Module 13: Aircraft Aerodynamics, Structures and Systems (B2)

Module Overview

Module 13 for the EASA Part-66 B2 licence category covers the fundamental aircraft systems that an avionics certifying staff member must understand to safely maintain and certify modern transport category aeroplanes. This module bridges the gap between theoretical aerodynamic principles and the complex, interconnected systems installed on contemporary aircraft. The syllabus encompasses electrical power generation and distribution, flight control systems (both mechanical and fly-by-wire), fuel systems, hydraulic power, environmental control systems (air conditioning and pressurisation), communication and navigation systems, and the various warning and protection systems that ensure safe operation.

For the B2 licence holder, a detailed understanding of these systems is essential, as avionics technicians are responsible for the troubleshooting, testing, and certification of the electronic and electrical components that control and monitor these systems. The knowledge level required ranges from a general overview of mechanical systems (to understand their interaction with avionics) to a detailed, component-level understanding of electronic systems, sensors, computers, and display systems.

1. Aircraft Electrical Power Systems

1.1 DC Power Generation and Distribution

1.1.1 Nominal Voltages and Standards

Aircraft DC electrical systems operate at a nominal 28 V DC. This is the standard for most general aviation and transport category aircraft. Some older light aircraft may use a 14 V DC system, but 28 V DC has become the industry standard due to its ability to deliver twice the power for the same current, reducing wiring weight.

The normal operating range for a 28 V DC system is 22 V to 30 V, with 30 V being the maximum allowable continuous voltage. Voltages above this can damage sensitive electronic equipment, while voltages below 22 V may cause relays and contactors to drop out or electronic equipment to malfunction.

1.1.2 DC Generators and Alternators

DC generators produce direct current through the rotation of an armature within a magnetic field. The output voltage is regulated by controlling the field current. A voltage regulator senses the generator output voltage and adjusts the field current to maintain a constant voltage regardless of generator speed or load. If the output voltage drops, the regulator increases field current; if it rises, the regulator decreases field current.

A reverse current relay (or reverse current cut-out) is a critical protection device in DC generator systems. It automatically disconnects the generator from the bus when the generator voltage falls below the battery voltage. Without this device, the battery would discharge through the generator, driving it as a motor and potentially causing serious damage.

Modern aircraft increasingly use DC alternators with solid-state rectification, which are lighter and more reliable than traditional DC generators.

1.1.3 Battery Systems

The aircraft battery serves multiple purposes:

Provides power for engine/APU starting
Supplies essential power when generators are not available
Acts as an emergency power source in the event of total generator failure
Provides a stabilising influence on the DC bus, absorbing voltage transients

Battery charging systems ensure the battery is charged correctly, preventing overcharging (which can cause thermal runaway and gassing) or undercharging (which reduces capacity and can cause sulphation). Modern charging systems use constant-voltage/current-limited charging profiles and include temperature compensation.

Battery contactors (isolators) allow the pilot or the electrical system to disconnect the battery from the aircraft's electrical system. This is essential for safety, allowing the battery to be isolated in the event of a fault or fire.

1.2 AC Power Generation and Distribution

1.2.1 Three-Phase AC Systems

Aircraft AC power systems use three-phase, 115/200 V AC at 400 Hz. The three phases are electrically separated by 120 degrees, providing balanced power delivery. The 400 Hz frequency is used because it allows transformers, motors, and other magnetic components to be significantly smaller and lighter than their 50/60 Hz counterparts.

The phase relationship between the three voltages is critical:

Phase A, Phase B, and Phase C are each 120° apart
The line-to-neutral voltage is 115 V AC
The line-to-line voltage is 200 V AC (115 V × √3)

1.2.2 Constant Speed Drives (CSD) and Integrated Drive Generators (IDG)

Modern aircraft engines operate at varying speeds, but AC generators must rotate at a constant speed to produce a constant 400 Hz frequency. The Constant Speed Drive (CSD) is a hydraulic-mechanical device that maintains a constant output speed regardless of input speed variations from the engine.

