Study Notes: IR Part-66 Module 15 - Gas Turbine Engine

Fundamentals of Mechanical Energy

Energy is defined as the capacity to do work, with both work and energy measured in Joules. The principle of conservation of energy states that it cannot be created or destroyed, although it frequently converts into less useful forms like waste heat due to friction, conduction, convection, and radiation. Mechanical energy exists primarily as potential energy ($PE$) and kinetic energy ($KE$). Potential energy results from the position of a mass under gravity or its state, such as elastic energy in a spring or pressure energy in a compressed gas. The potential energy due to gravity is calculated by multiplying mass in kilograms by the acceleration due to gravity ($9.81\,m/s^2$) and the height in metres. For example, raising a 1 metric tonne mass to 20 metres results in a $PE$ of $1000\,kg \times 9.81\,m/s^2 \times 20\,m = 196,200\,J$ or $196.2\,kJ$.

Kinetic energy results from the motion of a mass at a specific velocity. To reach this state, energy must be expended to accelerate the mass. The formula for kinetic energy is $KE = \frac{1}{2}mv^2$. For instance, an aircraft with a mass of 20 tonnes travelling at $20\,m/s$ possesses a $KE$ of $0.5 \times 20,000\,kg \times (20\,m/s)^2 = 4,000,000\,J$ or $4\,MJ$. Energy types convert readily; a mass falling from height loses $PE$ and gains $KE$, which then converts entirely into heat upon impact with the ground (provided there is no bounce). Total energy equilibrium in the universe is eventually reached as all energy levels find their way back to dissipated heat.

Newton’s Laws of Motion

Newton’s First Law states that a body remains at rest or in uniform motion in a straight line unless acted upon by an external force. Friction and air resistance typically oppose this motion on Earth, but in the vacuum of space, a probe will travel unassisted indefinitely until a force like vectored thrust is applied. The Second Law dictates that the rate of change of momentum is proportional to the applied force and occurs in the direction of that force. A force is the product of mass ($kg$) and acceleration ($m/s^2$), expressed in Newtons. For example, accelerating $10\,kg$ of air at $20\,m/s^2$ requires a force of $200\,N$. The Third Law states that for every action, there is an equal and opposite reaction. In a jet engine, the rearward acceleration of air mass produces forward reaction thrust on all internal static and rotating parts.

The Brayton and Otto Cycles

The operation of a heat engine requires raising the energy level of a working fluid, typically air, by releasing heat from fuel. In a reciprocating piston engine, this occurs via the Otto Cycle, a constant volume process. Air is compressed and heated at a constant volume, causing a pressure rise that forces the piston down to do work. In contrast, the gas turbine engine operates on the Brayton Cycle, a constant pressure process. Here, induction, compression, combustion, and expansion occur simultaneously and continuously. Heat is added at a constant pressure in the combustion chamber to replace the energy being converted to kinetic energy during expansion. The theoretical cycle involves adiabatic compression in the intake and compressor, heat addition in the combustor, and adiabatic expansion through the turbine and exhaust nozzle.

Power, Work, and Thrust Relationships

Work is the product of force ($N$) and distance ($m$), expressed in Joules ($J$). Power is the rate of doing work, expressed in Watts ($W$), and is calculated as $(F \times d) / t$. For example, moving a book with a steady force of $2\,N$ over $5\,m$ in $3.5\,s$ produces a work of $10\,J$ and a power output of $2.86\,W$. In the context of a turbo-jet, thrust is the forward reaction force generated by accelerating air mass. If a stationary engine consumes $260\,kg/s$ of air and exhausts it at $500\,m/s$, the force is $260\,kg/s \times 500\,m/s = 130,000\,N$ or $130\,kN$. However, if the aircraft is stationary, the work and power are zero. As forward speed ($v$) increases, the net thrust decreases because the change in velocity is smaller ($v_j - v$), but the engine gains power. At a flight speed of $300\,m/s$, the net thrust drops to $52\,kN$, while the power rises to $15.6\,MW$.

The energy required for thrust production varies significantly between engine types. The formula for energy requirement is $\frac{1}{2}mv^2$. Accelerating a small mass to high velocity (turbo-jet) requires far more heat energy than accelerating a large mass to slow velocity (turbo-prop). A high by-pass turbo-fan engine compromises these approaches by using a fan to move a massive amount of cold-stream air and a core to move a smaller, high-velocity hot-stream, making it efficient for subsonic cruise speeds around Mach 0.75 to 0.85.

Constructional Arrangements of Turbine Engines

A simple turbo-jet engine consists of five main sections: intake, compressor, combustion chamber, turbine, and exhaust nozzle. The cycle is continuous, and thrust acts internal static and rotating components. Low by-pass turbo-jets add a second spool (LP and HP) and divert a portion of LP air around the core to mix in the exhaust, increasing propulsive efficiency for high transonic flight. High by-pass turbo-fans extend this ratio (e.g., $5:1$), with the fan providing around 70% of total thrust. Some designs incorporate a three-spool arrangement, allowing the fan to rotate at its optimum speed through an independent shaft (LP fan, IP compressor, HP compressor).

