SH-60B HELICOPTER. FLIGHT MANUAL (2008)

A1-H60BB-NFM-000

PROFILE

At 94 percent N_g, the aircraft will be light on its

wheels. Be vigilant for signs of dynamic

rollover; maintain a centered cyclic and be

prepared to lower collective quickly. Sideward

tip path may increase possibility of dynamic

rollover.

The WOW switch may trigger, enabling AFCS

functions associated with flight; keep collect-

ive trim switch depressed. Emergency jettison

and fuel dump functions are enabled; keep

personnel clear of the aircraft.

For shipboard operations, request amber deck,

open RSD beams and slacken Main Landing

Gear chains. Ensure chains do not become taut;

dynamic instability may result.

CAUTION

When attempting engine crossbleed starts with

the engine intake cowling removed, a hot start

may be experienced if the bleed air plug is not

installed in the anti--ice bleed air line.

For a crossbleed start, the good engine should

indicate the maximum N_g safely attainable.

Donor N_g less than 94 percent may result in hot

starts.

Note

A full fuel load is recommended when conduct-

ing a crossbleed start to mitigate effects of a

high power setting on deck.

Fireguard posted, area clear.

AIR SOURCE ECS/START switch — ENG.

APU CONTR switch — OFF.

ENGINE IGNITION switch — NORM.

Fuel Selectors — XFD/DIR (as required).

Collective — Increase to set 94 percent N_g on operating engine.

ENG — START (normal procedures apply).

Collective — FULL DOWN.

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A1-H60BB-NFM-000

PROFILE

i. PCL — FLY.

j. Fuel selectors — DIR.

k. Repeat steps f. to j. for opposite engine, as required.

Windshield anti--ice operation/interlock — CHECK.

Note

APU GEN must be ON to provide AC power

for the following checks.

a. APU — START.

b. NO. 1 and NO. 2 GENERATOR switches — OFF.

c. PILOT/COPILOT WINDSHIELD ANTI--ICE switches — ON.

d. Note increase in temperature only on the pilot windshield.

e. NO. 1 or NO. 2 GENERATOR switches — ON.

f. Note increase in windshield temperature on copilot windshield

and pilot

windshield

remains heated.

Note

When windshield anti--ice is turned on, it may

take up to 60 seconds until a noticeable change

in windshield temperature can be detected.

g. PILOT/COPILOT WINDSHIELD ANTI--ICE switches — OFF.

h. NO. 1 and NO. 2 GENERATOR switches — ON.

10. Electrical systems — CHECK.

Power transfer from the APU generator to the

NO. 1 generator may cause disengagement of

SAS 1, SAS 2, TRIM, AUTO PLT, and

Stabilator which may cause rotors to dip as low

as 4 feet.

Note

The APU GEN must be on to ensure power is

supplied to the VIDS in the event of generator

dropoff.

a. APU — VERIFY ON.

b. APU generator — VERIFY ON.

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A1-H60BB-NFM-000

PROFILE

Underfrequency protection — TEST.

(1)

NO. 1 and NO. 2 PCLs — RETARD SLOWLY.

(2)

#1 GEN and #2 GEN cautions — Appear between 92 percent and 97 percent N_r. Allow

a 3--second time delay. Note N_r.

(3)

NO. 1 and NO. 2 PCLs — FLY. Note #1 GEN and #2 GEN cautions are OUT.

AC system bus tie connector — TEST.

(1)

NO. 1 GENERATOR switch — OFF.

(a)

#1 GEN caution — APPEARS.

(b) #1 CONV caution — REMAINS OFF.

(2)

NO. 1 GENERATOR switch — ON, #1 GEN caution OUT.

(3)

Repeat steps (1) and (2) for NO. 2 generator.

DC system bus tie connector — TEST.

(1)

NO. 1 CONVERTER POWER circuit breaker (CENTER, NO. 1 AC PRI BUS, ROW

1, CB 14) — PULL.

(a) SAS 1 pushbutton — NOTE OFF, THEN RESET ON.

(b) #1 CONV caution — APPEARS.

(d) COPILOT Mode selector panel lights — ON.

(2)

NO. 1 CONVERTER circuit breaker — RESET. #1 CONV caution OUT.

(3)

Repeat steps (1) and (2), pulling NO. 2 CONVERTER POWER circuit breaker

(CORNER, NO. 2 AC PRI BUS, ROW 2, CB 4). SAS 1 pushbutton should remain ON.

AFCS voltage relay check.

Note

During this check, it is acceptable for some

AFCS fail advisory lights to illuminate.

(1) AC ESNTL BUS SUPPLY circuit breaker (NO. 1 AC PRI, CORNER, ROW 1,

CB 12) — PULL, pilot’s AI OFF flag should remain out of view.

(2) CMPTR PWR/RESET pushbutton — CHECK ON.

(3) AC ESNTL BUS SUPPLY circuit breaker — RESET.

C D

11. Flight controls — CHECK.

a. Disengage AUTO PLT, TRIM, SAS 2, and SAS 1 pushbuttons noting no jump in cyclic or

tip--path plane.

b. Engage SAS 1, SAS 2, TRIM, and AUTO PLT pushbuttons.

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A1-H60BB-NFM-000

PROFILE

A B C D

12. Ground vibration analysis (tail, oil cooler, input, NSV) — AS REQUIRED.

A B C D

13. Alternate tooth wear — AS REQUIRED.

Be vigilant for signs of dynamic rollover;

maintain a centered cyclic and be prepared to

lower collective quickly. Sideward tip path

may increase possibility of dynamic rollover.

Maintain a centered cyclic equivalent to a level

tip path on level ground. Forward tip path may

allow tail wheel to become light and allow tail

wheel to slide sideways.

The WOW switch may trigger, enabling AFCS

functions associated with flight; keep collect-

ive trim switch depressed. Fuel Dump, Rescue

Hoist, Rescue Hoist Shear, and Cargo Hook

Shear functions are enabled; keep personnel

clear of the aircraft.

Tie--downs should be removed prior to pulling

torque.

Note

The following procedures are for shore--based

operations only. No attempt should be made to

perform these procedures aboard ship.

A full fuel load is recommended when conduct-

ing these procedures to mitigate the effects of

a high power setting on deck.

Collect ATABS data in accordance with the

A1--H60CA--VIB--100 manual.

With the NO. 1

ENG operating in FLY and NO. 2 ENG shutdown:

a. Allow 5 min for stabilization at 100 percent N_r, flat pitch.

After 5 min:

b. Collective — INCREASE TORQUE TO 75--80 percent.

c. ATABS — COLLECT DATA.

d. Collective — FULL DOWN.

e. Ignition switch — NORM.

f. NO. 2 Fuel Selector — DIR.

g. NO. 2 ENG — START.

h. NO. 2 PCL — FLY.

i. Allow 5 min for stabilization at 100 percent N_r, flat pitch.

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A1-H60BB-NFM-000

PROFILE

After 5 min:

j. Collective — INCREASE TORQUE TO 35--40

percent.

k. ATABS — COLLECT DATA.

l. Collective — FULL DOWN.

m. NO. 1 ENG — SHUTDOWN.

(1) ENG IGNITION switch — OFF.

(2) NO. 1 PCL and Fuel Selector — OFF.

n. Allow 5 minutes for stabilization at 100 percent N_r, flat pitch.

After 5 min:

o. Collective — INCREASE TORQUE TO 75--80

percent.

p. ATABS — COLLECT DATA.

q. Collective — FULL DOWN.

14. Main rotor track and balance — AS REQUIRED.

CAUTION

Care must be taken when reseating struts.

Helicopter damage may occur if excessive

control inputs are made.

A B

15. Engine anti--ice system — CHECK (as required).

a. PCL (engine not being checked) — AS REQUIRED.

Failure of the ENG ANTI--ICE ON advisory to

illuminate when the ENG ANTI--ICE switch is

selected to ON, or when the N_g is below

approximately 88 percent regardless of the

switch position; or constant illumination of the

ENG ANTI--ICE ON advisory with N_g greater

than 90 percent (94 percent if OAT is 15 °C or

greater) and the ENG ANTI--ICE switch OFF

are indicative of a malfunctioning anti--ice/start

bleed valve. This condition may cause engine

flameout during low power settings such as

quick stops and autorotative flight.

Note

D Engine Anti--ice Check should be performed

into the wind on deck, in a hover over a suitable

landing pad, or in a stable, level--flight regime.

D On deck or in a hover, it may be necessary to

retard the PCL on the engine not being checked

to prevent a vertical climb.

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A1-H60BB-NFM-000

PROFILE

b. ENG ANTI--ICE switch (engine being checked) — OFF.

c. Collective

— RAISE TO INCREASE N_g OF ENGINE BEING CHECKED TO

90 PERCENT OR ABOVE (94 percent or above if OAT is above 15 ºC).

d. ENG ANTI--ICE ON advisory (engine being checked) — OFF.

e. ENG ANTI--ICE switch (engine being checked) — ON.

Confirm:

(1) ECS SHUTDOWN caution (if ECS on and AIR SOURCE ECS/START

switch — ENG).

Note

ECS will not shut down if AIR SOURCE

ECS/START switch is in APU.

(2) TGT increases 30 °C to 100 °C.

(3) ENG ANTI--ICE ON advisory appears.

(4) ENG INLET ANTI--ICE ON advisory:

(a) OAT above 13 ºC — OFF.

(b) OAT above 4 ºC up to 13 ºC — May appear.

A TGT increase greater than 100 ºC and/or

appearance of either ENG INLET ANTI--ICE

ON advisory when OAT is above 13 ºC may

result in a loss in available torque at intermedi-

ate power up to 49 percent when the engine

anti--ice system is activated. If any part of the

engine anti--ice check fails, maintenance action

is required prior to flight into icing conditions.

Note

ENG INLET ANTI--ICE ON advisory illumin-

ates when engine inlet temperature reaches 93 °C

which may take approximately 90 seconds.

f. ENG ANTI--ICE switch (engine being checked) — OFF.

Confirm:

(1) TGT decreases.

(2) ENG ANTI--ICE ON advisory — DISAPPEARS.

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A1-H60BB-NFM-000

PROFILE

(3) ECS — ON (if ECS on and AIR SOURCE ECS/START switch — ENG).

(4) ENG INLET ANTI--ICE ON advisory — DISAPPEARS (after approximately

seconds).

g. PCL (engine not being checked) — ADVANCE TO FLY (as required).

h. Repeat steps a. to g. for other engine.

[16. Blade de--ice system — CHECK (as required).

Note

A TR DE--ICE FAIL caution may appear

during the blade de--ice test while in an

electromagnetic environment.

BLADE DE--ICE TEST selector switch — NORM.

BLADE DE--ICE POWER switch — TEST.

CAUTION

Leaving the BLADE DE--ICE POWER switch

in the TEST position may cause blade damage.

PWR MAIN RTR and TAIL RTR lights — CHECK. PWR MAIN RTR light may illuminate

for 2 to 4 seconds.

CAUTION

IfPWRMAINRTRorTAILRTRlightsremain

on for more than 10 seconds, main rotor or tail

rotor damage may result. Turn BLADE

DE--ICE POWER switch OFF. If either light

remains illuminated, pull the appropriate

circuit breaker

(DE--ICE CONTR MB,

COPILOT/TAIL BLADE DE--ICE). If either

light remains illuminated, secure electrical

power.

TESTINPROGRESSlight—CHECK.Thelightshouldilluminatefor2minutes.TheICE

DETECTED caution will appear for approximately 6 seconds. No other blade de--ice

system cautions should appear. PWR MAIN RTR and TAIL RTR lights should illuminate

for 2 to 4 seconds near end of test. The TEST IN PROGRESS light should then extinguish.

BLADE DE--ICE POWER switch — OFF.

BLADE DE--ICE TEST selector switch — SYNC 1.

BLADE DE--ICE POWER switch — TEST. MR DE--ICE FAIL caution appears.

BLADE DE--ICE POWER switch — OFF. MR DE--ICE FAIL caution disappears.

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A1-H60BB-NFM-000

PROFILE

BLADE DE--ICE TEST selector switch — SYNC 2.

BLADE DE--ICE POWER switch — TEST. MR DE--ICE FAIL caution appears.

BLADE DE--ICE POWER switch — OFF. MR DE--ICE FAIL caution disappears.

BLADE DE--ICE TEST selector switch — OAT.

m. BLADE DE--ICE POWER switch — TEST. MR DE--ICE FAIL and TR DE--ICE FAIL

cautions appear.

BLADE DE--ICE POWER switch — OFF. MR DE--ICE FAIL and TR DE--ICE FAIL

cautions disappear.

CAUTION

EOT test should not be performed at ambient

temperatures above 38 °C. Between 22 °C and

38 °C, allow 5 minutesat 100 percent N_r before

this test is attempted. Between 10 °C and 22 °C,

rotormustbeturningat100percent.Exceeding

temperature parameters may cause blade

damage.

BLADE DE--ICE TEST selector switch — EOT.

BLADE DE--ICE MODE selector switch — MANUAL (M).

BLADE DE--ICE POWER switch — ON. TR DE--ICE FAIL caution appears after

approximately 15 to 30 seconds. MR DE--ICE FAIL caution appears after approximately

50 to 70 seconds. PWR MAIN RTR and/or PWR TAIL RTR lights may flash.

BLADE DE--ICE POWER switch — OFF. (TR DE--ICE FAIL and MR DE--ICE FAIL

cautions disappear).

BLADE DE--ICE TEST selector switch — NORM.

BLADE DE--ICE POWER switch — OFF.

BLADE DE--ICE MODE selector switch — AUTO.

APU generator backup — CHECK.

(1) NO. 1 or NO. 2 GENERATOR — OFF. (Applicable GEN caution light should be on).

(2) BLADE DE--ICE MODE SELECT — MANUAL (M).

(3) APU GENERATOR — ON.

(4) BLADE DE--ICE POWER — ON. (Wait 30 seconds, no de--ice lights should be on).

(5) Generator turned off in step (1) — ON. (Applicable GEN caution light should go off).

(6) BLADE DE--ICE POWER — OFF.

(7) BLADE DE--ICE MODE — AUTO.

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A1-H60BB-NFM-000

PROFILE

A B

17. HIT baseline check (on deck) — AS REQUIRED.

Note

D HIT baseline should be established on deck or

in flight. Subsequent HIT checks should be

done in the same regime as HIT baseline.

D If icing conditions exist, consideration should

be given to performing HIT Check on deck.

D HIT Checks shall be performed on the first

flight of the day and recorded on the HIT/

Power Check Log (Figure10-3) formonitoring

trends in engine performance.

ENG ANTI--ICE switches — OFF.

CAUTION

If icing conditions exist, do not keep ENG

ANTI--ICE OFF longer than necessary to

perform check.

