 Open Access
 Total Downloads : 3658
 Authors : K. Sreenivasarao, S.K. Bhatti
 Paper ID : IJERTV2IS1478
 Volume & Issue : Volume 02, Issue 01 (January 2013)
 Published (First Online): 30012013
 ISSN (Online) : 22780181
 Publisher Name : IJERT
 License: This work is licensed under a Creative Commons Attribution 4.0 International License
CFD Modeling Of An Aero Gas Turbine Combustor For A Small Gas Turbine Engine
K. Sreenivasarao, Manager (Design), Aero Engine Research and Design Centre, Hindustan Aeronautics Limited,
Bangalore560093.
Prof. S.K. Bhatti, Department of Mechanical engineering, AU College of Engineering, Vishakapatnam.
Abstract
This paper discusses CFD applications to the analysis in small, high intensity complex flow field gas turbine combustors. Performance of combustor has been
predicted using a commercially available CFD code ANSYS CFX. The entire flow field from the inlet to exit of the combustor, including the liner injection holes, has been modeled with reasonable accuracy. Comparison of the quality of temperature distribution and combustion efficiency at combustor exit are presented. A close agreement is observed between the predictions and experiments. Computational Fluid Dynamic analysis of 3D combustor flows and the empirical validation of results are used to demonstrate the effectiveness of CFD as a design tool in this challenging environment. It is demonstrated that
commercial CFD codes provide qualitative and reasonable quantitative information that can be effectively used in the combustor design optimization
A reverse flow annular combustor for a small gas turbine engine have been modeled. The simulations were performed using commercial Computational fluid dynamics code ANSYS CFX, in which a three dimensional compressible k turbulent flow model and a onestep overall chemical reaction between JETA / air were used. Reverse flow annular combustor with simplex type of fuel injectors were used.
Keywords: Aero Engine Annular, Design, Combustion Chamber, CFD, ANSYS CFX.

Introduction
The conventional approach to the design and development of combustion systems for gas turbine engines involves extensive
use of empirical correlations, derived form experimental investigations and component development tests. The designer of a combustion system for a gas turbine engine is given the task of achieving the desired goals of the combustor, namely, complete combustion, low total pressure loss, proper temperature distribution at exit with no hot spots, highly stable combustion, freedom from flameout, good relight capability and operation over a wide range of mass flow rates, pressure and temperatures in small size.
The main objective of the present work is to select a suitable combustor for a small gas turbine engine using empirical correlations and a commercial CFD code CFX from ANSYS by modeling the complete geometry.