Turbo-propeller engines exploit almost all gas energy in the turbines to drive a propeller through reduction gearing, typically at a ratio of $1:16$. These are rated in shaft horse-power ($shp$) and are unmatched for efficiency at speeds between $350$ and $400\,kts$. Free turbine designs decouple the power turbine from the gas generator, facilitating easier starting and allowing the propeller to spin at optimum RPM. Turbo-shaft engines are similar but drive a specific load like a helicopter transmission or an auxiliary power unit ($APU$) rather than a propeller. They produce no residual thrust, converting all gas energy into mechanical work.

Engine Performance and Thrust Indicators

Gross thrust is the product of mass airflow and exit velocity, ignoring retarding forces. Net thrust ($T_n$) accounts for momentum drag ($MV$), which is the force needed to move the intake air at the aircraft's speed. At high speeds, the propelling nozzle often reaches a "choked" condition where gas velocity reaches the local speed of sound (Mach 1). Because the gas cannot expand further within the nozzle, it retains residual pressure ($p_j$) above ambient ($p_{\text{amb}}$), producing additional pressure thrust calculated as $A(p_j - p_{\text{amb}})$.

Thrust is distributed throughout the engine components: forward-acting forces are generated in the compressor, diffuser, and combustion chamber, while rearward-acting forces occur in the turbines and the propelling nozzle. Resultant thrust is the sum of these forces. Turbo-props use equivalent shaft horse-power ($eshp$), which adds the residual jet thrust to the shaft power. Static sea level ISA conditions allow a conversion where approx $2.5\,lb$ of jet thrust equals $1\,shp$. Specific Fuel Consumption ($SFC$) measures efficiency as the mass of fuel needed to produce $1\,lb$ of thrust ($lb/hr/lbt$) or $1\,eshp$ for one hour. $SFC$ generally improves with altitude as inlet temperatures drop, until around $36,000\,ft$ where decreasing air density requires higher RPM to maintain mass flow.

Thermodynamics and Engine Efficiency

Thermal efficiency (internal) is the ratio of fuel energy converted into kinetic energy in the jet stream, usually around 45% for modern turbo-fans. Propulsive efficiency (external) is the ratio of kinetic energy converted into propulsive work. It is highest when aircraft speed ($v$) is closest to jet velocity ($v_j$. The formula for propulsive efficiency is $PE = 2v / (v + v_j)$. Overall efficiency is the product of thermal and propulsive efficiencies.

Engine Pressure Ratio ($EPR$) is the primary thrust setting indicator, representing the ratio of turbine exit pressure to intake pressure ($P_7/P_1$). High by-pass engines may use Integrated $EPR$ ($IEPR$), which integrates fan and hot-stream pressures. Thermodynamic behavior follows the Combined Gas Law: $(P \times V) / T = \text{Constant}$. Subsonic flow in a convergent duct results in increased velocity and decreased pressure/temperature, while the opposite occurs in a divergent duct (diffusion). At supersonic speeds, these relationships reverse: air slows and compresses in a convergent duct and accelerates in a divergent duct. The continuity of mass flow ($\rho AV = \text{Constant}$) dictates that area changes must accommodate density and velocity changes to keep flow rates stable.

Engine Ratings and Ambient Factors

Engines are rated by the maximum static thrust they produce at sea level ISA conditions ($15^{\circ}C$ and $1013.25\,mb$). Rated thrust is the maximum thrust certified by the manufacturer. Modern engines are often "Flat Rated," meaning they can deliver maximum rated thrust up to a specific ambient temperature (e.g., $30^{\circ}C$) by using a reserve of power. Beyond this temperature, the engine becomes EGT-limited and performance falls. High air temperature and high altitude (low pressure) both reduce air density and thus mass flow and thrust. Humidity slightly reduces air density, marginally lower thrust.

Forward aircraft speed Initially causes thrust to fall due to momentum drag, but this is eventually overcome by "ram recovery," where the intake boosts compressor entry pressure. Overall engine limitations include shaft speeds ($N_1$, $N_2$, $N_3$), exhaust gas temperature ($EGT$), and oil limits. Exceeding these triggers mandatory inspections for over-speed or over-temperature, often involving time-logged "creep" assessments.