ECS — OFF.

BAR ALT — SET TO 29.92.

N_r — 100 percent.

PCL (Engine not being checked) — IDLE.

Collective — Increase to maintain 60 percent TRQ for at least 30

seconds.

Record OAT, PA, and TGT.

Note

When using HIT TGT Reference Table, round

up temperature to the nearest value.

Repeat steps f. and g. resetting TRQ for a total of three times.

PCL — FLY.

Repeat steps e. to i. for the other engine as required.

Calculate average indicated of three indicated TGT readings.

Determine table TGT from TGT Reference Table (Figure 10-4) for recorded OAT and

pressure altitude. Record TGT on baseline worksheet.

Note

When using HIT table, round OAT up, and

pressure altitude to the nearest value.

ORIGINAL

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A1-H60BB-NFM-000

PROFILE

m. Subtract average indicated TGT from table TGT to establish TGT margin.

n. Establish TGT upper and lower limits by adding 20 °C to TGT margin and subtracting

20 °C from TGT margin (step m.). Record upper and lower limits on HIT/Power Check

Log (Figure 10-3) in helicopter ADB.

18. HIT CHECK — AS REQUIRED (see paragraph 7.17.1).

19. ENGINE ANTI--ICE, PITOT HEAT, BLADE DE--ICE, WINDSHIELD ANTI--ICE, DE--ICE

MASTER switches — AS REQUIRED.

20. FMCP — AS REQUIRED.

21. APU — SHUTDOWN (as required).

10.2.5

Mission/Weapons Systems Checklist

MSN PWR — PRI.

DPLR, RDR, DATA LINK — RADIATE, as required.

ALQ 205 — AS REQUIRED.

AAR 47 — AS REQUIRED (2--minute warm--up).

Mission systems — IPL and key.

a. SO — IPL.

b. ATO MPD — TEST.

c. Secure equipment — KEY.

d. ATO MPD — ON (after test pattern appears).

e. Secure communications — CHECK.

Mode select panels, BDHI, AI — AS REQUIRED.

7. BDHIs and AIs — CHECK.

[8.

TACAN/IFF — TEST, SET.

AOP — INITIALIZE.

[10.

NAV — INITIALIZE.

If GPS is to be used as the primary navigation mode:

a. GPS — KEY.

b. GPS — INIT/TEST (via equipment status table once GPS is available).

CAUTION

Options 1 (Get GPS PRESET) and 8 (Send Init

Data) of the GPS Data Table should not be

performed when the aircraft NAV mode is GPS,

as large and unpredictable aircraft LAT/LONG

errors may occur.

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A1-H60BB-NFM-000

PROFILE

Note

Performance of INIT SYNC to initialize the

GRP is strongly recommended. This will avoid

problems such as uncapturable fly--to points.

See NTRP 3--22.2--SH60B for establishing a

data link with the ship.

[11. Initialize FHS — AS REQUIRED.

a. FLIR EU PWR — ON.

When the FLIR EU PWR switch is ON, the

FLIRTurretautomaticallymovestotheSTOW

position. Ensure personnel and equipment are

clear of the FLIR Turret area before applying

power.

b. Hellfire PWR — ON (as required).

c. ACRT/FLIR/M299 — INIT/TEST (via equipment status table).

d. FLIR Turret — Verify operation.

e. Hellfire Attack Menu — SELECT. Verify correct missile symbology.

If missile symbology incorrect:

f. RESET LAUNCHER — SELECT.

g. FLIR Turret — STOW.

h. Hellfire PWR — OFF.

12. RAD ALTs — TEST. Press and hold the RAD ALT PUSH--TO--TEST (PTT) button.

Note

Note altitude indicator pointers indicate 100

±10 feet, the green self--test light illuminates,

and aural radar altitude warning system

(RAWS) tone heard in pilot and ATO headsets.

Release the PTT button and indicator should

return to zero.

13. EMCON — AS REQUIRED.

14. External power/data link hardwire — DISCONNECT (as required).

10.2.6

Taxi Checklist

Crew — SET FOR TAXI.

Lights — AS REQUIRED.

ORIGINAL

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A1-H60BB-NFM-000

PROFILE

3. Ordnance/Pins/Covers — ARM/REMOVE (as required).

4. Chocks — OUT.

With missiles uploaded, exercise extreme care

when chocking and chaining to avoid damage

to missile seeker or fins or injury to personnel.

Brakes — RELEASE AND CHECK.

TAIL WHEEL switch — UNLOCK.

BDHI, standby compass, AIs turn and slip — CHECK.

10.2.7

Takeoff Checklist

Lights — AS REQUIRED.

AFCS — AS REQUIRED.

STABIILATOR AUTO CONTROL pushbutton — ON.

CONTGCY PWR switch — AS REQUIRED.

PCLs — FLY.

Instruments, Caution/Advisories — CHECK.

Communication and navigation — AS REQUIRED.

a. Transponder — AS REQUIRED.

b. TACAN — AS REQUIRED.

FLIR turret — STOW (as required).

TAIL WHEEL switch — LOCK.

10. Harness — LOCKED.

11. Aircrewman Pretakeoff Checklist — COMPLETE.

12. Ordnance/pins/covers — ARM/REMOVE (as required).

13. Chocks/chains (Main, High Points, and Tail) — REMOVED.

With missiles uploaded, exercise extreme care

when chocking and chaining to avoid damage

to missile seeker or fins or injury to personnel.

14. Brakes — AS REQUIRED.

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A1-H60BB-NFM-000

PROFILE

15. RDR/BAR ALT pushbutton — AS REQUIRED.

Note

Do not engage RDR ALT HOLD until clear of

deck edge.

10.2.8 Hover Checks

A B C D

1. Torque matching — CHECK. Note any abnormal torque matching on initial collective pull to

a hover. Abnormal torque matching is indicated by a torque split of 5 percent to 10 percent,

which may match in a stable hover.

Abnormal torque matching may be indicative

of an LDS malfunction. Do not fly the

helicopter until maintenance action is

performed.

C D

Hover controllability — CHECK.

a. Hover into wind.

b. Cyclic position approximately centered.

c. Left pedal should be slightly forward of neutral.

AFCS hover — CHECK.

a. SAS — CHECK.

Note

Properly functioning SAS will dampen in-

duced pitch/roll rates, but will not necessarily

return aircraft to trimmed attitude. The tend-

ency to return to trimmed position should not

be confused with a damped response.

Cyclic control inputs should be rapid, smooth, and produce desired change in approximately

one second. While guarding cyclic, make input against trim and allow cyclic to return to

trimmed position.

(1) SAS 2, AUTO PLT pushbuttons — OFF. SAS 1, TRIM, and SAS/BOOST

pushbuttons — ON.

(2) Controllability — CHECK.

(a) Apply cyclic pulse to induce ±3° pitch attitude change. The helicopter should

exhibit a damped response to the change.

(b) Applycyclicpulsetoinduce±5°rollattitudechange.Thehelicoptershouldexhibit

a damped response to the change.

response in yaw. (The helicopter should maintain heading within ±15°.)

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A1-H60BB-NFM-000

PROFILE

(d) SAS 2 pushbutton — ON. SAS 1 pushbutton — OFF.

(e) Repeat steps (a) to (c). Damped responses of the helicopter should be noticeably

increased due to addition of hover augmentation of SAS 2.

b. Autopilot — CHECK.

(1) SAS/BOOST, SAS 1, SAS 2, TRIM, and AUTO PLT pushbuttons — ON.

(2) Momentarily remove hands and feet from flight controls and monitor. Attitude

retention in pitch and roll should be ±1° (calm wind). Heading hold should be ±2°.

(3) Induce cyclic pulse ±5° in pitch. Helicopter should return to trimmed attitude with no

more than one overshoot of 3°.

(4) Repeat step (3) for cyclic roll pulse of ±10°.

(5) With feet off pedals, make a ±20 percent torque change. Helicopter should retain

heading within ±2°.

4. Generator underfrequency — CHECK.

Power transfer from the APU generator to the

NO. 1 generator may cause disengagement of

SAS 1, SAS 2, TRIM, AUTO PLT, and

Stabilator which may cause rotors to dip as low

as 4 feet.

Note

The APU generator should be on to ensure

power is supplied to the VIDS in the event of

generator dropoff.

a. APU — START, AS REQUIRED.

b. Retard NO. 1 and NO. 2 PCLs to slowly reduce N_r to 90 percent. Allow a 3--second time

delay and note LOW ROTOR RPM warning light flashing below 96 percent N_r.

c. #1 GEN and #2 GEN cautions should not appear.

d. PCLs — FLY.

e. APU — SHUTDOWN (if not required).

Main Transmission oil pressure — RECORD (minimum 45 PSI).

RAST system main probe — CHECK.

a. Crewman open cargo hook door.

b. RAST MASTER — ON.

c. MAIN PROBE — DOWN. Status light indicates DN.

d. MESSENGER CABLE — DOWN. Status light indicates OUT.

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A1-H60BB-NFM-000

PROFILE

e. Crewman verify main probe down, messenger cable out.

f. Raise messenger cable. Position switch to OFF when status light indicates messenger cable

IN.

g. MAIN PROBE — UP. Status light indicates UP.

h. RAST MASTER — OFF.

i. Crewman close cargo hook door.

Main rotor track and balance — AS REQUIRED.

10.2.9

Climb Checks

C D

Stabilator position — CHECK. PNAC shall monitor and call out stabilator indicator position

versus airspeed.

a. STAB POS indication should begin moving up at approximately 30 KIAS.

b. STAB POS indication should be approximately 25° by 60 KIAS.

Flight instruments — CHECK.

a. Airspeed indicators — CHECK PILOT AND COPILOT INDICATORS. Maximum

difference between indicators should not be over 5 knots. Note difference.

b. VSIs — CHECK.

c. Altimeters — CHECK.

d. BDHIs and compasses — CHECK.

10.2.10 Post Takeoff Checks

1. Instruments, Caution/Advisories — CHECK.

2. RDR/BAR ALT pushbutton — AS REQUIRED.

3. CONTGCY PWR switch — AS REQUIRED.

4. Lights — AS REQUIRED.

5. Aircrewman Post Takeoff Checklist — COMPLETE.

6. COMP controllers — SLAVED and aligned.

7. Manual fuel transfer — CHECK AS REQUIRED. (Short manual transfers from each auxiliary

fuel tank should be checked. Main fuel level shall be 3,700 pounds or less prior to each

transfer.)

8. HIT check — AS REQUIRED (see paragraph 7.17.1).

Note

D At a minimum, the HIT check shall be

performed on the first flight of the day.

D For FCFs, HIT check is not required if

performing HIT Baseline check.

ORIGINAL

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PROFILE

9. Engine anti--ice check — AS REQUIRED (see paragraph 7.17.10).

10. Power Available — CHECK (as required).

a. ECS/ANTI-ICE and CONTGCY PWR switches — OFF.

b. Stabilize aircraft at intended operating altitude, level the VSI, ball centered, 100 to 130

KIAS. Airspeed is dependent on environmental conditions and gross weight.

c. Gradually increase collective until N_p begins to droop on either engine or maximum

dual--engine torque limits are reached. Stabilize for 5 seconds and record indicated torque.

11. Tactical/Combat Checklist — AS REQUIRED (see paragraph 7.8).

10.2.11 Flight Checks

C D

In--flight controllability — CHECK.

a. Increase airspeed to 100 KIAS and stabilize.

(1) Cyclic position — APPROXIMATELY CENTERED.

(2) Pedals — RIGHT PEDAL, CENTERED TO SLIGHTLY FORWARD OF NEUTRAL.

b. Airspeed indicators — CHECK PILOT AND COPILOT INDICATORS. Difference in

indicators should not be over 5 KIAS. Note difference.

c. Vibrations — NOTE ANY ABNORMAL VIBRATION LEVEL.

d. VSI — CHECK FOR DIFFERENCES.

e. Hold a steady state autorotative descent at 100 KIAS with collective full down.

Stabilator position should be 3° to 7° trailing--edge up. Record position.

A B

Engine power checks.

Note

Rapid reduction in PCL may result in rapid N_r

decay.

a. Intermediate range power (IRP) limiter check/pre--maximum power check.

Note

D If N_r droops before 839° ±10 °C, the engine is

N_g/fuel flow limited.

D If N_g/fuel flow or torque limited, the ENG

ANTI--ICE switch for the engine being

checked may be turned ON to increase TGT

and verify limiter operation.

D The 839° ±10 °C limiter should be obtainable

in OATs above --10 °C. If TGT exceeds 851 °C,

discontinue IRP limiter and maximum power

check.

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A1-H60BB-NFM-000

PROFILE

(1) Set altimeter to 29.92; establish level flight at a convenient 1,000--foot interval.

(2) Verify ENG ANTI--ICE, CONTGCY PWR, AIR SOURCE/ECS START, and ECS

switches OFF.

(3) Retard PCL on engine not being checked to 20 percent TRQ or less.

(4) While maintaining altitude, slowly increase collective until 2 percent N_r droop is

obtained. Airspeed should be adjusted to maintain level flight. Ensure torque of engine

not being checked remains below 20 percent as collective is increased.

(5) Stabilize for 10 seconds; TGT limit is 839° ±10 °C.

(6) Note TGT for IRP limiter check.

(7) With ENG ANTI--ICE switch OFF, record maximum TGT and maximum N_g for

maximum power check.

Maximum power check.

(1) Maintain level flight at current altitude.

(a) If TGT limited, reduce collective to regain 100 percent N_r. Increase airspeed as

necessary to obtain maximum TGT noted during step a.(7) while maintaining 100

percent N_r.

(b) If N_g/fuel flow limited, reduce collective to regain 100 percent N_r. Increase

airspeed as necessary to obtain maximum N_g noted during step

a.(7) while

maintaining 100 percent N_r.

(2) Stabilize for 10 seconds and record PA, OAT, TRQ, TGT, N_g, and engine oil pressure

and temperature.

Contingency range power (CRP) limiter check.

(1) CONTGCY PWR switch — ON.

(2) Increase airspeed as necessary; altitude may vary.

(3) Slowly increase collective until 2 percent N_r droop is obtained.

Note

D If N_r droops before 891° ±10 °C, the engine is

N_g/fuel flow limited.

D If N_g/fuel flow or torque limited, the ENG

ANTI--ICE switch for the engine being

checked may be turned ON to increase TGT

and verify TGT limiter operation.

(4) Stabilize for 10 seconds; TGT limit is 891° ±10 °C.

Note

The 891° ±10 °C limiter should be obtainable

in OATs above --10 °C. If TGT exceeds 903 °C,

discontinue the CRP limiter check.

(5) Record TGT.

(6) Advance PCL to FLY.