The Integrated Drive Generator (IDG) combines the CSD and the AC generator into a single, compact unit. The IDG is mounted directly on the engine accessory gearbox and contains:

A hydraulic constant-speed drive mechanism
An oil-cooled AC generator
Integral oil system for cooling and lubrication

1.2.3 Voltage Regulation and Generator Control

The Generator Control Unit (GCU) performs multiple functions:

Regulates generator output voltage by controlling field excitation
Monitors frequency and voltage
Controls the generator line contactor (GLC) to connect/disconnect the generator from the bus
Provides protection against over/under voltage, over/under frequency, and differential faults

1.2.4 Transformer-Rectifier Units (TRUs)

Transformer-Rectifier Units (TRUs) convert AC power to DC power. A TRU steps down the 115 V AC to approximately 28 V AC using a transformer, then rectifies it using solid-state diodes to produce 28 V DC. TRUs are used to supply the DC buses from the AC generation system, eliminating the need for separate DC generators.

1.2.5 Static Inverters

A static inverter uses solid-state electronics to convert DC power (e.g., 28 V DC) into AC power (e.g., 115 V AC, 400 Hz). Static inverters are used to supply AC power to loads that require it when the main AC generation is unavailable, or in smaller aircraft that do not have engine-driven AC generators.

1.2.6 Autotransformers

An autotransformer has a single continuous winding tapped at different points to provide voltage transformation. Unlike a conventional transformer, there is no electrical isolation between primary and secondary. Autotransformers are used in aircraft AC systems to provide a different voltage level (e.g., 115 V to 26 V) without the weight of a fully isolated transformer.

1.3 Electrical Distribution and Protection

1.3.1 Buses and Bus Tie Contactors

An electrical bus is a common distribution point that supplies power to various branch circuits. Aircraft typically have multiple buses:

Main AC buses (left and right)
Essential AC bus
DC buses (fed by TRUs)
Battery bus (directly connected to the battery)
Hot battery bus (always energised)

A Bus Tie Contactor (BTC) allows two separate electrical buses to be connected, enabling power transfer from one source to another during normal operation or in the event of a generator failure. For example, if the left generator fails, the left main bus can be fed from the right generator through the bus tie contactor.

1.3.2 Circuit Protection Devices

Circuit breakers are protective devices that interrupt excessive current flow to prevent damage to wiring and components. They are resettable and are typically located in the cockpit or an electrical equipment bay. Circuit breakers are rated by current capacity and trip characteristics.

Current limiters are protective devices that restrict current flow to a safe level. They are typically used in high-current circuits where a conventional circuit breaker would be impractical. Current limiters are often non-resettable and must be replaced after activation.

Fuses are one-time protective devices that melt when excessive current flows. They are less common in modern aircraft but may be found in some older designs or specific applications.

1.3.3 Load Shedding

Load shedding is a control function that disconnects non-essential loads to preserve power for critical systems when generation capacity is reduced. When a generator fails, the electrical system automatically sheds loads in a predetermined sequence:

First shed: galley loads, passenger entertainment
Second shed: non-essential air conditioning, some utility systems
Essential loads are never shed: flight instruments, flight controls, navigation, communication

1.3.4 External Power

The external power receptacle allows ground power units (GPUs) to supply electrical power to the aircraft, reducing the need to run the APU or engines for maintenance and ground operations. External power is typically 115/200 V AC, 400 Hz, or 28 V DC.

1.4 Emergency Power Systems

1.4.1 Ram Air Turbine (RAT)

The Ram Air Turbine (RAT) is a small turbine that deploys into the airstream in an emergency to generate hydraulic or electrical power. The RAT ensures continued operation of critical flight instruments and flight controls in the event of a total loss of engine-driven generation. The RAT is typically deployed automatically or manually and provides power for:

Essential flight instruments
Flight control actuators (in fly-by-wire aircraft)
Essential communication and navigation

1.4.2 APU Generator

The Auxiliary Power Unit (APU) generator can supply electrical power if main engine generators fail or for ground operations without external power. The APU generator is typically of similar capacity to the main engine generators and can power the entire aircraft electrical system.

1.5 Electrical System Annunciations

GEN OFF / GEN FAIL: Indicates that the generator is not supplying power to the bus, often due to a fault or manual disconnection
ELEC / MASTER WARNING: Alerts the flight crew to a significant electrical system fault requiring attention
BUS TIE OPEN: Indicates that the bus tie contactor is open, isolating the buses

2. Air Data Systems and Flight Instruments

2.1 Pitot-Static System

The pitot-static system is fundamental to the operation of air data instruments. It supplies pitot (ram) pressure and static pressure to the air data computers and instruments.

