M — Powerplant (Groupe motopropulseur)Chapter 6 · 20 practice questions

Chapter 6: Turbine Engine — Theory & Construction

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Turbine Engine — Theory & Construction

Overview

This chapter covers the fundamental principles of gas turbine engine operation, including the thermodynamic cycle, major engine components, and their functions. It explores the construction and operational characteristics of compressors, combustion chambers, turbines, and supporting systems such as fuel control, ignition, and thrust reversers. The material also addresses common inspection findings, performance anomalies, and the differences between engine configurations like single-spool, dual-spool, turbojet, and turbofan designs.

Key Concepts Explained

Diagram — Turbine Engine — Theory & Construction Cycle de Brayton — Moteur à Turbine à Gaz Diagramme de principe — Théorie TEA (Transport Canada) ENTRÉE Air (P₁, T₁) COMPRESSEUR Air comprimé (P₂ > P₁, T₂ > T₁) CHAMBRE COMBUSTION Carburant + Air (T₃ >> T₂) TURBINE Détente (P₄ < P₃, T₄ < T₃) TUYÈRE Poussée Légende du Cycle de Brayton (Cycle à Pression Constante) 1. Compression adiabatique (Air → P₂, T₂) 2. Combustion isobare (Ajout de carburant → T₃ max) 3. Détente adiabatique (Turbine → Travail mécanique) 4. Échappement (Tuyère → Poussée nette) Rendement η = (Travail sortie) / (Chaleur ajoutée) Configurations: Simple corps / Double corps — Turboréacteur / Turbofan Schéma pédagogique TEA — Transport Canada — Révision 2025

The Brayton Cycle

Animated Brayton cycle Animated Brayton Cycle — Aircraft Turbine Engine INTAKE Ambient air P ≈ 1 atm 1. COMPRESSION P ↑ T ↑ Axial/centrifugal compressor 2. COMBUSTION Fuel P constant, T ↑↑ 3. EXPANSION Shaft — drives the compressor P ↓, T ↓ — Work extracted 4. EXHAUST P atm Brayton Cycle — 4 continuous processes (constant pressure) 1. Compression: P↑, T↑ — air is drawn in and compressed 2. Combustion: P constant, T↑↑ — fuel is injected and burned 3. Expansion: P↓, T↓ — gases produce work (turbine) 4. Exhaust: exhaust gases are discharged to the atmosphere Compressor efficiency affects the EGT. A surging compressor reduces airflow. Single-spool engine: 1 shaft. Twin-spool engine: 2 concentric shafts. Turbofan: bypass flow bypasses the core. Continuous flow — all 4 processes occur simultaneously

The gas turbine engine operates on the Brayton cycle, also known as the constant-pressure cycle. This thermodynamic cycle consists of four continuous processes:

  1. Compression – Air is drawn into the compressor and its pressure is raised significantly.
  2. Constant-pressure heat addition – Compressed air enters the combustion chamber where fuel is injected and burned continuously, raising the temperature and energy of the gas.
  3. Expansion – The high-energy, high-temperature gases expand through the turbine, extracting energy to drive the compressor and accessories.
  4. Constant-pressure heat rejection – The remaining gases exit through the exhaust nozzle, producing thrust.

Unlike reciprocating engines, the Brayton cycle is continuous rather than intermittent, allowing for smooth, high-power output with fewer moving parts.

Compressor Types and Functions

Axial compressor stages Axial Compressor Stages Schematic cross-section — Multi-stage axial compressor AIR Axial flow ROTOR 1 STAT. 1 P↑ ROTOR 2 STAT. 2 P↑ ROTOR 3 STAT. 3 TO CHAMBER Rotor (rotating blades) Stator (stationary blades) Axial air flow P↑ = Increasing pressure Each rotor/stator stage increases pressure and straightens the flow for the next stage.

The primary purpose of the compressor is to draw in air and compress it to a high pressure before it enters the combustion chamber. This compression is essential for the Brayton cycle to achieve efficient combustion and energy extraction.

Axial flow compressors are the most common type used in modern turbofan and turbojet engines. They consist of alternating rows of rotating blades (rotors) and stationary vanes (stators). The rotors accelerate the air, while the stators decelerate it, converting velocity into pressure. Axial compressors offer high efficiency and can handle large airflow volumes, making them ideal for high-thrust applications.

Variable inlet guide vanes (VIGVs) are fitted to some compressors to adjust the angle of air entering the first stage. Their purpose is to match the airflow to rotor speed, preventing stall and improving efficiency, particularly at low RPM. By optimizing the angle of attack on the first rotor blades, VIGVs help maintain stable compressor operation across a wide range of power settings.

