Turbine Engine Combustion Section

The combustion section houses the combustion process, which raises the temperature of the air passing through the engine. This process releases energy contained in the air/ fuel mixture. The major part of this energy is required at the turbine or turbine stages to drive the compressor. About 2/3 of the energy is used to drive the gas generator compressor. The remaining energy passes through the remaining turbine stages that absorb more of the energy to drive the fan, output shaft, or propeller. Only the pure turbojet allows the air to create all the thrust or propulsion by exiting the rear of the engine in the form of a high-velocity jet. These other engine types have some jet velocity out the rear of the engine but most of the thrust or power is generated by the additional turbine stages driving a large fan, propeller, or helicopter rotor blades.

The primary function of the combustion section is, of course, to burn the fuel/air mixture, thereby adding heat energy to the air. To do this efficiently, the combustion chamber must:
  • Provide the means for proper mixing of the fuel and air to assure good combustion,
  • Burn this mixture efficiently,
  • Cool the hot combustion products to a temperature that the turbine inlet guide vanes/blades can withstand under operating conditions, and
  • Deliver the hot gases to the turbine section.

The location of the combustion section is directly between the compressor and the turbine sections. The combustion chambers are always arranged coaxially with the compressor and turbine regardless of type, since the chambers must be in a through-flow position to function efficiently. All combustion chambers contain the same basic elements:
  1. Casing
  2. Perforated inner liner
  3. Fuel injection system
  4. Some means for initial ignition
  5. Fuel drainage system to drain off unburned fuel after engine shutdown

There are currently three basic types of combustion chambers, variations within type being in detail only. These types are:
  1. Can type
  2. Can-annular type
  3. Annular type

The can-type combustion chamber is typical of the type used on turboshaft and APUs. [Figure 1] Each of the can-type combustion chambers consists of an outer case or housing, within which there is a perforated stainless steel (highly heat resistant) combustion chamber liner or inner liner. [Figure 2] The outer case is removed to facilitate liner replacement.

Aircraft Gas Turbine Engine Combustion Section
Figure 1. Can-type combustion chamber

Aircraft Gas Turbine Engine Combustion Section
Figure 2. Inside view of a combustion chamber liner

Older engines with several combustion cans had each can with interconnector (flame propagation) tube, which was a necessary part of the can-type combustion chambers. Since each can is a separate burner operating independently of the other cans, there must be some way to spread combustion during the initial starting operation. This is accomplished by interconnecting all the chambers. As the flame is started by the spark igniter plugs in two of the lower chambers, it passes through the tubes and ignites the combustible mixture in the adjacent chamber, and continues until all the chambers are burning.

The flame tubes vary in construction details from one engine to another, although the basic components are almost identical. [Figure 3] The spark igniters previously mentioned are normally two in number, and are located in two of the can-type combustion chambers.

Aircraft Gas Turbine Engine Combustion Section
Figure 3. Interconnecting flame tubes for can-type combustion chambers

Another very important requirement in the construction of combustion chambers is providing the means for draining unburned fuel. This drainage prevents gum deposits in the fuel manifold, nozzles, and combustion chambers. These deposits are caused by the residue left when the fuel evaporates. Probably most important is the danger of afterfire if the fuel is allowed to accumulate after shutdown. If the fuel is not drained, a great possibility exists that, at the next starting attempt, the excess fuel in the combustion chamber will ignite and exhaust gas temperature will exceed safe operating limits.

The liners of the can-type combustors have perforations of various sizes and shapes, each hole having a specific purpose and effect on flame propagation within the liner. [Figure 1] The air entering the combustion chamber is divided by the proper holes, louvers, and slots into two main streams—primary and secondary air. The primary or combustion air is directed inside the liner at the front end, where it mixes with the fuel and is burned. Secondary or cooling air passes between the outer casing and the liner and joins the combustion gases through larger holes toward the rear of the liner, cooling the combustion gases from about 3,500 °F to near 1,500 °F. To aid in atomization of the fuel, holes are provided around the fuel nozzle in the dome or inlet end of the can-type combustor liner. Louvers are also provided along the axial length of the liners to direct a cooling layer of air along the inside wall of the liner. This layer of air also tends to control the flame pattern by keeping it centered in the liner, thereby preventing burning of the liner walls. Figure 4 illustrates the flow of air through the louvers in the annular combustion chamber.

