Pilatus PC-24 Notes

This is my overview primer and basic study guide for the Pilatus PC-24. It tries to provide general coverage of systems, functionality, and limitations.

The information in this primer is sourced from the Pilatus PC-24 FCOM, AFM, QRH, maintenance manual, and FlightSafety Intl. Pilot Training Manual. This is not an official document and does notreflect any official procedures or opinions from any manufacturers, operators, or training providers; It is solely meant to condense about 4,200 pages of material into around 30 pages of information for study.

If you find any incorrect/inaccurate information or think of any good additions or changes that ought be made then please contact me!

🛩️ General Information

The Pilatus PC-24 is a conventional, semi-monocoque aircraft (its external skin supports part of the structural load, working together with an internal framework to provide strength). This construction method offers a balance between durability and weight efficiency.

The PC-24 is built from electronically-bonded aluminum and composite materials.

Key design features of the PC-24 include:

The PC-24 has:

The PC-24 is certificated for operation by either a single pilot or a two-person crew and includes seating for up to 10 passengers.

The PC-24 is certificated for:

Performance

Dimensions

The PC-24 can complete a 180 ° turn within 50 ft and a wheel clearance of 26 ft, 7 in.

Emergency Equipment

The PC-24 is equipped with:

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🖥️ Avionics

The current (at time of writing) Pilatus PC-24 avionics package is provided by the Honeywell Primus Epic 2.0 Advanced Cockpit Environment (ACE).

Advanced Cockpit Environment (ACE)

ACE is comprised of 2x Modular Avionics Units (MAUs) which, along with a Utility Management System (UMS), connect to and process data for various aircraft systems. Each MAU contains an Advanced Graphics Module (AGM) which is capable of generating 2 displays to one of the 4x identical Display Units (DUs).

The entire ACE system is spread across a redundant electrical system and series of separate physical components: MAU 1 is in the nose bay, MAU 2 is in the avionics rack behind the co-pilot seat, and UMS is spread across 4x Data Concentration and Processing Units (DCPUs) which control 4x Electrical Power Distribution Units (EDPUs) which each contain 41x Electronic Circuit Breakers (ECBs).

MAU + UMS Breakdown

Display Redundancy

Each DU has equal access to all primary flight and aircraft-related data provided via AGMs from MAUs + UMS. In the event of a DU failure, DUs can be reconfigured to revisionary formats to display flight-critical data. In the event of an AGM failure, DUs typically controlled from one AGM can display data from the other AGM.

Electronic Standby Instrument System (ESIS)

In addition to the 4x DUs with data provided via AGMs by MAU + UMS, there exists 1--2 self-contained Electronic Standby Instrument Systems (ESISs) which can display basic flight instrumentation in the event of a full avionics or power failure. ESIS utilizes its own magnetometer for heading data and its own pitot-static static probe for airspeed and altimeter data.

ACE Breakdown

So, ACE is comprised of:

Therefore:

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🧠 Utility Management System (UMS)

Utility Management System (UMS) is a multi-processor computing platform functioning as the “brain” of the PC-24’s systems. Together with the Modular Avionics Units (MAUs), UMS can be described as a phantom "third (or, second, if operating single-pilot) pilot". During normal operations, UMS controls and monitors almost all of the PC-24’s onboard systems. UMS is designed to improve efficiency, reliability, and pilot workload management by consolidating the operation and supervision of various systems into a single interface. UMS is not flight-critical, meaning it's not essential for safe flight and its failure does not affect core flight systems. Systems controlled by the UMS can often be manually operated in case of failure.

The UMS manages and monitors:

UMS is comprised of 4x dual-channel Data Concentration and Processing Unit (DCPUs) spread throughout the aircraft. The DCPUs can be considered “lobes” of UMS’s “brain”. Each DCPU is powered by either the Left Electrical Bus, Right Electrical Bus, or Essential Electrical Bus.

In the event of power loss or disruption, DCPUs are able to utilize power from different electrical buses through various Cross-Tie Contacts (XCs). In the event of total generated power loss only DCPU 1 is powered via the Essential Bus for up to 60 min.

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🎛️ Flight Instruments

Flight Instrument Components

Raw flight instrumentation data is collected by the following components:

Processing and Display

The system uses a left/right split where MAU 1 processes and delivers left-side data and MAU 2 processes and delivers right-side data: each PFD displays airspeed, altitude, vertical speed, and AOA information based on the data from their on-side sensors.