Variable Inlet Guide Vanes (VIGV) Mechanism Variable Inlet Guide Vanes (VIGV) CUTAWAY VIEW — FIRST STAGE OF THE AXIAL COMPRESSOR Airflow VIGV Rotor 1 Stator 1 Rotor 2 Stator 2 To subsequent stages Engine centerline VARIABLE PITCH MECHANISM Control ring ACTUATOR (hydraulic or pneumatic) Ring rotation EFFECT ON ANGLE OF ATTACK LOW SPEED (high pitch) 25° Optimal angle of attack for low speed HIGH SPEED (low pitch) 5° Reduces friction losses at high flow rates VIGV FUNCTIONS Prevent compressor stall Improve efficiency at low speeds Optimize the angle of attack of air Reduce the risk of surging LEGEND VIGV vanes (variable pitch) Rotor blades (rotating) Stator vanes (fixed) Incoming airflow Flow straightened by VIGV The VIGVs are controlled by the fuel control unit (FCU) based on rotor speed and throttle position. VARIABLE PITCH

Combustion Chamber

Combustion chamber zones Combustion Chamber Zones Dilution zone (between casing and liner) Chamber head Primary zone Intermediate zone Dilution zone Primary air Secondary air Dilution air INJ Fuel Flame stabilized Cooling film (protects the liner) Hot gases to turbine Zone legend: Primary zone: Air-fuel mixing, combustion, stabilized flame Intermediate zone: Completes combustion, gas mixing Dilution zone: Cools gases to turbine allowable temperature Film cooling: Cold air film on the liner Fuel is injected under pressure and atomized finely for optimal mixing with primary air.

The combustion chamber (combustor) is where fuel is injected and burned continuously with the compressed air, raising the temperature and energy of the gas. The design must ensure stable combustion, minimal pressure loss, and uniform temperature distribution at the turbine inlet.

Combustor liners are subject to extreme thermal stress and may develop cracks over time. Maintenance manuals specify allowable crack limits; if a crack is within these limits, no action is required. Sooty deposits on combustor components indicate incomplete combustion, typically caused by a rich mixture. This can lead to hot section fouling and reduced efficiency.

Turbine Section

The turbine nozzle (stator vanes) accelerates the hot gases from the combustor and directs them at the optimal angle onto the turbine blades. This maximizes energy extraction by ensuring the gas stream strikes the blades efficiently.

Turbine Nozzle (Stator) — Gas Acceleration onto Blades Turbine Distributor: Gas Acceleration COMBUSTION CHAMBER High-temperature and high-pressure gas DISTRIBUTOR (STATOR) Fixed distributor vanes CONVERGENT GAS ACCELERATION Velocity increases, pressure decreases GAP TURBINE BLADES Energy extraction OPTIMAL ATTACK ANGLE α OUTLET Expanded gas reduced velocity High-pressure hot gas Fixed vanes (distributor) Moving blades (rotor) Expanded gas The distributor's convergent profile converts pressure into velocity, directing the flow at the optimal attack angle for the turbine blades.

The turbine extracts energy from the gas stream to drive the compressor and accessories. Turbine blades operate in extreme conditions and may develop cracks. As with combustor liners, maintenance manuals define allowable crack limits for turbine blades. A technician finding a crack during borescope inspection must refer to the manual before deciding on replacement.

Ignition System

Igniter plugs produce a high-energy spark to ignite the fuel-air mixture during engine start. Once the engine is running, combustion is self-sustaining, and the igniters are no longer required. The ignition system is typically used only during the start sequence and for re-light in flight if necessary.

Fuel Control Unit

The fuel control unit (FCU) meters the correct amount of fuel to the combustor based on power lever position, altitude, and other parameters to maintain the desired thrust. It is a critical component that ensures the engine receives the appropriate fuel flow for all operating conditions.

Accessory Gearbox

The accessory gearbox is driven by the high-pressure spool and provides power to drive essential engine and aircraft accessories, such as fuel pumps, oil pumps, hydraulic pumps, and electrical generators. It is a key interface between the engine core and the aircraft systems.

Thrust Reverser

Thrust reversers deploy to redirect the exhaust flow forward, creating reverse thrust to assist in decelerating the aircraft after landing. They are a critical safety feature that reduces landing distance and brake wear.

