Engines: Applied Laws Flashcards
What is thermodynamics?
Thermodynamics is the study of heat/pressure energy or the behavior of gases (including air) and
vapors under variations of temperature and pressure.
Explain Bernoulli’s theorem
Bernoulli’s theorem is that the total energy in a moving fluid or gas is made up of three forms of
energy:
1. Potential energy (the energy due to the position)
2. Pressure/temperature energy (the energy due to the pressure)
3. Kinetic energy (the energy due to the movement)
When considering the flow of air, the potential energy can be ignored; therefore, for practical
purposes, it can be said that the kinetic energy plus the pressure/temperature energy of a smooth flow
of air is always constant. Thus, if the kinetic energy is increased, the pressure/temperature energy
drops proportionally, and vice versa, so as to keep the total energy constant. This is Bernoulli’s
theorem
Explain a venturi
A venturi is a practical application of Bernoulli’s theorem, sometimes called a convergent/divergent
duct.
A venturi tube has an inlet that narrows to a throat, forming a converging duct and resulting in (1)
velocity increasing, pressure (static) decreasing, and (3) temperature decreasing. The outlet section is
relatively longer with an increasing diameter, forming a diverging duct and resulting in (1) velocity
decreasing, (2) pressure (static) increasing, and (3) temperature increasing.
For a flow of air to remain streamlined, the mass flow through a venturi must remain constant. To
do this and still pass through the reduced cross section of the venturi throat, the speed of flow through the throat must be increased. In accordance with Bernoulli’s theorem, this brings about an accompanying drop in pressure and temperature. As the venturi becomes a divergent duct, the speed reduces, and thus the pressure and temperature increase.
What is the theory of a jet/gas turbine engine?
Frank Whittle described the theory behind the jet engine as the balloon theory: “When you let air out
of a balloon, a reaction propels the balloon in the opposite direction.” This, of course, is a practical
application of Newton’s third law of motion.
A jet/gas turbine produces thrust in a similar way to the piston engine/propeller combination by
propelling the aircraft forward as a result of thrusting a large weight of air rearward.
Thrust = air mass × velocity
Early jet engines adopted the principle of taking a small mass of air and expelling it at an
extremely high velocity. Later gas turbine engines have evolved into taking and producing a large
mass of air and expelling it at a relatively slow velocity (e.g., high-bypass engine).
What is specific fuel consumption (SFC)?
Specific fuel consumption is the quantity/weight (lb) of fuel consumed per hour divided by the thrust
of an engine in pounds:
fuel burn (lbs) per hour/engine thrust (lbs)
What is the combustion cycle of a jet/gas turbine engine?
The combustion cycle of a jet/gas turbine engine is induction, compression, combustion, expansion,
and exhaust. In a jet/gas turbine engine, combustion occurs at a constant pressure, whereas in a piston engine, it occurs at a constant volume.
Why was the jet/gas turbine engine invented?
Frank Whittle invented the jet aircraft engine as a means of increasing an aircraft’s attainable altitude,
airspeed, reliability, and, to a lesser extent, maneuverability for the military.
Frank Whittle designed the jet/gas turbine engine for two main reasons:
1. To achieve higher altitudes and thus airspeed because propeller aircraft had limited altitude and
speed capabilities.
2. As a more simplistic and therefore reliable engine because the piston engine was a very
complicated engine with many moving parts and thus was unreliable.
Describe how a jet/gas turbine engine works.
The jet engine or aerothermodynamic duct, to give it its real name, has no major rotating parts and
consists of a duct with a divergent entry and convergent or convergent/divergent exit. When forward
motion is imparted to it from an external source, air is forced into the engine intake, where it loses
velocity or kinetic energy and therefore increases its pressure energy as it passes through the
divergent duct. The total energy is then increased by the combustion of fuel, and the expanding gases
accelerate to atmosphere through the outlet converging duct, thereby producing a propulsive jet.
A jet engine is unsuitable as an aircraft power plant because it is incapable of producing thrust at
low speeds. That is, it requires a forward motion itself before it produces any thrust.
The gas turbine engine has avoided the inherent weakness of the jet engine by introducing a
turbine-driven compressor that produces thrust at low speeds. Therefore, the aircraft power plant is
in fact a gas turbine engine and subsequently will be referred to as such.
The gas turbine is essentially a heat engine using air as a working fluid to provide thrust by
accelerating air through the engine and increasing its kinetic energy. To obtain this increase, the
pressure energy is increased first by a compressor, followed by the addition of heat energy in the
combustion chamber, before its final conversion back to kinetic energy in the form of a high-velocity
jet efflux across the turbine. (This provides extra shaft power to either drive a conventional frontal
propeller or fan or to compress extra air to provide more jet flow, as in a ducted fan and bypass
engine.) The airflow is then finally exhausted through the exhaust nozzle duct.
What is the mechanical arrangement of the gas turbine,
compressor, combustion, turbine, exhaust, is in
series so that the combustion cycle occurs continuously at a constant pressure.