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⚡ Electrical System

The Pilatus PC-24 operates on a 28 v DC electrical system (max. 29.5 v DC) with 2x power-lines in a left/right split.

Stable 28 v DC power is provided down the power-lines by 2x Power Conversion Units (PCUs) from 2x Starter Generator Units (SGUs) mounted underneath each engine that produce unregulated AC power.

The "Left" + Right Power-Lines are connected via 3x primary electrical busses:

(There is no real Left Bus. The concept of the "Left Bus" is an approved training concept, but it doesn't really exist. The "Left Bus" is just an extended section of the Essential Bus that has been bifurcated.)

These power-lines and busses provide electrical power to the 4x Electrical Power Distribution Units (EDPUs), 2x Batteries, and various systems connected to the electrical busses. Each EPDU contains 41 Electronic Circuit Breakers (ECBs) automatically controlled by Utility Management System (UMS).

When N2 > 53% --> SGU generates unregulated AC output --> delivered to associated PCU --> PCU converts to stable 28 v DC output --> powers respective power-line, EDPUs, electrical busses, and batteries.

Electrical Breakdown

The electrical system may be powered by a single generator via Cross-Tie Contacts (XCs). In the event of dual generator failure both Battery 1 and Battery 2 will provide emergency power to the Essential Bus for up to 60 min.

Distribution Network

Power distribution divided into Primary and Secondary systems:

Batteries

The Pilatus PC-24 has 2 batteries: Battery 1 (BAT 1), located in the nose compartment, and Battery 2 (BAT 2), located in a compartment in the right wing fairing.

In normal operation with both generators online, BAT 1 is charged from the "Left Bus" and BAT 2 is charged from the Right Bus.

When only one source of generated power is available cross-tie contacts are closed together to allow both BATs to remain charged from their respective buses by a single available power source. BAT 1 + BAT 2 will power Essential Bus for up to 60 min with total generated power loss. Beware of switching off BAT 1 during a L BUS FAIL CAS: EC2 from BAT 2 will not close to the Essential Bus, leading a total loss of Essential Bus components.

BAT 2 is the only battery utilized for engine start.

Battery Types and Limitations

The two types of batteries used on the Pilatus PC-24 are Nickel-Cadmium (NiCad) and Lithium-Ion (LiIon). LiIon is more popular. Older aircraft utilized NiCad batteries. The two battery series have different specifications and limitations.

Lithium-Ion Batteries

LiIon batteries contain a Battery Management System (BMS) that monitors, optimizes, and protects the battery from failure. BMS also contains battery heaters that help maintain proper battery temperature for normal operation.

If LiIon battery charge status indicates yellow then a maintenance procedure is required to properly recharge batteries to a usable state. If LiIon battery charge status indicates red then the batteries must be factory serviced.

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🚂 Powerplant

Williams International FJ44-4A-QPM

The Pilatus PC-24 is powered by 2x Williams International FJ44-4A-QPM turbofan engines. The FJ44-41-QPM is a medium-bypass jet engine with Quiet Power Mode (QPM) for ground power operations in lieu of dedicated auxiliary power unit. The exhaust of the FJ44-4A-QPM features a Coanda Effect that provides a + 3 ° thrust vector at takeoff power. The FJ44-4A-QPM provides the PC-24 with:

FJ44-4A-QPM Specifications

Engine is comprised of N1 (Low Pressure Rotary Group) and N2 (High Pressure Rotary Group) sections in a 4 + 2 + 2 group configuration. N1 extends throughout the entire engine, beginning at engine intake with 1x Low Pressure Compressor + 3x Intermediate Pressure Compressors and ending at engine exhaust with 2x Low Pressure Turbines. N2 comprises the center portion of the engine with a hollow shaft overlapping the N1 shaft spinning 1x High Pressure Compressor and 1x High Pressure Turbine.

[4x N1] + [2x N2] + [2x N1]

Air bypass surrounds N1 + N2, containing a combustion chamber around and behind N2’s High Pressure Compressor. The Combustion Chamber includes 1x Fuel Start Nozzle and 2x Dual Spark Igniters. Fuel and air are mixed and ignited resulting in gas expansion which is forced over N2’s High Pressure Turbine and N1’s 2x Low Pressure Turbines.

