Engines and Systems Flashcards
EXPLAIN Bernoulli’s Equation, given dynamic pressure, static pressure, and total pressure
Total pressure = Static pressure + Dynamic pressure (velocity)
Any incompressible fluid that passes through a convergent opening, velocity increases and pressure decreases
DESCRIBE the behavior of airflow in a nozzle
Velocity increases, pressure decreases
DESCRIBE the behavior of airflow in a diffuser
Velocity decreases, pressure increases
DESCRIBE the Brayton Cycle
The operating cycle of a gas-turbine engine, which consists of four events occuring simultatneously: intake, compression, combustion, and exhaust
DESCRIBE a gas generator
All gas turbine engines at a minimum will include a compressor, combustion chamber, and turbine. Theses components together are known as the gas generator and produce the high-energy airflow necessary for creating thrust.
DESCRIBE how airflow properties change through each section of a gas turbine engine
- Inlet: temp (+), press (+), velocity (-)
- Compressor: temp (+), press (+), velocity (+)
- Diffuser: temp (constant), press (+), velocity (-)
- Burner: temp (+), press (-), velocity (+)
- Turbine: temp (-), press (-), velocity (+)
- Exhaust: temp (-), press (-), velocity (+)
DESCRIBE engine thrust
Thust that a gas turbine engine develops is essentilaly the result of many pressure, temperature and velocity changes as airflow passes through an engine.
Gross Thrust is a measurement of thrust soley from the velocity of exhaust gases while the engine or aircraft is stationary, ignoring the velocity of air at the inlet.
Net Thrust is thrust that corrects for the effects of inlet airflow velocity
Net thrust = mass x (V final - V initial)/time
DESCRIBE the effects of airflow properties on thrust in a gas turbine engine
- Density (mass per unit volume): higher density will increase thrust
- Air Temperature: As temperature increases, density decreases, therfore thrust decreases
- Air Pressure: an increase in pressure, generally results in an increase in density, therefore thrust increases
- Altitude: Although both pressure and temperature decrease with altitude, the decrease in thrust due to decreased pressure is greater than the increase in thrust due to decreased temperature. Therefore as altitude increases, thrust decreases. At appox. 36,000 ft, temperature stabilizes and thrust decreases more rapidly.
- Airspeed: Theoretically, as airspeed increases, thrust decreases due to the decrease in acceleration of the air.
EXPLAIN ram effect in a gas turbine engine
While a decrease in velocity by itself would cause a decrease in thrust, ram effect, increases the mass and pressure of the inlet air. This offsets the decrease in acceleration and results in a neutral effect or slight increase in thrust at subsonic speeds.
At supersonic speeds, airflow becomes compressible and mass due to ram effect increases at an increasing rate. This results in significant overall increase in thrust.
DESCRIBE the cockpit thrust measuring devices
Engine Pressure Ratio (EPR) indicates the pressure ratio between the inlet and exhaust airflow. It is used by aircraft with turbojets and turbofans, which rely on the propulsive power of exhaust gases.
Torquemeter indicates shaft horsepower available to drive a propeller or rotor. Propeller or rotor driven aircraft use it to indicate power available
Tachometer is the most common gauge used by a pilot to determine engine performence. It measures speed in revolutions per minute. Gas turbine engine tachometers are calibrated in percent RPM, where 100% represents full power.
DESCRIBE inlet ducts
Designed to provide the proper amount of high pressure, turbulence-free air to the compressor and must operate with high efficiency from ground idle to possible supersonic speeds at a variety of altitudes and attitudes.
- Normally designed to act as a diffuser
- Must minimize drag
- Must minimize the intake of a boundary layer
Two basic designs:
Single entrance duct simplest and most effective inlet duct design
Divided-entrance inlet duct allows the pilot to sit lower in the fuselage and reduces friction losses due to length.
Variable geometry inlet duct uses mechanical devices such as ramps, wedges, or cones to change the shape of the inlet duct as the aircraft speed varies between subsonic and supersonic
DESCRIBE compressors
The primary function of the compressor is to supply enough air to satisfy the requirements of the combustion section. It increases the pressure of the airflow from the inlet and directs it to the burners in the quantity and at the pressures required.
Secondarily, it supplies bleed air to operate various components throughout the engine and aircraft.
