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JET ENGINE COMBUSTOR


Combustor is a component or area of a gas turbine, RAMJET, or SCRAM-JET engine where combustion takes place. It is also known as a burner, combustion chamber or flame holder. The  objective of the combustor in a gas turbine is to add energy to the system to power the turbines, and produce a high velocity gas to exhaust through the nozzle in aircraft applications





PURPOSE

The primary function of the combustion section is to burn the fuel/air mixture, thereby adding heat energy to the air (i.e. complete combustion). To maintain Low pressure loss across the combustor. It should capable of operating at wide range. Most combustors must be able to operate with a variety of inlet pressures, temperatures, and mass flows. These factors change with both engine settings and environmental conditions

PROCESS
The process of combustion in a gas turbine involves following four steps
1. Formation of reactive mixture
2. Ignition
3. Flame propagation
4. Cooling of combustion products with air

High temperature and high pressure is necessary for complete combustion of fuel
air mixture. This process will takes place in 3 stages. They are
  1. About 15-20% of air is introduced around the jet of fuel in the primary zone to provide the necessary high temperature for rapid combustion.
  2. Some 30% of total air is then introduced through holes in flame tube in secondary zone to complete the combustion. For high combustion efficiency, the air must be injected carefully at right points in the process to avoid chilling the flame locally and drastically reducing the reaction rate in that neighborhood.
  3. In the tertiary zone the remaining air is mixed with products of combustion to cool them down to temperature required at inlet to turbine. Sufficient turbulent must be promoted so that hot and cold streams are thoroughly mixed to give the desired outlet temperature distribution, with no heat streaks which would damage the turbine blades.
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Constructional Features




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.     Ignition system
5.     Swirler

CASING:
                        The case is the outer shell of the combustor, and is a fairly simple structure. The casing generally requires little maintenance. The case is protected from thermal loads by the air flowing in it, so thermal performance is of limited concern. However, the casing serves as a pressure vessel that must withstand the difference between the high pressures inside the combustor and the lower pressure outside.

LINER

    The liner contains the combustion process and introduces the various airflows (intermediate, dilution, and cooling) into the combustion zone. The liner must be designed and built to withstand extended high temperature cycles. For that reason liners tend to be made from superalloys like Hastelloy X. Furthermore, even though high performance alloys are used, the liners must be cooled with air flow. Some combustors also make use of thermal barrier coatings. However, air cooling is still required. In general, there are two main types of liner cooling; film cooling and transpiration cooling.
          Film cooling works by injecting cool air from outside of the liner to just inside of the liner. This creates a thin film of cool air that protects the liner, reducing the temperature at the liner from around 1800 Kelvin (K) to around
830 K.
          Transpiration cooling, is a more modern approach that uses a porous material for the liner. The porous liner allows a small amount of cooling air to pass through it, providing cooling benefits similar to film cooling.
 
FUEL INJECTION SYSTEM

                             The fuel injector is responsible for introducing fuel to the combustion zone and, along with the Swirler is responsible for mixing the fuel and air. There are four primary types of fuel injectors; pressure-atomizing, air blast, vaporizing, and premix/prevaporizing injectors.

IGNITION SYSTEM

   Most igniters in gas turbine applications are electrical spark igniters, similar to automotive spark plugs. The igniter needs to be in the combustion zone where the fuel and air are already mixed, but it needs to be far enough upstream so that it is not damaged by the combustion itself. Once the combustion is initially started by the igniter, it is self-sustaining and the igniter is no longer used. In can-annular and annular combustors, the flame can propagate from one combustion zone to another, so igniters are not needed at each one.

SWIRLER/DOME

The dome and Swirler are the part of the combustor that the primary air flows through as it enters the combustion zone. Their role is to generate turbulence in the flow to rapidly mix the air with fuel. Early combustors tended to use bluff body domes which used a simple plate to create wake turbulence to mix the fuel and air. Most modern designs, however, are swirl stabilized (use swirlers). The Swirler establishes a local low pressure zone that forces some of the combustion products to recirculate, creating the high turbulence. However, the higher the turbulence, the higher the pressure loss will be for the combustor, so the dome and Swirler must be carefully designed so as not to generate more turbulence than is needed to sufficiently mix the fuel and air.

Types
                   There are three types of combustion chamber. They are
1.CAN
2.CANNULAR
3.ANNULAR 


CAN TYPE:
                   Can combustors are self-contained cylindrical combustion chambers. Each "can" has its own fuel injector, igniter, liner, and casing 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. The outer case is removed to facilitate liner replacement. The primary air from the compressor is guided into each individual can, where it is decelerated, mixed with fuel, and then ignited. The secondary air also comes from the compressor, where it is fed outside of the liner (inside of which is where the combustion is taking place). The secondary air is then fed, usually through slits in the liner, into the combustion zone to cool the liner via thin film cooling. In most applications, multiple cans are arranged around the central axis of the engine, and their shared exhaust is fed to the turbine.