FJ44-4A-QPM Breakdown

Accessory Gear Box (AGB)

An Accessory Gear Box (AGB) is secured to the bottom of each engine. Each AGB is connected to N2 with a Tower Gearbox Assembly via which it siphons rotational energy to energize various systems. The Tower Gearbox Assembly also connects the Starter Generator Unit (SGU) to N2. The AGB provides energy for the operation of:

Fuel Control Unit (FCU)

The Fuel Control Unit (FCU) contains necessary components to deliver fuel to the engine arriving from the fuel tanks.

Starter Generator Unit (SGU)

With engine running, Starter Generator Units (SGUs) connect to N2 via the Tower Gearbox Assembly and produce unregulated AC power. This unregulated AC power is converted to stable 28 v DC power by Power Conversion Units (PCUs) for use by onboard electrical systems. SGUs + PCUs maintain nominal operation at N2 ≥ 53%.

For engine start, SGU + PCU operation is reversed: 26.4 v DC power is delivered from BAT 2 via the R PCU to the R SGU which spins up N2 via the Tower Gearbox Assembly until combustion takes over (approximately 17% N2).

Permanent Magnetic Alternator (PMA)

Within each Fuel Control Unit (FCU) exists a self-contained Permanent Magnetic Alternator (PMA). PMAs provide a constant source of independent power to each each side's respective Full Authority Digital Engine Control (FADEC) unit. PMAs begins producing power as the engine spins up (approximately N2 ≥ 80%). PMAs are not dependent on aircraft electrical power ensuring FADEC operation even if main power is lost.

Full Authority Digital Engine Control (FADEC)

Full Authority Digital Engine Control (FADEC) is a system that contains multiple components to monitor, manage, and control each engine. FADEC utilizes an array of temperature, pressure, speed, and position sensors to schedule engine speeds between ground idle and takeoff power settings as a function of throttle lever angle.

Each FADEC is primarily controlled by dual-channel Engine Control Units (ECUs) mounted on the pressure bulkhead. Each ECU is connected to a Fuel Control Unit (FCU) on the Accessory Gear Box (AGB) and additionally connected to the opposite ECU (ECU 1 + 2). Each ECU receives startup power from the Essential Bus and the Electronic Power Distribution Units (EPDUs) until nominal operation of the Permanent Magnetic Alternators.

Ignition

The ignition system is comprised of 2x Dual Spark Igniter Plugs located at the 5 and 7 o'clock positions in the combustion chamber and 2x Ignition Exciter Units (IEUs) installed midway along the upper outer engine case. Igniter plugs are energized by the exciter units via dedicated electrical leads.

Electrical power for Ignition Exciter Units is split between the Essential Bus and Right Bus (via EDPU 4):

Except during manual activation or air starts, only 1 igniter plug is utilized per engine start alternated by FADEC. Dual Spark Igniters are always on during:

Quiet Power Mode (QPM)

Quiet Power Mode (QPM) is a reduced idle setting provided by the FJ44-4A-QPM to supply electrical power to the aircraft without running the main engines at full idle or using a ground power unit. QPM on the PC-24 uses only the right engine. Use of QPM comes with some caveats:

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🌬️ Pneumatic System

The Pilatus PC-24 utilizes engine bleed air for its pneumatic system. The pneumatic system manages high pressure bleed air and moderates its pressure and temperature to provide pressurized air to:

The pneumatic system operates automatically based on various sensor inputs and functions as two independent systems in a left/right split where each system supplies air to its respective on-side systems.

Pneumatic System Components

The pneumatics system is comprised of various components, including:

Pneumatic System Breakdown

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💨 Environmental Control System (ECS)

The Environmental Control System (ECS) is an automated system managed by UMS to provide a comfortable and safe environment to the cabin with air from the pneumatic system.

Air from the pneumatic system passes through the Bleed Check Valve (BCV) to be routed to either the ECS or the Ice Protection System (IPS). ECS air off the pneumatic system is in a left/right split: left engine bleed air for the ECS goes to the cockpit, right engine bleed air for the ECS goes to the cabin. Regardless of left/right source both ECS lines fork to either pass through a Dual Heat Exchanger (DHS) via a Temperature Control System (TCS) that controls a Temperature Control Valve (TCV) or continue straight on to the cockpit/cabin.