Types of compressors:
- Centrifugal: consists of an impeller, diffuser, and manifold. The impeller is driven at high speeds by the turbine and accelerates the air outward toward the diffuser. The diffuser, which is stationary, decreases the velocity and increases pressure. The manifold directs the airflow to the combustion chamber
- Axial: Consists of rotor blades and stator vanes. Rotor blades are rotating, airfoil shaped blades. Stator vanes are stationary airfoil shaped blades. Each set of rotors and stators make up a stage. Axial compressors may be dual spool, with a low pressure compressor followed by a high pressure compressor, each driven by seperate turbines. Higher compression ratios can be attained with minimum total compressor weight and frontal area with a dual spool compressor.
- Axial-centrifugal: uses a combination of the axial and centrifugal flow compressor. Advantage is its small size.
In addition to rotors and stators, the compressor utilizes inlet and exit guide vanes.
- Inlet guide vanes impart a swirling motion to the air entering the compressor in the direction of engine rotation.
- Exit Guide Vanes are located at the discharge end of the compressor and straighten airflow for the diffuser.
Diffuser is located after the compressor and prepares the airflow for the burner chamber by decreasing the velocity and increasing pressure
DESCRIBE the burner section of a gas turbine engine
Airflow entering the burner consists of two types:
- Primary air: (25%) mixed with fuel for combustion
- Secondary air: (75%) flows around the chamber to cool the thin walls and control the flame. Also may be used to cool the turbine, and for afterburner operation.
The burner section contains the combustion chamber and provides the means for proper mixing of the fuel and air to assure good combustion. The chamber must add sufficient heat energy to accelerate the air and produce desired thrust and power the turbines.
DESCRIBE combustion chambers
- Can: typically found on older centrifugal compressor engines. Consists of a fuel nozzle, burner liner, and casing.
- Annular: Consists of a continuous, circular, inner and outer shroud. Fuel is introduced through a series of nozzles where it is mixed and ignited with the incoming air. Allows for uniform heat distribution and better mixing of the air and fuel.
- Can-Annular: Primarily used on high performance engines. Combines the ease of maintenance of the can type with the excellent thermodynamics of the annular type.
DESCRIBE the turbine section of a gas turbine engine
Consisting of stators and rotors, the turbine section drives the compressor and accessories. It is designed to increase airflow velocity.
Stators prepare the airflow from the combustion chamber and deflect it at a specific angle in the direction of turbine wheel rotation.
The rotor converts heat energy into mechanical energy. About 75% of the energy is used by the turbine to drive the compressor and accessories, while the remaining 25% is used for thrust.
May be single or multistage, and may have independent shafts with a low pressure and high pressure turbine.
The turbine section is the most highly stressed part of the engine. It operates at temperatures nearing 2,000 °F and rotates over 10,000 RPM.
DESCRIBE the phenomenon of creep in a gas turbine engine
Blades undergo elongation, or creep, as they are heated. Excessive temperatures over long periods of time may result in permanent blade deformation, which could cause them to fail catastrophically.
DESCRIBE the exhaust section of a gas turbine engine
The exhaust section directs the flow of hot gases rearward to cause a high exit velocity to the gases while preventing turbulence. It consists of an outer duct, and inner cone, and three or four hollow radial struts.
Two types:
- Convergent: takes relatively slow gases from the turbine section and gradually accelerates them
- Convergent-Divergent: used to accelerate the air into supersonic flow. Used after the afterburner in most fighter aircraft. Typically variable geometry.
DESCRIBE the afterburner section of a gas turbine engine
The afterburner is a method of thrust augmentation used in turbojets and turbofans to increase the maximum thrust from an engine by 50% or more. Fuel consumption may increas by 300%. Secondary air from the burner section is used for combustion in the afterburner.
Consists of many parts, but the ones you need to know are:
- Spray Bars introduce fuel, and are located in the forward section of the duct.
- Flame Holder provides a region in which airflow velocity is reduced and turbulent eddies are formed, allowing for proper mixing of fuel and air for combustion. Consists of several concentric rings with a V cross-sectional shape
- Screech Liner consists of inner sleeves that are corrogated and perforated with thousands of holes that allow the liner to reduce pressure fluctuations and vibrations.
