Module 8: Basic Aerodynamics

Module 8: Basic Aerodynamics – B2 Licence Category Study Material

1. Module Overview

Module 8 of the EASA Part-66 basic knowledge syllabus provides the fundamental aerodynamic principles required for aircraft maintenance certifying staff. For the B2 (Avionics) category, this module is essential because avionic systems—particularly pitot-static systems, air data computers, stall warning systems, and flight control systems—are directly influenced by aerodynamic phenomena. A B2 technician must understand not only how these systems operate electronically but also why they are designed and calibrated based on aerodynamic principles.

The module is structured into several sub-topics:

• 8.1: Physics of the Atmosphere – International Standard Atmosphere (ISA), pressure, density, temperature, and their variations with altitude.
• 8.2: Aerodynamics – Airflow around bodies, lift generation, drag, stall, and the lift equation.
• 8.3: Theory of Flight – Straight-and-level flight, climbs, descents, turns, and load factor.
• 8.4: Flight Stability and Dynamics – Static and dynamic stability, control surfaces.
• 8.5: High-Speed Aerodynamics – Compressibility effects, Mach number, shock waves, and critical Mach number.

The knowledge level required for B2 is generally Level 2 (general knowledge) for most topics, with some areas requiring Level 3 (detailed theory), particularly those directly related to avionic systems.

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2. Key Concepts Explained in Detail

2.1 The International Standard Atmosphere (ISA)

The atmosphere is the medium in which an aeroplane operates. Its properties—pressure, temperature, and density—vary with altitude and directly affect aerodynamic forces and instrument indications.

Standard Sea-Level Conditions (ISA):

• Pressure (P₀): 1013.25 hPa (hectopascals) or 101.325 kPa
• Temperature (T₀): 15 °C (288.15 K)
• Density (ρ₀): 1.225 kg/m³
• Speed of sound (a₀): 340.3 m/s (approximately)

Temperature Lapse Rate:

• Up to the tropopause (11,000 m / 36,090 ft): Temperature decreases at a rate of 1.98 °C per 1,000 ft, or approximately 6.5 °C per 1,000 m.
• Above the tropopause (stratosphere): Temperature remains constant at −56.5 °C up to 20,000 m.

Pressure Variation with Altitude:

Pressure decreases with altitude, but not linearly. At sea level, pressure is 1013.25 hPa. At 18,000 ft (approximately 5,500 m), pressure is approximately half of the sea-level value (about 506 hPa). At 30,000 ft (approximately 9,100 m), pressure drops to roughly 300 hPa.

Density Variation:

Density also decreases with altitude, following a similar exponential trend. At 30,000 ft, air density is approximately 0.46 kg/m³—about 38% of the sea-level value.

Why This Matters for the B2 Technician:

• The altimeter is essentially a barometer calibrated to read altitude based on static pressure, assuming ISA conditions.
• The airspeed indicator (ASI) reads dynamic pressure, which depends on air density. At constant indicated airspeed (IAS), true airspeed (TAS) increases with altitude because density decreases.
• Air data computers (ADCs) use ISA models to compute pressure altitude, true airspeed, and Mach number.

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2.2 The Pitot-Static System and Pressure Measurement

The pitot-static system is fundamental to flight instrumentation. It comprises:

Pitot (Impact) Pressure:

• Measured by the pitot tube, typically mounted on the fuselage nose or wing leading edge, facing directly into the airflow.
• The pitot tube senses total pressure (also called stagnation or pitot pressure), which is the sum of static pressure and dynamic pressure.
• Total pressure = Static pressure + Dynamic pressure

Static Pressure:

• Sensed by static ports, typically located on the fuselage side, away from local flow disturbances.
• Static ports are small, flush-mounted openings that sense the ambient atmospheric pressure at the aircraft's altitude.

Placement of Static Ports:

Static ports are deliberately positioned on the fuselage side, away from:

• The wing (to avoid the accelerated airflow over the upper surface)
• The fuselage boundary layer (to avoid local pressure variations)
• Propeller wash or engine exhaust (to avoid pressure fluctuations)

This placement minimises position error—the difference between the pressure sensed at the port and the true ambient static pressure. Misalignment of a static port with the local airflow direction disrupts the sensed static pressure, causing errors in all instruments that use static pressure: the altimeter, vertical speed indicator (VSI), and airspeed indicator.

