The combustion chamber, or combustor, must contain the burning mixture of air
which is being passed from the compressor
and fuel, from the fuel spray nozzles, in order to generate the maximum heat release at
a substantially constant pressure
so that the turbine receives a uniformly expanded heated and accelerated stream
of gas.
This is not an easy task but advancements are constantly being made
in combustion chamber design to enable more efficient use of fuel with less and
less pollution of the atmosphere.
Efficient combustion has been made increasingly more important because of
the rise in the cost of the fuel and also the increasing awareness of the
aviation industry and the general public of the dangers of atmospheric pollution
from the exhaust smoke.
There is a limit to the maximum temperature of the gas exiting from the
combustion chamber this is imposed by the materials from which the nozzle
guide vanes
the turbine blades are manufactured.
The slightest excursion which limit will mean distortion of the
turbine blades and the possible disintegration of the turbine with
probably catastrophic results. Modern turbine and nozzle Guide vane
materials will allow against temperature not exceeding 1700 degrees Celsius when
it exits the combustion chamber. Considering that the air which leaves
the high-pressure compressor may already have been heated to around 550 degrees
Celsius during his compression. We can add sufficient fuel to raise the
temperature of the gas exiting the combustion chamber by a further 1150
degrees Celsius before we exceed the limit temperature of 1700 degrees
Celsius.
It must also be remembered that the combustion chamber has to be capable of maintaining
stable and efficient combustion over a wide range of engine operating
conditions.
Of course, 1700 degrees Celsius would be the temperature of the
gas exiting the chamber with full power selected lower power settings would
require lower fuel flows and consequently would generate lower gas
temperatures.
You'll see from this graph that the air enters the combustion chamber
at a slightly slower rate than which it enters the intake of the engine
speeds of around 500 feet per second are not unusual.
The flame rate of kerosene that is the speed at which the leading edge of a
flame would travel through kerosene vapor is approximately 30 feet per
second.
If burning kerosene was exposed to an airstream which was travelling at 500
feet per second it would be extinguished immediately.
So something must be done to slow down the speed of the airflow after it leaves
the compressor and before it reaches the combustion chamber. Otherwise the flame
would not be sustainable.
At this position on the graph we see how this reduction in the velocity of the
air is achieved. The air is slowed down
and an added bonus it's pressure is increased after it leaves the compressor.
By passing it through a divergent duct immediately before it enters the
combustion chamber.
In fact the pressure retained at this point at the end of this divergent duct
just before the air goes into the combustion chamber is the highest in the
whole of the engine.
The reduction in velocity is still not enough however further decreases must be
achieved if the flame is not to blow out.
The air is divided after it exits the high-pressure compressor into primary,
secondary, and tertiary air flows. We will now examine each of them.
This diagram shows how the air entering the flame tube
through the snout before being divided to go through the
perforated flare
and the swell veins and into the primary zone.
The primary zone is a region of lower velocity recirculation positioned
immediately downstream of the fuel spray nozzle. It's within the zone that stable
combustion is achieved.
The primary air is approximately 20% of the total airflow coming from the
high-pressure compressor and into the combustion chamber. This is the air which
is mixed in a ratio of approximately 15 to 1 by weight with the fuel and burnt.
By being passed through the flair and the swell veins the velocity of the
primary air is reduced which must happen if the flame is not to be extinguished
and the shape and position of the flare and swirl vanes also starts the air
recirculating within the region.
The remaining 80% of the output of the high-pressure compressor air which has
not been directed through the snout goes into the space between the flame tube
and the air casing. Some of this remaining air, approximately
another 20% of the output of the high-pressure compressor, is allowed into
the flame tube through secondary air holes. This air is called secondary air
and it reacts with the primary air which is flowing through the swirl vanes to
form a toroidal vortex. The toroidal vortex stabilizes and anchors the flame
and prevents it being moved through the flame tube away from the fuel nozzle
area.
The temperature of the gas is at the center of the primary zone reaches about
2,000 degrees Celsius. This is far too hot for the materials of the nozzle
guide veins and turbine blades so a further drop in temperature is required
before the gases can be allowed to exit the combustion chamber.
The remaining 60% of the total air coming out of the high-pressure
compressor is progressively introduced into the flame tube through corrugated
joints and dilutions air holes in the frame tube. This air is called tertiary
air.
Tertiary air is used to cool both the air casing and the gas exiting the
chamber.
The type of combustion chamber which we have used to illustrate the interaction
of the flame and the cooling air is representative of those which would have
been used in an early multiple combustion chamber system.
Different methods of keeping the combustion chambers from overheating are
used. Some flame tubes have ceramic coated tiles fixed to a skin on their
interior walls. Cooling air passes through the holes in the skin and flows
between the skin and the tile which has a ridged surface. The ridged surface
improves the heat transfer between the tile and the air. The air finally enters
the flame tube at the front and rear of the tile and forms an insulating film
for the tile as it flows over it.
Other engine designs use a different method of cooling the air casing which
is called transpiration cooling where a film of air flows between laminations
which form the flame to wall and then exits the laminations to form an
insulating film of air within the flame tube.
This picture illustrates several other features of the type of combustion
chamber which would be used in a multiple combustion chamber system.
Although the engine would probably start quite readily with only one igniter
operating most gas turbine engines have two
igniters.
However, because there are only two igniters another means of passing the
starting frame between the combustion chambers in this type of system has to
be found.
This is called the inter connector.
Immediately after light up the flame in the combustion chambers which have the
igniters causes an increase in the pressure within those chamber. The
pressure differential between the chamber which has a flame in it and
those are joining it which have no flames drives the burning gases through
the interconnected pipe work. When the burning gases come into contact with any
unlit mixture in the adjacent combustion chamber they ignite that mixture.
This process is continued around the engine until the contents of all of the
chambers is burning. Whereupon the pressures within them are equalized and
the flow through the interconnector ceases.
The ceiling ring at the turbine end of the combustion chamber allows for
elongation of the chamber due to expansion.
The chamber is fixed at the compressor end by being bolted on to it and it cannot
expand in that direction.
The sealing ring allows the chamber to expand into the nozzle box,
which is the portion of the engine immediately preceding the nozzle guide
vanes, while maintaining a gas-tight seal between the chamber and the atmosphere.
The straight-through flow multiple combustion chamber system was developed
from Sir Frank whittles original design. Which was supplied with air by a
centrifugal compressor.
The street through combustion chamber system was later used on some earlier
types of axial flow engines and is still in use on centrifugal compressor engines,
such as the Rolls-Royce Dart.
A multiple combustion chamber system consisted of a number of the individual
combustion chambers, which are shown here. Each combustion chamber consists of a
frame tube, which has its own air casing. The combustion chambers are disposed
around the engine just to the rear of the compressor section.
This picture shows a multiple combustion chamber system similar to that used in
the Rolls-Royce Aven which was a powerful for its time axial flow
compressor engine used on many different types of aircraft both military and
commercial fighter and transport for a considerable number of years.
Shown here are the snout, the primary air scoop,
the interconnectors,
and the drain tubes. The drain tubes allow the drainage of
excess fuel from the combustion chambers in the unlikely event of the engine
failing to start. An event which is more commonly called a wet start. A wet start
happens when the mixture of air and fuel in the combustion chamber fails to
ignite during a start.
A considerable amount of fuel will have been fed into the combustion chamber
during the attempt to start. If that field is not removed before the
next attempt to start and if that attempt is successful the result will be
a combination of excessively high gas temperatures in the turbine region
and torching. Which is the name for a very long very hot and very dangerous
jet of flame issuing from the rear of the engine.