Compressor Inlet Ducts and Shock Waves

Intakes must decelerate free-stream air to approx Mach 0.4 at the compressor face to prevent tip stall and shock waves on rotor blades. Subsonic aircraft use divergent Pitot intakes, which are efficient up to Mach 1.2. Transonic intakes vary operation based on engine demand: critical (shock wave at lip), sub-critical (spillage/drag), and super-critical (shocks inside duct). Supersonic flight requires variable geometry intakes (like on Concorde) that use ramps to position oblique and normal shock waves, gradually slowing air to subsonic velocities while recovery pressure reaches up to 200\% at Mach 2.

Ice protection is vital for intakes. Ground suction creates a depression that lowers temperature; moisture freezes at ambient temperatures up to $10^{\circ}C\, (50^{\circ}F)$. Common systems include Hot Air (HP bleed air distributed via a perforated ring), Electrical (heating mats for turbo-props), and Hot Oil (using scavenge oil to heat vanes/struts). Hot air anti-icing typically raises $EGT$ and reduces net thrust.

Compressor Principles and Construction

Compressors supply high-pressure air to the combustion chamber. Centrifugal compressors (impeller, diffuser, manifold) are robust and provide high pressure rise per stage (max approx $4:1$) but have high frontal areas. Axial compressors use alternating rows of rotating blades and stationary vanes. Each rotor-stator stage provides a ratio of approx $1.2:1$. Total pressure ratio is the product of all stage ratios. Modern axial compressors achieve ratios of $30:1$ or more. Axial blades are twisted (stagger angle) to maintain uniform airflow and prevent tip stall.

Multi-spool (two or three spool) designs allow the rear compressor stages to spin faster than the front, improving efficiency, reducing starting loads, and allowing for by-pass operation. Stator vanes often have "end bend" to energize boundary layers. Titanium is used for lower-temperature stages, while nickel-based alloys like Inconel are used for high-temperature HP stages. Balanced rotors are critical; out-of-balance forces increase by the square of the speed. Fan blades are matched by "moment weight" (marked on the root) to ensure single-plane (static) and two-plane (dynamic) balance.

Stall, Surge, and Airflow Control

Stall is a localized aerodynamic breakdown on rotor blades, while surge is a complete flow reversal through the compressor. Causes include icing, FOD, rapid throttle movements, or off-schedule control systems. A "hung stall" is a sustained stall that can progress to surge. Both lead to high $EGT$, vibration, and rhythmic "thumping" or loud bangs. Control methods include Bleed Valves (opening at low RPM or during deceleration to reduce volume and prevent choking the rear stages) and Variable Vanes (VIGVs and VSVs). Variable vanes change the angle of attack at low RPM to prevent front-stage stalling and progressively open as speed increases.

Combustion Systems and Processes

The combustion section raises the air's energy via fuel release at constant pressure. Efficiency is nearly 100% at sea level. The stoichiometric air:fuel ratio for kerosene is $15:1$, but the overall engine ratio is often $60:1$ to $130:1$. Only 20-25% of air is used for primary combustion; the rest is used for cooling the flame tube and diluting gases to a temperature acceptable to the turbine (approx $900-1100^{\circ}C$).

Swirl vanes create a toroidal vortex to anchor the flame. Systems include: Multiple-can (individual tubes), Tubo-annular (can-annular with common casings), and Annular (single continuous chamber). Annular chambers are the most compact and efficient. Dual-dome systems use primary and secondary regions to manage wide air:fuel ratios and reduce emissions like carbon monoxide ($CO$) and nitrogen oxides ($NO_x$). Vaporizing systems mix fuel with air in a heated tube before injection. Protection against fuel accumulation is provided by a spring-loaded combustion chamber drain valve that opens upon shutdown.

Turbine Section and Materials

Turbines extract energy to drive compressors and accessories. One stage consists of a static ring of nozzle guide vanes ($NGVs$) and a row of rotor blades. Most blades are Impulse-Reaction type: the root is impulse-sectioned (parallel) and the tip is reaction-sectioned (convergent), ensuring uniform exit velocity. High-pressure ($HP$) turbines experience the highest temperatures and are cooled internally by HP compressor air via impingement, convection, and film cooling.

Materials are typically Nimonic high-nickel alloys. Blades can be Equi-axed (standard crystal), Directionally Solidified (elongated crystals for better creep resistance), or Single Crystal (purest structure for highest heat resistance). Shrouded blades have a "T" tip that prevents gas leakage and reduces vibration. Active Clearance Control ($ACC$) manages the gap between blade tips and the static shroud by cooling the casing with bleed air. Blades suffer from stress (centrifugal and bending) and creep (permanent extension over time). Creep has three stages: primary, secondary (predictable), and tertiary (failure). Over-limit temperatures accelerate creep dramatically.

Exhaust Systems and Noise Suppression

Exhaust units decelerate gas slightly (diffusion) to reduce friction before acceleration through the propelling nozzle. Inner cones and aerodynamic struts straighten flow. Propelling nozzles are typically convergent; if the nozzle "chokes" at Mach 1, it produces pressure thrust. Supersonic aircraft use Convergent-Divergent (Con-di) nozzles to allow gas to expand and accelerate beyond Mach 1. Concorde used secondary flaps to form a divergent section aft of the primary nozzle.

