Pilatus Aircraft Ltd
PO Box 992
CH-6371 Stans, Switzerland
Tel: + 41 41 619 62 96
Fax: + 41 41 619 62 24
Pilatus Business Aircraft, Ltd.
11755 Airport Way
Broomfield, Colorado 80021 USA
Tel: + 1 303 465 9099
Fax: + 1 303 465 9190
Pilatus Australia Pty Ltd
17 James Schofield Drive
Adelaide Airport SA 5950, Australia
Tel: + 61 8 8234 4433
Fax: + 61 8 8234 4499
www.pilatus-aircraft.com
General Description . . . . . . . . . . . . . . . . . . . . 1
Aircraft Data . . . . . . . . . . . . . . . . . . . . . . . . . 3
External Dimensions . . . . . . . . . . . . . . . . . . . . . . 3
Cabin and Entry Dimensions . . . . . . . . . . . . . . . . . . 3
Weights . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3
Capacities . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3
Power Plant . . . . . . . . . . . . . . . . . . . . . . . . . . . 3
Propeller . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3
Specific Loadings . . . . . . . . . . . . . . . . . . . . . . . . 3
Limitations . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3
Airspeed . . . . . . . . . . . . . . . . . . . . . . . . . . . 3
Altitude . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4
Operations . . . . . . . . . . . . . . . . . . . . . . . . . . 4
Flight Crew . . . . . . . . . . . . . . . . . . . . . . . . . . 4
Outside Air Temperature . . . . . . . . . . . . . . . . . . 4
Flight Load Factor . . . . . . . . . . . . . . . . . . . . . . 4
Propeller . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11
Electrical . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11
General . . . . . . . . . . . . . . . . . . . . . . . . . . 11
Junction Boxes . . . . . . . . . . . . . . . . . . . . . . . 12
Circuit Breakers . . . . . . . . . . . . . . . . . . . . . . 13
Controls and Indicators — Overhead Panel . . . . . . . 13
Fuel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13
Flight Controls . . . . . . . . . . . . . . . . . . . . . . . . . 13
Landing Gear . . . . . . . . . . . . . . . . . . . . . . . . . . 15
Hydraulic Power . . . . . . . . . . . . . . . . . . . . . . . . 16
Environmental Control . . . . . . . . . . . . . . . . . . . . 17
Oxygen . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 17
Fire Protection . . . . . . . . . . . . . . . . . . . . . . . . . 17
Ice and Rain Protection . . . . . . . . . . . . . . . . . . . . 17
Lighting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 18
Internal . . . . . . . . . . . . . . . . . . . . . . . . . . . 18
External . . . . . . . . . . . . . . . . . . . . . . . . . . 19
Airframe/Structures . . . . . . . . . . . . . . . . . . . 5
General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5
Fuselage . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5
Doors and Emergency Exit . . . . . . . . . . . . . . . . . . . 6
Windows . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6
Seats/Restraint Systems . . . . . . . . . . . . . . . . . . . . 6
Equipment and Furnishings . . . . . . . . . . . . . . . . . . . 7
Cabin Volume . . . . . . . . . . . . . . . . . . . . . . . . . . 7
Baggage Compartment . . . . . . . . . . . . . . . . . . . . . 7
Empennage . . . . . . . . . . . . . . . . . . . . . . . . . . . 7
Wings . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7
Crew Alert System (CAW) . . . . . . . . . . . . . . . . . . . 19
Stall Warning . . . . . . . . . . . . . . . . . . . . . . . . . . 19
Figures
Page
1. Interior Configurations . . . . . . . . . . . . . . . . . . . 1
2. Aircraft Planform . . . . . . . . . . . . . . . . . . . . . . 2
3. Ground Turning Radius . . . . . . . . . . . . . . . . . . . 4
4. Fuselage Structure Showing Cabin and Cargo Doors . . . 5
5. Standard and Executive Seats . . . . . . . . . . . . . . . 6
6. Example Cockpit Arrangement . . . . . . . . . . . . . 8–9
7. PT6A-67P Engine . . . . . . . . . . . . . . . . . . . . . 10
8. Electrical Power System . . . . . . . . . . . . . . . . . 12
Aircraft Systems . . . . . . . . . . . . . . . . . . . . . . 8
9. Fuel System . . . . . . . . . . . . . . . . . . . . . . . . 14
Cockpit Arrangement . . . . . . . . . . . . . . . . . . . .8–10
10. Landing Gear System . . . . . . . . . . . . . . . . . . . 15
Power Plant . . . . . . . . . . . . . . . . . . . . . . . . . . 10
11. Brake System . . . . . . . . . . . . . . . . . . . . . . . 18
The PC-12 is a large single-engine turbo-
grass and prepared, unpaved runways.
prop utility aircraft designed to perform
• Complete icing protection for flight
a wide range of missions, including the
into known icing conditions.
transport of passengers, cargo or combina-
• Advanced Honeywell Primus Apex
tions of both.
Design features include:
configuration can also be converted to a
complete cargo role by removing all nine
passenger seats.
The optional Executive configuration
cockpit design for reduced workload,
consists of two crew seats plus seating
single-pilot operations.
for six to eight passengers. Seating is
• A Pratt & Whitney Canada PT6A-67P
The PC-12 can be outfitted in a Stan-
turboprop engine, flat-rated at 1,200
dard or optional Executive configuration.
SHP for takeoff and climb.
arranged to offer the choice of relaxing privately or working, with conference seating
for four. This configuration includes six ex-
• A pressurized, large volume cabin
The Standard configuration consists of
ecutive chairs, two refreshment cabinets,
quickly convertible from an all-
two crew seats plus seating for up to nine
three folding tables, in-flight accessible
passenger to an all-cargo or a
passengers with an in-flight accessible
baggage area and a private chemical flush-
combination of passenger/cargo
baggage area. The Standard configuration
ing toilet. The Executive configuration can
configurations.
can easily be converted to a combination
be converted to a combination passenger/
passenger/cargo role by removing the aft
cargo role by removing the aft two to four
three to five passenger seats. This results
passenger seats.
• A high-lift wing for exceptional short
field performance.
• Separate forward passenger and aft
cargo doors.