What is a fuel injection system, and what are its advantages and disadvantages?
A fuel injection system delivers metered fuel directly into the induction manifold and then into the
combustion chamber (or cylinder of a piston engine) without using a carburetor. Normally, a fuel
control unit (FCU) is used to deliver metered fuel to the fuel manifold unit (fuel distributor). From
here, a separate fuel line carries fuel to the discharge nozzle in each combustion chamber (or cylinder head in a piston engine, or into the inlet port prior to the inlet valve). With fuel injection, a separate fuel line can provide a correct mixture.
Advantages:
- Freedom from vaporization ice (fuel ice)
-More uniformed delivery of the fuel-air mixture around the combustion chamber section or each
cylinder
-Improved control of fuel-air ratio
-Fewer maintenance problems
-Instant acceleration of the engine after idling, i.e., instant response
-Increased engine efficiency
Disadvantages:
-Starting an already hot fuel injection engine may be difficult due to vapor locking in the fuel lines.
-Having very fine fuel lines, fuel injection engines are more susceptible to contamination (i.e., dirt or
water) in the fuel.
-Surplus fuel provided by the fuel injection system will pass through a return line, which is usually
routed to only one of the fuel tanks. This may result in either the fuel being vented overboard (thus
reducing fuel available) or asymmetric (uneven) fuel loading.
What are thrust reverses, and how do they work?
Thrust reverses on jet/gas turbine engine reverse the airflow forward, thereby creating a breaking
action. There are two types of thrust reverses: (1) blockers or bucket design and (2) reverse flow
through the cascade vane.
Describe maximum takeoff thrust and its limitations.
Maximum takeoff thrust is simply the maximum permissible engine thrust setting for takeoff,
expressed either as an N1 or engine pressure ratio (EPR) figure.
Maximum takeoff thrust is the highest thrust setting of the aircraft’s engine when the highest
operating loads are placed on the engine. However, as a protection to the engine, maximum takeoff
thrust settings have a time limit on their use, namely, 5 minutes for all engines working and 10 minutes
with an engine failure.
Note: Some authorities allow a 10-minute time limit with all engines operating
Describe maximum continuous thrust.
Maximum continuous thrust is simply the maximum permissible engine thrust setting for continuous
use, expressed either as an N1 or engine pressure ratio (EPR) figure.
What is the compression ratio of a gas turbine engine?
The compression ratio of a gas turbine engine is a ratio measure of the change in air pressure between
the inlet and outlet parts of either an individual compressor stage or the complete compressor section
of the engine.
Individual compressors, either centrifugal or axial-flow types, are placed in series so that the
power compression ratio accumulates.
What is the principle of the bypass engine?
The principle of the bypass engine is an extension of the gas turbine engine that permits the use of
higher turbine temperatures to increase thrust without a corresponding increase in jet velocity by
increasing the air mass/volume intake and discharge to atmosphere via the bypass ducts. Remember,
Thrust = air mass × velocity
The bypass engine involves a division or separation of the airflow. Conventionally, all the air
entering into the engine is given an initial low compression, and a percentage is then ducted to bypass the engine core. The remainder of the air is delivered to the combustion system in the usual manner. The bypass air is then either mixed with the hot airflow from the engine core in the jet pipe exhaust or immediately after it has been discharged to atmosphere to generate a resulting forward thrust force.
The term bypass normally is restricted to engines that mix the hot and cold airflow as a combined
exhaust gas. This improves (1) propulsive efficiency and (2) specific fuel consumption and (3)
reduces engine noise (this is due to the bypass air lessening the shear effect of the air exhausted
through the engine core).
What is bypass ratio?
Bypass ratio in an early single- or twin-spool bypass engine is the ratio of the cool air mass flow
passed through the bypass duct to the air mass flow passed through the high-pressure system.
Typically, this early evolution of the bypass engine has a low bypass ratio, i.e., 1:1.
Alternatively, bypass ratio for a fan-ducted bypass engines is the ratio of the total airmass flow
through the fan stage to the airmass flow that passes through the turbine section/high-pressure (engine core) system. A high bypass ratio, i.e., 5:1, is usually common with ducted fan engines.
Describe the fan engine and its advantages.
The fan engine can be regarded as an extension of the bypass engine principle with the difference that it discharges its cold bypass airflow and hot engine core airflow separately.
The turbine-driven fan is in fact a low-pressure axial-flow compressor that provides additional
thrust. Normally, the fan is mounted on the front of the engine and is surrounded by ducting that
controls the high supersonic airflow speeds experienced at the blade tips, preventing them from
suffering from compressibility effect losses.
The fan is either coupled to the front of a number of core compression stages (twin spool engine),
which restricts the width size of the fan, and its bypass air is ducted overboard at the rear of the
engine through long ducts, or it is mounted on a separate shaft driven by its own turbine (triple spool
engine) where the bypass airstream is ducted overboard directly behind the fan through short ducts, hence the term ducted fan. For example, the CFM56-3 is a twin spool fan engine, and the Rolls
Royce RB211 is a triple spool fan engine.