Aircraft Gas Turbine Engine Combustion Section
Figure 4. Annular combustion chamber liner

Some provision is always made in the combustion chamber case for installation of a fuel nozzle. The fuel nozzle delivers the fuel into the liner in a finely atomized spray. The more the spray is atomized, the more rapid and efficient the burning process is.

Two types of fuel nozzle currently being used in the various types of combustion chambers are the simplex nozzle and the duplex nozzle. The construction features of these nozzles are covered in greater detail in Engine Fuel and Fuel Metering Systems section.

The spark igniter plugs of the annular combustion chamber are the same basic type used in the can-type combustion chambers, although construction details may vary. There are usually two igniters mounted on the boss provided on each of the chamber housings. The igniters must be long enough to protrude from the housing into the combustion chamber.

The burners are interconnected by projecting flame tubes which facilitate the engine-starting process as mentioned previously in the can-type combustion chamber familiarization. The flame tubes function identically to those previously discussed, differing only in construction details.

The can-annular combustion chamber is not used in modern engines. The forward face of each chamber presents six apertures, which align with the six fuel nozzles of the corresponding fuel nozzle cluster. [Figure 5] These nozzles are the dual-orifice (duplex) type requiring the use of a flow-divider (pressurizing valve), as mentioned in the can-type combustion chamber discussion. Around each nozzle are preswirl vanes for imparting a swirling motion to the fuel spray, which results in better atomization of the fuel, better burning, and efficiency. The swirl vanes function to provide two effects imperative to proper flame propagation:
  1. High flame speed—better mixing of air and fuel, ensuring spontaneous burning.
  2. Low air velocity axially—swirling eliminates overly rapid flame movement axially.

Aircraft Gas Turbine Engine Combustion Section
Figure 5. Can-annular combustion chamber components and arrangement

The swirl vanes greatly aid flame propagation, since a high degree of turbulence in the early combustion and cooling stages is desirable. The vigorous mechanical mixing of the fuel vapor with the primary air is necessary, since mixing by diffusion alone is too slow. This same mechanical mixing is also established by other means, such as placing coarse screens in the diffuser outlet, as is the case in most axial-flow engines.

The can-annular combustion chambers also must have the required fuel drain valves located in two or more of the bottom chambers, assuring proper drainage and elimination of residual fuel burning at the next start.

The flow of air through the holes and louvers of the can-annular chambers, is almost identical with the flow through other types of burners. [Figure 5] Special baffling is used to swirl the combustion airflow and to give it turbulence. Figure 6 shows the flow of combustion air, metal cooling air, and the diluent or gas cooling air. The air flow direction is indicated by the arrows.

Aircraft Gas Turbine Engine Combustion Section
Figure 6. Airflow through a can-annular combustion chamber

The basic components of an annular combustion chamber are a housing and a liner, as in the can type. The liner consists of an undivided circular shroud extending all the way around the outside of the turbine shaft housing. The chamber may be constructed of heat-resistant materials, which are sometimes coated with thermal barrier materials, such as ceramic materials. The annular combustion chamber is illustrated in Figure 7. Modern turbine engines usually have an annular combustion chamber. As can be seen in Figure 8, the annular combustion chamber also uses louvers and holes to prevent the flame from contacting the side of the combustion chamber.

Aircraft Gas Turbine Engine Combustion Section
Figure 7. Annular combustion with chamber ceramic coating

Aircraft Gas Turbine Engine Combustion Section
Figure 8. Combustion chamber louvers and holes