• Retractable landing gear with trailinglink main landing gear capable of
in a large, usable cargo area in the aft
The PC-12 is also well suited for a vari-
cabin. An optional Combi Conversion Kit
ety of special operations, such as air am-
provides the means to secure cargo during
bulance transport, for which dedicated or
passenger/cargo operations. The Standard
quick change configurations are available.
Figure 1 Interior Configurations
PC-12
PC-12
PC-12
PC-12
PC-12
Executive
Standard
Combi
Cargo
Air Ambulance
(other options
Commuter
available)
September 2007
(other options
available)
All information subject to change without notice
1
PC-12 TECHNICAL DESCRIPTION
Figure 2 Aircraft Planform
NOTES
1. Dimensions shown
are based on standard
empty weight and
proper nose gear and
tire inflation.
2. Propeller ground clearance: 12.5 in (32.0
cm) with nose tire fully
inflated and airplane
at maximum takeoff
weight and most forward C.G. loading. 1.9
in (4.9 cm) with nose
tire fully deflated and
nose gear strut bottomed out.
3. Total wing area is 277.8
ft2 (25.81 m2).
2
All information subject to change without notice
September 2007
External Dimensions
Capacities
Wing Span
53 ft 4 in
16.28 m
Length
47 ft 3 in
14.40 m
Height
14 ft
4.26 m
Wing Area
277.8 ft
2
25.81 m2
Horizontal Tail Span
17 ft 1 in
5.20 m
Propeller Diameter
8 ft 9 in
2.67 m
Propeller Ground Clearance
12.5 in
32.0 cm
14 ft 10 in
11 ft 5 in
4.53 m
3.48 m
Nose Tire
6.25 in
6.0 in
17.5 in
60 psi
Main Tires
8.5 in
10.0 in
22.0 in
60 psi
Landing Gear
Track
Wheelbase
Tire dimensions
Width
Inner Diameter
Outer Diameter
Pressure
16 ft 11 in
5 ft 0 in
4 ft 10 in
4 ft 3 in
5.16 m
1.53 m
1.47 m
1.30 m
Forward Cabin Door
Width
Height
2 ft 0 in
4 ft 5 in
0.61 m
1.35 m
Cargo Door
Width
Height
4 ft 5 in
4 ft 4 in
1.35 m
1.32 m
Overwing Emergency Exit
Width
Height
1 ft 6 in
2 ft 2 in
0.46 m
0.66 m
Compartment Volume
Baggage
Cabin*
40 ft3
330 ft3
1.13 m3
9.34 m3
Total Quantity
Consumption
Normal
Maximum
3.6 US gal
13.6 liters
0.2 lb/h
0.3 lb/h
0.1 kg/h
0.14 kg/h
Power Plant
Manufacturer
Pratt & Whitney Canada
Model
PT6A-67P
Takeoff Power
Thermodynamic
Flat-rated
1,845 ESHP (un-installed)
1,172 SHP flat rated at sea level
up to 33° C
1,745 ESHP (un-installed)
1,172 SHP flat rated at sea level
up to 26° C
Cruise Power
Thermodynamic
Flat-rated
1,745 ESHP (un-installed)
1,000 SHP flat rated at sea level
up to 28° C
TBO
3,500 hours
Manufacturer
Hartzell
Model
HC-E4A-3D/E10477SK
Type
Constant speed, fully reversing
Number of Blades
4 aluminum
Diameter
8 ft 9 in (2.67 m)
Operating Speed
1,700 rpm
Specific Loadings
Wing Loading
Power Loading
Weights
Max Ramp Weight
10,495 lb
4,760 kg
Max Takeoff Weight
10,450 lb
4,740 kg
Max Landing Weight
9,920 lb
4,500 kg
Max Zero Fuel Weight
9,040 lb
4,100 kg
Standard Empty Weight (equipped)
6,186 lb
2,806 kg
Standard Operating Weight*
6,386 lb
2,897 kg
Max Useful Load
4,309 lb
1,955 kg
Max Payload**
2,654 lb
1,204 kg
September 2007
1,540 liters
1,522 liters
Propeller
*Cockpit/cabin divider to rear pressure bulkhead
*Includes 200 lb pilot
407 US gal
402 US gal
Oil
Climb Power
Thermodynamic
Flat-rated
Cabin and Entry Dimensions
Cabin
Length*
Width
Height
Width at Floor
Fuel
Total
Total Usable
37.6 lb/ft2
183.7 kg/m2
10.45 lb/shp
4.74 kg/shp
Limitations
Airspeed
Max. Operating (Vmo/Mmo)
236 KIAS/0.48 M
Max. Flap Extended (Vfe)
15° (Takeoff)
40° (Landing)
163 KIAS
130 KIAS
Max. Landing Gear Operating (Vlo)
177 KIAS
Max. Landing Gear Extended (Vle)
236 KIAS
**Pilot not included
All information subject to change without notice
3
PC-12 TECHNICAL DESCRIPTION
continued
Altitude
Max. Operating Altitude
30,000 ft
9,144 m
Flight Crew
Minimum required flight crew is one pilot.
Cabin Altitude
at 25,000 ft
at 30,000 ft
8,000 ft
10,000 ft
2,438 m
3,048 m
Outside Air Temperature
Minimum
Maximum
Operations
The PC-12 is certified for flight under the following operations:
VFR Day and Night
IFR Day and Night
Flight into known icing conditions
– 67° F
+122° F
– 55° C
+50° C
+ 3.3 g
+ 2.0 g
– 1.32 g
– 0.00 g
Flight Load Factor
With Flaps Up
With Flaps Down
Figure 3 Ground Turning Radius
$
"
!
#
#
!
0 )6 / 4 0 / ). 4
"
NOTE: Ground turning radius using nosewheel steering,
inside brake and partial power. Pivot point is the
inboard main landing gear.
A. Wing Tip Radius
32 ft 2 in
B. Nosewheel Radius
13 ft 9 in
C. Outside Main Gear Radius 14 ft 10 in
D. Tail Radius
4
30 ft 2 in
$
9.83 m
4.2 m
4.53 m
9.2 m
All information subject to change without notice
September 2007
GENERAL
The aircraft primary structure is conven-
FUSELAGE
The fuselage is a semi-monocoque
tional aluminum alloy in sheet or extruded
design consisting of frames, Iongerons,
form. State-of-the-art corrosion proofing is
and stringers. The pressurized cockpit
used on all internal and external struc-
and cabin are contained within a fire-proof
tures. Epoxy resin-based composites are
titanium forward pressure bulkhead and an
used for secondary structures. Numerous
access panels facilitate ease of inspection
aluminum alloy aft pressure dome.