Above image shows the Arrangement of can-type combustors for a gas turbine engine, looking axis on, through the exhaust. The blue indicates cooling flow path, the orange indicates the combustion product flow path. The small orange circles are the fuel injection nozzles.

ADVANTAGES
  • Ease of design and testing
  • Easy to maintain, as only a single can needs to be removed, rather than the whole combustion section
  • Air flow pattern can be controlled easily
 
DISADVANTAGES
  • The combustion is inefficient
  • Combustor is structurally weaker that other forms of combustors
  • More weight
  • Pressure drop is higher than other types
 
CANNULAR TYPE



Above image shows the Arrangement of cannular type combustors for a gas turbine engine, looking axis on, through the exhaust. The blue indicates cooling flow path, the orange indicates the combustion product flow path. The small orange circles are the fuel injection nozzles.
          Like the can type combustor, can annular combustors have discrete combustion zones contained in separate liners with their own fuel injectors. Unlike the can combustor, all the combustion zones share a common ring (annulus) casing. Each combustion zone no longer has to serve as a pressure vessel. The combustion zones can also "communicate" with each other via liner holes or connecting tubes that allow some air to flow circumferentially. The exit flow from the cannular combustor generally has a more uniform temperature profile, which is better for the turbine section. It also eliminates the need for each
chamber to have its own igniter. Once the fire is lit in one or two cans, it can easily spread to and ignite other cans

ADVANTAGES

  • Low weight compared to can type
  • Pressure loss is minimum
  • More uniform at temperature at exit

      DISADVANTAGES

 more difficult to maintain than a can combustor
 it is less efficient than the annular combustor

      ANNULAR TYPE

                             Most commonly used type of combustor is the fully annular combustor. 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. There are many advantages to annular combustors, including more uniform combustion, shorter size (therefore lighter), and less surface area. Additionally, annular combustors tend to have very uniform exit temperatures.



      Above image shows the Arrangement of annular type combustors for a gas turbine engine, looking axis on, through the exhaust. The small orange circles are the fuel injection nozzles

     ADVANTAGES

  • Ø More uniform combustion
  • Ø Simple design
  • Ø Low pressure drop
  • Ø shorter size (therefore lighter), and less surface area

    DISADVANTAGES

  • Ø repairs or replacement will necessitate a whole engine disassembly; thereby, time consuming and expensive
  • Ø structural problems due to thin casing

    FACTORS AFFECTING COMBUSTION CHAMBER PERFORMANCE

                             The main function of gas turbine combustor is to effect the chemical combination of oxygen of air supplied by compressor with carbon and hydrogen components of fuel in such a manner that a steady gas at uniform temperature is produced. Factors affecting combustion chamber performance is

    PRESSURE LOSS

                It is evident that turbulence is necessary for rapid combustion. However this will cause some pressure drop in combustion chamber .This loss is usually regarded as parasitic loss and hence should be minimized. The pressure losses are caused by two factors 1. friction 2. Acceleration accompanying heat addition. The combined pressure loss due to both heating and friction is sum of pressure losses determined  as cold loss and hot loss

    COMBUSTION INTENSITY

                                      The size of combustion chamber is determined by the rate of heat release required. The heat release rate is based on mass flow rate, fuel air ratio and calorific value of fuel. If the volume is enough larger it is easy to achieve low pressure drop, high efficiency good outlet temperature distribution. The design problem is also easied by increase in pressure and temperature of entering chamber. By increasing the temperature , time necessary for preparation of fuel air mixture (evaporation of droplets of fuel) will be reduced. 

     If inlet pressure is high then rate of chemical reaction will increase because at high pressure collision between molecules will be more

    Combustion Intensity = Heat Release Rate / (Combustion Volume * Pressure)

       Lower the combustion intensity the easier it is design a combustion system will meet all requirements. For aircraft the combustion intensity in the region  2-5*10^4 kW/m^3 atm.

  

3. COMBUSTION EFFICIENCY

                                   The combustion efficiency is computed with chemical analysis of gases

    Combustion efficiency = Actual temperature rise / ideal temperature rise

        The actual temperature rise is obtained by direct measurement of inlet and outlet of chamber and this ideal temperature rise is found by already available curves, measured values of air fuel ratio and inlet temperature.














 






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