Important Formulas, Regulations, and Procedures

Borescope Inspection Procedures

Borescope Inspection Points — Engine Cutaway Borescope Inspection Points — Turbine Engine Air inlet AXIAL COMPRESSOR BORESCOPE POINT Compressor blades COMBUSTION CHAMBER BORESCOPE POINT Liner TURBINE BORESCOPE POINT Turbine blades Hot gases MAINTENANCE MANUAL Acceptable limits — Compare: ACCEPTABLE NOT ACCEPTABLE Cracks, nicks, deposits → see chapter 6 → compare to limits → decide: repair or replace GOLDEN RULES OF BORESCOPE INSPECTION: 1. Compare ANY anomaly to manual limits — never estimate. 2. Black soot deposits on turbine blades = incomplete combustion (rich mixture). 3. Small nicks on compressor blades → blending if permitted. 4. Liner cracks → tolerated only within specified limits. LEGEND Borescope inspection point Air / gas flow Acceptable zone Non-acceptable zone BORESCOPE PROBE

Borescope inspections are performed to examine internal engine components without disassembly. Key inspection points include:

  • Combustor liners – Check for cracks, distortion, and soot deposits.
  • Turbine blades – Check for cracks, erosion, and deposits.
  • Compressor blades – Check for dents, nicks, and leading edge damage.

Critical rule: Always refer to the maintenance manual for allowable damage limits before deciding on corrective action. Cracks, dents, and deposits may be acceptable within specified parameters.

Compressor Blade Damage Limits

Small dents on compressor blades, particularly on the leading edge, can often be blended out to restore aerodynamic shape and prevent stress risers. The maintenance manual specifies allowable limits for dent depth, location, and number. Blending must be performed using approved techniques and tools.

Performance Anomalies

Slow acceleration when the power lever is advanced may be caused by:

  • A bleed valve stuck closed, which can cause compressor stall during acceleration.
Compressor Stall — Airflow Separation Mechanism Compressor Surge: Causes and Prevention CAUSES OF SURGE 1. Disturbed Inlet Aerodynamic stall: airflow separates from blades (vortices, gusts, angle of attack) 2. Rapid Throttle Movement Abrupt acceleration: fuel flow too high relative to airflow 3. Damaged Blades Nicks, corrosion, erosion: disruption of the blade aerodynamic profile 4. Blocked Bleed Valve Bleed valve stuck closed: no air evacuation during acceleration FLOW THROUGH THE COMPRESSOR AIR INLET Normal flow Attached flow SURGE: separated airflow Damaged blade BLEED VALVE Bled air TO COMBUSTION CHAMBER SURGE PREVENTION Role of Bleed Valves • Evacuate inter-stage air • Reduce upstream pressure • Maintain attached flow • Active during accelerations Normal Operation Valve closed in steady state Full pressure to combustion Surge Risk Valve stuck closed: excessive pressure, separated flow Solution Valve open: air evacuation stabilizes the flow Maintenance Inspection Check blade condition and valve operation Normal flow Surge / separation Bleed valve Transport Canada — AME Training — M-POWERPLANT ch6
  • A faulty fuel control unit.

High exhaust gas temperature (EGT) at a given power setting is most likely caused by a lean mixture, which burns hotter. Possible causes include:

  • Fuel control issues.
  • Clogged fuel nozzles.
  • Air leaks in the combustion section.

Sooty deposits on turbine blades or combustor components indicate a rich mixture and incomplete combustion, leading to hot section fouling.

Metallic particles in the oil filter most likely originate from the accessory gearbox or bearings. Turbine blade failure would produce different debris and other symptoms such as vibration or performance loss.

Compressor stall during high-power operation is caused by airflow separation from the compressor blades. Common causes include:

  • Disturbed inlet airflow.
  • Rapid throttle movements.
  • Compressor blade damage or fouling.

