The compressor section of the turbo-jet engine has many functions. Its primary purpose is to supply air in sufficient quantity to satisfy the requirements of the combustion process. Specifically to fulfil its purpose, the compressor must increase the pressure of the mass of air received from the air inlet duct and then to discharge it to the burners in the combustion chambers, in the mass and pressures required.
Compressors may be identified by the direction of the airflow through them. The two basic types have a centrifugal flow or an axial flow. Some engines may use both types on one compressor assembly.
CENTRIFUGAL COMPRESSORS
These compressors receive the air at their centre and accelerate it outwards by centrifugal force. The air is then expelled into a divergent duct, called a diffuser, where velocity is exchanged for energy.
A complete centrifugal compressor assembly consists of an impeller rotor, a diffuser and a manifold. The impeller can be single or double sided and can be installed in either one or two-stage assemblies. Whilst this type of compressor can generate a high mass-flow from a small diameter engine, it cannot take advantage of ram effect due to the tortuous route that the airflow has to follow through the
compressor.
Whilst compression ratios in the vicinity of 5:1 were the norm on earlier designs, it is now possible to produce centrifugal compressors with compression ratios of 15:1, which are quite competitive with axial compressors. It is very rare to find more than two stages of compression, due to the huge losses caused by the continued re-direction of the airflow through the stages, the added weight of the impellers and the power required from the turbine to drive the compression stages.
The illustrations below show a single stage, dual-sided compressor and a two-stage, single sided centrifugal compressor.
Single Stage Dual Sided Compressor Two Single Sided Compressors Figure 4.1
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AXIAL FLOW COMPRESSORS
The axial flow compressor has two main components, a rotor and stator. The rotor has blades attached to a spindle or drum, which impels the air rearwards in the same manner as a propeller. The stator blades act as diffusers at each stage, partially converting high velocity to pressure.
A set of rotor and stator blades constitutes a pressure stage, each stage being capable
of producing a pressure rise of about 1.25:1. The number of these stages is dictated by
the amount of air and the pressure rise that is required. A normal maximum number of stages to be found is between 16 and 18.From the front to the rear of an axial compressor, the space between the rotor shaft and the stator casing becomes smaller. This is necessary to maintain a near constant axial velocity of the air as the density increases with compression. This is shown in the illustration below.
Axial compressor Figure 4.2
ADVANTAGES AND DISADVANTAGES ADVANTAGES
CENTRIFUGAL COMPRESSOR AXIAL COMPRESSOR
High pressure rise per stage High ram effect efficiency Good efficiency over wide speed range High peak pressures Simplicity of manufacture, low cost Small frontal area Low weight
Low power for starting Damage tolerant
DISADVANTAGES
CENTRIFUGAL COMPRESSOR AXIAL COMPRESSOR
Large frontal area Complex manufacture
Limited to two stages Relative high weight High starting power Low pressure rise/stage
Mod15 Gas Turbine Engines by COBC 47
COMBINED COMPRESSORS
There are a number of gas turbines that use both types of compressor by having an axial compressor, followed immediately by a centrifugal compressor. The
aerodynamic advantages of this arrangement are too complex to discuss at this stage but, suffice to say that this layout can generally be found on turbo-shaft
engines, which power helicopters. The example, illustrated below, is of the Lycoming T-55 engine that powers the Chinook helicopter. It can be seen that it has seven axial stages followed by a single centrifugal stage.
Combined Axial/Centrifugal Compressors Figure 4.3
OPERATING PRINCIPLES CENTRIFUGAL
The impeller is rotated at high speed by the turbine and the air is continuously induced into the centre of the impeller. Centrifugal action causes it to flow outwards along the vanes to the tip, which causes it to accelerate and the pressure to rise.
Once it leaves the impeller it passes through the diffuser section, which is divergent, causing the pressure to rise again. This demonstrates how this arrangement has half the compression occurring in the impeller and half in the diffuser.
This type of compressor works best at high rotational speeds. It is normal for a centrifugal compressor to have impeller tip speeds of around 1,600 ft per second, (well over Mach1). This is one of the reasons why centrifugal compressors generate a high level of noise when operating.