ORIGINAL

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A1-H60BB-NFM-000

PROFILE

(7) Reduce collective, check TGT below IRP limiter, CONTGCY PWR switch — OFF.

(8) ENG ANTI--ICE switch — OFF (as required).

(9) AIR SOURCE ECS/START switch — ENG.

Note

If contingency power is secured with TGT

greater than 839° ±10 °C, N_r will droop.

Repeat steps a. to c. for other engine as required.

Engine calculations.

(1) Determine torque adjusted (TRQADJ):

(a) For OAT greater than --20 °C, TRQADJ = TRQMEAS. (Torque measured during

Max Power Check).

(b) For OAT less than --20 °C, TRQADJ = TRQMEAS + [0.2 X (TGTREF --

TGTMEAS)]. TGTREF is obtained from the Torque Factor TGT Reference Chart

(Figure 10-5) for the OAT recorded during the maximum power check.

(2) Determine the target torque value (TTV) using the Maximum Power Check Chart

(Figure 10-6) for the OAT and pressure altitude recorded during the maximum power

check.

(3) Specification torque ratio (STR) = TRQADJ ÷ TTV.

(4) Determine engine torque factor (ETF). Enter Torque Factor Chart (Figure 10-7) with

calculated STR and OAT recorded during maximum power check.

(a) If OAT is below --5 °C, use the -5 °C line.

(b) If OAT is above 35 °C, use the 35 °C line.

(5) IfETFisatorabove.90,engineperformanceissatisfactory.IfETFisbelow.90,engine

performance is unsatisfactory.

(6) Record ETF on FCF record card and HIT/Power Check Log (Figure 10-3).

A B

HIT baseline check (in flight) — AS REQUIRED.

Note

D HIT baseline may be established on deck or in

flight. Subsequent HIT checks should be done

in the same regime as HIT baseline.

D If icing conditions exist, consideration should

be given to performing HIT Check on deck.

D HIT Checks shall be performed on the first

flight of the day and recorded on the

HIT/Power Check Log

(Figure

10-3) for

monitoring trends in engine performance.

10-39

ORIGINAL

A1-H60BB-NFM-000

PROFILE

ENG ANTI--ICE switches — OFF.

CAUTION

If icing conditions exist, do not keep ENG

ANTI--ICE OFF longer than necessary to

perform check.

ECS — OFF.

BAR ALT — SET TO 29.92.

N_r — 100 Percent.

Collective — INCREASE TO MAINTAIN 60 PERCENT TRQ FOR AT LEAST 30

SECONDS.

Record OAT, PA, and TGT.

Repeat steps e. and f. resetting TRQ for a total of three times.

Repeat steps e. to g. for the other engine as required.

Calculate average indicated of three indicated TGT readings.

Determine table TGT from the HIT TGT Reference Table (Figure 10-4) for recorded OAT

and pressure altitude. Record TGT on baseline worksheet.

Note

When using HIT table, round OAT up, and

pressure altitude to the nearest value.

Subtract average indicated TGT from table TGT to establish TGT margin.

Establish TGT upper and lower limits by adding 20 °C to TGT margin and subtracting 20 °C

from TGT margin (step k.). Record upper and lower limits on HIT/Power Check Log

(Figure 10-3) in helicopter ADB.

A B

ENG ANTI--ICE system — CHECK (as required).

a. PCL (engine not being checked) — AS REQUIRED.

Failure of the ENG ANTI--ICE ON advisory to

illuminate when the ENG ANTI--ICE switch is

selected to ON, or when the N_g is below

approximately 88 percent regardless of the

switch position; or constant illumination of the

ENG ANTI--ICE ON advisory with N_g greater

than 90 percent (94 percent if OAT is 15 °C or

greater) and the ENG ANTI--ICE switch OFF

are indicative of a malfunctioning anti--ice/start

bleed valve. This condition may cause engine

flameout during low power settings such as

quick stops and autorotative flight.

ORIGINAL

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A1-H60BB-NFM-000

PROFILE

Note

D Engine Anti--ice Check should be performed

into the wind on deck, in a hover over a suitable

landing pad, or in a stable, level--flight regime.

D On deck or in a hover, it may be necessary to

retard the PCL on the engine not being checked

to prevent a vertical climb.

ENG ANTI--ICE switch (engine being checked) — OFF.

Collective — RAISE TO INCREASE N_g OF ENGINE BEING CHECKED TO

PERCENT OR ABOVE. (94 percent or above if OAT is above 15 °C).

ENG ANTI--ICE ON advisory (engine being checked) — OFF.

ENG ANTI--ICE switch (engine being checked) — ON.

Confirm:

(1) ECS SHUTDOWN caution (if ECS on and AIR SOURCE ECS/START

switch — ENG).

Note

ECS will not shut down if AIR SOURCE

ECS/START switch is in APU.

(2) TGT increases 30 ºC to 100 ºC.

(3) ENG ANTI--ICE ON advisory appears.

(4) ENG INLET ANTI--ICE ON advisory:

(a) OAT above 13 ºC — OFF.

(b) OAT above 4 ºC up to 13 ºC — MAY APPEAR.

A TGT increase greater than 100 ºC and/or

appearance of either ENG INLET ANTI ICE

ON advisory when OAT is above 13 ºC may

result in a loss in available torque at

intermediate power up to 49 percent when the

engine anti--ice system is activated. If any part

of the engine anti--ice check fails, maintenance

action is required prior to flight into icing

conditions.

Note

ENG INLET ANTI--ICE ON advisory

illuminates when engine inlet temperature

reaches 93 ºC which may take approximately

90 seconds.

10-41

ORIGINAL

A1-H60BB-NFM-000

PROFILE

f. ENG ANTI--ICE switch (engine being checked) — OFF.

Confirm:

(1) TGT decreases.

(2) ENG ANTI--ICE ON advisory — DISAPPEARS.

(3) ECS — ON (if ECS on and AIR SOURCE ECS/START switch — ENG).

(4) ENG INLET ANTI--ICE ON advisory — DISAPPEARS (after approximately

seconds).

g. PCL (engine not being checked) — Advance to FLY (as required).

h. Repeat steps a. to g. for other engine.

A B

ECS — CHECK.

AIR SOURCE ECS/START switch — ENG.

APU CONTROL switch — OFF.

ECS MODE switch — AUTO.

ECS TEMP knob — COLD. (Cold air should flow from vents.)

ECS TEMP knob — HOT. (Hot air should flow from vents.)

NO. 1 or NO. 2 ENG ANTI--ICE switch — ON.

ECS SHUTDOWN caution — APPEARS. (No airflow from vents.)

ENG ANTI--ICE switch — OFF. (ECS SHUTDOWN caution OUT and hot air

flows from vents.)

ECS MODE switch — MAN.

ECS HOT/COLD switch — COLD. Hold until cold air flows from vents, then switch to

HOT. Hold until hot air flows from vents, then release.

CONTGCY PWR switch — ON.

ECS SHUTDOWN caution and #1 ENG CONT PWR and #2 ENG

CONT PWR advisories — ON. No airflow from vents.

m. CONTGCY PWR switch — OFF. ECS SHUTDOWN caution and #1 ENG CONT PWR

ON and #2 ENG CONT PWR ON advisories OUT. Hot air flows from vents.

ECS MODE switch — AS DESIRED.

6. Main rotor track and balance (120 KIAS) — AS REQUIRED.

A B C D

7. Vibration analysis (cabin health, nose absorber) (120 KIAS) — AS REQUIRED.

C D

8. Stabilator — CHECK.

a. Trim 120 KIAS in level flight.

b. Check stabilator position 1° to 7° trailing--edge down.

c. Enter a sideslip by applying left pedal against trim until ball is displaced one ball width to

the right. Check stabilator position moves down 3° from trimmed position.

ORIGINAL

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A1-H60BB-NFM-000

PROFILE

d. Release pedals and check that ball returns to one--half ball width of center, and helicopter

returns to within ±1° of trimmed heading.

e. Repeat step c. with right pedal and one ball width to the left. Check stabilator position moves

up 3° from trimmed position.

f. With fixed collective in balanced flight and feet off pedals, roll into a 45° AOB turn against

trim. Maintain 120 KIAS (this check will result in a loss of altitude).

Check:

(1) Maneuvering stability - aft cyclic is required to maintain 120 KIAS.

(2) Stabilator position moves 1° down from trim.

g. Release cyclic and allow aircraft to return to straight and level flight.

Check:

(1) No adverse pitch movements.

(2) Stabilator returns to trimmed position.

AFCS in--flight check.

Trim 120 KIAS in level flight.

SAS check.

Note

D This check should be performed in smooth air.

Cyclic control inputs should be rapid, smooth,

and produce desired attitude change in

approximately

second. While guarding

cyclic,makeinputagainsttrimandallowcyclic

to return to trim position.

D Properly functioning SAS will dampen

induced rates, but will not necessarily return

aircraft to trimmed attitude. The tendency to

return to trimmed position should not be

confused with a damped response.

(1)

SAS/BOOST, SAS 1, and TRIM pushbuttons — ON.

(2)

SAS 2 and AUTO PLT pushbuttons — OFF.

(3)

Apply cyclic pulse to induce ±3° pitch attitude changes against trim. Helicopter attitude

should exhibit a damped response.

(4)

Repeat step (3) for 5° roll attitude changes against trim.

(5)

Repeat steps (3) and (4) with SAS 2 pushbutton ON and SAS 1 pushbutton OFF. Note

an improved response with SAS 2 pushbutton ON attributed to the addition of gust

alleviation.

(6)

SAS 1 and AUTO PLT pushbuttons — ON.

10-43

ORIGINAL

A1-H60BB-NFM-000

PROFILE

Autopilot check.

(1)

Attitude and heading retention — CHECK.

Note

Steady state attitude/heading retention

tolerance in smooth air should be ±1° pitch,

roll, and heading. Airspeed retention should be

±10 KIAS.

(a)

Apply a forward cyclic pulse against trim to change pitch attitude 5°.

(b)

Attitude should return to trim with no more than one overshoot.

(c)

Repeat steps (a) and (b) for aft cyclic.

(d)

Apply a right cyclic pulse against trim to change roll attitude 10°.

(e)

Attitude should return to trim with no more than one overshoot.

(f)

Repeat steps (d) and (e) for left cyclic.

(g)

Without pressing trim release switches, press pedals to displace ball one width from

center.

(h)

Release pedals. Ball should return to 1/2 ball width of center and heading should

return within ±1° of trimmed heading.

(i)

Repeat steps (g) and (h) for opposite direction.

(j)

With pedal trim switches depressed and no force on pedals, roll helicopter against

trim into a 30° AOB and release. Helicopter should return to its original attitude

with only a slight overshoot and without adverse pitch.

(k)

Repeat step (j) for opposite direction. Response should be the same.

Coordinated turn check.

(1) With feet off pedals, roll helicopter into a 45° AOB.

(2) Ball should stay within 1/2--width of center.

(3) Roll wings level.

(4) Ball should return to center.

(5) Repeat steps (1) to (4) for opposite direction.

Cyclic and yaw trim check.

(1) Using the four--way trim switch, roll helicopter into 30° AOB.

(2) Roll should be smooth at approximately 6° per second.

(3) Release the four--way trim switch at 30° AOB. Helicopter should retain this AOB with

a slight overshoot, cyclic should return to center, and ball should stay within 1/2

ball--width of center.

(4) Return helicopter to level attitude using four--way trim switch, noting smooth response.

(5) Repeat steps (1) through (4) in opposite direction.

ORIGINAL

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A1-H60BB-NFM-000

PROFILE

(6) Beep cyclic forward for 2 seconds. Airspeed should increase 12 ±5 KIAS.

(7) Beep cyclic aft for 2 seconds. Airspeed should decrease 12 ±5 KIAS.

(8) Apply collective HDG TRIM switch to change helicopter heading 5°. The first degree

should be a flat turn for 1 second. The remaining 4° should be coordinated at 1° per

second.

(9) Repeat step (8) for opposite direction. Response should be the same.

Airspeed hold check.

(1) Trim aircraft to 120 KIAS.

(2) Slowly increase airspeed 10 KIAS by pushing cyclic against trim.

(3) Note force required and release cyclic to trim position.

(4) Airspeed should return to 120 ±5 KIAS with no more than one overshoot.

(5) Decrease airspeed 10 KIAS by pulling slowly aft on cyclic against trim.

(6) Note force required and release cyclic to trim position.

(7) Airspeed should return to 120 ±5 KIAS with no more than one overshoot.

(8) Press cyclic trim release to increase airspeed to 130 KIAS. When airspeed reaches 130

KIAS, release cyclic trim release.

(9) Airspeed should hold at speed when trim release is released.

(10) Repeat steps (8) and (9) decreasing airspeed to 120 KIAS.

g. Barometric and radar altimeter hold/heading hold check.

(1) SAS/BOOST, SAS 1, SAS 2, TRIM, and AUTO PLT pushbuttons — ON.

(2) BAR ALT pushbutton — ON.

(3) Steady--state altitude retention should be ±10 feet.

(4) Enter a standard--rate turn for 180°. Steady--state altitude retention should be ±10 feet

with ±30 foot transient.

(5) Increase AOB to at least 30°. Steady--state altitude retention should be ±30 feet with

±60 foot transient.

(6) Repeat steps (4) and (5) in opposite direction.

(7) With helicopter in trimmed straight and level flight, enter a 500--fpm rate of descent

against collective and release. Heading retention should be ±1°. Helicopter should

return to within ±10 feet of initial altitude.

(8) Repeat step (7) for climb.

(9) Press collective TRIM RELEASE and establish a 500--fpm rate of descent.

(10) Release collective TRIM RELEASE. Note BAR ALT pushbutton is ON and altitude

is within ±10 feet of designated altitude.

(11) BAR ALT pushbutton — OFF, RDR ALT pushbutton — ON.

10-45

ORIGINAL

A1-H60BB-NFM-000

PROFILE

(12) Repeat steps (3) to (10) using RDR ALT hold. Altitude retention tolerances are the same

as with the BAR ALT.

(13) RDR ALT/BAR ALT pushbutton — OFF.

10. Flight checks (140 KIAS):

a. Increase airspeed to 140 KIAS and stabilize.

b. Airspeed indicators — CHECK. Pilot and copilot indicator difference should not be more

than 5 KIAS. Note difference.

c. Vibrations — NOTE ANY ABNORMAL VIBRATIONS.

11. Main rotor track and balance (140 KIAS) — AS REQUIRED.

12. Vibration analysis (4--per, cabin absorbers) (140 KIAS) — AS REQUIRED.

C D

13. Flight checks (Vh):

a. Increase airspeed in level flight to Vh (106 percent TRQ or 839 °C ±10 TGT, whichever

occurs first).

