ENGINE FIRE)

Main rotor transmission,

ПОЖАР РЕДУК. АИ-9 (MAIN ROTOR АИ-9

Service fuel cell

(AI-9) FIRE)

АИ-9В (AI-9V) engine

ПОЖАР РЕДУК. АИ-9 (MAIN ROTOR АИ-9

(AI-9) FIRE)

184

КО-50 Kerosene-

ПОЖАР КО-50 (КО-50 FIRE)

combustion heater

7.6.4. Fire extinguishing system

Fixed fire extinguishing system comprises two УБШ-4-4 (UBSH-4-4) type ball shaped

ballons with extinguishing agent served for two orders discharging, Fig. 7.30. The

ballon is charged extinguishing agent comprised of Freon gas 114В2 weighing

5,640 kg and, for providing this agent supply, it is charged air mixture or nitrogen

weighing 0,180 kg to the pressure (10500...11500) kРа [(105...115) kgf/cm2] at a

temperature of 15...20 °С. Each ballon is equipped with four initiators (as per the

number of protected units). Each initiator is opened by two pyros. Ballons are located

in the unit of the main rotor transmission on the starboard side.

Fig. 7.30. УБШ-4-4 (UBSH-4-4) Fire extinguisher

1. Branch pipe to the unit of main rotor

5. ПГКц (PGKts) Initiator

transmission, the fuel consumed tank and АИ-

6. Siphon tube

9В (AI-9V) engine

7. Collector component

2. Branch pipe outboard

8. Branch pipe to the КО-50 unit

3. Pressure gauge

9. Branch pipe to the left engine unit

4. УБШ-4-4 (UBSH-4-4) Fire extinfuisher

10. Branch pipe to the right engine unit

ballon, collector component

Please note, that during fire extinguishing in the unit of the main rotor transmission,

the fuel consumed tank and the АИ-9В (AI-9V) engine, extinguishing agent spreads

through three protected sections (the section of the main rotor transmission, the

section of the fuel consumed tank as well as the section of the АИ-9В (AI-9V)

engine), in spite of there was a fire in the section of the fuel consumed tank (for

example).

7.6.5. Normal Operation

Operating Conditions (stages) and required actions

1. CHECK OF SERVICEABILITY OF ELECTRICAL POWER SYSTEM FIRE PROTECTION EQUIPMENT.

After energizing the electrical power sources, it is required to:

 set the РАДИО (RADIO) breaker to off

;

 set the ПРОТИВОПОЖАРНАЯ СИСТЕМА (FIRE PROTECTION SYSTEM)

breaker to on

;

 set the КОНТРОЛЬ ПИРОПАТРОНОВ (PYROS MONITORING) switch to the

positions I and II alternately

, make sure that pyros electric circuits

are serviceable (no yellow fire extinguisher indicator lamps should come

on)

;

 set the КОНТРОЛЬ ДАТЧИКОВ-ОГНЕТУШЕНИЕ (MONITORING OF

SIGNAL DEVICES - FIRE EXTINGUISHING) switch to the КОНТРОЛЬ

(MONITORING) position

. At that, the КОНТРОЛЬ ДАТЧИКОВ

(MONITORING OF SIGNAL DEVICES) indicator lamp should come on

;

 set the rotary switch to positions of six channels alternately

When setting the switch to positions 1 and 2, the following lamps should come on:

ПОЖАР ЛЕВ. ДВ. (LEFT ENGINE FIRE), ПОЖАР ПРАВ.ДВ. (RIGHT ENGINE FIRE),

ПОЖАР КО-50 (КО-50 FIRE), ПОЖАР РЕДУК.АИ-9 (ROTOR TRANSMISSION АИ-9

186

(AI-9) FIRE)

. When setting the switch to

the position 3, the КО-50 fire indicator lamp should go off. When setting the switch

to the positions 4, 5 and 6, the ПОЖАР РЕДУК.АИ-9 (ROTOR TRANSMISSION АИ-9

(AI-9) FIRE) indicator lamp should come on, ahother lamps should go off.

2. SYSTEM PREPARATION FOR WORK

 set the rotary switch to the ВЫК (OFF) position

;

 set the КОНТРОЛЬ ДАТЧИКОВ-ОГНЕТУШЕНИЕ (MONITORING OF

SIGNAL DEVICES - FIRE EXTINGUISHING) switch to the ОГНЕТУШЕНИЕ

(FIRE EXTINGUISHING) position

. At that, the КОНТРОЛЬ

ДАТЧИКОВ (MONITORING OF SIGNAL DEVICES) indicator lamp should go

off

 The system is ready to automatic operation or manual actuation.

3. SWITCHING THE SYSTEM OFF

After taxiing in and shutting down the engines, for switching the system off, set the

ПРОТИВОПОЖАРНАЯ СИСТЕМА (FIRE PROTECTION SYSTEM) breaker to off.

WARNING. To avoid operating of the first (automatic) order fire extinguishers of

the system, do not set the rotor switch from the ВЫКЛ (OFF) position if the

КОНТРОЛЬ ДАТЧИКОВ-ОГНЕТУШЕНИЕ (MONITORING OF SIGNAL DEVICES - FIRE

EXTINGUISHING ) switch is in the position ОГНЕТУШЕНИЕ (FIRE EXTINGUISHING),

and do not set the КОНТРОЛЬ ДАТЧИКОВ-ОГНЕТУШЕНИЕ (MONITORING OF

SIGNAL DEVICES - FIRE EXTINGUISHING ) switch to the ОГНЕТУШЕНИЕ (FIRE

EXTINGUISHING) position before the rotor switch is set to the ВЫКЛ (OFF) position.

The system operation and crew procedure

with fire protection equipment in case of fire

If there is any protected units fire (for example, the left engine fire) and thermal

electromotive force on any signal devices channels, the following lamps come on:

 the ПОЖАР ЛЕВ ДВ (LEFT ENGINE FIRE) lamp on the fire protection

system switchboard (Fig. 7.27, 1-a);

 the ПОЖАР (FIRE) lamp on the left dashboard (Fig. 7.28).

Simultaneously, pyro of extinguishing liquid supply line of the first order ballon

detonates. At that, the 1 ОЧЕРЕДЬ (1 ORDER) indicator lamp of the first order

operation on the fire protection system switchboard comes on (Fig. 7.27, 2-a).

When fire is extinguished, during not more than 10 s, thermal electromotive force on

any signal devices channels disappeares and the ПОЖАР ЛЕВ ДВ (LEFT ENGINE

FIRE) lamp goes off, but the 1 ОЧЕРЕДЬ (1 ORDER) lamp continues to come on.

When the lamps do not go off in 10 s, press the ВЫК. СИГНАЛИЗАЦИИ ПОЖАРА

(ALARM SILENCE PUSH BUTTON) button (Fig. 7.27, 6).

When the first order ballon is not operated automatically, the ПОЖАР ЛЕВ ДВ (LEFT

ENGINE FIRE) lamp on the fire protection system switchboard and the ПОЖАР

(FIRE) lamp on the left dashboard come on, but the 1 ОЧЕРЕДЬ (1 ORDER) lamp

does not come on), it is required to detonate the pyro of extinguishing liquid supply

line to the left engine unit of the first order ballon manually by pressing the

ОЧЕРЕДЬ (1 ORDER) button for the left engine unit (Fig. 7.27, 4-a).

When the left engine fire is extinguished by the first order ballon, and then there is

another unit fire (for example, the unit of the main rotor transmission, the fuel

consumed tank and APS (auxiliary power supply)), that is required to detonate the

pyro of the second order ballon manually by pressing the 2 ОЧЕРЕДЬ (2 ORDER)

button for the unit of the main rotor transmission, the fuel consumed tank and APS

(Fig. 7.27, 5-d). After that, the 2 ОЧЕРЕДЬ (2 ORDER) lamp, located under the

ПОЖАР РЕД.АИ-9 (ROTOR TRANSMISSION АИ-9 FIRE) lamp on the fire protection

system switchboard, comes on (Fig. 7.27, 3-d).

NOTE. After fire extinguishing system operating in the unit of the main rotor transmission, the fuel

consumed tank and APS operates, the АИ-9В (AI-9V) engine shuts down, if it operates, and its start is

blocked. That is similar to the КО-50 unit.

See also Onboard fire

7.7. Environmental and heating system

7.7.1. Brief description

The environmental and heating system is used for creating and maintaining

comfortable environmental conditions for crew and passengers. It provides:

 supply of heated and atmospheric air to the cockpit and cargo cabin;

 airflow for blowing on front winshields and cockpit blisters.

This system consists of the KO-50 combustion heater and two DV-302T cockpit fans

for right and left pilot.

The heater is operating in the following way: after heater was started, in the

combustion camera the kerosene-air mixture is being burned and exhaust gas is

being removed through the exhaust nozzle. Heat from the combustion, is warming

up the calorifer. The airflow from a fan is moving through calorifer, air is being

warmed up and fed to the helicopter’s cockpit.

The heater can operate either in automatic, manual or ventilation mode. When

heater is operating in automatic mode, the temperature, set by the temperature

knob, is being maintained constant. Manual control is used for maximum or medium

heating output modes. The recirculation mode is used to speed up the heating of the

cabins during winter conditions, by using air from the cargo cabin. In the game, this

188

behavior of shut-dumper is not modelled. In the ventilation mode, this system

provides cooling of the heater and cabins ventilation (through air ducts) during

warmer seasons.

In the electrical circuit of the heater, the relay, which disconnects electrical power

supply in case of a fire in the KO-50 bay, is installed.

The KO-50 heater is located in front of the right external fuel tank, Fig. 7.31:

Fig. 7.31. Location of the KO-50 heater on the helicopter

Heat ouptut of the KO-50 heater is equal to 50,000 kcal/h, in case of temperature

difference on the ground of 130°С. Air consumption is 1,760 kg/h, if pressure behind

the heater is

100 mm H₂O. Fuel consumption is

8.7 kg/h. Electrical power

consumption of the fan is 2.5 kW.

Fuel for the heater is supplied from the right engine’s fuel pipes, when the 610200А

КО-50 electro-magnetic valve is opened. This valve is opened by the heater starter

system, from the KO-50 control panel, Fig. 7.32.

7.7.2. Enabling the KO-50 heater in automatic mode

 Enable the КО-50 ACB on the right over head circuit breaker panel

;

 set the mode switch on the heater control panel into the АВТОМ.

(AUTOMATIC) position (Fig. 7.32, 3);

 set required temperature by the temperature knob (Fig. 7.32, 6);

 press the ЗАПУСК (START) button (Fig. 7.32, 1), as a result on the heater

control panel the ПОДОГРЕВАТЕЛЬ (PREHEATER) signal panel goes on.

When fuel temperature reaches

70±5°С, the ПОДОГРЕВАТЕЛЬ

(PREHEATER) panel goes off and the ЗАЖИГАНИЕ (INGITION) and КО-50

РАБОТАЕТ (KO-50 IS OPERATING) lights are on. After not more than 40

seconds, the ЗАЖИГАНИЕ (IGNITION) panel should go off, this will

indicate sustainable combustion of the fuel in the KO-50 heater.

7.7.3. Enabling the KO-50 heater in manual mode

 Enable the KO-50 ACB on the right over head circuit breaker panel panel;

 set the mode switch on the KO-50 control panel to the РУЧН (MANUAL)

position;

 set the heat output mode switch to the ПОЛН.РЕЖИМ (HIGH OUTPUT) or

СРЕДН.РЕЖИМ (MEDIUM OUTPUT) position;

 press the ЗАПУСК (START) button.

The following stages of operation are equal to those in automatic mode.

NOTE: Before disabling the heater, to cool it faster, set the ВЕНТИЛ. (VENTILATION) switch to the

ВКЛ. (ENABLE) position for 3..10 minutes (no need to use it in game).

7.7.4. Using the KO-50 heater in ventilation mode

When the heater is supposed to be used during summer time for cabins ventilation,

enable the KO-50 ACB and set the ВЕНТИЛ. (VENTILATION) switch to the ВКЛ.

(ENABLE) position. Air is supplied, using the same air channels as during winter time.

7.7.5. Switching OFF the KO-50 heater

 Set the РУЧН.- АВТОМ. (MANUAL-AUTO) switch into the neutral position;

 after landing, disable corresponding ACB.

5. Left pilot’s FPP-7M light

Electrical circuits of the exterior lighting system are connected via ACBs, located on

the right over head circuit breaker panel, Fig. 7.34:

Fig. 7.34. The ACBs of the exterior lighting system

1. Left search/landing light, light control circuit

4. Right search/landing light, incandescent

2. Left search/landing light, incandescent light

light bulb circuit

bulb circuit

5. Navigation lights’ supply circuit

3. Right search/landing light, light control

6. Formation lights’ supply circuit

circuit and incandescent light bulb circuit of

7. МИГАЛКА (FLASH) circuit

taxi light

8. PRF-4 lights (on the main landing gear

struts, not implemented)

The ACB of the anti-collision light is located on the right side console (see below).

FPP-7M search/landing light

There are two FPP-7 search/ landing lights, mounted under the helicopter nose

section to the right and left of the longitudinal centerline (Fig. 7.33 , 4, 5)

. The lights are designed for search operations during low

visibility conditions, searching and illuminating landing sites and taxi areas at night.

The lights can be elevetade at angle from 0 (down) to up to 120° and turned at any

angle in azimuthal direction or can be retracted.

The right FPP-7M lights is connected to the battery circuit, and the left one to the

rectifier bus. In the light control and incandescent light bulb circuits are automatic

circuit breakers, located on the right over head circuit breaker panel

Turning on the lights and supplying electrical power to the extension/retraction and

rotation switches is done with help of two ФАРЫ СВЕТ - ВЫК - УБРАНА (LIGHT ON

- OFF - RETRACTED) switches, when they are set to the СВЕТ (LIGHT) position

[LShift + L]

[RShift + L]

. These switches are located on separated

panels, attached to the left of the pilot’s instrument panel and to the right of the co-

pilot’s one (Fig. 7.35, 2 ,3):

Fig. 7.35. Lights controls

194

1. Left pilot’s FR-100 taxi light switch

3. Right pilot’s FPP-7M search/landing light

2. Left pilot’s FPP-7M search/landing light

control switch

Extension, retraction (beam elevation control), rotation to the left and to the right

(azumith control) is performed with help of two five-position hats, located on the

collectives, Fig. 7.36:

Fig. 7.36. Search/taxi light beam controls

1. Light (beam) forward (up)

[LShift + 8]

(left

3. Light (beam) backward (down)

light) /

[RShift + 8]

(right light)

[LShift + 7]

[RShift + 7]

2. Light (beam) rotation to the right

4. Light (beam) rotation to the left

[LShift + 0]

[RShift + 0]

[LShift + 9]

[RShift + 9]

When the ФАРЫ СВЕТ - УБРАНА (LIGHT ON -OFF - RETRACTED) switches are set

to the УБРАНА (RETRACTION) position

[not assigned]

, lights turn off

and retract automatically. During retraction, the lights are automatically returning to

their initial position in relation to the helicopter centerline. Fully extended and

retracted positions are locked with help of limit switches.

The ВЫК (OFF) position

(neutral)

[LShift + ;]

(first press) is used for

preserving position of the light, when light is OFF. In game, this is a default position.

FPP-7M FEATURES are shown in Table 7.2

Table 7.2

Nominal supply voltage, V

Current, consumed by the each electrical motor, А, not

0,7

Consumed power, Watt, not more

480

Nominal luminous intensity, cd

300000

Extension elevation angle, degrees

from 0 to 120

Rotation angle (azumith)

not limited

Weight, kg, not more

FR-100 taxi light

The light is used for illuminating surface while taxiing. The taxi light is located on the

lower part of the fuselage

(in real life, used for saving service

life of the landing lights).

It is supplied by the battery bus via the ФАРЫ ПРАВАЯ УПРАВЛЕНИЕ (LIGHTS

RIGHT CONTROL) ACB

, and turns on by the ФАРЫ

РУЛЁЖНАЯ (LIGHTS -TAXI) switch, located near the left FPP-7 switch, (Fig. 7.35, 1)

[RCtrl + L]

MSL-3 anti-collision light

The MSL-3 anti-collision light is designed to indicate the direction of travel and the

position of the helicopter at night, during inclement weather, and in poor visibility. It

is used to indicate the place of helicopter in case of emergency landing. The MSL-3

light is located on the top of the tail,

. This light is turned on by

196

the ПРОБЛЕСК МАЯК (ANTI-COLLISION LIGHT) ACB, located on the right electrical

panel

[RCtrl + 6]

This light is connected to the battery bus (on some models, for example Mi-8MTVx -

to the rectifier bus).

To avoid overheating, do not operate the anti-collision light for more than 10 minutes

on the ground if the main rotor is not turning (not implemented).

Navigation lights

Navigation lights are used for indication of helicopter position.

On the helicopter the BANO-45 type navigation lights and the KhS-39 tail light are

installed.

The red and green navigation lights are installed on the weapon rack tips

or on the starboard and port sides of the nose part of the fuselage

(if weapon racks and armor are not mounted), KhS-39 is

installed on top of the tail

Navigation lights are supplied via the "АНО" (NAVIGATION LIGHTS) ACB, located on

the right over head circuit breaker panel

. Navigation

OPS-57 Formation lights

The OPS-57 formation lights are designed to assist pilots in maintaining formation

while flying at night and in poor visibility conditions.There are three formation lights

on the helicopter

The formation lights are supplied by the recitifier bus via the СТРОЕВ ОГНИ

(FORMATION LIGHTS) ACB, which is installed on the right over head circuit breaker

panel

. These lights are controlled (enabling and

selection of illimination intensity) with help of the ОГНИ СТРОЕВ ТУСКЛО

(нейтраль)- ЯРКО (NAVIGATION LIGHTS DIMINISHED

-(neutral) -BRIGHT),

installed

the

electrical

panel

the

right

side

console

ТУСКЛО (DIMINISHED) -

[RCtrl

(down), ЯРКО (BRIGHT) -

[RCtrl + 4]

(up) . Neutral position turns of all formation lights.

Blade tip lights

The blade tip lights are used to indicate the position of the main rotor blades for

movement of the helicopter at night or in poor visibility. A blade tip light is installed

inside a removable fairing on each

main rotor blade tip cap.

. This lights are enabled by the ОГНИ

КОНТУР (BLADE TIP) switch, which is installed on the electrical panel of the right

side console

[RCtrl + 5]

Blade tip lights are powered by the 115 VAC bus through the TR 115/7.5V

transformer.

