Gas Turbine Engine - Lecture Flashcards

Main types of gas turbine engines

Turbojet: original gas turbine engine; now rarely used in aircraft. Core components (engine core) include intake, compressor, combustion, turbine, and exhaust.
Turbofan: same basic core as turbojet but adds a large fan in the intake. Not all air from the intake enters the core; most air is bypass air that flows around the core, improving efficiency and reducing noise for subsonic flight.
Turboprop: retains the turbojet core but drives a propeller via a shaft; most thrust comes from the propeller, similar to a piston-engine setup.
Subsonic focus: this chapter concentrates on subsonic aircraft, with notes that supersonic engines share the same elements but require adjustments for very fast airflow.

Air intake (inlet)

Function: scoop in as much air as possible and direct it smoothly toward the compressor.
Poor intake design increases turbulence, reducing thrust and performance.
Divergent duct: air intake area increases toward the compressor/fan, causing a slight slowdown and a pressure rise known as ram recovery, which improves efficiency and flow through the engine.
Divergent duct example:
ext{divergent duct}
ightarrow ext{increases area}
ightarrow ext{ram recovery}
Intake designs depend on engine location on the aircraft:
- Pitot type (single entrance): simplest, used on wing or tail-mounted engines. Air flows straight to the compressor; relatively short intake.
- Pitot limitations: performance issues at extreme attitudes (e.g., high angles of attack); most commercial airliners do not operate in those conditions.
- Detoured intake: air takes a slight detour via ducting; may include a front-screen to prevent debris entry.
- Divided entrance intake: air enters from two sides and merges before the compressor; used in some single-engine fighter jets (air from fuselage sides, converges toward center of engine).
Ice protection: intake is exposed to the environment; ice can restrict airflow. Ice protection is essential in cold weather; protection systems often use bleed air from the engine for anti-ice/anti-icing tasks.
Bleed air: hot air from the compressor used for ice protection and other systems.
Important case study: 13 January 1982, Boeing 737 at Washington National Airport
- Snowstorm conditions with lingering ice; aircraft waited 49 minutes before takeoff.
- Ice/snow on wings and around engines; engine ice protection system was not activated.
- The aircraft struggled during takeoff roll; the pilots did not abort and subsequently crashed into the Potomac River.
- Main issue: ice in the intake restricted air flow and, more critically, ice blocked sensors used to set engine thrust (Engine Pressure Ratio, EPR).
- EPR (Engine Pressure Ratio) is the ratio of the pressure of the air exiting the engine to the intake pressure:
  EPR = rac{P{ ext{exit}}}{P{ ext{inlet}}}
- In this incident, the EPR sensor inside the air intake was blocked by ice, leading to false thrust readings and significantly reduced thrust relative to what EPR indicated.

Compressor (air compression stage)

Role: first moving part the air encounters; increases air pressure substantially to enable high-thermal-energy combustion.
Two main compressor types:
- Centrifugal compressor
- Axial flow compressor
In practice: some engines combine both (axial front end, followed by centrifugal or vice versa).
Influence on flow: centrifugal compressors redirect air, causing a change in flow direction; axial flow compressors maintain a mostly straight flow.

Centrifugal compressor

Structure: impeller (rotating) and diffuser (stationary outer ring).
Operation: air is directed toward the center of the rotating impeller and accelerated outward by centrifugal force.
Diffuser: converts some kinetic energy into pressure energy; collects compressed air in a compressor manifold.
Spool connection: impeller is driven by a shaft connected to a turbine at the rear; part of a broader spool assembly.
Typical pressure rise: around PR_{ ext{centrifugal}} \approx 7:1 (i.e., air pressure increases about sevenfold).
Stages: multiple centrifugal compressors can exist, each called a stage; typically up to two stages are common.
Advantages: simple, robust, relatively cheap to manufacture; resistant to ice/debris.
Disadvantages: efficiency drops when flow direction changes; less suitable for very large thrust requirements.
Common use: now mainly in smaller turboprop engines.
Key figure reference: Figure 8.3 illustrating impeller, diffuser, and compressor manifold.

Axial flow compressor

Structure: rotor (spinning blades) and stator (non-moving blades) arranged in alternating stages.
Operation: rotor blades accelerate air rearward; stator blades behind convert the increased velocity into pressure (diffusion).
Each rotor/stator pair constitutes a pressure stage; typical stage pressure ratio is about PR_{ ext{stage}} \approx 1.2:1.
Multistage design: engines use multiple rotor/stator stages to achieve a large overall pressure rise across the entire compressor.
Airflow characteristics: as air passes through the compressor, the flow area reduces progressively; velocity remains relatively steady while pressure increases.
Spooling: driven by a turbine via a central shaft; larger engines may have multiple spools with separate shafts.

Multi-spool architectures

Large gas turbine engines often have two or more compressor sections, each with its own turbine section (two-spool or three-spool configurations).
Two-spool engine: low-pressure compressor (LPC) and high-pressure compressor (HPC) each driven by a separate turbine section; allows each compressor to operate at its optimum speed.
Three-spool engine: adds an intermediate pressure compressor/turbine stage.
Figures referenced: Figure 8.6 (two-spool engine) and Figure 8.4 (definition of a spool).
Practical implication: increased complexity but better efficiency and performance across operating regimes.

Combustion section

Primary role: burn fuel with compressed air to generate high-energy exhaust gases for thrust.
Distinguishing factor: combustion is continuous in a gas turbine engine (not a series of discrete explosions as in piston engines).
Key design challenge: maintain stable, efficient combustion across a wide range of conditions (altitude, temperature, pressure, g-load).
Diffuser before combustion: slows down the air to a suitable entry speed to avoid flame blowout when entering the combustion chamber.
Air distribution in combustion: only about 20% of the air entering the combustion zone passes the fuel nozzle initially (primary zone) to ensure proper fuel-air mixing without becoming too lean.
Ignition: a spark is usually required only at engine start; after ignition, combustion is self-sustaining.
Primary zone vs dilution zone:
- Primary zone: initial combustion area where fuel mixes with a portion of air.
- Dilution zone: the remaining 80% of air flows around the main combustion area, gradually mixing and cooling burnt gases before entering the turbine.
Air distribution varies by engine and design; fuel-air mixture control is critical for stable operation and emissions control.
Figure reference: Figure 8.7 shows airflow and components in a typical combustion chamber.

Types of combustion chambers

There are three main chamber designs, each with a thermal lining to protect engine components from extreme temperatures:

Can (Can-type): multiple cans arranged around the shaft; each can has its own duct from the compressor; post-combustion exhausts merge before reaching the turbine. Pros: easy to manufacture and service; fault in one can can be replaced easily. Cons: inefficient use of space; mostly found in older engines.
Can-Annular (Cannular): multiple cans within a larger annular chamber; improves inter-can interconnections and ignition reliability while maintaining some can-level modularity.
Annular: combustion chamber encircles the shaft, using the full annular space; highly efficient use of space and supports higher overall air flow; maintenance is more challenging.
These designs correspond to Figure 8.8.

Turbine section

Purpose: extract some energy from the hot combustion gases to drive the compressor and other components; energy taken by the turbine reduces the energy available to the exhaust, producing thrust.
Energy distribution: in modern turbofan engines, typically over 70% of the energy from combustion is used to drive the compressors and the fan; in turboprops, the turbine absorbs most energy to drive the propeller and compressor.
Structure: sequence of stationary blades (nozzle guide vanes) to direct flow and rotating turbine blades mounted on discs connected to the compressor via a central shaft.
Spool associations: multiple turbine sections may exist, corresponding to multiple spools (e.g., high-pressure turbine for HPC, low-pressure turbine for LPC, etc.).
Additional turbine sections can drive other components (e.g., fan in a turbofan, propeller in a turboprop).
Temperature challenges: turbine stage components operate in extreme temperatures, often exceeding 10^3\,^{\circ}\mathrm{C} (approximately 1800\,^{\circ}\mathrm{F}).
Material and cooling: turbine blades require heat-resistant materials; cooling channels and systems are used to protect blades, as discussed in later chapters.
Figure reference: Figure 8.9 shows a turbine section on a two-spool engine.

Turbine temperatures and cooling

Turbine section temperatures frequently exceed 1000°C (1800°F).
Materials: special heat-resistant alloys and coatings are used.
Cooling: blades may incorporate cooling passages; Chapter 10 covers turbine blade cooling in more depth.

Exhaust section

Function: maximize thrust by shaping and accelerating gas flow as it exits the engine.
Components: tailpipe, tail cone, and exhaust nozzle.
Flow management: collect and straighten exhaust gases after the turbine, then slightly slow the flow in a wider section to reduce friction losses against the tail cone.
Nozzle effect: near the end of the exhaust, a convergent nozzle (narrowing) increases gas velocity to help increase thrust.
Bypass air interaction: in low bypass turbofan engines, cold bypass air is often mixed with hot exhaust within the exhaust section; in high bypass turbofans, this mixing typically occurs after the nozzle (in the nozzle area or downstream of it).
Figure reference: Figure 8.10 shows the exhaust section arrangement.

Overall design considerations and real-world relevance

Each section is carefully designed to maximize thrust while withstanding extreme conditions (high temperatures, high speeds, and strong mechanical stresses).
The chapter focuses on subsonic aircraft but notes that supersonic engine designs share core elements with additional adjustments.
The evolution from turbojets to turbofans and turboprops reflects efficiency, noise, and performance trade-offs across real-world applications.
EPR (Engine Pressure Ratio) is a key instrument for assessing thrust command and engine health; misreadings can lead to unsafe performance, as illustrated by the 1982 Boeing 737 case study.
Practical implications include ice protection systems, bleed air usage, and the need for reliable instrumentation (e.g., intact EPR sensors) to ensure safe operation in adverse weather conditions.

Quick reference formulas and numerical notes

Engine Pressure Ratio (EPR):
EPR = rac{P{ ext{exit}}}{P{ ext{inlet}}}
Centrifugal compressor typical overall pressure rise:
PR_{ ext{centrifugal}} \approx 7:1
Axial-flow compressor stage pressure ratio:
PR_{ ext{stage}} \approx 1.2:1
Typical turbine temperatures: often exceed
T \gtrsim 1000^{\circ}C \(\approx 1800^{\circ}F)
Multistage compressor concept: engines may have multiple stages; two-spool or three-spool configurations distribute loads for optimal speed matching across compressors and turbines.

Connections to real-world practice and safety implications

Intake design (divergent vs straight) affects ram recovery and overall engine efficiency; proper icing protection is essential to prevent power loss or flameout.
Can, Cannular, and Annular combustion chamber designs reflect trade-offs between ease of maintenance, space efficiency, ignition reliability, and flow capacity.
Bleed air use for anti-ice and other aircraft systems demonstrates how core engine physics support auxiliary functions, requiring robust thermal management.
Ice-induced sensor blockage (as in the 1982 Boeing 737 incident) highlights the importance of redundant instrumentation and proper operation of ice protection systems for safe thrust management and climb performance.