Environmental Control System Components

Heating and Cooling Systems

When the engines are not running the Temperature Control System (TCS) controls cabin cooling through a Vapor Cycle Cooling System VCCS consisting of a condenser and compressor. After engine start, TCS controls cabin heating and cooling through 5x Electric Heaters across the cabin/cockpit floors and 2x Cooling Evaporators inside the cabin in addition to hot + cold bleed air entering the cabin from the ECS.

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🍃 Cabin Pressure Control System (CPCS)

The Cabin Pressure Control System (CPCS) maintains safe cabin environmental pressure. The system can maintain a maximum cabin pressure of 8,000 ft at FL450 at 8.78 psi.

Pressurization is maintained by allowing Environmental Control System (ECS) air to continuously flow into the cabin from the pneumatic system while modulating air discharge via an Outflow Valve.

The Outflow Valve (OV) mounted underneath the cargo compartment floor is controlled by the Utility Management System (UMS) via an Electronic Control and Monitoring Unit (ECMU) to provide a comfortable and safe cabin altitude against ECS pressurization and maintain cabin pressure to 8.78 psi. The ECMU mode incorporates a dedicated Cabin Pressure Transducer to measure cabin pressure. A spring-loaded, pressure bulkhead mounted Pressure Relief Valve (PRV) begins discharging cabin air at approximately 9.07 psi if an over-pressure state exists, while 2x spring-loaded, pressure bulkhead mounted Negative Pressure Relief Valves (NPRVs) allow ambient outside air to enter the cabin if a negative pressure state exists around - 0.3 psi.

ECMU + CPCS will issue a CABIN ALTITUDE CAS at cabin pressure altitudes ≥ 9,500 ft.

The PRV should begin to open at 9.07 psi and be fully open by 9.27 psi to maintain a maximum of 9.27 psi. The NPRVs will be fully open at -0.3 psi.

A forward-mounted Emergency Ram Air Scoop (ERAS) can be deployed to allow ambient outside air to enter an unpressurized cabin at cabin altitudes ≤ 14,500 ft. Outside air via the ERAS will not enter the cabin if cabin altitude exceeds 14,500 ft.

Electronic Control and Monitoring Unit (ECMU) power is supplied from EPDU 4, while power for the monitoring/manual channel is supplied by the Essential Electrical Bus. Cabin Pressure Control System will not function (cabin pressure will be lost) with a loss of all generated power (GEN 1 + 2 Offline).

Cabin Pressure Control System Breakdown

Oxygen Mask Deployment

The Electronic Control and Monitoring Unit (ECMU) automatic channel transmits a passenger oxygen mask deployment signal to the Mask Deployment System (MDS) at cabin altitudes of 12,850 ft (± 150 ft). If the ECMU deployment trigger fails then the Advanced Cockpit Environment (ACE) avionics system also transmits its own secondary signal to MDS at cabin altitudes of 14,800 ft (± 500 ft).

Cabin Altitude Limiting

Unless a catastrophic failure of the pressure vessel has occurred, ECMU will automatically limit cabin altitudes to ≤ 14,500 ft by closing the Outflow Valve (OV), overriding AUTO, DUMP and MAN CPCS modes.

Override of Automatic Cabin Altitude Schedules

In normal mode Cabin Pressurization Control System (CPCS) operates either in the climb or descent mode and regulates cabin altitude according to a set pressurization schedule. However, normal pressurization schedule can be overridden.

CAB LO Mode

CAB LO Mode is a control mode that allows a lower, more comfortable target cabin altitude to be selected instead of relying on the predefined climb and descent schedules normally used. During CAB LO Mode, the altitude target maintains a maximum cabin pressure differential of 8.83 psi instead of the normal 8.78 psi. CPCS can maintain Landing Field Elevation (LFE) cabin altitude of 0 MSL @ 8.83 psi until 23,350 ft.

FIELD HI Mode

FIELD HI Mode is automatically set by the CPCS when the aircraft operates at high elevation airports (e.g. normal cabin altitude is 8,000 ft @ 8.78 psi, whereas KTEX sits at 9,069 ft MSL.). FIELD HI is necessary to change the pressurization schedule of the aircraft to facilitate operations at high airfields.

