experiment 2 ramjet

8/10/2019 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.

experiment 2 ramjet

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