Noise suppression utilizes lobed or corrugated nozzles to increase airflow mixing, or internal mixing of hot and cold streams in high by-pass engines. Multi-layer acoustic linings (honeycomb sandwich) absorb core and fan noise. Thrust reversers include: Hot-stream deflector doors, Clamshell doors (rotated into stream), and Cold-stream reversers (translating cowls with blocker doors and cascade vanes). Reversers are generally inhibited in flight and utilize pneumatic or hydraulic power with safety interlocks and feedback mechanisms.

Bearings and Seals

Main bearings are either ball (absorb radial and axial thrust loads; locate shafts) or roller (absorb radial loads; allow axial expansion). Squeeze-film bearings use an oil film around the outer race to dampen vibration. Bearings are grouped in chambers and kept cool with sealing air and oil spray. Seals include Labyrinth (series of rotating fins near abradable linings), Hydraulic (oil troughs), Carbon (rubbing surface for high integrity), and Brush (metal bristles against ceramic). Sealing air pressure must always exceed bearing chamber pressure to prevent oil leakage.

Lubricants and Fuels

Gas turbines use synthetic ester-based oils. Type 1 (alkyl diester) and Type 2 (polyester) are common; they are not compatible and must never be mixed. Key properties include low viscosity, high flash point, and stability against oxidation. Fuels include Jet A, Jet A-1 (kerosene, freezing point $-44^{\circ}C$ to $-50^{\circ}C$), and Jet B (wide-cut/gasoline blend). Kerosene is preferred for its lower volatility (low fire risk) and higher calorific value per unit volume. Additives include Icing Inhibitors ($FSII/PRIST$), Biocides to prevent Cladosporium Resinae fungal growth, and Static Dissipators to prevent electrical arcing during refuelling.

Lubrication and Fuel Metering Systems

Lubrication systems are Dry Sump, Re-circulatory systems. Pressure Relief systems maintain constant pressure from idle to max, whereas Full Flow systems match pump output to RPM (allowing for increased bearing demands at high speed). Components include oil tanks with de-aerators, spur gear or gerotor pumps, and Fuel Cooled Oil Coolers ($FCOC$). Filters use micronic ratings (e.g., $15\,microns$ for disposable). Magnetic Chip Detectors ($MCDs$) provide on-condition monitoring of ferrous wear.

Fuel metering systems provide precise control through hydro-mechanical units or electronic controllers ($EEC, FAFC, FADEC$). Pressure Control systems keep upstream pressure constant, while Flow Control systems maintain a constant pressure drop across the throttle valve. Acceleration Control Units ($ACU$) schedule fuel flow against HP compressor delivery pressure ($P_3$) to prevent stall. $FADEC$ (Full Authority Digital Engine Control) replaces traditional units, managing everything from start cycles to turbine clearance control via high-speed digital processors and torque motors. Nozzle types range from Simplex (fixed area) and Duplex (primary/main orifices) to advanced Air-spray nozzles.

Starting, Ignition, and Indication

Starting cycles involve motoring the engine via electric or air turbine starters, followed by high-energy ignition and fuel introduction. Self-sustaining speed is roughly 30-50% RPM. Potential faults include Hot Starts (insufficient air), Hung Starts (stalled acceleration), and Wet Starts (ignition failure). Ignition units use storage capacitors to produce discrete sparks of approx $4-16\,Joules$. Units have lethal potential and require documented discharge times (approx 1-5 mins) before servicing. Indication systems use digital and analogue displays for performance ($EPR$, $N_1$, $N_2$, $EGT$, Fuel Flow) and condition (Oil $P/T$, Vibration). Exhaust temperature is sensed by Chromel-Alumel thermocouples wired in parallel for averaging. Vibration is monitored by electro-magnetic or piezo-electric accelerometers.

Maintenance, Storage, and Preservation

Maintenance involves Hard Time (fixed lives), On-Condition (inspections like boroscoping and oil analysis), and Condition Monitoring (reliability stats). Ground operations require strict hazard zone awareness and aircraft preparation (weight and wheel chocking). Storage and preservation are critical for engines out of use for over 7 days. Procedures include fuel system inhibiting (replacing fuel with inhibiting oil), fitting desiccant (silica gel, blue when active), and sealing in Moisture Vapour Proof ($MVP$) bags. Humidity indicators must be checked periodically (e.g., every 24 hours Initially, then fortnightly or monthly). Depreservation involves removing all covers, blowing through fire spray rings, and draining inhibiting fluids.", "title": "Study Notes: IR Part-66 Module 15 - Gas Turbine Engine"}