• Divider with curtain between cockpit
and cabin.
• Baggage safety barrier between cabin
and baggage area.
• Individual passenger reading lights
and air vents.
• Carpet on the entire floor area.
The engine cowling is constructed from
The cabin contains a forward cabin door
carbon fibre/nomex honeycomb sandwich
and aft cargo door located on the left fuse-
The entire aircraft is fatigue-tested
material. The engine mount is welded steel
lage side and an emergency overwing exit
to the equivalent of six lifetimes. Aircraft
tubing bolted to the firewall. The firewall is ti-
located on the right fuselage side. Aircraft
fatigue life is based on 20,000 flying hours
tanium and protected by insulation material.
avionics are cockpit panel mounted with
and maintenance.
and 27,000 landings, assuming a typical
The Standard aircraft interior includes:
some remote units located under the cabin
flight length of 45 minutes.
• Full interior lining, soundproofing and
floor. All remote avionic units are readily ac-
insulation.
cessible through quick release panels.
Figure 4 Fuselage Structure Showing Cabin and Cargo Doors
September 2007
All information subject to change without notice
5
PC-12 TECHNICAL DESCRIPTION
continued
DOORS AND EMERGENCY EXIT
defogging and anti-ice capability. Both side
The forward cabin door is located in the
windows are stretched acrylic. A separate
front left fuselage, immediately aft of the
pilot side window, also stretched acrylic,
cockpit, and is 4 ft 5 in (1.35 m) high by
can be opened to provide pilot visibility/
2 ft 0 in (0.61 m) wide. The door can be
smoke evacuation during emergencies and
opened or closed from inside or outside
can be used to provide additional cockpit
the aircraft and is secured by six locking
airflow during ground operations.
pins which can be visually checked to verify
engagement. The PASS DOOR warning light
Figure 5 Standard and Executive Seats
The cabin area contains a total of nine
dual-layer windows with four on the left
on the Crew Alerting System (CAS) display
side and five on the right side. All nine
window comes on when the forward cabin
windows are stretched acrylic with integral
door is unlocked. The door is hinged at the
sliding shades.
bottom and incorporates an integral folding
step/handrail assembly which automatically extends and retracts as the door is
opened or closed.
SEATS/RESTRAINT SYSTEMS
Standard aircraft seating includes two
crew seats and nine passenger seats. The
The structural cargo door is located in
crew seats are adjustable both fore and aft
the aft left fuselage and is 4 ft 4 in (1.32
and vertically. They can recline and provide
m) high by 4 ft 5 in (1.35 m) wide. This size
adjustable thigh and lumbar support. Arm-
accommodates standard pallets. The door
rest height is adjustable and the headrest
is opened hydraulically and closed electri-
has six position adjustments. Each crew
cally with the push of a button. The door
seat is equipped with a four-point restraint
is secured by locking pins which can be
system consisting of an adjustable lap belt
visually checked from the outside to verify
and a dual-strap inertia reel-type shoulder
engagement. The CAR DOOR warning light
harness.
on the CAS display window comes on when
Passenger seats are quickly and easily
the cargo door is unlocked. The cargo door
removed or installed. Each passenger seat
is hinged at the top and swings up out of
is equipped with a three-point restraint
the way to facilitate loading and unloading.
system consisting of an adjustable lap belt
The overwing emergency exit is located
and an inertia reel-type shoulder harness,
over the right wing and is 2 ft 2 in (0.66
reclining backrest, sliding headrest and a
m) high by 1 ft 6 in (0.46 m) wide. This exit
folding inner armrest.
contains a window and can be quickly and
easily removed when required.
Executive seats are available as an
option. Each Executive seat has leather upholstery, fore/aft and lateral tracking, swiv-
WINDOWS
A two-piece windshield and two side
eling, reclining backrest, sliding headrest
and sliding armrest. The Executive seats
windows provide cockpit visibility. Both
can be arranged in forward or aft facing
pilot and copilot windshields are laminated
configurations. Each seat is equipped with
twin-layer mineral glass with an embedded
a three-point restraint system, consisting
polyvinyl butyrol (PVB) layer. The pilot and
of an adjustable lap belt and an inertia
copilot windshield incorporates triple-zone
reel-type shoulder belt.
dual-level electric heating elements for
6
All information subject to change without notice
September 2007
EQUIPMENT AND FURNISHINGS
A divider with center curtain is installed
BAGGAGE COMPARTMENT
The baggage compartment is aft of the
WINGS
The wing is a two-piece, cantilever
between the cockpit and cabin while the in-
seating area and is accessible in flight.
structure consisting of a main spar, a rear
flight accessible baggage area is separated
The baggage area holds 40 cu ft (1.13
auxiliary spar, ribs and stringer reinforced
from the cabin by a safety barrier net.
cu m) of baggage weighing up to 400 lb
skin. The wing root is a NASA LS(1)-417
(181 kg). This increases to 80 cu ft (2.26
Mod design and the wing tip incorporates a
a forward closet with private lavatory en-
cu m) of volume and 500 lb (227 kg) with
winglet. Each wing contains an integral fuel
closure, two forward refreshment cabinets
incorporation of a larger baggage net
tank, aileron, flap, deicing boots, and main
and LED cabin illumination. A stereo/CD
(SB25-010).
landing gear and is attached to the cabin
The optional Executive interior includes
player system with six headphones and
jacks is an additional option.
CABIN VOLUME
The cabin volume is 330 cu ft (9.34 cu
In addition, the baggage compartment
can be increased in the Combi configura-
The ailerons are conventional aluminum
tion by removing up to five rear seats of
construction with a single spar and ribs.
the standard seating configuration and
They are mass balanced and the aileron/
reinstalling the safety barrier at frame 24.
wing gap is sealed. Each wing incorporates
meters) measured from the cockpit/cabin
divider back to the rear pressure bulkhead.
carry-through spar attach fittings.
an electrically actuated Fowler flap of con-
EMPENNAGE
Installation of air conditioning does not
The tail assembly is a T-tail configura-
infringe on the 330 cu ft cabin volume.
tion. The vertical and horizontal stabilizer
ventional aluminum construction with flap
tracks and support arms.
