TURBINE COMBUSTION SECTION

Today, three basic combustion chambers are in use. They are the annular combustion chamber, the can type, and the combination of the two called the can-annular. Variations of these basic systems are used in a number of engines. The three systems are discussed individually in the following subparagraphs. The most commonly used gas turbine engine in Army aircraft is the annular reverse-Row type. The combustion section contains the combustion chambers, igniter plugs, and fuel nozzles or vaporizing tubes. It is designed to burn a fuel-air mixture and deliver the combusted gases to the turbine at a temperature which will not exceed the allowable limit at the turbine inlet.
Fuel is introduced at the front end of the burner in a highly atomized spray from the fuel nozzles. Combustion air flows in around the fuel nozzle and mixes with the fuel to form a correct fuel-air mixture. This is called primary air and represents approximately 25 percent of total air taken into the engine. The fuel-air mixture which is to be burned is a ratio of 15 parts of air to 1 part of fuel by weight. The remaining 75 percent of the air is used to form an air blanket around the burning gases and to lower the temperature. This temperature may reach as high as 3500° F. By using 75 percent of the air for cooling, the temperature operating range can be brought down to about half, so the turbine section will not be destroyed by excessive heat. The air used for burning is called primary air- and that for cooling is secondary air. The secondary air is controlled and directed by holes and louvers in the combustion chamber liner.
Igniter plugs function only during starting, being cut out of the circuit as soon as combustion is self-supporting. On engine shutdown, or, if the engine fails to start, the combustion chamber drain valve, a pressure-actuated valve, automatically drains any remaining unburned fuel from the combustion chamber. All combustion chambers contain the same basic elements: a casing or outer shell, a perforated inner liner or flame tube, fuel nozzles, and some means of initial ignition. The combustion chamber must be of light construction and is designed to burn fuel completely in a high velocity airstream. The combustion chamber liner is an extremely critical engine part because of the high temperatures of the flame. The liner is usually constructed of welded high-nickel steel. The most severe operating periods in combustion chambers are encountered in the engine idling and maximum rpm ranges. Sustained operation under these conditions must be avoided to prevent combustion chamber liner failure.

• The annular-type combustion chamber shown in figure 1.18 is used in engines of the axial-centrifugal-flow compressor design. The annular combustion chamber permits building an engine of a small and compact design. Instead of individual combustion chambers, the primary compressed air is introduced into an annular space formed by a chamber liner around the turbine assembly. A space is left between the outer liner wall and the combustion chamber housing to permit the flow of secondary cooling air from the compressor. Primary air is mixed with the fuel for combustion. Secondary (cooling) air reduces the temperature of the hot gases entering the turbine to the proper level by forming a blanket of cool air around these hot gases.

1. ANNULAR TYPE COMBUSTION CHAMBER LINER
2. COMBUSTION CHAMBER HOUSING ASSEMBLY
Figure 1.18. Annular-type Combustion Chamber.

The annular combustion chamber offers the advantages of a larger combustion volume per unit of exposed area and material weight, a smaller exposed area resulting in lower pressure losses through the unit, and less weight and complete pressure equalization.

• The can-type combustion chamber is one made up of individual combustion chambers. This type of combustion chamber is so arranged that air from the compressor enters each individual chamber through the adapter. Each individual chamber is composed of two cylindrical tubes, the combustion chamber liner and the outer combustion chamber, shown in figure 1.19. Combustion takes place within the liner. Airflow into the combustion area is controlled by small louvers located in the inner dome, and by round holes and elongated louvers along the length of the liner. Airflow into the combustion area is controlled by small louvers located in the inner dome, and by round holes elongated louvers along the length of the liner.

Figure 1.19. Can-type Combustion Chamber (Cutaway).

Through these openings flows the air that is used in combustion and cooling. This air also prevents carbon deposits from forming on the inside of the liner. This is important, because carbon deposits can block critical air passages and disrupt airflow along the liner walls causing high metal temperatures and short burner life.
Ignition is accomplished during the starting cycle. The igniter plug is located in the combustion liner adjacent to the start fuel nozzle. The Army can-type engine employs a single can-type combustor.

• Can-annular combustion chamber. This combustion chamber uses characteristics of both annular and can-type combustion chambers. The can-annular combustion chamber consists of an outer shell, with a number of individual cylindrical liners mounted about the engine axis as shown in figure 1.20. The combustion chambers are completely surrounded by the airflow that enters the liners through various holes and louvers. This air is mixed with fuel which has been sprayed under pressure from the fuel nozzles. The fuel-air mixture is ignited by igniter plugs, and the flame is then carried through the crossover tubes to the remaining liners. The inner casing assembly is both a support and a heat shield; also, oil lines run through it.

Figure 1.20. Can-Annular Combustion Chamber.