Mitsubishi Heavy Industries Technical Review Vol. 48 No. 4 (December 2011)44

The “Greening” of Spacecraft Reaction Control Systems

NOBUHIKO TANAKA*1

TETSUYA MATSUO*1KATSUMI FURUKAWA*2 MITSURU NISHIDA*3SHIGENORI SUEMORI*3

AKINORI YASUTAKE*4

Recently, higher performance (i.e., reduction of propellant consumption), operability
improvement/safer handling (i.e., lower toxicity) and cost reduction are required for thrusters used
for the orbital maneuvering and attitude control of spacecraft such as rockets, artificial satellites
and space probes. For this reason, reaction control systems (RCSs) which use lower toxicity
propellants called “Green Propellants” instead of the conventional toxic propellant hydrazine are
attracting attention as next generation fuels. Mitsubishi Heavy Industries, Ltd. (MHI) has
continued the development of monopropellant thrusters that use SHP, a green hydroxylammonium
nitrate (HAN)-based propellant that has the potential to achieve higher performance thanhydrazine. This report describes the results of the evaluation of system advantages and safety(detonability), as well as the firing test for evaluating lifetime.

|1. Introduction
Spacecraft such as rockets, artificial satellites, and space probes have small rocket engines(thrusters) for orbital maneuvering and attitude control. These thrusters produce thrust using thereaction force by exhausting gas. As shown in Figure 1, for supplying propellants to thrusters,tanks, valves and heat control materials are necessary, which are assembled to form the Reaction
Control System (RCS).

Figure 1*1

Reaction control system

Maritime & Space Systems Department, Aerospace Systems*2

Administrator, Maritime & Space Systems Department, Aerospace Systems*3

Nagasaki Research & Development Center, Technology & Innovation Headquarters*4Research Manager, Nagasaki Research & Development Center, Technology & Innovation Headquarters

Mitsubishi Heavy Industries Technical Review Vol. 48 No. 4 (December 2011)45

Hydrazine-based fuel is used in the RCSs of many spacecraft. Generally, bipropellantthrusters, which produce thrust by the combustion of hydrazine-based fuel and an oxidizer such as
dinitrogen tetraoxide, are used when a large amount of velocity increment is required such as inorbit insertion or planetary exploration. On the other hand, monopropellant thrusters, whichproduce thrust by decomposition of the propellant to generate high-temperature gas, are used whenless velocity increment is needed.
For these kinds of thrusters, a higher specific impulse*1 (reduction of propellant
consumption), operability improvement/safer handling (use of lower toxicity propellants) and cost
reduction are preferred. For this reason, RCSs using lower toxicity propellants called “GreenPropellants” instead of toxic hydrazine are attracting attention. Based on the above background,
this report describes the status of our latest research and development on green monopropellant
thrusters using a hydroxylammonium nitrate (HAN)-based propellant.
*1 Specific impulse is a measure of rocket engine efficiency obtained by dividing the thrust by
the propellant mass flow rate.

|2. Comparison of Green Propellants(Physical Properties, Performance and Toxicity)
As shown in Table 1, green propellants developed so far are roughly classified into fourgroups: HAN-based propellant, ammonium dinitramide (ADN)-based propellant, hydrazinium
nitroformate (HNF)-based propellant and hydrogen peroxide. Moreover, HAN-based propellantscan be classified into four subgroups depending on the solvent used. We selected SHP, a mixture ofHAN, ammonium nitrate (AN), water and methanol, as the propellant in our research projects.
From Table 1, it can be seen that SHP has a lower freezing point, higher density and higher specificimpulse, which provides significant advantages in spacecraft applications as described later.
Table 1 Comparison of physical properties and performance among green propellants
Current
propellant
HydrazineFreezing point (°C)
Density (g/cm3)
Theoretical specificimpulse*8 (s)Density specific impulse(g/cm3·s)
Adiabatic flame
temperature (K)21Green propellants*1

HAN-based
ADN-based*2
Hydrogen
HNF-based*3
peroxide
LP1846*4 SHP*5 LTHG*6 HAN/HN-based*7 LMP-103S-100 ≤-30 -35-35-7?
-6
1.4
1.4
1.3
1.4
1.3
1.4
1.4

239

262

276

191

210

255

260

182241

376

396

254

294

332

354

256

1,183

2,171

2,401 1,251

1,455

2,054

2,218

1,154

*1: NH3OHNO3: hydroxylammonium nitrate*2: N2H4(NO2)2: ammonium dinitramide*3: N2H5C(NO2)3: hydrazinium nitroformate*4: Composed of HAN, TEAN and water

Figure 2*5: Composed of HAN, AN, water and methanol*6: Composed of low-temperature HAN/glycine HAN, glycine and water*7: Composed of HAN, HN, TEAN and water*8: Calculated based on the frozen flow assumption, with a combustion
chamber pressure of 0.7 MPa and a nozzle open area ratio of 50