Flush mounted, pneumatically operated
assemblies are conventional aluminum
deicing boots are attached to the leading
(5.16 m) from the cockpit/cabin divider
alloy construction of spars, ribs, stringers
edge of each wing. The winglets incorpo-
to the rear pressure bulkhead. Cabin floor
and skin panels.
rate position/strobe light assemblies and
The cabin floor measures 16 ft 11 in
width is 4 ft 3 in (1.3 m).
The horizontal tail assembly consists of a
are constructed of carbon laminate with
A safety barrier net can be installed be-
trimmable stabilizer incorporating a separate
fiberglass layers and metal strips for light-
hind the overwing emergency exit at frame
two-piece elevator, deicing boots and redun-
ning protection.
24, in front of the cargo door at frame 27
dant load paths and attach points (fail-safe
and in front of the baggage compartment
design). The dorsal and ventral fin fairings
at frame 34. All barriers are secured by at
are a kevlar honeycomb sandwich covered
least ten separate cabin attach points.
with a fiberglass ply to resist moisture.
September 2007
All information subject to change without notice
7
PC-12 TECHNICAL DESCRIPTION
Figure 6 Example Cockpit Arrangement
1
10
4
5
6
7
8
9
11
12
13
3
2
28
27
16
1
17
1
6
26
1
8
18
19
19
5
25
20
0
21
1
4
24
22
23
8
All information subject to change without notice
September 2007
15
1. Parking Brake Handle
2. ECS Side Air Outlet
3. Emergency Standby Instrument System (ESIS)
4. Primary Flight Display (PFD)
5. Master Caution and Warning Lights
6. Audio/Marker Panel
7. PFD & Radio Control Panel
8. Situation Awareness MFD
13
9. Autopilot Control Panel
10. Overhead Electrical Control Panel
14
11. Optional Co Pilot PFD & Radio Control Panel
12. Optional Co Pilot Audio/Marker Panel
13. Master Caution and Warning Lights
2
14. Optional Co Pilot PFD
15. Control Wheel
16. ACS and CPCS Control Switches
17. Systems MFD
18. Multi Function Controller
19. Display Reversionary Control Panel
20. Flap/Trim Interrupt, ALT STAB Trim Switches
21. Flap Selector
22. Condition Lever
23. Cockpit/Cabin Lighting Controls
24. Manual Override Fuel Control
25. ELT Remote Control Panel
26. Power Control Lever
27. Ice Protection Switches
28. Control Wheel
15
September 2007
All information subject to change without notice
9
PC-12 TECHNICAL DESCRIPTION
continued
COCKPIT ARRANGEMENT
The cockpit is designed for single pilot
POWER PLANT
control and audio marker panel.
The overhead panel contains all electri-
The PC-12 is powered by the Pratt
reduced workload operations. The left side
cal power management controls, indica-
& Whitney PT6A-67P, reverse flow, free
of the instrument panel contains the pilots
tors, engine start and ignition switches,
turbine engine. The five-stage compressor
primary flight display (PFD), landing gear
fuel pump switches, and external lighting
section consists of four axial stages and
controls and electronic stand-by instru-
switches.
a single centrifugal stage. A compression
The left cockpit side-wall contains the
mentation.
The central panel and console area
left front and rear circuit breaker panels,
ratio of 12 to 1 is achieved as air exits the
centrifugal compressor.
houses two multi function displays (MFD’s)
hand microphone, PARK BRK handle,
the multi function controller and the power
oxygen mask/mic switch, oxygen pressure
of a single-stage compressor turbine and
control lever (PCL). It also contains the flap
gauge and left-hand power junction box.
a two-stage power turbine. The two-stage
controls, condition lever, manual override
The right cockpit side-wall contains the
lever, emergency gear extension and fuel
right front and rear circuit breaker panels,
shut-off valves.
co-pilot mask/mic com switch, right-hand
The three-stage turbine section consists
power turbine drives the propeller through
the propeller reduction gearbox.
Engine power is controlled by the
The right side of the instrument panel
power junction box and hand microphone.
POWER CONTROL and CONDITION levers,
contains the optional co-pilots primary flight
Refer to Figure 6, Example Cockpit
located on the center console. The POWER
display along with the associated radio
Arrangement, for detailed cockpit layout.
CONTROL lever controls engine power
Figure 7 PT6A-67P Engine
Fuel Manifold
Adapter and Nozzle
Assembly
Propeller
Governor
Mounting
Pad
Exhaust
Duct
Turbine Rotor
Assembly
Accessory
Gearbox
Combustion
Chamber Liner
Gas Generator
Case
Oil
Tank
Bleed
Adapter
Propeller
Shaft
Magnetic Chip
Detector
Propeller
Reduction
Gearbox
10
Fuel Drain
Valve
Fuel Flow Divider
and Purge Valve
Fuel
Drain
Centrifugal
Impeller
Magnetic Chip
Detector
Compressor
Rotor Assembly
Starter/Generator 2
Drive Pad
Generator 1
Drive Pad
All information subject to change without notice
September 2007
during ground, flight and propeller reverse
Ignition is provided by an ignition exciter
operations. The CONDITION lever sets the
and two spark igniter plugs. A two-position,
• Standby power bus
fuel control unit operating condition for
ON/AUTO, IGNITION switch controls engine
• Emergency Power Supply (EPS) bus
operation in fuel cut-off/feather, ground
ignition. When set to ON, ignition will oper-
• Hot battery bus
or flight modes. In case of power control
ate continuously.
• Avionics 1 bus
system malfunction, the MANUAL OVERRIDE lever allows the pilot to control engine
power output through manual manipulation
of the fuel control unit.
• Essential bus
Electrical power from Generator 2 is
PROPELLER
The propeller assembly is a constant
distributed through the:
• Generator 2 bus
speed, fully feathering, reversible unit.
• Non-Essential bus
The computer controlled Engine
Nominal propeller rpm during all flight
• Cabin bus
Information Display, shown on the PFD,
operations is 1,700 rpm and propeller
• Avionics 2 bus
shows engine torque, turbine temperature,
pitch is regulated by engine oil through the
• Main bus
gas generator speed, oil temperature and
propeller governor.
Generator 2 is also the engine starter
pressure, fuel flow, and propeller speed.
