Experiment 2 Ramjet
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Transcript of Experiment 2 Ramjet
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Abstract
Hot water rocket extract external energy of applied heat to provide energy to the propellant.
Water as the propellant is pressurised by heating it to high temperature in a closed tank. Due to
pressure, the boiling temperature is raised to be more than 100 C. Once the temperature reaches
saturation temperature, then water which has now turned into water vapour is ejected out of the
tank and accelerated by a convergent-divergent nozzle. This ejection of exhaust gas produced
thrust in forward direction, opposite of the ejection direction. The maximum thrust is found to be
49.4424 N and the propellant mass flow rate is 0.3172 kg/s. Exit velocity is then calculated to be
103.44816 and the corresponding specific impulse is 10.54 seconds.
Introduction
The word propulsion comes from the Latin propulsus,which is the past participle of the verb
propellere, meaning to drive away. In a broad sense propulsion is the act of changing the motion
of a body. Propulsion mechanisms provide a force that are initially at rest, changes a velocity, or
overcomes retarding forces when a body is propelled through a medium. Jet propulsion is a
means of locomotion whereby a reaction force is imparted to a device by the momentum of
ejected matter (Sutton & Biblarz, 2010). There are many type of engine that applies jet
propulsion principle. Some of them uses need the presence of air for operation while some can
operate in vacuum without the need of air.
For air breathing engine, there are gas turbine powered engine and ram powered engine.
Gas turbine engine uses turbine powered compressor to compress air before the air is being fed
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into the combustion chamber for combustion. On the other hand, ramjet uses shockwave to
compress air that is going to be combusted later.
Rocket, rocket engine or rocket motor is a non-air breathing propulsion system (does not
need surrounding air for combustion, and thus propulsion) that produces forward thrust by
accelerating mass through a nozzle (typically convergent-divergent type).Rocket propulsion is a
class of jet propulsion that produces thrust by ejecting matter stored in a flying vehicle called the
propellant (Ward, 2010).There are several energy sources applicable in rocket propulsion such
as chemical combustion, solar radiation, nuclear reaction, and thermal energy. These energy
applications have sub-divided the rocket propulsion into chemical propulsion, solar propulsion,
nuclear propulsion and thermal propulsion.
Thermal propulsion rocket engine uses propellant (working fluid) that is superheated by
external heat source to flow out of the system through a nozzle. Nuclear thermal rocket, solar
thermal rocket, laser thermal rocket, and steam rocket uses this principle to produce thrust for
propulsion. Steam rocket or hot water rocket has water as its propellant. The water is kept in a
closed pressure vessel, and then heated by external heat source to increase its saturated vapour
pressure to be higher than the ambient pressure. Once the water vapour inside the tank reaches
the designed temperature and pressure, it is then allowed to escape at high velocity from the tank
through a nozzle by opening the tank's valve. The release of high energy vapour through the
nozzle produces thrust that propel the rocket forward.
The thrust can be calculated by using this formula
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Objectives
1. To measure the thrust produced by the steam jet leaving the tank through the exit nozzle
2. To determine the mass flow of the hot water rocket
3. To calculate the exit velocity of the steam jet and hence to determine the performance of
the hot water rocket by calculating the specific impulse
Methods and Procedures
1.
A container is filled with 8-liter tap water and the water was poured into the hot water
rocket tank. This step is repeated for 4 times so that the total volume of water filled inside
the rocket tank is 32 litre.
2. The water is heated by external heat source which are two gas stoves. It is heated until it
reaches 120C.
3. Data Acquisition System (DAQ) is executed to record the readings of the water
temperature and to operate the solenoid valve.
4. When the temperature reaches 120C, one camera is positioned to record the thrust
measurement on the digital spring balance and another camera is positioned to take the
photograph of the steam jet. The digital spring balance is rezero-ed.
5. The solenoid valve is opened by DAQ switch and the measurement of the chamber
temperature is recorded until the spring balance stopped taking measurements.
6. Thrust readings are translated from video recording to Microsoft Excel and graph of
thrust against time is plotted. The temperature readings recorded by DAQ are also
exported to Microsoft Excel and plotted in graphical form.
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7. The important parameters, which are mass flow rate and stream jet exit velocity are
calculated to calculate the engine's specific impulse that will determine the performance
of the hot water rocket.
