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.