(1) Cyclic — CENTERED LATERALLY. (At least 2 inches of forward cyclic should

remain.)

(2) Pedals — RIGHT PEDAL. (Should be no more than 1 inch forward of neutral.)

(3) Collective — Should not be against upper stop.

b. Stabilator position - 0° to 4° — Trailing--edge down.

c. Vibrations — NOTE ANY ABNORMAL VIBRATIONS.

d. Airspeed indicators — CHECK. Pilot and copilot indicator difference should not be more

than 5 KIAS. Note difference.

14. Main rotor track and balance (Vh) — AS REQUIRED.

A B C D

15. Absorber Tuning — AS REQUIRED.

CAUTION

Do not exceed Torque, N_p, or TGT limitations.

Note

Collect ATABS data in accordance with the

A1--H60--CA--VIB--100 Manual.

a. Increase airspeed to 140 KIAS and stabilize.

b. ENG SPD Trim switch — SET N_g to 97 percent.

c. ATABS — Collect Data.

d. Repeat steps b. and c. for 98, 99, 99.5 100 and 100.5 percent N_r.

e. Decrease airspeed to within the calculated single engine envelope.

ORIGINAL

10-46

A1-H60BB-NFM-000

PROFILE

f. ENG SPD trim switch — SET N_r to 98 percent.

Adjustment of the PCLs outside of the

helicopter single engine envelope may result in

an unrecoverable loss of N_r following an

engine malfunction.

g. PCL (engine with lowest ETF) — MOMENTARILY ADVANCE TO LOCKOUT.

RETARD TO SET 100 PERCENT N_r.

h. Increase airspeed to 140 KIAS and stabilize.

i. PCL — ADJUST AS REQUIRED TO MAINTAIN N_r AT 100 PERCENT.

j. ENG SPD Trim switch — SET N_r TO 101 PERCENT.

k. ATABS — COLLECT DATA.

l. Repeat steps j. and k. for 102 percent N_r.

m. Decrease airspeed to within the calculated single engine envelope.

n. PCL (engine in Lockout) — RETARD TO IDLE THEN ADVANCE TO FLY.

o. ENG SPD Trim switch — SET N_r TO 100 PERCENT (as required).

16. Automatic approach, hover coupler, and departure check.

a. Automatic approach — CHECK.

(1) Automatic Approach Checklist — COMPLETE. HVR ALT rotary knob set to 80 feet,

LONG VEL and LAT VEL controls set to 0, and SAS 2, TRIM, AUTO PLT

pushbuttons — ON.

(2) Trim helicopter for 100 KIAS and 200 feet altitude, heading into wind.

(3) Automatic approach — ENGAGE. Check for steady RDR ALT hold light. Helicopter

should decelerate at 1 knot per second.

(4) As helicopter decelerates below 80 knots groundspeed, radar altitude hold disengages

and a 120 feet--per--minute rate of descent will be commanded.

(5) At 60 KIAS, transition is made from heading hold to Doppler mode, eliminating lateral

drift.

(6) Helicopter should approach a hover at 80 feet and approximately 10 knots longitudinal

groundspeed.

(7) At ±2 feet of HVR ALT setting, RADALT hold engages.

(8) At 1 KGS, HVR mode engages. HVR pushbutton illuminates.

10-47

ORIGINAL

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PROFILE

Note

If water is smooth, the Doppler may go into

memory mode and deceleration and drift will

have to be controlled manually. As the

helicopter approaches

feet, the radar

altimeter hold will engage. When approaching

a hover, the rotor wash will disturb the water

surface enough to receive a Doppler return,

however, pilot must reengage Approach/

Hover.

Hover coupler — CHECK.

(1)

Helicopter should hover within ±4 feet of HVR ALT setting and ±2 KGS with the

LONG VEL and LAT VEL rotary knobs set at 0.

(2)

Start a 10--knot lateral drift against trim. Cross--check with hover bars and Doppler.

Release cyclic.

(3)

Helicopter should return to a stable hover.

(4)

Repeat steps (2) and (3) for opposite lateral and longitudinal drift directions.

(5)

LAT VEL rotary knob — Set to establish a 10--knot drift. Cross--check with hover bars

and Doppler. Return LAT VEL rotary knob to 0. Helicopter should return to and

maintain previously established hover.

(6)

Repeat step (5) for LAT VEL rotary knob in opposite direction.

(7)

Repeat steps (5) and (6) for LONG VEL in forward and aft directions.

(8)

Repeat steps (5) to (7) using four--way TRIM switch. Cross--check with hover bars and

Doppler. Once 10--knot drift is established press cyclic TRIM REL button and release.

Helicopter should return to a stable hover.

(9)

Move collective HDG TRIM switch to L and hold.

(10)

Helicopter should change heading to the left at 3° per second. When switch is released,

helicopter should synchronize on the heading when released within ±2°.

(11)

Repeat steps (9) and (10) to the right.

(12)

Reestablish a stable hover into the wind.

(13)

HVR ALT rotary knob — 150 FT.

(14)

Helicopter should climb to 150 ±4 feet with only one overshoot. Rate of climb should

not exceed 1,000 fpm. Upward collective movement should cease when torque reaches

a maximum of 116 percent; however, transients above 116 percent are possible.

(15)

RADAR ALTIMETER variable index — SET TO 80 FT.

(16)

HVR ALT rotary knob — 70 FT. Check for RAWS tone descending through 80 feet

(check for both pilot and copilot).

(17)

Helicopter should descend to 70 ±4 feet with only one overshoot and torque limits as

shown above. Rate of descent should not exceed 250 fpm.

ORIGINAL

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A1-H60BB-NFM-000

PROFILE

(18) HVR ALT control — 30 feet. Check for continuous RAWS tone and light below 35

feet.

(19) HVR ALT control — RESET to desired altitude.

Crew hover — CHECK.

(1) CREW HVR pushbutton — ON.

(2) HOVER TRIM light on Crew Hover Control Panel should illuminate.

(3) Hover trim grip TRIM switch — PRESS FOR FORWARD DRIFT.

Note

The control authority of the crewman is limited

to 5 knots of groundspeed in longitudinal and

lateral directions in either axis around the

lateral and longitudinal velocity set on the

AFCS control panel. The HOVER TRIM

switch output is pressure sensitive. Increase

pressure to command an increase in

groundspeed.

(4) Release switch pressure. Helicopter should return to stable hover.

(5) Repeat steps (3) and (4) for aft, left, and right.

(6) CREW HVR pushbutton — OFF.

(7) HOVER TRIM light on Crew Hover Control Panel should extinguish.

Departure — CHECK.

(1) HVR pushbutton — ILLUMINATED.

(2) Cyclic DEPART HOVER button — PRESS. DPRT light on AFCS CONTROL panel

should illuminate.

(3) Helicopter should make smooth transition to 480 feet--per--minute rate of climb and 2

knots--per--second acceleration, and fly to 500 feet altitude for RDR ALT hold and 100

KIAS for airspeed hold trimmed.

(4) Automatic Approach — ENGAGE.

(5) During deceleration, beep cyclic aft for 2 seconds to verify deceleration increase of 1

knot/second.

(6) At airspeed less than 50 KIAS, press cyclic DEPART HOVER button.

(7) Prior to reaching 500 feet AGL and 100 KIAS, momentarily press collective TRIM

RELEASE switch and cyclic TRIM RELEASE button. The AFCS should stabilize the

helicopter at the altitude and airspeed at the point of release. RAD ALT hold remains

ON.

17. Navigation, mission, and communication equipment — TEST/CHECK OPERATION.

a. UHF--1 and --2.

b. UHF backup (pilot/copilot).

c. HF.

d. TACAN.

10-49

ORIGINAL

A1-H60BB-NFM-000

PROFILE

e. UHF/ADF.

f. Doppler navigation.

g. IFF/interrogator.

h. MAD.

i. RADAR.

j. Equipment status table.

k. GPS.

18. Fuel dump — CHECK.

Fuel dumping should be manually terminated

at no less than 600 pounds total fuel. With less

than 600 pounds of fuel, fuel starvation may

occur when balanced flight is not maintained

and/or pitch attitudes exceed 15° nose up or

nose down.

a. Determine the amount of fuel to be dumped.

b. BuNo 162349 and subsequent, FUEL MGT control panel — TRANSFER/MANUAL

OVRD.

c. ALE--39 power — OFF.

d. FUEL DUMP switch — DUMP.

e. FMCP — FUEL DUMP light illuminated, as applicable.

After the desired quantity has been dumped:

f. FUEL DUMP switch — OFF.

Internal wear of the fuel dump switch guard

may not move the switch to the OFF position.

Aircrew must verify switch position when

completing dumping procedures.

g. Observe the fuel readout to ensure that dumping has ceased.

If fuel dumping continues:

h. FUEL DUMP circuit breakers — PULL.

(1) FUEL DUMP NO. 1 (CENTER, NO. 1 AC PRI, ROW 1, CB 13).

(2) FUEL DUMP PUMP (CORNER, NO. 2 AC PRI, ROW 1, CB 1).

(3) FUEL DUMP CONTR (ATO OVERHEAD, DC ESNTL, ROW 3, CB 1).

ORIGINAL

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A1-H60BB-NFM-000

PROFILE

i. FUEL DUMP switch — CYCLE.

j. Land as soon as practical.

19. FUEL MGT control panel — CHECK. (Short manual transfer from each auxiliary fuel tank

should be checked if installed. Main fuel level shall be 3,700 pounds or less prior to each

transfer.)

a. FMCP MASTER/MODE switches — TRANSFER/MANUAL OVRD.

b. External AUX tank pushbutton — SELECT.

c. Confirm FMCP FLOW light illuminates for selected tank.

d. FMCP MASTER — STOP FLOW.

e. Confirm FMCP FLOW light extinguishes.

f. Repeat steps a. to e. for remaining AUX tanks.

g. FMCP MASTER/MODE switches — AS REQUIRED.

20. Autorotation rpm check.

a. Climb to a suitable autorotation altitude.

b. When climbing through check altitude, record OAT, pressure altitude (set 29.92 on BAR

ALT), and gross weight.

c. Level off and stabilize at 75 KIAS.

D Transient N_p up--speeds of several seconds

duration are typically encountered during

autorotation entry and recovery. During entry,

the N_p should split (decrease) away from N_r

prior to reaching 109 percent and reduce to a

value of 100--103 percent N_p (PCLs in FLY).

When the main rotor has split away from N_p,

transient values as high as 110--114 percent N_p

may be seen when collective pull is initiated.

This behavior is normal. If the N_p follows N_r in

a steady manner for N_r values exceeding 109

percent, or recovery N_p peak exceeds

114

percent, an LDS malfunction is indicated.

Discontinue further autorotational flight.

Lowering the collective to minimum with an

engine LDS malfunction will cause N_p and N_r

to rise rapidly and may cause activation of N_p

overspeed resulting in engine shutdown.

D Conduct autorotation rpm check at an altitude

that will allow for power recovery at a safe

altitude. If possible, have a suitable forced

landing area within range.

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d. Slowly reduce collective to minimum and maintain 75 KIAS in balanced flight.

CAUTION

If N_r is less than 90 percent or greater than 120

percent, initiate a power recovery.

e. Vibrations — NOTE ANY ABNORMAL LEVEL.

f. Airspeed indicators — CHECK. Note difference. Pilot and copilot difference should not

be greater than 5 KIAS.

g. Record N_r when passing through the desired pressure altitude.

h. Execute a power recovery and verify proper N_p/N_r response.

i. Compute DA at which stabilized N_r was recorded (Figure 10-8).

j. Verify autorotational N_r satisfies requirements of Autorotation RPM Correction Chart

within 2 percent (Figure 10-9).

10.2.12 Landing Checklist

1. CONTGCY PWR switch — AS REQUIRED.

2. Lights — AS REQUIRED.

3. Brakes — AS REQUIRED.

4. TAIL WHEEL switch — LOCK.

5. Instruments, Caution/Advisories — CHECK.

6. BAR ALT/RAD ALT — AS REQUIRED.

7. HF Radio — OFF.

8. Armament — SAFE.

9. Harness — LOCKED.

10. Aircrewman Before Landing Checklist — COMPLETE.

11. FLIR — STOW.

12. Return to Force Checklist — AS REQUIRED.

10.2.13 Post Landing Checklist

1. Lights — AS REQUIRED.

2. Tail Wheel/Brakes — AS REQUIRED.

3. CONTGCY PWR switch — OFF.

4. Ordnance pins/covers — DEARM/SAFE/INSTALL (as required).

CAUTION

Install Hellfire missile seeker covers

(if

available) before all ground operations and

before all chock and chain operations.

ORIGINAL

10-52

A1-H60BB-NFM-000

PROFILE

5. M299 launcher SAFE/ARM switch — SAFE.

6. Chocks and Chains — AS REQUIRED.

With missiles uploaded, exercise extreme care

when chocking and chaining to avoid damage

to missile seeker or fins or injury to personnel.

7. Transponder — STBY.

8. ENGINE ANTI--ICE/PITOT HEAT/BLADE DE--ICE — OFF.

Thepitotandstaticportsbecomeextremelyhot

during operation. Care should be taken to

prevent ground personnel from touching these

surfaces. Failure to turn off pitot heat could

burn personnel whenever external

power is

applied.

10.2.14 Shutdown

TAIL WHEEL switch — LOCK.

Parking brake — SET.

Chocks — IN.

APU — START.

a. ECS — OFF.

b. AIR SOURCE ECS/START switch — APU.

c. FUEL PUMP switch — APU BOOST.

d. APU CONTR switch — ON.

APU GENERATOR switch — ON.

IFF MODE 4 CODE — HOLD (as required).

Avionics.

a. DOPPLER — AS REQUIRED.

b. MPDs — AS REQUIRED.

c. SAC power — OFF.

d. FLIR power — OFF.

e. MSN power — OFF.

10-53

ORIGINAL

A1-H60BB-NFM-000

PROFILE

f. RDR — STBY.

g. Data link — STBY.

h. TACAN — OFF.

i. IFF MASTER — OFF.

j. RADALTs — OFF.

8. BACKUP HYD PMP switch — ON.

Note

FLIR EU PWR should be secured prior to

cycling the BACKUP HYD PMP switch

because the associated power surge may cause

a gimbal lock.

9. Flight controls — POSITION AND HOLD.

10. NO. 1 and NO. 2 GENERATOR switches — OFF.

Power transfer from the APU generator to the

NO. 1 generator may cause disengagement of

SAS 1, SAS 2, TRIM, AUTO PLT, and

Stabilator which may cause rotors to dip as low

as 4 feet.

11. ENGINE IGNITION switch — OFF.

12. Lights — AS REQUIRED.

13. PCLs — IDLE.

CAUTION

Enginesshouldbecooledfor2minutesatanN_g

of 90 percent or less before moving PCLs to

OFF. If an engine is shut down without being

cooled, it should not be restarted for 4 hours

unless restart is performed within 5 minutes.