The fan design reflects the specific requirements of the engine’s airflow cycle and gives an initial
compression to the intake air before it is split between the engine core and the bypass duct. The fan is
capable of handling a larger airflow volume than the high-pressure compressor. Therefore, a fan
engine normally will have a high bypass ratio, which means the engine’s resultant thrust properties
are dominated by the large bypass air mass; therefore, the advantages of the bypass engine are
increased further for the fan engine.
The following are some of the main advantages of a fan engine:
1. Smaller engine size.
2. Better propulsive efficiency.
3. Better specific fuel consumption.
4. Reduction in engine noise.
5. Contamination (i.e., bird strikes, heavy water) are centrifugally discharged through the bypass duct,
therefore protecting the main engine core from damage and even a flame out from water
contamination.
What are the advantages of a wide-chord fan engine?
The advantages of a wide-chord fan engine are better fuel economy, more thrust, and less weight and noise. A wide-chord fan engine is a term used to describe a modern turbofan jet engine having a ducted
fan with specific blade geometry, namely wider blades. This technology was pioneered by Rolls
Royce in the 1970s.
Designers refined the blade design by making the blade chord wider, altering the blade geometry,
manufacturing them with hollow cross-sections, and by using lighter materials, such as titanium, to
extract more thrust for any given fan area.
Describe a triple-spool turbofan engine, e.g., the RB211, and its advantages.
A triple-spool turbofan engine such as the Rolls Royce RB211 is a further development of the fan
engine that has two distinct differences from the twin-spool fan engine (see Q: Describe the fan
engine and its advantages, page 66):
1. The triple-spool turbofan engine has three independent compressor spools:
N1
, the low-pressure compressor spool or fan
N2
, the intermediate-pressure compressor
N3
, the high-pressure compressor spool and they are each driven by their own turbine and
connecting shafts.
2. The front turbofan, or N1
low-pressure compressor spool, is not connected to any other
compression stages.
The turbofan on a triple-spool engine is further improved because it is not restricted to the size of
other compressor spools (as it is on a twin-spool engine) and it is driven at its optimal speed by its
own turbine. This allows it to have a larger frontal area that consists mainly of a giant ring of large
blades, which act more like a shrouded prop than a fan. It is responsible for producing an even larger
bypass ratio (i.e., 5:1), which generates approximately 75 percent of the engine’s thrust in the form of
bypass airflow delivered to the atmosphere via the engine’s bypass ducts behind the fan.
The one part of air that flows through the engine N2 and N3 compressors becomes highly
compressed, of which one-third is used for combustion and two-thirds is used for internal engine
cooling.
Advantages of the triple-spool fan engine, including the RB211, are twofold:
1. Particular to the triple-spool configuration, including the RB211:
a. The N1
fan compressor can be built closer to its optimal design, namely, a wider chord, because
it is not restricted by any connection to booster compressors. (See Q: What are the advantages
of a wide- chord fan? page 67.)
b. The N1
, N2
, and N3 compressor sections all work closer to their optimal performance levels
because they have their own independent turbine connecting shafts, especially the N1 spool.
c. The triple-spool configuration allows more flexibility due to the aerodynamic matching at part
load and lower inertia of the rotating parts.
d. Higher engine thrust output due to the improved fan section (point a) and the improved
independent spool configuration (points b and c).
e. Easier to start because only one spool needs to be turned by the starter.
f. The triple spool’s modular assembly makes it easier to build and in particular to maintain; i.e., if
the N3 compressor suffers a fault, then the modular assembly of the engine allows for the N3
section of the engine alone to be removed for repair.
2. Advantages common to fan engines. (See Q: Describe the fan engine and its advantages, page
66.)
Why is a fan engine flat rated?
The fan engine is flat rated to give it the widest possible range of operation, keeping within its
defined structural limits, especially in dense air.
Note: Flat rating guarantees a constant rate of thrust up to a fixed temperature, namely, the warmest temperature at which the engine can produce its maximum-rated thrust. This temperature usually corresponds to the performance TREF
temperature.
When and where is a jet/gas turbine (bypass) engine at its most efficient, and why?
At high altitudes and high rpm speeds
Why does a jet aircraft climb as high as possible?
Jet aircraft climb as high as possible (i.e., to their service ceiling) because the gas turbine (bypass)
engines are most efficient when their compressors are operating at a high rpms—approximately 90 to
95 percent. This high-rpm speed results in the engine’s optimum gas flow condition that achieves its
best specific fuel consumption (SFC). However, this high-rpm speed can be achieved only at high
altitudes because only at high altitudes, where the air density is low, will the thrust produced be low
enough to equal the required cruising thrust.
The primary reason for designing an engine’s optimum operating condition at approximately 90 to
95 percent rpms is to make it coincident with the best operating conditions of the airframe, namely,
minimum cruise drag.