FIELD HI mode will be set whenever the aircraft is landing or departing from airfields ≥ 8,300 MSL. FIELD HI mode will transition from, or transition back into, FIELD HI mode through 24,800 ft.

FIELD HI mode changes CABIN ALTITUDE CAS activation from 9,500 ft to 14,200 ft and changes the Electronic Control and Monitoring Unit's (ECMU) mask deployment signal to "descent mode" only (LFE + 850 ft ± 150 ft).

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🧊 Ice Protection System (IPS)

The Pilatus PC-24 Ice Protection System (IPS) is monitored and controlled automatically by the Utility Management System (UMS). UMS will autonomously detect the accumulation of airframe icing and automatically activate IPS. IPS also retains manual control options through UMS via physical controls and the IPS synoptic page within ACE.

Ice Protection System Components

Stall Warning Protection System (SWPS)

The Stall Warning Protection System (SWPS) is a function of IPS automatically controlled by UMS. It comprises 3 modes:

When IPS/SWPS is operating in an Ice Mode stick shaker, stick pusher, and and low-speed awareness (LSA) airspeed activation schedules are conservatively biased to provide acceptable stall protection margins. Each Ice Mode escalation increases the conservatism of SWPS biases. It takes approximately 25 sec for SWPS to transition between Ice Modes.

IPS Requirements + Considerations

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🛞 Hydraulic Power Generation System

Hydaulics on the Pilatus PC-24 are primarily utilized for Main Landing Gear (MLG) braking. The central hydraulic system is the Hydraulic Power Generation System which consists of:

This system is governed by the Utility Management System (UMS) via the Digital Antiskid Control Unit (DACU).

Hydraulic Pump actuation takes place automatically to maintain system pressure between 3,100 psi -- 2,700 psi. A high pressure relief valve is installed to prevent over-pressure. Ideally, the system will maintain Main Brake Accumulator pressure to 3,000 psi to support the Hydraulic Pump during transient periods of high demand. System pressure may drop during flight but Hydraulic Pump actuation should occur automatically before landing to return system pressure to 3,000 psi.

Digital Antiskid Control Unit (DACU)

The Digital Antiskid Control Unit is a redundant, dual-channel, computerized anti-skid and braking system governed by UMS which controls the hydraulic power generation system and independently controls the left/right brakes on each MLG. DACU provides:

Parking/Emergency Brake

In addition to the primary Hydraulic Power Generation System, the PC-24 includes a separate, independent hydraulic pressure source in the Parking/Emergency Brake. This system operates at 3,000 psi and provides braking redundancy in the event of a normal brake failure and also provides hydraulic pressure for the parking brake. Braking with the Parking/Emergency Brake system must be accomplished with the Parking/Emergency Brake handle and no DACU protection is provided. The system consists of:

UMS/DACU should automatically recharge the Parking/Emergency Brake Accumulator through the Parking/Emergency Brake Valve by running the Hydraulic Pump and opening the Parking/Emergency Brake Isolation Valve (IV) whenever Parking/Emergency Brake Accumulator pressure drops below 2,800 psi and the aircraft is stationary on the ground.

Hydraulic Notes

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⛽ Fuel System

The Pilatus PC-24 fuel system comprises 2x identical and normally independent systems. The fuel system stores fuel, gauges fuel, balances fuel, and delivers fuel to each engine. Fuel can be delivered to the PC-24 either from a single-point Refuel/Defuel System or through over-wing Gravity Refuel Points.

The fuel system comprises Wing Fuel Tanks, Fuel Pumps, Fuel Lines, and Fuel Valves.

Fuel Tanks

Each wing contains 3x Wing Fuel Tanks:

There are 4x underside Drain Valves at the lowest points of the fuel system between the Inner Cell of the Main Tank and the Collector Tank, and the Collector Tank.

Fuel System Operation

Fuel is delivered to each respective Main Tank Outer Cell through a Pressure Refueling Pipe from the Refuel/Defuel System. Fuel in the Main Tank Outer Cell progressively flows via gravity through the Main Tank Middle Cell and Main Tank Inner Cell to the Collector Tank through a series of one-way flapper valves.

A Climb Vent Line in the Main Tank Inner Cell allows for system pressure equalization via the Vent Tank's NACA Vent.

