CFD Modeling Of An Aero Gas Turbine Combustor For A Small Gas Turbine Engine

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,

Bangalore-560093.

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 one-step overall chemical reaction between JET-A / 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.

1. 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 re-light 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.

2. 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.

   Table-1: 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

   Table-2: 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

   Table-3: 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 zone-1	2*144		2
   Dilution zone-2	244	307	0.4
   Liner cooling-1	1040	1008	0.6
   Liner cooling-2	848		1

   Fig.2: Reverse Flow Combustor Configuration

   CFD Modeling Of The Combustor

   The challenging task in aero gas turbine combustor CFD modeling is two-fold, 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 22.5° 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 22.5° 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 table-3

3. 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.

4. Boundary Conditions

   The boundary conditions used for the analysis are given in Table-1, The simulation was performed using commercial Computational fluid dynamics code ANSYS-CFX, in which a three-dimensional compressible k- turbulent flow model and a one-step overall chemical reaction between JET-A / air was used. The fuel used in the present study is Jet-A 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.

   Table-4: 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

5. 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.

6. Basic Assumptions

   • Geometrical Approximations

   • 3D model analysis

   • Thickness of the sheet element is ignored

7. 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 pressure-swirl 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

8. 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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