The propeller hub and four propeller
The Engine Condition Monitoring System
blades are aluminum. Each propeller
(ECMS) is displayed on the MFD.
blade incorporates an electric deicing
Outside air is ducted to the engine inlet
through the air induction system. A circular
cross-section intake under the aircraft
motor. System status is displayed on the
lower MFD and the CAS.
Under normal power generation
boot. The propeller diameter is 8 ft 9 in
distribution system operating conditions,
(2.67 m).
the systems and circuits powered from
The propeller blades will be set to mini-
Generator 1 are designated channel 1.
nose directs airflow past the inertial
mum pitch and reverse thrust when the
The systems and circuits powered from
separator to the engine inlet. The inertial
POWER CONTROL lever is in the GROUND
Generator 2 are designated channel. 2.
separator, electrically controlled by a cock-
range. The propeller is fully reversing with
The channels operate independently. The
pit switch, separates moisture and foreign
operation in the Beta range permitted dur-
only connection is through a bus tie in the
objects from incoming air to avoid engine
ing ground operations only.
event of component failures. Automatic
intake icing and damage.
The engine air inlet lip is continuously
heated by engine exhaust gas during engine operation to prevent ice accumulation.
An oil filler neck with quantity dipstick
switching and load shedding take place
ELECTRICAL
General
The PC-12 incorporates five independent
power generation sources consisting of:
should component failures occur.
Should either the Generator 1 or Generator 2 fail, the control relays in the Power
Generation Distribution System automati-
and cap are located on top of the acces-
• Generator 1 – 28V, 300A generator
cally change and connect the remaining
sory gearbox. The OIL QTY warning light,
• Generator 2 – 28V, 300A starter/
generator and both batteries to the Power
on the Crew Alerting System (CAS) display
generator
and Secondary Power Lines. A caution will
panel, comes on when oil quantity is low
• Battery 1/Battery 2 – 24V, 42Ah
display on the Crew Alert System (CAS)
and the engine is shut down.
• Emergency Power Supply (EPS) – 24V,
window.
The starter/generator unit is a starter
until engine speed (Ng) stabilizes at 50%.
Above 50% Ng the starter/generator functions as a 300 amp generator.
September 2007
5Ah lead-acid battery
Electrical power from Generator 1 is
distributed through the:
• Generator 1 bus
All information subject to change without notice
Battery 1 and Battery 2 are installed
in the rear fuselage. Battery 1 provides
power to maintain the essential systems
during engine start. Battery 2 provides the
11
PC-12 TECHNICAL DESCRIPTION
continued
power for engine start. Should both gen-
The external 28VDC power can be used
There is a Battery and External Power
erators fail, the batteries supply power to
for engine starting, ground operation of the
Junction Box (BEPJB) containing compo-
the essential electrical systems after auto-
optional air conditioning system, ground
nents for the batteries, external power
matic load shedding for a maximum range
operation of cabin heaters or maintenance
functions, hot battery bus and associ-
glide and one attemped engine start.
requirements.
ated circuit breakers. It also contains
Junction Boxes
the necessary components to permit
In the event of a total power loss, the
There are two Power Junction Boxes
optional nickel cadmium battery instal-
provide sufficient power through the EPS
(PJB), one for each generator. Generator 1
lation. The BEPJB is located in the rear
bus to the backup systems for 30 minutes.
PJB is located on the cockpit lower left wall.
fuselage. There is also a Relay Module
Emergency Power Supply (EPS) battery will
On the ground, the DC system can
Generator 2 PJB is located on the cockpit
Panel (RMP) for power Channel 1 and 2
be powered by an external power source
lower right wall. They contain the principal
which contains terminal blocks and re-
connected under the rear fuselage of the
contactors, relays and other circuit protec-
lays and is located under the cabin floor
PC-12 on the left side.
tion devices.
on the left and right sides.
Figure 8 Electrical Power System
12
All information subject to change without notice
September 2007
Circuit Breakers
ity of approximately 201 US gal (761 liters).
Circuits supplied from the Bus Bars
Fuel is supplied to the engine in excess
have circuit breakers on color coded pan-
of that required for all ground and flight
els. These panels are located on both the
operations. The transfer and delivery of
left and right cockpit walls.
fuel is achieved through a triple-redundant
Controls and Indicators — Overhead Panel
configuration consisting of a motive flow
The electrical system is controlled from
system, boost pumps and engine driven
the Electrical Power Management section
pumps. Fuel balancing is managed by auto-
of the overhead control panel. The panel
matic activation of the boost pumps.
has controls for the:
The distribution system transfers fuel
• Avionics busses (AV1 and AV2)
between left and right wing tanks and deliv-
• Generators (GEN1 and GEN2)
ers fuel to the engine through transfer-and-
• Batteries (BAT1 and BAT2)
delivery ejector pumps, shutoff valves, a
• External power (EXT PWR)
fuel filter, an air separator, and a low pres-
• Standby bus (STBY BUS)
sure engine driven pump. Electric boost
• Cabin bus (cabin bus)
pumps provide fuel for engine starting and
• Master power (master power)
are used as a fuel delivery backup.
• Emergency Power System (EPS)
• Bus Tie (bus tie)
Fuel delivery pressure is initially
provided by electrically operated boost
pumps while motive flow ejector pumps
FUEL
are the main source of fuel flow during
Total fuel capacity is 407 US gal (1,540
normal operation. A transfer ejector pump
liters) with total usable fuel being 402 US
transfers fuel from the main fuel tank to
gal (1,522 liters). Fuel quantity is deter-
the collector tank.
mined by a fuel indication system, fuel flow
Each wing tank contains four capaci-
and a low-level warning system. Fuel quan-
tance-type fuel quantity probes that are
tity is displayed in the Fuel System Status
connected to the fuel computer part of
Window of the systems MFD and refueling
the Fuel Control and Monitoring System
is accomplished using overwing filler caps.
(FCMU).
The firewall fuel shutoff valve is me-
Total fuel quantity used and fuel
chanically connected to the FUEL SHUTOFF
quantity remaining are displayed in the fuel
handle in the aft end of the center console
window of the systems MFD. A fuel quantity
and is used to shut off fuel flow to the
reset function is available so that reference
engine in an emergency. A maintenance
fuel quantities can be reset.
shutoff valve is used to isolate the fuel
system during system maintenance.
