MAE 423 Prject Report
-
Upload
deepak-kumar -
Category
Documents
-
view
227 -
download
1
Transcript of MAE 423 Prject Report
MAE 423Contemporary Issue Project
Deepak Kumar
Dr Paul E. DesJadin
Table of Contents
Problem statement ................................................................................................ 2
Introduction ........................................................................................................... 3
Assumptions .......................................................................................................... 4
Method of solution ............................................................................................... 5
Discussion and results ........................................................................................... 7
Summary & conclusion ........................................................................................
16
References ........................................................................................................... 17
1
Problem Statement
A non-ideal turbojet is to be studied at varying Mach numbers. Specific thrust, thrust specific
fuel consumption, propulsion efficiency, thermal efficiency, and overall efficiency will be
analyzed at the given set of Mach numbers. We will also be studying about the effects of
compressor ratios with the above parameters. The Mach numbers used in this analysis will be
0.2, 0.4, 0.6, 0.8, 1.4, 1.8, and 2.0. The compressor ratio that will be used for the study will be
between 2 and 100, in an increment of 0.5. Finally the stoichiometric ratio range is set to be
0.5 < ф <1.5.
The turbojet engine to be analyzed is given by figure 1
Figure 1
2
Introduction
This project is an analysis of an air breathing turbojet engine. The ambient air from the free
stream flow is drawn in trough the diffuser into the compressor. The air velocity is decreased as
the air is carried to the compressor. The air is then compressed in a dynamic compressor. The
compressor increases the pressure and the temperature of the air. Work is done by the
compressor to obtain the required compression ratio, the resulting temperature change is
dependent on the efficiency of the compressor. The air is once again heated in the combustion
chamber by burning fuel in an air and fuel mixture. The high temperature and pressure gas is
allowed to expand through a turbine to generate the necessary power needed to drive the
compressor. During this process there is a loss in the temperature and pressure of the gas. As
the gas leaves the turbine the gas is still at a higher temperature compared to the ambient
temperature, as a result the turbine inlet temperature is high. The air is finally accelerated and
exhausted through the nozzle. Engine cooling system uses the relatively cool air from the
compression system that bypasses the combustor via air system flow paths to cool the turbine
nozzle guide vanes and blades to ensure acceptable metal temperatures at very high gas
temperatures.
3
Assumptions
The following are the assumption used to calculate all data for this study.
Table 1
M1 2πb 0.93ηt 0.9Pe Pa
Pa (P1) 20000 paTa (T1) 216 K
Qr 42000000 J/Kgηc 0.85ηb 0.85ηn 0.95
Tturb inlet 1750 Kϒ 1.4
Ua=M2sqrt(ϒRTa) 589.198778 m/s
ϒc 1.4ϒt 1.3R 287 J/ Kg-Kπc 2τλ 8.101851852
Cpc 1004.5Cpt 1243.666667
Cpc/Cpt 0.807692308φstc 0.069097569JP8 166 MWAir 2402.4 MW
4
Method of solutions
Air to fuel ratio
The Chemical reaction of JP8 and air:
Molecular Weights:
Carbon = 12 g/mol
Oxygen = 16 g/mol
Nitrogen = 14 g/mol
Hydrogen = 1g/mol
Therefore the fuel to air ratio is as follows
5
Specific Thrust
I = specific thrust
Thrust Specific Fuel Consumption
Propulsion Efficiency
Thermal Efficiency
Overall Efficiency
6
Discussion and Results
a)
1 10 1000.0000
200.0000
400.0000
600.0000
800.0000
1000.0000
1200.0000
Specific Thrust vs πc
0.2 mach0.4 mach0.6 mach0.8 mach1.4 mach1.8 mach2.0 mach
Compressor Ratio
Spec
ific t
hrus
t
Graph 1
The graph above represents the effects of increasing Mach number and compression ratio on
the Specific Thrust. It is clearly seen from the graph that the specific thrust reduces as the Mach
number increases. The increase in the compression ratio increases the specific till it reaches a
max specific thrust as the components of the engine reach performance limits. All the values
are constrained by the stoichiometric ratio range.
7
1 10 1000.0000
0.0500
0.1000
0.1500
0.2000
0.2500
0.3000
TSFC vs πc
0.2 mach0.4 mach0.6 mach0.8 mach1.4 mach.8 mach2.0 mach
Compressor ratio
TSFC
Graph 2
From the graph above we can see that the TSFC increases as the Mach number increases within
the same range of the stoichiometric ratio. From the graph it can also be deduced that the inlet
pressure ratio reduces to produce the same thrust with the increasing fuel flow.
