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as necessary to achieve the desired pressure ratio
and all the airflow from the intake passes through the
compressor.
16. The multi-spool compressor consists of two or
more rotor assemblies, each driven by their own
turbine at an optimum speed to achieve higher
pressure ratios and to give greater operating
flexibility.
17. Although a twin-spool compressor (fig. 3-7) can
be used for a pure jet engine, it is most suitable for
the by-pass type of engine where the front or low
pressure compressor is designed to handle a larger
airflow than the high pressure compressor. Only a
percentage of the air from the low pressure
compressor passes into the high pressure
compressor; the remainder of the air, the by-pass
flow, is ducted around the high pressure compressor.
Both flows mix in the exhaust system before passing
to the propelling nozzle (Part 6). This arrangement
matches the velocity of the jet nearer to the optimum
requirements of the aircraft and results in higher
Compressors
24
Fig. 3-8 Typical triple spool compressor.
propulsive efficiency, hence lower fuel consumption.
For this reason the pure jet engine where all the
airflow passes through the full compression cycle is
now obsolete for all but the highest speed aircraft.
18. With the high by-pass ratio turbo-fan this trend
is taken a stage further. The intake air undergoes
only one stage of compression in the fan before
being split between the core or gas generator system
and the by-pass duct in the ratio of approximately
one to five (fig. 3-8). This results in the optimum
arrangement for passenger and/or transport aircraft
flying at just below the speed of sound. The fan may
be coupled to the front of a number of core
compression stages (two shaft engine) or a separate
shaft driven by its own turbine (three shaft engine).
Principles of operation
19. During operation the rotor is turned at high
speed by the turbine so that air is continuously
induced into the compressor, which is then
accelerated by the rotating blades and swept
rearwards onto the adjacent row of stator vanes. The
pressure rise results from the energy imparted to the
air in the rotor which increases the air velocity. The
air is then decelerated (diffused) in the following
stator passage and the kinetic energy translated into
pressure. Stator vanes also serve to correct the
deflection given to the air by the rotor blades and to
present the air at the correct angle to the next stage
of rotor blades. The last row of stator vanes usually
act as air straighteners to remove swirl from the air
prior to entry into the combustion system at a
reasonably uniform axial velocity. Changes in
pressure and velocity that occur in the airflow
through the compressor are shown diagrammatically
in fig. 3-9. The changes are accompanied by a
progressive increase in air temperature as the
pressure increases.
20. Across each stage the ratio of total pressures of
outgoing air and inlet air is quite small, being
between 1:1 and 1:2. The reason for the small
pressure increase through each stage is that the rate
of diffusion and the deflection angle of the .blades
must be limited if losses due to air breakaway at the
blades and subsequent blade stall are to be avoided.
Although the pressure ratio of each stage is small,
every stage increases the exit pressure of the stage
that precedes it. So whilst this first stage of a
compressor may only increase the pressure by 3 to
4 lb. per sq. in., at the rear of a thirty to one
compression system the stage pressure rise can be
up to 80 lb, per sq. in, The ability to design multistage
axial compressors with controlled air velocities
and straight through flow, minimizes losses and
results in a high efficiency and hence low fuel
consumption. This gives it a further advantage over
the centrifugal compressor where these conditions
are fundamentally not so easily achieved.
21. The more the pressure ratio of a compressor is
increased the more difficult it becomes to ensure that
it will operate efficiently over the full speed range.
This is because the requirement for the ratio of inlet
area to exit area, at the high speed case, results in
an inlet area that becomes progressively too large
relative to the exit area as the compressor speed and
hence pressure ratio is reduced. The axial velocity of
the inlet air in the front stages thus becomes low
relative to the blade speed, this changes the
incidence of the air onto the blades and a condition
is reached where the flow separates and the
　
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