Figure 1 Experimental setup
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Figure 2 Heating process
Figure 3 Hot water rocket exhaust gas
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Figure 4 National Instrument DAQ module
Figure 5 Lab View Data Acquisition System circuit diagram
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Experiment Setup and Initial Condition Information
Table 1 Nozzle specifications (Zulfikli, 2011)
SpecificationsValues
Dt=0.010m
Diameter inlet, Di(m) 0.0254
Area inlet, Ai(m2) 5.067 x 10
-4
Throat diameter, Dt(m) 0.010
Throat area, At (m
2
) 7.85 x 10
-5
Exit diameter, De(m) 0.0142
Exit area, Ae(m2
) 1.59 x 10-4
Nozzle area expansion
ratio2.02
Nozzle contraction ratio 6.452
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Results
Table 2 Thrust and Chamber Temperature
Time (s) Thrust (kg) Thrust (N) Chamber Temperature (C)
0 0 0 110.198
1 5.04 49.4424 110.174
2 5.04 49.4424 110.134
3 5.04 49.4424 110.102
4 5.04 49.4424 110.059
5 5.04 49.4424 110.033
6 5.04 49.4424 109.976
7 5.04 49.4424 109.9338 5.04 49.4424 109.838
9 5.04 49.4424 109.77
10 5.04 49.4424 109.67
11 5.04 49.4424 109.548
12 5.04 49.4424 109.428
13 5.04 49.4424 109.357
14 5.04 49.4424 109.207
15 5.04 49.4424 108.989
16 5.04 49.4424 108.715
17 5.04 49.4424 108.32418 5.04 49.4424 108.214
19 5.04 49.4424 108.026
20 5.04 49.4424 107.686
21 5.04 49.4424 107.496
22 5.04 49.4424 107.251
23 3.66 35.9046 106.941
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24 3.48 34.1388 106.599
25 3.48 34.1388 106.384
26 3.48 34.1388 106.14
27 3.48 34.1388 105.719
28 3.48 34.1388 105.363
29 3.48 34.1388 104.992
30 3.48 34.1388 104.593
31 3.48 34.1388 104.302
32 3.48 34.1388 103.857
33 3.48 34.1388 103.433
34 3.48 34.1388 102.981
35 3.48 34.1388 102.578
36 3.48 34.1388 102.284
37 3.48 34.1388 101.812
38 3.48 34.1388 101.45
39 3.48 34.1388 101.059
40 3.48 34.1388 100.642
41 3.48 34.1388 100.414
42 3.48 34.1388 99.7611
43 2.84 27.8604 99.531
44 2.82 27.6642 98.7936
45 2.8 27.468 98.134
46 2.78 27.2718 97.6697
47 2.78 27.2718 97.2978
48 2.78 27.2718 96.738349 2.78 27.2718 96.2712
50 2.78 27.2718 95.9347
51 2.78 27.2718 95.4768
52 2.78 27.2718 95.1945
53 2.78 27.2718 94.7004
54 2.78 27.2718 93.9944
55 2.78 27.2718 93.4975
56 2.78 27.2718 92.9626
57 2.78 27.2718 92.4506
58 2.78 27.2718 91.986159 2.78 27.2718 91.4636
60 2.78 27.2718 90.9794
61 2.78 27.2718 90.4708
62 2.1 20.601 90.0217
63 2.08 20.4048 89.4852
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Figure 6 Thrust versus time graph
Figure 7 Chamber temperature versus time graph
0
10
20
30
40
50
60
0 10 20 30 40 50 60 70
Thrust(N)
Time (s)
Thrust Vs. Time
85
90
95
100
105
110
115
0.00 10.00 20.00 30.00 40.00 50.00 60.00 70.00
ChamberTemperature
(C)
Time (s)
Chamber Temperature Vs Time
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Calculation
Mass flow rate can be calculated by using formula
[ ( )] (1)
* +
(2)
Equation (1) will be calculated to get the value of Xe. In equation (2), the value of Xeis then used
to calculate Meusing numerically. Speed of sound at nozzle exit is assumed to be 340.29 m/s. It
is not possible to calculate the speed of sound at nozzle exit as the nozzle thermocouple was not
functioning.
(3)To calculate Ve, equation (3) is used.
Thrust produced by the rocket is calculated by using the formula
(4)However, it is assume that the exit pressure is equal to the ambient pressure thus reducing
equation (4) to
or (5)
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Specific impulse is then calculated using the following equation.
(6)
From Table 2, the maximum thrust produced by the engine is stated to be 49.4424 N. This value
is the net thrust after considering the friction between wheel and desk. However, the frictional
force is unknown as it is not stated in the final year project report "Development and Testing of a
Hot Water Rocket". Total firing time was 63 seconds.
The average mass flow rate is = 0.3172 kg/s
The exit velocity is 103.44816 m/s
Therefore, the specific impulse is 10.54 s
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Discussion
The experiment was designed to find out several important parameters that determine the
performance of the rocket engine. From the experiment, generated thrust, chamber temperature
and nozzle temperature are obtained. However, the nozzle temperature set of data obtained is
found to be decreasing once the steam is released while it suppose to be increasing. It is found
later that the thermocouple at the nozzle is not functioning well. The firing time is 63 seconds
and the thrust produced is ranging from 20.4048 N to 49.4424 N.
Figure 6 shows the thrust measurements are constant for certain time, then the thrust
reading suddenly drop and start to be constant again. This cycle is consistent throughout the
firing. Actually this is not the real thrust profile. The thrust should be gradually reduce over time.
This stair-like profile is obtained due to the spring balance do not display the instantaneous
thrust. This is the limitation of the spring balance that hold the reading if the thrust is decreasing
slowly that it appears to be constant. Electronic load cell would be a better choice for thrust
measurement.
Conclusion
All the objectives of the experiment are accomplished. The hot water rocket engine was
successfully fired. Maximum thrust produced by the hot water rocket engine is 49.4424 N,
average mass flow rate is 0.3172 and specific impulse is 10.54 s.
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References
Sutton, G., & Biblarz, O. (2010). Rocket Propulsion Elements 8th Edition.United States of America: John
Wiley and Sons Inc.
Ward, A. T. (2010).Aerospace Propulsion Systems.Singapore: John Wiley & Sons (Asia) Pte Ltd.
Zulfikli, S. (2011). Development and Testing of A Hot Water Rocket.Kuala Lumpur: Department of
Mechanical Engineering, IIUM.