8
1 10 1000.0000
0.1000
0.2000
0.3000
0.4000
0.5000
0.6000
0.7000
Propulsion efficiency vs πc
M=0.2M=0.4M=0.6M=0.8M=1.4M=1.8M=2
Compressure pressure ratio
Prop
ulsi
on E
ffici
ency
Graph 3
It is obvious from the above graph that the propulsion efficiency is approximately 0.65 at Mach
2.0 and for Mach 0.2 is between 0.2 and 0.1. The propulsion efficiency decreases with
increasing compressor ratio. The propulsion efficiency also tends to remain constant after
reaching its critical conditions.
9
1 10 1000.1000
0.1500
0.2000
0.2500
0.3000
0.3500
0.4000
0.4500
Thermal efficiency vs πc
M=0.2M=0.4M=0.6M=0.8M=1.4M=1.8M=2
Compressure pressure ratio
Ther
mal
Effi
cienc
y
Graph 4
From the graph above, the thermal efficiency increases with the increasing compressor ratio.
The max thermal efficiency at all Mach number seem to be approximately close to each other.
10
1 10 1000.0000
0.0500
0.1000
0.1500
0.2000
0.2500
0.3000
Overall efficiency vs πc
M=0.2M=0.4M=0.6M=0.8M=1.4M=1.8M=2
Compressure pressure ratio
Ove
rall
Efficie
ncy
Graph 5
As the Mach number increases, the overall efficiency increases with increasing compressor ratio.
11
b) Maximum values of I and TSFC
Table 2
M πc I [Ns/Kg]TSFC [Kg/N
hr] ηp ηth ηo
0.2 35.0 1058.577 0.253 0.159 0.445 0.04910.4 32.5 1004.676 0.260 0.281 0.446 0.09370.6 28.5 954.070 0.258 0.370 0.448 0.13510.8 24.0 906.695 0.251 0.436 0.449 0.17381.4 12.0 782.889 0.227 0.635 0.445 0.27091.8 7.0 714.236 0.216 0.824 0.441 0.30242.0 5.0 683.784 0.314 0.944 0.439 0.3136
c) We can conclude from the table given in part b, that max specific thrust reduces as the Mach
number climbs. The thrust specific fuel consumption reaches maximum with the as the Mach
number increases.
d) Lean fuel stability is constrained by the operating range as the flight Mach number is
increased such that the operating range of the compressor ratio decreases. Therefore the
turbojet engine will be operating at a much smaller range of operating conditions as the Mach
number increases. Essentially the turbojet engine will stall at very high Mach numbers. Thus
proving that the ram jet engines are more effective at very high Mach number than a turbojet
engine.
12
e)
1 10 1000
0.05
0.1
0.15
0.2
0.25
0.3
Propulsion Efficiencyvs πc
Mach = 1.8
Mach = 2.0
Single Shock Mach = 1.8
Single Shock Mach = 2.0
Compressor Ratio
Prop
ulsio
n Effi
cienc
y
Graph 6
Highest propulsion efficiency occurs at two-shock systems. The higher the Mach number the
higher propulsion efficiency.
13
1 10 1000
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0.8
0.9
Thermal Efficiencyvs πc
Mach = 1.8Mach = 2.0Single Shock Mach = 1.8Single Shock Mach =2.0
Compressor Ratio
Ther
mal
Effi
cienc
y
Graph 7
This graph shows the thermal efficiency of a single shock system has optimum efficiency. Higher
than the two shock system. As the Mach number increases the thermal efficiency increases.
14
1 10 1000
0.02
0.04
0.06
0.08
0.1
0.12
0.14
0.16
0.18
Overall Efficiency vs πc
Mach = 1.8Mach = 2.0Single Shock Mach = 1.8Single Shock Mach = 2.0
Compressor Ratio
Ove
rall
Efficie
ncy
Graph 8
This graph settles the results without a doubt that the two-shock systems is greater than the
overall efficiency of single shock systems.
15
Summary and Conclusion
From the analysis above it is needless to say that the higher the Mach number and
compression ratio is the lower the specific thrust and thrust specific fuel consumption. From
the efficiency graphs we can say that at higher Mach number the more efficient the turbojet
engine is.
From the shock analysis the two shock system is more efficient and preferred. This clear
shows that it is better to use oblique shocks. In conclusion the Turbojet engines perform more
efficiently at greater Mach numbers. At high supersonic speed it is preferable to have oblique
shocks than a single shock system. The range of the compressor points reduce with the given
range of the stoichiometric ratio.
16
Reference
http://mit.edu/16.unified/www/FALL/thermodynamics/notes/node85.html
http://en.wikipedia.org/wiki/Overall_pressure_ratio
Mechanics and Thermodynamics of Propulsion 2nd edition Philip Hill ,Carl Peterson
http://www.grc.nasa.gov/WWW/k-12/airplane/oblique.html
http://www.oocities.org/siliconvalley/7116/jv_aerom.html
17