Pressure Relationships:

• Dynamic pressure (q) = Total pressure − Static pressure
• q = ½ρV² (in incompressible flow)

Example Calculation:

If total pressure = 100 kPa and static pressure = 80 kPa, then:

Dynamic pressure = 100 − 80 = 20 kPa

Blocked Static Port:

If a static port becomes blocked, the static pressure trapped in the system remains at the value present at the time of blockage. Consequently:

• The altimeter reading becomes frozen at the altitude where the blockage occurred.
• The VSI indicates zero (no rate of change).
• The ASI continues to respond to changes in pitot pressure but reads incorrectly (it will over-read during climb and under-read during descent).

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2.3 Airspeed Definitions and Relationships

Understanding airspeed definitions is critical for B2 technicians working with air data systems.

Indicated Airspeed (IAS):

• The reading directly from the airspeed indicator, based on the difference between pitot and static pressure.
• The ASI is calibrated to read correctly under ISA sea-level conditions.

Calibrated Airspeed (CAS):

• IAS corrected for instrument and position errors.

Equivalent Airspeed (EAS):

• CAS corrected for compressibility effects at high speeds and altitudes.

True Airspeed (TAS):

• The actual speed of the aircraft through the air.
• TAS = EAS × √(ρ₀/ρ), where ρ₀ is sea-level density and ρ is the density at altitude.

Relationship Between IAS and TAS:

At constant IAS, dynamic pressure (q) remains constant. Since q = ½ρV²:

• As altitude increases, density (ρ) decreases.
• To maintain the same q, true airspeed (V) must increase.
• Therefore, TAS increases with altitude at constant IAS.

Mach Number:

Mach number (M) is the ratio of true airspeed to the local speed of sound:

M = TAS / a

Where a (speed of sound) = √(γRT), with γ = 1.4 for air, R = 287 J/(kg·K), and T = absolute temperature in Kelvin.

Example Calculation:

If an aircraft flies at Mach 0.8 and the speed of sound is 295 m/s:

TAS = 0.8 × 295 = 236 m/s

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2.4 Lift Generation and the Lift Equation

Lift is the aerodynamic force perpendicular to the relative airflow. It is generated primarily by the pressure difference between the upper and lower surfaces of the wing.

Bernoulli's Principle:

For incompressible, inviscid flow:

P + ½ρV² = constant (along a streamline)

Where:

• P = static pressure
• ½ρV² = dynamic pressure

Over the upper surface of a cambered aerofoil, airflow accelerates, causing a decrease in static pressure. Over the lower surface, airflow decelerates, causing an increase in static pressure. The resulting pressure difference produces lift.

The Lift Equation:

L = ½ρV² × S × CL

Where:

• L = lift force (Newtons)
• ρ = air density (kg/m³)
• V = true airspeed (m/s)
• S = wing reference area (m²)
• CL = lift coefficient (dimensionless)

Lift Coefficient (CL):

• CL is a dimensionless number that represents the lifting efficiency of the wing at a given angle of attack.
• CL increases with angle of attack up to the critical angle of attack (typically 15–20° for conventional aerofoils).
• At the critical angle, CL reaches its maximum value (CLmax).
• Beyond the critical angle, CL decreases rapidly—this is the stall.

Application: Weight Change and Airspeed:

In straight and level flight, lift equals weight:

W = ½ρV² × S × CL

If weight increases by 20% and CL and density remain constant:

V_new² = 1.2 × V_old²

V_new = V_old × √1.2 = V_old × 1.095

Example:

If V_old = 120 m/s:

V_new = 120 × 1.095 = 131.5 m/s

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2.5 The Stall and Stall Warning Systems

The Aerodynamic Stall:

A stall occurs when the critical angle of attack is exceeded. At this point:

• The airflow over the upper surface can no longer remain attached.
• Flow separation occurs, causing a sudden loss of lift.
• The wing produces less lift despite an increase in angle of attack.

Key Principle: Stall is Angle-of-Attack Dependent, Not Airspeed Dependent:

The stall occurs at a specific angle of attack (the critical angle), which corresponds to CLmax. This critical angle is a fixed aerodynamic property of the wing for a given configuration (flap setting) and Mach number.