7.8.2. Interior lighting equipment

The interior lighting system includes interior illumination of the cockpit, cargo cabin

and various compartments, and in-helicopter light signalization system.

Cockpit illumination

Cockpit illumination consists of:

 player flashlight;

 illumination of pilots’ working space;

 red backlight system.

PLAYER’S FLASHLIGHT

200

Fig. 7.37. Backup dome lights

RED LIGHT BACKLIGHTING SYSTEM

Red light backlighting system is used for backlighting various instruments and

gauges on the instrument panels, central panel and electrical panel (right rear

console).

Red light backlighting system of the central panel and electrical panel is divided on

two groups. Group I is connected to the rectifier bus and group II to the battery bus.

Red light backlighting system has no ACBs. Turning on of both groups and backlight

intensity adjustment is performed with help of rheostats. To limit maximum voltage

applied to red backlight system, rheostats have limiters.

Red light backlighting for the left overhead circuit breaker panel, left-side console,

left triangular panel, left overhead console, pilot’s instrument panel, KI-13 compass,

left pilot's and cargo cabin’s intercoms and left FPP-7M control panel is controlled by

the КРАСНЫЙ ПОДСВЕТ - ГРУППА 1 (RED BACKLIGHT -GROUP1) and КРАСНЫЙ

202

ПОДСВЕТ - ГРУППА 2 (RED BACKLIGHT - GROUP2) rheostats, installed on the left

side console

Red light backlighting for the weapon console, R-863 radio control panels, right over

head circuit breaker panel, right side console, right triangular panel, copilot’s

instrument panel and right FPP-7 control panel is controlled by the КРАСНЫЙ

ПОДСВЕТ - ГРУППА 1 (RED BACKLIGHT -GROUP1) and КРАСНЫЙ ПОДСВЕТ -

ГРУППА 2 (RED BACKLIGHT - GROUP2) rheostats, installed on the right side console

Red light backlighting of the central autopilot console, center overhead console, AC

power control console, R-828, UV-27, Yadro-1M control panels, and panels installed

in the doorway, is controlled by the КРАСНЫЙ ПОДСВЕТ - ГРУППА 1 (RED

BACKLIGHT -GROUP1) and КРАСНЫЙ ПОДСВЕТ - ГРУППА 2 (RED BACKLIGHT -

Fig. 7.38. Instruments backlighted by the "Подсвет 5,5V" (Backlight 5.5V) system

Supply circuits of the red 5.5V backlight are connected to the 115 VAC generator

bus.

Illuminaton of cargo cabin and technical compartments

This function is not implemented.

In-helicopter light signalization system

Monitoring of the helicopter’s systems and units, besides dedicated instruments, is

performed with help of in-helicopter light signalization system.

Signalization is done with help of light panels with light filters of red, yellow and

green color, located on the instrument panels, central panel and electrical panel. For

some light panel groups, some special operating modes are intended (FLASHING,

DAY-NIGHT) and lamps checking.

The “MIGALKA” (FLASH) system is developed for attracting pilot’s attention to the

light panel, indicating malfunction or failure of some system, unit and informing him

about emergency situations (fire, icing, excessive engine vibration, operation of

backup systems, 270 liters of fuel remains in fuel tank). The operating principle of

the “MIGALKA” (FLASH) is based on lamps, operating in impulse mode. Flashing

mode is enabled with help of a switch, located on the right side console

[RCtrl + -]

THE "ДЕНЬ-НОЧЬ" (DAY-NIGHT) system is intended to reduce illumination intensity

for some intruments and light panels. This is done by setting the "ТАБЛО ДЕНЬ-

НОЧЬ" (PANEL DAY-NIGHT) switch, located on the right side console

, to the "НОЧЬ" (NIGHT) (up) position. As a result,

light panels, connected to the ДЕНЬ-НОЧЬ" (DAY-NIGHT) system will illuminate with

half-intensity.

WARNING LAMPS CHECKING SYSTEM

To enable warning lamps checking system pilot must enable:

 ACB

"ПРОВЕРКА

ЛАМП

МИГАЛКА"(LAMP

TEST

FLASH)

 set the "ПРОВЕРКА СИГНАЛ ЛАМП - МИГАЛКИ" (WARN LTS - TEST-

FLASH) switch, located on central panel, into "СИГНАЛ ЛАМП" (WARN

206

Fig. 7.39. Light panels, connected to the MIGALKA (FLASH) system

1. ОБЛЕДЕНЕНИЕ (ICING)

3. ОСТАЛОСЬ 270л (270 L

(started/continuing)

REMAINS)(emergency fuel)

2. Fire in compartments (from left to right):

4. ДУБЛИР ВКЛЮЧЕНА (BACKUP IS ON)

 ПОЖАР ЛЕВ ДВ (FIRE LEFT ENGINE)

(indicates that backup hydraulic system is

(in left engine's compartment)

operating, it automatically means failure in the

 ПОЖАР ПРАВ ДВ (FIRE RIGHT

utility hydraulic system)

ENGINE) (in the right engine’s

5. Fire annunciator ПОЖАР (FIRE), then -

compartment)

ВЫКЛЮЧИ ЛЕВ ДВ (STOP LEFT ENGINE),

 ПОЖАР КО-50 (FIRE KO-50) (fire in

ВЫКЛЮЧИ ПРАВ ДВ (STOP RIGHT ENGINE)

the heater)

(signalization of excessive engine vibration,

 ПОЖАР РЕД АИ-9 (FIRE

that can damage corresponding engine)

TRANSMISSION AI-9) (in the

transmission compartment of the APU

and in service fuel tank)

Circuits for checking warning lamps functionality, "МИГАЛКА" (FLASH) and "ДЕНЬ-

НОЧЬ" (DAY-NIGHT) systems are connected to battery bus via the ПРОВЕРКА ЛАМП

МИГАЛКА (LAMP TEST FLASH) ACB on the right over head circuit breaker panel.

7.9. Registration of the flight parameters and voice recording

7.9.1. SARPP-12DM flight data recorder

This system is not implemented, but corresponding light panel and switch are

animated.

The SARPP-12DM flight data recorder is designed to record the flight parameters of

the helicopter under normal and emergency conditions.

208

The system continuously records the helicopter’s barometric altitude, indicated

airspeed, the position (tilt and height) of the main rotor swashplate, the main rotor

RPM, and the helicopter pitch and roll angles. The system also records nine event-

driven parameters. The trigger events include low fuel emergency, failure of any fuel

boost pump, engine emergency power engagement, detection of a fire in any deck

compartment, main hydraulic system failure, backup hydraulic system failure, loss of

oil pressure in the main transmission, engagement of engine and particle separator

anti-ice system (right engine only), and engagement of rotor deice system. All data is

provided to the recorder via sensors, pressure switches, and/or transducers installed

in the monitored systems.

Fig. 7.40. SARPP-12D1M FDR power switch and indicator light, left side panel

The recorder may be turned on manually or set for automatic operation using the

"САРПП-12Д1М "РУЧН - АВТОМ"" (FLIGHT RECORDER) switch on the pilot’s left

side panel. When the switch is placed in the AUTO (down) position, the system is

turned on automatically just after takeoff, when the AM-800K microswitch in the left

main landing gear strut is activated. The microswitch activates when the weight is

removed from the main landing gear. If the film transport and light-beam in the K12-

51DM data storage unit are operational, the "САРПП РАБОТАЕТ" (SARPP WORKING)

indicator light, located near the switch, will blink. In AUTO mode, the system

activates only if there is sufficient pressure in the main or reserve hydraulic system

The flight recorder system consists of:

 K12-51DM data storage unit;

 УсС-4-1M UsS-4-1M matching device;

 sensors.

The SARPP-12 system is enabled by the "САРПП-12Д РУЧН-АВТОМ" (SARPP-12D

MANUAL-AUTO) switch on the left side console

To enable the system before engine start, set the corresponding switch to the

"АВТОМ." (AUTO) or РУЧН (MANUAL) position

[LAlt + LCtrl + LShift + 6]

When this switch is set to the АВТОМ. (AUTO) position, the flight recorder begins

operating only if there is enough pressure in the utility or backup hydraulic system,

or if limit switches in the main landing gear struts trigger when helicopter is airborne.

If set to the РУЧН. (MANUAL), recorder engages immediately, indendently on

pressure in the hydraulic systems.

Data storage unit in armored container and matching device are installed in the tail

boom.

The SARPP-12DM system is supplied by the 27-29V onboard DC bus, and in

emergency mode by the battery bus, through PM-10 SARPP ACB, located on the ACB

panel.

Continuous operating time of the SARPP-12 system is not less than 5.5 hours.

7.9.2. P-503B voice recorder

Not implemented, but corresponding switch is animated.

Voice recorder is designed for recording voice information of the helicopter’s pilot,

transmitted over radios and intercoms, as well as information given to the pilot.

210

 fuselage attach points;

 release barrier;

 pulley block and cable (not modelled);

 extension cables for 5, 10, 15, 20, 30, 40, 50 и 65 m (up to 30m in

game);

 cargo cable set (4х4m), also known as "spider" (any length is possible in

the game);

 various additional details (swivel hook, master link of cargo cable set), are

not modelled.

The main load-bearing elements of the outer sling cable are four 16-mm cables. On

top of each cable there is a hinge by which this cable is attached to a special

fuselage attach point, located under helicopter’s ceiling (directly under the main rotor

transmission) on frames № 7 and 10. In the lower part all four cables are connected

with the scale of the DG-64M and then cable goes through the access hole. Sling

load equipment is shown on Fig. 7.42.

Fig. 7.42. Sling load equipment diagram

1. Four load-bearing cables

4. Access hole

2. Release barrier

5. Extension cables

3. DG-64M cargo hook

7.10.2. Cargo hook controls

The DG-64M cargo hook lock is controlled electrically:

212

 supply voltage is 27-29V;

 minimum voltage required for hook opening is 20V.

Controling cargo hook lock means choosing the unhook method: either automatic

opening (or unhook) when cargo touches the ground or manually opening the lock in

flight (and cargo release).

To power up electrical circuits, controlling hook lock, pilot must enable the

"УПРАВЛЕНИЕ ОТКРЫТИЕМ ЗАМКА" "ОСНОВНОЕ", "ДУБЛИР" (CARGO HOOK MAIN

AUX)

the

right

overhead

circuit

breaker

panel:

Automatic opening

Automatic cargo unhook is activated before landing when cargo is already hooked.

To do that, one must enable the ВНЕШНЯЯ ПОДВЕСКА АВТОМ СБРОС (EXTERNAL

LOAD AUTO RELEASE) switch (UP), located on the left side panel:

[RCtrl + RShift + K]

When cargo touches ground and load on lock lever reduces to less than 25 kg, hook

opens and cargo remains on the ground. When lock is opened, the ЗАМОК ОТКРЫТ

(HOOK OPEN) light panel is on

Manual cargo release

Cargo hook can be opened during flight (cargo release). Two procedures are used

for that: normal (tactical) release and emergency cargo release, if emergency

situation occurs during the flight.

To open hook lock with help of electromagnetic trigger, pilot has to press tactical

cargo release button or emergency cargo release button. Both buttons are located on

collective, Fig. 7.43.

Fig. 7.43. Cargo release buttons on the collective

1. Safety cover of the АВАР СБРОС

3. Tactical release button

(EMERGENCY RELEASE) button

[RCtrl +

[RCtrl + Rshift + Ralt +L]

Rshift + Ralt + R]

[RCtrl + Rshift + Ralt + Q]

(only if safety

2. Emetgency release button

cover is opened)

[RCtrl + Rshift + Ralt + P]

4. Safety cover of the ТАКТ СБРОС ГРУЗА

[RCtrl + Rshift + Ralt + A]

(only if safety

(TACTICAL CARGO RELEASE) button

[RCtrl + Rshift + Ralt + T]

cover is opened)

There is no any difference for player in tactical or emergency cargo release during

flight, result is completely the same. The only difference between tactical and

emergency release is that release signal goes via different electrical circuits to open

the DG-64 cargo hook. In both cases, when hook is opened, the ЗАМОК ОТКРЫТ

(HOOK OPEN) light panel is on

Cargo hook is closed manually (in this game it closes automatically, if cargo is

selected with help of radio menu and player hovers over the cargo for some amount

of time, see here).

Sling load operation (selection of the cargo, attachment and release) is described

here.

7.11. АП-34Б (AP-34B) autopilot system

7.11.1. General description

The AP-34B four-channel electrical-hydraulical autopilot system is designed to

stabilize the helicopter in roll, pitch, heading, altitude, and airspeed. The autopilot is

214

a system, receiving information about changes in angular positions of the helicopter,

barometric altitude and instrumental airspeed. It has four independent autopilot

channels, controlling corresponding helicopter’s controls:

 yaw channel - tail rotor pitch;

 roll channel - swashplate in lateral direction;

 pitch channel- swashplate in longitudinal direction;

 altitude channel - main rotor pitch;

When the altitude channel is engaged, the pitch channel receives correction signals

from the КЗСП (KZSP) airspeed correction unit to stabilize the airspeed.

The four autopilot channels (roll, pitch, yaw, altitude) provide:

 stabilization of helicopter’s position in three axes (longitudinal, lateral,

vertical);

 stabilization of altitude, during sustained level flight and hover;

 stabilization of indicated airspeed.

AP-34B autopilot system specifications:

Supply voltage, DC

+28.5 V

Supply voltage, AC

~36 V (3Ph), 400 Hz

Time to readiness

under 2 min

Controls travel range in percents of full travel range, given to the

20%

autopilot:

Autopilot stability in calm atmosphere conditions:

yaw channel

±1°

roll channel

±0,5°

pitch channel

±0,5°

altitude

±10 m

airspeed

±10 km/h

Max altitude

10000 m

Weight

under 25 kg

7.11.2. Autopilot system’s unit

The AP-34B autopilot system includes the following units:

 control panel

;

 control unit, it is needed for transforming, accumulating and amplifying of

the control signals (fully modeled);

 yaw, roll and pitch angular velocity sensors (fully modeled);

 amplifiers unit (fully modeled);

 ИН-4 (IN-4) zero indicator

;

 three compensation sensor (in each channel, fully modeled);

 КВ-11 (KV-11) altitude corrector (fully modeled).

The АП-34Б (AP-34B) autopilot and control system mechanisms are powered by 27V

DC from the onboard battery and rectifier buses bus via the АВТОПИЛОТ-ОСНОВН.

(AUTOPILOT-MAIN), АВТОПИЛОТ-ФРИКЦ. (AUTOPILOT-FRICT) and АВТОПИЛОТ-

ЭЛЕКТРОМУФТЫ (AUTOPILOT-SOLENOID CLUTCHES)

circuit

breakers. All circuit breakers are located on the right circuit breaker console.

Additionally, system is powered by the 36 VAC 400 Hz from 36 V three-phase

generator buses.

7.11.3. Autopilot controls and indicators

Control panel

The autopilot control panel is located on the center console and provide the following

functions:

 zeroing of autopilot control input prior to engaging the autopilot;

 individual engagement/disengagement of autopilot channels;

 entering of small corrections (±10%) using the centering knobs for pitch,

roll, and yaw channels;

 testing of the altitude channel using the "КОНТРОЛЬ" (TEST) switch.

216

Fig. 7.44. Autopilot control panel

1. Lamp-button, enabling yaw channel

6. Rotating scales, indicating mismatch

2. Lamp-button, disabling yaw channel

between signals from yaw, roll and pitch

3. Lamp-button, enabling pitch and roll

sensors and actual controls position

channels

(1 mark corresponds to 1°)

4. Lamp-button, disabling altitude channel

7. КОНТРОЛЬ (CONTROL) button for checking

5. Lamp-button, enabling altitude channel

the altitude channel

8. Pitch channel centering knob

9. Roll channel centering knob

10. Yaw channel centering knob

ИН-4 (IN-4) zero indicator

The ИН-4 (IN-4) zero indicator unit shows the relative displacement of the flight

control servo spindle for each autopilot channel. It is located on the center console.

Fig. 7.45. ИН-4 (IN-4) zero indicator

1. Maximum deflection marks

4. Pitch channel indicator

2. Yaw channel indicator

5. Altitude channel indicator

3. Roll channel indicator

In other words, these indicators show center position of the control range, given to

the autopilot, relative to the helicopter controls such as pedals, collective or cyclic.

See below for explanations.

7.11.4.

Equipment interacting with the autopilot system



The GMK-1A gyro compass system provides signals to the autopilot

control panel to stabilize a given heading of the helicopter. Besides that,

this system gives signal to the control panel to engage

alignment/matching mode for the yaw channel, when course set mode

and course system check are enabled. This functionality is enabled, when

the НАПРАВЛЕНИЕ (YAW) channel is engaged.



AGB-3K ADI of the co-pilot is a roll and pitch sensor for the autopilot

system. It is used when the ROLL-PITCH channel is engaged.



The КЗСП (KZSP) airspeed correction unit is designed to provide an

electrical signal, proportional to the deviation of the actual speed from the

required one. The КЗСП (KZSP) affects pitch channel. It is used, when the

altitude channel is engaged.



The БСГ (BSG) ready signal unit operates together with the КЗСП (KZSP)

and provides ready signal, when КЗСП (KZSP) correction calculation

system is operational.



Combined KAU-30B control units (control actuator) and RA-60B steering

unit (control actuator) are executive units for both helicopter and

autopilot. They affect helicopter’s controls (swashplate of main rotor and

pitch control of tail rotor). Four steering units are installed in lateral and

longitudinal channels, collective channel and pedals channel. These

steering units have built-in feedback sensors, with help of which autopilot

receive feedback, see below.



Electromagnetic valves (GA-192), using signals from the autopilot, engage

steering units in autopilot mode (details described below).



Two ВЫКЛ. АП (AUTOPILOT OFF) buttons are installed on cyclics

[LWin

+ LShift + A]



Two buttons disengaging altitude channel, installed on the collectives

(their press is emulated when position of collective has been changed or

collective control buttons have been pressed -

[Num - ]

[Num +]



Four small-sized switches on the pedals. When pilot puts feets on the

pedals, these switches trigger and alignment mode engages for the yaw

channel.



Small-sized switch, located on mechanical rocker of the SPUU-52 tail rotor

pitch limit system, which is used to engage alignment mode for yaw

channel, when mechanical rod in pedal channel is close to the mechanical

stop, position of which is defined by the SPUU-52 system (i.e. yaw

channel disengages when mechanical stop is reached).



Time relay with

0.5s delay, included in yaw channel, to prevent

oscilations, when autopilot switches to alignment mode.