AXIAL
The rotor is rotated at high speed by the turbine, continuously drawing air into the front of the compressor. After each rotor stage, which has caused the pressure to rise, the air passes through a stator stage, which diffuses, (decelerates), the air and causes the pressure to rise yet again. This process continues throughout the number of stages of the compressor, each stage comprising a rotor and a stator, each stage achieving a compression ratio of approximately 1.25:1
The stators have a second duty, which is to straighten out the ‘swirl’ which is the result of axial compression. As the air leaves each rotor stage with increased
velocity, it also has a rotary motion that, if not corrected, will reduce the efficiency of each progressive stage. The stator turns the air in the reverse direction, resulting in the airflow flowing axially through each stage.
When axial compressors are required to produce a high level of compression, it becomes very difficult to control the air throughout all of the stages. This is due to the variables that any aircraft can meet. These include the speed of the compressor, due to throttle demand from the flight deck; the speed of the aircraft, especially in the climb or descent and the density of the air or altitude at which the aircraft is operating.
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CONSTRUCTIONAL FEATURES CENTRIFUGAL
The centrifugal compressors are usually mounted on ball or roller bearings and are driven by the turbine stage(s). The connecting shaft may be manufactured in two parts, to allow engine disassembly, whilst having a self-aligning coupling to join the parts together.
The discs are forged with the vanes straight for ease of manufacture. Normally a separate set of rotating guide vanes, which cannot be easily forged, are attached to the front of the impeller. These draw the air into the impeller unit.
Diffuser assemblies are often part of the compressor case, with integrally cast vanes to act as both diverging ducts and to direct the airflow into the elbows and the
combustion chambers.
AXIAL
The Axial compressor consists of firstly, the rotating rotor, made up from the main shaft supported by ball and roller bearings and either separate discs or a drum assembly, to which are affixed the blades of differing sizes. Secondly, the casing assembly, in a number of pieces (to allow splitting, for access to the rotors), contains all of the stator vanes attached to the inside face of the case. The case also provides part of the strength of the complete engine and, on some designs, has attachments or mounting points built into the case design.
The vanes are affixed to the rotor discs and stator case(s) by a variety of methods, all giving positive retention against centrifugal force, (rotors) and rotation, (stators).
The rotor blades are of aerofoil section and are twisted, much the same as a propeller, to give an even thrust along their length. This is shown by the different stagger angles between the root and tip of the blades. The roots of the blades are formed into a shape that matches the recesses in the rotor disc and they only have to be retained on the disc by plates that restrict fore and aft movement. This can be seen in the illustration overleaf.
Blade Details Figure 4.4 1. STATOR VANES
Stator vanes are also of aerofoil section and are located in slots around the
compressor casing. There is no chance for the blades to move fore and aft due to the retention of the grooves, but there is a tendency for the blades to slide radially around the grooves. This tendency is caused by the air loads, generated by the blades straightening the airflow after each rotor stage.
This movement is prevented by retaining set screws, which hold a number of blades in place, preventing any movement by the others. This is shown in Figure 4.5, where a retaining ring, held by the screw, holds the blades in place.
Stator Blade Retention Figure 4.5
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FAN BALANCING
The fan consists of the single front stage of the compressor. Normally, it is the low- pressure (LP) compressor and is part of a twin or triple-spool engine. It will usually consist of a small number of blades that can be removed, often individually, if they become damaged in service. The engine shown below, a Rolls Royce Tay, has a wide chord fan which can be both repaired, (by blade replacement) and balanced in situ.
Fan blades may be manufactured from Titanium, sometimes as a skin with a
honeycomb core, although some have been manufactured from composite materials.
Titanium is used normally because of the bird strike requirements that dictate very strong blades on the first stage of the engine.
Rolls Royce Tay Fan Figure 4.6
The need for the engine to be precisely balanced because of its high rotational speed, means that the replacement of individual blades must be undertaken with care. In most cases, the blades will have been pre-weighed by the manufacturer and the value engraved upon the blade. The blades will be divided into weight “groups”
so that, providing the replacement blade is of the same “group” as the one removed, there should be no need to balance the assembly.
