190 International Journal of Advances in Arts, Sciences and Engineering, Volume 4 Issue 9 Sep 2016 2320-6144 (Online)

Design and Analysis Of Can-Type Combustion Chamer To Enhance Its

Performance

1.T.V.S. MANIKANTA 2.M.GOWRI SANKHAR

1. PG Scholar (Aerospace Engineering), NIMRA Institute of Science and Technology, , AP, India

2. Asst.Professor (Aerospace Engineering), NIMRA Institute of Science and Technology, , AP, India

E-Mail :-tvs.manikanta4@gmail.com

ABSTRACT:

The project entitled "DESIGN AND ANALYSIS OF CAN-TYPE COMBUSTION CHAMBER

TO ENHANCE ITS PERFORMANCE" is to design can type combustion chamber which gives

efficient results.

In this project, the role of combustion chamber in gas turbines and different types of

combustion chambers are studied. Among those can type combustion chamber is taken for the

project to increase its performance from basic type.

In basic can type combustion chamber the temperature and velocity of outflow is less.

Also it's wall temperature is more. As to overcome these drawbacks the new can type

combustion chamber is designed which is having blade vanes arrangement around it's

combustion chamber. Through these vanes the bleed air is made to flow. Because of this

process the wall temperature of combustion chamber is decreased.

Two types of can type combustion chambers are designed and analyzed in ANSYS

software. Their pressure, temperature and velocity results are compared.

The software used in this project is ANSYS ICEM CFD.

1. INTRODUCTION

The combustor in a gas turbine is to

add energy to the system to power the turbines,

and produce high velocity gas to exhaust

through the nozzle in aircraft applications.

Combustion chambers must be designed to

ensure stable combustion of the fuel injected

and optimum fuel utilization within the limited

space available and over a large range of

air/fuel ratios. In a gas turbine engine, the

combustor is fed by high pressure air by the

compression system. A combustor must

contain and maintain stable combustion despite

very high air flow rates. To do so combustors

are carefully designed to first mix and ignite the

air and fuel, and then mix in more air to

complete the combustion process.

To be more competent, the combustor model is

chosen from the literature work of Reddy and

Kumar [1]. This paper presents the

191 International Journal of Advances in Arts, Sciences and Engineering, Volume 4 Issue 9 Sep 2016 2320-6144 (Online)

experimental and numerical results for a two

stage combustor capable of achieving flameless

combustion. The concept of high swirl flows

has been adopted to achieve high internal

recirculation rates in flameless combustion

mode. Computational analysis of the flow

features shows that decrease in the exit port

diameter of the primary chamber increases the

recirculation rate of combustion products and

helps in achieving the flameless combustion

mode. Detailed experimental investigations

show that flameless combustion mode was

achieved with evenly distributed combustion

reaction zone and uniform temperature

distribution in the combustor. The preference of

natural gas is chosen from this investigation.

Industrial gas turbines have a wider scope of

fuel. Ghenai [2] has done numerical

investigation of the combustion of syngas fuel

mixture in gas turbine can combustor to

understand the impact of the variability in the

alternative fuel composition and heating value

on combustion performance and emissions. The

composition of the fuel burned in can

combustor was changed from natural gas

(methane) to syngas fuel with hydrogen to

carbon monoxide (H2/CO) volume ratio

ranging from 0.63 to 2.36. Results show the

changes in gas turbine can combustor

performance with the same power generation

when natural gas or methane fuel is replace by

syngas fuels. The gas temperature for the all

five syngas shows a lower gas temperature

compared to the temperature of methane. The

gas temperature reduction depends on lower

heating value and the combustible and non-

combustible constituents in the syngas fuel

which results in less emission.

2.LITERATURE REVIEW

2.1 COMBUSTION

Combustion is a chemical process in

which a substance reacts rapidly with oxygen

and gives off heat. The original substance is

called the fuel, and the source of oxygen is

called the oxidizer. The fuel can be a solid,

liquid, or gas, although for airplane propulsion

the fuel is usually a liquid. A combustor is a

component or area of a gas turbine, ramjet,or

scramjet engine where combustion takes place.

It is also known as a burner, combustion

chamber or flame holder. In a gas turbine

engine, the combustor or combustion chamber

is fed high pressure air by the compression

system. The combustor then heats this air at

constant pressure. After heating, air passes

from the combustor through the nozzle guide

vanes to the turbine. In the case of a ramjet or

scramjet engines, the air is directly fed to the

nozzle.