Pitot pressure: Ram air pressure measured at the pitot probe, used for airspeed computation
Static pressure: Ambient atmospheric pressure measured at static ports, used for altitude and vertical speed computation

2.1.1 Air Data Modules (ADMs)

Air Data Modules (ADMs) are smart sensors that convert pitot and static pressures into digital data for the air data computers and other systems. ADMs contain:

Pressure transducers
Analogue-to-digital converters
Microprocessors for signal processing
Digital data bus interfaces (ARINC 429, etc.)

ADMs are typically mounted close to the pitot/static probes to minimise pneumatic line length and improve accuracy.

2.1.2 Air Data Computer (ADC)

The Air Data Computer (ADC) receives pitot and static pressures and computes:

Indicated airspeed (IAS)
True airspeed (TAS)
Mach number
Pressure altitude
Vertical speed
Static air temperature (SAT) and total air temperature (TAT)

2.1.3 Pitot-Static System Leak Testing

Pitot-static system leak tests are performed to detect leaks that would cause erroneous instrument readings. A leak in the static system allows ambient pressure to enter, causing incorrect static pressure readings and thus erroneous altitude and airspeed indications.

Typical leak test limits (varies by aircraft):

Pitot line: Maximum leak rate of approximately 200 feet per minute (or equivalent in hPa) over a specified test period
Static line: Maximum leak rate of approximately 100 feet per minute (or equivalent in hPa)

The test is performed by applying a vacuum to the static system or pressure to the pitot system and monitoring the decay rate.

2.2 Angle of Attack (AoA) Systems

Angle of attack is the angle between the wing chord line and the relative airflow. The angle of attack indicator displays this value to the pilot.

The stall warning system uses AoA inputs to determine when the aircraft is approaching a stall condition. The primary sensor is an AoA vane mounted on the fuselage, which senses the local airflow direction. The stall warning computer compares the local vane angle to the aircraft's longitudinal axis to derive the actual angle of attack and triggers the warning when a predetermined threshold is exceeded.

The AoA vane must be checked for freedom of movement and correct alignment during maintenance. A stiff or misaligned vane would not move freely, failing to trigger the warning at the correct angle of attack.

2.3 Inertial Reference Systems (IRS)

2.3.1 Alignment Mode

The Inertial Reference System (IRS) must be aligned on the ground before flight to establish a reference frame. During alignment:

The system's gyros and accelerometers are initialised
The aircraft's position is entered (manually or automatically via GPS)
The system detects the Earth's rotation and gravity to establish true north and local vertical

Inertial alignment requires the gyroscopes to detect the Earth's rotation and gravity. If the gyros are faulty, alignment will fail. GPS is not required for initial alignment, though it can aid it. Modern ADIRUs can level themselves and do not require the aircraft to be perfectly level.

2.3.2 Air Data Inertial Reference Unit (ADIRU)

The ADIRU combines air data and inertial reference functions in a single unit. It provides:

Air data: airspeed, altitude, Mach, vertical speed
Inertial data: attitude, heading, acceleration, angular rates

The ADIRU provides data to multiple systems, including the EFIS displays, flight management system, autopilot, and flight control computers.

2.4 Electronic Flight Instrument System (EFIS)

2.4.1 Primary Flight Display (PFD)

The PFD displays critical flight information:

Attitude
Airspeed
Altitude
Vertical speed
Heading
Flight director commands
Navigation information

2.4.2 Symbol Generators

The symbol generator receives data from various systems and creates the visual representation (symbols, text, and graphics) on the display units. It is a core component of EFIS, converting digital data into video signals for the display.

2.4.3 Reversionary Modes

If the AHRS fails, the PFD has a reversionary mode that can use data from the Integrated Standby Instrument System (ISIS), which contains its own AHRS and air data sensors, to continue displaying critical flight information.

The ISIS is designed to be independent of the aircraft's main electrical system, with its own internal battery providing power in the event of a total electrical failure.

2.5 Flight Data Recorder (FDR)

The Flight Data Recorder (FDR) records a specific set of parameters for accident investigation. The Flight Data Acquisition Unit (FDAU) collects and digitises data from various sensors and sends it to the FDR.

Underrange and overrange checks are performed to ensure that the sensors and recording system are functioning correctly and not recording out-of-limit values. If one parameter is missing, the fault could be in the sensor, the data bus, the acquisition unit, or the FDR itself. A systematic approach starts by verifying the source data is present at the FDR input.