14. NO. 2 PCL and fuel selector — OFF.

15. Droop stops — IN.

16. NO. 1 PCL and fuel selector — OFF.

17. Rotor brake — ON (between 30 percent and 50 percent N_r).

18. TGT — MONITOR.

CAUTION

If TGT rises above 540 °C, perform Internal

Engine Fire emergency procedure.

ORIGINAL

10-54

A1-H60BB-NFM-000

PROFILE

A B

19. DECU codes — CHECK.

20. Engine cleaning — AS REQUIRED (see paragraph 7.17.12).

Note

If engine cleaning is required, proceed to step

14 of Engine Cleaning Procedures.

21. Blade fold — AS REQUIRED.

CAUTION

Should the blade fold system stall during fold,

cycling the BLADE FOLD switch to SPREAD

should return the rotor blades to the spread

position.

Shut down the APU immediately if a rotor

blade remains stalled in the APU exhaust.

When the ROTOR SPREAD light is not

illuminated,pressingtheRDRALTpushbutton

during blade fold system operations may cause

failure of the automatic system. Do not press

this pushbutton during blade fold operations.

Simultaneous folding of main rotor blades and

tail pylon is prohibited.

Note

Failure to suppress the DECU numerical fault

codes on the PDU will prevent the automatic

blade fold from operating due to the torque

signal being relayed to the AFCS computer.

Codes can be suppressed by pressing either

OVSP TEST A or B buttons for the affected

engine.

Area — CLEAR (wing walkers positioned as required).

BACKUP HYD PMP switch — ON.

STABILATOR AUTO CONTROL — OFF.

SAS 1 and SAS 2 — OFF, TRIM — ON, AUTO PLT — OFF.

SERVO switch — 1ST OFF or 2ND OFF.

BLADE FOLD MASTER switch — ON.

BLADE FOLD switch — FOLD.

Rotor brake — OFF.

ROTOR INDEXED light — ILLUMINATED.

10-55

ORIGINAL

A1-H60BB-NFM-000

PROFILE

Note

Blades may be manually indexed if the main

rotor index actuator/gust lock fails. Cycling the

BLADE FOLD switch OFF, pulling the RTR

HEAD INDEX MOTOR circuit breaker (NO.

2 AC PRI, SO OVHD, ROW 3, CB 5), and

cycling the BLADE FOLD switch to FOLD

may disengage the indexer. Rotate the rotor

system until the INDEXED light illuminates,

then continue with the Blade Fold Checklist.

Rotor brake — APPLY.

BAR ALT — FLASHING.

Collective, cyclic, and pedals — FREE TO POSITION.

Note

If computer is unable to null after 30 seconds,

the AFCS DEGRADED caution will appear.

To attempt another cycle, turn BLADE FOLD

switch OFF, press any FAIL ADVISORY

MODE RESET pushbutton, and repeat blade

fold sequence.

m. BAR ALT pushbutton — PRESS.

Note

Thefollowingbladestatuspanellightsequence

indicates proper operation of the fold cycle:

TRIM light flashing (blades positioned for

pitch lock insertion) and PITCH LOCKED

light illuminated (last pitch lock in). Blades

willbeginfoldingfollowingtheilluminationof

the PITCH LOCKED light. Should the

INDEXED light flicker or extinguish during

folding (indicating a loss of index), the blade

fold sequence will stall. Cycling the BLADE

FOLD switch to SPREAD should clear the

stall. When the SPREAD light illuminates, the

rotor head may be re--indexed and another fold

cycle attempted.

ROTOR FOLDED light — ILLUMINATED.

BLADE FOLD switch — OFF.

BLADE FOLD MASTER switch — OFF.

SERVO switch — CENTER.

22. BACKUP HYD PMP switch — OFF.

23. ECS — OFF.

24. Exterior/interior/NVD lights — OFF.

ORIGINAL

10-56

A1-H60BB-NFM-000

PROFILE

25. APU GENERATOR switch — OFF.

26. APU — SHUTDOWN.

a. AIR SOURCE ECS/START switch — OFF.

b. APU CONTR switch — OFF.

c. FUEL PUMP switch — OFF.

27. BATT switch — OFF.

10.2.15 Post Flight Checks

1. Interior equipment.

2. Tires/struts.

3. Leaks.

4. Missing panels.

10.3

ENGINE POWER CHECKS

The Engine Torque Factor method provides an accurateindication of available power by incorporating

ambient temperature effects into the power available calculation. An in--flight check will be made

under the following conditions:

1. When a new engine is installed.

2. When an engine is reinstalled.

3. When an engine fails the HIT check for other than a faulty anti--icing start/bleed valve or a dirty

compressor.

10.3.1 Torque Factor Terms

The following terms are used when determining the maximum torque available:

1. Torque Ratio (TR): The ratio of torque available to specification torque at the desired ambient

temperature.

2. Engine Torque Factor (ETF): The ratio of an individual engine’s torque available to

specification torque at a reference temperature of 35 °C.

3. Aircraft Torque Factor (ATF): The ratio of an individual aircraft’s power available to

specification power at a reference temperature of 35 °C. The ATF is the average of the ETFs

of both engines.

4. Specification Torque Ratio (STR): The ratio of the adjusted actual torque to the specification

torque or target torque at ambient conditions.

10.3.2 Maximum Power Check Chart

The Maximum Power Check Chart (Figure 10-6) presents the Target Torque Value (TTV) at 120 KIAS

and 100percent N_r for theoperational rangeof PAand OAT.The single-- and dual--enginetransmission

limits for continuous operation are shown and should not be exceeded.

10-57

ORIGINAL

A1-H60BB-NFM-000

PROFILE

10.3.3 Torque Factor Chart

The Torque Factor Chart (Figure 10-7) is used for obtaining the ETF by normalizing the STR to a

referenced temperature of 35 °C. For temperatures below --5 °C use --5 °C, and for temperatures above

35 °C use 35 °C.

10.3.4 Torque Factor TGT Reference

The Torque Factor TGT Reference Chart (Figure 10-5) is used to compute TRQ adjusted when OAT

is below --20 °C.

10.3.5 Torque Factor Procedure

The objective of the torque factor procedure is to obtain an ETF for each engine. This is accomplished

bytakingdataduringanin--flightmaximumpowercheckandthenusingthisdataalongwiththeTorque

Factor charts to calculate an ETF. The ETF is then displayed on the HIT/Power Check Log (Figure

10-3) for use by operational pilots and maintenance personnel.

10.3.6 Torque Factor Procedure In Flight

The inflight check will be conducted to establish/reestablish the ETF and to ensure that the engine

meets minimum power requirements. Performance data will be taken at an engine--limiting condition,

while maintaining constant altitude and adjusting airspeed as required. The ENG ANTI--ICE, ECS, and

AIR SOURCE ECS/START switches must be OFF and the altimeter will be set at 29.92 In--Hg. Data

will be taken one engine at a time. If the maximum power check is being performed because of a single

engine installation/reinstallation or failed HIT check, it is at the maintenance officer’s discretion to

obtain new ETF data for the other engine.

ORIGINAL

10-58

A1-H60BB-NFM-000

PROFILE

10.3.7

Sample Problem 1

Recorded during maximum power check, N_g limited. SH--60B.

PA = 2,000 feet.

OAT = --30 °C.

TGT measured = 790 °C.

TRQ measured = 126 percent.

From Figure 10-6 TTV = 130 percent.

From Figure 10-5 TGT ref = 802 °C.

Calculate TRQ adj:

(1) TRQ adj = TRQ measured + 0.2 (TGT ref) -

0.2

(TGT measured).

(2) TRQ adj = 126 + 0.2 (802-790).

(3) TRQ adj = 126 + 2.4.

(4) TRQ adj = 128.4 percent.

Calculate STR:

(1) STR = TRQ adj/TTV.

(2) STR = 128.4/130.

(3) STR = 0.987.

Determine ETF:

(1) From Figure 10-7 (using -5 °C line) ETF = 0.957 (satisfactory performance).

10.3.8

Sample Problem 2

Recorded during maximum power check, N_g limited. SH--60B.

a. PA = 1,000 feet.

b. OAT = 20 °C.

c. TGT measured = 840 °C.

d. TRQ measured = 112 percent.

e. From Figure 10-6, TTV = 122 percent.

f. Calculate TRQ adj:

(1) TRQ adj = TRQ measured.

(2) TRQ adj = 112.

10-59

ORIGINAL

A1-H60BB-NFM-000

PROFILE

g. Calculate STR:

(1) STR = TRQ adj/TTV.

(2) STR = 112/122.

(3) STR = 0.918.

h. Determine ETF:

(1) From Figure 10-7 (using 20 °C line), ETF = 0.878 (unsatisfactory performance).

10.4

FUNCTIONAL CHECKFLIGHT CHARTS AND TABLES

TEMPERATURE—°C

-40

-30

-20

-10

+10

+20

+30

+40

+50

TIME TO IDLE-

SECONDS USING

10,000 FT

*50

*46

*44

MIL-L-23699

(*USING MIL-l-7808)

*50

*46

*44

IDLE SPEED

MAXIMUM

66.1

66.7

67.1

67.6

68.0

68.5

68.9

69.4

69.8

70.3

^(Ng%)

MINIMUM

63.2

63.7

64.1

64.7

65.1

65.5

65.9

66.2

66.7

67.0

NOTE

D Prior to flight. Idle N

gspeedmustbegreaterthantheminimumvaluelistedforthecurrentOAT.

D This is the maximum time allowed between advancing the PCL to IDLE and N

glevelingoffatidle

speed.

D N

gidlespeedhigherthanthemaximumlimitmaybeanindicationofLDSrollpinfailureormisrigging.

Flight with N

gidlespeedhigherthanthemaximumlimitisauthorizedprovidedthehighNg idlespeed

troubleshooting procedures are completed satisfactorily.

Figure 10-2. Time to Idle/Idle Speed Chart

ORIGINAL

10-60

A1-H60BB-NFM-000

CHAPTER 11

Aerodynamic Characteristics and

Techniques in Various Regimes

11.1

GENERAL FLIGHT CHARACTERISTICS

The normal speed range extends from a rearward/lateral speed of 35 knots to a maximum forward speed of 180 KIAS.

Normally, with the stabilator in the full trailing--edge--down position, the aircraft will hover approximately 4° to 5°

noseup and 2° to 3° left wing down. During approach or slow flight (approximately 15 knots), a translational

lift--induced vibration will be felt.

11.1.1 Hover/Slow Speed Flight (At or Below Translational Lift)

In a steady, no--wind hover, the main rotor experiences a symmetrical distribution of lift dictated by the rotational

velocity and constant pitch of the rotor blades. The blade tips are moving at 725 feet per second or Mach 0.65 (65

percent of the speed of sound). Since the airflow is subsonic, the movement of the blades through the air is felt

upstream (Figure 11-1), resulting in an upward movement of air prior to coming in contact with the blade. This

induced flow causes the lifting force to be shifted aft, resulting in the generation of a drag component referred to as

induced drag.

11.1.1.1 Ground Effect

A helicopter is said to be in ground effect when the rotor disk is within one rotor diameter of the ground. Ground effect

causes the main rotor thrust vector to shift forward so that it is more vertical (more lift/less induced drag). Therefore,

less power is required to hover in ground effect than at higher altitudes. These effects are strongest close to the ground

and dissipate rapidly as altitude above the ground is increased. The SH--60B is considered to be hovering in ground

effect at radar altimeter altitudes at or below 45 feet.

For an SH--60B at the gross weight of 21,700 pounds on a standard sea level day, downwash below the rotor can

exceed 150 knots. This results in the generation of a ground vortex that surrounds the aircraft just outside the rotor

arc. It is important to consider the effects of rotor downwash and the ground vortex on personnel and other aircraft,

particularly much lighter civil aircraft.

Figure 11-1. Induced Flow

11-1

ORIGINAL

A1-H60BB-NFM-000

11.1.1.2 Transition to Forward Flight

In flight regimes other than a hover, the rotor blades, as they move around the rotor head, experience different relative

velocities. Hence, an asymmetrical distribution of lift is created. To compensate for this dissymmetry, the blades on

theadvancing sideoftherotordisk rise(flaps up), decreasing theAOA and reducing thelift generated. Theretreating

blade flaps down, increasing the AOA and generating additional lift. This process of flapping causes the pitch of the

blade to be continuously changing in a cyclic manner.

The flapping nature of the rotor system results in approximately 90°of phaselag between where inputs are madeand

their effects are felt.

11.1.1.3 Blowback

When hovering in a windless environment, the main rotor disk will be level. If the aircraft is exposed to a headwind

gust, the retreating blade sees less relative wind velocity and the advancing blade sees more relative wind velocity.

This causes the rotor disk to be tilted aft or blown back. Blowback of the main rotor disk tilts the main rotor thrust

vector aft, causing the nose of the helicopter to pitch up. This reaction is countered by pilot input in the long term

and by the hover augmentation and gust alleviation feature of SAS--2, and attitude hold feature of the autopilot, in

the short term. When transitioning to forward flight, blowback results in more forward cyclic being required to

continue helicopter acceleration. The magnitude of blowback is proportional to airspeed and lift.

11.1.2 Slow-Speed Flight

11.1.2.1 Stabilator Effect in a Hover/Slow Speed Flight Below 30 Knots

Below 30 knots, the stabilator is full trailing edge down. Increasing collective in a hover increases the amount of

downwash on the stabilator and pushes the nose up. The effect of varying collective position on nose attitude during

hoveriscompensatedforbycollective--to--longitudinalmechanicalcontrolmixing.Theairspeed,collectiveposition,

lateral acceleration, and pitch rate inputs to the stabilator system have no effect below 30 knots. In transition to

forward flight, the full--down position of the stabilator will cause a nosedown pitch until programming begins at 30

knots. Aft cyclic will be required to counter this pitching moment until programming begins.

11.1.2.2 Translational Lift

Forward flight is initiated by displacing the cyclic forward. This tilts therotor thrust vector forward. Tilting themain

rotor thrust vector forward reduces the vertical lift component. Therefore, additional increase in collective pitch may

benecessary as thehelicopterbegins to translateforward to keep it from descending. With furtherincreasein forward

speed, the mass flow rate of air through the rotor system increases, resulting in greater lift production and a rapid

decrease in induced power required for level flight. Although profile power (power required to spin the blades) and

parasite power (power required to drag nonlifting parts of the helicopter through the air) are both steadily increasing,

the reduction in induced power required results in an overall reduction in total power required. Maintaining hover

power will result in approximately a 500 foot--per--minute rate of climb at 80 KIAS.

