Development of a 1 lbf Hydrogen Peroxide Monopropellant Thruster Constructed by Using Composite...

7/26/2019 Development of a 1 lbf Hydrogen Peroxide Monopropellant Thruster Constructed by Using Composite Catalyst Pa

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Development of a 1 lbf Hydrogen Peroxide Monopropellant Thruster Constructed

by Using Composite Catalyst Packing

Hung-Wei Hsu1)

, Wei-Kang Chen2)

, Jian-An Chen1)

, Gung-Bang Chen3)

and Yei-Chin Chao2)

1)Aerospace Science and Technology Research Center, National Cheng Kung University, Tainan, Taiwan2)Department of Aeronautics and Astronautics, National Cheng Kung University, Tainan, Taiwan

3)Research Center for Energy Technology and Strategy, National Cheng Kung University, Tainan, Taiwan

For increasing global environmental and safety concerns, the concept of green, environment-friendly and economical propulsion

technology is becoming one of the most important topics in modern space propulsion development.Hydrogen peroxiderecently re-attractintensive attention and considered as a greenpropellant due to its outstanding features of being non-toxic to human and environment,

relatively safe to store and easy to produce. However, hydrogen peroxide monopropellant thruster suffers from the problem of catalystdurability which significantly affects the thruster performance and lifespan.

In the study, silver is selected as the active catalyst for

hydrogen peroxide decomposition and it is alternately packing with support materials.

By hierarchical deployment of particle size of

support materials, the active area of the catalyst bed can be adjusted and the structural strength of the catalyst bed is enhanced. This

composite catalyst bed configuration is used in the development of a 1lbf-level hydrogen peroxide monopropellant thruster with a showerhead injector. The results of ground and vacuum tests demonstrate that the current catalyst bed design significantly improve the thruster

performance, especially in low-temperature start. The measured static thrust under atmosphere condition is about 0.56 lbf (Isp ~105 s,

with a mass flow rate of 2.44g/s of H2O2) and it is about 0.92 lbf( Isp = ~ 157 s) under 10-4torr vacuum condition for repeated and long

duration tests. This result confirms the reliability of the composite catalyst configuration. In addition, the 1lbfhydrogen peroxide thrusterwith composite catalyst configuration is constructed as a propulsion system onboard a Sounding Rocket as the payload to verify the key

techniques and the high altitude performance of the thruster system. The Sounding Rocket will be launched in May, 2013.

Keywords: Hydrogen Peroxide, Monopropellant, Catalyst, Silver, Composite Catalyst

1. Introduction

Reactive control system (RCS), that uses chemical reactionto achieve the purpose of attitude control and orbit

maintenance/transfer of a spacecraft, is one of the most

important sub-systems onboard a spacecraft that may

determine the success of the mission and the life-span of a

spacecraft. Monopropellant propulsion systems are most

attractive reactive control systems due to their simplicity,

which translates into cost reduction and less complexity

compared with bipropellant systems.High test peroxide (HTP)

or high concentration hydrogen peroxide has a long history in

aerospace propulsion. Research on hydrogen peroxide

propulsion can be dated back to 1930s in Germany for

example by Walter1)

and much more researches on hydrogen

peroxide thrusters / rockets were performed by NASA in the

1960s2,3)

. In recent years hydrogen peroxide, being

re-considered as a green propellant, has become more

attractive as a viable alternative to hydrazine monopropellant

due to its high density, non-toxicity, environment-friendliness,

and ease of handling and storage.

As a monopropellant, it usually takes catalysts to

decompose hydrogen peroxide into oxygen and steam with

high temperature, according to the reaction :

2H2O2(l)2H2O(g)+O2(g)+Heat (1)

It can release a large amount of heat and then translate thermalenergy into kinetic energy by nozzle expansion. Adiabatic

decomposition of pure H2O2 can heat the product gases to

1267K4)

. When the concentration goes below 64 wt%, the heat

release by H2O2decomposition is not enough to evaporate allwater in the product of the H2O2. The adiabatic temperature of

H2O2 decomposition with the concentration above 64 wt%

increases linearly with concentration. In general, H2O2 with

concentration above 85 wt% is defined as the propellant for its

thrust performance.