2.2 TYPES OF COMBUSTION IN

VARIOUS ENGINES

2.2.1 PISTON ENGINE

A piston is a component of

reciprocating engines,

reciprocating pumps,

gas

compressors and pneumatic cylinders, among

other similar mechanisms. It is the moving

component that is contained by a cylinder and

is made gas-tight by piston rings. In an engine,

transferred from the crankshaft to the piston for

the purpose of compressing or ejecting the fluid

in the cylinder. In some engines, the piston also

192 International Journal of Advances in Arts, Sciences and Engineering, Volume 4 Issue 9 Sep 2016 2320-6144 (Online)

acts as a valve by covering and uncovering

ports in the cylinder wall.

A reciprocating engine, also often known as a

piston engine, is a heat engine (usually,

although there are also pneumatic and hydraulic

reciprocating engines) that uses one or more

reciprocating pistons to convert pressure into a

rotating motion. These engines are also

classified in two ways: either a spark-ignition

(SI) engine, where the spark plug initiates the

combustion; or a compression-ignition (CI)

engine, where the air within the cylinder is

compressed, thus heating it, so that the heated

air ignites fuel that is injected then or earlier.

2.2.2 JET ENGINE

Jet engines move the airplane forward

with a great force that is produced by a

tremendous thrust and causes the plane to fly

very fast.

All jet engines, which are also called gas

turbines, work on the same principle. The

engine sucks air in at the front with a fan. A

compressor raises the pressure of the air. The

compressor is made with many blades attached

to a shaft. The blades spin at high speed and

compress or squeeze the air. The compressed

air is then sprayed with fuel and an electric

spark lights the mixture. The burning gases

expand and blast out through the nozzle, at the

back of the engine. As the jets of gas shoot

backward, the engine and the aircraft are thrust

forward. As the hot air is going to the nozzle, it

passes through

Figure 2.1– Basic components of Jet Engine.

another group of blades called the turbine. The

turbine is attached to the same shaft as the

compressor. Spinning the turbine causes the

compressor to spin.

2.2.3 ROCKET ENGINE

A rocket engine is a type of jet engine

that uses only stored rocket propellant mass for

forming its high speed propulsive jet. Rocket

engines are reaction engines, obtaining thrust in

accordance with Newton's third law. Most

rocket engines are internal combustion engines,

although non-combusting forms also exist.

Vehicles propelled by rocket engines are

commonly called rockets. Since they need no

external material to form their jet, rocket

engines can perform in a vacuum and thus can

be used to propell spacecraft and ballistic

missiles.

2.3 TYPES OF COMBUSTIONS IN JET

ENGINE

2.3.1 LOW SPEED COMBUSTION

Low-speed (low Mach number)

combustion is important in a variety of

contexts, including furnaces, spark ignition

engines, forest fires, and rocket motors.

Fundamentally it is concerned with problems

of ignition and with flames. This workshop will

focus on two areas in which a large number of

fundamental combustion topics are relevant,

193 International Journal of Advances in Arts, Sciences and Engineering, Volume 4 Issue 9 Sep 2016 2320-6144 (Online)

namely the burning of solid propellants, and of

liquid fuel sprays.

2.3.2 SUBSONIC COMBUSTION

Subsonic aerodynamics studies fluid motion in

flows which are much lower than the speed of

sound everywhere in the flow. There are several

branches of subsonic flow but one special case

arises when the flow is in-viscid,

incompressible and ir-rotational. This case is

called potential flow and allows the differential

equations used to be a simplified version of the

governing equations of fluid dynamics, thus

making available to the aerodynamicist a range

of quick and easy solutions

2.3.3 SUPERSONIC COMBUSTION

Generally, in an air-breathing engine,

flow speed in a combustor becomes faster as

flight speed increases. The flow speed in the

combustor can be supersonic (over Mach

number 1) when the flight speed becomes

hypersonic (over Mach number 5). The air-

breathing engine being that flow speed in the

combustor is supersonic, called Scramjet

engine, is one of the key technologies to

develop hypersonic air-breathing propulsion

system.

Figure 2.2- Schematic of SCRAMJET

engine.

2.4. COMBUSTION SECTION

The combustion section contains the

combustion chambers, igniter plugs, and fuel

nozzle or fuel injectors. It is designed to burn a

fuel-air mixture and to deliver combusted gases

to the turbine at a temperature not exceeding

the allowable limit at the turbine inlet.

Theoretically, the compressor delivers 100

percent of its air by volume to the combustion

chamber. However, the fuel-air mixture has a

ratio of l5 parts air to 1 part fuel by weight.

Approximately 25 percent of this air is used to

attain the desired fuel-air ratio. The remaining

75 percent is used to form an air blanket around

the burning gases and to dilute the temperature,

which may reach as high as 3500º F, by

approximately one-half. This ensures that the

turbine section will not be destroyed by

excessive heat.