As the airspeed reaches approximately 17 knots, a noticeable vibration will be felt as the aircraft encounters its own

ground vortex. The ground vortex is rolled up under the aircraft as speed continues to increase and dissipates as the

aircraft reaches approximately 30 knots.

11.1.2.3 Tail Winds in Transition to Forward Flight

Normally, a helicopter transitioning to forward flight from a hover is moving toward a state of less power required.

This is not the case for an aircraft transitioning to forward flight with a tailwind. When the helicopter is motionless

over a spot, the rotor disk does not care which direction the wind is coming from. Therefore, a helicopter with a

tailwind requires less power to hover than one in calm winds. As the helicopter moves forward, the rotor will reach

a condition of zero relative wind when helicopter speed matches tailwind speed. The helicopter moves into a state

of more power required for level flight. Therefore, even if the aircraft had enough power to hover in a tailwind, it

may not have enough power to continue in forward flight and reach translational lift.

11.1.3 Forward Flight

Following translational lift, the aircraft will accelerate through 30 KIAS, at which point the stabilator will begin

programming the trailing edge upward, requiring forward cyclic movement to continue the helicopter acceleration.

ORIGINAL

11-2

A1-H60BB-NFM-000

When the aircraft passes through 50 KIAS, the DAFCS will level the wings to maintain heading in balanced flight.

Above 50 KIAS, the beeper trim or trim release button must be used to establish the desired forward airspeed. The

directional control pedals will automatically move toward the position required to maintain balanced flight. Forces

opposing incorrect pilot directional flight control input will be felt. An increase in speed is accomplished by using

the beeper trim switch, or depressing the trim release button and displacing the cyclic forward until the desired

airspeed is attained. This, in turn, tilts the rotor disc forward. As it tilts forward, a greater percentage of the lift being

produced by the main rotor is being used to increase the forward airspeed of the helicopter. An increase in power is

required to restore the vertical lift component to maintain altitude. The stabilator programs to counter the nosedown

attitude experienced as the rotor disc and fuselage tilt forward and will maintain an approximately level nose attitude

up to approximately 130 KIAS. The automatic flight control system (AFCS) will maintain theheading, altitude, and

airspeed in balanced flight as selected by the pilot.

11.2

BLADE STALL

11.2.1 Blade Stall Causes

The tendency of the retreating blade to stall in forward flight limits the high--speed potential of the helicopter,

increases component stresses, and decreases component life. The retreating blade has a tendency to stall because the

blade tip is traveling at the rotational velocity minus the forward speed of the helicopter. As the in--air velocity of

the retreating blade decreases, the blade angle of attack must be increased to equalize lift to provide stabilized flight.

As the angle of attack increases, the blade will eventually stall (lost lift and increased drag). The increased drag will

cause loss of rotor speed, unless power is increased. The advancing blade, on the other hand, is traveling at a

substantially higher speed, has relatively uniform low angles of attack, and is not subject to blade stall. Blade stall

will first occur at the blade root and is most likely to occur when operating at high values of speed, gross weight,

density altitude, and power. Any of these conditions is especially aggravated by low rotor rpm. Maneuvers,

acceleration, or turbulent air, all of which increase G--load factors, will induce blade stall by reducing the airspeed

at which blade stall will occur. The blade stall chart in Chapter 22 portrays the airspeeds at various pressure altitudes,

temperatures, gross weights, rotor speeds, and load factors (angle of bank), as limited by blade stall. The blade stall

chart establishes the maximum airspeeds to allow for turbulence, mild maneuvers, and necessary control inputs to

maintain the desired flight attitude. At these speeds, roughness is encountered, but reasonable maneuvers or mild

turbulence can be tolerated. Severe turbulence or abrupt control maneuvers at this point will increase the severity of

the stall, and the helicopter will become more difficult to control. In the blade stall condition, each main rotor blade

will stall as it passes through the stall region and create vibrations per revolution equal to the number of blades. If

a stall is allowed to develop fully, loss of control will be experienced, and the helicopter will pitch upward and to

the left. The use of forward cyclic to control this pitch up is ineffective and may aggravate the stall as it increases

the blade angle of attack of the retreating blade.

11.2.2 Methods of Eliminating Roughness Caused by Blade Stall

If blade stall is causing roughness in the helicopter during high--speed flight or when maneuvering, either condition

may be eliminated by accomplishing one or any combination of the following:

1. Decrease collective pitch.

2. Decrease severity of maneuver.

3. Gradually decrease airspeed.

4. Increase rotor rpm.

5. Decrease altitude.

6. Decrease gross weight.

11-3

ORIGINAL

A1-H60BB-NFM-000

11.3

SH-60B TAIL ROTOR CHARACTERISTICS

11.3.1 Tractor Tail Rotor

A tractor tail rotor is mounted on the side of the vertical fin where the rotor slipstream is directed away from the

vertical fin, thus pulling the tail and providing an antitorque reaction for helicopter directional control. Additionally,

the tail rotor is designed to provide 2.5 percent of the total lift in hovering flight. This is required due to the SH--60B

having a relatively aft center of gravity. Having 2.5 percent of the total lift aft of the center of gravity helps lower

aircraft nose attitude in a hover. To provide this lift, the tail rotor is canted 20° from the vertical plane. The effect of

varying tail rotor thrust on aircraft nose attitude is compensated for by yaw--to--longitudinal control mixing.

11.3.2 Tail Rotor Considerations in Low-Speed Flight

11.3.2.1 Loss of Tail Rotor Authority

Tail rotor authority is limited by the maximum pitch available on the tail rotor, DA and N_r. At high torque settings

and/or high DA, the maximum pitch on the tail rotor blades may only be sufficient to provide sluggish left pedal

response. If N_r droops, the maximum tail rotor thrust available decreases rapidly, potentially leading to uncontrolled

right yaw. Maximum tail rotor authority is proportional to the square of N_r, that is at 90 percent N_r, tail rotor thrust

at maximum pitch/left pedal will be 81 percent of that at 100 percent N_r and maximum pitch/left pedal.

11.3.2.2 Loss of Translational Lift

Loss of translational lift results in increased power demand and additional antitorque requirements. If the loss of

translational lift occurs when the aircraft is in a right turn, the right turn rate will be accelerated if corrective action

is not taken. When operating near maximum power available, the increased power demand could result in rotor rpm

decay. Insufficient attention to wind direction and velocity can lead to unexpected loss of translational lift. Aircraft

heading, ground track, and groundspeed must be evaluated continually.

11.3.2.3 Hover/Air Taxi

Right, sideward flight, oraright crosswind, increases airflow across thetail, resultingin areduction inangleofattack

(AOA) for a set pedal position and a reduction in tail rotor thrust. If increased left pedal is applied, a right yaw will

occur. Yaw rate will be further amplified by increased airflow over the tail pylon, which will tend to streamline the

aircraft. When the aircraft is operated at low wheel heights, main rotor tip vortex can produce an area of downwash

turbulence that may interact with the tail rotor. Tail rotor thrust variations may require rapid pedal inputs to maintain

directional control.

11.3.2.4 Loss of Tail Rotor Effectiveness

Although the SH--60B tail rotor system is a fairly robust rotor against loss of tail rotor effectiveness (LTE), an

understanding of the principles is important to safe helicopter operation and diagnosis of tail rotor system modes of

failure. Tail rotor loss of effectiveness is caused by various relative wind conditions acting on the tail rotor blades.

The regions of tail rotor loss of effectiveness are shown in Figure 11-2 and are described below.

11.3.2.5 Factors Increasing the Likelihood of Loss of Tail Rotor Effectiveness

Recovery from a high yaw rate is more difficult in conditions requiring higher main rotor power (e.g., high gross

weight, high DA, or arresting a high descent rate).

Low airspeeds require more power to maintain flight and increased antitorque requirements. Also, streamlining effect

isreducedatlowerairspeeds.Rapidapplicationofcollectivemaycausetransientrotorrpmdrooptooccur. Adecrease

in main rotor rpm causes a greater proportional decrease in tail rotor rpm/thrust. Low N_r with left pedal application

can cause loss of directional control as tail rotor rpm decays.

ORIGINAL

11-4

A1-H60BB-NFM-000

Figure 11-2. Wind Effects on the Tail Rotor

11.3.2.6 Recovery from LTE

Should LTE occur, correct and timely response is critical. If the response in incorrect or slow, the yaw rate may

accelerate to a point where it is extremely difficult to recover. One or more complete revolutions may be experienced.

The appropriate responses to LTE can be achieved by:

1. Altitude permitting, lowering the collective to reduce torque and assist in arresting right yaw; however, if a

significant rate of descent is established, the additional power required to arrest the rate of descent may

aggravate or reinitiate loss of tail rotor effectiveness.

2. Using forward cyclic to increase airspeed and, if necessary, turning in the direction of rotation. This results

in a reduction in tail rotor thrust required and produces a streamlining effect.

3. At very low speeds or in a hover, application of full left pedal may arrest the right yaw. Understand that the

control inputs may take several seconds/revolutions to take effect partially due to the effects of momentum

and ambient conditions. Neutralizing the pedals, adding right pedal, or increasing collective will only

accelerate the yaw rate.

11.3.2.6.1 Angle of Attack (AOA)

Reduction region (060° to 120° relative). Relative winds in this region decrease the effective AOA on the tail rotor

blade element, thereby decreasing tail rotor thrust and resulting in the nose spinning to the right. This condition can

bedemonstrated in sideward flight to theright, which requires moreleft pedal input as sideslip speed increases.Also,

the right--wing down attitude required for sideslip makes the plane of the tail rotor blades more vertical so that more

of the relative wind is perpendicular to the tail rotor disk.

11-5

ORIGINAL

A1-H60BB-NFM-000

11.3.2.6.2 Weather Vaning (120° to 240° relative)

Relative winds in this region can blow on the tail pylon and cause the helicopter to spin right or left in an attempt

to place relative wind on the nose. This condition can be demonstrated in rearward flight where the nose appears to

swing randomly left and right and it is difficult to keep heading constant.

11.3.2.6.3 Vortex Ring State (210° to 330° relative)

Relative winds in this region can cause the tail rotor to ingest its own vortex (as in the main rotor at high rates of

descent and low forward airspeed). This causes a loss of tail rotor thrust and the helicopter will spin to the right. The

20° tilt of the tail rotor disk tends to help prevent this problem because the relative wind would need an upward

component to blow the vortex directly on the plane of the blades. This effect is not very apparent in sideward flight

to the left because the left--wing--down attitude required places the plane of the tail rotor blades flatter with respect

to the horizon. The vortex is blown off the right side of the tail rotor disk.

11.3.2.6.4 Main Rotor Disk Vortex (280° to 330° relative)

Relative winds in this region can cause the main rotor disk retreating blade vortex to impinge on the tail rotor blades.

This can result in sudden losses and surges to tail rotor thrust at aconstant pedal setting. It can result in thehelicopter

spinning right without enough left pedal authority to stop it. Surges in tail rotor thrust are easily controlled by

reducing left pedal input during the surge.

11.4

MANEUVERING FLIGHT

11.4.1 Maneuvering Flight Characteristics

Maneuvering flight is accomplished by moving the cyclic in the direction of the desired turn. Above 50 KIAS, the AFCS

will assist in moving the directional control pedals as required to maintain balanced flight. Above 30° angle of bank, aft

cyclic will be required to maintain the desired airspeed. In a decelerating turn, the AFCS will switch from a turn

coordination mode to a heading hold mode as the aircraft passes through 50 KIAS. Normal maneuvers are restricted to

the maneuvers and airspeed limitations as depicted in Chapter 4. A portion of the information in this section particularly

applies to highly dynamic maneuvering flight such as might be executed during evasive maneuvering.

11.4.2 Coordinated Flight

The slip/skid indicator (ball) is designed to indicate lateral acceleration on the aircraft due to sideslip in steady

zero--net--acceleration flight. In slow speed (less than 60 KIAS), high AOB (greater than 25°), nonsteady (transient)

turns, the predominant component of lateral acceleration is the weight vector, not sideslip--induced acceleration.

During slow speed, high AOB, transient turns, balanced flight (minimal sideslip) can best be maintained by

maintaining pedal position that centers the ball. Intentional sideslip within limits is permissible during maneuvering

flight; however, to prevent occurrence of high tail rotor component loads, left pedal application beyond that required

for balanced flight should only be attempted at midposition or lower collective settings.

CAUTION

Centering the ball in slow speed, high AOB, and transient turns induces

large values of aircraft sideslip and may cause high tail rotor component

loads in left turns.

Note

When a requirement to achieve a slow--speed, high--rate turn to the left

exists, such as during evasive maneuvering, lowering the collective may be

necessary to reduce main rotor torque; associated tail rotor antitorque

requirement will greatly facilitate the turn, resulting in higher turn rates.

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A1-H60BB-NFM-000

11.4.3 Tail Rotor Spar Loads in Maneuvering Flight

Counterclockwise turning single main rotor helicopters exhibit transient torque increases in forward flight with roll

rates to the left. Left roll rate increases retreating blade AOA, driving torque up, and main rotor precession loads

contribute further to this effect. Left roll rates (above approximately 30° per second in forward flight above 75 KIAS)

can combine with induced tail rotor gyroscopic and flapping loads to cause excessive tail rotor spar loading.

CAUTION

When executing high roll rate maneuvers to the left, collective should be

lowered concurrently with maneuver initiation to control transient torque

increases and reduce high tail rotor spar loads. Left roll rates in excess of

30° per second should be avoided in forward flight above 75 KIAS to

prevent damage to the tail rotor spar.

11.4.4 Main Rotor Vertical 4/Rev Vibration Cueing

During maneuvering flight, main rotor component fatigue damage occurs simultaneously with an increase in main

rotor vertical 4/rev vibration level. As an increasing G level is placed on the aircraft, the 4/rev onset will appear and

is noticed as an increasein aircraft roughness similarto that of the4/rev shudderexperienced when flying theaircraft

through transitional lift. The4/rev vibration onset is an indication that the lift--generating capability ofthe main rotor

has been exceeded and that a main rotor stall region has been created. Further attempts to increase load factor will

only increase the blade stall region, resulting in reduced maneuverability and increased component fatigue damage.

If a noticeable increase in main rotor 4/rev vibration level is observed, relax G level slightly until 4/rev vibrations

decrease to a normal level. The most effective method of reducing the stall--related main rotor 4/rev vibration level

is to reduce collective.

CAUTION

Significant increased main rotor 4/rev vibration level during maneuvering

flight indicates the onset of retreating blade stall. Main rotor component life

is being reduced and aircraft maneuvering performance degraded.