- Variable Exhaust Nozzle converts from a convergent nozzle for subsonic operations to a convergent-divergent nozzle for supersonic operation. Commonly called “turkey feathers.”
DESCRIBE the angle of attack of compressor blades
Within the compressor, the relative wind is formed by combining the compressor rotation (RPM) and inlet airflow. The angle between the relative wind and the rotor blade’s chord line makes up the angle of attack (AOA).
DESCRIBE a compressor stall
A stall occurs when airflow over an airfoil breaks away causing the airfoil to lose lift due to excessive angle of attack. A high AOA on the compressor blades may result from increasing the rotation speed or decreases the velocity of the inlet airflow, possibly causing compressor stall.
Airflow distortions that may induce compressure stall include:
- Abrupt change in aircraft attitude
- Encountering air turbulence
- Deficiency of air velocity or volume, caused by atmospheric conditions
- Rapid throttle movement
DESCRIBE four mechanical malfunctions that can lead to a compressor stall
- Variable inlet guide vane and stator vane: Failure to change angles will cause too much or too little airflow at low engine speed.
- Fuel Control Unit: A sudden increase in fuel can cause excessive burner pressure and a back flow of air into the compressor resulting in a compressor stall.
- Foreign Object Damage: FOD that damages the blades will change their aerodynamic properties
- Variable exhaust nozzle: If the nozzle failes to open an excessive back pressure could lead to compressor stall
DESCRIBE appropriate actions a pilot can take regarding compressor stalls
- Erratic or abrupt Power Control Lever (PCL) movements should be avoided, especially at low airspeeds or high angles of attack. Advance the PCL in a smooth fashion.
- Maintain at least the prescribed minimum airspeed
- Avoid abrupt changes in aircraft attitude
- Avoid flight through severe weather and turbulence
DESCRIBE four engine design features that can be incorporated into a gas turbine engine design to minimize the potential for a compressor stall
- Variable inlet guide vane and stator vanes are installed so the AOA is changed at low engine speed. They are automatically positioned by the stator vane actuator (SVA) and controlled by the fuel fontrol unit (FCU).
- Dual/twin/split-spool axial flow compressors allow the front rotor to turn at a slower RPM than the rear rotor.
- Bleed Valves, installed near the middle or rear of the compressor, vent air, increasing airflow in the front of the compressor at low engine RPMs.
- Variable Exhaust Nozzle unloads the pressure during afterburner operation.
DESCRIBE a turbojet engine
A turbojet engine is the simplest form of gas turbine engine, consisting of the basic gas generator (compressor, combuster, and turbine) with an inlet and exhaust. It creates thrust by highly accelerating a small mass of air through the engine. 75% of the energy is used to drive the compressor and accessories, while 25% is used for thrust.
DESCRIBE the characteristics of a turbojet engine
Advantages:
- Higher specific weight (weight per pound of thrust produced)
- Higher and faster than any other engine
Disadvantages:
- Low propulsive efficiency at low speeds
- Relatively high TSFC at low altitude and low airspeeds
- Long takeoff roll required
DESCRIBE a turbofan engine
A turbofan engine is similar to a turbojet engine, but also has a ducted fan that is driven by the gas generator. The fan provides additional thrust by accelerating a fairly large mass of air around the gas generator, called bypassed or ducted air.
On average, the bypassed air produces between 30-60% of the total thrust of a turbofan engine, while the gas generator exhaust provides the remaining 40-70%.
A free or power turbine is a turbine that drives the fan, but is not connected to the gas generator,
The amount of air that bypasses the gas generator in comparison with the amount of air that passes through the gas generator is called the bypass ratio. A higher bypass ratio yeilds a lower TSFC. (more fuel-efficient)
DESCRIBE the characteristics of a turbofan engine
Advantages
- Higher thrust at low airspeeds
- Lower TSFC
- Shorter takeoff distance
- Considerable noise reduction, 10 to 20 percent over the turbojet
Disadvantages
- Higher specific weight
- Large frontal area
- Inefficient at higher altitudes
DEFINE thrust specific fuel consumption
The amount of fuel required to produce one pound of thrust.
COMPARE the thrust specific fuel consumption of turbojet engines
Relatively high TSFC at low altitude and low airspeeds
