Performance Test of Mono-Propellant Propulsion System for a Japanese Microsatellite “Hodoyoshi-1”

2013-a-01

Performance Test of Mono-Propellant Propulsion System for a Japanese Microsatellite “Hodoyoshi -1” 2) By Shutaro N ISHIKIZAWA1) Takehiro OHIRA1) Hironori SAHARA1) Naoki M IYASHITA2) and Yusuke K URAMOTO URAMOTO 1)

 Department of Aerospace Aerospace Engineering, Engineering, Tokyo Metropolitan Metropolitan University, Tokyo, Tokyo, Japan 2)  AXELSPACE  AXELSPACE Corporation, Corporation, Tokyo, Japan

Development of under-50kg-class microsatellites is recently increasing and their new utilization has also been proposed  by universities and non-governmental associations, where propulsion system for microsatellites tends to be eagerly desired for their missions with orbit transfer such as formation flight, constellation, and end-of-life de-orbiting. As for propulsion, chemical propulsion is the most suitable to microsatellite with progress in miniaturization because of its high thrust density, short term injection, and easiness to handle. The conventional propulsion for satellite, however, is difficult to handle its  propellant, hydrazine, due to its toxicity and high cost of catalysts, so that universities and companies developing their microsatellites have not installed such propulsion so far. Accordingly, we have been developing a propulsion system for microsatellites based on 60wt% hydrogen peroxide solution because of its little toxicity, low cost, and handling properties compared to the conventional one. Thus, we completed a mono-propellant propulsion system for microsatellites with the  policies of SAFTY FIRST and EFFECTIVE COTS. Now we are planning to demonstrate our propulsion system in a Japanese microsatellite, Hodoyoshi-1, to execute its phase shift in orbit, and it has a mono-propellant thruster with 500mN of thrust and 80 seconds of specific impulse. We have already developed its flight model and carried out its performance test and mechanical environment test such as vibration test for its launch. This propulsion system has blow down system which had a feed system equipped with tanks installing respective bladders pressured by pressurant gas, a set of electromagnetic valves, and other elements and is suitable to adopt the feed system to small volume of microsatellite. In addition, we developed a low-powered electronic substrate for our propulsion system to control its injection and to monitor temperature and pressure. In this paper, we present the results of performance test with flight model, the electronic substrate, and control program. Key Words: Mono-Propellant Propulsion, Hydrogen Peroxide, Microsatellite

considered that chemical propulsion is very suitable for  propulsion system of the microsatellites du e to its high thrust density and little power consumption. Practicable propulsion system for microsatellites, however, does not exist so far while the conventional propulsion is awfully expensive due to its unique and identifiable product characteristics developed  by proper specialists with quite high reliability. Moreover, hydrazine is conventionally treated as the most principal  propellant for chemical propulsion, but the developers of microsatellites are required to handle it under special treatments with particular equipment such as hazmat suit and gas mask because of its high toxicity, so that it is too difficult to handle a hydrazine thruster in the universities and non-governmental associations. Accordingly, we have been developing a propulsion system for microsatellites based on hydrogen peroxide solution because of its little toxicity, low cost, and handling properties compared to the conventional  propulsion system. Thus, we completed a mono-propellant  propulsion system for microsatellites with the policies of SAFTY FIRST and EFFECTIVE COTS. Now we are  planning to demonstrate our propulsion system in a Japanese microsatellite, Hodoyoshi-1, to execute its phase shift in orbit. The propulsion system has a mono-propellant thruster with 500 mN of thrust and 80 seconds of specific impulse. In this  paper, we present the mono-propellant propulsion system and

Nomenclature C 

     

T   M 

k   

: : : : : : :

Mach number Density Ratio of specific heat of mixed gas Boiling point of water Molecular weight of mixed liquid Bulk modulus of mixed liquid void fraction

: : :

mixed flow mixed gas mixed liquid

Subscripts S  G  L

1.

Introduction

The development of under-50kg-class microsatellites is recently increasing and its new usage also has been considered at universities and non-governmental associations. The  propulsion system for microsatellites tends to be eagerly desired for their missions with orbit transfer such as formation flight, constellation, and end-of-life de-orbiting. We

1

result of its efficiency test. 2.

Propellant

We considered that hydrogen peroxide is suitable for  propulsion system of the microsatellites. The hydrogen  peroxide was strenuously researched as propellant in rocket application before 1970, which has high concentration from of 70 to 99wt%. However, it is rare to use the high-concentration hydrogen peroxide after 1970 due to the difficulty to handle and store it. In the case of such a high-concentration hydrogen  peroxide, heat of decomposition is quite larger than that of its vaporization (Fig. 1), so that its decomposition is accelerated to explode storage tank once a little decomposition started. That is why we chose the hydrogen peroxide with concentration of up to 60wt% because its heat of decomposition is lower than that of its vaporization and sufficient stabilizer are contained in it to control its self-decomposition. In addition, there is an advantage that it is easy to purchase the up-to-60wt% hydrogen peroxide as a COTS product.

Fig. 2. Feed system in propulsion

4.

Design of Catalyst Bed

We designed a catalyst bed due to realize a mono-propellant thruster with 500 mN of thrust and 80 seconds of specific impulse. The thruster obtains a thrust by heat of decomposition of 60wt% hydrogen peroxide by a catalyst of  platinum. Thermochemical equation of hydrogen peroxide is followed by  H 2O2

1 2

  

O2

 96140  J 

(1)

To design a catalyst bed, its volume was determined by calculating the mass flow rate, as may be required. In addition, oxygen and water are generated after the decomposition of hydrogen peroxide. So the mixed flow consists of oxygen, water, and hydrogen peroxide without being decomposed. We calculated Mach number of this mixed flow by using the void fraction α of the mixed flow.