218

 Autopilot’s distribution box (not modelled).

Due to fact that the AP-34B autopilot operates together with the KAU-30B and RA-

60A electro-hydraulical executive units (control actuators), controls of the helicopter

can be controlled by the pilot, using normal helicopter controls (cyclic and collective)

and by the autopilot simultaneously (so called combined mode, when pilot can

overrule autopilot at any time while autopilot is engaged). In this mode, autopilot

affect small cylinder of the steering unit, which is, in its way, an adjustable rod,

included into the control system. Due to that, resulting movement of the controls is a

sum of pilot’s and autopilot’s control inputs. Movements of the executive steering

units, due to signals from the autopilot, are not transferred back to cyclic and

collective.

Besides that, the autopilot can operate in the automatic helicopter stabilization mode

(when all four channels are engaged). Peculiarities in operation of the hydraulic units

in cases of autopilot and manual control are described below.

When the altitude channel is engaged, the pitch channel receives correction signals

from the КЗСП (KZSP) airspeed correction unit.

Autopilot roll, pitch, and altitude correction signals are limited to a maximum of 20%

of control travel for flight safety in the event of false signals or system failure. In

addition, the pilot may intervene at any time while the autopilot is engaged to make

manual corrections by operating the flight controls.

7.11.5. Use of autopilot in different flight phases

The autopilot system is engaged for all normal flight operations. The pitch and roll

channel is normally engaged throughout the flight from takeoff to landing. Yaw and

altitude channels are used rarely.

The autopilot is engaged before takeoff, by pressing lamp-buttons, corresponding to

the required channels

. The altitude channel can not

be engaged without engaging “ROLL-PITCH” channel.

When performing vertical takeoff, one must enable “ROLL-PITCH” channel

(pilots, operating Mi-8 in Russian Federation, normally do not engage “YAW”

channel before takeoff).

When performing rolling takeoff, engage only “ROLL-PITCH” channel. Engaging

must be monitored with help of lamp-buttons, which should be green for engaged

channels.

When in hover, the autopilot stabilizes the helicopter in pitch and roll, as well as

heading when the pedals are released (feet off the pedals). Autopilot functionality in

hover mode can be verified by checking the zero indicator unit for fluctuations in the

"К" (roll channel) "Т" (pitch channel) "Н" (yaw channel) servo displacement indicator

needles around neutral positions.

In sustained flight conditions such as level flight, climb or descent with flight

controls released by the pilots, the autopilot will stabilize the helicopter attitude while

slowly decreasing airspeed, because the system is maintaining a pitch angle and not

airspeed (up to 150 km/h).

Piloting the helicopter with help of autopilot without moving controls

To adjust a helicopter heading, within ±5° range during sustained level flight, while

autopilot is engaged, one must rotate the centering knob on the autopilot control

panel clockwise or counterclockwise

Complete rotation of this knob

results in heading change of 10°. There are ten ticks on the scale, each tick

corresponds to 1° . In a similar way, the roll and pitch can be adjusted.

The altitude channel can be engaged after establishing level flight at an altitude of

not less than 50 m. Altitude channel operation can be verified by fluctuations of the

"В" (altitude channel) servo displacement indicator needle on the zero indicator unit,

changing of the main rotor collective pitch angle on the UP-21-15 gauge, and vertical

displacement of the helicopter as the autopilot system maintains altitude in turbulent

air.

Landing approach, braking, and landing are normally performed with the autopilot

channels “ROLL” and “PITCH” engaged. After landing, the autopilot system is

disengaged with the "ВЫКЛ. АП" (Autopilot OFF) button, located on both cyclics

7.11.6. Explanations of the AP-34B peculiarities and key commands

Combined operation of hydraulic boosters and the AP-34B

autopilot in case of manual control

In case of automatic helicopter stabilization, the actuating rods of the KAU-30B combined hydraulic

boosters can move themselves within 20% of their full travel range, wherein control sticks (cyclics) do

not move and are hold in the same fixed positions by the spring load mechanism. Collectives are hold

by the friction mechanism.

Limitation of 20% (by design of steering units/ control actuators) of full travel range of is needed to

provide flight safety in case of autopilot failure, because the most of the failures are accompanied with

appearance of an one-way signal on the autopilot output and, as a result, fast reaction of the actuator

units.

220

At the same time, this limited operating range of actuator units using autopilot signals is sufficient for

compensation of real-life distrubances, affecting helicopter and for stabilization of angular positions of

the helicopter.

The pilot may intervene at any time while the autopilot is engaged. For this purpose are centering

knobs for pitch, roll and yaw channels, full rotation of which corresponds to 10° change. There are ten

ticks on the scale, each tick corresponds to 1° .

By rotating correspondinig centering knob clockwise or counterclockwise, pilot force the helicopter to

turn. When pilot intervene roll or pitch, it is necessary to remove impact of signals from angular and

angular velocity sensors on hydraulic boosters of roll and pitch channels. For this purpose the

compensation sensors, kinematically connected with lateral and longitudinal controls, which provide

signals, equal to ones received from attitude indicator, but with opposite sign, are installed. In other

words, when pilot moves cyclic, the reference value of roll and pitch (which autopilot will maintain) is

constantly being updated.

In this way pilot controls helicopter’s roll and pitch, without disengaging autopilot, which is constantly

trying to maintain helicopter’s position, given by the pilot. For the pilot, to control yaw, on the pedals

are triggers and microswitches, which, when pressed, enable alignment mode for yaw channel. When

maneuver is finished and pedals are released, the autopilot for yaw channel engages automatically.

Game peculiarities, when pedals are used and "НАПРАВЛЕНИЕ"

(YAW) channel is engaged

To control yaw in the Mi-8MTV2, the RA-60A steering unit is used. Unlike the KAU-

30B, it has a special mechanism, allowing autopilot to move pedals, if limited travel

range of 20% is not sufficient for maintaining given direction.

If, for example, yaw channel is engaged and player during vertical ascend is not

compensating reactive moment of the main rotor with pedals, then after using 20%

of limited travel range, the RA-60A will be moving the right pedal forward without

any actions from player. Player pedals will remain in the same position. To align

position of player pedals with position of virtual pedals in Mi-8 model, perform the

following actions:

 return gaming pedals to a neutral position (if pedals have springs - simply

remove feets from the pedals);

 reset trimmer

[LCtrl + T]

After this procedure the pedals in the model return to a neutral position and they will

be aligned with gaming pedals.

NOTE. Take into account that this procedure returns cyclic into neutral position, as well.

Trimmer features when autopilot is engaged

As in real life, in our model interaction between trimmer and autopilot in the Mi-8 is different from the

one, implemented in the Ka-50. In the Mi-8 depression and holding of the trimmer button do not

engage alignment mode of the autopilot, as it is done in the Ka-50. I.e. trimmer as object is not

interacting with autopilot at all.

Autopilot controls roll (pitch) within 20% of full cyclic travel range for this channel. The edge positions

on the IN-4 zero indicator (Fig. 7.45, 1) tell us that hydraulic boosters of roll or pitch (more precisely

the small cylinders of the steering unit) reached 20% of travel range, given for automatic stabilization

of roll and pitch angles. This appears as incrased sensitivity in helicopter reaction on small cyclic

movements to the side where limit was reached and creates certain discomfort for piloting.

The deviation of the center of the 20% zone, given to autopilot to stabilize helicopter, from current

position of the cyclic, can be estimated with help of IN-4 (Fig. 7.46, 2). If pilot, using trimmer, moved

cyclic outside this zone (according to indications on the IN-4), then autopilot is not able to

“comfortably” stabilize roll (pitch) anymore, as before. Of course, pilot feels discomfort as well, even

when applied forces were removed from cyclic by the trimmer. In practice, it requires increased

amount of small cyclic corrections, needed for maintaining desired roll or pitch. To ease piloting, it is

necessary align the center of autopilot’s 20% of range with the current position of the cyclic. In other

words,“point” the autopilot to a new reference position. This can be done in the two following ways:

1. By Centering knobs on the AP-34B control panel. In real life, it is a crew chief, who

normally, by rotating these knobs, removes misalignment, by setting roll and pitch according to

horizontal marks. Pilot, during this process, must slowly move cyclic towards new position to maintain

sustained flight, i.e. moving away from the position where autopilot was engaged and then trimming

the cyclic once again.

2. By disabling autopilot and balance correction followed by the autopilot engagement.

The first approach is useful, because it gives an opportunity to smoothly remove misalignment, the

second allows to do it fast without involving other crew members.

To better understand mentioned above features, see below.

Positions of Cyclic and zero indicators on IN-4 during Hover and

Level fligh (Roll-Pitch channel)

Below, the positions of cyclic, depending on moment when the “ROLL-PITCH”

channel was enabled, are shown for hover and level flight.

HOVER

Fig. 7.46. Hover

1. Hover.

2. Hover.

AP-34B (Roll-Pitch channel) is OFF

AP-34B (Roll-Pitch channel) was turned ON while

on the ground, when Cyclic is in neutral position

222

Fig. 7.48. The "Autopilot Ajustment" checkbox in special settings

When AI crew chief receives a command from the player to adjust the autopilot it

will “report” about actions taken.

AP-34B Key Commands

LAlt + LWin + A

Autopilot Altitude Channel OFF Button

LAlt + A

Autopilot Altitude Channel ON Button

LAlt + V

Autopilot Altitude Control Switch - Down

LAlt + F

Autopilot Altitude Control Switch - Up

LShift + LWin + S

Autopilot Heading Adjustment Knob - CCW/Left

LShift + LWin + D

Autopilot Heading Adjustment Knob - CW/Right

LCtrl + LWin + A

Autopilot Heading Channel OFF Button

LCtrl + A

Autopilot Heading Channel ON Button

LCtrl + LShift + S

Autopilot Pitch Adjustment Knob - CCW/Left

LCtrl + LShift + D

Autopilot Pitch Adjustment Knob - CW/Right

LWin + RCtrl + S

Autopilot Roll Adjustment Knob - CCW/Left

LWin + RCtrl + D

Autopilot Roll Adjustment Knob - CW/Right

LWin + A

Autopilot Roll/Pitch Channel ON Button

7.12. Exhaust IR suppression devices

Installs exhaust IR suppression devices (Fig. 7.49). It decreases engines’ IR

signature by approximately two times. This reduces lock on range for IR SAMs and

increases probability of successful flight through SAM protected areas.

224

DIGITAL COMBAT SIMULATOR Mi-8МТV2

8. OPERATING LIMITS AND RESTRICTIONS

8.1. Calculating maximum takeoff weight

Maximum takeoff weight for out of ground effect vertical takeoff (landing) (OGE

maximum hover weight) is displayed by Fig. 8.1. Maximum takeoff weight for in

ground effect vertical takeoff (landing) (IGE maximum hover weight) is displayed by

Fig. 8.2.

The maximum hover weight charts display maximum takeoff weight in relation to the

pressure altitude of the landing field and free air temperature (FAT) assuming calm

winds, 93% main rotor RPM, disengaged PZU Air Inlet Particle Separator System

(PSS), disengaged anti-ice systems.

Fig. 8.1. OGE maximum hover weight chart (hover altitude 20 m).

PZU and anti-icing disabled

N O T E . If EGS installed, reduce maximum weight indicated in chart by 300 kg.

With PZU (PSS) turned on, reduce maximum takeoff weight indicated in chart by 200

kg. With engine and rotor anti-ice systems turned on, reduce maximum takeoff

weight indicated in chart by 1000 kg.

When Exhaust IR suppression devices (7.12) are fitted reduce maximum takeoff

weight indicated in chart by 300 kg.

Fig. 8.2. IGE maximum hover weight chart (hover altitude 3 m).

PZU and anti-icing disabled.

Any headwind increases maximum takeoff weight: +200 kg at 5 m/s; +1200 kg at

10 m/s.

Crosswind up to 5 m/s reduces performance by affecting the tail rotor and increasing

engine power requirements. Reduce maximum takeoff weight by 200 kg in the

presence of a crosswind of up to 5 m/s. At greater crosswind speeds, translational lift

effects become more dominant.

Performance reduction in tailwind conditions (blowback of hot exhaust gases into the

exhaust system) is not modeled in the simulation.

When calculating wind corrections for maximum hover weight, consider that wind

speed and direction may vary during takeoff/landing. Assume the lowest maximum

hover weight corresponding with possible wind variance.

If wind conditions cannot be determined, assume poor hover conditions of 4-6 m/s

tailwind.

228

EXAMPLE:

Fig. 8.2 includes a solution (orange arrows) to the following example problem: determine the

maximum hover weight for vertical takeoff in ground effect from an airfield located at an altitude of

2,300 m and +30°C FAT.

SOLUTION:

Using the IGE maximum hover weight chart Fig. 8.2, enter the graph from the left at the point of the

desired pressure altitude of 2,300 m. Draw a line horizontally to intersect the desired temperature of

+30°C. From the intersection point, draw a vertical line down to find the maximum hover weight

value, in this case 11,680 kg.

To determine the maximum takeoff weight for a vertical takeoff out of ground effect, perform the

same process using the OGE maximum hover weight chart Fig. 8.1.

MAXIMUM TAKEOFF WEIGHT FOR A RUNNING TAKEOFF:

To determine the maximum takeoff weight for a running takeoff, utilize the IGE

maximum hover weight chart Fig. 8.2, but add an additional 500 kg to the solution.

Prior to performing a running takeoff, execute a test hover to an altitude of no less

than 1 m to verify correct maximum weight calculation.

MAXIMUM TAKEOFF WEIGHT FOR A NOSE WHEEL RUNNING TAKEOFF:

Use nose wheel running takeoff maximum takeoff weight chart Fig. 8.3 to determine

the maximum takeoff weight for a nose wheel running takeoff.

Fig. 8.3. Nose wheel running takeoff maximum takeoff weight chart

Execute a test hover to verify correct maximum weight calculation prior to

performing a nose wheel running takeoff. The takeoff can be performed if the

helicopter is able to lift off the ground during the test hover.

In all cases, the maximum takeoff weight should never exceed the helicopter

maximum gross weight of 13,000 kg.

The calculate fuel and cargo weight limitations, individual helicopter component

weights are provided in the Table 8.1:

Table 8.1

Weight, included in

simulation model (as

Mi-8MTV2

May be

"equipped empty

weight ")

Empty helicopter weight

7200

Cabin ladder

7.3

Safe

4.5

Not modeled

Auxiliary fuel tank

Not modeled

Troop seats (30)

58.7

Cargo ramps

31.6

Not modeled

ЛПГ-150м (LPG-150M) hoist

33.5

Mounts for assault rifle

19.8

Waste equipment

93.6

Not modeled

Assault ropes

Not modeled

Unusable fuel

Main fuel 0.775 kg/L

counted in the mission

- service tank

322

editor

counted in the mission

- main tanks

1608

editor

- auxiliary tanks

1388

Not modeled

Armor plating

419

- cockpit

332

- cargo cabin

- hydraulics compartment

Weapons stations racks

401

Nose PKT 7.62-mm machine gun w/

38.8

Not modeled

ammo

Tail PKT 7.62-mm machine gun

counted in the mission

26.9

w/ammo (3 box at 250 cartridges)

editor

Door Machine gun 'KORD" 12.7-mm with

mounting device and on-board

counted in the mission

138

ammunition supplies (12 box at 50

editor

cartridges)

UV-26 (self-defense EW system) with

26.3

128 flares

counted in the mission

B8-V20A launchers w/t ammo

100

editor

counted in the mission

Rockets (20 x 80-mm S-8OM)

242

editor

GUV-1 (12.7-mm +7.62-mm machine

counted in the mission

452

gun)

editor

counted in the mission

GUV-1 (30-mm grenades)

274

editor

counted in the mission

UPK-23-250 (23-mm mashin gun)

230

editor

Minelaying container (empty)

Not modeled

Crew O₂ equipment

19.3

Not modeled

Rope ladder

19.7

Not modeled

PKV sight

2.5

OPB-1r sight

8.2

Not modeled

Junction box of weapon with Bombing

15.9

230

weapon Control Panel

Rescue eqiupment:

216

Not modeled

- LPG-300 hoist

- hoist installation

95.5

- Junction box with fixing system

15.6

- spot light

20.5

- rescue basket

- boat-hook

0.9

- hand spot light (2)

- rescue belt (2)

- crew cheif belt

1.8

- rubber mat

6.6

- rescue seat

- other

5.4

External sling system:

160.5

- keylock DG-64M

21.3

- Safeguard (skirt) of bottom

9.3

flap

- main ropes (4x1.11m)

Releasable components:

119.9

- sling cable type 1.7m

3.15

- sling cable type 5m

7.3

- sling cable type 10m (2)

27.3

- sling cable type 20m (2)

52.5

- lower strap (4x4m), or cargo

hook

- minor components

5.6

Rotor blade mooring equipment

43.3

Not modeled

L166V1A (self-defense EW system)

Not equipted

Total equipment weight, kg

1164.5

Lubrication oil

71.7

Crew of three

270

Total helicopter weight (oil, crew) WITHOUT fuel

8706.2

and armament, kg

8.2. Calculating flight range, radius, and time

This section provides the required data for calculating flight navigation planning.

Flight range (radius) and flight time depend on fuel quantity and consumption rate,

which in turn depends on helicopter weight, payload (which affects aerodynamic

performance (in particular drag)), flight altitude and airspeed.

The effects of these factors on flight distance and time are examined below.

ALTITUDE. Helicopters are generally flown at low altitudes. However if long range

operations are required, flight at altitudes of 2000 - 3000 m result in approximately

15% greater range than low altitudes.

AIRSPEED. Greatest flight range is achieved at or near optimum cruise speed (±20

km/h)

CRUISE SPEED - optimum speed that provides greatest flight range (minimum fuel

consumption rate) is provided in Table 8.2.

Table 8.2

Helicopter weight 11 100 kg

Helicopter weight greater than

or less

11 100 kg

Altitude, m

Airspeed (km/h)

indicated

true

indicated

true

100

230

233

215

219

500

225

233

210

218

1000

220

233

205

218

2000

210

234

195

218

3000

195

230

160

190

4000

170

213

120

154

5000

120

163

6000

100

145

AERODYNAMIC FACTORS

When Exhaust IR suppression devices (7.12) are fitted, fuel consumption rates per

kilometer and per hour provided in Table 8.4 increase by 6%.

With armament fitted, fuel consumption rates are as indicated in Table 8.4.

ENGINE AIR BLEED FACTORS

With anti-icing and particle separator system engaged, fuel consumption rates

indicated in the tables below increase as follows:

 engine anti-icing: 3%

 main and tail rotor anti-icing: 2%

With the PZU particle separator system engaged, fuel consumption rates per hour

provided in Table 8.4 increase by 3%.

MINIMUM FUEL QUANTITY

To ensure flight safety, a minimum fuel quantity is provided, which equals 260 L for

the Mi-8MTV2.

GROUND RUNUP FUEL CONSUMPTION (G_T3) includes:

 fuel required for engine start, warm-up, and taxi: 30 kg/5 min (6 kg/min)

 fuel required by the APU while powering the electrical systems prior to

engine start: 1.25 kg/min)

GROUND TARGET ATTACK RUNS consume 12 kg of fuel/min on the first run. With 4-minute

repeat attack runs, each is estimated to consume 50 kg, which is equivalent to a

reduction of flight radius by approximately 10 km.

232

Fuel consumption rates required from takeoff to altitude are provided below in Table

8.3 (nominal engine power setting).

Fuel consumption, distance, and time required from takeoff to altitude:

Table 8.3

Helicopter weight, kg

11000

12000

13000

Indicated

Altitude, m

airspeed,

km/h

Takeoff

and climb

100

120

1.5

500

120

1000

120

2.5

2000

120

4.5

5.5

3000

110

100

4000

110

7.5

115

140

11.5

4800

100

215

5000

100

115

9.5

155

6000

170

FUEL CONSUMPTION RATES AT VARIOUS WEIGHT, ALTITUDES AND AIRSPEEDS FOR BOTH MAXIMUM

RANGE AND DURATION'S LEVEL FLIGHT, see Table 8.4

DIGITAL COMBAT SIMULATOR Mi-8МТV2

Per kilometer and per hour fuel consumption rates at various altitudes and airspeeds for maximum rage depending on helicopter

weight.

Main rotor RPM 95%

Table 8.4

Fuel consumption rate vs helicopter weight (kg)

9000

10 000

11 000

12 000

13 000

Н, m

Qmin кgf/hr/

at Vind,

kgf/k

kgf/h

at Vind,

kgf/

kgf/k

kgf/h

kgf/km

kgf/hr

at Vind, km/h

kgf/km

at Vind, km/h

km/h

Transport configuration (w\o Weapons stations racks)

100

2.66

620

445 / 120

2.69

627

470 / 120

2.75

641

495 / 120

2.84

621

520 / 110

2.93

640

550/ 120

500

2.55

593

445 / 120

2.6

605

455 / 110

2.67

621

485 / 110

2.76

601

515 / 110

2.84

623

545 / 110

1000

2.44

569

425 / 120

2.49

580

450 / 120

2.57

599

475 / 120

2.66

587

505 / 120

2.77

614

540 / 120

2000

2.24

525

400 / 100

2.33

546

425 / 100

2.42

572

455 / 110

2.56

559

490 /130

2.72

592

530 / 120

3000

2.11

485

380 / 100

2.23

510

410 / 110

2.36

540

445 / 120

2.65

500

480 / 120

2.94

554

535 / 120

4000

426

370 / 100

2.14

455

400 / 110

2.36

502

445 / 120

3.23

487

495 / 120

3.85

575

580 / 110

5000

2.09

354

360 / 100

2.37

406

400 / 110

2.8

488

470 / 110

Combat configuration (with Weapons stations racks) without armament

100

2.75

643

445 / 100

2.81

660

475 / 100

2.87

676

500 / 115

2.95

651

530 / 120

3.03

673

555 / 120

500

2.67

630

435 / 110

2.73

646

460 / 110

2.79

663

490 / 110

2.85

640

515 / 110

2.94

660

545 / 115

1000

2.55

613

425 / 110

2.6

629

450 / 110

2.66

648

480 / 110

2.75

627

505 / 110

2.9

651

540 / 110

2000

2.34

570

400 / 100

2.42

586

425 / 100

2.53

610

460 / 100

2.63

599

500 / 115

2.81

638

540 / 115

3000

2.21

515

385 / 105

2.29

540

415 / 115

2.44

581

445 / 120

2.7

527

490 / 120

579

550 / 120

4000

2.07

447

370 / 120

2.23

477

405 / 110

2.5

523

450 / 115

3.46

522

525 / 100

4.17

596

630 / 100

5000

2.12

375

360 / 100

2.45

431

415 / 100

3.03

507

520 / 100

Combat configuration with armament

100

2.83

692

450 / 110

2.91

697

480 / 115

2.99

706

505 / 115

3.07

673

530 / 110

3.15

686

560 / 120

500

2.75

679

440 / 110

2.83

681

460 / 110

2.91

684

490 / 110

2.99

649

520 / 110

3.07

662

550 / 115

1000

2.66

645

430 / 110

2.75

650

455 / 110

2.82

657

480 / 100

2.9

624

510 / 110

2.99

642

545 / 120

2000

2.45

586

405 / 100

2.52

593

433 / 100

2.6

607

465 / 110

2.7

594

505 / 130

2.86

625

550 / 120

3000

2.26

522

385 / 100

2.36

536

415 / 100

2.48

559

450 / 120

2.75

537

495 / 120

3.13

617

560 / 110

4000

2.12

464

375 / 120

2.28

481

405 / 110

2.6

548

455 / 110

3.65

553

545 / 100

4.5

684

695 / 110

5000

2.15

376

365 / 105

2.61

445

425 / 100

3.6

553

595 / 100

q, kgf/km - fuel consumption in kg for 1 km; Q, kgf/hr - fuel consumption in kg per hour; Vind - indicated airspeed

DIGITAL COMBAT SIMULATOR Mi-8МТV2

FUEL CONSUMPTION, DISTANCE, AND TIME REQUIRED IN LANDING APPROACH

Table 8.5

Indicated

Fuel

Starting altitude, m

airspeed,

Descent rate, m/s

consumption,

Distance, km

km/h

kgf

Deceleration, hover,

and landing

100

120-130

2-4

500

140-150

5-6

1000

140-150

5-6

2000

140-150

5-6

3000

140-150

5-6

4000

120

3-4

5000

120

3-4

130

FUEL CONSUMPTION PER HOUR IN HOVER (KGF/HR) OUT OF GROUND EFFECT

Table 8.6

Fuel consumption per hour (kgf/hr) vs airfield altitude, m

Helicopter weight

500

1000

2000

3000

9000

700

660

640

630

610

10000

730

710

700

690

11000

790

770

760

12000

850

840

13000

920

To calculate the required fuel for a given mission range, calculate the fuel required

for non-navigation phases of the mission (start, taxi, target attack run(s)), add the

minimum fuel quantity, then subtract this sum from the total fuel quantity on board.

Multiply the remainder by 0.95 to add 5% navigational error factor and another 0.95

to add 5% for formation keeping.

8.3. Engines and transmission limits

8.3.1. Engines limits

TV3-117VM Maximum Operating Limits (all Altitude, all Ambient temperature) see in

Table 8.7

Power setting

MAX PTIT (°C)

MAX N1 (%)

Max Rated

990

101.0

Take Off

990

101.0

MAX LTD Cruise

955

99.0

LTD Cruise

910

97.5

Cruise

870

95.5

IDLE

780

See Fig. 4.3

TV3-117VM оperating range see in Table 8.8

Table 8.8

RPM

Oil Temp °С

Oil

POWER

MAX Time

Pressure

SETTING

Allowed, minutes

(kgf/cm2)

IDLE

See Fig. 4.3

40-55

55-70

CRUISE

95 ± 2

3,5±0,5

150

80-140

No limit

in accordance

with the IR-117,

Fig. 8.4, but

does not exceed

LTD CRUISE

the values of the

95±2

3,5±0,5

150

80-140

No limit

Table 8.7

MAX LTD

95±2

3,5±0,5

150

80-140

CRUISE

TAKE OFF

93±1

3,5±0,5

150

80-140

MAX RATED*

93±1

3,5 ±0,5

150

80-140

(see **)

*When one engine has failed, the operating engine automatically elevates power to MAX Rated

available. MAX Rated Power operating mode can not be activated for both engines simultaneously.

Other translate: MAX Rated Power operating mode one of two engine can be activated only when the

other engine failure (ie any action of the crew with (for) two simultaneously operating engines can not

be set MAX Rated Power).

**Exceeding 6 minutes of operating time in the EMER (MAX RATED) /Take Off settings or the time

limits for other power settings, will result in a reduction in engine service life.

N1% Limits Adjusted for Ambient Temperature see in Fig. 8.4

236

Fig. 8.4. N1% Limits Adjusted for Ambient Temperature

1. Maximum allowed N1 at standard atmospheric

4. MAX LTD CRUISE Power area

pressure

5. LTD CRUISE Power area

2. MAX RATED Power area

6. CRUISE Power area

3. TAKE OFF Power area

Example. Determine Maximum and Minimum N1 of Take-Off Power Setting mode for Ambient

Temperature at +24°C.

Key (for Minimum N1): enter the graph from the bottom applying the reported ambient temperature

(+24, a). Proceed vertically to the MIN Take Off power setting diagonals (b). Continue from the

intersecting point to the left to obtain the Minimum N1 (97.2%, c). Maximum N1 is 99.4%.

NOTE. This chart applies to standard atmospheric pressure. Apply the N1% established in this chart to

Fig. 8.5, N1 Adjusted for Barometric Pressure to find the N1 for the power setting required.

Fig. 8.5. N1% Adjusted for Barometric Pressure

8.3.2. Transmition limits

Main Transmition Maximum Operating Limits:

a) oil pressure:

 IDLE mode - > 0,5 kg/sm²;

 other Power Setting mode - 3.5±0,5 kg/sm²;

b) oil temperature:

 MAX 90°С;

 recommended 50 - 80°С;

 MIN Initial Oil Temp (for operating more IDLE) -15°С;

 MIN Oil Temp Continuous Operation +30°С.

c) Main Rotor RPM limitations see Table 8.9

Table 8.9

Absolute Limits:

RPM, %

MAX Time Allowed

MAX Rated & Take-Off Power

103%Maximum

10 Seconds

MAX Rated & Take-Off Power

88% Minimum

30 Seconds

All Settings above LTD Cruise

101%Maximum

20 Seconds

All Settings below LTD Cruise

103%Maximum

20 Seconds

Normal Operating Limits:

RPM, %

MAX Time Allowed

IDLE

55 to 70 (40 to 55 Single Engine)

20 Minutes

CRUISE

97% Maximum

Not Limited

LTD CRUISE

97% Maximum

60 Minutes

MAX LTD CRUISE

97% Maximum

60 Minutes

TAKE OFF

94% Maximum

6 to 15 Minutes

MAX RATED POWER

94% Maximum

6 to 60 Minutes

MAX oil temperature in Intermediate gearbox and Tail rotor gearbox is 110°С.

238

 Set the pressure altimeter pointers to zero and check the barometric

pressure display for compliance with the actual aerodrome pressure with

an accuracy of ±1.5 mm Hg

[LShift + B]

[LCtrl + B]

 energize the helicopter electrical systems

 check crew intercommunication over the SPU-7 ICS (in case multicrew

game)

 check windshield wiper operation

 check and set the clock;

 check fuel quantity on the fuel gauge and set the fuel gauge selector to

"РАСХ" (SVC CELL) [2] (move to the co-pilot's seat),

[RCtrl + RShift + V]

for CW turn

[RCtrl + RShift + B] for CCW turn.

 release the rotor brake

[RCtrl + R]

(down)

 make sure the collective pitch control lever is at the lower stop [Num -]

and the throttle control twist grip is turned fully to the left

[PgDn]

, the

engines throttle levers are set in the neutral detent

[RCTRL + HOME /

RCTRL + END]

, the control stick is in a position close to neutral and the

fuel shutoff levers are in the aft (closed) position

[ RCtrl + PageUp..

PageDown]

9.2. Preparation and equipment check procedures preceding

APU start

The power supply is required to start the APU. APU start can be made with or using

batteries (A), or an external power supply (B).

9.2.1. Simplified procedure

A. Turn on onboard electrical power:

1. Right Side Panel - DC POWER:

АККУМУЛ I, II (Battery 1, 2) - ON (UP),

check the voltage of the battery bus by setting

the DC selector knob to "ШИНЫ АКК." (BATT

BUS). The voltage should be no less than 24 V.

242

[\]

(Radio Menu),

[F8]

[F2]

[F1]

(Connect Ground Electric Power).

6. Once the "АЭР. ПИТ. ВКЛЮЧЕНО" (EXT PWR

ON) light illuminates (after successful connection

to an external power source), and check the DC

ground power source voltage by setting the DC

selector knob to "АЭРОДРОМ. ПИТАН." (EXT

PWR). The voltage should be within the limits of

27 - 29 V.

7. Follow the procedures, described in the

chapter A

9.2.2. Full procedure

While performing complete preparation procedure, one must connect power sources

and continue extensive control and preparation of equipment for start.

Connection of power supplies

A. Turning on the batteries:

 "АККУМУЛ. I" и "II" (BATT I, II) to "ВКЛ." (ON (up)) [LCtrl + LShift

+ 1..2]

;

 check the voltage of the battery bus by setting the DC selector knob to

"ШИНЫ АКК." (BATT BUS). The voltage should be no less than 24 V.

 check the condition of the batteries as follows:

244

a) "АЭРОДР. ПИТАН." (EXT PWR (external power)) selector to "ВЫКЛ."

(OFF) [LCtrl + LShift + 7]

;

b) DC selector knob to "АККУМУЛ. I" (BATT I)

[LCtrl + LShift + 9..0]

c) "АККУМУЛ. II" (BATT II) to "ВЫКЛ" (OFF (down))

[LCtrl + LShift +2]

d) any fuel boost pump to ON and check Volts (no less than 24 V)

[RShift

+ 1]

e) DC selector knob to "АККУМУЛ. II" (BATT II)

[LCtrl + LShift + 9..0]

f) АККУМУЛ. II" (BATT II) to "ВКЛ" (ON (up))

[LCtrl + LShift +2] and

"АККУМУЛ. I" (BATT I) to "ВЫКЛ" (OFF (down))

[LCtrl + LShift +1]

Check voltage (no less than 24 V).

g) fuel boost pump to OFF

[RShift + 1]

;

h) "АККУМУЛ. I" и "II" (BATT I, II) to "ВКЛ." (ON (up))

[LCtrl + LShift

+ 1..2]

;

i) DC selector knob to "ШИНЫ АКК." (BATT BUS) [LCtrl + LShift + 9..0]

B. Connecting to external electrical power:

To connect to an external power source on the ground:

 DC EXTERNAL POWER: once the "АЭР. ПИТ. ВКЛЮЧЕНО" (EXT PWR ON)

light illuminates, then

[LCtrl + LShift + 7]

, check the DC ground power

source voltage by setting the DC selector knob to "АЭРОДРОМ. ПИТАН."

(EXT PWR). The voltage should be within the limits of 27 - 29 V.

 AC EXTERNAL POWER (in DCS this occurs simultaneously with the DC power

supply): once the "АЭР. ПИТ. ВКЛЮЧЕНО" (EXT PWR ON) light

illuminates (after successful connection to an external power source), then

[LAlt + LShift + `]

, check the ground power source voltage by setting the

AC selector knob to "АЭРОДРОМ. ПИТАН." (EXT PWR). The voltage

should be within the limits of 200 - 205 V. Set the "АЭРОДРОМ. ПИТАН."

(EXT PWR) switch to "ВКЛ." (ON). Set the "ПО-500А ~ 115" (Inverter 1)

and "ПТ-200 ~ 36" (Inverter 2) switches to the "АВТОМАТ" (AUTO)

(down) postion. Set the "ВЫПРЯМИТЕЛИ I, II, III" (RECTIFIERS 1, 2, 3)

to the "ВКЛ." (ON) (up) position. Check the rectifier bus voltage by setting

the DC selector knob to "ШИНЫ ВЫПР." (RECT BUSES). The voltage

should be within the limits of 27 - 29 V.

o set the AC selector knob to the "~115" position. The voltage should be

115 V.

o Set the "АЭРОДР. ПИТАН." (EXT PWR) switch to "ВКЛ." (ON) and the

Inverter 115 switch to "РУЧНОЕ" (MAN).

o Check the inverter output voltage by setting the AC selector knob to

"~115". The voltage should be 115 V.

Equipment preparation and check procedures, continued

 switch ON all the circuit-breakers and switches required for starting the

APU and main engines (the starting system, ignition systems for the APU

and main engines, fire protection system, hydraulic systems, trim

actuators, fuel tank pumps, fuel quantity gauge, engine anti-icing system,

friction clutch, electric clutch, gyro correction cutout switch, attitude

indicator, directional gyro, autopilot, voice warning system, tail rotor pitch

limit system, cockpit voice recorder, anti-collision light

[RCTRL + RSHIFT

+ 1... 9]

 make sure the AC generator switches are set to "ВЫКЛЮЧЕНО" (OFF)

(down)

[LAlt + LShift + 1..2]

 make sure that РИ-65 (RI-65) switch is in the ВЫКЛ. (OFF) position and

the corresponding light panel ВКЛЮЧИ РИ-65 (ENABLE RI-65) is on;

 make sure that the СПУУ-52 (SPUU-52) switch is in the ВЫКЛ. (OFF)

position and the corresponding red light-button on the center console is

on;

 check that helicopter has enough fuel, using fuel meter, after check

procedure, set the switch to the "РАСХ." (FEED TANK) position (player

must take the co-pilot seat) [RCtrl + RShift + V]

rotation clockwise,

[RCtrl + RShift + B]

rotation counterclockwise;

 check that squibs of the fire extinguishers and fire signalization system are

operational according to the procedure described in the chapter 7.6.3

For night time operations:

 flashlight

[LAlt + L]

(mouse control)

 turn on red lighting of the instrument and control panels according 7.8.2

 turn on the navigation lights

[RCtrl + 1..2]

(dimmed), the MSL-3 flasher

[RCtrl + 6]

and rotor tip lights

[RCtrl + 5]

(they must be turned on for

safety of the ground crew, in game it is not necessary).

9.3. Starting the APU and main engines

After preflight cockpit check and energizing electrical power sources, is required

preparing for APU and main engines start:

246

 "РЕЗЕРВН. ГЕНЕРАТ." (STBY GEN) switch

is set to "ВЫКЛ." (OFF) (down)

[LCtrl +

LShift + 3]

The APU must run for a minimum of 1 minute before attempting to start the main

engines.

