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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

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