In some cases, due to the engine having built-in vibration sensors, it will be possible to carry out balancing ground-runs. This will allow the engineer, following the
maintenance manual, to add or remove small balance weights, at specific points around the fan assembly, until the assembly is in perfect balance.
Shown below is the fan assembly of the Tay engine, showing the use of weights to give a balanced assembly.
Tay Fan Assembly Figure 4.7
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STALL AND SURGE
Each stage of a multi-stage compressor possesses certain airflow characteristics that are dissimilar from those of its neighbour thus; to design a workable and efficient compressor, each stage must be matched to the next stage. This matching is fairly simple when the engine is running on a test bed, it is much more difficult when
speed, altitude, temperature, etc. are included, such as when the aircraft is operating normally.
In extreme conditions, the airflow through the compressor can become disturbed and vibration can be set-up. This stalling of the blades can either be positive or
negative, depending whether the fault is at the intake, (front), or at the high compression, (rear), end of the engine.
If the engine demands a pressure rise from the compressor greater than the blades can sustain, surge will occur. This is an instantaneous breakdown of flow through the engine and high-pressure air in the combustion system is expelled forwards through the compressor,T.G.T. would rise and may be accompanied with a loud bang, resulting in a loss of thrust. To overcome this problem, engines have a declared ‘safety margin’ to ensure the area of instability is avoided. This is shown graphically below.
Surge Margin Diagram Figure 4.8
To control these disturbances, which occur most often on single shaft engines with high compression ratios, a variety of methods are used on different engines. This control can take the form of variable inlet guide vanes for the first stage and variable stator vanes for other stages. As the compressor slows from its optimum, the blades change their angle of attack to vary the airflow on to the rotor blades, so that they do not stall and remain at their optimum angle of attack.
BLEED BANDS/VALVES
In addition, an interstage bleed may be fitted to the compressor casing, usually located at the higher compression stages, permitting excessive pressure to be bled overboard. This avoids the choking which may occur during rapid acceleration. Due to the loss of performance during normal operations, bleed valves will usually only be opened during starting and acceleration. The operation of these air bleed systems can either be actuated by hydraulic, pneumatic or electronic methods
VARIABLE INLET GUIDE VANES (VIGVS)
The number of stages that have variable incidence vanes depends on the design of the engine. Some may only have the first stage inlet guide vanes moveable, whilst others can have four or more stages that are variable. The illustration below shows an engine with variable inlet guide vanes and three variable stator vanes.
Guide Vanes and Stator Vanes Figure 4.9
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OVERFUELLING SURGE
All engines have to be over fuelled by a small margin to cause them to accelerate. If the over fuelling is above the correct figure, due perhaps to a badly adjusted fuel control unit, then the inertia of the rotating parts of the engine will resist acceleration.
The excessive fuel will cause choking at the turbine; this will cause a slowing of the compressor air velocity, resulting in a progressive stall through the engine from the front. The resulting reversal of the airflow is a surge.
TWIN SPOOL AXIAL FLOW COMPRESSORS
Relief from surging troubles can be obtained from the devices described earlier. A better solution is the twin-spool axial flow compressor, part of the twin spool engine type described earlier. The compressor has two sections, each section is completely independent from the other and driven by its own turbine assembly, each mounted on its own co-axial shaft. The LP compressor is driven by the aft, LP turbine and the HP compressor is driven by the forward, HP turbine. Each shaft assembly will be rotating at its optimum speed.
Whether the engine is at high or low altitude or whether it is moving through the air at high or low speeds, the two spools will be matched to the external atmosphere
parameters and aircraft performance.
At idle, for example, the HP system is doing most of the work whilst the LP spool runs slower, this makes its angle of attack of the airflow on to the first stage much better and, due to the faster moving HP spool, there is less chance of ‘choking’.
Equally, at higher altitudes, when the LP spool rotates faster, due to the reduced air density, the greater mass airflow to the HP section restores some of the losses that a single spool engine would suffer at this altitude.
INTENTIALLY LEFT BLANK
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