Combustor Types and Design Considerations

The evolution of gas turbine combustor technology has resulted in a number of popular design configurations developing. There are five basic concepts that are integral to any combustor design, these being, a primary reaction zone, a secondary

Compressor blades

Combustion chambers

Turbine blades

Exhaust nozzle

Compressor blades

Exhaust nozzle Turbine blades

Fan blades

Combustion chambers

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reaction zone, a dilution zone, various wall jets, and management of heat transfer at the combustor boundary, as illustrated in Figure 2.8. In traditional gas turbine combustors fuel and air are injected into the combustion chamber individually. This type of combustor design is known as a diffusion combustor. Although the fuel and air are not pre-mixed prior to entering the primary zone, the combustion reaction does not occur at the interface between the fuel and air as with a tradition diffusion flame.

Instead the injection of the reactants, the mixing of the reactants, the entrainment and mixing of the energetic species, and combustion reaction are occurring simultaneously throughout the volume of primary zone [21].

Alternatively, combustor designs can be of a pre-mixed variety in which the reactants have undergone some level of mixing prior to entering the reaction zone. Only systems where the fuel and air have undergone pre-mixing before injection into the primary zone can be classified as truly pre-mixed. Other pre-mixed system variants include;

rapidly mixed, non-premixed - where reactants are injected separately but undergo intense mixing preceding the onset of combustion, and spatially injected, non-premixed - where fuel and air are injected at multiple discrete points within the combustion chamber and spread out over a larger area promoting a more disperse mixture region. It is these latter variations which have seen most use in pre-mixed aviation gas turbines designs.

Figure 2.8: Basic combustor features, reproduced from [21].

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In a gas turbine combustor the role of the primary zone is to establish a stable combustion region by controlling air-fuel mixing and flame structure. Prior to entering the primary zone the air exiting the compressor is passed through over a series of vanes to induce a turbulent swirling flow. This not only aids the initial mixing process immediately following fuel injection but also creates recirculation zones near to the fuel and air inlet ducts. Recirculation zones help maintain continuous combustion by drawing hot combustion products back towards the reactant injection point.

Additional air can also be introduced - labelled as ‘primary jets’ in Figure 2.8 - to help promote mixing, influence stoichiometry and improve the flame structure within the combustor. These air jets are also designed to direct the flow of the reacting mixture up and downstream of the primary zone in the correct proportions. This ensures that the hot mixtures are entrained in the recirculation zones and that the mixing and combustion process continues into the secondary zone.

The main purpose of the secondary zone is to oxidise CO to carbon dioxide (CO2) via the kinetic reaction shown in Equation 1.

𝐶𝑂 + 𝑂𝐻 → 𝐶𝑂₂+ 𝐻

Equation 1: Secondary zone mechanism for the oxidation of CO to CO2.

In order to maximise the oxidation rate an elevated temperature is maintained, whilst the residence time within the secondary zones is designed to be as long as possible and surplus air is provided into the reaction region to create an overall lean mixture.

Lastly the combustion products enter the dilution zone where the main goal is to cool the mixture to a temperature which will not damage the turbine blades of the engine.

Cooling is achieved using dilution air which is injected into the region, leading to a reduction in the mean exhaust temperature and making the final exhaust extremely lean.

There are three main combustor configurations which are generally used in aviation gas turbine engines. These options are illustrated in Figure 2.9, and are known as tubular or ‘can’, tubo-annular, and annular. The choice of which combustor type and

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layout is implemented in a gas turbine design is largely determined by the intended purpose of the engine, and the availability of space in the engine housing.

A tubular combustor design consists of a cylindrical liner mounted concentrically inside a cylindrical casing. The design is rarely used in modern aircraft engine designs due the excessive weight and length required, however for small scale uses it has the benefit of being easier to maintain when compared to the other combustor choices.

In a tubo-annular combustor design tubular liners are arranged in a single annular casing combining the compactness of an annular chamber with the mechanical strength of the tubular chamber. This system became popular as the pressure ratios of modern aviation gas turbines increased.

Figure 2.9: Illustration of the main combustor designs adopted in aviation gas turbines, adapted from [22].

Eventually combustor design has developed to a point where fully annular designs have become the standard for civil and commercial aviation transport. The clean aero-dynamic layout allows for a compact combustion chamber with low pressure losses.

Combustion systems such as those shown in Figure 2.10 employ advanced fuel injection and wall cooling techniques to further improve fuel and combustion efficiencies thereby reducing operation instability and pollutant emissions.

Tubular Tubo-annular Annular

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Figure 2.10: Schematic drawings of (a) General Electric CF6-50 annular combustor and (b) Rolls Royce RB211 annular combustor, reproduced from [22].

The selection of a particular combustor technology or design configuration used in any aviation gas turbine is of course not based solely on individual advantages or disadvantages one option may have over another. There is however an accepted general ranking of the areas which are considered most important in any design. Table 2.1, which is adapted from a study by Rhode in 2002 [23], shows a prioritised list of combustor design considerations.

Table 2.1: Prioritised combustor design considerations for aviation gas turbine engines, adapted from [23].

Design Consideration Criteria

Safety

Operability Take-off to 45,000 ft

Efficiency 99.9%

Durability 6,000 cycles/36,000 hours

Emissions Best available emissions at 7%, 30%, 85%

and 100% power (LTO Cycle) and cruise Downstream

- Turbomachinery - Thermal and Life - Integrity

Minimise the impact of the combustion output

(a) (b)

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From this ranking structure as would be expected safety and operational robustness are the main consideration in a gas turbine design. The rank given to achieving low exhaust pollutant emissions in aviation applications although not a primary factor in combustor design, due to the necessary importance which is placed upon human safety and accident mitigation, is still something which is crucial to the competiveness and success of an engine design for OEMs as regulatory standards and public awareness of environmental sustainability increase.