2.4.1 CAN-TYPE COMBUSTION

CHAMBER

2.4.1.1 Basics of Can-Type Combustion

Chamber

The combustion chamber (Fig.1) has the

difficult task of burning large quantities of fuel,

supplied through the fuel spray nozzles, with

extensive volumes of air, supplied by the

compressor and releasing the heating such a

manner that the air is expanded and accelerated

to give a smooth stream of uniformly heated gas

at all conditions required by the turbine. This

task must be accomplished with the minimum

loss in pressure and with the maximum heat

release for the limited space available. The

amount of fuel added to the air will depend

upon the temperature rise required. However,

the maximum temperature is limited to within

the range of 850 to 1700 deg C. by the materials

from which the turbine blades and nozzles are

made. The air has already been heated to

between 200 and 550 deg. C. by the work done

during compression, giving a temperature rise

requirement of 650 to 1150 deg.

194 International Journal of Advances in Arts, Sciences and Engineering, Volume 4 Issue 9 Sep 2016 2320-6144 (Online)

Figure 2.3- Basic Construction Of

Combustion Chamber.

3.1 INTRODUCTION TO ANSYS ICEM

Meeting the requirements for

integrated mesh generation and post processing

tools for today’s sophisticated analysis.

ANSYS ICEM CFD provides geometry

acquisition, mesh generation, mesh

optimization, and post processing tools.

Maintaining a close relationship with

the geometry during mesh generation and post

processing, ANSYS ICEM CFD is used

especially in engineering applications such as

computational fluid dynamics and structural

analysis

ANSYS ICEM CFD’s mesh

generation tools offer the capability to

parametrically create meshes from in numerous

formats

Multiblock structured

Unstructured hexahedral

Cartesian with h grid refinement

Hybrid Meshes comprising hexahedral,

tetrahedral, pyramidal and/or prismatic

elements

Quadrilateral and triangular surface meshes

ANSYS ICEM CFD provides a direct link

between geometry and analysis. In ANSYS

ICEM CFD, geometry can be put just above

any format, whether it is form a commercial

CAD design package, 3rd party universal

database, scan data or point data.

Beginning with a robust geometry module

which supports the creation and modification

of surfaces, curves and points, ANSYS ICEM

CFD’s open geometry database offers the

flexibility to combine geometric information in

various formats for mesh generation. The

resulting structured or un structured meshes,

topology, inter-domain connectivity and

boundary conditions are then stored in database

where they can easily be translated to input files

formatted for a particular solver.

Mesh visualization tools, including

solid/wireframe display, 2D cut planes,

color coding and node display is provided.

ANSYS ICEM CFD visual provides

easy-to-use powerful result visualization

features for structured, unstructured and

hybrid grids, both steady-state and

transient.

3.2 DESIGNING OF BASIC CAN-TYPE

COMBUSTION CHAMBER

Figure 3.1- Geometry for Basic Can type

combustion chamber

3.2.2 DESIGNING IN ICEM

Start menu > all programs > ansys > meshing

> ICEM

Figure 3.2- Ansys ICEM CFD Home Page.

INLET

Create two circles of radius 5 and 10. By

creating surface using two circles inlet area

can be created and by extruding 15 units

195 International Journal of Advances in Arts, Sciences and Engineering, Volume 4 Issue 9 Sep 2016 2320-6144 (Online)

two cylinders can be create as shown in

figure.

Figure 3.3- Inlet section for Basic Can-Type.

FUEL INLET

At first the surface has to be created

and by using fuel inlet lines the surface

created can trimmed.

Figure 3.4- Fuel Inlets for Basic Can-Type

VANES

Creating vanes by using co-ordinates.

Figure 3.5- Vanes for Basic Can Type.

COMBUSION CHAMBER WALLS

Creating combustion chamber walls by

driving a circular ark and a straight line

along a circle.

Figure 3.6- Combustion Chamber wall for

Basic Can-Type.

NOZZLE

Creating nozzle according to points in

geometry.

Figure 3.7- Nozzle section for Basic Can

Type.

COMPLETE BASIC CAN-TYPE

COMBUSTION CHAMBER

Figure 3.8- Complete Design of Basic Can-

Type Combustion Chamber.

4.MESHING

4.1 MESHING OF BASIC CAN-TYPE

COMBUSTION CHAMBER DESIGN

The partial differential equations that govern

fluid flow and heat transfer are not usually

196 International Journal of Advances in Arts, Sciences and Engineering, Volume 4 Issue 9 Sep 2016 2320-6144 (Online)

amenable to analytical solutions, except for

very simple cases. Therefore, in order to

analyze fluid flows, flow domains are split into

smaller subdomains (made up of geometric

primitives like hexahedra and tetrahedral in 3D

and quadrilaterals and triangles in 2D). The

governing equations are then discredited and

solved inside each of these subdomains.