Therefore at high altitudes, there are two main consequences:
1. Minimum cruise airframe drag. This is experienced at high altitudes because drag varies only
with equivalent airspeed (EAS), i.e., as EAS decreases, drag decreases. At very high altitudes,
i.e., above 26,000 ft, the Mach number (MN) speed becomes limiting and therefore EAS and true
airspeed (TAS) are reduced for a constant MN with an increase in altitude. (See Q: Describe
Mach number? page 122.) Therefore, the lowest cruise EAS is at the highest attainable altitude
(service ceiling), and because drag varies only with EAS, airframe drag is also at its lowest value
at high altitudes. Consequently, the thrust requirements are lower at high altitudes because the
thrust value must only be equal to the drag value.
2. Best engine SFC. This is experienced at high altitudes due to the engine’s ability to operate at its
optimum high-rpm condition because of the low atmospheric air density/temperature. The engine’s
best SFC operating conditions are a function of its internal aerodynamic design. This reflects the
optimization of the engine to be generally at its best under the conditions where it will spend most
of its working life, i.e., high altitudes, high-speed conditions, and comparatively high engine rpm
setting.
High-altitude conditions optimize the engine’s design by using the reduction in the atmospheric air
density as a reduced airflow mass into the engine for a given engine rpm speed with an increase in
altitude. The fuel control system adjusts the fuel delivery to match the reduced mass airflow to
maintain a constant mixture and so maintains a constant engine speed. This causes the thrust to fall for
a given rpm speed and requires an increase in compressor rpms to maintain its thrust values with an
increase in altitude until its optimum high-rpm speed is reached.
In addition, the required thrust is lower with an increase in altitude because the EAS and airframe
cruise drag reduce with altitude. Therefore, it follows that only at high altitudes will the thrust be low
enough to equal the required thrust at the engine’s optimal normal cruising high engine rpm, which
achieves its best SFC. (See Q: What advantages does a jetengined aircraft gain from flying at high
altitudes? page 71.)
What advantages does a jet-engined aircraft gain from flying at a high altitude?
The advantages a jet engine gains from flying at high altitudes are
1. Best specific fuel consumption (SFC)/increased (maximum) endurance
Note: Endurance is the need to stay airborne for as long a time as possible for a given quantity of
fuel. Therefore, the lowest SFC in terms of pounds of fuel per hour is required.
2. Higher true airspeed (TAS) for a constant indicated airspeed (IAS), providing an increased
(maximum) attainable range
Note: Maximum attainable range is the greatest distance over the ground flown for a given quantity
of fuel, or the maximum air miles per gallon of fuel.
1. Best SFC and thus increased (maximum) endurance are achieved at high altitudes because of two
effects:
a. Minimum cruise drag is experienced at high altitudes because the Mach number (MN) speed
becomes limiting above approximately 26,000 ft, and for a constant MN (as is the normal
operating practice), the TAS and equivalent airspeed (EAS) decrease with altitude, and drag
varies only with EAS. As such, the EAS is reduced progressively to a level closer to the
aircraft’s best endurance speed, the higher the altitude, which is obviously where drag is least.
Note: Minimum drag speed (VIMD), broadly speaking, remains constant with altitude. And
because minimum aircraft drag requires minimum thrust (i.e., thrust = drag), and given that thrust is a
product of engine power and fuel consumption is a function of engine power used, the aircraft thus has
its lowest fuel consumption in terms of fuel used per hour. Hence it produces the maximum endurance
flight time for a given quantity of fuel when flying at its lowest cruise EAS (best endurance speed),
which the highest attainable altitude will achieve for a constant MN.
b. In addition, an aircraft’s fuel consumption decreases slightly at high altitudes because of the
higher propulsive efficiency of the engine (see preceding question), which brings it closer to its
best SFC operating condition. Therefore, an aircraft’s best SFC and maximum endurance are
attained by flying at
(1) The best endurance speed (VIMD)
(2) Highest possible altitude where the engine achieves its best propulsive efficiency
2. The higher the altitude, the greater is the TAS for a constant IAS, which provides an increased
(maximum) attainable range. Maximum range is also defined by an EAS that is a slightly higher
speed than the best endurance speed because the benefits of the increased IAS (i.e., greater ground
speed/distance covered) outweighs the associated increased drag and higher fuel consumption.
(See Q: Define maximum endurance and range with reference to the drag curve, page 209.)
The higher TAS for a constant IAS results simply from the reduction in air density at higher
altitudes. Thus ground speed/ground distance covered and range are increased (or the flight time is
reduced for a given distance) at high altitudes for a constant IAS that gives a higher TAS.
Thus the rule of thumb for best range is: The higher the better. Just how high depends on other
factors, such as winds at different levels and sector lengths.
Note: Normal operating practice for jet aircraft will be to have a limiting MN speed above
approximately 26,000 ft. Therefore, a constant MN is flown above 26,000 ft, which would see TAS
decrease with altitude because IAS and the local speed of sound (LSS) decrease to maintain a
constant MN. (See Qs: Describe local speed of sound (LSS) and Mach number (MN), page 122;
How does temperature af ect local speed of sound (LSS), page 122; A flight carried out below
optimal altitude has what result on jet performance? page 211.)