Design of Annular Combustion Chamber
Saravanamutto et al.
/19/ has given cycle analysis and performance characteristics of individual components of gas turbine engine. The data for designing Annular straight and reverse flow type combustion chambers is presented in Table I.
The design of the combustion chamber is carried out as outline in the literature
/1,2,3,6,10,11,12,13,18/.
Table 2 gives the – dimensional details of combustion chamber designed. Figure 2 shows the sector models of combustion chambers.
Table1: Design data for combustion chamber
S.No
Description
Value
1
Entry Pressure
10 bar
2
Entry temperature
640 K
3
Entry Flow
Air
4.7 kg/s
4
Assumed pressure loss
5 %
5
Fuel ratio
/
Air
0.018786
6
Fuel flow
303
kg/hr
7
Turbine entry Temperature
1300 K
8
Limiting value of liner temperature
1000 K
Table2: Geometrical parameters
S.No
Description
Value
1
Combustion chamber casing outer diameter (mm)
330
mm
2
Combustion chamber inner casing
diameter (mm)
105
mm
3
Flame tube
inner liner diameter (mm)
304
mm
4
Flame tube
outer liner diameter (mm)
224
mm
5
Flame tube length (mm)
105
mm
Table3: Air distribution holes
Number of holes
dia (mm)
inner
outer
Dome cooling
200
0.4
Primary zone
48
80
5
Secondary zone
80
128
3.5
Dilution zone1
2*144
2
Dilution zone2
244
307
0.4
Liner cooling1
1040
1008
0.6
Liner cooling2
848
1
Fig.2: Reverse Flow Combustor Configuration
CFD Modeling Of The Combustor
The challenging task in aero gas turbine combustor CFD modeling is twofold, namely, the development of models to describe the real world component geometries and an accurate description of the coupled interacting physical and chemical phenomena. An accurate modeling of the combustor geometry is required for arriving at realistic predictions. Initially, combustor simulations were limited to the flow field inside the liner and the diffuser / combustor annulus region. The solutions obtained by these simulations are strongly dependent on the accuracy levels associated with the description of boundary conditions. However, from the past few years, simulations are being carried out by these simulations are strongly dependent on the accuracy levels associated with the
description of boundary conditions. However, from the past few years, simulations are being carried out by modeling the entire flow field from compressor exit to the turbine entry.
Physically phenomena that have to be addressed are turbulent transport, fuel spray atomization and vaporization and finite rate chemistry of the reactants. Based on the regime of operation of an aircraft engine, different components of these phenomena become more important while the others are less significant. Further, there are two main requisites for a correct evaluation of CFD models: the availability of a set of experimental data accurate and detailed
enough to
enable
the
specification
of
the
boundary cond
itions, and a
comparison
between
experimental
and
numerical
values.
Grid Generation
A 22.5Â° Sector 3D model was generated by using UNIGRAPHICS. Model was meshed with around 30,00,000 tetrahedral elements by using ICEM TETRA.
Fig.3: Mesh generation
The combustor considered in the present analysis is annular reverse flow with 16 swirlers equally spaced along the circumferential direction. Hence, the calculations are conducted for the flow in a 22.5Â° sector of the combustor with periodic boundary conditions being imposed on the two side planes. The outer and inner liners are provided with one row each of primary, secondary and dilution zones. In addition, the liners are cooled by effusion cooling. The air distribution through the liner holes is given in table3

CFD Code
The commercially available CFD code ANSYS CFX is used for the analysis. ANSYS CFX solves a set of coupled transport equations for fluid, heat and mass transport processes in laminar or turbulent, compressible or incompressible, transient or steady and flows at any speed. It can also cater
for chemical reactions, combustion, thermal radiation, soot formation, droplet transport, embedded moving boundaries, complex fluid property relations and a number of other fluid flow phenomena.

Boundary Conditions
The boundary conditions used for the analysis are given in Table1, The simulation was performed using commercial Computational fluid dynamics code ANSYSCFX, in which a threedimensional compressible k turbulent flow model and a onestep overall chemical reaction between JETA / air was used. The fuel used in the present study is JetA aviation kerosene; It is assumed to be in a fully vaporized state and a single phase calculation has been performed. The fuel is injected uniformly through an annular opening inside the swirler thus closely representing the actual atomizer. No slip condition is imposed at all the internal as well as the external walls of the computational domain.
Table4: Boundary conditions
Inlet
Outlet
Fuel Inlet
Total Pressure =
10 bar
Mass flow rate of air
= 4.73
kg/sec
Mass flow rate of fuel
=0.73 kg/sec
Total Temperature
= 640 K
–
Total Temperature
=303 K

Numerical Simulations: There are two ways to analyze the combustion chamber numerically. One way is to give input conditions at inlet and all the air admission holes as per the design conditions. But in actual case, the flow distribution in different zones" cannot be controlled. This is the biggest drawback of providing different inputs at different air admission holes.
The second way of analyzing the combustion chamber is to provide only one inlet at the inlet of the diffuser and let the flow divide by itself into liner and casing, and from casing into different zones through air admission holes and cooling slots. Such condition is the exact replica of the real case experimentation, in which the air is supplied at the inlet diffuser with known conditions of pressure, temperature and velocity, and then, allowed to divide between the casing and the
liner with fuel injection at liner entry. Pressure swirl Atomizer is used for fuel injection in the present study. A three dimensional atomizer is modeled as shown in Figure 4.