Motive Flow

The fuel returned to the engine from the Scavenge Ejector and Engine Return Motive Flow Lines create circulatory "motive flow" within the system which maintains positive pressure within in the Collector Tank. By N2 > 40% this pressurized circulatory motive flow is able to provide the engines with fuel alone and the Fuel Booster Pump turns off.

Fuel Balancing

Fuel quantity is measured from 6x capacitance-type fuel quantity probes in each wing monitored by Utility Management System (UMS). If UMS detects a fuel imbalance of 220 lb it will automatically initiate fuel balancing, pumping fuel from the higher-quantity tank to the lower-quantity tank by the higher Fuel Booster Pump through a Balance Line that feeds into the Pressure Refueling Pipe.

Automatic fuel balancing will cease with:

The maximum allowable fuel imbalance is 330 lbs (and the maximum tested imbalance is 1,100 lb).

Cross-Feed

Cross-feed is used to transfer fuel supply directly from one wing to the opposite side engine or to facilitate fuel balancing in the event of a failure of a Refuel/Defuel Solenoid Shutoff Valve. The Cross-Feed Line connects across between each respective Engine Feed Line through a remote drive motorized Cross-Feed Valve. When the Cross-Feed Valve is opened, the Fuel Booster Pump on the higher side is automatically turned on to start moving fuel.

With a failure of generated power only left Fuel Booster Pump is operable. Therefore, cross-feed is required to air-start right engine.

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🪽 Flight Control System

The flight controls on the Pilatus PC-24 can be divided into 2 groups:

Primary Flight Controls are either manually or automatically actuated via carbon steel cables, push-pull rods, actuators, and bell cranks. Some Primary Flight Control surfaces incorporate balance tabs that provide control input assistance and reduce resistance pressure. Primary stops on all Primary Flight Controls limit travel to prevent exceeding operating ranges:

Secondary Flight Controls are electronically actuated to assist the function of the Primary Flight Control System in both normal (e.g. flaps, roll assist) and emergency operations (e.g. rudder bias).

Flap System

Double-slotted flaps on each wing provide necessary lift augmentation at low speeds during takeoff and landing. The flaps are electromechanically actuated from motors within flap actuators governed by an Actuator Power and Control Module (APCM) which is controlled by Utility Management System (UMS). Monitoring sensors report position of each flap set directly to UMS. Flap settings are:

Flaps 8 º is used for normal takeoff. Flaps 33 º is used for normal landings. Flaps 15 º may be used for contaminated runway takeoffs, approach, and landing. Flaps 15 º is the maximum allowable flap setting for landing in icing conditions or with the Stall Warning Protection System + Ice Protection System in any Ice Mode > 0. Flaps should not be retracted past 8 º if ice accumulation is suspect.

The flaps system is normally powered by the Right Electrical Bus. However, in the event of a power failure, the flaps may be powered through the Essential Electrical Bus by pressing the FLAP EMER PWR pushbutton within the cockpit. It will take 2 min for FLAP EMER PWR to transiton power through the Essential Bus to the flap actuator motors.

Pitch Trim System

Pitch Trim System functions are accomplished by moving the horizontal stabilizer with a dual-channel, dual-motor Pitch Trim Actuator (PTA) through a range of + 10 º to - 5 º. Pitch trim can be manually activated or controlled by autopilot input from Modular Avionics Unit (MAU). Additionally, automatic Pitch Trim Compensation (PTC) can be activated by UMS to help guard against over-speed by commanding pitch-up at speeds > 0.785 ma.

In the event of a STAB Trim Fail CAS (failure or runaway of the primary Pitch Trim System channel or motor) the Pitch Trim System can be controlled through the secondary channel and secondary motor in a low rate via the SEC STAB TRIM controls on the center console. ENABLEing the SEC STAB TRIM controls deenergizes the control wheel stabilizer trim switches. SEC STAB TRIM ENABLE push-button will permanently deactivate the Primary Trim System and cannot be reset in flight.