HTP can be decomposed by various of catalysts, such as

silver2,5,6,7)

, permanganates8)

, manganese oxides (MnO2 and

Mn2O3), platinum, ruthenium dioxide, and lead oxide9)

. The

activity of these catalysts for hydrogen peroxide

decomposition has been classified by Laurence Pirault-Roy10)

.

In his research, it was found that silver has the best activity for

HTP decomposition. However, silver, as active catalyst for

HTP decomposition in monopropellant thrusters, suffers from

its inherent low melting point to survive from decomposition

processes of H2O2 over 92 wt%. Silver screens are commonly

used in the catalyst bed design for high concentration H 2O2.

However, severe sintering can be found in catalyst bed of

silver screen when used in high concentration hydrogen

peroxide or HTP environment. This phenomenon not only

affects the decomposition rate in the catalyst bed but also

reduces the chamber pressure of the monopropellant thruster.

In this study, a new and novel method for silver catalyst

packing for monopropellantthrusters is proposed, developed,

tested and will be demonstrated onboard a Sounding Rocket as

a payload to verify the key techniques and its high altitudeperformance.

2013-a-02


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2. Catalyst Bed

The proposed packing of the silver catalyst bed is

conceptually shown in Figure 1. The silver flickers are subject

to effectively decompose HTP due to its highly active reaction

with hydrogen peroxide, and ceramic materials are used to

separate silver flakes and support flow channels for HTP. Inaddition, ceramic materials are highly heat resistant so that it

can sustain during the decomposition process.

The current design of 1 lb f HTP thruster is based on the

previous experiment of 1N HTP thruster using scaling

principle. In the previous 1N thruster experiment, the mass

ratio of silver flakes and ceramic materials was 1, which could

completely decompose the HTP in the 10mm catalyst bed of a

total weight of 1 g for the mass flow rate of 0.6g/s HTP. Based

on this composition condition, although the HTP can be

decomposed, the pressure build-up is slow during cold start.

Raising the amount of active catalyst materials or increasing

the contact area of the active catalyst materials can overcomethe pressure build-up problem, but that would face the

challenge of sintering. The ignition delay of the thruster

chamber pressure rise can be significantly reduced by

increasing the mass of silver flake, but that would lead to

thruster instability due to catalyst sintering. The mass ratio of

the silver flake catalyst relative to ceramic materials is an

important trade-off factor for pressure build-up and catalyst

sintering. As mentioned the references11)

, when mass ratio is

below 0.33, silver catalyst would not sinter, but reaction

would be very slow for cold start. The previous experiments

show that silver flakes would sinter when the mass ratio is

over 1, and when the mass ratio sets between 0.5 and 1, with

probability the silver catalyst would sinter depending on the

HTP flow rate and operation duration. By past successful

experiment, taking 0.5 for mass ratio and utilizing pulse mode

to preheat the catalyst bed can effectively enhance the

performance and avoid sintering. Figure 2 shows the sintering

and the improving catalyst after experiments.

3. The Thruster and Payload Design

3.1. Design and analysis of the thruster

The study is based on the results of the lab-developed

composite catalyst bed and applied to HTP monopropellant

thruster design. The thrust design criterion for the thrusterunder vacuum is set at 1 lbfwith a chamber pressure of 200

psi. The concentration of HTP is 95 wt% or lower. The fuel

mass flow rate, throat area and the exit area in the expansion

section can then be analyzed during the design of the thruster.

Some assumptions are made for easily estimating the design

parameters and performance, such as steady state, isentropic

flow between catalyst bed and outlet, ideal gases and all

products in gas phase, neglecting friction and boundary effects

before throat, exhaust gases leaving the nozzle in axial

direction, and in chemical equilibrium. According to those

assumptions and conditions, the mass flow rate is estimated to

be 2.68 g/s, the throat diameter be 1.5 mm, and the expansion

angle and area ratio be 15 degree and 44:1, respectively.