Fig. 1. Heat of Decomposition/Vaporization (h2o2.com)

3.

     H 2 O

  

Feed System

The typical feed system in propulsion consists of propellant line and pressuring line. However, it is difficult to adopt the feed system to microsatellite because it has too many elements to install them into small volume of microsatellite. So we introduced blow down system which had feed system equipped with tanks installing respective bladders pressured  by pressurant gas, a set of electromagnetic valves, and other elements as shown in Fig. 2. When electromagnetic valves open, the bladders pressurized by the gas pressure supply hydrogen peroxide to the thruster, then, hydrogen peroxide is decomposed with a catalyst there to generate high-temperature vapor and oxygen gas. There has a high-temperature mixture of vapor, water liquid droplet, and oxygen gas, which is injected from a nozzle to generate thrust. We already evaluated its performance in injection tests on ground and in vacuum, and are planning to conduct the detailed vacuum test and mechanical environment test such as vibration test for its launch. In this way, we succeeded to reduce the number of elements again compared to the typical feed system in  propulsion.

C S 



V G

V G  V  L 

 

(2)

1

 1         S     2 2    LC  L   G C G 

  S 



(1   )   L

C G



C  L



  G

  RGT 

 M 

k 

 

 

 

(3)

(4)

(5)

(6)

   L

We calculated the catalyst bed volume from Mach number of the mixed flow, specific impulse, and thrust obtained. For the outline of the catalyst had already been decided, we decided the catalyst bed from the inner diameter and length of catalyst. The catalyst bed and tanks are shown in Fig. 3 and Fig. 4.

2

Fig. 3. Catalyst Bed and Nozzle

Fig. 6.

A example of results of performance test (duty ratio 80%)

Table 1. The Result of Injection Test of Mono-Propellant

Fig. 4. Tanks and fees system

5.

Duty ratio[%]

Thrust [mN]

Isp[s]

20 40

346±0.01 410±0.02

82.3±3.4 88.8±3.3

60

410±0.02

92.8±3.6

80

467±0.02

89.3±3.0

100

491±0.02

88.2±2.8

Injection Test

We have been conducting the injection tests and measurements of performance of the propulsion system we designed. Propellant tank were pressured at 5 atm with  pressured nitrogen gas to maintain the test condition at a constant. We used the feed system of blow down system to  provide hydrogen peroxide to the thruster as propellant, and measured pressure, temperature and mass flow rate at the respective positions with sampling rate of 10 Hz. An electromagnetic valve was operated by PWM control with the cycle of 200 ms and the variable open duty ratio of 20~100 % (every 20%) to control the mass flow rate of hydrogen  peroxide. The result of mass flow rate is shown in Fig.5. And the example of results of performance test (duty ratio 80%) is shown in Fig. 6. In estimation of thrust performances, we supposed specific heat ratio, C* efficiency and nozzle efficiency as 1.3, 0.9, and 0.8, respectively, as shown in Table 1.

Fig. 5.

6.

Consideration

In the result of injection tests of the mono-propellant  propulsion system, its thrust became large relatively as the duty ratio grew. When the duty ratio is 100%, its thrust was approximately 500 mN. And specific impulse always achieved over 80 sec. These indicate that propulsion performance  becomes the best when the duty ratio becomes the highest in the mono-propellant system. And when the duty ratio is 60%, its mass flow rate is slightly low. This reason we thought is the catalyst is slightly blocked. However it is solved by continuing injection.

7.

Conclusions

It is one of the urgent issues to develop a suitable  propulsion system for microsatellite. We had started to develop such a propulsion system since 2004, and completed it based on the SAFETY FIRST POLICY and EFFECTIVE COTS at once, by the beginning of 2008. We have been improving the propulsion system since then, and we have been using hydrogen peroxide with a concentration of 60wt% for  propellant as ever for safety and handleability. And we introduced blow down system which had a feed system equipped with tanks installing respective bladders. So we succeeded to reduce mass and volume of propulsion system. In the result of injection tests, we obtained thrust of approximately 500 mN with the specific impulse of over 80 sec in the mono-propellant. Our propulsion system and the

Mass flow rate

3

concerning technologies will lead to establish a way to form an on-orbit constellation with plural microsatellites for some earth observation mission in the future, and to produce in a new trend in space utilization with microsatellites.

References 1)

2)

Acknowledgments

This research is granted by the Japan Society for the Promotion of Science (JSPS) through the "Funding Program for World-Leading Innovative R&D on Science and Technology (FIRST Program)," initiated by the Council for

3) 4) 5)

Science and Technology Policy (CSTP).

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Suzuki, N. and Sahara, H.: Generalized Mono-/Bi-Propellant Propulsion System for Microsatellite Based on Non-Toxic Propellant Technology, Proc. 46th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, 2010, AIAA2010-6805. Suzuki, N. and Sahara, H.: Ignition Test of Bi-Propellant Propulsion System Based on Green Propellants for Microsatellite, Proc. 47th  AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit , AIAA, 2011-6805. Sutton, G. P. and Biblarz, O.: Rocket Propulsion Elements, 7th ed. Willey-Interscience, New York, (2000), pp 253. H2O2.com, http://h2o2.com/ [cited 15 July 2011] Shutaro, N., Sahara, H., Naoki, M., and Yusuke K.: Development of Mono-Propellant Propulsion System for A Japanese Microsatellite “Hodoyoshi-1”, Proc. 46th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, 2012, AIAA2012-3757.