The fuel system consists of storage, distribution and indication/warning systems.
Fuel is contained in two integral wing
tanks. Each wing has a usable fuel capac-
September 2007
The CAS window of the systems MFD
contains five fuel system-related cautions:
Amber Caution
LH + RH Fuel Low
less than 20 US gal (75
liters)
Fuel Pressure Low
The fuel system pressure
less than 2 psi (0.14 bar)
Fuel Balance Fault
FCMU automatic fuel
balancing is not successful
Fuel Imbalance
(on ground)
A fuel imbalance of more
than 178 lbs. between LH
and RH fuel quantity. Takeoff
is prohibited until balanced.
The CAS window of the systems MFD
displays the following status for the fuel
system, on the ground:
WHITE
FCMU Fault
The FCMU has detected an
internal fault
Low Lvl Sense Fault
The FCMU has detected a
fault with fuel low level
sensing
FLIGHT CONTROLS
The flight control system is conventional, using push-pull rods and carbon steel
cables connected to the pilot and copilot
control wheels and rudder pedals. Internal
gust locks are set from the cockpit.
The ailerons are connected to cockpit
control wheels by cables in the fuselage
which are connected to push-pull rods in
the wings. Both ailerons are mass balanced and include a flettner tab on the
outboard side. The flettner tab of the left
aileron is electrically trimable from both
control wheels. Roll trim is pilot or copilot
controlled using a control-wheel-mounted
trim switch.
The elevator is a two-piece unit attached to the horizontal stabilizer and is
connected to the cockpit control wheels
Fuel Quantity Fault
The FCMU is unable to determine fuel quanity
by cables.
LH Fuel Low
RH Fuel Low
The fuel quantity in left, right
or both tank(s) has reached
der pedals by cables. Both pilot and copilot
All information subject to change without notice
13
The rudder is connected to cockpit rudrudder pedals are adjustable fore and aft.
PC-12 TECHNICAL DESCRIPTION
continued
Electric triple-trim systems assist the
pilot in controlling the aircraft around the
Figure 9 Fuel System
pitch, roll and yaw axes. A three-axis tripletrim position indicator is located in the
TRIM window of the systems MFD. It shows
individual trim position of the ailerons,
horizontal stabilizer and rudder. If an emergency condition arises, all trim systems
can be disconnected by pressing the trim
interrupt switch.
Pitch and roll trim are controlled by
a switch on the outboard horn of each
control wheel. Rudder trim is controlled
by a switch incorporated in the POWER
CONTROL lever.
Pitch trim is accomplished by an
electrically controlled actuator connected
to the moveable horizontal stabilizer. The
secondary trim motor, installed in the
same actuator, is controlled by the autopilot and can also be used as a backup system (alternate stabilizer trim) by the pilot.
Alternate pitch trim can be accomplished
by pressing the Alternate Stab Trim switch
in the desired direction. The color of the
symbolic aircraft views of trim positions
for aileron, rudder and horizontal stabilizer
change based on trim position. An invalid
trim status will show with an amber cross.
The aircraft symbols change to green
when each trim position is correctly set
for takeoff.
The aircraft incorporates Fowler flaps
that have four primary positions: 0° (up),
15° (takeoff), 30° (short takeoff) and
40° (landing). The flaps are electrically
operated by a handle located to the
right of the power controls on the center
14
All information subject to change without notice
September 2007
console. A flap position indicator is located
Figure 10 Landing Gear System
in the FLAP window of the systems MFD. A
single electric drive unit, connected to four
screw actuators by flex shafts, drives the
flaps into the selected position.
Flap movement will stop automatically
if out-of-limit conditions are detected. The
CAS Flaps caution will then come on. If the
airspeed goes above the maximum limit for
the current flap setting, the Flight Alerting
System (FAS) will initiate an “Overspeed”
warning on the PFD and a “Speed” voice
callout will be heard.
A flap interrupt switch, on the center
console, allows the pilot to interrupt flap
operation if the main flap control unit
malfunctions. Once this emergency system
is actuated, the normal control system is
locked out to prevent any uncommanded
inputs.
LANDING GEAR
The landing gear is a conventional
tricycle arrangement capable of operations
on prepared, unpaved runways. The landing
gear is extended, retracted and held in
the retracted position by hydraulic pressure produced by an electrically-powered
hydraulic pump. No mechanical up locks
are required.
A nitrogen-charged accumulator is
used to maintain the hydraulic pressure
necessary to hold the landing gear in the
retracted position following a hydraulic
failure. If required, the landing gear can be
lowered manually through a combination of
free falling and the emergency landing gear
hand pump.
September 2007
All information subject to change without notice
15
PC-12 TECHNICAL DESCRIPTION
continued
Nosewheel steering is by rudder pedals
linked to the nosewheel. Differential braking contributes to nosewheel steering. Use
of rudder pedals only will turn the nosewheel ± 12° from center while differential
wheel well. Each main gear retracts inward
message on the PFD when the following
into the main gear wheel well.
Cockpit controls and indicators consist
• A LANDING GEAR control handle,
located on the lower center section
center.
of the instrument panel, to initiate
The tires are a low-pressure type that
proved fields.
Aircraft braking is controlled by toe ped-
conditions are met:
• flaps set to more than 15°; or
of the following:
braking will turn the nosewheel ± 60° from
allows operations from grass and unim-
The CAS will initiate a Gear warning
extension or retraction of the landing
• airspeed less than 130 KIAS; and
• power setting less than 10% of maximum torque.
With aircraft weight on wheels, a
solenoid locks the LANDING GEAR control
gear.
• An emergency landing gear hand
handle in the down position to prevent
pump, located in the center console,
inadvertent landing gear retraction during
als operating brake assemblies attached to
can be used to manually lock
ground operations. An emergency hydraulic
the left and right main landing gear. Propel-
the landing gear in the down position.
fluid line provides hydraulic fluid during
ler reversing is used to enhance aircraft
braking and aircraft ground operations.
The brake system is separate and independent from the aircraft hydraulic system.
A separate brake master cylinder is mechanically connected to each toe pedal.