Toxicity assessment of green propellants

Mitsubishi Heavy Industries Technical Review Vol. 48 No. 4 (December 2011)46

Figure 2 shows the toxicity of propellants including hydrazine, with the x-axis representing
the oral median lethal dose (LD50, the dose required to kill half the members of a tested population
after a specified test duration) and the y-axis representing the carcinogenicity based on theevaluation of the International Agency for Research on Cancer (IARC). A smaller x value meansthat the propellant is more acutely toxic, while a larger y value indicates higher carcinogenicity. So
the upper left region in the figure shows toxic propellants and the lower right region shows green
propellants. The green propellants in Table 1 all appear in the lower right area of Figure 2. SHP,
which was used in our research projects, also has low carcinogenicity and low acute toxicity.

|3. Summary of our Latest Research Results, Technical Challenges,
and Applications
Table 1 shows that SHP is a green propellant with excellent potential. For use in thrusters,
SHP has the following advantages over the widely used hydrazine:
・ Lower toxicity=> Cost reduction due to operability improvement and safety
・ Higher performance (higher specific impulse and higher density specific impulse)=> Propellant consumption reduction, Weight reduction due to tank size reduction,and increase of the number of mission devices installed
・ Lower freezing point=> Reduction of power consumption of heaters required to prevent propellant
from freezing
On the other hand, the technical challenges are as follows:・ Improvement of the ignition response・ Controlling the detonability (i.e., difficulty in stable combustion)
・ Improvement of heat resistance of catalyst and catalyst support structures in response toincreased combustion gas temperature
・ Necessity of catalyst bed heating before firing*2*2: No catalyst bed heating before firing is preferred in thrusters applied forrocket attitude control systems.

So far, we clarified the reaction mechanism and selected the suitable catalyst through basic
reactivity tests such as the drop test, and developed an injector by atomization testing through theuse of water as a working fluid. From these tests, we confirmed that the use of iridium-based
catalysts improved the ignition response, combustion without detonation was demonstrated during
firing tests (Figures 3 and 4) and the effect of the catalyst bed/combustion chamber length onperformance was evaluated. Because the fundamental data have thus been obtained, sections 4 to 6
address the quantitative evaluation of system advantages in applying SHP to an RCS used for asmall earth orbit satellite, the evaluation of the safety of the propellant, and system lifetime.
RCSs with HAN-based propellants can be applied to rockets and manned missions, in which
low toxicity is a definite advantage, as well as small satellites with a relatively short operating time,in which a reduction in size and cost is highly important.

Figure 3 Thruster tested in the firing test

Figure 4 Pressure history of the combustion
chamber when using SHP, a HAN-based
propellant

Mitsubishi Heavy Industries Technical Review Vol. 48 No. 4 (December 2011)
47|4. System Advantage
Application

EvaluationofHAN-based

Propellant

As described in Section 3, SHP has the advantages of low toxicity, high performance (highspecific impulse and high-density specific impulse), and low freezing point. In addition, SHP canbe ignited by catalysts, which means no special ignition system for thrusters is required. These
allow the use of the same structural configuration as the RCSs used with comparable hydrazine
monopropellant thrusters.
In this study, an RCS for a small earth orbit satellite is considered. Based on thespecifications of the satellite described as follows, we compared the system weight and power
consumption of an SHP RCS with the corresponding hydrazine RCS.1
・Propellant: hydrazine or SHP・Initial satellite weight: 300 kg
・Velocity increment: 100 m/s
・Thruster: 3 N-class thrusters (4 units)
・Gas supply method: blowdown
・Piping system: See Figure 5

Figure 5 Schematics of the RCS underthe subject

Table 2 shows the results. The SHP RCS has the following advantages compared to thehydrazine RCS:・ Higher performance: Reduction of propellant weight by 2kg
Reduction of the propellant tank weight by 0.5kg
・ Lower freezing point: Reduction of the thermal instrumentation weight by 0.5 kg
Reduction the heater power by 16 W
Overall, the system weight can be reduced by approximately 14% in the SHP RCS systemand power consumption can be reduced by about 33%.
Table 2Item

Specification comparison between hydrazine and HAN-based propellant reaction
control systems
Propellant
Specific impulse (s)Physical properties Density (g/cm3)of propellant
Freezing point (°C)
Density specific impulse(g/cm3·s)
Dry weight of reaction
control system (kg)Weight of reaction Propellant weight (kg)control system Pressurized gas (kg)Wet weight of reaction
control system (kg)For temperature
Quantity of heat maintenance (W)generated by
For heating of catalysts (W)heaters (at peak)
Total (W)