3. Flight Control Systems

3.1 Mechanical Flight Control Systems

3.1.1 Control Surface Balancing

Aerodynamic balancing reduces the control force required by the pilot. A trim tab deflecting opposite to the control surface (e.g., aileron) produces an aerodynamic force that helps move the control surface. This is a form of aerodynamic balancing.

3.1.2 Trim Systems

Trim tabs are used to relieve control forces in steady flight. The aileron trim tab is deflected in the opposite direction to the aileron movement to assist the pilot in moving the control surface.

3.2 Fly-by-Wire (FBW) Systems

3.2.1 Principles of Operation

In a fly-by-wire system, the pilot's side stick controller sends electrical signals to the Flight Control Computers (FCCs). There is no direct mechanical link to the control surfaces. The FCCs process pilot inputs and aircraft state data to command the hydraulic actuators that move the control surfaces.

3.2.2 Redundancy and Dissimilarity

Redundant, dissimilar FCCs are used to ensure that a single hardware or software fault cannot cause a total loss of flight control. If one computer fails, the other(s) continue to command the actuators seamlessly, maintaining normal aircraft handling. There is no mechanical reversion in a pure fly-by-wire system.

3.2.3 Feel Systems

In fly-by-wire systems, there is no direct mechanical link to the control surfaces, so the pilot feels no aerodynamic feedback. The feel system provides a synthetic force to the sidestick or yoke to give the pilot a sense of control authority.

3.3 Autopilot Systems

3.3.1 Control Wheel Steering (CWS) Mode

In CWS mode, the pilot can move the controls, and the autopilot follows, providing assistance and holding the new attitude when the pilot releases the controls. This is a "hands-on" mode where the pilot remains in the control loop.

3.3.2 Mode Control Panel (MCP)

The Mode Control Panel (MCP) is the pilot interface for the autopilot/flight director system. It contains knobs and buttons for selecting modes and setting targets (heading, altitude, speed, vertical speed, etc.).

3.3.3 Gain and Stability

Oscillation in an autopilot-controlled axis is a classic sign of excessive gain in the control loop. The autopilot over-corrects, leading to a continuous oscillation. An inoperative trim servo would cause a different symptom, such as a steady pitch offset.

3.3.4 Autopilot Disengagement

Autopilot systems rely on air data inputs (airspeed, altitude) and inertial data. Erroneous inputs can cause the autopilot to disengage to prevent unsafe control.

3.4 Yaw Damper

The yaw damper automatically moves the rudder to dampen Dutch roll, an oscillatory motion combining yaw and roll. This improves passenger comfort and aircraft stability. The yaw damper uses a yaw rate sensor to detect the oscillation and commands the rudder to counteract it.

3.5 Stall Warning Systems

The stick shaker is a mechanical device that vibrates the control column to warn the pilot of an approaching stall. It is a key part of the stall warning system.

The stall warning system primarily relies on AoA inputs. An erroneous aural warning during normal flight is typically caused by a faulty AoA vane or its signal processing.

4. Hydraulic Power Systems

4.1 System Components

4.1.1 Variable Displacement Pumps

A variable displacement pump uses a swashplate to vary the pump output. The pressure compensator senses system pressure and adjusts the pump's swashplate angle to maintain pressure within a deadband, reducing flow when demand is low. This provides:

Constant system pressure
Reduced power consumption when demand is low
Reduced heat generation

4.1.2 Pressure Relief Valves

The pressure relief valve is a safety device that opens at a preset pressure to prevent damage to the system. It protects the system from overpressure caused by:

Pump failure (e.g., stuck swashplate)
Thermal expansion
External loads

4.2 Hydraulic Actuators

Hydraulic actuators convert hydraulic pressure into mechanical force. They are used for:

Flight control surface actuation
Landing gear extension/retraction
Thrust reverser actuation
Cargo door operation

5. Landing Gear Systems

5.1 Extension and Retraction

Landing gear systems use hydraulic actuators for extension and retraction. The system includes:

Uplock mechanism: Mechanically holds the landing gear in the up position, preventing unintended extension
Downlock mechanism: Mechanically holds the landing gear in the down position
Sequence valves: Ensure correct sequence of gear door and gear operation

5.2 Anti-Skid System

The anti-skid system prevents wheel lockup during braking. It relies on wheel speed sensors to detect impending skid conditions. If a sensor fails, the system will typically flag an inoperative condition and illuminate the warning light.