Wing Spoiler System

The Wing Spoiler System is an electromechanical system consisting of 4x spoilers on each wing separated into inboard and outboard groups: Multifunction Spoilers (MFS) and Ground Spoilers (GS). Each spoiler group is actuated by 4x motorized actuators commanded by Actuator Power and Control Modules (APCMs) all governed by UMS. The spoiler system provides:

Rudder Travel Limiter System (RTLS)

Rudder Travel Limiter System (RTLS) automatically limits rudder movement range to prevent excessive yaw or side-to-side movement during flight. The RTLS sets rudder movement limitations according to flap position via a UMS-controlled rotary type electromechanical actuator attached to a cam end stop with 3 positions:

There is no RTLS protection in the event of an Essential Electrical Bus + "Left" Electrical Bus failure.

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🧯 Fire Protection System

The Pilatus PC-24 Fire Protection System comprises fire protection, fire detection, and fire extinguishing capabilities.

Fire Protection

Fire protection in is provided by the Engine Pylon and a nacelle Engine Fire Zone that comprises the region between the outermost engine case and the inside wall of the engine nacelle. The Engine Fire Zone is is encapsulated by 3 firewalls: front, rear, and pylon. This Engine Fire Zone is meant to help contain engine fires. The Engine Pylon provides physical separation and distance from the engine and the fuselage and contains its own firewall. Fuel and Bleed Air Shutoff Valves allow systems isolation of the engine from the rest of the aircraft.

Fire Detection System

Detector Loop

A fire Detector Loop of sealed, pressurized helium with a segmented, heat-activated, hydrogen-producing core material is installed around each engine within the nacelle Engine Fire Zone. A fire detect signal is sent to Utility Management System (UMS) from hydrogen responding elements when any 5 in segment of the detector reaches 510 º c, ± 40 º c, a temperature above the highest possible bleed air temperature.

Cabin Smoke Detector

The baggage compartment of the PC-24 is equipped with a multi-sensor Optical Smoke Detectordesigned to trigger a smoke detect signal to UMS within 1 min of the beginning of any smoke or fire event.

Fire Extinguishing

2x spherical, dual outlet Fire Extinguisher Bottles each containing 1.5 lb of Halon 1301 are mounted in the upper rear fuselage. These Fire Extinguisher Bottles are connected together through a Double Check Tee Valve and connected to each engine via Fire Extinguishing Plumbing terminating at Extinguisher Discharge Nozzles.

When activated, 2x electronically-activated Explosive Cartridges rupture disks within the Fire Extinguisher Bottle outlet fittings releasing 600 psi of Halon 1301 through the Fire Extinguishing Plumbing (directed to the appropriate engine via the Double Check Tee Valves) and out the Fire Discharge Nozzles.

NACA Vents on each engine's nacelle help Halon 1301 distribution throughout the nacelle Engine Fire Zone with ram air.

In normal function, dedicated Engine Isolation Switches within the cockpit isolate affected engines by closing fuel and bleed Shutoff Valves before sending a discrete arming signal to the forward Fire Extinguisher Bottle. Discharge of the forward Fire Extinguisher Bottle is available 5 sec after pressing an Engine Isolation Switch. After discharge, arming of the rear Fire Extinguisher Bottle is handled by Utility Management System (UMS) with discharge becoming available after 30 sec.

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🛟 Engine Emergency Protections

The Pilatus PC-24 includes some automatic mechanisms to assist in controllability and safety during engine failure or One Engine Inoperative (OEI) operations.

Rudder Bias (RB)

Rudder Bias (RB) is an Automated Flight Control System (AFCS) function that assists in compensating for yawing moments from OEI scenarios. RB function is fully automatic and engages when a significant difference between N1 values is observed. RB is armed whenever:

Rudder Bias disengages with manual rudder trim input, YD activation, or elimination of N1 discrepancy.

Automatic Thrust Reserve (ATR)

Automatic Thrust Reserve (ATR) is a feature within the FADEC software which automatically increases thrust output of a remaining operative engine to the 3,600 lbf ATR limit in the event of a loss of power on one engine during takeoff. ATR engages on the active channel of each respective engine's FADEC if KIAS ≥ V1 and an uncommanded drop or split of ≥ 15% N1 is detected.

ATR is automatically armed for takeoff on the ground when:

Engine Automatic Shutdown

The engine system is designed to automatically shutdown engines in events that would otherwise jeoprodize the integrity of the engines or aircraft. Engine Automatic Shutdown is accomplished by closing the fuel Solenoid Shut Off Valve (FSOV) via direct mechanical linkage or from commands issued by FADEC. Engine Automatic Shutdown will activate when: