Unit 1: Fundamentals of Gas Turbine - Subjective Questions

A simple gas turbine engine consists of an intake, compressor, combustion chamber, turbine, and exhaust nozzle.

• Intake: Captures atmospheric air and delivers it smoothly to the compressor.
• Compressor: Raises the pressure of the incoming air by consuming shaft power.
• Combustion chamber: Fuel is injected, atomized, and burned approximately at constant pressure. This greatly increases the temperature and energy of the gas.
• Turbine: Hot gases expand through the turbine and produce shaft work. A major portion of this work drives the compressor and accessories.
• Exhaust nozzle: The remaining gas energy is converted into kinetic energy, producing a high-velocity exhaust jet.

The flow sequence is:

Atmospheric air → Intake → Compressor → Combustion chamber → Turbine → Exhaust nozzle

Forward thrust is produced because the engine accelerates air rearward. According to Newton's third law, an equal and opposite reaction acts on the engine.

The ideal gas turbine operates on the Brayton cycle, which consists of four internally reversible processes:

1. Isentropic compression: Air is compressed from state 1 to state 2 in the compressor.
2. Constant-pressure heat addition: Heat is supplied from state 2 to state 3, representing combustion.
3. Isentropic expansion: Hot gas expands from state 3 to state 4 through the turbine.
4. Constant-pressure heat rejection: Heat is rejected from state 4 to state 1.

For a constant-specific-heat ideal cycle, its thermal efficiency is

where is the compressor pressure ratio and is the ratio of specific heats.

Increasing pressure ratio improves ideal-cycle efficiency. In a real engine, compressor and turbine inefficiencies, combustion pressure loss, mechanical loss, and variable specific heats reduce performance.

Consider a control volume surrounding an engine operating steadily. Let be the airflow rate, the fuel-flow rate, the flight velocity, and the exhaust velocity.

The rate of rearward momentum increase is

If the exhaust pressure differs from ambient pressure , an additional pressure force acts over nozzle exit area :

Therefore, the gross thrust equation is

Since , it may also be written as

If fuel mass is neglected and the nozzle is perfectly expanded, then and , giving

Thus, thrust depends mainly on mass flow, change in velocity, and nozzle-exit pressure mismatch.

• Gross thrust: Forward force created by the momentum and pressure of the exhaust stream:

• Ram drag: Momentum per unit time carried into the engine by incoming air:

• Net thrust: Useful propulsive force remaining after subtracting ram drag from gross thrust:

• Pressure thrust: Force resulting from a difference between nozzle-exit and ambient pressures:

When , the nozzle is perfectly expanded and pressure thrust is zero. Net thrust then depends entirely on the difference between exhaust and flight momentum.

The principal factors affecting thrust are:

• Air mass flow: Greater airflow generally produces greater momentum thrust.
• Exhaust velocity: Increasing jet velocity relative to flight velocity increases thrust.
• Flight speed: Ram drag rises with aircraft speed, although inlet ram compression may partly compensate for it.
• Ambient temperature: High inlet temperature reduces air density and mass flow, usually decreasing thrust.
• Ambient pressure and altitude: Lower pressure at altitude reduces density and mass flow.
• Compressor pressure ratio and efficiency: These determine the pressure and temperature available for combustion.
• Turbine inlet temperature: A higher permissible value generally increases specific thrust.
• Nozzle pressure ratio and efficiency: These influence exhaust velocity and pressure thrust.
• Fuel-air ratio: More heat addition can increase gas energy, subject to temperature and stability limits.
• Installation losses: Intake distortion, inlet pressure loss, exhaust loss, and accessory power extraction reduce installed thrust.

For a given inlet area and velocity, airflow is approximately

and the ideal-gas relation gives

Therefore, an increase in compressor-inlet pressure at nearly constant temperature increases density and mass flow. This generally causes:

• Increased compressor airflow
• Increased fuel flow at a maintained fuel-air ratio
• Increased exhaust mass flow
• Increased gross and net thrust

A reduction in pressure, such as that caused by high altitude or inlet pressure loss, has the opposite effect. Flight-speed ram compression can raise the total pressure at the compressor face. However, inlet friction, shock waves, boundary-layer separation, and distortion reduce total-pressure recovery and consequently reduce thrust.

An increase in flight velocity has competing effects:

• It increases inlet ram pressure, which can improve compressor-face total pressure and increase engine mass flow.
• It increases ram drag, represented by .

• Net momentum thrust is approximately

  when fuel mass and pressure thrust are neglected.

• If does not increase sufficiently, the increase in reduces the velocity difference and hence specific thrust.
• At high Mach numbers, shock and inlet losses may lower total-pressure recovery.

Turbojets retain thrust comparatively well at high speed because of ram compression and high exhaust velocity. Turboprops lose propulsive effectiveness at high speed because propeller compressibility losses become severe.

At constant pressure, air density varies inversely with absolute temperature:

When inlet temperature increases:

• Air density and mass flow decrease.

• Compressor work per unit mass increases because

• More turbine work is required to drive the compressor.

• Corrected speed and corrected flow change, moving the compressor operating point.

• Net thrust and thermal margin usually decrease.

Cold air is denser, so the engine can ingest a greater mass flow and usually generate greater thrust. Very low temperatures, however, may introduce icing risks or mechanical and control limitations. Inlet-air cooling or water injection may be used temporarily to offset hot-day thrust loss.

Increasing altitude normally lowers ambient pressure, temperature, and density. The lower density reduces engine mass flow and therefore decreases thrust. The reduction in temperature partly helps by increasing density relative to a constant-temperature case and reducing compressor work, but it does not usually overcome the effect of falling pressure.

Increasing flight Mach number produces ram compression, raising inlet stagnation pressure and temperature:

For an ideal inlet,

Ram compression can improve mass flow and gross thrust, but ram drag also rises. Real inlet shock and friction losses become important at high Mach numbers. Consequently, installed net thrust depends on the balance among ambient-density reduction, ram-pressure recovery, inlet losses, and increasing ram drag.

An afterburner or reheat system injects additional fuel into the exhaust gas downstream of the turbine. Because turbine exhaust still contains oxygen, this fuel can burn and raise the exhaust temperature substantially.

The increased temperature raises gas volume and available nozzle enthalpy, producing a higher exhaust velocity. A variable-area nozzle is normally required to accommodate the increased mass flow and prevent excessive back pressure on the turbine.

Advantages:

• Large and rapid thrust increase
• Relatively simple method of increasing thrust without enlarging the core
• Useful for take-off, combat, acceleration, and supersonic flight

Limitations:

• Very high specific fuel consumption
• Increased noise and infrared signature
• High thermal loading
• Need for flame holders and a variable nozzle
• Inefficient for continuous subsonic operation

Afterburning is therefore mainly used for short-duration thrust augmentation.

Water injection augments thrust by cooling the air entering the compressor or by cooling the mixture in the combustion region.

• Compressor-inlet injection: Evaporation lowers inlet temperature, increases air density and mass flow, and reduces compressor work.
• Combustor injection: Added water increases the mass flowing through the turbine and permits additional fuel burning while controlling turbine inlet temperature.
• Water-methanol mixture: Methanol prevents freezing and also contributes heat when burned.

The result is a temporary increase in thrust, especially during take-off in hot or high-altitude conditions.

Its disadvantages include the added mass of fluid and equipment, limited operating duration, possible corrosion or deposits, increased maintenance, and reduced efficiency. Modern high-bypass turbofans generally rely less on this method because they can provide high take-off thrust more efficiently.

Major thrust-augmentation methods include:

• Afterburning or reheat: Additional fuel is burned downstream of the turbine to increase exhaust temperature and jet velocity.
• Water or water-methanol injection: Evaporative cooling increases inlet density, reduces compressor work, and may permit greater fuel flow.
• Inlet-air cooling: Refrigeration, evaporative cooling, or fogging lowers compressor-inlet temperature and improves mass flow.
• Mass or fluid injection: Additional fluid increases the exhaust mass flow and can raise momentum thrust.
• Ejector augmentation: The primary jet entrains secondary ambient air, increasing the total accelerated mass; it is useful mainly in specialized applications.
• Variable-cycle operation: Changes in bypass flow, fan pressure ratio, or flow-path geometry allow the engine to provide higher specific thrust when required.

Each technique increases either the accelerated mass flow, the exhaust velocity, or both, but normally involves penalties in fuel consumption, mass, complexity, or component life.

Afterburning and water injection both provide temporary additional thrust, but their principles differ.

Afterburning provides a larger increase in specific thrust, whereas water injection is especially useful for recovering thrust lost because of high ambient temperature.

A turbojet passes nearly all inducted air through the core. The turbine extracts only enough work to drive the compressor and accessories, leaving substantial energy for acceleration through the exhaust nozzle.

Characteristics:

• Low mass flow compared with a turbofan of similar thrust
• High exhaust velocity and high specific thrust
• Small frontal area
• Good performance at high speed and high altitude
• Relatively low bypass ratio, ideally zero

Advantages:

• Simple flow path
• Compact engine diameter
• Suitability for fast aircraft
• Effective operation where high jet velocity is required

Limitations:

• Low propulsive efficiency at ordinary subsonic speeds
• High specific fuel consumption compared with modern turbofans
• High exhaust noise
• High exhaust temperature

Turbojets are best suited to applications that value compactness and high-speed performance more than subsonic fuel economy.

A turboprop uses a gas-generator core to drive a propeller through a turbine and reduction gearbox. Most of the useful gas energy is extracted as shaft power, while the residual exhaust produces only a small fraction of total thrust.

Equivalent power may be expressed approximately as

where is propeller shaft power and is residual jet thrust.

Characteristics:

• Accelerates a large mass of air through a relatively small velocity change
• High propulsive efficiency at low and medium flight speeds
• Strong take-off and low-speed performance
• Good fuel economy for regional and transport operations
• Requires a reduction gearbox and propeller control system
• Suffers from propeller-tip compressibility, drag, and noise at high speed

Turboprops are commonly most effective below approximately to , although the exact limit depends on propeller design.

A turbofan contains a gas-turbine core and a front-mounted fan. The fan divides inlet air into core flow and bypass flow. The core flow passes through the compressor, combustor, turbine, and nozzle, while bypass air flows around the core and is discharged through a separate or mixed nozzle.

Characteristics:

• Produces thrust by accelerating a larger mass of air through a smaller velocity increase than a turbojet.
• Has better subsonic propulsive efficiency and lower specific fuel consumption.
• Produces less jet noise because its effective exhaust velocity is lower.
• A high-bypass turbofan is suitable for subsonic transport aircraft.
• A low-bypass turbofan has higher specific thrust and is suitable for high-speed or military aircraft.

The turbine must supply power to both the compressor and fan, often through separate low- and high-pressure spools.

The fundamental distinction is the balance between mass flow and velocity rise. Turbojets produce thrust using a small mass flow with a large velocity increase, while turboprops use a large mass flow with a small velocity increase. Turbofans provide an intermediate and highly adaptable solution.

The bypass ratio is the ratio of the air mass flow passing around the core to the air mass flow passing through the core:

An increase in bypass ratio generally:

• Increases the total mass of air accelerated by the engine
• Reduces the required exhaust velocity difference for a given thrust
• Improves propulsive efficiency at subsonic speed
• Reduces specific fuel consumption
• Reduces jet noise
• Increases fan diameter, frontal area, nacelle drag, and engine mass

Low-bypass engines have greater specific thrust and smaller frontal area, making them suitable for high-speed aircraft. High-bypass engines offer superior subsonic economy and are widely used in commercial transport aircraft.

For an ideal propulsor with perfectly expanded flow and negligible fuel mass, thrust is

Useful propulsive power is

The rate of increase of jet kinetic energy is

Propulsive efficiency is therefore

Substitution and factorization give

Hence,

For a given thrust, propulsive efficiency improves when is closer to . This means that accelerating a large mass flow through a small velocity increase wastes less energy in the exhaust wake. This principle explains the high subsonic efficiency of turboprops and high-bypass turbofans.

The net thrust equation is

The exhaust momentum term is

The inlet momentum or ram-drag term is

Therefore, the momentum-thrust component is

The pressure-thrust component is

Thus, total net thrust is

Hence, the engine develops of momentum thrust, of pressure thrust, and of net thrust.