6. Fuel Systems

6.1 Fuel Tanks and Pumps

6.1.1 Boost Pumps

Boost pumps are located in the fuel tanks and ensure a continuous supply of fuel under pressure to the engine's fuel pump, preventing cavitation and ensuring reliable engine operation.

6.1.2 Cross-Feed Valves

The cross-feed valve allows either engine to be fed from any tank, providing flexibility and redundancy in fuel management. This is essential for:

Balancing fuel loads
Managing fuel in the event of an engine failure
Correcting lateral imbalance

6.2 Fuel Quantity Indicating System (FQIS)

6.2.1 Capacitance Probes

The FQIS uses capacitance probes in the fuel tanks to measure fuel level. The capacitance of the probe changes with the fuel level, as fuel has a different dielectric constant than air.

6.2.2 Compensator Probes

The compensator probe is used to correct for changes in fuel dielectric constant (which varies with fuel type and temperature). If it fails, the system may not properly compensate, leading to inaccurate readings.

6.2.3 Fault Conditions

A short circuit or moisture in the probe wiring can cause an erroneous high reading. A faulty compensator is a classic cause of erroneous FQIS readings.

6.2.4 Capacitance Testing

Capacitance testers simulate fuel levels by applying a known capacitance to the FQIS computer and verifying the indicated quantity. This is a standard test procedure.

7. Air Conditioning and Pressurisation

7.1 Bleed Air System

7.1.1 Bleed Air Sources

Bleed air is extracted from the engine compressor stages or the APU. It is used for:

Air conditioning
Pressurisation
Engine and wing anti-icing
Hydraulic reservoir pressurisation
Water system pressurisation

7.1.2 Precooler

The precooler heat exchanger reduces the temperature of the bleed air to a safe level for downstream systems, particularly the air conditioning pack. It uses fan air or ram air as the cooling medium.

7.1.3 Check Valves

Check valves are one-way valves that prevent bleed air from flowing back into a failed or lower-pressure source, ensuring system integrity.

7.2 Air Cycle Machine (ACM)

The Air Cycle Machine (ACM) is the core of the air conditioning pack. It uses a compressor, turbine, and heat exchangers to cool and dehumidify bleed air. The ACM operates on the reverse Brayton cycle:

Bleed air is compressed (raising its temperature)
The hot air is cooled in a primary heat exchanger
The air is further compressed and cooled in a secondary heat exchanger
The air expands through a turbine, dropping its temperature significantly
The cold air is used for cabin conditioning

7.3 Water Separators

The water separator removes moisture that condenses as the air is cooled in the ACM. This prevents fogging and water ingress into the cabin. The water separator typically uses a cyclone or coalescer principle.

7.4 Cabin Pressurisation

7.4.1 Pressurisation Control

The cabin pressure controller maintains cabin altitude within acceptable limits. As aircraft altitude increases, ambient pressure drops. To keep the differential pressure from exceeding the structural limit, the controller must allow the cabin altitude to rise (i.e., cabin pressure to decrease) while remaining at or below the maximum certified differential pressure.

7.4.2 Outflow Valve

The outflow valve regulates cabin pressure by controlling the rate at which air exits the cabin. If the outflow valve is stuck fully open, the cabin cannot pressurise.

7.4.3 Safety Valve

The safety valve is a pressure relief device that opens if the cabin pressure exceeds a preset limit, preventing structural damage.

7.5 Windshield Heating

The windshield heating system uses a conductive coating on the windshield. The current draw at a given voltage indicates the resistance of the coating. A lower than expected current draw indicates a higher resistance, often caused by damage to the conductive coating, reducing the heating effect.

7.6 Engine Anti-Ice

Engine anti-ice uses bleed air to heat the engine inlet cowl. This bleed air is extracted from the engine, causing a slight decrease in engine performance. To maintain the selected thrust, the engine fuel control unit will increase fuel flow. A typical indication of correct operation is an increase in fuel flow with the system selected ON.

8. Communication and Navigation Systems

8.1 Communication Systems

8.1.1 VHF Transceivers

VHF communication transceivers operate in the 118-137 MHz range. A typical fault scenario: the transceiver receives but does not transmit. If the microphone and PTT switch are verified functional, the failure is in the transmit path. The transmitter circuit within the transceiver, or the antenna coupler (if used), are common failure points.