Figure 3 shows the main components and assembly of the

thruster. The main body of the HTP thruster is made of

SS316stainless steel. The thruster includes injector, catalyst

reaction chamber and nozzle. There are two major design

considerations factors for the injector. One is for proper

distribution of HTP for smooth HTP decomposition, and the

other is to stabilize the pressure oscillation, if occurs byproviding sufficient pressure impedance. For improving the

pressure impedance, an injector plate and plenum assembly is

used. This injector plate is of shower-head type, which is 17

mm in diameter, 1 mm thickness and 28 holes with 0.4 mm in

diameter each. This injector plate design can uniformly

distribute the HTP in the catalyst bed during tests. For the

catalyst reaction chamber, the chamber diameter is 15 mm and

26mm in length including the injector assembly, catalyst bed,

distributing plate and spring. In the catalyst bed, there are

stainless steel meshes at upstream and downstream locations

to separate from the injector and distributing plate. It is also

used to hold the catalyst in position in the reaction chamber.Downstream of the distributing plate, heat-treated Inconel 718

spring is used to press the catalyst bed in position and to absorb

pressure oscillation in the chamber. On the end of the chamber,

pressure transducer and thermocouple are inserted for

acquiring the chamber data. The nozzle is designed for ground

and vacuum performance tests. In addition, there is a orfice sat

between the thruster control valve and the thruster for

restricting the fuel mass flow rate. This plate is 6.3 mm in

diameter, 2 mm in thickness, with a central orifice of 0.3 mm

in diameter made of aluminum alloy (6061-T6).

3D numerical simulation using Fluent commercial code

with turbulence model is also performed to evaluate the design

parameters. Unconstructed mesh grids are used in the

simulation. Variable inlet pressures at the computational

domain, i.e., at the exit of the thruster valve are tested, and the

results indicate that it required 300 psi at the exit of the

thruster valve to reach the required mass flow rate of 2.65

g/sec. The simulation results agree with to the theoretically

estimated values of the design. Figure 4 shows the numerical

simulation results of the velocity profile in the injector.

3.2. Configuration for sounding rocket flight tests

For further tests of the high-altitude performance of the

thruster, a sounding rocket flight mission is planned for the

HTP monopropellant thruster system as a propulsion payload.Being limited by the total weight of the payload for this

mission, all the mechanism and major components in this

system must be simplified. This system employs nitrogen (N2)

as pressurant to push the HTP into the thruster to decompose

to generate thrust.

This monopropellant thruster system for flight tests is

composed with two major parts. One is propellant feeding

subsystem and the other is the thruster subsystem. In this

section we will focus only on the feeding subsystem and

related components. The feeding subsystem is used to steadily

provide the propellant to the thruster. A piston type of

propellant tank is designed to ensure steady and smooth

operation of propellant supply in low gravity and vacuum

conditions, the tank can accommodate a volume of 437.5 c.c.


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with the pressure of 500 psi. For weight restriction of the

payload and the stability of propellant supply, a lightweight

and high-pressure proof vessel with a regulator is used. The

pressure vessel, made of SS304L by Swagelok, can

accommodate 300 c.c. N2pressurant and can sustain 1800 psi

in pressure. A regulator by Tescom Company is used for

adjusting the operating pressure. It is qualified for pressureproof up to 6000 psi. An isolation valve (latch valve) made of

SS316 is used for operational during launch process. Utilizing

a filter in the pipeline to avoid dust or impurities entrained by

the flow is deemed necessary. In this study we choose the 30

micrometer filter made of SS316 by Swagelok. The pressure

drop across the filter is less than 10 psi. Two quick-connect

work valves made of SS316 by Swagelok are used for filling

N2pressurant and HTP from both ends of the propellant tank.

The response time is important for the RCS thruster control

valve (solenoid valve). In this system, the response time for

the thruster valve is less than 5 micro-seconds and the internal

and external leakage rate is very low.Figure 5. shows the schematic diagram of the current HTP

monopropellant thruster system for flight tests.Figure 6 is the

picture of the thruster system for flight tests.