Common Relationships Between Concepts

Single-Spool vs. Dual-Spool Engines

Single-Spool vs Dual-Spool Engine Cutaway Single-spool vs dual-spool engine SINGLE-SPOOL ENGINE Incoming air COMPRESSOR (rotors + stators) COMBUSTION CHAMBER TURBINE (nozzle + blades) Exhaust gas Single shaft — compressor and turbine linked together N = single speed (compressor = turbine) One shaft connects the compressor and turbine — identical rotation speed DUAL-SPOOL ENGINE Incoming air LOW-PRESSURE COMPRESSOR (LP / N1) HIGH-PRESSURE COMPRESSOR (HP / N2) COMBUSTION CHAMBER HIGH-PRESSURE TURBINE (HP / N2) LOW-PRESSURE TURBINE (LP / N1) LP shaft (outer) — connects LP compressor to LP turbine HP shaft (inner) — connects HP compressor to HP turbine N1 (LP) and N2 (HP) rotate at distinct optimal speeds Two concentric shafts — improved efficiency, wider operating range TECHNICAL COMPARISON SINGLE-SPOOL DUAL-SPOOL Single shaft — compressor and turbine linked together Two concentric shafts — independent LP and HP Single speed N — compromise for all operating conditions N1 and N2 optimized separately for each stage Acceptable efficiency, limited operating range Better overall efficiency, wider range Higher surge risk at low speeds Reduced surge risk — better stability vs Chapter 6 — Turbine engine: Theory and construction | Transport Canada — AME Training

A single-spool engine has one shaft connecting the compressor and turbine. All compressor stages rotate at the same speed.

A dual-spool engine has two concentric shafts:

  • The low-pressure spool connects the low-pressure compressor (or fan) to the low-pressure turbine.
  • The high-pressure spool connects the high-pressure compressor to the high-pressure turbine.

This arrangement allows each spool to operate at its optimal speed, improving efficiency and stall margin across the operating range.

Turbojet vs. Turbofan

Turbojet vs turbofan Turbojet vs turbofan — Core flow vs fan flow comparison TURBOJET (pure turbojet) Air inlet Comp. axial Combustion Brayton Turbine + nozzle Exhaust hot gases Single flow: all air passes through the core Thrust = gas flow × exhaust velocity Efficiency: low at subsonic speed Noise: high (high-speed jet) Fuel consumption: high TURBOFAN (dual flow) Fan blower Secondary flow (bypass) — bypasses the core Comp. LP + HP Combustion Brayton Turbine HP + LP Exhaust core + bypass Thrust = core flow + fan flow (bypass) Efficiency: high at subsonic speed Noise: lower (jet slowed by bypass) Fuel consumption: reduced (better economy) Comparison — Core flow vs fan flow: thrust, efficiency, noise Criterion Turbojet Turbofan Maintenance impact Thrust Very high-speed gas Low mass flow rate Gas + fan at moderate speed High mass flow rate Fan and bearing wear to monitor Efficiency Optimal at supersonic speed Optimal at subsonic speed (M < 0.85) Check VIGV and bleed valves Noise Dominant jet noise Very high Fan noise + attenuated jet Quieter Fan blade inspection and casing Fuel consumption High (high TSFC) Reduced range Reduced (low TSFC) Better range FCU control and fuel injection TSFC = Thrust Specific Fuel Consumption (lb/lbf·h)

The primary advantage of a turbofan engine over a turbojet is the large fan that bypasses the core. This produces thrust more efficiently at subsonic speeds, resulting in better fuel economy. The bypass air also reduces noise and exhaust velocity.

Compressor and Turbine Relationship

Compressor-turbine energy balance Compressor-Turbine Energy Balance 1. POWER EXTRACTED BY THE TURBINE Drives the compressor and accessories COMPRESSOR Shaft ACCESSORY GEARBOX Alternators, pumps hydraulic, FCU ≈ 60-70% of energy 2. COMBUSTION CHAMBER Heating at constant pressure (Brayton cycle) COMBUSTION Compressed air + fuel → hot pressurized gases Compressed air Hot gases 3. GAS EXPANSION IN THE TURBINE Energy extraction to drive the compressor TURBINE DISTRIBUTOR Hot gases Shaft ≈ 30-40% remaining 4. NOZZLE AND THRUST PRODUCTION Remaining energy converted into thrust NOZZLE Gas acceleration THRUST Exhaust gases at high velocity → reaction (Newton's 3rd law) Air Recycled mechanical energy Brayton cycle: Compression → Combustion → Expansion → Exhaust. The turbine extracts just enough energy to drive the compressor and accessories.

The compressor and turbine are mechanically linked by one or more shafts. The energy extracted by the turbine must be sufficient to drive the compressor and accessories. Any imbalance in this relationship—such as from bleed valve malfunctions, fuel control issues, or component damage—will affect engine performance and stability.

Fuel-Air Mixture and Temperature

The fuel-air mixture ratio directly affects combustion temperature:

  • Lean mixture (excess air) → higher EGT.
  • Rich mixture (excess fuel) → lower EGT but incomplete combustion, leading to soot deposits.

The fuel control unit maintains the correct mixture across all operating conditions to balance performance, temperature limits, and emissions.

Diagram

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