8.1.2 Cabin Interphone and Passenger Address

The cabin interphone system allows crew communication. The passenger address (PA) system uses a separate amplifier to drive the cabin speakers. If the crew interphone works but the PA doesn't, the amplifier is likely faulty.

8.1.3 Emergency Locator Transmitter (ELT)

ELTs have a test mode that transmits a short signal to verify operation. The test should be brief and performed according to the AMM to avoid triggering a real distress signal.

8.2 Navigation Systems

8.2.1 Automatic Direction Finder (ADF)

The ADF operates in the 190-1750 kHz range and receives signals from Non-Directional Beacons (NDBs). The ADF needle indicates the bearing to the selected NDB station. Correct operation is confirmed when the needle points to the known station's bearing.

8.2.2 Instrument Landing System (ILS)

The ILS provides precision approach guidance. The localizer provides lateral guidance, and the glideslope provides vertical guidance. If a test signal only contains localizer information, the glideslope receiver will not receive a valid signal, and the glideslope deviation indicator will show a full-scale deflection (or an 'off' flag).

8.2.3 Weather Radar

The weather radar system uses a transmitter/receiver unit and a stabilised antenna. The antenna sweep is controlled by a drive motor. If the antenna is not moving, the motor or its control circuit is likely faulty.

The tilt system is a closed-loop servo. If the motor receives power and signals but does not move, a faulty feedback potentiometer can cause the servo loop to be unstable or not respond, as the system cannot sense the antenna's position.

If the system passes its self-test but shows no returns, the radar may be in standby mode, which prevents transmission.

8.2.4 Traffic Alert and Collision Avoidance System (TCAS)

The TCAS self-test is designed to check the system's internal circuitry and the transponder's ability to respond. During the test, the TCAS sends an internal interrogation and the transponder replies, which is a normal part of the self-test sequence. This is independent of the transponder's mode setting, as the test is internal.

A successful TCAS self-test typically includes a visual and aural indication.

8.3 Warning Systems

8.3.1 Ground Proximity Warning System (GPWS)

The GPWS self-test checks the internal computer logic and verifies that the warning outputs (both visual and aural) are functioning. It does not require a valid radar altimeter reading, nor does it test computational functions against terrain.

GPWS Mode 4 provides a warning when the aircraft is too close to terrain with an excessive closure rate.

8.3.2 Fire Detection Systems

In a continuous loop fire detection system, the sensor's resistance changes with temperature. A short circuit (low resistance) is interpreted as a fire or overheat condition.

The test switch simulates a fire condition by applying a test signal to the loop. If the loop is shorted to ground, the control unit may not respond correctly to the test signal, failing to energise the alarm.

8.3.3 Aural Warning Systems

The 'gear not down' warning is triggered by the landing gear position sensors. If a sensor is faulty, it may not signal the down position, causing a false warning.

9. Flight Management Systems

9.1 Flight Management Computer (FMC)

The Flight Management Computer (FMC) integrates navigation sensors, flight plan data, and aircraft performance to provide guidance and predictions. It is a central component of the flight management system.

9.2 Navigation Database

The navigation database is a key component of the FMS, containing the route information necessary for the FMS to compute a flight plan and provide guidance. The database contains:

Waypoints and intersections
Airways and routes
Airports and runways
NAVAIDs (VOR, NDB, DME, ILS)
Holding patterns and procedures

9.3 Central Maintenance Computer (CMC)

The Central Maintenance Computer (CMC) collects BITE data from various systems, allowing maintenance to quickly identify failed LRUs and run tests. It is a diagnostic tool, not a flight control or flight data system.

10. Engine Systems

10.1 FADEC/EEC

FADEC (Full Authority Digital Engine Control) or EEC (Electronic Engine Control) continuously monitors and controls engine parameters such as fuel flow, compressor speed, and temperatures to ensure safe and efficient operation. The FADEC:

Controls fuel flow based on throttle position and ambient conditions
Monitors engine limits (N1, N2, EGT, etc.)
Provides engine protection (overspeed, overtemp)
Communicates with aircraft systems (autothrottle, flight management)

11. Regulatory Framework and Maintenance Practices

11.1 Part-66 Requirements

Part-66.A.25 specifies that an applicant for an Aircraft Maintenance Licence must be at least 18 years of age.