In case of an inadvertent shutdown of the APU, press the "ВЫКЛЮЧЕНИЕ АИ-9В"

(APU OFF) button for 2 to 3 seconds in order to cut off fuel supply to the APU

[End]

The APU start can be aborted at any time by pressing the "ВЫКЛЮЧЕНИЕ АИ-9В"

(APU OFF) button for 2 to 3 seconds.

In case of an unsuccessful APU start, crank the APU as follows:

 set the "ЗАПУСК - ПРОКРУТ. - ЛОЖНЫЙ ЗАПУСК" (START-CRANK-

FALSE START) selector switch to (CRANK) [RCtrl + E]

[RAlt + E]

;

 press the "ЗАПУСК" (START) button

[RShift + Home]

and check that the

"АВТОМАТ. ВКЛЮЧЕН" (AUTO IGNITION) and "ДАВ. МАСЛА НОРМА"

(OIL PRESS NORM) lights illuminate.

Restart attempts must be 3 minutes apart. Three attempts can be made. If the unit

does not start after three attempts, a 15 minute shut-down/cooling period must

follow before another start is attempted.

Continuous APU operation is limited to

30 minutes. In "РЕЗЕРВН. ГЕНЕРАТ"

(STANDBY/GEN) mode, the APU cannot be operated beyond 30 minutes, after which

a 15 minute shut down/cooling period is required. Cool down the APU 15 minutes

between shutdown and restart. Run the APU a minimum of 1 minute before

shutdown.

Three consecutive attempts to start the main engines via APU bleed air are allowed.

The duration of each air bleed cycle should not exceed 45 seconds with intervals

between the air bleed cycles no less than 1 minute, during which the APU is run at

250

5. If engine startup is performed with use of

onboard batteries only, then after the first

engine has been started, connect onboard

power consumers to the APU generator by

setting the "РЕЗЕРВН. ГЕНЕРАТ." (STANDBY

GENERATOR)

[LCtrl + LShift + 3] и "ПРОВЕРКА

ОБОРУД.") (EQUIPMENT TEST)

[LCtrl + LShift +

into the "ВКЛ." (ON) positions.

After completion of the automatic starting cycle the lights should turn off (AUTO

IGNITION ON light in 30 seconds; STARTER ON light upon N1 RPM reaching 60-

65%).

Unusual thumps or impact noises during main engine start and rotor spin up indicate

the main rotor blades centrifugal droop limiters are hitting their stops. Carefully

adjust the cyclic control stick position until the noise is eliminated.

Set the "ЛЕВ. - ПРАВ." (LEFT - RIGHT) engine selector switch to the second starting

engine (upwind side)

[RAlt + RShift + E]

[RCtrl + RShift + E] and repeat the

starting procedure for the second engine.

With both engines started and running at idle speed, the Nr (main rotor) RPM should

stabilize within 55-70%.

8. If engine startup is performed from batteries

only, then do not turn of the AI-9V until engines

have been warmed up and main rotor has

reached RPM of 88%. To supply onboard

consumers during the idle mode and when the

first engine has been started and second has

not, enable the STG-3 (see phase 5.) generator

by switching the "РЕЗЕРВН. ГЕНЕРАТ."

(STANDBY GENERATOR)

[LCtrl + LShift + 3]

"ПРОВЕРКА ОБОРУД.") (EQUIPMENT

TEST)

[LCtrl + LShift + 8]

to the "ВКЛ." (ON)

positions. Note. If before APU start the transfer

pumps were off, to safe batteries power, now it

is time to enable them.

9.4. Engines warm up, Flight controls and Hydraulic systems

checks

Engine warm up is performed at idle power with collective lowered to minimum

[Num - ]

, throttle turned fully left

[PgDn]

, engine condition levers set in the idle

detent (middle) position

[RCtrl + Home / RCtrl + End]

Monitor the powerplant instrumentation during engine warm up. Warm up should not

exceed 1 minute.

Test the flight controls and hydraulic systems at idle power as follows:

 *alternatively move the cyclic and pedals to make sure the controls move

smoothly without jamming;

 *as the controls are moved, the main hydraulic system pressure should

vary within a range of 45±3 to 65±⁸²kg/cm². Pressure in the reserve

hydraulic system should be approximately 5 kg/cm².

The throttle can be set from full left to full right to accelerate the engines out of idle

power once the engine outlet oil temperature reaches +30°C and the main gearbox

oil temperature reaches at least -15°C.

Additional checks at idle (for in-depth study of the helicopter):

1. Check the utility and standby hydraulic systems;

2. Check the engine anti-icing system at idle;

3. Check the electronic engine governor, using channels ST1 and ST2:

 switch on the "ЭРД ЛЕВ. ПРАВ" (ELECTRONIC ENGINE GOVERNOR LEFT RIGHT);

 set the "КОНТРОЛЬ СТ-1 - РАБОТА - КОНТРОЛЬ СТ-2" (CHECK ST1 - OPERATION -

CHECK ST2) into the "СТ-1" (ST1) positiion;

 gradually increase the main rotor RPM by twisting the throttle grip to the right (main

rotor pitch should be in the lower stop position), if nesessary, use the ECL handles, until

the yellow warning panels "ПРЕВ. Nст ЛЕВ. ДВ" (EXCEEDING RPM LEFT ENGINE),

"ПРЕВ. Nст ПРАВ (EXCEEDING RPM RIGHT ENGINE) go on. It should happen when

Nr _{(main rotor)}= 91.5±2%;

 gradually reduce Nr at 5÷7%, but not less than 88%, warning panels, mentioned above,

should remain on;

 set the switch to the "СТ-2" (ST-2) position and repeat the check.

WARNING: When switching from position "СТ-1" (ST-1) to "СТ-2" (ST-2) one must hold

them in intermediate state for some time. Fast toggling from "СТ-1" (ST-1) to "СТ-2"

(ST-2) can cause engines shut off (not implemented in the game).

9.5. Engine run up, switching ON generators and rectifiers.

Avionics checks

А. Switching ON generators and rectifiers

1. Rotate the throttle full right

[PgUp]

2. Set the AC generators 1 and 2

[LAlt + LShift

+ 1..2] and rectifiers switches (1..3) to ON

[LCtrl + LShift + 4..6]

256

6. Set (or check) the ПО-500 ~115 V (Inverter

[LAlt + LShift + 3] and ПТ-500 ~36 V

(Inverter 2)

[LAlt + LShift + 5]

inverters to the

АВТОМАТ (AUTO (down)) position.

Additional checks

at with fully opened (all the way to the right) throttle (for in-depth study of the helicopter):

1. Check the anti-icing system of the engines at right throttle. After 25-40 seconds from the moment

when the system was enabled, can increase: the gas temperature in front of compressor’s rotor (not

more than 60⁰C) and N_tc (Turbine compressor RPM) (not more than 2%).

2. Check engines response:

 rotate the throttle twister all the way to the right (clockwise) and operate in this mode

for 1 min;

 remember RPM of the turbine compressor;

 rotate twister all the way to the left (counterclockwise);

 rotate throttle twister all the way to the right within 1-2 seconds and measure time from

the beginning of rotation to the moment when the RPM of turbine compressor reaches

RPM, which is 1-1.5% less than one, observed on the fully opened throttle;

 engines response should be within 3-6 seconds range.

3. Transition to the generator mode: 1^st generator -"вкл" (on), using the rotating switch, measure the

interphase voltage, 2^nd generator -"вкл" (on), measure the interphase voltage, set the rotating switch

to the "115 в" (115 V) position.

4. Enable and check rectifiers (ВУ) - "вкл" (ON), enabling them one by one and checking current

using ampermeter;

5. Check the system adjusting main rotor RPM:

 set the collective to 3^o according to the main rotor pitch angle indicator, when throttle is

twisted all the way to the right, oil temperature in the main driveshaft must be not less

than 30⁰ C;

 deflect the N2 trim INCR-DEC switch downward and after main rotor RPM has settled,

check the value of main rotor RPM, it should be within 91±2%;

 deflect the switch upward and make sure that the main rotor RPM is within 96..99%;

 if the upper limit of the main rotor RPM has not been reached, warm up oil in the main

driveshaft to 40-60°С and repeat the check;

 after checking the main rotor RPM range, by deflecting the N2 trim INCR-DEC switch, set

a main rotor RPM of 94..95% and lower the collective all the way down.

5. Check anti-icing system of main rotor, tail rotor, windshileds and dust protection devices (if

necessary) according to 7.4.6

7. "Обогрев ПВД" (Pitot tube heating) - check (if necessary);

8. SPUU-52- check:

 enable the circuit breaker on the right circuit breaker console, enable the "СПУУ-52"

(SPUU-52) on the left triangular panel;

 return pedals to the neutral position;

 press the "откл" button (lamp goes on), set the indicator to the middle position;

 by keeping the the "откл" button pressed, set the spring loaded switch to the “t”

position, pointer on the indicator deflects to the right mark, then to the “P” position -

pointer deflects to the left mark;

 release the button and the switch - lamp button goes off, and pointer on the indicator

returns to the middle position;

 press the "откл" button and while keeping it pressed, by rotating the "Контроль"

(CHECK) knob, deflect the pointer all the way to the right;

 disable the the "СПУУ-52" (SPUU-52) and release the lamp-button, pointer on the

indicator should deflect all the way to the left, while lamp-button continues being on;

 enable the the "СПУУ-52" (SPUU-52), press the lamp-button and while keeping it

pressed, by the "Контроль" (CHECK) knob set the pointer to the middle position;

 during flight the zero pointer on the indicator can deflect itself to the left, when ambient

temperature increases or air pressure reduces.

B. Avionics check

Switch on all the flight, navigation, radio communication and electronic equipment

required for the flight, and test for proper functioning.

7. Prior to powering up the АГБ-3К (AGB-3K)

attitude indicator, cage the device by pressing

the "АРРЕТИР" (CAGE) button: pressing the

"АРРЕТИР" (CAGE) button

[LCtrl + LShift + N]

(

[RCtrl + RShift + N] for right) and switch on

[LAlt + LShift + G]

(

[RAlt + RCtrl + P]

for right)

and test the attitude indicator

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11*. When directional gyro heading arrow

settles on the starting ground course, set the

ГМК-1А (GMK-1A) gyromagnetic compass to

ГПК (directional gyro) mode

[RCtrl + RShift + O]

12*. Test the autopilot by pressing:

 yaw channel - the "НАПРАВЛЕНИЕ"

(YAW)

[ LCtrl + A]

 roll-pitch chanel - "КРЕН-ТАНГАЖ"

(ROLL- PITCH) -

[LWin + A]

 altitude chanel - "ВЫСОТА"

(ALTITUDE)

[LAlt + A]

button-lamps on the autopilot control panel.

13*. With feet off the pedals, press the "3K"

momentary switch on the PU-26 control panel

for a short time to the left or right [not

assigned]. The YAW channel scale on the

autopilot control panel should rotate in response

to the manual heading change input.

Another translate var

13*. Deflect the spring-loaded "ЗК" switch to

the left for or to the right momentary. Due to

that the "НАПРАВЛЕНИЕ" (DIRECTION) scale

should begin rotating to the corresponding

direction.

262

11*. Set the "САРПП-12Д1М "РУЧН - АВТОМ"

(FLIGHT RECORDER) switch to the "АВТОМ"

(AUTO) position (down)

[LAlt + LCtrl + LShift + 6]

9.7. Preparing for taxi and taxiing

А. Prior to taxi, perform a taxi check:

 external power cables are disconnected:

[\]

(Radio Menu),

[F8]

[F2]

;

 fuel pumps are on (check switch positions and indicator lights);

 APU is shut down;

 all circuit breakers are switched ON.

B. Check the taxiway is clear of obstructions and proceed to:

 set throttle full right;

 switch ON PZU particle separators;

 check main rotor RPM to be within 95±2%;

 request clearance to taxi;

 release the wheel brakes [W]

Increase collective to set 1-2° of collective pitch and slightly push forward on the

cyclic to begin forward movement.

C. Maintain weight on wheels during taxi.

If the ground surface prevents safe taxiing, perform hover taxi using low speed/low

altitude flight.

D. Taxi speed should not exceed 15 - 20 km/h. Perform taxi turns using smooth

pedal input and small moving the Cyclic to side of turn. Avoid completely unloading

weight off the nose wheel shock strut.

E. Wind speed during taxi must not exceed 15 m/s. In crosswind conditions, the

helicopter tends to turn into the wind. Correct any uncommanded turning tendency

with slight opposite pedal and any uncommanded roll with slight opposite cyclic.

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F. Upon reaching the takeoff position, check the flight and navigation indicators,

ensure the attitude indicator is powered up (no warning flag present), the heading

indicator shows correct bearing to the selected ADF beacon, and the compass set is

slaved and indicating correct takeoff heading.

Start the flight timer on the AChS-1 cockpit clock by pressing the left red knob

[RALT + RCTRL + RSHIFT + C]

9.8. HOVER

А. The following maximum hover altitude limitations apply depending on helicopter

gross weight (GW):

 GW ≤ 11,100 kg: 10 m

 GW > 11,100 kg: 5 m

Hover altitudes greater than above limitations are allowed when carrying external

sling loads or if dictated by tactical requirements.

B. Yaw rate in hover must not exceed 12°/sec.

C. To perform a hover:

 position the helicopter into the wind if possible

 check throttle set full right

 smoothly increase collective to set collective pitch to 3°

 check main rotor RPM to be 95%. If necessary adjust main rotor RPM

using the N₂ TRIM INCR-DECR switch on the collective control handle

 turn off the autopilot (all channel)

[LWin + LShift + A] and then turn on

the autopilot "КРЕН-ТАНГАЖ" (ROLL-PITCH)

[LWin + A]

channel by

pressing the corresponding button-lamps and checking for green light

illumination

 continue to smoothly increase collective to lift the helicopter off the

ground and climb until reaching desired hover altitude

D. Increase of collective control during liftoff must be smooth and gradual, allowing

no less than 5 seconds for the engines to attain Takeoff power, ensuring main rotor

RPM is maintained within normal limits of 92 - 94%.

E. During liftoff, the helicopter tends to drift forward and left.

F. The cyclic control stick deflection in hover is approximately:



1/4 stick travel aft when helicopter CG is at normal to aft limit position;

1/2 stick travel aft when helicopter CG at the forward limit position



1/4 stick travel right regardless of CG position

9.9. Shifts and hops at low aititude

А. Shifts and hops at low aititude may be performed for training purposes, special

purpose operations, and in cases where the ground surface conditions do not allow

for safe ground taxi.

B. Lateral and reverse hover flight speed may not exceed 10 km/h. Use the ground

for visual reference and ensure that the flight path is clear of obstacles.

C. Forward hover flight altitude may not exceed 10 m and speed not exceed 20

km/h. Use the ground for visual reference and the stationary flight indicator of the

Doppler system for precise flight control.

With wind speeds of up to 10 m/s, shifts and hops at low aititude can be performed

into the wind or at 90° to the wind. With wind speeds greater than 10 m/s, hover

taxi should only be performed into the wind.

D. Perform low level flight over uneven terrain (gullies, ditches, drop-offs, etc.) at

altitudes of no less than 20 m and speeds no less than 60 km/h.

9.10. Takeoff

Takeoff is performed using one of the following procedures:

 vertical takeoff with acceleration in ground effect

 vertical takeoff with acceleration out of ground effect

 running takeoff

 running nose wheel takeoff

The minimum dimensions of the airfield required for a takeoff or landing at altitudes

of up to 1500 m are as follows:



50 x 50 m for a vertical takeoff/landing free of obstacles



50 x 120 m for a vertical takeoff/landing with an obstacle height of 15 m

at the airfield outer edge



50 x 160 m for a running takeoff/landing free of obstacles



50 x 200 m for a running takeoff/landing with an obstacle height of 15 m

at the airfield outer edge

Vertical takeoff with acceleration in ground effect

A vertical takeoff with acceleration in ground effect may be performed when the

helicopter hovers at an altitude of no less than 3 m with the engines set to Takeoff

power.

Fig. 9.1. Vertical takeoff with acceleration in ground effect

Position the helicopter into the wind (4) and perform a hover safety check to ensure

the helicopter is ready for takeoff (1). Confirm normal indicator readings and a

sufficient hover altitude for a vertical takeoff. Reduce altitude to 0.5-1 m (2) and

begin the takeoff by smoothly pushing the cyclic forward while simultaneously

270

advancing power as required up to Takeoff setting to avoid main rotor RPM drooping

below 92%. Accelerate in ground effect in a shallow climb to reach 60-70 kph at an

altitude of 20-30 m (3). Transition to a climbout attitude while accelerating to 120

kph.

Vertical takeoff with acceleration out of ground effect

A vertical takeoff with acceleration out of ground effect must be performed when the

takeoff area is confined and surrounded with obstacles, and the helicopter’s takeoff

weight allows for a hover out of ground effect.

Fig. 9.2. Vertical takeoff with acceleration out of ground effect

Position the helicopter into the wind (4). Perform a vertical takeoff while minimizing

drift (1) to an altitude of at least 10 m above obstacle height. In the vertical climb,

monitor the main rotor RPM to ensure it does not droop below 92%. Having attained

a hover altitude sufficient for a safe transition to forward flight above obstacle

height, smoothly push the cyclic forward to accelerate up to 20-50 kph

(2).

Transition to a climbout attitude while accelerating up to 120 kph (3).

Running takeoff

A running takeoff may be performed if the helicopter hovers at an altitude of no less

than 1 m with the engines set to Takeoff power. For a running takeoff, only the

ROLL-PITCH channel of the autopilot should be engaged.

Fig. 9.3. Running takeoff

Perform a hover check then land the helicopter (1). Reduce collective until the

helicopter rests on the ground with weight on wheels.

Push the cyclic forward and simultaneously increase collective to establish forward

acceleration up to 20-50 kph (2). Increase collective further to attain Takeoff power

and lift the helicopter off the ground (3).

In the takeoff run, the helicopter tends to lift off the main wheels first, followed by

the nose wheel. Compensate for this tendency with slight pull aft of the cyclic at the

moment of liftoff.

After liftoff, continue to accelerate up to 120 kph in a shallow climb, followed by a

transition to a climbout attitude (4).

The takeoff run requires 250 - 300 m. If the takeoff area is limited or blocked by

obstacles, the transition to climbout can be made at 50 - 60 kph.

Running nose wheel takeoff

A running nose wheel takeoff may be performed to increase takeoff performance

with a high takeoff weight or to reduce the distance of the takeoff run on airfields

that provide for a safe ground run.

Fig. 9.4. Running nose gear takeoff

Perform a hover safety check then land the helicopter (1).