Typically, one of three methods is used to solve

the approximate version of the system of

equations: finite volumes, finite elements, or

finite differences. Care must be taken to ensure

proper continuity of solution across the

common interfaces between two subdomains,

so that the approximate solutions inside various

portions can be put together to give a complete

picture of fluid flow in the entire domain. The

subdomains are often called elements or cells,

and the collection of all elements or cells is

called a mesh or grid.

4.1.1 EDITING PART MESH SETUP

Figure 4.1- Editing part mesh setup file in

ICEM CFD.

4.1.2 COMPUTING MESH

Figure 4.2- Computing mesh file in ICEM

CFD.

4.1.3 EXPOTING MESH FILE

Figure 4.3- Exporting mesh file in ICEM CFD

4.2 MESHING OF MODIFIED CAN

TYPE COMBUSTION CHAMBER

DESIGN

4.2.1 EDITING PART MESH SETUP

197 International Journal of Advances in Arts, Sciences and Engineering, Volume 4 Issue 9 Sep 2016 2320-6144 (Online)

Figure 4.4- Meshing file for Modified Can

Type in ICEM CFD.

4.2.2 COMPUTING MESHING

Meshing of inner surfaces

Figure 4.5- Computing

meshing file for Inner

surfaces. Meshing of

total surface

Figure 4.6- Computing

meshing file for Total

surface

5.1 RESULTS FOR BASIC CAN-TYPE

5.1.1 INLET FLOW VELOCITY AT 10

m/sec

CH 4 Mass fraction CO2 Mass fraction

Figure 4.1- CH 4

Mass fraction for for

Basic Can Type at 10

m/sec

Figure 4.2- CO2

Mass fraction Basic

Can Type at 10

O2 Mass fraction Temperature

Figure 4.3- O2 Mass

fraction for

Basic Can Type at 10

m/sec.

Figure 4.4-

Temperature for

Basic

Can Type at 10

m/sec.

Pressure Velocity

Figure 4.5- Pressure

for Basic

Can Type at 10

m/sec.

Figure 4.6- Velocity

for Basic

Can Type at 10

m/sec.

5.1.2 INLET FLOW VELOCITY AT 30

m/sec

198 International Journal of Advances in Arts, Sciences and Engineering, Volume 4 Issue 9 Sep 2016 2320-6144 (Online)

CH 4 Mass

fraction

CO2 Mass fraction

Figure 5.7- CH4

mass fraction for

Basic Can Type

at 30 m/sec

Figure 5.8- CO2 mass

fraction for

Basic Can Type at 30

O2 Mass fraction Temperature

Figure 4.9- O2

mass fraction for

Basic Can Type

at 30 m/sec.

Figure 4.10-

Temperature for Basic

Can Type at 30 m/sec.

Pressure Velocity

Figure 4.11-

Pressure for

Basic

Can Type at 30

m/sec.

Figure 4.12- Velocity

for Basic

Can Type at 30 m/sec.

CONCLUSION

The design of Can-type combustion

chamber, modified can-type combustion

chamber geometry and numerical

investigations is carried out. The k-ω model

used for analysis and also the mean

temperature, reaction rate, and velocity fields

are almost insensitive to the grid size.

Numerical investigation on Can-type

combustion chamber and a modified can-type

combustion chamber geometry is gives less NO

emission as the temperature at the exit of

combustion chamber is less. For methane as

fuel and with initial atmospheric conditions, the

theoretical flame temperature produced by the

flame with a fast combustion reaction is 1950

K. The predicted maximum flame temperature

is 1850 K of the combustion products compares

well with the theoretical adiabatic flame

temperature. Temperature profiles shows

increment at reaction zone due to burning of

air-methane mixture and decrement in

temperature downstream of dilution holes

because more and more air will enter in

combustion chamber to dilute the combustion

mixture along center line . Specie namely NO

is increasing and achieving peak point at

reaction zone because they are products of

combustion along center line. Due to increase

in equivalence ratio, temperature and mass

fraction of NO increases because more fuel is

utilized. There in not much variation in

temperature and NO emission by shifting the

axial location of dilution holes. In modified can

-type combustion chamber geometry

Temperature profiles shows increment at

reaction zone along the axis due to burning of

air-methane mixture and decrement in

temperature downstream the walls. In modified

can-type combustion chamber clearly shows

199 International Journal of Advances in Arts, Sciences and Engineering, Volume 4 Issue 9 Sep 2016 2320-6144 (Online)

that temperature and pressure profiles decrease

and contribute to cool the chamber walls but the

exit velocity profile contributes for some

losses. The streamline wall cooling is provided

by installing the vanes in the combustion

chamber at different position which enables the

proper wall cooling for the design considered.

REFERENCES

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for achieving flameless combustion,”

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Advances in Mechanical Engineering,

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