The airspeed at which the stall occurs varies with:

• Weight: Higher weight requires higher lift, which requires higher CL, which is achieved at a higher angle of attack. Therefore, the stall speed increases with weight.
• Altitude: At higher altitudes, air density is lower, so a higher TAS is required to generate the same lift. The stall speed (TAS) increases with altitude.
• Load factor: In turns, the load factor increases, requiring more lift, which increases the stall speed.

Stall Warning Systems:

• Stall warning devices (e.g., stall warning vanes, angle-of-attack sensors, or stick shakers) are calibrated to activate at an angle of attack slightly below the actual stall angle.
• This provides the pilot with a warning before the stall occurs, allowing recovery action.
• Because the stall is fundamentally an angle-of-attack phenomenon, stall warning is based on angle of attack (AoA), not airspeed.
• AoA sensors directly measure the angle of attack, providing a direct indication of proximity to the critical angle.
• Airspeed alone cannot provide a universal stall warning because the stall speed varies with weight, altitude, and load factor.

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2.6 High-Speed Aerodynamics and Compressibility

At high subsonic speeds (typically above Mach 0.3), air can no longer be treated as incompressible. Compressibility effects become significant.

Mach Number Effects:

As the free-stream Mach number increases, the airflow accelerates over the upper surface of the wing. At a certain free-stream Mach number, the local airflow over the wing first reaches Mach 1.0. This free-stream Mach number is called the critical Mach number (M_crit).

Critical Mach Number:

• The free-stream Mach number at which the local airflow first reaches sonic speed (Mach 1.0) at some point on the aircraft.
• For typical transport aircraft, M_crit is approximately 0.7–0.8.

Flying Above M_crit:

When the aircraft flies above M_crit:

• Local supersonic regions form over the upper surface.
• A normal shock wave forms at the rear of the supersonic region.
• Across a normal shock, static pressure increases significantly, and velocity decreases abruptly.
• This causes wave drag (a sharp increase in drag).
• Flow separation may occur behind the shock, causing buffeting and potential control issues.

Example:

An aircraft flying at Mach 0.85 with local airflow over the wing accelerating to Mach 1.2 is flying above M_crit. The free-stream Mach number is 0.85; the local Mach number is 1.2. The aircraft is in the transonic regime.

Effect on Pitot-Static System:

In high-speed flight, local airflow over the fuselage can accelerate to supersonic speeds, forming a normal shock. If static ports are located behind the shock:

• The sensed static pressure is higher than the true ambient static pressure.
• The altimeter reads low (indicates lower altitude than actual).
• The airspeed indicator reads low (under-reads) because the pressure difference (pitot − static) is reduced.

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2.7 Turns and Load Factor

Coordinated Turn:

In a coordinated turn, the lift vector is tilted inward, providing a horizontal component (centripetal force) that causes the aircraft to turn.

Load Factor (n):

Load factor is defined as:

n = Lift / Weight

In straight and level flight, n = 1.

In a coordinated turn:

• The vertical component of lift must equal weight to maintain altitude.
• The horizontal component provides centripetal force.
• Therefore, total lift must be greater than weight.
• Load factor n > 1.

Relationship:

n = 1 / cos(φ)

Where φ is the bank angle.

Example:

At 60° bank, n = 1 / cos(60°) = 1 / 0.5 = 2. The wings must produce twice the weight in lift.

Effect on Stall Speed:

Stall speed increases with load factor:

V_stall(n) = V_stall(1) × √n

At 60° bank (n = 2), the stall speed increases by √2 ≈ 1.41, or 41%.

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2.8 Boundary Layer and Trailing Edge Effects

Boundary Layer:

The boundary layer is the thin layer of air adjacent to the wing surface where viscous effects are significant. It is divided into:

• Laminar boundary layer: Smooth, orderly flow with low drag but prone to separation.
• Turbulent boundary layer: Chaotic, mixing flow with higher drag but more resistant to separation.

Adverse Pressure Gradient:

Over the rear portion of the aerofoil, pressure increases (adverse pressure gradient). This decelerates the boundary layer and can cause flow separation.

Trailing Edge Deformation:

The trailing edge of an aerofoil controls pressure recovery and boundary layer behaviour. A deformation (e.g., a dent) at the trailing edge:

• Disturbs the smooth airflow.
• Causes an adverse pressure gradient.
• Promotes premature boundary layer separation.
• Reduces the effectiveness of high-lift devices (flaps).
• Lowers CLmax.
• Increases stall speed.