Disable the autopilot by pressing the AUTOPILOT OFF button on the cyclic control

stick. Reduce collective pitch to minimum by lowering the collective control handle to

the lower stop (full down). Apply the wheel brakes. Push the cyclic control stick to

the forward limit and proceed to turn ON the autopilot ROLL-PITCH channel by

pressing the corresponding button-lamp on the autopilot control panel and checking

for green light illumination. Pull the cyclic control stick to the aft limit and release the

stick force by pressing the TRIM button on cyclic control stick. The forward and aft

deflection limits of the cyclic are determined by the absence of thumping noise as the

main rotor blades strike against the flapping hinges of the rotor assembly.

Release the wheel brakes. Smoothly increase collective until the main landing gear

begins to lift off the ground. Maintain the nose gear on the ground. Smoothly (3 - 5

sec) push the cyclic forward 1/2 to 2/3 stick travel to begin the ground run and

establish a nose-down pitch attitude of -8 to -9° below standing pitch (2). Control

the pitch angle during the ground run by maintaining the blade tips of the rotor disc

on the horizon line. Smoothly increase engine power to Takeoff setting in the ground

run.

272

At approximately 40 kph the helicopter exhibits a tendency to pitch up and sink onto

the main gear, followed by a reversal and an energetic pitch down. These tendencies

must be countered with corresponding forward and aft cyclic control adjustments.

Lift off 1 - 2 seconds after the pitch down with a smooth cyclic pull aft as ground

speed reaches 60 - 65 kph (3). Continue to accelerate to 70 - 80 kph up to an

altitude of 10 m. Proceed with climbout at a speed of 120 kph (4).

With a maximum takeoff weight of 1300 kg, a ground run of 150 m is required for a

paved runway or 340 m for a field airstrip. If the helicopter’s CG is close to the aft

limit, the ground run distance increases by a factor of 1.5.

After leveling off at the desired altitude in stabilized in level flight, switch OFF the

autopilot by pressing the AUTOPILOT OFF button on the cyclic control stick, stabilize

the controls and switch ON the autopilot ROLL-PITCH and YAW channels by pressing

the corresponding button-lamps on the autopilot control panel and checking for

green light illumination.

9.11. Climb to altitude

А. The optimal climb speed is 120 kph in altitudes up to 2000 m; 110 kph in altitudes

of 2000 - 4000 m; and 100 kph above 4000 m. Climbs are normally performed in

maximum continuous engine power. If required, the climb may be performed in

Takeoff power (limited to 6 minutes) as well as power settings below maximum

continuous.

After completing the takeoff, establish the desired climb rate and switch OFF the PZU

particle separators.

Fig. 9.5. Engine pressure ratio (EPR) indicator

B. Current engine power setting is monitored on the engine pressure ratio (EPR)

indicator up to an altitude of 2500 m and is determined by the position of the side

indices with respect to the "Н" and "К" markers:

 Takeoff power: side indices above "H"

 Maximum continuous power: side indices above "K" up to "H"

 Cruise power: side indices aligned with or below "K"

C. At altitudes greater than 2500 m, engine power setting is determined based on

corresponding performance charts.

D. In a climb at maximum continuous power with a constant collective pitch angle,

the main rotor RPM is automatically maintained at 95±2% up to a limited altitude.

Further climb will result in the main rotor RPM drooping as engine power output is

reduced due to compressor RPM limits imposed by the engine governor system.

Maintain main rotor RPM above 92% by gradually reducing collective pitch as main

rotor RPM begins to droop. The maximum continuous power limitations begin to

affect main rotor RPM at 1000 - 1500 m.

In a climb at cruise power with a constant collective pitch angle, the main rotor RPM

is automatically maintained constant up to an altitude of 2000 - 2500 m.

In a climb at takeoff power with constant collective pitch angle, the main rotor RPM

is not maintained automatically. Maintain main rotor RPM in the 92-94% range by

gradually reducing collective pitch as altitude increases.

9.12. Level flight

The recommended airspeed for flying an airfield pattern is 160 kph.

Roll angles are limited to 30° at normal takeoff weight and 20° at maximum takeoff

weight.

9.13. Transitional maneuvers

А. To transition from a vertical climb to a hover after reaching the desired altitude,

stop the climb by smoothly reducing collective and maintain altitude with slight

collective adjustments.

B. To transition from a hover to a vertical descent, reduce collective and ensure the

descent rate does not exceed 0.2 m/s near the ground prior to touchdown.

C. To transition from a hover to level flight, push the cyclic forward to establish an

accelerating attitude. Simultaneously adjust (slightly increase) collective to maintain

altitude and counter any lateral drift and yaw with opposite cyclic and pedal input.

When approaching the desired airspeed, pull the cyclic aft to stabilize in a level flight

attitude and maintain airspeed.

D. To transition from level flight to a hover while maintaining altitude, smoothly

reduce collective and pull the cyclic aft to reduce airspeed. Upon reaching airspeed of

50 - 60 kph, the helicopter exhibits a tendency to descend. Counter this tendency by

increasing collective.

At airspeeds below

50 kph, the helicopter develops

vibrations, which disappear as airspeed is reduced further. At airspeeds below 40 -

20 kph, the helicopter exhibits a tendency to yaw left. Timely application of cyclic

and right pedal input is required to avoid uncommanded roll and left yaw in the

transition to hover.

E. To transition from level flight to a power-on glide, reduce collective and apply

cyclic to establish a desired glide speed attitude.

274

F. To transition from a power-on glide to level flight, apply collective to set engine

power as required for level flight and apply cyclic to establish a desired airspeed.

G. In transitional maneuvering, the main rotor RPM is automatically maintained at

95± 2% only within a limited rate of collective application:

 when increasing collective, no less than 5 seconds from 1 - 3° collective

pitch up to the pitch angle establishing takeoff power.

 when reducing collective, no more than 1°/sec from any starting collective

pitch angle

Collective input rates above these limits can lead to main rotor RPM drooping below

the minimum allowable limit (88%) when increasing collective or overspeed the main

rotor above the maximum allowable limit (103%) when reducing collective.

If main rotor RPM runs outside 95±2%, adjust collective to return RPM to the normal

range.

Large deflections of the cyclic can lead to main rotor RPM drooping in accelerations

and increasing RPM in decelerations. The range of main rotor RPM divergence is

proportional to the rate of cyclic deflection.

When performing transitional maneuvers, unload the forces on the controls with

short presses of the TRIM button on the cyclic control handle as the flight controls

are adjusted.

9.14. Descent

Depending on altitude, power-on descent may be performed either vertically or on a

glideslope. Autorotation may only be performed on a glideslope descent.

Power-on vertical descent

A power-on vertical descent from am altitude of 10 m down to the ground is

permissible in all conditions. From an altitude of 110 m down to 10 m, a vertical

descent is only permissible when a glideslope approach cannot be performed due to

obstacles or out of tactical considerations. Descent from the helicopter’s service

ceiling down to 110 m must be performed on a glideslope and within airspeed

limitations.

In a vertical descent from an altitude of 110 m down to 10 m, the descent rate may

not exceed 3 m/s. If the descent rate increases beyond 3 m/s, smoothly increase

collective to arrest the descent rate. If engine power is insufficient to arrest the

descent and maintain main rotor RPM within limits, transition out of the vertical

descent to a glideslope descend or forward flight to gain airspeed.

From an altitude of 10 m down to the ground, continually reduce the rate of descent

so it does not exceed 0.2 m/s at touchdown.

Power-on gliding descent

In a power-on glideslope descent, maintain main rotor RPM within 95±2% with

collective input as required. Gradual reduction of collective pitch down to the

minimum setting is permissible to maintain a desired descent rate as altitude

decreases as long as main rotor RPM is maintained within limits.

The recommended glide speed at altitudes below 2000 m is 120 - 180 kph. The rate

of descent at this speed should be 3 - 5 m/s.

9.15. Autorotation descent

А. An autorotation descent is used in case of dual engine failure in flight. To perform

an autorotation landing:

 establish a desired descent airspeed prior to initiating the descent

 reduce collective down to the lower stop (full down) and check the main

rotor RPM to be within normal limits (95±2%)

 counter the any tendency to yaw right and pitch down with opposite pedal

and cyclic

 set throttle to full left

 upon transitioning to an autorotation descent, adjust collective as required

to maintain main rotor RPM within limits

B. Maintain the following airspeeds in a power-on autorotation descent:

 altitude 2000 m and greater: 100 - 120 kph

 altitude below 2000 m: 120 - 190 kph

The optimum gliding speed for altitudes below 2000 m is 180-190 kph.

C. Maintain a descent rate of 10 - 12 m/s. The minimum rate of descent of 10 m/s

corresponds with a gliding speed of 120 kph.

D. Avoid roll angles of greater than 20° during autorotation descents.

E. To recover from a power-on autorotation descent:

 smoothly set throttle to full right while monitoring the engine and main

rotor RPM

 at altitudes above 1500 m, counter main rotor overspeeding beyond

maximum limits by increasing collective to set a collective pitch angle of 3

- 4°

 at altitudes below 1500 m, increase collective pitch only after setting the

throttle to full right. Avoid drooping the main rotor RPM below 92% by

raising collective gradually

F. Maintain 100 - 120 kph in a power-off autorotation descent

9.16. Landing

Landing is performed using one of the following procedures:

 vertical landing from a hover in ground effect

 vertical landing from a hover out of ground effect

 power-on running landing

 single engine landing

 power-off autorotation landing (only in emergency situations)

276

When performing any landing with forward airspeed, including autorotation landing,

disengage the autopilot YAW and ALTITUDE channels.

Vertical landing from a hover in ground effect

Perform a glideslope approach at 120 kph. At an altitude of 100 m, smoothly pull

the cyclic aft to begin reducing forward airspeed to attain 60 - 50 kph at an

altitude of 60 - 50 m.

At an altitude of 5 - 8 m, smoothly pull the cyclic further aft and increase

collective as required to establish a hover at a altitude of 2 - 3 m.

The minimum glideslope approach length:

 with GW ≤ 11,100 kg is 1000 -1200 m at altitude of 100 m and

airspeed 100-120 kph

 with GW > 11,100 kg is 1400 -1500 m at altitude of 100 m and

airspeed 100-120 kph

While performing the deceleration and transitioning to a hover, release the forces

on the controls with frequent presses of the TRIM button on the cyclic control

stick.

Upon stabilizing in a hover, smoothly decrease collective to perform a vertical

descent while gradually reducing vertical speed such that it does not exceed 0.2

m/s at touchdown.

Avoid lateral drifting while in the vertical descent. Reduce collective to minimum

only when certain that the helicopter is firmly on the ground with weight on

wheels.

In a crosswind landing, apply cyclic opposite of the wind direction to maintain

position over the landing point until the helicopter is firmly on the ground with

weight on wheels.

Vertical landing from a hover out of ground effect

Perform a vertical landing from a hover out of ground effect only when obstacles

make it impossible to perform a landing from a hover in ground effect.

The procedure for a vertical landing from a hover out of ground effect is identical

to a vertical landing from a hover in ground effect.

Begin the deceleration at an altitude of 50 m above obstacle height such that a

hover position over the landing point is attained at an altitude of no less than 5

m above obstacle height.

Power-on running landing

A power-on running landing may be performed in cases where engine power is

insufficient to ensure a hover and vertical landing (such as high gross weight, high

altitude, high temperatures).

The landing can be performed on a prepared runway or a an unprepared area known

to be safe for such an approach (must be sufficiently level and large) provided a

clear approach path is available.

Execute the final approach on a glideslope with an airspeed of 120 kph.

Maintain the glideslope such that airspeed is maintained 20 kph faster than current

altitude, i.e. 100 kph at 80 m down to an altitude of 40 m.

Perform the remainder of the descent with a continual reduction of airspeed and rate

of descent such that at 0.5 - 1 m above the ground, the airspeed is reduced to 50 -

40 kph and the rate of descent is reduced to 0.1 - 0.2 m/s.

Perform a smooth touchdown on the main gear and reduce collective to minimum.

Allow the nose gear to touch down. Set throttle to full left and apply the wheel

brakes to brake the helicopter. Anticipate a landing run of 20 - 30 m. The total field

distance for safe operations should be no less than 100 m.

If the airfield dimensions do not allow for a landing run of 20 - 30 m, but it is

necessary to perform a running landing, execute a running landing with a short

landing run.

Begin a smooth reduction of forward airspeed and rate of descent at an altitude of

40 - 50 m above the field by increasing collective and pitching the helicopter up with

aft cyclic while maintaining the main rotor RPM within allowed limits. Perform a

landing deceleration maneuver so as to attain near takeoff engine power at an

altitude of 5 - 10 m with a ground speed of 20 - 40 kph. At an altitude of 5 - 10 m,

push the cyclic forward to

bring the helicopter to a landing attitude while

avoiding a tail boom strike against the ground, but ensuring continued reduction of

ground speed down to 10 - 15 kph for touchdown. At an altitude of 5 - 10 m, raise

collective at a rate of 2 - 4°/sec to reduce the rate of descent such that it is no

greater than 0.2°/sec at touchdown. Upon touchdown, push the cyclic forward 1/3 -

1/4 stick travel, reduce collective down to minimum, set throttle to full left, and apply

the wheel brakes to brake the helicopter.

Single engine landing

Perform a single engine landing onto a flat landing area that provides for a clear

approach or onto a prepared runway. The gross weight of the helicopter for a single

engine landing may not exceed 10,000 kg.

Perform a single engine landing into the wind if possible or with a crosswind not

exceeding 5 m/s.

At an altitude of 300 m, before starting the APU, switch OFF the engine anti-icing

system and PZU particle separators if these were previously switched on. Start the

APU. Confirm successful APU start and illumination of the OIL PRESS NORMAL

(ДАВЛ. МАСЛА НОРМА), NORMAL SPEED (ОБОРОТЫ НОРМА) lights.

Execute a single engine landing so that the helicopter touches down at 10 - 20 kph

or 50 kph (as decided by the pilot in command) as follows:

 control collective pitch to maintain main rotor RPM within 95±2%

278

 ensure the operating engine attains emergency power setting

 maintain an airspeed of 20 kph higher than current altitude (in meters)

during the approach

 establish a landing attitude at an altitude of 5- 7 m

 from an altitude of 3 - 5 m, reduce the rate of descent by increasing

collective. Simultaneously apply smooth right pedal to counter torque-

induced yaw from increased collective pitch and use cyclic control to

maintain the landing attitude. When increasing collective, ensure that

main rotor RPM does not fall under 88%

 upon touchdown immediately decrease collective smoothly and push the

cyclic forward 1/3 - 1/4 stick travel to prevent a tail boom strike

 apply wheel brakes after nose gear touchdown to brake the helicopter

Using this procedure the helicopter touches down at a speed of 10 - 20 kph and the

landing run is 5 - 20 m. To touch down at a speed of 50 kph, perform the approach

such that the airspeed is maintained 20 kph higher than current altitude down to an

altitude of 40 m, then maintain 60 km/h down to 5 - 7 m. Perform the touchdown as

described above, which results in a landing run of 80 - 100 m due to the higher

landing speed.

9.17. Search and Rescue (SAR) operations

А. Prior to departure for a SAR mission:

 turn on the VHF ADF (РАДИОКОМПАС УКВ) circuit breaker on the right

circuit breaker panel of the overhead console.

 on the VHF ADF (АРК - УД) control panel, set the mode selector switch to

NS (ШП), the frequency selector switch to VHF (УКВ), and the CHANNELS

(КАНАЛЫ) selector switch to 4.

 on the intercom control box, set the selector switch to PK2 and the INT -

RADIO (СПУ-РАДИО) selector switch to RADIO (РАДИО)

 fly the helicopter to enter the search area bearing in mind that the VHF

ADF (АРК - УД) detection and homing ranges increase with altitude (at an

altitude of 500 m the coverage is no less than 25 km)

 with the ADF operating in standby reception mode, positive reception of a

beacon signal will be indicated by the corresponding indication light.

B. After detection and identification of the beacon signal, determine its location as

follows:

 set the mode selector switch to a position corresponding to the indication

light: narrow band NB (УП) or broadband BB (ШП) if the NB light is on,

set the mode selector switch to NB

 test the pointer arrow by pressing the ANT L (AHT. Л) or R (АНТ. П)

buttons to manually turn the arrow and make sure that it returns to the

signal bearing when the buttons are released

 turn the helicopter so that the pointer arrow points to "0" and continue to

fly the helicopter to maintain the pointer arrow in this position. At long

ranges to the beacon, begin homing in narrow band NB (УП) mode. As

signal strength increases (indicated by increasing volume in the headset),

select the broadband BB (ШП) mode. The VHF ADF operation is more

reliable in broadband mode.

9.18. Flight (hover) over featureless terrain using the Doppler

Navigation System

А. The stationary flight indicator provides visual indication of the ground speed in the

following speed ranges: 0 - 50 kph in forward flight, 0 -25 kph in reverse flight, 0 -

25 kph in lateral flight.

B. The forward and lateral speeds are indicated by corresponding moving indexes

against a numerical scale (up and down for forward and reverse flight, left and right

for lateral flight). Vertical velocity is indicated by a moving triangular index along the

left side of the indicator.

Before takeoff, turn on the Doppler system by setting the DOPPLER (ДИСС) circuit

breaker on the overhead console and the DOPPLER (ДИСС) switch on the right

switch panel of the overhead console to ON (BKЛ).

C. When hovering, observe the indicators of the Doppler system. Apply cyclic control

opposite of the movement of the line indicators to maintain a hover position by

keeping the indicators inside the center ring position. Maintain the vertical velocity

indicator at “0” to maintain altitude.

In limited visibility conditions when the natural horizon cannot be seen, control the

helicopter’s attitude using the Attitude Indicator and other flight instruments. Control

hover altitude using the radar altimeter. The radar altimeter provides accurate

altitude indication up to 1000 m above ground level (AGL). If airspeed reaches 50

kph, the stationary flight indicator will turn off and the OFF (ВЫКЛ.) light will

illuminate.

9.19. Night operations in visual meteorological conditions

(VMC)

А. The start, ground test, and shutdown procedures for nighttime operations are

identical to daytime operations except the following: additionally turn on the LAND

LIGHTS (ФАРЫ), NAV LTS (АНО), FORM LIGHTS (СТРОЕВ. ОГНИ) and CHK BLINKER

(ПРОВЕРК. ЛАМП-МИГАЛКА) circuit breakers on the overhead console and set the

DOME LT RED - WHITE (ПЛАФОН КРАСНЫЙ - БЕЛЫЙ) selector switches to WHITE

(БЕЛЫЙ) on the left and right switch panels of the overhead console. Turn down the

red lighting rheostats on the left and right side panels of the overhead console and

the flight compartment doorway. Turn on ФР-100 taxi light.