11.4.5 Main Rotor Flapping Margin

Main rotor flapping margin, a measurement of the amount of blade spindle displacement remaining in the flapping

(vertical) axis before blade motion stops are contacted, may be reduced to zero by maneuvers involving large and

rapid application of forward cyclic. Main rotor flapping margin is especially reduced when rapid forward cyclic is

coupled with low collective settings and/or aft longitudinal cg.

CAUTION

Inducement of less than 1 g flight by rapid application of forward cyclic will

result in decreased cyclic authority and may result in exceeding main rotor

flapping margin limits and droop stop pounding.

11.4.6 High AOB Turns

Figure 11-3 shows how the vertical (lift) component of main rotor thrust decreases with increasing AOB. In order

for the aircraft to maintain level flight, main rotor thrust must be increased so that lift will remain equal to weight.

For example, if a pilot does not apply additional collective in a 45° AOB turn at 300 feet, the aircraft will crash in

less than five seconds. Application of additional collective or a reduction of airspeed pitch allows the aircraft to

perform level turns.

11-7

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A1-H60BB-NFM-000

11.4.7 G-Loading in Turns

Accelerated or turning flight (G--loads) can be established by using aft cyclic and/or collective control. Cyclic

maneuvering provides a transient maneuvering capability because forward airspeed will decrease. As airspeed

decreases, transient rotor thrust decays as a result of less mass flow through the rotor disk; therefore, there is less

energy to complete the maneuver. Sustained maneuvering must be accomplished by application of collective power

so that theaircraft speed and energy areconserved. Transient G--loads applied using aft cyclic result in airspeed bleed

and eventual flight at speed less than bucket speed. Sustained G--loads applied with collective increase power required

for level flight at that airspeed (due to AOB), which results in a decrease in excess power. Maneuvers at slow speeds

are incapable of resulting in structural damage because the aircraft will encounter an aerodynamic limit of rotor thrust.

High--speed maneuvering can result in main rotor transient (cyclic) and sustained (collective) power exceeding

structural limitations. G--loads encountered in level coordinated turns at various AOB are presented in Figure 11-4.

11.4.8 Rolling Pullouts

Another situation where an aircraft can generate high G--loading is during a rolling pullout (Figure 11-5). Due to

centrifugal acceleration (G--loading), the weight vector of the aircraft increases. Lift produced by the rotor system

must be increased proportionally to the G--load to arrest the descent and establish level flight. Power can be applied

by transient powerinput (aft cyclic)and sustained powerinput (collective). This can resultin asituation wherepower

required for recovery greatly exceeds total power available in the rotor system and “mushing” occurs. During

“mushing,” the aircraft will continue to descend rapidly even though maximum power may be applied; longitudinal

cyclic control will feel sluggish, a noticeable increase in Main Rotor Vertical 4/Rev Vibrations (see paragraph 11.4.4),

and retreating blade stall may occur.

Figure 11-3. Main Rotor Thrust vs. AOB

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A1-H60BB-NFM-000

In situations where loss of lateral control is experienced in a steep turn and

asymmetrical stores load/shift in lateral cg is the suspected cause,

consideration should be given to jettisoning the stores. This should shift the

lateral cg sufficiently to provide enough control authority to recover from

the turn.

11.4.10 Power Required Exceeds Power Available

At high density altitudes, high gross weights, or when operating with reduced power, power required may exceed

poweravailable.Itmaynotbepossibletomaintainlevelflightduetolackofpower, whichwill causesettling tooccur.

The attendant loss of altitude is of minor consequence except in certain situations where sufficient altitude is not

available to achieve the airspeed necessary to maintain level flight. Careful preflight analysis of engine performance

and hover charts in Chapters 22, 23, 24, 25, 26, and 27 will aid in avoiding extreme situations. To recover from this

condition, complete the Emergency Malfunction In Flight procedures.

Note

The NATOPS density altitude calculation chart and charts indexed for

pressure altitude corrected to nonstandard temperature do NOT account for

the effect of humidity. High humidity has negligible effect on power

available, but a great effect on power required. The effects of humidity on

power required become apparent above 40 percent relative humidity. A

good rule of thumb is to add 100 feet to DA for every 10 percent of relative

humidity above 40 percent. Some weather briefs use “virtual temperature”

or temperature corrected for humidity to calculate density altitude. In this

case, no correction is necessary.

NTRP 3--22.2.4--SH60B contains a detailed description of Energy Maneuverability. Also included are a series of

charts relating speed, bank angle, and rate of climb/descent for various combinations of aircraft weight and

environmental conditions. These charts provide an invaluable preflight guide to maneuvering limitations.

11.5

DESCENDING FLIGHT AND AUTOROTATION

There are four flow states of a rotor system: Normal Thrusting, Vortex Ring, Autorotative, and Windmill Brake. Each

flow state represents a larger rate of descent relative to the induced velocity of the rotor system. In the normal thrusting

state of the rotor system, vortices are concentrated at the blade tips. The velocity profile of air relative to the rotor

is downward across the entire rotor disk. This is the condition encountered in hover, forward flight, climbing flight,

and slow rates of descent.

11.5.1 Vortex Ring State/Power Settling

Vortex ring state describes an aerodynamic condition where a helicopter may be in a vertical descent with maximum

power applied and little or no cyclic authority. The term “power settling” comes from pilot observations that the

helicopter keeps settling even though full engine power is used. In a normal, out of ground effect hover, the helicopter

is able to remain stationary by propelling a large mass of air down through the main rotor. Nearthe tips of theblades,

some of the air is recirculated, curling up from the bottom of the rotor system and rejoining the air entering the rotor

from the top. This phenomenon is common to all airfoils and is known as tip vortices. Tip vortices consume engine

power but produce no useful lift. As long as the tip vortices are small, their only effect is a small loss in rotor

efficiency; however, when the helicopter begins to descend vertically, it settles into its own downwash, which greatly

enlarges the tip vortices. This is the vortex ring state where most power developed by the engines is wasted in

accelerating the air in a doughnut pattern around the rotor while N_r remains at 100 percent.

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A1-H60BB-NFM-000

The effect is measurable at descent rates greater than 700 fpm and airspeeds between 0 and 20 KIAS and is the worst

at descent rates of approximately 1,500 fpm with airspeeds of 5 to 10 KIAS. Fully developed vortex ring state is

characterized by an unstable condition where the helicopter experiences uncommanded pitch and roll oscillations,

has little or no cyclic authority, and achieves a descent rate that may approach 6,000 fpm. It is accompanied by

increased levels of vibration.

Flight conditions causing vortex ring state should be avoided at low

altitudes because of the attendant loss of altitude necessary for recovery.

Recovery from fully developed vortex ring state may require entering

autorotation before regaining airspeed.

Note

Vortex ring state may also be entered during any dynamic maneuver which

places the main rotor in condition of high upflow and low longitudinal

airspeed. This condition is frequently seen during “quick stop” type

maneuvers or during autorotational recoveries.

To recover from this condition:

1. Decrease collective pitch.

2. Move cyclic forward.

3. Enter autorotation if altitude permits. A considerable loss of altitude may occur before the condition is

recognized and recovery is completed. During approach for landing, conditions causing vortex ring state

should be avoided.

11.5.2 Autorotative State

The autorotative state of the rotor system results in a lift--producing rotor with sufficient steady--state driving forces.

The autorotative state can be achieved at descent velocities between approximately 3,125 fpm and 4,450 fpm at

19,000 pounds. gross weight. A rotor disk in a steady--state autorotation is shown in Figure 11-6. The prop region

is 30 percent ofthe rotordisk, theautorotative region is 45 percent, and the stall region is 25 percent. The prop region

creates usable lift. The auto region produces forward--tilting force that creates both lift and a pro--rotational force that

overcomes bladedrag and keeps therotorspinning at aconstant rpm. Thestall region creates only drag (Figure11-6).

A steady--state autorotation is a balance between lift, drag, and rotational forces on the rotor system. An autorotation

is a dynamic exchange of potential and kinetic energies. Energy needed to keep the rotor turning at 100 percent N_r

and producing useful lift and rotational forces is gained from a decrease in helicopter potential energy (i.e.,

descending flight).

11.5.3 Autorotational Entry

The major variables affecting autorotational entry are altitude, airspeed, N_r, power state before entry, AOB, balanced

flight, yaw, and rotor inertia. The SH--60B has relatively low rotor inertia. Slow pilot reaction time can lead to N_r

decaying to unrecoverable levels in as little as two seconds. N_r decay rate will be greater when the helicopter is at

a high power condition before entry (i.e., climbing or at high speed). In general, more airspeed on entry is better up

to the maximum autorotational airspeed of 100 KIAS because the kinetic energy of forward flight can be translated

into spinning the rotor system as the aircraft slows down. The higher the altitude, the more time is available to

establish a steady--state autorotation.

11-11

ORIGINAL

A1-H60BB-NFM-000

Figure 11-6. Stall Region in an Autorotative Descent

11.5.3.1 Autorotational Descent

The major variables affecting autorotational descent are airspeed, N_r, AOB, yaw, gross weight, and density altitude.

Minimum rate of descent airspeed in an autorotation is approximately 75 KIAS. A decrease or increase in airspeed from

this value results in an increase in rate of descent. Increasing airspeed above 75 KIAS increases glide distance up to the

maximum glide airspeed of approximately 95 KIAS. Practice autorotations are typically shot at five knots faster than the

optimum descent and distance values. This allows the aircraft speed to slow toward an ideal condition instead of away

from it. For a given gross weight, there is one ideal rotor rpm that will provide the minimum rate of descent. As rotor

rpm is allowed to build above ideal, rate of descent increases. This is because the increase in rotor speed comes at the

expense of aircraft potential energy (i.e., a higher rate of descent). As rotor rpm decreases from ideal, the rate of descent

increases. A heavy aircraft actually will descend slower in a steady--state autorotation than a lighter one. This is because

a heavier aircraft can have a greater rate of exchange of potential energy (altitude and weight) into kinetic energy (rpm)

of the rotor system. A rotor system with more kinetic energy needs more collective pitch to govern it to the optimum rpm.

More collective pitch means more lift, therefore a smaller rate of decent for heavier aircraft. AOB increases G--loading,

which equates to a larger weight vector. A larger weight vector means that more potential energy can be traded for rotor

system kinetic energy. Rotor rpm will increase in a turn. Governing N_r at optimum will require more collective pitch (more

lift); however, the lift vector is not vertical, so rate of descent will still increase in the turn. Yaw attitude affects autorotation

due to the amount of energy transferred to or taken from the drive system due to relative wind effects on the tail rotor

blades. A higher density altitude allows the blades to spin with less resistance; therefore, more collective pitch is required

to govern to a given rpm at a higher density altitude.

Note

Aircraft descend at a higher rate at lower gross weights. Executing an

autorotation at lower gross weight allows management of momentum

(mass times velocity) during the recovery and lowers the power required

to hover upon completion. A heavier aircraft has more momentum (both

downward and forward) than a lighter aircraft, requiring more power to

arrest aircraft rate of descent and velocity during the final flare and

collective pull, as well as requiring more power to maintain hover at

recovery altitude.

11.5.4 Autorotational Recovery

When executing a flare at the bottom of an autorotation, rotor rpm increases due to the trading of aircraft forward

kinetic energy (airspeed) for rotor system kinetic energy (N_r). This increase in rpm is governed to optimum by

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A1-H60BB-NFM-000

adjusting collective pitch. The flare slows the rate of descent and since the rotor is now tilted aft, it helps slow the

forward speed of the aircraft. The aircraft attitude presents a more blunt cross section to the relative wind which also

helps slow the forward speed. In addition, the stabilator programs down, which helps slow the aircraft even more.

Final recovery is made by leveling the aircraft and applying collective pitch to cushion a tail wheel--first landing at

slow to zero groundspeed.

Note

In an actual autorotational flare, the pilot may choose to allow N_r to increase

above optimum for descent. The higher the N_r when executing the final

recovery, the more kinetic energy is available in the rotor system for the

collective pull that will cushion the landing as N_r decays.

11.6

ROLLOVER CHARACTERISTICS

11.6.1 Static Rollover

Static rollover angle is the angle at which the helicopter will tip over on its side if it is parked on a very steep

embankment. The static rollover angle of the SH--60B is approximately 28°. It occurs when the cg of the helicopter

is directly over the longitudinal axis passing through the tail gear and either main gear ground contact points.

11.6.2 Critical Rollover

Critical rollover angle is the maximum lateral angle of slope that can be negotiated in a takeoff or landing. At this

angle, full lateral cyclic input is required to trim the wheels level with the slope without sliding. With left wheel uphill

and brakes on, this angle is approximately 12°.

11.6.2.1 Dynamic Rollover

Dynamic rollover is an insidious dynamic condition that can occur during takeoff or landing with one wheel on the

ground and can result in destruction of the helicopter. It is not definable by a single number, nor is it simply a function

of slope angle or lateral control authority. These will aggravate a rollover condition, but the main contributor to

dynamic rollover is the buildup of angular velocity of the helicopter cg about the wheel touching the ground. When

the angular velocity about the wheel is greater than can becountered with full oppositecyclic, thehelicopter will roll

over. This situation can happen in less than two seconds. This is illustrated in Figure 11-7.

In this condition, the upsetting rolling moment is caused by the helicopter acting about the wheel ground--contact

point. Thus, the wheel restraint converts the lateral translation (drift) into an angular motion (roll rate) or angular

momentum. The roll rate can be very large depending on sink and drift speeds and the degree of wheel restraint. In

addition, a rolling motion to the right will be made worse by the thrust produced from the tail rotor. Conversely, a

roll to the left will decrease as the tail rotor thrust acts to provide deceleration. Upon contact with the ground, the

roll center is transferred from the helicopter cg to the touchdown wheel, resulting in higher roll inertia. The roll inertia

about the ground contact wheel can be five times greater than the roll inertia about the cg. The resulting roll rate can

be checked with opposite lateral cyclic, but lateral control will be only 1/5 of fully airborne, and the aircraft response

will be sluggish and limited by maximum cyclic displacement. If lateral cyclic cannot be displaced far enough to tilt

the rotor lift vector outside the wheel tread, then rotor lift adds to the rolling moment. Since rotor lift opposes the

only restoring force remaining, lift should be reduced.

Note

If encountering a situation where only one wheel is in contact with the

ground and a rolling moment is present, smooth reduction of collective is

the most effective corrective action the pilot can take to prevent dynamic

rollover.

11-13

ORIGINAL

A1-H60BB-NFM-000

Figure 11-7. Dynamic Rollover

11.6.3 Slope Landing/Takeoff Considerations

There are three major forces in slope landings and takeoffs (Figure 11-8). These major forces are helicopter weight,

rotor lift, and ground reaction (normal and friction forces). In general, the helicopter is controlled using collective

to balance weight with rotor lift and tilting the rotor lift into theslope, this adds to wheel friction and prevent sliding.