A parking brake valve, connected to the
cockpit PARKING BRK handle, secures the
• Landing gear position icons (one for
emergency landing gear hand pump, in the
systems MFD indicate landing gear
aft center console, can be used to lower
position and status.
and lock the landing gear in the extended
position. The landing gear can be retracted
The three individual landing gear
position icons, in the GEAR window of the
normally if the landing gear was manually
systems MFD, can show gear displays for
extended for training purposes.
the following conditions:
HYDRAULIC POWER
brakes when the aircraft is parked.
The nose gear is a hydraulic-fluid and
nitrogen-filled shock strut. Nose gear doors
are spring loaded to the open position and
are mechanically closed during nose gear
retraction. The nose gear is locked in the
extended position by the folding strut being
put in an over-center position. The nose
gear retracts rearward into the nosewheel
well and is completely enclosed by the gear
doors when retracted.
Condition of left main
gear, right main gear,
and nose gear
Color
and Font
State is
‘undetermined’
Amber cross on
black
background
State is
‘Gear Up’ normal
White UP with
white box
outline
UP
State is ‘Gear Up’
(after 20 sec. with
flaps up)
Grey UP with
grey box outline
UP
State is
‘Gear Up’ warning
White UP
in red box
State is ‘Gear Down’
Black DN with
green
background
The main landing gear is a trailing
link design. A hydraulic fluid and nitrogen
filled shock strut connects the trailing link
to the main leg hinge point. Each main
gear actuator incorporates a mechanical
down-lock and a gear down and locked
micro-switch. The main landing gear door
is a single unit attached to the main gear
State is
‘Gear in Transit’
White hatched
lines with black
background
State is ‘Gear in
Transit Warning’
White hatched
lines with red
background
Gear
Display
The hydraulic system is rated at 3,000
psi (207 bar) and its only function is to extend and retract the landing gear. Maintaining hydraulic system pressure is completely
UP
automatic and requires no pilot inputs.
Hydraulic pressure is supplied by a
hydraulic power package. The hydraulic
power package incorporates an electricallypowered variable displacement motor/
pump, a reservoir with sight gauge and a
low pressure filter.
DN
Hydraulic fluid level in the reservoir
can be checked visually using the sight
gauge. Hydraulic power package electric
motor/pump operation is controlled by the
pressure limit switch and the motor overtemperature switch.
A nitrogen-charged accumulator is used
to minimize system pressure loss, allow for
leg and the outside edge of the main gear
16
emergency landing gear extension. The
each gear) in the GEAR window of the
All information subject to change without notice
September 2007
thermal expansion, and maintain system
any electrical power loss will not depressur-
with built-in microphones. The passenger
pressure in case of hydraulic motor/pump
ize the cabin. The system has independent
oxygen masks are constant-flow types and
failure. Servicing is done through a nitro-
safety valves that open if the cabin differ-
are stored within easy reach of all seats.
gen-charge valve located in the hydraulic
ential pressure exceeds 6.35 psi (0.43 bar)
servicing bay.
or if a negative cabin pressure differential
the pilot on a gauge in the left cockpit side
exists, such as cabin altitude exceeding air-
panel, forward of the Test Panel.
The CAS will show a hydraulic caution
if the hydraulic system fluid pressure falls
craft altitude. The pressurization warning
below 1,800 psi (124 bar).
in the CAS window on the system MFD, will
come on to indicate ECS malfunctions.
ENVIRONMENTAL CONTROL
Primary heating and cooling for cockpit
Oxygen system quantity is displayed to
FIRE PROTECTION
The engine fire detection circuit is a
continuous-loop corrision-resistant steel
The environmental control system (ECS)
and cabin is by air bled from the engine
capillary tube that is routed throughout the
provides aircraft heating, cooling and pres-
compressor and cooled to the required
engine bay. The system activates at 932°
surization. ECS control is automatic, follow-
temperature by the air-cycle cooling pack,
F (500° C) for localized fire conditions and
ing pilot inputs, with manual operation as
located in the engine bay. This bleed-air
at 675° F (357° C) for general fire condi-
a backup. The cockpit and cabin heating
can be cooled to any temperature between
tions. The CAS warning “Engine Fire” red
and cooling requirements are individu-
37° F (2° C) and 194° F (90° C) and is set
will illuminate.
ally set and controlled using the control
using controls on the ECS control panel.
panel and the environment window of the
Engine bleed-air is routed through a heat
systems MFD.
exchanger where it is precooled by ambient
Aircraft pressurization is achieved by
using pre-conditioned engine bleed-air.
A maximum cabin pressure differential
air and delivered to the cockpit and cabin
by the air cycle cooling pack.
Additional aircraft cooling can be
A portable halon-type fire extinguisher,
in the cockpit area, is standard equipment.
ICE AND RAIN PROTECTION
The aircraft is certified for flight into
known icing conditions and is equipped
of 5.8 psi (0.40 bar) provides a 10,000
provided by an optional electrically driven
with icing protection on the wing leading
ft (3,050 m) cabin altitude at 30,000
air conditioning system. This system can
edges, horizontal stabilizer leading edge,
ft (9,150 m). Cabin pressurization is
be automatically controlled based on the
windshield, engine inlet, propeller blades,
controlled by the Cabin Pressurization
temperature demands as set on the sys-
pitot-static and AOA probes.
Control System (CPCS) control panel and
tems MFD. Cooling is provided to cockpit
Inflatable deicing boots are installed
monitored based on the inputs defined on
and cabin areas through three separate
on the leading edges of the wings and
the systems MFD. The CPCS will maintain
evaporators and individual passenger over-
horizontal stabilizer. All deicing boots have
cabin altitude at the selected level and
head outlets. The aircraft can be precooled
suction applied during engine operation,
will increase or decrease cabin altitude at
when the engine is not operating by power-
regulated by deice boot-ejector valves, to
any set rate. Manual control switches are
ing the vapor cycle cooling system from the
maintain the aerodynamic profile when
located at the cross bar panel.
external electrical DC power receptacle.
not in use. The wing boots are separated
into inboard, outboard, upper and lower
A dump valve, controlled by the weighton-wheels switch and the condition lever
OXYGEN
sections. The horizontal stabilizer boots are
position, will prevent inadvertent cabin
An emergency oxygen system is in-
pressurization with the aircraft on the
stalled which provides oxygen to two crew
ground. This dump valve can also be
and up to ten cabin outlets. The system
by the pilot to inspect the left wing leading
opened by the pilot if conditions warrant
is certified for use up to 30,000 ft (9,150
edge for ice.
the immediate dumping of cabin pressure.
m). Both pilot and copilot oxygen masks
Deicing boot inflation is controlled by
This dump valve is spring-loaded closed so
are quick-donning diluter-demand types
a boot cycle timer. Inflation cycles of one
September 2007
All information subject to change without notice
one-piece sections.