B. After starting the engines and disconnecting the external electrical power source,

switch off the white dome lights, set the DAY - NIGHT (ДЕНЬ - НОЧЬ) selector

switch to NIGHT (НОЧЬ), and turn on the BLINKER

(МИГАЛКА),

ANTI-COLL

LIGHT (ПРОБЛЕСК) and BLADE TIP (КОНТУР. ОГНИ) switches. Set the navigation

and formation lights

switches to BRIGHT (ЯРКО) or DIM (ТУСКЛО) as desired.

280

Taxi with the ФР-100 taxi light on. Use the ФПП-7 landing/search lights only when

required for improved forward visibility or taxi turns. Limit operation of the ФПП-7

landing/search lights to 5 minutes followed by a 5 minute cooling period.

C. Take off with both the ФР-100 taxi and ФПП-7 landing/search lights turned on.

Adjust the direction of the light beams in a hover at an altitude of 3 - 5 m by

operating the corresponding switches on the collective control handle.

D. Accelerate and climb out to an altitude of 50 m more gradually than in daytime

operations. At an altitude of 30 - 50 m, transition to instrument flight and turn off the

taxi and landing/search lights.

E. For night time flying, refer primarily to the flight instruments with occasional

checking of the outside airspace.

F. Perform approach and landing maneuvers as during daytime operations. At an

altitude of 50 - 70 m, turn on the ФПП-7 landing/search lights. If the landing/search

lights make visual perception of the ground more difficult, turn the lights off and use

other light sources for ground reference, such as ground-based light projectors if

available. Use the radar altimeter to control altitude with visual ground checks using

available light source references.

After landing, taxi with the ФР-100 taxi light turned on.

9.20. Day or night operations in instrument meteorological

conditions (IMC)

А. Prior to embarking on any flight in IMC, carefully examine the weather conditions

in the area of operations, paying special attention to possible icing conditions, wind

speeds and directions. Flight inside clouds is permissible up to an altitude of 3500 m.

B. Before taxiing out, check that all circuit breakers and switches required for flight

are turned on and set correctly. Ensure normal operation of the autopilot control

channels, attitude indicators, turn indicators, ADFs, compass system, radar

altimeter, windshield wipers, and the Doppler system speed and drift

indication. Check that the clock is running and set for the correct time. The pressure

on the barometric altimeter should correspond to the actual airfield pressure when

the altimeter is set to 0 altitude. Check that the compass system is turned on and

operating normally, the setting of the latitude correction, selection of magnetic

compass (MK) mode, and the Doppler system is operating normally as indicated by

the FUNC (РАБОТА) light on the control panel.

C. For ambient temperatures of +5°C or below, turn on the pitot tube heaters before

taxiing out of the parking area and turn them off after taxiing to the parking area.

Before taxiing out of the parking area in ambient temperatures of +5°C or below,

turn on the engine anti-ice systems to prevent icing in the air intakes and ingestion

of ice into the engines by setting the ANTI-ICING SYSTEM. ENG DUST PR-LEFT and

ANTI-ICING SYSTEM. ENG DUST PR-RIGHT (ОБОГРЕВ. ДВИГ. ПЗУ ЛЕВ. (ПРАВ))

switches to ON (ВКЛ.).

D. Set the chart angle for the selected route on the coordinate indicator of the

Doppler system, set the range and angle error to 0, and turn off the control panel

using the OFF (ВЫКЛ.) button.

E. After taxiing to the takeoff position, slave the compass system, set heading

pointer of the heading indicator to the takeoff magnetic heading.

F. Request permission for takeoff from the controller. Upon receiving clearance,

proceed with takeoff. Maintain visual contact with the ground in hover.

G. After takeoff and before entering the cloud cover, establish a climbing profile at

an airspeed of 150 kph and a climb rate of 3 - 4 m/s. Transition to complete

instrument flight 25 - 30 m below cloud cover.

When flying in clouds, the following flight profiles are recommended:

 climb speed of 150 kph at climb rate of 3 - 4 m/s

 descent speed of 120 - 200 kph at rate of descent of 3 - 4 m/s

 horizontal flight speed of 160 - 180 kph for a standard approach

Prolonged route flying at altitudes of up to 1000 m at the following airspeed:



220 kph at normal takeoff weight



200 kph at maximum takeoff weight

9.21. Wide rectangle pattern

The wide rectangle pattern is a convenient landing pattern when the approach to the

locator middle marker (LMM) is made within 60° of the magnetic landing course.

Perform an IMC approach and landing by referencing the LMM, which is positioned at

a distance of 1300 m from the landing point. The recommended pattern altitude is

300 m, airspeed is 160 kph, roll angle in turns is 10°.

If flying the pattern after takeoff, establish a climbing profile for an airspeed of 150

kph and a climb rate of 3 - 4 m/s.

Perform the first turn to the crosswind leg at an altitude of no less than 150 m at

distance of 3500 m from the reference startling line (runway takeoff position) or

when the calculated flight time for the initial leg has expired (1 min 32 sec in calm

weather). Upon reaching an altitude of 300 m, establish level flight at 160 kph. In a

missed approach or practice approach without landing, perform the first turn to the

crosswind leg 2 minutes after passing over the LMM.

Perform the second turn to the downwind leg when the LMM relative bearing (RB

(bearing from current heading to marker) equals 240° ± drift angle (DA) (120° ± DA

for a right-hand pattern) or upon reaching the required magnetic radio bearing

(MRB, bearing to marker from due North and indicated on the heading indicator

compass card by the bearing pointer) 3 min 27 sec after takeoff time.

Perform the third turn to the base leg when RB = 240° ± DA (120° ± DA for a right-

hand pattern) or upon reaching the required MRB.

On the base leg, descent at a rate of 2 - 3 m/s and establish an airspeed of 155 kph.

Descend to an altitude of 200 m.

282

On the base leg nearing the fourth turn to final approach, the bearing needle of the

directional gyro will be moving toward the desired course needle (set to magnetic

runway heading). At the start of the turn, the angle between the desired course and

bearing needles should be 15°. When the turn is executed correctly, the two needles

will align approximately 30° prior to reaching final approach course.

Continue the final turn with the desired course and bearing needles aligned.

If during the first half of turn the angle between the bearing needle and the desired

course needle is constant or increasing, the angle of roll should be decreased. If

after the needles align the bearing needle starts falling behind the desired course

needle, the angle of roll should be increased, but by no more than 15°.

After recovering from the final turn, begin to descent at a rate of 2 - 3 m/s and

reduce airspeed to pass over the LMM at 100 - 140 kph at an altitude of 100 m. If an

altitude of 100 m is reached prior to passing over the locator middle marker,

transition to level flight.

If the final turn is recovered on a heading different from the landing course, perform

a course correction while on the descent by checking the course deviation angle

when the bearing needle is centered directly ahead to the LMM. If the difference

exceeds 5°, correct the heading error by turning toward the bearing needle (away

from the desired heading needle) such that the bearing needle is set midway

between the heading index on the compass card and the desired heading needle. If

the course indicator reads a magnetic heading exceeding the landing course, perform

the course correction to the right, otherwise perform the course correction to the

left.

After starting the course correction, maintain the corrective heading until the bearing

needle aligns with the desired heading needle, then turn the helicopter so as to align

the bearing needle and the desired heading needle on the landing course over the

compass card (or with an angle-off the landing course to account for drift).

In a right drift scenario, maintain the bearing and the desired heading needles

aligned along the landing course, but offset to the right along the compass card to

correspond to the drift angle.

Maintain the current landing course after passing over the LMM.

When flying the wide rectangle pattern for landing, maintain the landing course after

passing over the LMM and execute the first turn to the crosswind leg when the

calculated flight time has expired (2 min in calm weather).

9.22. Tight rectangle pattern

When the LMM is approached at an angle of greater than 60°, but less than 120° off

from the landing course, use the tight rectangle landing pattern.

284

Fig. 9.7. Tight rectangle pattern diagram

After crossing the LMM, turn to a heading perpendicular to the landing (runway)

heading (estimating the drift angle). After the estimated time has been reached (1

min 15 sec for calm weather conditions), turn to a heading opposite the landing

heading and estimate the drift angle. Afterwards, the pattern is to be completed like

the big pattern approach.

9.23. Straight in approach with teardrop procedure turn

If the LMM is approached on a magnetic heading opposite to the landing course

(MHLO) or if the difference does exceed 60°, perform a straight in approach with a

teardrop procedure turn.

The magnetic heading, flight time (HFT), and estimated turn angle (ETA) calculations

for performing the teardrop procedure turn are prepared in advance taking into

consideration approach altitude and anticipated drift angles based on wind

conditions. The resulting values are entered into a reference chart:

Procedure

Н (altitude), m

elements

300

400

500

600

700

800

900

1000

ETA, deg

HFT, min:sec

1:30

2:15

3:00

3:45

4:30

5:15

6:00

6:45

NOTE. chart assumes calm weather:

V_gs= 160 км/ч; V_s= 2-3 m/s; AGS = 150 км/ч;  = 10°

Fig. 9.8. Straight in approach with a teardrop procedure turn diagram

When performing a straight-in teardrop approach, navigate to the LMM at a selected

altitude. After passing the LMM, execute a right-hand or left-hand procedure turn

corresponding to the calculated turn angle (including drift calculation) and continue

flying on this heading until reaching the estimated turn to final approach position.

When the calculated horizontal leg flight time expires (HFT), proceed to descend at

an airspeed of 150 kph and a vertical speed of 2 - 3 m/s, turn to the landing course

with a roll angle of 10° and a descent of 200 m. In the final approach, compensate

for the drift angle, maintain the desired condition of flight and avoid sideslipping and

skidding. Having reached an altitude of 100 m, transition to horizontal flight, passing

the LMM at 100 m and an airspeed of 100 - 140 kph.

After breaking through the cloud cover, adjust the helicopter’s position on the

descent as required and proceed to land.

286

9.24. Special considerations for takeoff and landing operations

at high altitudes

The minimum field dimensions for a vertical takeoff or landing in ground effect with

no obstacles are 50 x 50 m. With a 15 m obstacle height, minimum dimensions for

the field depend on altitude as follows:

 up to 1500 m: 50 x 120



2000 m: 50 x 165



3000 m: 50 x 255



3500 m: 50 x 300



4000 m: 50 x 345

The minimum field dimensions for a running takeoff or landing are:

 up to 1500 m: 50 x 200



2000 m: 50 x 225



3000 m: 50 x 350



3500 m: 50 x 410



4000 m: 50 x 475

The minimum dimensions of a field for a single engine running landing at altitudes

up to 1500 m are 50 x 190 m if landing speed is 10 - 20 kph and 50 x 360 m if

landing speed is 50 kph.

Perform a vertical takeoff with acceleration out of ground effect in cases where the

field is of limited dimensions and surrounded with obstacles, and available engine

power is insufficient to allow a hover out of ground effect.

Perform a vertical landing with a hover out of ground effect when landing on a field

of limited dimensions at helicopter gross weights allowing a hover out of ground

effect.

Perform a vertical takeoff with acceleration in ground effect in cases where the

available engine power is sufficient to allow a hover at an altitude of no less than 3

m and the field dimensions allow for acceleration in ground effect.

Perform a vertical landing with a hover in ground effect in cases where the field

dimensions and approach paths, as well as the available engine power allow for a

deceleration and hover in ground effect.

Perform a running takeoff in cases where the available engine power is sufficient for

a hover at an altitude of no less than 1 m, and the field surface properties allow a

safe takeoff run over a distance of 80 to 100 m and subsequent acceleration in

ground effect.

Perform a running landing in cases where the field surface conditions and dimensions

allow for a safe execution of this procedure.

In a landing approach for a vertical landing in ground effect, plan the deceleration

such that the outer edge of the landing point is reached at an altitude of 2 - 3 m at 5

- 10 kph. Attempt to reach a hover position in the center of the landing point with

subsequent assessment of the ground surface and ideal touchdown point. If

necessary, approach the desired touchdown point at an airspeed of 5 - 10 kph.

In case of obstacles on the landing approach, plan the approach to allow a minimum

altitude of 10 m above obstacle height.

When performing an approach to deliver external sling cargo, plan and begin the

deceleration in advance. A high altitude approach with external sling load requires

1.5 - 2 times more distance than a low altitude approach. Aggressive deceleration

leads to more difficult flight control and load instability.

9.25. Takeoff and landing on an incline

Landing fields in mountainous terrain typically consist of inclines of various grades.

Incline grade limits for vertical takeoff and landing operations without engine

shutdown are:

 nose up incline: 7°

 nose down incline: 5°

 left side up incline: 7°

 right side up incline: 2°30'

Incline grade limits for vertical takeoff and landing operations with engine shutdown

are:

 nose up incline/nose down incline/left side up incline: 3°

 right side up incline: 2°30'

Vertical takeoff and landing on an incline in wind speeds up to 5 m/s are permissible

from any wind heading. In wind conditions above 5 m/s, vertical takeoff and landing

on an incline is permissible only into the wind and within above grade limitation

guidelines. Always attempt to perform takeoff and landing from an incline in either a

nose or right side up incline position. A nose up incline position is best.

When landing in a down incline position, hover at an altitude of no less than 3 m

above ground to avoid striking the tail boom on the ground. Descend to touchdown

strictly vertically and avoid any drift, in particular backward.

As the helicopter is hovered 3 m above ground, the height of the tail boom tip from

the ground is 0.6 - 0.8 m. At touchdown and liftoff, the height of the tail boom tip is

0.3 m. Aggressive deceleration just prior to achieving a hover position during landing

or excessive reduction of collective pitch may lead to the tail boom striking the

ground.

When descending to touchdown or taking off from an incline, keep the wheel brakes

depressed.

When landing on an incline in a position perpendicular to the slope, adjust cyclic in

the up slope direction to avoid drifting down slope. A hover over an incline is

maintained with some roll angle.

When lifting off from an incline in a position perpendicular to the slope, climb strictly

vertically and avoid any drift or yaw. As the helicopter lifts off, it will exhibit a

tendency to roll in the up slope direction, requiring opposite cyclic to compensate

and maintain position.

288

DIGITAL COMBAT SIMULATOR Mi-8МТV2

10.

EMERGENCY PROCEDURES

10.1. Single engine failure

Symptoms:

 Uncommanded right yaw, the severity of which depends on the engine

power setting and airspeed at the moment of failure (the higher the

engine power setting and the lower the airspeed, the stronger the effect);

 Drop of compressor RPM and EGT of the failed engine;

 Increase of compressor RPM of the operating engine;

 Drooping of main rotor RPM;

 Illumination of the ЧР. ЛЕВ. (ПРАВ.) ДВ. (EMER PWR LFT (RGT) ENG)

light depending on the helicopter’s weight and altitude at the moment of

failure.

10.1.1. Crew actions in case of a single engine failure at an altitude above

100 m:

 Reduce the collective pitch to maintain main rotor RPM no less than 92%;

 Use cyclic and pedal control to correct uncommanded roll and yaw;

 Accelerate or decelerate as required to an airspeed of 120 kph;

 Determine which engine has failed by observing the instrumentation;

 Order the crew chief to close the fuel shutoff lever and the fire valve of

the failed engine;

WARNING

When closing the fuel shutoff levers and fire valves, use extreme caution not to shut

down the operating engine!

 Having reached an airspeed of 120 kph, operate the collective to set the

operating engine to a power setting sufficient to maintain level flight;

 Make sure engine performance is adequate and sufficient to maintain level

flight. Navigate to the nearest airfield or find a suitable landing location;

 Prior to landing, check the helicopter weight.

NOTE

1. In case of a single engine failure, the power setting of the remaining engine is

automatically increased by the electronic engine governor (EEG) system all the way

up to emergency power, depending on helicopter weight.

2. In case of EEG failure in flight, the remaining engine is not automatically set to

emergency power.

Crew actions if helicopter weight is under 12000 kg

 Performing a gliding descent, maintain airspeed 20 kph higher than

current altitude until reaching an altitude of 40 m.

 At an altitude of 40 m, begin to reduce airspeed by pulling the cyclic aft to

attain an airspeed of 40 kph at an altitude of 5 m with a vertical descent

rate of 2 - 3 m/s;

 At an altitude of 5 - 7 m, establish a landing attitude;

 From an altitude of 3 - 5 m, reduce vertical speed by increasing collective

pitch at a rate of 2 - 4°/sec. When increasing collective pitch, softly press

the right pedal to counter induced left yaw. Use the cyclic to maintain the

landing pitch angle. While increasing collective pitch, do not allow rotor

RPM to droop below 70%;

 Land at a speed of 30 kph;

 After landing, immediately proceed to smoothly lower collective to

minimum and simultaneously push the cyclic 1/3 - 1/4 travel forward to

prevent the main rotor blades from striking the tail boom.

 After nose wheel touchdown, apply the wheel brakes.

Crew actions if helicopter weight is greater than 12000 kg

The following particulars must be taken into account:

 The airspeed on the glideslope must be controlled such that 60 - 70 kph is

maintained as the helicopter reaches an altitude of 5 - 10 m;

 Land at a speed of 50 kph;

 Before landing, make sure the rotor RPM is no less than 88%.

NOTE

If at the moment of engine failure the airspeed is less than 120 kph and during

acceleration to an altitude of 10 - 20 m the helicopter does not reach sufficient

airspeed to allow level flight with a single engine operating in emergency power

setting, transition to a rapid vertical and forward deceleration to perform a landing as

described above.

10.1.2. Crew actions in case of a single engine failure at an altitude below

100 m

 Reduce collective pitch to maintain rotor RPM no less than 92% and

ensure the remaining engine attains increased (emergency) power

setting;

 Use cyclic and pedal control to correct uncommanded roll and yaw;

 If airspeed is above 120kph, begin to decelerate and climb by increasing

helicopter pitch to 10 - 15°;

 Determine which engine has failed by observing the instrumentation;

 Order the crew chief to close the fuel shutoff lever and fire valves of the

failed engine;

 Having reached an airspeed of 120 kph, operate the collective to set the

operating engine to a power setting sufficient to maintain level flight;

 Make sure engine performance is adequate and sufficient to maintain level

flight;

 When the airspeed is stabilized, navigate to the nearest airfield or find a

suitable landing location;

If at the moment of engine failure the airspeed is less than 80 kph, actions are as

follows:

 Reduce collective pitch to maintain main rotor RPM no less than 92% and

ensure the remaining engine attains increased (emergency) power

setting;

 Use cyclic and pedal controls to correct uncommanded roll and yaw;

 Accelerate or decelerate as required to an airspeed of 40 - 60 kph,

depending on helicopter weight;

 Begin to descend with a vertical speed not exceeding 3 - 4 m/s;

 Descend to the chosen airfield;

 Land.