4. Performance Tests

4.1. Thrust test

The propellant flow rate must be calibrated by adjusting the

propellanttank pressure before the test. It is estimated that in

ideal conditions when the upstream HTP pressure in the

pipeline reaches 245 psi, the chamber pressure can achieve the

design goal of 200 psi with a static thrust of 0.72 lb f at

atmospheric condition. In reality, one has to consider pressure

loss of the pipeline, connecting valves and filters. Also during

the tests, the propellant tank pressure, upstream HTP liquid

pressure, chamber pressure, chamber temperature and thrust

are monitored simultaneously. The test results are shown in

Fig. 7. The results show that whenpropellanttank pressure is

set at 250 psi, the upstream HTP pressure is 225 psi and the

chamber pressure reaches 198 psi with the static thrust of

0.595 lbf at atmospheric condition. The thruster chamber

pressure and thrust reaches the design criteria.

In Fig. 7 two firing tests are shown, the first one on the left

hand side is the cold start test, and the second one on the right

side is the hot start test. Under the cold start condition, theignition delay time is longer than expect and the system takes

about 2 s to reach the final steady thrust performance with an

Isp value of 104 s at atmospheric condition. In comparison,

for the hot start the ignition delay time is less than 150 ms,

pressure rise is quick, the chamber pressure oscillation is less

than 5%, and the Isp is 107 s at atmospheric condition.

Chamber preheating can effectively improve overall

performance and stability of the thruster. Preheat can be

achieved by continuously firing 5 times of 0.1 Hz pulses

before the test. The results are shown in Fig. 8.

As shown in Fig. 8, with 5 preheating pulses, the ignition

delay time is reduced to 130 ms, the chamber pressure is 192

psi, pressure oscillation is less than 5%, static thrust is 0.564

lbf and Isp is 105 s at atmospheric condition. All the values

imply that the thruster reaches the design goal. Reliability of

repeated tests is also performed for 5 times, and the average

results for each test are shown in Fig. 9. Since it is difficult to

set identical propellant tank pressure for each test, the

upstream HTP pressure is seen to vary accordingly on each

test.The mass flow rate of HTP varies between 2.3~2.8 g/s.

The 5 tests show that the chamber pressure is between190~200 psi, pressure oscillation is less than 10%, static thrust

is about 0.55~0.63 lbf and Isp is 95~108 s at atmospheric

condition. After repeated tests for almost 15 min and over

2250 c.c. of HTP consumed, the catalyst is still in good

condition and the thruster performance agrees with

expectation.

4.2. Flight mission requirements test

To accommodate the test sequence of the flight mission of

the Sounding Rocket experiment, the performance tests of the

thruster under atmospheric and vacuum conditions before

launch are also performed. Figure 10. shows the

scheduledRCS propulsion system test sequenceof the sounding rocket

experiment after launch.

Figure 11. shows the results of the test according to the

flight test sequence under atmospheric condition. The test

results for the chamber pressure is 206 psi, static thrust is

0.602 lbfand Isp is 103 s at atmospheric condition.

Figure 12. shows the chamber pressure of the flight test

sequence under vacuum condition. The vacuum test is

performed in in the vacuum chamber of 10-4

torr. Chamber

pressure of 206 psi, vacuum static thrust of 0.92 lbfand Isp of

157 s in vacuum condition are achieved. The hot fire tests

before launch show that the overall performance of the

thruster satisfies mission requirements.

5. Conclusion

Development of a 1 lbfHTP Monopropellant system to be

used onboard the Sounding Rocket for high-altitude flight

tests is conducted in this research. The proposed new and

novel concept of composite silver catalyst bed packing for

hydrogen peroxide monopropellant thruster is the key

technique. By repeated tests, the results confirm the catalyst

bed and thruster design and better and stable thruster

performance is achieved. Through ground and vacuum tests,

the measured static thrust under atmospheric condition isabout 0.56 lbfcorresponding to an Isp of 105 s and the mass

flow rate of 2.44g/s H2O2and it is about 0.92 lbfand Isp of ~

157 s under 10-4

torr vacuum condition. The results meet the

performance requirements and objectives of the mission. The

Sounding Rocket will be launched in May of 2013.