Therefore, because TAS decreases at high altitude for a constant MN, this results in the ground
speed being reduced slightly, and thus obtainable range also will be reduced slightly below its
maximum range and/or its flight time will increase for still-air conditions. To counter this, a slightly
higher long-range cruise MN speed (and thus IAS) can be selected that increases the TAS for the
altitude, although this would be detrimental to endurance.
In basic terms of best SFC and endurance and greater TAS and range, an aircraft should remain as
high as possible for as long as possible. (See Q: What results does a flight carried out below its
optimum altitude have on a jet performance? page 211.)
Explain the jet/gas turbine engine’s thrust-to-thrust lever position.
The thrust lever produces more engine thrust from its movement near the top of its range than the
bottom.
An engine’s operating cycle and gas flow are designed to be at their most efficient at a high-rpm
speed, where it is designed to spend most of its life. (See Q: When and where is a jet/gas turbine
engine at its most ef icient and why? page 69.) Therefore, as rpms rise, mass flow, temperature, and
compressor efficiency all increase, and as a result, more thrust is produced, say, per 100 rpm, near
the top of the thrust lever range than near the bottom.
In practical terms, this translates to differing thrust output per inch of thrust lever movement; i.e., at
low-rpm speed (near the bottom of its range), an inch movement of the thrust lever could produce
only 600 lb of thrust, but at a high-rpm speed (near the top of its range), an inch movement of the
thrust lever typically could produce 6000 lb of thrust.
For this reason, if more power is required at a low thrust lever setting, then a relatively large
movement/opening of the thrust levers is required, i.e., when initiating a go-around/overshoot.
However, if operating at the top of the thrust lever setting, a large reduction or increase in thrust
would only require a relatively small movement of the thrust lever.
Obviously, an appreciation of the jet/gas turbine engine’s response characteristics helps the pilot’s
understanding and operation of his or her engines.
What are the main engine instruments?
The main primary engine instruments usually are
1. Engine pressure ratio (EPR) gauge (thrust measurement)
2. N1 gauge (low compressor rpms)
3. Exhaust gas temperature (EGT) or total gas temperature (TGT) (engine temperature)
Other possible primary engine instruments are
4. N2 gauge (intermediate compressor rpms)
5. Fuel flow (fuel flow indicator)
Secondary engine instruments usually are
6. Oil temperature gauge, pressure gauge, and quantity gauge
7. Engine vibration meter
What is engine pressure ratio (EPR)?
Engine pressure ratio (EPR) is the ratio of air pressure measurements taken from two or three
different engine probes and displayed on the EPR gauge for the pilot to use as a parameter for setting
engine thrust.
The EPR reading is the primary engine thrust instrument, with the temperature of the turbine stage
governing the engine’s maximum attainable thrust.
Normally, EPR on a gas turbine–powered aircraft is a ratio measurement of the jet pipe pressure
to compressor inlet ambient pressure or sometimes the maximum compressor cycle pressure to
compressor inlet ambient pressure. However, on a fan engine, the EPR is normally a more complex
ratio measurement of an integrated turbine discharge and fan outlet pressure to compressor inlet pressure.
What is exhaust gas temperature (EGT), and why is it an important engine parameter?
EGT is exhaust gas temperature and is an important engine parameter because it is a
measure/indication of the temperatures being experienced by the turbine.
The only real operating threat to the engine’s life is excessive turbine temperatures. The maximum
temperature at the turbine is critical because if the EGT limit is exceeded grossly on startup (hot
start), the excessive temperatures will damage the engine, especially the turbine blades. Also, if the
cruise EGT limit is exceeded slightly for a prolonged period, this will shorten the engine’s life.
Describe an engine wet start and its causes, indications, and actions.
An engine wet start is otherwise known as a failure to start after the fuel has been delivered to the
engine. The cause of a wet start normally is an ignition problem.
Indications of a wet start are
1. Exhaust gas temperature (EGT) does not rise.
2. Revolutions per minute (rpms) stabilize at starter maximum.
Actions required for a wet start include
1. Close the fuel lever/supply as soon as a wet start is diagnosed (usually at the end of the starter
cycle).
2. Motor over the engine to blow out the fuel (approximately 60 seconds).
Describe an engine hung start and its causes, indications, and actions.
A hung start occurs when the engine ignites but does not reach its self-sustaining rpms. (Selfsustaining speed is an rpm engine speed at and above which the engine can accelerate on its own
without the aid of the starter motor.)
The cause of a hung start is insufficient airflow to support combustion due to the compressor not
supplying enough air because of one or a combination of the following, but not restricted too:
1. High altitude, low-density air
2. Hot conditions, low-density air
3. Inefficient compression
4. Low starter rpms
Indications of a hung start include
1. High exhaust gas temperature (EGT), above normal
2. Engine rpm below normal self-sustaining speed
Actions required for a hung start are
1. Close fuel lever/stop fuel delivery.
2. Motor over the engine to blow out the fuel (for approximately 60 seconds)
Note: To gain a successful start in hot and high conditions, you have to introduce more air into the
engine. Adjusting the fuel supply does not help.