This is why trailing edge damage is critical and requires repair per structural repair manuals (SRMs).

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2.9 Control Surface Forces

Aerodynamic Force on Control Surfaces:

The force required to move a control surface (elevator, aileron, rudder) is proportional to dynamic pressure:

F ∝ ½ρV²

As airspeed increases:

• Dynamic pressure increases.
• The aerodynamic force on the control surface increases.
• The control feel becomes "heavier."

This explains why elevator control feels heavy at high speeds. Aircraft may incorporate aerodynamic balancing (e.g., horn balances, balance tabs) or hydraulic/power-assisted controls to reduce pilot effort.

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2.10 Static Discharge Wicks

Static discharge wicks (static wicks) are small, flexible rods installed on trailing edges (wingtips, elevators, rudders) to dissipate static electricity into the atmosphere. Their purpose is to:

• Prevent static charge buildup.
• Reduce interference with radio communications and avionics.

Aerodynamic Effect:

Static discharge wicks are not aerodynamic devices. Their removal does not alter airflow or aerodynamic performance. They are non-structural, non-aerodynamic items, and their absence affects only static discharge capability, not flight characteristics.

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3. Important Formulas and Relationships

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4. Common Relationships Between Concepts

Pitot-Static System ↔ Atmosphere:

• Static pressure decreases with altitude → altimeter reads altitude.
• Dynamic pressure depends on air density → ASI reads IAS, which relates to TAS through density.

Stall ↔ Angle of Attack ↔ Airspeed:

• Stall occurs at critical AoA (fixed for a given configuration).
• Stall speed varies with weight, altitude, and load factor.
• Stall warning is AoA-based, not airspeed-based.

High-Speed Flight ↔ Pitot-Static Errors:

• Above M_crit, shock waves form over the fuselage.
• Static ports behind a shock sense elevated static pressure.
• This causes altimeter and ASI to under-read.

Lift Equation ↔ Performance:

• Weight increase → higher required lift → higher airspeed at same CL.
• Altitude increase → lower density → higher TAS for same IAS.
• Load factor increase → higher required lift → higher stall speed.

Boundary Layer ↔ High-Lift Devices:

• Trailing edge damage → premature separation → reduced CLmax → higher stall speed.
• Flaps increase CLmax by increasing camber and delaying separation.

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5. Typical Exam Focus Points

For the B2 Module 8 examination, candidates should focus on:

1. ISA values and variations – Sea-level pressure (1013.25 hPa), temperature (15 °C), density (1.225 kg/m³), and lapse rates.
2. Pitot-static system principles – Total vs. static pressure, dynamic pressure calculation, placement of static ports, effects of blockage or misalignment.
3. Airspeed relationships – IAS vs. TAS vs. Mach number, and how they change with altitude.
4. Lift equation applications – Calculating new airspeeds for weight changes, understanding CL and angle of attack.
5. Stall principles – Critical angle of attack, CLmax, stall warning systems, and why AoA-based warning is preferred.
6. High-speed aerodynamics – Critical Mach number, shock waves, compressibility effects on instruments.
7. Load factor and turns – Relationship between bank angle, load factor, and stall speed.
8. Boundary layer and trailing edge effects – Impact of damage on aerodynamic performance.
9. Control surface forces – Relationship between dynamic pressure and control feel.
10. Identification of aerodynamic vs. non-aerodynamic components – E.g., static discharge wicks.

Exam Strategy:

• Understand the physical principles, not just memorise facts.
• Be able to perform simple calculations using the lift equation, dynamic pressure, and Mach number relationships.
• Recognise how aerodynamic principles directly affect avionic systems (pitot-static, air data computers, stall warning).
• Know the SI units used in all calculations (m/s, kPa, hPa, kg/m³, Newtons).

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References

• Regulation (EU) No 1321/2014, Annex III (Part-66), Appendix I – Basic Knowledge Syllabus, Module 8: Basic Aerodynamics.
• Acceptable Means of Compliance (AMC) to Part-66 – Guidance on knowledge levels and examination standards.
• EASA CS-25 (Certification Specifications for Large Aeroplanes) – For aerodynamic performance and systems requirements (reference only).