11.2 Maintenance Data

Maintenance must be performed in accordance with approved data. If a task is not in the AMM or other approved documents, it cannot be performed. The technician must report the issue so that the appropriate data can be obtained or the task can be properly scheduled. (Reference: Part-145.A.45, Maintenance Data)

11.3 B1 vs B2 Scope

The B2 (Avionics) licence covers electronic and electrical systems, while B1 (Mechanical) covers mechanical systems (engines, airframe, landing gear). The FMC is a core avionics computer system. Troubleshooting and replacing it falls squarely within the B2 avionics scope.

12. Common Relationships Between Concepts

12.1 System Interdependencies

Electrical → Avionics: All avionics systems require electrical power. Generator failure leads to load shedding, which may affect non-essential avionics.
Pneumatic → Air Conditioning: Bleed air from engines/APU feeds the air conditioning packs.
Hydraulic → Flight Controls: Hydraulic power actuates flight control surfaces, landing gear, and thrust reversers.
Air Data → Autopilot: Autopilot relies on air data (airspeed, altitude) and inertial data for guidance.
Fuel → Engines: Fuel system provides pressurised fuel to engines; FADEC controls fuel flow.

12.2 Sensor-to-Display Path

Sensors (pitot/static, AoA vanes, temperature probes) → ADMs/transducers → Air Data Computers → Display processors/symbol generators → PFD/ND displays.

12.3 Fault Isolation Logic

When troubleshooting, follow a systematic approach:

Verify power is present
Check circuit breakers
Run BITE tests
Verify sensor inputs
Check signal paths
Test output devices

13. Typical Exam Focus Points

13.1 Electrical Systems

28 V DC nominal voltage; 22-30 V operating range
115/200 V AC, 400 Hz three-phase systems
Phase separation of 120 degrees
Function of voltage regulators, GCUs, TRUs, static inverters
Purpose of bus tie contactors, current limiters, circuit breakers
Load shedding priorities
CSD/IDG function for constant frequency

13.2 Air Data and Instruments

Pitot-static system function and leak test procedures
ADM function (converting pressure to digital data)
IRS alignment requirements
EFIS reversionary modes
FDR parameter verification

13.3 Flight Controls

Fly-by-wire redundancy and dissimilarity
CWS mode operation
Yaw damper function (Dutch roll damping)
Stick shaker function
Control surface balancing (trim tabs)

13.4 Fuel Systems

Cross-feed valve purpose
Boost pump function
FQIS capacitance probe operation
Compensator probe function
Capacitance testing procedures

13.5 Hydraulic Systems

Pressure compensator function in variable displacement pumps
Pressure relief valve purpose
Uplock mechanism function

13.6 Air Conditioning and Pressurisation

Precooler function
ACM operation
Water separator purpose
Cabin pressure controller logic
Safety valve function

13.7 Communication and Navigation

VHF transceiver troubleshooting
PA amplifier function
ILS test signal interpretation
Weather radar antenna drive and tilt
TCAS self-test procedure
GPWS modes and self-test

13.8 Warning Systems

Fire detection loop principles (short circuit = fire)
Gear warning sensor operation
Emergency lighting duration (90 minutes)

13.9 Regulatory

Part-66.A.25 minimum age (18 years)
Part-145.A.45 maintenance data requirements
B1 vs B2 task allocation

14. Key Formulas and Values

Parameter	Value
DC system nominal voltage	28 V DC
DC system operating range	22-30 V
AC system voltage	115/200 V AC
AC system frequency	400 Hz
Three-phase phase separation	120°
Emergency lighting duration	90 minutes
Pitot line leak rate limit	~200 ft/min (varies by aircraft)
Static line leak rate limit	~100 ft/min (varies by aircraft)
Minimum age for AML	18 years

15. Conclusion

Module 13 for the B2 licence encompasses a broad range of aircraft systems with a focus on the electronic and electrical aspects. The B2 certifying staff must understand not only the individual systems but also their interconnections and dependencies. A systematic approach to troubleshooting, familiarity with BITE and self-test procedures, and a thorough understanding of the regulatory framework are essential for safe and effective maintenance practice.

The knowledge areas covered in this module form the foundation for the practical skills required in daily maintenance activities, from routine functional tests to complex fault isolation. Mastery of these concepts ensures that the B2 certifying staff can safely certify aircraft systems and maintain the highest standards of airworthiness.