Increasing fuel = rpm decreases & EGT increases
Decreasing fuel = rpm increases & EGT decreases
Describe an engine hot start and its causes, indications, and actions.
A hot start is one in which the engine ignites and reaches self-sustaining rpms, but the combustion is
unstable and the exhaust gas temperature (EGT) rises rapidly past its maximum limit.
Causes of a hot start include
1. Overfueling (throttle open)
2. Air intake/exhaust blocked
3. Tailwind, causing the compressor to run backward
4. Seized engine, e.g., ice blockage
Indication of a hot start is an EGT rising rapidly toward its maximum limit.
Actions required for a hot start are
1. Close fuel lever/stop fuel delivery before the EGT limit has been reached.
2. When the engine rpms have slowed to the reengagement speed, motor over the engine to blow out
the fuel (approximately 60 seconds).
What is a variable/reduced thrust takeoff (flex)?
A variable/reduced thrust takeoff uses the takeoff thrust (EPR/N1
) required for the aircraft’s actual
takeoff weight, which is a reduced thrust value from the maximum takeoff weight thrust value that
meets the aircraft’s takeoff and climb performance requirements with one engine inoperative.
The full takeoff thrust is calculated against an aircraft’s performance-limited (either field length,
WAT, tire or net flight path, obstacle clearance climb profile) maximum permissible takeoff weight
and not its actual takeoff weight. The reduced takeoff thrust is the correct thrust setting for the
aircraft’s actual takeoff weight that achieves the aircraft’s takeoff and climb, one engine inoperative,
performance requirements.
A reduced/variable takeoff thrust is calculated by using the assumed/flexible temperature
performance technique (see Q: What is an assumed/flexible temperature? page 201).
Using this variable takeoff engine pressure ratio (EPR) means that the aircraft is now operating at
or near a performance-limiting condition, whereby following an engine failure the whole takeoff
would still be good enough in terms of performance. In fact, a lower weight aircraft using a variable
thrust will mirror the takeoff run/profile of a maximum takeoff weight aircraft using a full thrust
setting.
Note: A function of reduced thrust (assumed/flexible) takeoff is as follows. A higher
assumed/flexible temperature relates to a lower air density delivered to the engine. The fuel flow is
reduced correspondingly to maintain the correct air-fuel mixture. This produces a lower theoretical
thrust from the engines, and practically this is even lower because the actual air density is greater and
therefore the mix is air dominant and less explosive. This lower thrust therefore makes the aircraft’s
momentum more dominant in reaching its V1 and VR
speeds. This greater dependence on momentum
requires a longer runway (TOR/D) for a given weight, or it relates to the same takeoff run (TOR)
performance (i.e., rotation point) for a lower weight aircraft using a reduced takeoff thrust against a
maximum takeoff weight aircraft using a maximum thrust setting.
A variable/reduced thrust takeoff can be employed on most occasions when the aircraft’s takeoff
weight is lower than the performance-limited maximum takeoff weight for the ambient air
temperature. Its use is optional, but the higher EPR required for the maximum takeoff weight always
must be available whenever a variable/reduced thrust takeoff is being conducted. It is recommended that full takeoff power be restored in the event of an engine failure above V1 or whenever deemed necessary. The use of maximum takeoff weight (MTOW) EPR for an aircraft with a lower actual weight obviously will provide the aircraft with a better all-round takeoff performance.
It should be clearly understood that a reduced/variable thrust takeoff is a reduced thrust in relation to the thrust setting required for the maximum permissible takeoff weight. It is in fact the correct thrust takeoff setting for the aircraft’s actual weight that meets its performance requirements.
Can a maximum takeoff weight aircraft use a reduced takeoff technique?
Yes, a reduced-thrust takeoff can be used even when an aircraft is at its maximum takeoff structural weight, providing the TOR/D is not limiting. This is so because you can trade momentum gained from a longer TOR/D to achieve the V1 and VR speeds at the performance-limiting conditions for a lower thrust setting.
Why do you use variable/reduced thrust (flex) takeoffs in a jet aircraft?
There are two main reasons for using a variable/reduced thrust takeoff:
1. To protect engine life and to improve engine reliability. Variable/reduced thrust takeoffs reduce
the stress and attrition of the engine during the takeoff period when the highest loads are placed on
the engine.
2. To reduce the noise generated by the aircraft. (Noise suppression of this type normally is used
for takeoff and occasionally on approaches over noise abatement areas, as well as for nighttime
flying noise restriction.)
Why is the risk per flight decreased with a reduced-thrust takeoff?
When a reduced thrust is used for takeoff, the risk per flight is decreased because of the following
main reasons:
1. The assumed/flexible temperature method of reducing thrust to match the takeoff weight does so at a constant thrust-weight ratio, making the actual takeoff distance and takeoff run distance from the
reduced-thrust setting less than that at full thrust and full weight by approximately 1 percent for
every 3°C that the actual temperature is below the assumed temperature.