Basic Assumptions

Geometrical Approximations

3D model analysis

Thickness of the sheet element is ignored


CFD Models

3D turbulent, steady and compressible

Shear stress transport turbulence model

Jet A liquid (C12 H23) as the fuel

Eddy dissipation method for combustion modeling

Solver ANSYS CFX
CFD Modeling of Jet A liquid (C12 H23) Fuel
The combustor modeling is carried out using eddy dissipation model based on the concept that chemical reaction is fast relative to the transport processes in the flow. When reactants mix at the molecular level, they instantaneously form products.
The liquid evaporation model is a model for particles with heat
transfer and one component of mass transfer, and in which the continuous gas phase is at a higher temperature than the particles. The model uses two mass transfer correlations depending on whether the droplet is above or below the boiling point. This is determined through the Antoine equation.
The LISA (Linearized Instability Sheet Atomization) model is able to simulate the effects of primary breakup in pressureswirl atomizers. With pressure swirl injectors, the fuel is set into a rotational motion and the resulting centrifugal forces lead to a formation of a thin liquid film along the injector walls, surrounding an air core at the center of the injector. Outside the injection nozzle, the tangential motion of the fuel is transformed into a radial component and a liquid sheet is formed. This sheet is subject to aerodynamic instabilities that cause it to break up into ligaments. For secondary breakup, TAB (Taylor Analogy Breakup) Model is used. This spray breakup models will lead to drop diameter distribution depending upon the
injection pressure
differential (Atomizer model is shown in Figure 4).
The exit conditions of Coefficient of discharge, drop diameter distribution, velocity and film thickness as obtained from CFD analysis of pressure swirl atomizer is given at a point in the present combustion chambers (Figure 5).
Fig.4: Atomizer model Fig.5: Atomized Liquid


Results and Discussions:
The quality of combustor exit temperature distribution is generally expressed in terms of non dimensional parameter viz., Pattern factor, which can be defined as follows:
Pattern Factor = (T4max T4avg) / (T4avg T3avg)
The CFD analysis was carried out with defined boundary conditions described in previous section. The temperature and velocity distribution as found using Commercial CFD Code CFX for gas turbine combustion chambers designed at design condition are given in Figure 6 to Figure 9. Temperature profile and efficiency of combustion chamber are given in Figure
10 and 11. Number of configurations with varied hole distribution in the combustion liner were analyzed. Circumferential and Radial Pattern factors achieved 0.235 and 0.1. Combustor pressure loss achieved 6.52%.
Fig.6: Total Temperature Distribution at outlet
Fig.7: Total Temperature Distribution On XY Plane
NGV RADIAL HEIGHT (
NGV RADIAL HEIGHT (
RADIAL PATTERN FACTOR VS NGV RADIAL HEIGHT
100
90
80
70
60
50
40
30
20
10
0
0.2
0.15 0.1 0.05 0 0.05 0.1
0.15
0.2
RADIAL PATTERN FACTOR
CFD Result Experimental result
RADIAL PATTERN FACTOR VS NGV RADIAL HEIGHT
100
90
80
70
60
50
40
30
20
10
0
0.2
0.15 0.1 0.05 0 0.05 0.1
0.15
0.2
RADIAL PATTERN FACTOR
CFD Result Experimental result
Fig.8: Total Pressure Distribution On XY Plane
Fig.9: Velocity Vector Distribution On XY Plane
Fig.9: Mach number variation On XY Plane
Fig.10: Temperature profile distribution
Combustion Efficiency Variation
Experimental data
Combustion Efficiency Variation
Experimental data
100
100
80
60
40
20
0
80
60
40
20
0
0
0.5
1
1.5
2
2.5
0
0.5
1
1.5
2
2.5
Air Loading Parameter
Air Loading Parameter
Combustion Efficiency
Combustion Efficiency
Fig.10: Combustion efficiency variation

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