10.2. Dual engine failure (Autorotation landing)

Symptoms:

 Uncommanded right yaw, the severity of which depends on the airspeed

at the moment of failure (the higher the engine power setting and the

lower the airspeed, the stronger the effect);

 Change in the cockpit sound of the powerplant;

 Rapid drop of the main rotor RPM;

 Drop of RPM and EGT of both engines.

10.2.1. Crew actions in case of dual engine failure at an altitude above

100 m:

 Immediately reduce collective pitch to minimum;

 Use cyclic and pedal control to correct uncommanded roll and yaw;

 Close the engine fuel shutoff levers. Order the crew chief to close the fire

valves, turn off the boost and transfer pumps;

 Accelerate or decelerate to reach the indicated glide speed of 100 - 120

kph;

 Maintain main rotor RPM at maximum, i.e. 90 - 100% by observing the

indicator and adjusting collective control to avoid peaks over 110% and

droops under 88%;

 Jettison all external payload in order to reduce weight;

 Trim the helicopter to set a shallow dive, use cyclic control to counter the

rolling moment;

 Find a suitable airfield and perform an upwind approach if possible;

 If altitude allows, the approach can be adjusted using collective control

while maintaining rotor RPM within permissible limits;

 At an altitude of 70 - 100 m, slightly and smoothly adjust cyclic control to

set and maintain a constant airspeed of 100 kph for a running landing or

70 kph for a vertical landing;

 Starting from an altitude of 50 - 70 m, use the ground to visually gauge

and control altitude above the landing point. Use cyclic control to maintain

helicopter pitch;

292

 Starting from an altitude of 10 - 15 m for a running landing or 15 - 20 m

for a vertical landing, increase collective pitch to 7 - 8° (perform a flare

within approximately 1 sec) and maintain it for 0.5 - 1 sec. If this is not

sufficient to reduce the vertical speed, increase collective pitch to 12°

(within 1 - 1.5 sec) to reduce vertical speed further;

 During the flare and with a collective pitch increase rate of 10°/sec,

increase the pitch angle to 5 - 6° in order to reduce forward airspeed for a

running landing or to 8 - 10° for a vertical landing. Maintain the pitch

angle by slightly pushing the cyclic forward;

 After landing, set collective pitch to 7 - 8° and maintain it until the landing

run is complete and the helicopter is stopped;

 Pull back the cyclic to maintain the required pitch angle during landing

until nose gear touchdown, then smoothly push the cyclic forward 1/3 -

1/4 travel and apply the wheel brakes.

NOTE

1. If the selected landing field is off course from the flight path or the approach heading must be

changed due to wind conditions, perform the required maneuver (provided sufficient altitude is

available).

2. For an autorotation landing with a 180° turn (with a roll angle of 15°) altitude must be at least

650 m.

10.2.2. Crew actions in case of dual engine failure at an altitude of 100 m

and below

If the airspeed is close to 70 kph at the moment of dual engine failure, immediately

reduce collective pitch to maintain rotor RPM of 90 - 100%. Set a gliding speed of 70

kph for a running landing or as described above for a flare maneuver for a vertical

landing in case of a dual engine failure at an altitude above 100 m. After touchdown,

during the landing roll, order the crew chief to close the fuel shutoff levers and fire

valves, switch off the boost and transfer pumps, and turn off all electrical power;

If the airspeed is more than 120 kph at the moment of dual engine failure,

immediately reduce collective pitch to maintain main rotor RPM of 88% and

simultaneously decelerate by setting helicopter pitch up to 20° depending on

airspeed and altitude (the greater the airspeed and the lower the altitude, the higher

the pitch angle) above which the engines had failed. Increase collective pitch to keep

main rotor RPM under 110%.

Close the fuel shutoff levers of both engines. Use cyclic and pedal control to correct

any uncommanded roll and yaw.

If altitude allows for a quick deceleration to 70 kph, balance the helicopter at this

speed, then follow the above recommendations. If altitude is insufficient, maintain

helicopter pitch until ready to flare (fast pitch-up) at an altitude of 15 - 20 m.

If both engines fail in a hover, crew actions are the same as in the case of a single

engine failure, but keep in mind that in a hover, yaw instability, drop of main rotor

RPM, and transition to an uncontrolled descent are more abrupt due to the high

engine power settings required for hover.

CAUTION

In case of dual engine failure, a safe landing can be performed only on a firm and

level surface. Landing in any other conditions may cause damage to the helicopter.

10.3. Onboard fire

Symptoms:

 Illumination and flashing of red fire lights ПОЖАР (FIRE) on Left

Instrument Panel;

 Illumination and flashing of red fire lights ПОЖАР ЛЕВ. ДВ. (FIRE LFT

ENG) or ПОЖАР ПРАВ. ДВ. (FIRE RGT ENGINE), ПОЖАР К0-50 (FIRE

КO-50) or ПОЖАР РЕД. ВСУ (FIRE XMSN/APU), depending on where

the fire has been detected;

 Presence of smoke, fire (not implemented in simulation);

 Illumination of the amber light I ОЧЕРЕДЬ (1^ST/MAIN DISCH)

corresponding to the compartment where a fire has been detected, upon

automatic initiation of discharge of the first extinguisher.

10.3.1. Crew actions in case of onboard fire:

а) In case of fire in the left (right) engine:

 Close the fuel shutoff lever of the affected engine;

 Order the crew chief to close the fire valve of the left (right) engine and

confirm that the first fire extinguisher has been automatically discharged;

 Abort the mission and find an airfield to land.

WARNING

After the fire has been extinguished, attempting to restart the affected engine is

prohibited.

b) In case of a gearbox or APU fire:

 If the APU is operating, shut it down;

 Confirm that the first fire extinguisher has been automatically discharged.

c) If the automatic system fails (the I ОЧЕРЕДЬ

(1ST DISCH) light is not

illuminated), engage it manually by pressing the РУЧНОЕ ВКЛЮЧЕНИЕ I

ОЧЕРЕДЬ (MANUAL 1ST DISCH) button for the corresponding compartment.

If the fire has been eliminated, then within 10 seconds from the moment of the first

discharge the ПОЖАР ЛЕВ.ДВ (FIRE LFT ENG) (or ПОЖАР ПPAВ.ДВ. (FIRE RGT

ENG), ПОЖАР КО-50 (FIRE KO-50), ПОЖАР РЕД. ВСУ (FIRE XMSN/APU) -

depending on where the fire has occurred) lights will go off while the I ОЧЕРЕДЬ

(1^ST/MAIN DISCH) light will remain on confirming that the first extinguisher has been

discharged.

If the fire has not been eliminated by the first discharge (the ПОЖАР ЛЕВ ДВ.

(FIRE LFT ENG) or ПОЖАР ПРАВ. ДВ. (FIRE RGT ENG), etc. light remains on),

manually initiate the second discharge by pressing the РУЧНОЕ ВКЛЮЧЕНИЕ II

ОЧЕРЕДЬ (MANUAL 2ND DISCH) button for the corresponding compartment.

294

If the first discharge eliminated the fire, but left signs of fire in another

compartment, press the РУЧНОЕ ВКЛЮЧЕНИЕ II ОЧЕРЕДЬ (MANUAL 2ND

DISCH) for the corresponding compartment.

NOTE

To turn off all electrical power:

Set the ГЕНЕРАТОРЫ

2 (GENERATOR 1, 2) switches to neutral (center

position);

Turn the switches АККУМУЛЯТОРЫ 1.2 (BATTERY 1, 2) to ОТКЛ. (OFF);

If the СТГ-3 starter generator is ON, set the РЕЗЕРВ ГЕНЕР (STBY GEN) switch to

ОТКЛ. (OFF).

CAUTION

After eliminating of fire, abort the mission. If the fire persists, start immediate

landing with a parachute (depending on the situation)

- not implemented in

simulation.

10.4. Engine Malfunction

10.4.1. Automatic Control System Malfunction

There are two types of Engine Automatic Control System Malfunctions.

Type 1: N1 split of 2% or more, 3% at maximum power with temp regulator

operating, and/or engine power fails to change during collective movement with Nr

stable at 95 ± 2%.

1. Land as Soon As Practical

Type 2: N1 split greater than 2% with spontaneous increase of Nr above 95 ±2%.

If on takeoff or final approach:

1. Smoothly rotate throttle to left and increase collective to maintain 95 ± 2%

Nr.

For other phases of flight.

1. Increase collective to maintain 95 ± 2% Nr.

2. Determine malfunction by moving collective down then up (maintain Nr <

98%). Resulting engine indications determine what procedures to follow:

 Case 1: N1 varies in both engines, Nr maintains 95 ± 2%.

1. Maintain 55-80 KIAS (100-150 kph)

2. Land as Soon as practical

 Case 2: N1 varies in one engine, other engine at takeoff power, Nr

maintains 95 ± 2%.

1. Reduce ECL of engine at takeoff power

2. Maintain 55 - 80 KIAS (100-150 kph)

3. Land as Soon as practical

 Case 3: N1 varies in one engine, other engine at take off power, Nr does

not maintain 95 ± 2%. (increases as collective is lowered and droops as

collective is raised)

1. Adjust throttle to maintain 95 ± 2% Nr

2. Maintain 55 - 80 KIAS (100 - 150 kph)

3. Land as Soon as practical

10.4.2. Abnormal Engine Vibration

Engine vibration is a signigicant indicator of potential catastrophic engine failure.

Verify which of the following indications are applicable and respond accordingly

A. Indications:

1. Amber HI VIBE light is illuminated

2. Red SHUT DOWN ENGINE light is illuminated.

B. Procedure:

1. HIGH VIBE LFT/RGT ENG Light On

o a. Monitor engine

o b. continue with mission

2. SHUT OFF LFT/RGT ENG Light On

o a. Reduce power

 If light goes out:

o b. Establish 70 - 75 KIAS (130-140 kph)

o c. Land as soon as practical

 If light stays on:

o d. Shut down engine (If you do not, possible engine fire after 30s )

o e. Refer to Single Engine Failure Procedure (10.1)

10.4.3. Low Engine Oil Pressure

If oil pressure is between 2-3 kgf/cm²:

1. ECL - Reduce (Affected engine)

2. Engine oil pressure - Monitor

3. Land As Soon As Practical

If oil pressure drops below 2 kgf/cm² or oil temp exceeds 150°C:

1. Shut down engine

2. Refer to Single Engine Failure Procedure

10.4.4. Failure of electronic engine governor (EEG) power turbine

channels

Symptoms:

296

 Illumination of the yellow ПРЕВ.nст ЛЕВ ДВ. (HIGH N2 LFT ENG) or

ПРЕВ. nст ПРАВ ДВ (HIGH N2 RGT ENG) light on the left instrument

panel;

 the engine does not shut down.

Crew actions:

 Temporarily switch off the EEG of the affected engine;

 Check the ПРЕВ.nст ЛЕВ. (ПРАВ.) ДВ. (HIGH N2 LFT (RGT) ENG) light

to extinguish;

 Switch on the EEG;

 If the EEG light does not illuminate after resetting the EEG, check engine

operation indicators for normal readings and if no further faults are found,

continue the mission paying particular attention to the engine

performance parameters;

 If the EEG light illuminates once again and remains on, abort the mission

and land at the nearest airfield, paying particular attention to the engine

performance parameters.

10.4.5. Failure of electronic engine governor (EEG)

Symptoms:

 Illumination of the ОТКЛ.ЭРД ЛЕВ ДВ. (GOV OFF LEFT ENG) or

ОТКЛ.ЭРД ПРАВ.ДВ. (GOV OFF RGT ENG).

Crew actions:

 Switch OFF the failed EEG;

 Continue the mission paying particular attention to engine performance

parameters.

 With a failed (switched off) EEG, the compressor RPM must not exceed

102.5%.

10.5. Main/Intermediate/Tail Rotor Gear Box Malfunction

Main Transmission Malfunction or Failure Main transmission malfunctions normally

procede catastropic failure. Early failure indications may very, depending on the

nature of the problem and the severity of failure. If a malfunction is encountered,

begin immediate decent at 70 - 75 KIAS and execute either a vertical or roll-on

landing as soon a possible. Indications and actions to be taken in the event of a main

transmission malfunction are as follows:

INDICATIONS:

 Unusual Noise or Shaking of Aircraft

 Transmission Oil Temperature above normal range (see Table 8.8)

 Transmission CHIP light illuminated

 Transmission Oil Pressure below normal operating range

 Transmission Oil Pressure warning light illuminated.

Malfunctions in the drive line can be experienced in either the electornic

synchronisation system or as a mechanical problem in any of the dynamic

components. Nr irregularities, vibrations, oil over temperature or excessive main

gearbox pressure are the basic indications of drive line problems.

10.6. Yaw contol failure

Symptoms:

 If the tail rotor or its transmission are damaged in flight, the helicopter

exhibits an abrupt left yaw, right roll, and negative pitch.

Crew actions:

 Immediately reduce collective pitch and, if altitude is sufficient, order the

crew to eject from the helicopter.

 If altitude is not sufficient for the crew to eject, crew actions are:

o Begin an autorotation descent; maintain heading by setting a roll angle

to the side opposite of the turning tendency;

o Balance the helicopter with sideslip as required; compensate the yaw

moment with lateral cyclic control;

o Find an airfield to land;

o Shut down the engines by closing the fuel shutoff levers;

o Close the fire valves, close the boost and transfer pumps (if possible);

o Perform an autorotation landing. Before landing, set a zero roll attitude

for touchdown.

If the actuator of the tail rotor is operative, but the cockpit controls are damaged

(the helicopter does not respond to pedal inputs), establish an airspeed of 120 - 130

kph, adjust collective as required to establish level flight or a shallow descent and

proceed to an airfield that can be used for a safe landing. Balance the helicopter with

sideslip as required. Execute a running landing. It is PROHIBITED to increase

collective pitch before touchdown to avoid unbalancing the helicopter.

If yaw control fails in hover or while moving at low altitude, crew actions are:

 Immediately smoothly reduce collective pitch and descend to touchdown;

 During the descent, press the right pedal and deflect the cyclic to the right

to counter left turn and left drift; pull the cyclic aft to counter the negative

pitch;

 At the moment of touchdown by the main wheels, immediately and rapidly

minimize collective pitch and shut down the engines.

10.7. Hydraulic system failure

10.7.1. Failure of main hydraulic system

Symptoms:

 The red ДУБЛИР.ВКЛЮЧЕНА (B/U SYS ON) light on the center

overhead console illuminates and starts flashing while the ОСНОВН.

ВКЛЮЧЕНА (MAIN SYS ON) extinguishes;

298

 The pressure in the main hydraulic system drops down below 42 kgf/cm²,

while the pressure in the backup hydraulic system grows up to 42 - 73

kgf/cm².

Crew actions:

 Set the ОСНОВН ГИДРОСИСТЕМА (MAIN HYD) switch to to ВЫК

(OFF).

NOTE

Switching to the backup system disengages the AP-34B autopilot and the collective

clutch release system. In this case, collective handle friction force can be adjusted

manually using the friction hand wheel (not implemented in simulation).

 Disengage the autopilot by pressing the autopilot disconnect button on

the cyclic control stick;

 Abort the mission. With a particular attention to the hydraulic system,

perform landing in the nearest airport or onto a chosen airfield.

10.7.2. In case of failure of both main and backup hydraulic systems

 Eject from the helicopter (not implemented in model).

10.8. Uncommanded left yaw during takeoff or landing

10.8.1. During a hover check before takeoff

Symptoms:

 The helicopter does not respond to right pedal input up to full deflection

and continues to yaw left;

 Main rotor RPM droops below the lower limit due to a sharp increase of

collective pitch angle.

Crew actions:

 Immediately decrease collective pitch by 1 - 2° to counter uncommanded

roll and pitch;

 Disengage the SPUU-52 tail rotor pitch limit system;

 Descend to touchdown;

 At the moment of touchdown rapidly lower collective to minimum and

order the crew chief to shut down the engines, close the fire valves, and

turn off all electrical power.

10.8.2. During a hover before landing

Symptoms:

 The helicopter does not respond to right pedal input up to full deflection

during final deceleration to hover prior to landing and continues to yaw

left;

 Main rotor RPM droops below the lower limit due to a sharp increase of

the collective pitch angle during final deceleration to hover prior to

landing;

 While turning, the helicopter loses altitude with uncontrollable evolutions

in roll and pitch.

Crew actions:

In a hover at an altitude below 10 m

 Immediately decrease collective pitch by 1 - 2° to counter uncommanded

roll and pitch;

 Disengage the SPUU-52 tail rotor pitch limit system;

 Descend to touchdown;

 At the moment of touchdown rapidly lower collective to minimum and

order the crew chief to shut down the engines, close the fire valves, and

turn off all electrical power.

In hover at an altitude above 10 m

 While applying full right pedal, quickly decrease collective pitch by 1 - 2°

and simultaneously push the cyclic forward and left to correct

uncommanded roll and pitch angles; transition to forward flight;

 Disengage the SPUU-52 tail rotor pitch limit system;

 Execute a go-around;

 Repeat the landing approach and perform a running landing.

10.9. Vortex ring state (VRS)

The helicopter is susceptible to VRS in a vertical descent of greater than 3 m/s or in

a powered glide with a forward airspeed of less than 40 kph and a vertical speed of

more than 4 m/s.

Symptoms:

 Uncommanded rapid increase of vertical speed;

 Abrupt attitude changes and severe vibrations;

 Loss of flight control effectiveness;

 Unstable readings of airspeed and vertical speed indicators.

Crew actions:

 Push the cyclic forward to establish a nose-down pitch of -10 to -20°,

increase collective to set engine power up to takeoff setting and begin to

transition into forward flight while avoiding drooping main rotor RPM

below 92%;

 Having reached an airspeed of 60 - 80 kph, transition to level flight.

NOTE

1. A transition out of a vortex ring state to level flight may require 50-200 m of

altitude loss, depending on:

 initial forward speed at the moment of entry into VRS;

 the vertical descent rate at the moment of exiting VRS;

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Mi-8MTV2 Magnificent Eight. PC flight simulator manual (2016) - page 2