2. The acceleration-stop distance is further improved by the increased effectiveness of full-reverse thrust at the lower temperature.
3. The continued takeoff after engine failure is protected by the ability to restore full power on the
operative engine.
What are the limitations of a variable/reduced thrust (flex) takeoff?
Clearly, a variable/reduced (flex) thrust takeoff can only be used when full takeoff thrust is not
required to meet the various performance requirements on the takeoff and initial climb-out. Therefore,
the limitations of using a reduced thrust takeoff are
1. Not maximum takeoff weight limited by
a. Takeoff field length
b. Takeoff weight-altitude-temperature (WAT) curve (engine-out climb gradient at takeoff thrust)
c. Net takeoff flight path (engine-out obstacle clearance)
2. Maximum outside air temperature (OAT) limitation
Where the proposed takeoff weight is such that none of the preceding considerations are limiting,
then the takeoff thrust may be reduced until one of the considerations listed below becomes limiting.
Note: Based upon one engine inoperative.
3. Furthermore, reduced (flex) thrust takeoffs are limited by
a. The reduced (flex) thrust must not be reduced by more than a set amount (engine specific), e.g.,
25 percent below the maximum takeoff weight full-rated takeoff thrust.
Note: This is normally restricted by a maximum flex or assumed temperature constraint, i.e., Tmax
flex. That is,
Flex temperature < Tmax
b. The reduced (flex) temperature cannot be lower than TREF
(flat rating cutoff temperature that
guarantees a constant rate of thrust at a fixed temperature, which equates to the climb thrust); i.e.,
flex temperature > TREF
c. The reduced (flex) thrust cannot be lower than the maximum continuous thrust used for the final
takeoff flight path.
d. The actual OAT; i.e., flex temperature > OAT (see Q: What is an assumed/flexible
temperature? page 201).
4. Reduced/variable/flexible takeoffs should not be used on the following occasions:
a. On an icy or very slippery runway
b. On contaminated runways, i.e., precipitation covered, etc.
c. When reverse thrust is inoperative
d. When the antiskid system is inoperative
e. When an increased V2 procedure is used in order to improve an obstacle-limited takeoff weight
What happens to engine pressure ratio (EPR) on the takeoff roll?
For the purpose of this question, let us assume that EPR is a measure of the jet pipe exhaust pressure
(P7
) against compressor inlet ambient pressure (P2
). Prior to opening the throttle levers, the EPR
reading will be very low, if not a 1:1 ratio. Now, when the throttle levers are advanced, the EPR
reading will decrease initially because the P2 pressure increases against a constant or slowly
increasing P7 value; therefore, the ratio decreases before steadily increasing to its takeoff setting.
There are two reasons for this. First, the engine compressor, combustion, and turbine stages are in
series, and this leads the engine to suffer from a slow response (or lag) to a throttle input. This is so
because the greater amount of air induced into the engine takes time to move through the compressor,
combustion, and turbine stages before it is expelled from the engine with a resultant/reaction forward
force.
Second, a consequence of the engine’s slow response rate or lag is its effect on the EPR reading
because the reading of the (P7) jet pipe exhaust pressure is taken from the rear of the engine and the (P2) compressor inlet pressure reading is taken from the front of the engine. Therefore, as the engine throttles are opened up, initially the compressor inlet air pressure (psi) will increase before the jet pipe exhaust air pressure (psi) increases proportionally, thus creating an increased or initially quickly increasing P2 value for a constant or slowly increasing P7 value. This results in an initial decrease in the EPR reading. Then, as the engine accelerates along its entire length, the engine turbine and compressor rpm speeds increase, and the EPR reading increases steadily to its takeoff setting.
What is an engine windmill start, and when is it used?
A windmill start occurs when the engine is started without the aid of the starter because the
compressors are being turned by a natural airflow when airborne. This delivers the air charge to the
combustion chambers, where fuel and an ignition spark are introduced as normal for a stable engine
relight. This effect is known as windmilling, and as such, windmill starts are used to relight an engine
when airborne.
What is the purpose of engine relight boundaries?
The purpose of engine relight boundaries is to ensure that the correct proportion of air is delivered to
the engine’s combustion chamber to restart the engine in flight. For this reason, the aircraft’s flight
manual outlines the approved relight envelope of airspeed against height. This ensures that within the
limits of the envelope, the airflow ingested into the engine will rotate the compressor at a speed that
generates and delivers a sufficient volume of air into the combustion chamber to relight the engine
successfully.
Note: The approved flight envelope usually will be subdivided into starter assist and windmill
boundaries.
What causes a jet/gas turbine upset, and how do you correct it?
Disturbed or turbulent airflow will cause a jet/gas turbine engine to be upset and to stall.
This occurs because a jet/gas turbine engine is designed to operate using a clean uniform airflow
pattern that it obtains within the aircraft’s normal operating attitude. However, beyond the aircraft’s
normal angles of incidence and slip and/or in extremely severe weather turbulence, the engines can experience a variation in the ingested air’s pressure/density, volume, angle of attack, and velocity properties. This changes the incidence of the air onto the compressor blades, causing the airflow over the blades to break down and/or inducing aerodynamic vibration. This upsets the operation of the engine causing it to stall.
