COMBUSTION CHAMBER AERODYNAMICS

13 Appendix D: Combustion chamber performance characteristics

13.3 COMBUSTION CHAMBER AERODYNAMICS

As shown in Figure 13.3, the air will be introduced into the liner in stages. The first stage called the primary stage, uses 15 to 20% of the incoming air to create the stable burning zone. The o ther zone, shown in Figure 13.3, is known as the dilution zone, where the leftover air is

introduced into the liner. This air is used to cool down the cornbusted gas to temperatures that are acceptable for the turbine. The mixing of hot and cold streams is essential for reducing the occurrence of hot streaks that can damage the turbine.

Since it is possible that some of the fuel has not reacted during the short time in the primary zone, in practice, it is necessary to introduce a secondary zone where more fresh air enters the liner

Secondarv 3

r

Cooling slots

Dilution zone

Figure 13.1: Illustration of the zonal liner air introduction method.

Flame temperature for stoichiometric mixtures is in the region of 1600 "C, which is higher than the material limits of the liner. The liner therefore has to be cooled down to ensure that no wall failures occur and is done by cooling slots in the liner wall. Cooling slots introduce a cooling barrier layer between the hot gas and the liner wall. This cooling air should be directed into the same general direction as the gas stream, not to penetrate the gas stream like the dilution holes.

Distinction between these two introduction methods can be seen in the illustration in Figure 13.2.

Secondary inlets

____, 1

7

Figure 13.2: Illustration of liner inlet flow direction.

The aerodynamic process plays a vital role in the design process of gas turbine combustion chambers and demands knowledge of flow recirculation, jet penetration and mixing, and discharge coefficients for all the air inlets in the liner. Some variables need to be defined before any of the combustor parameters are discussed: The reference velocity

[u,,],

which is the mean velocity across the plane of maximum cross-sectional area of the ~ a s i n ~ [ ~ ~ ~ ~ ] in the absence of the liner and the dynamic pressure q,,

.

They are defined as

and

13.3.1 Pressure loss parameters

Both Lefebvre (1998) and Cohen et a1.(1996) noted two dimensionless pressure-loss parameters i.e.: (1) The total pressure loss to inlet pressure ratio (also known as the overall pressure loss)

- and (2) the total pressure loss to dynamic pressure ratio (also known as the pressure Pi,,

loss factor[PLu) ( A P ; , , - ~ ~ ~ ,

.

The two parameters are related:

The conceptual design for development of a micro gas turbine generator.

Appendix D: Combustion chamber performance characteristics

Lefebvre (1998) tabulated typical pressure losses in combustion chambers. These values are, however, for cold losses only.

Table 13.1 : Pressure-losses in combustors.

Losses due to combustion are normally not included in the overall pressure loss and values are in the range of 4 to 8 % depending on operational conditions. Flow resistance introduced into the air stream between the recuperator outlet and the turbine inlet is represented by the PLF and is a fixed property of the combustor chamber and corresponds to the sum of the pressure drop in the diffuser and the drop across the combustor liner:

Combustor type Tubular

Tubo-annular Annular

qrd 4Wf 9nf

The pressure loss over the diffuser will be discussed later in this chapter under the liner geometry section. Pressure drop over the liner has to be sufficiently high as it promotes high injection air velocities, steep penetration angles, high turbulence levels that encourage good mixing, which can result in a shorter liner. An optimum pressure loss over the liner has to be established as a high pressure loss over the combustor reduces some work done by the compressor, and is thus a disadvantage to the system output.

The PLF over the liner is determined by the total effective hole area [Ah[effl ] in the liner.

i n -

)

Pin 0.07 0.06 0.06

Lefebvre (1998) derived that the effective inlet hole area in the liner is determined by the casing

( b i n - o u r

)

qrd 37 28 20

0.0036 0.0039 0.0046

13.3.2 Relation between size and pressure loss

By inspection it seems that only the maximum casing area [A,] is variable since all the other values are fixed by the combustor design. This casing area needs to be optimized in order to achieve the best fuel consumption, while keeping the overall pressure loss at a minimum. One percent increase in pressure loss results into either half a percentage point reduction in thrust, or a quarter percentage point increase in fuel consumption. For use in the TCIR cycle, a large diameter and low-pressure-loss is required and this will result into low fuel consumption. The overall pressure loss dictates the casing size, and Aref is obtained from

Optimization of Are, can be done by providing as large as possible liner cross sectional area. This will result in lower velocities and longer combustion times, which are beneficial to ignition and combustion efficiency. However, the only possibility of obtaining a larger liner diameter is to enlarge the annulus area. This will raise the annulus velocity and lower the annulus static pressure, thus reducing the pressure drop over the liner holes. A high static pressure drop over the longer holes is needed to ensure adequate penetration and sufficient turbulence intensity to promote rapid mixing with the combustion products. Lefebvre (1998) suggested that the optimal value for the ratio of liner cross-sectional area to casing cross-sectional area [k] can be derived from:

r 1

L

^qr,/

J

with

A

the diffuser pressure loss and r the ratio of casing area to combustor inlet area and results k = l -

The value of k may be assumed as 0.7.

(1-+A

i n - A r 2

--

13.3.3 Flow in the annulus and line inlet holes

Flow conditions in the annulus have an effect on the airflow pattern within the liner and the level and distribution of liner temperatures. Gas velocity in the annulus is governed by the combustion reference velocity as well as the ratio of liner area to casing area. Therefore, a change in the velocity profile further along the liner occurs due to air that is drawn through the liner holes and cooling ducts from the annulus. Gas flow through the liner inlet holes depends on factors shush as the size and pressure drop over the inlet hole, and on duct geometry and flow conditions in the vicinity of the inlet hole:

lJ =- m

ref P i n

Lefebvre (1 998) expressed the equation for flow through a hole as

The conceptual design for development of a micro gas turbine generator.

Appendix D: Combustion chamber performance characteristics

with

4

the total pressure upstream of the inlet hole, p j the static pressure downstream of the inlet hole.

The discharge coefficient [Cd] of a hole is the relation between the smallest area of the flow and the actual area of the flow. The discharge coefficient is used to describe the inlet hole's characteristics and the flow through the hole. Only plain, circular liner inlet holes will be utilized in this study, and for non-swirling flow, the coefficient of discharge for plain circular holes is defined by Lefebvre (1 996) as

with

a

the relation of mass flow through the hole [h,] to the annulus mass flow rate[m,,,] ratio and K the ratio of the jet dynamic pressure to the annulus dynamic pressure upstream of the holes.

K should vary between 2 and 6.

\ \

Figure 13.3: Illustration o f flow through liner wall.

Other factors that influence the combustor chamber's aerodynamics that will not be discussed in this study include jet trajectories and duct geometry.