A wing-ice inspection light can be used
17
PC-12 TECHNICAL DESCRIPTION
continued
or three minutes are selectable by the
whenever the engine is operating, by routing
with the engine running. The PROPELLER
pilot, using the BOOTS switch on the ICE
engine exhaust gas through the inlet lip. An
advisory on the ice protection portion of
PROTECTION control panel. A green BOOTS
inertial separator is built into the aft inlet
the systems MFD will come on when the
advisory is shown in the ICE PROTECTION
area and is opened or closed by the INERT
propeller deice system is on.
window of the systems MFD.
SEP switch on the ICE PROTECTION panel.
The pitot-static and AOA probes are
The pilot and copilot windshield is
The INERT SEP green advisory, on the ice
deiced using integral electric heating ele-
deiced using integral electrically heated
protection portion of the systems MFD,
ments. These elements are controlled by
elements. The pilot windshield can be
will illuminate when the inertial separator
the PROBES switch on the ice protection
deiced by setting the LH WSHLD switch on
bypass system is activated.
control panel. With the PROBES switch
the ice protection portion of the pilots sub
The propeller blades are deiced by
ON, all pitot-static and AOA probes will be
panel to “heavy” or “light.” The copilot’s
electrically-heated boots attached to each
windshield can be deiced by setting the
blade. The boots are controlled by the PRO-
RH WSHLD switch on the pilot’s side sub
PELLER switch on the pilot side right-hand
display panel will come on when a deicing
panel to “heavy” or “light.”
sub panel. Propeller deicing is activated
malfunction is detected in the AOA deice
when the PROPELLER switch is set to ON
system. The “Pitot heat 1” and “Pitot heat
The engine inlet is continuously deiced,
heated.
The AOA deice caution light on the CAS
2” warning will illuminate when a malfunction has occurred in either the Pitot 1 or
Figure 11 Brake System
Pitot 2 systems, respectively.
BRAKE FLUID
RESERVOIR
LIGHTING
Internal
Cockpit lighting consists of internally lit
LEFT
R IG H T
LE F T
cockpit displays, controllers, switch pan-
R IG H T
els, instrument panel and circuit breaker
PILOT BRAKE/
MASTER CYLINDER
ASSEMBLIES
COPILOT BRAKE/
MASTER CYLINDER
ASSEMBLIES
panel mounted floodlights, map lights,
and a dome light. All cockpit lighting has
adjustable intensity levels to compensate
for day and night operations.
LEFT
SHUTTLE
VALVE
RIGHT
SHUTTLE
VALVE
SET
The cabin area is lit by ceiling lighting.
TO PARKING
BRAKE T-HANDLE
PARKING BRAKE VALVE
(SHOWN WITH PARKING
BRAKE NOT SET)
NOT SET
SET
TO LEFT BRAKE
ASSEMBLY
TO RIGHT BRAKE
ASSEMBLY
NOT SET
PARKING BRAKE VALVE
(SHOWN WITH PARKING
BRAKE SET)
KEY
- RESERVOIR BRAKE FLUID
- LEFT BRAKE SYSTEM
- RIGHT BRAKE SYSTEM
18
All information subject to change without notice
September 2007
A vestibule light illuminates the airstair.
Caution attention light and generate the ap-
action. A cyan advisory CAS message
There is also a baggage area light. Read-
propriate Crew Alerting System (CAS) alert.
will be displayed in reverse (cyan
ing lights are installed in the cabin ceiling
The CAS messages are displayed in the CAS
background) for 5 seconds. After 5
panels at each passenger seating location.
window of the systems MFD.
seconds they will show in the cyan
A separate dome light is installed in the
The CAS messages have four levels:
lavatory area.
• Warning (red) indicates a condition
External
External lighting consists of a LED posi-
advisory color.
• Status (white) are only displayed on
that requires an immediate corrective
the ground in white text and indicate a
action by the pilot. A red warning CAS
maintenance action is required.
tion light and strobe light unit mounted in
message will be displayed in reverse
each winglet tip, a white LED position light
(red background) until acknowledged
mounted in the tail, a 250 watt landing
by pressing the WARNING attention
light mounted to each main landing gear,
light. After which the CAS warning
tion, an audible tone, and a stick shaker
a 150 watt taxi light mounted to the nose
message text will be shown in the red
and stick pusher system.
landing gear, a wing deicing inspection
warning color. Some warnings also
light mounted in the left fuselage, one
include a voice call out.
recognition light mounted below each wing
• Caution (amber) indicates a condition
STALL WARNING
Stall warning is provided by PFD indica-
An angle-of-attack transmitter vane is located in the outboard leading edge of each
wing and each vane is electrically heated
and red flashing beacons located on the
that requires a pilot’s attention but not
for icing protection. The electrical output
tail and belly areas of the aircraft. The rec-
an immediate reaction. An amber cau-
from each transmitter vane is used by a
ognition lights and logo light are optional.
tion CAS message will be displayed in
processing unit to control the stall warning
Dual navigation lights are standard.
reverse (amber background) until ac-
tone and stick shaker/pusher system.
knowledged by pressing the CAUTION
CREW ALERTING SYSTEM
When approaching a stall condition,
attention light. After which the CAS
both the stall warning tone and stick
caution message text will be shown in
shaker are activated. If recovery from an
continuously monitors the interfaced air-
the amber caution color. Unacknowl-
impending stall is not initiated, the stick
craft systems and initiates the appropriate
edged reversed caution messages
pusher will automatically lower the nose by
cannot be scrolled off the CAS window.
pushing on the control wheel with a 60-65
The Monitor Warning Function (MWF)
warning, caution alert, advisory and status
messages to the crew when necessary.
• Advisory (cyan) indicates a sys-
lb force. The stick pusher will automatically
When the MWF detects an out of limits con-
tem condition, which requires pilot
disengage once a safe margin above stall
diton, it will illuminate either the Master or
awareness and may require crew
has been reached.
September 2007
All information subject to change without notice
19
PC-12 TECHNICAL DESCRIPTION
20
September 2007