Slope landings can be grouped into three separate classifications: cross--slope, up--slope and down--slope.

Consideration should be given to setting the parking brake prior to landing. With the parking brake applied, an

increased potential exists for dynamicrollover; without applying theparking brake, increased potential exists forthe

helicopter to slide from the landing site. Fly the appropriate approach to a hover. From a hover the aircrew should

continuously clear and direct the helicopter for landing gear placement. Slope landings should be executed as slow,

controlled, vertical descents. The PAC should descend slowly, placing the up--slope wheel on the ground first.

From the moment the up--slope wheel touches the ground and until it is lifted from the ground, the principal control

is collective. Attempt to maintain the rotor disc level with the horizon throughout the landing and takeoff. As

collective is decreased cyclic must be displaced toward the up--slope side. As collective is increased cyclic must be

moved toward the neutral position. Balance between the two controls is indicated by a rotor disc that is level with

respect to the horizon.

Aft cyclic positions, in conjunction with low or decreasing collective pitch,

may cause rotor blades to contact the tail pylon resulting in loss of tail rotor

drive.

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A1-H60BB-NFM-000

Note

D When landing site is a combination of cross--slope and up-- or down--slope,

use the most restrictive slope landing limit. Be prepared to execute a

combination of control inputs to maintain stability.

D Depending on slope and helicopter configuration, the tail wheel may touch

down prior to the up--slope wheel.

11.6.3.1 Cross--Slope Landing

In a cross--slope environment, lateral cyclic still provides roll control, but the control power is reduced to 1/5 of its

airborne value.

Once the up--slope wheel touches the ground slowly lower the collective, coordinating lateral cyclic into the slope

until the down--slope wheel is also on the ground. The controls should always be positioned to keep the helicopter

from drifting. Continue coordinated movement of the collective and lateral cyclic until all of the helicopter weight

is resting firmly on the ground. If lateral cyclic control contacts the stop, or if rotor--to--ground clearance becomes

marginal before the down--slope wheel is firmly on the ground, return to a hover by slowly raising collective and

centering cyclic. Select another landing site with a shallower slope. After the slope landing is completed, ensure that

the helicopter will not slide or sink while smoothly lowering the collective to a full down position. As collective is

lowered, adjust lateral cyclic input to maintain the rotor disc as level as possible. This has the effect of maintaining

the CG on the up--slope side of the helicopter.

During cross--slope landings, rotor clearance is decreased on the up--slope

side. Personnel should traverse the rotor arc on the down--slope side of the

helicopter.

CAUTION

During cross--slope landings, avoid rapid collective reductions to prevent

blades from striking the ground and high roll rates about the up--slope

wheel.

11.6.3.2 Up--Slope Landings

For up--slope landings where main mounts touch first, longitudinal cyclic (either forward or aft) may be required with

coordinated collective movement to safely lower the tail wheel to the ground. If the tail wheel touches down first,

thePACshallmakesmooth,coordinatedcyclicinputssimilartoanormallanding.Amoreforwardthan normalcyclic

position may be required to prevent rolling down the slope backwards. With the tail wheel on the ground, smooth

coordinated cyclic and collective movements are required to control pitch rate and aircraft position until main mounts

are on deck. As the helicopter nears max up--slope nose attitude, the PAC should arrest pitching moment to ensure

limit is not exceeded.

11.6.3.3 Down--Slope Landings

Once the tail wheel touches the ground, slowly lower the collective, coordinating aft cyclic until the main mounts

are on deck.

11-15

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A1-H60BB-NFM-000

Figure 11-8. Slope Landing Takeoff Forces

11.6.3.4 Take Off

Slope take offs should be made by first moving cyclic into the slope to prevent drift. Using coordinated collective

and cyclic, increase power while adjusting cyclic to raise the downhill wheel first until the helicopter becomes level

and lifts off the ground. Avoid large cyclic inputs when the helicopter becomes airborne, because the roll center will

now shift back to the helicopter’s CG and full control authority will be restored. If thehelicopter rolls past level (into

the slope) during takeoff, lower collective rapidly but smoothly to avoid dynamic rollover -- check for a blocked or

hung--up wheel.

When taking off or landing on a sloping or rough surface, the helicopter can

be subjected to conditions which, with the slightest inattention by the pilot

to developing roll rates, can result in almost certaindestruction bydynamic

rollover. Keep the helicopter under control at all times. When landing or

taking off with one wheel touching the ground, use smooth collective

motion to maintain low roll rates.

The following procedures shall be observed to ensure safe landing/take off in a slope environment:

1. Execute all landings/take offs smoothly and maintain low roll rates.

2. Take off slowly so that induced roll rates can be easily controlled.

3. Maintain wheels--level during all landings/take offs.

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A1-H60BB-NFM-000

4. Always lift the downhill wheel first during take off.

5. Do not allow helicopter to drift.

6. If lateral control feels sluggish, reduce collective and check for obstructions.

7. To control roll rate, reduce collective smoothly. Avoid fast collective reductions to prevent blade strike or

induced high rates of roll about up--slope wheel.

11.7

SALTWATER POWER DETERIORATION

Salt spray ingestion in the engine may result in a loss in performance. This deterioration will be noticed as an increase

in TGT foragiven torque. The circumstances under which power deterioration may occurduring saltwateroperation

vary with a number of factors. The flight regime, gross weight, wind direction and velocity, pilot technique, duration

of maneuver, salinity of the water, and the relative density of the salt spray all have a bearing on performance

deterioration, though wind velocity and hover altitude have the most significance. Intermittent operation in moderate

salt spray conditions could expose the engines to enough salt spray to cause noticeable performance deterioration.

During prolonged operations, power deterioration will be readily apparent. Maneuvers such as hovering close to the

water in light winds, or low flights at low speeds, will generate maximum rotor downwash spray conditions. The

amount of spray observed on the windshield is the best indication of spray ingestion.

While operating in a salt spray environment for any period of time, a TGT

rise of 20 °C or more for constant torque is an indication of engine

performance degradation and possible salt encrustation. A TGT rise of

greater than

40 °C for a constant torque is an indication of engine

performance degradation which may result in compressor stall(s).

CAUTION

D Take note of windshield spray deposits. The amount of saltwater spray

observed on the windshield is usually the best indication of the salt spray

environment to which the engines are being subjected. If the spray on the

windshield is sufficient to require use of the windshield wipers, the engines

are ingesting a very significant amount of salt water. In this condition, there

is a probability of performance degradation thus, an increase in altitude is

recommended.

D In the event that hovering is necessary in the above conditions, the best

indication of performance loss is the relationship between TGT and torque.

Note

Continued engine operation in clean air may dissipate some of the salt

buildup, but this cannot be assured. Flight through rain may also be

beneficial in reducing salt buildup, thereby improving performance.

11-17

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A1-H60BB-NFM-000

11.8

FLIGHT WITH EXTERNAL LOADS

It is very important that the pilot know the riding characteristics of certain loads and the associated flight control

applications. All maneuvers that are made with external cargo loads should be gradual and well coordinated. Care

must be taken when flying with external loads that have aerodynamic characteristics (i.e., light aircraft, wings, tail

sections, sheet metal, plywood, etc.). The aerodynamic lift capabilities of these loads may amplify any oscillation

and cause the load to contact the helicopter. Hovering and turns while hovering present no unusual problems;

however, some helicopter oscillation may be noticed with low density bulky cargo when in level flight. This

oscillation can beminimized by the useof smooth control movements. When making turns at higher airspeeds, more

than the normal opposite lateral cyclic displacement is necessary to prevent excessive rolling motion into the

direction of turn. This tendency increases with airspeed and requires a slightly larger turning radius than would be

required at the same gross weight with an internal cargo load. More than normal amounts of cyclic displacement are

necessary to overcome the external cargo inertia when initiating or stopping sideward flight; however, slow steady

sideward flight presents no problems. Experience has shown that for any type of external cargo load, there is an

airspeed best suited for that particular load. This speed may vary from very low speeds with some loads to cruise

speeds or above for others. There is no one rule for flying external loads as the combination of weight, dimension,

and shape all have a direct bearing on the action of the load during flight. Increased power requirements will be

necessary to hover large flat loads that create a vertical drag factor than to hover loads of the same weight but smaller

configuration. However, if control movements are smoothly applied to preclude oscillation, and airspeed is slowly

increased to determine the riding characteristics of the load and the best airspeed for it to be flown, external cargo

can be flown satisfactorily. If high airspeeds or turbulence should cause a load to oscillate, the oscillation can be

reduced by decreasing airspeed and increasing collective.

Note

The parasitic drag of an external load will increase fuel consumption

significantly.

11.9

HELICOPTER VIBRATION

The inherent vibrations in any helicopter are those created by the mechanical functions of the engines and

transmission systems, dynamic action of the main and tail rotors, and aerodynamic effects on the fuselage. The overall

vibration level is influenced by the many individual frequencies of vibration and combinations thereof. Many

multiples of a basic frequency are felt, and often two or more different superimposed frequencies create beats. The

overall magnitude is the resultant of the amplitudes of all the frequencies and it would be difficult for the pilot to

completely separate all the types of vibrations encountered. Generally, these are divided into three categories: low,

medium, and high frequencies. Varying magnitudes of all three types of vibrations are often present in an individual

helicopter. Only through experience will the pilot be able to judge what is normal to the model and what is abnormal

and correctable. Excessive or abnormal vibration levels should be noted on a VIDS/MAF.

11.9.1 Main Rotor Specific Vibrations

11.9.1.1 One-Per-Revolution (4.3 Hz)

This vibration emanates from the main rotor system and is generally caused by main rotor head or blade imbalances.

It produces a rotary excitation of the fuselage which feels like a lateral oscillatory roll to the pilot. At high speeds,

one--per--revolution vibration will most likely be felt as a vertical vibration. The most probable causes are:

1. Main rotor blades out of track. A blade track adjustment is not warranted even though the blades visually

appear to be slightly out of track if a one--per--revolution vibration is not present. Out--of--track condition could

be caused by:

a. Damaged main rotor blade trailing edges.

b. Main rotor blade dynamic balance beyond tolerances.

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2. Worn or loose control rod end bearings. If the vibration is present in a hover only, the cause could be the same

as above, plus:

a. Main rotor blade static balance beyond tolerances.

b. Rotor head out of balance.

3. Malfunctioning blade damper.

11.9.1.2 Ground One-per-Revolution

This is a one--per--revolution lateral roll of the helicopter which often occurs during rotor engagement and is due to

the in--place misalignment of the main rotor blades, causing an out--of--balance condition in the main rotor system.

When the rotor attains flying speed, centrifugal force normally aligns the blades and the vibrations disappear. If the

vibration continues with the rotor up to speed at flat pitch but disappears when the helicopter is lifted into a hover,

the cause could be one of the following:

1. Static imbalance of the main rotor blades or rotor head.

2. Improper servicing of the landing gear strut.

11.9.1.3 One-per-Revolution (6.5 Hz)

This vibration is most probably caused by an initial aerodynamic upset which is amplified and maintained as a result of

an SAS and pilot--induced oscillation (PIO) in the longitudinal cyclic.

This vibration will most likely be encountered during autorotations over 100 KIAS at approximately 3,500 fpm rate of

descent, turns above 45° AOB, diving recoveries, or sideslips.

11.9.1.4 Four-per-Revolution (17.2 Hz)

This most common inherent vibration is caused by the dynamic response of the main rotor blades to asymmetrical

blade loading. Its intensity is greatest at low--forward speeds and during transition to a hover. It is felt as alateral/roll

shake caused by the main rotor blades traversing the downwash of preceding blades. This vibration normal to the

helicopter when felt at the point where the collective pitch is increased to sustain a hover, or when air taxiing the

helicopter just prior to applying collective pitch, and by planning the approach so that the hover can be attained with

a slow rate of final pitch application. At high speeds, the difference in the lift distribution between the advancing and

retreating main rotor blades results in heavy vibratory loads on the rotor head. It is felt as a combination of vertical

and lateral shake at the same frequency. The primary sources of excessive four--per--revolution vibrations are loose

or worn vibration absorbers, main rotor head pressure plates, swashplate and associated hardware, damper servicing,

loose stabilator, and loose cabin equipment and main landing gear struts (if experienced on deck only).

Note

D Adjusting the engine rpm switch will result in changes to four--per--revolution

vibration levels.

D Alongitudinal reversal ofthecyclicat lowspeeds(60to 90knots) mayresult

in momentary increase in four--per--revolution vibration.

11.9.2 Ground Resonance

Ground resonance is a phenomenon of multibladed helicopters like the SH--60B and is due to the cg of the rotating

blades traversing off center. Typically, it can happen during startup, takeoff, or landing. For the condition to occur,

there must be some abnormal lead/lag blade condition that would cause the cg of the rotors to progress outward,

causing further outward movement of cg. Ground resonance can be caused by a blade being badly out of track, a

peculiar set of landing conditions, or a malfunctioning damper. Ground resonance can be pilot induced and may occur

when a landing is made with a large descent coupled with lateral drift. When a wheel reaction occurs, such as a hard

one--wheel landing that would cause out--of--phase main rotor blades to be aggravated to the point where maximum

lead and lag blade displacement is realized, ground resonance can occur. If ground resonance should occur, primary

consideration should be given to getting the helicopter airborne. If this is impossible, immediately reduce collective

pitch, place the PCLs to OFF, and apply the rotor and wheel brakes.

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ORIGINAL

A1-H60BB-NFM-000

11.9.3 Tail Rotor Specific Vibrations

11.9.3.1 Tail Shake (Five Cycles per Second)

Tail shake is an aerodynamic excitation of the first lateral bending mode of the tail pylon in certain flight regimes.

This vibration will be felt as a random impulse around the yaw axis and may be more apparent with an aft center of

gravity.

11.9.3.2 Tail Rotor Drive Shaft Vibrations (High Frequency)

Generally, these vibrations are caused by an unbalanced drive shaft, bad bearings, or a failing tail/intermediate

gearbox. These vibrations can be identified during a ground run by feeling the tail cone and can also be felt as a buzz

in the pedals or a tickling in the nose similar to that of a feather.

11.9.3.3 One Times Tail Rotor Speed (20 Hz)

This vibration (1,189 cycles per minute at 100 percent) may be due to tail rotor imbalance, damage, loose hardware, pitch

change link bearing wear, or loose tail or intermediate gearbox, and is not easily isolated by the pilot. Since this frequency

is close to four per revolution (1,032 cycles per second), the vibration may be difficult to distinguish from a four per

revolution. Also, they may modulate at a frequency of 170 cycles per minute and be hard to distinguish from a one per

revolution (258 cycles per minute).

ORIGINAL

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SH-60B HELICOPTER. FLIGHT MANUAL (2008) - page 8