Acknowledgments

This research would like to acknowledge the financial

support of Chung-Shan Institution of Science and Technology,

Taiwan, ROC, through projects, 96-NSPO(B)-SE-FA09-01,

and National Space Organization, Taiwan, ROC, through

projects,NSPO-S-099048.


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References

1) H. Walter: Hydrogen Peroxide Rockets, History of German Guided

Missile Developments, AGARDograph, edited by T. Benecke and A.

W. Quick, Vol. 20, Butterworths, London, 1956.G.

2) Runckel, J. F., Willis, C.M., Salters ,Jr. L. B.: Investigation of

Catalyst Beds for 98-Percent-Concentration Hydrogen Peroxide,

NASA TN D-1808, Langley Research Center, Hampton, Virginia.

1963.

3) Willis, C. M.: The Effect Of Catalyst-Bed Arrangement On Thrust

Buildup And Decay Time For A 90 Percent Hydrogen Peroxide

Control Rocket, NASA TN D-516, 1960.

4) Kuan, C.-K.: Indigenous Technology Development of an Advanced

100mN HTP Monopropellant Microthruster, Masters degree thesis,

National Cheng Kung University, 2006.

5) P. Morlan, P.Wu, A. Nejad, D. Ruttle and F. Fuller: Catalyst

Development for Hydrogen Peroxide Rocket Engines, AIAA paper

1999-2740, 35th AIAA/ASME/SAE/ASEE Joint Propulsion

Conference and Exhibit, Los Angeles, CA, June 20-24, 1999.

6) E. Wernimont, and P. Mullens: Capabilities of Hydrogen Peroxide

Catalyst Beds, AIAA paper 1999-2740, 36th

AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit,

Huntsville, AL, July 16-19, 2000.

7) Ventura, M. and Wernimont, E.: Advancements in High

Concentration Hydrogen Peroxide Catalytic Beds, AIAA Paper 01-

3250, July 2001.

8) A. J. Musker: Highly Stabilized Hydrogen Peroxide as a Rocket

Propellant, AIAA paper 2003-4619, 35th AIAA/ASME/SAE/ASEE

Joint Propulsion Conference and Exhibit, Huntsville, Alabama, July

20-23, 2003.

9) Rusek, J. J.: New Decomposition Catalysts and Characterization

Techniques for Rocket-Grade Hydrogen Peroxide, Journal of

Propulsion and Power, Vol. 12, No. 3, 1996, pp. 574580.

10) Pirault-Roy, L., Kappenstein, C., Guerin, C., Eloirdi, R., and Pillet,

N.: Hydrogen Peroxide Decomposition on Various Supported

Catalysts Effect of Stabilizers, Journal of Propulsion and Power, Vol.18, No. 6, 2002, pp. 12351241.

11) Chan, Y. A.: Development of a HTP Mono-propellant thruster by

Using Composite Silver Catalyst, Masters degree thesis, National

Cheng Kung University, 2006.

Figures

Fig. 1. The concept of composite silver catalyst bed with interaction of

HTP.

Fig. 2. The sintering (before) and the improving catalyst after

experiments.

Fig. 3. The diagram of thruster components and assemblies.

Fig. 4. Numerical simulation results of velocity profile in the injector.

Injector & Inject Plate

& Chamber

Distributing Plate Spring

Nozzle

Connect toPressure Transducer

Connect to

Thermocouple

Thruster


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Fig. 5. The schematic diagram of the HTP monopropellant system for

flight tests.

Fig. 6. The picture of the HTP monopropellant system for flight tests.

Fig. 7. The result of thruster performance test at a cold start and hot start.

The propellant tank pressure set 250 psi and the chamber pressure at 198

psi.

Fig. 8. The diagram of the performance test with 5 preheat pulses.

Fig. 9. The result of thruster performance test for five tests.

Fig. 10. The test sequence of the thruster onboard the Sounding Rocket

for flight test


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Fig. 11. Hot fire test according to flight test sequence in atmospheric

condition.

Fig. 12. The result of final hot fire test in vacuum chamber before

launching.

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