The stall can be identified by (1) increases in total gas temperature (TGT), (2) engine vibration,
and (3) rpm fluctuations
What is a jet engine surge, what causes it, and what are the indications?
A surge is the reversal of airflow through an engine, where the high-pressure air in the combustion
chamber is expelled forward through the compressors, with a loud bang and a resulting loss of engine
thrust.
A surge is caused when
1. All the compressor stages have stalled, e.g., bunt negative-g maneuver.
2. An excessive fuel flow creates a high pressure in the rear of the engine. The engine will then
demand a pressure rise from the compressors to maintain its equilibrium, but when the pressure
rise demanded is greater than the compressor blades can sustain, a surge occurs, creating an
instantaneous breakdown of the flow through the machine.
A surge is indicated by
1. Total loss of thrust.
2. A large increase in TGT.
The required actions in response to an engine surge are
1. Close the throttles smoothly and slowly.
2. Adjust the aircraft’s attitude to unstall the engines, which lead to the surge.
3. Slowly and smoothly reopen the throttles.
Why are bleed valves fitted to gas turbine engines?
Bleed valves are fitted to gas turbine engines for two main reasons:
1. To provide bleed (tap) air for auxiliary systems. For example:
a. Air-conditioning and cabin heating/pressurization/EFIS cooling/cargo heating
b. Engine cooling, especially
(1) The combustion chamber
(2) The turbine section
c. Accessory cooling (generator, gearbox, and other engine-driven systems)
d. Engine and wing anti-icing systems
2. To regulate the correct airflow pressures between different engine sections.
Why do gas turbine engines have auto igniters, and how do they work?
Auto igniters are used in gas turbine engines to protect against disturbed/turbulent airflow upsetting the engine. This condition is particularly common with rear-mounted engines during some abnormal and even some rather normal flight maneuvers because rear-mounted engines are placed ideally to catch any disturbed airflow generated by the wing when the airflow pattern brakes down as a result of either a high incidence of attack (e.g., prestall buffet), high-g maneuvers (e.g., steep turns), or high Mach number effect.
Auto igniters work by sensing a particular value of incidence of the aircraft, via the incidencesensing (probe) system (which is also used to activate the stick shaker and pusher), and automatically
signals on the ignition system before the disturbed airflow generated by the wing affects the engines,
thus ensuring that the engines at least continue to run, although in some cases they might surge a little.
What is FADEC?
FADEC is full authority digital engine control and is a system that automatically controls engine
functions, i.e., start procedures, engine monitoring, fuel flow, ignition system, and power levels
required. FADEC computers can be found on the A320 aircraft’s engine and on the EJ200 engines
used on military aircraft.
What fuels are used commonly for civil jet aircraft?
The fuels used for gas turbine civil aircraft engines are
1. Jet A1 (Avtar). This is a kerosene-type of fuel with a normal specific gravity (SG) of 0.8 at 15°C.
It has a medium flash point and calorific value, a boiling range of between 150 and 300°C, and a
waxing point of –50°C.
2. Jet A. This is similar to A1, but its freezing point is only –40°C.
Note: Jet A normally is available only in the United States
How do jet/gas turbine engines generate noise?
The noise generated by a jet/gas turbine engine is from the sheer effect of different displaced air
velocities. The sheer is the difference between the jet’s faster displaced air and the slower ambient
air around it.
How is jet/gas turbine engine noise controlled or reduced?
Jet/gas turbine engine noise can be reduced by the following:
1. Bypass engines. This reduces the sheer effect between the displaced slower bypass engine air and
the ambient air that reduces noise.
2. Reduced-thrust takeoff.
Jet/gas turbine engine noise can be controlled by the following:
3. Maximum-angle climb after takeoff. This allows the aircraft to get above any noise-control zones
Is there a critical engine on a jet/gas turbine aircraft (nonpropeller driven)?
There is no critical gas turbine engine if the engines are positioned symmetrically with opposing
revolution direction.
The term critical engine is historically associated with propeller-driven aircraft whose
mechanically driven propellers were not symmetrical (in particular blade effect) and gave rise to
different thrust line and moment arm values. Therefore if the jet/gas turbine engines are positioned symmetrically with opposing revolution direction, then they have no design in balance and therefore
no “mechanical” critical engine can exist.
However, aerodynamically, there can be a critical “failure” where a crosswind exists, or if you
like a critical failure engine.
Then yes, there is a critical engine, but technically this is an incorrect term, as it is a function of
environmental properties, i.e., crosswind effect, rather than the mechanical properties of the engine.
Hence the term critical failure is more appropriate. (See Q: How does a crosswind af ect the critical
engine/failure? page 60.)
Note: There can be a governor engine, i.e., an engine that is the master that sets the rpm speed for the others. But this is not a critical engine and should not be confused as such.
