Contents (This book);
1 Basic mechanics 1
2 Working cycle and airflow 11
3 Compressors 19
4 Combustion chambers 35
5 Turbines 45
6 Exhaust system 59
7 Accessory drives 65
8 Lubrication 73
9 Internal air system 85
10 Fuel system 95
11 Starting and ignition 121
12 Controls and instrumentation 133
13 Ice protection 147
14 Fire protection 153
15 Thrust reversal 159
16 Afterburning 169
17 Water injection 181
18 Vertical/short take-off and landing 187
19 Noise suppression 199
20 Thrust distribution 207
21 Performance 215
22 Manufacture 229
23 Power plant installation 243
24 Maintenance 251
25 Overhaul 263
Appendix 1; Conversion factors 277

Combustion Chamber ( Ruang pembakaran )

The combustion chamber (fig. 4-1) has the difficult task of burning large quantities of fuel, supplied through the fuel spray nozzles, with extensive volumes of air, supplied by the compressor and releasing the heat in such a manner that the air is expanded and accelerated to give a smooth stream of uniformly heated gas at all conditions required by the turbine . This task must be accomplished with the minimum loss in pressure and with the maximum heat release for the limited space available.
The amount of fuel added to the air will depend upon the temperature rise required. However, the maximum temperature is limited to within the range of 850 to 1700 deg. C. by the materials from which the turbine blades and nozzles are made. The air has already been heated to between 200 and 550 deg. C. by the work done during compression, giving a temperature rise requirement of 650 to 1150 deg. C. from the combustion process. Since the gas temperature required at the turbine varies with engine thrust, and in the case of the turbo-propeller engine upon the power required, the combustion chamber must also be capable of maintaining stable and efficient combustion over a wide range of engine operating conditions.
Efficient combustion has become increasingly important because of the rapid rise in commercial aircraft traffic and the consequent increase in atmospheric pollution, which is seen by the general public as exhaust smoke.

Fig. 4-1 An early combustion chamber.

Air from the engine compressor enters the combustion chamber at a velocity up to 500 feet per second, but because at this velocity the air speed is far too high for combustion, the first thing that the chamber must do is to diffuse it, i.e. decelerate it and raise its static pressure. Since the speed of burning kerosine at normal mixture ratios is only a few feet per second, any fuel lit even in the diffused air stream, which now has a velocity of about 80 feet per second, would be blown away. A region of low axial velocity has therefore to be created in the chamber, so that the flame will remain alight throughout the range of engine operating conditions.

In normal operation, the overall air/fuel ratio of a combustion chamber can vary between 45:1 and 130:1, However, kerosine will only burn efficiently at, or close to, a ratio of 15:1, so the fuel must be burned with only part of the air entering the chamber, in what is called a primary combustion zone. This is achieved by means of a flame tube (combustion liner) that has various devices for metering the airflow distribution along the chamber.

Approximately 20 per cent of the air mass flow is taken in by the snout or entry section (fig. 4-2). Immediately downstream of the snout are swirl vanes and a perforated flare, through which air passes into the primary combustion zone. The swirling air induces a flow upstream of the centre of the flame tube and promotes the desired recirculation. The air not picked up by the snout flows into the annular space between the flame tube and the air casing.

Through the wall of the flame tube body, adjacent to the combustion zone, are a selected number of secondary holes through which a further 20 per cent of the main flow of air passes into the primary zone. The air from the swirl vanes and that from the secondary air holes interacts and creates a region of low velocity recirculation. This takes the form of a toroidal vortex, similar to a smoke ring, which has the effect of stabilizing and anchoring the flame (fig, 4-3). The recirculating gases hasten the burning of freshly injected fuel droplets by rapidly bringing them to ignition temperature.

Fig. 4-2 Apportioning the airflow,

Fig. 4-3 Flame Stabilizing &General Airflow pattern

It is arranged that the conical fuel spray from the nozzle intersects the recirculation vortex at its centre. This action, together with the general turbulence in the primary zone, greatly assists in breaking up the fuel and mixing it with the incoming air.

The temperature of the gases released by combustion is about 1,800 to 2,000 deg. C., which is far too hot for entry to the nozzle guide vanes of the turbine. The air not used for combustion, which amounts to about 60 per cent of the total airflow, is therefore introduced progressively into the flame tube. Approximately a third of this is used to lower the gas temperature in the dilution zone before it enters the turbine and the remainder is used for cooling the walls of the flame tube. This is achieved by a film of cooling air flowing along the inside surface of the flame tube wall, insulating it from the hot combustion gases (fig. 4-4). A recent development allows cooling air to enter a network of passages within the flame tube wall before exiting to form an insulating film of air, this can reduce the required wall cooling airflow by up to 50 per cent. Combustion should be completed before the dilution air enters the flame tube, otherwise the incoming air will cool the flame and incomplete combustion will result.
An electric spark from an igniter plug initiates combustion and the flame is then selfsustained.

Fig. 4-4 Flame tube cooling methods.

The design of a combustion chamber and the method of adding the fuel may vary considerably, but the airflow distribution used to effect and maintain combustion is always very similar to that described.

There are three main types of combustion chamber in use for gas turbine engines. These are the multiple chamber, the tubo-annular chamber and the annular chamber.

Fig. 4-7 Multiple combustion chambers.

Fig. 4-8 Tubo-annular combustion chamber

Fig. 4-9 Annular combustion chamber

SOURCE: The jet engine , Rolls-Royce plc.


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