Hydrazine

SHP

215

248

12

1.4
-30

217

356

12.6

11.615.5
0.1

13.5
0.1

28.2

25.2

32

1616164832

Comments
Estimated from the theoretical specificimpulse assuming 90 % efficiency–1 kg (–0.5 kg for tank, –0.5 kg for thermalinstrumentation system)
Due to the improved specific impulse
Approx. –3 kg (–11 %)Reduction of power consumption formaintaining temperature of components
Both systems require catalyst heating
before firing–16 W (–33 %)Mitsubishi Heavy Industries Technical Review Vol. 48 No. 4 (December 2011)48|5. Safety Evaluation (Detonability Test)
We conducted detonability tests using steel tubes1 for evaluating the safety of the propellant.
As shown by the diagram and photographs in Figure 6, a glass test tube filled with propellant wasplaced inside a steel tube that had a No. 6 detonator and RDX explosive at the bottom. The
assembly was buried in the ground and an explosion was then triggered. The fragmented state of
the tested steel tube due to the explosion was an indication of the propellant’s detonability, basedon the following criteria (Figure 7)
・ Detonative: The steel tube was fragmented and high-velocity detonation was considered
to have occurred.
・ Suspicious detonative: The steel tube was deformed to some extent and low velocity
detonation or rapid deflagration was considered to have occurred.
・ Not detonative: The steel tube was not fragmented and no reaction or partial reaction
was considered to have occurred.

Figure 6 Diagram of a test sample and photographs taken
during the detonability tes

Figure 7
Detonability
evaluation criteria

Figure 8 shows the test results. The detonability test was conducted on 19 samples withvarying mixture ratios of SHP (HAN: AN: methanol: water). The theoretical specific impulse for agiven mixture ratio is indicated by the lines in Figure 8(a). The density, indicated by the lines inFigure 8(b), is based on the data obtained. From Figure 8, propellant compositions with non
detonability and at higher specific impulse and higher density than hydrazine were found.

(a) Specific impulse

Figure 8

Detonability test results(b) Density

Mitsubishi Heavy Industries Technical Review Vol. 48 No. 4 (December 2011)49|6. Lifetime Test
To demonstrate the applicability to small earth orbit satellites, a firing test of a thruster wasconducted for evaluating the lifetime aiming at a cumulative operating time of approximately
2,000s using SHP propellant with an iridium-based catalyst. The firing test was conducted underconditions where the catalyst bed temperature should not exceed the heat resistance temperature
based on the space velocity (the ratio of propellant flow rate to catalyst mass) data obtained in pastfiring tests, in order to avoid increasing the catalyst bed temperature by too much and to maintainthe same specific impulse as the hydrazine monopropellant thrusters. Figure 9 shows the results ofthe test. The total cumulative operating time of the thruster was 2,002s. The longest continuous
operating time was 200 s. Under the current thruster design, space velocity must be small from thepoint of the heat resistance temperature of the catalyst, accompanying the lower thrust. (In the caseof larger space velocity, the catalyst bed temperature increases and the specific impulse alsoincreases.) After the firing test, the reduction of catalyst weight, the partial adhesion of the catalyst
in the catalyst bed and dent deformation of the catalyst support structure were observed after
disassembling and checking the tested thruster. However, because the thruster performance after
2,000 s of operation was the same as initial thruster performance, we think that this type of thruster
can be applied to small satellites for which the operating time is about several thousand seconds.
Although an increase of space velocity is needed to achieve higher performance, the consequent
increase in combustion gas temperature may cause catalyst deterioration. Practical applications ofan SHP RCS will require the development of more thermostable catalysts and the design of acatalyst support structure that can withstand temperatures higher than that of the combustion gas.
On the other hand, higher thrust can be obtained by increasing the scale of the thruster (byincreasing the amount of catalyst) without changing the space velocity(i.e., not changing either thecatalyst used, or the design or materials of the catalyst support structures).

Figure 9

History of thruster performance at continuous firing tests|7. Future Plans
The development of green propellant RCSs is currently the subject of increasing researchinterest around the world after a period of stagnation in the 1,990 s. For example, in Sweden, thepractical use of a thruster operating with an ADN-based propellant in space was demonstrated with
the PRISMA satellite in June, 2010. Because SHP has a theoretical specific impulse higher thanADN-based propellants, success in the development of SHP RCSs should provide an advantage inthe development of green propellant RCSs. As shown in Figure 10, our latest results indicate that
SHP monopropellant thrusters can be applied to small satellites with a relatively short cumulative
operating time of about several thousand seconds. In the future, we will accelerate the development

Mitsubishi Heavy Industries Technical Review Vol. 48 No. 4 (December 2011)
50of SHP RCSs and overcome technical challenges such as the development of more thermostable
catalysts and catalyst support structures with greater heat resistance. Then, we will demonstrate thepractical use of SHP RCSs in space to establish the technique of HAN-based thrusters.

Figure 10Future development target|8. Conclusions
MHI has developed monopropellant thrusters and RCSs with SHP that have higher
performance and lower freezing points compared with commonly used hydrazine systems. This
report describes our latest research and development results on green propellant RCSs with SHP.
・ The advantages of the green propellant RCSs were evaluated quantitatively under theassumption of applying green propellant RCS with SHP to small earth orbit satellites.
The results show that the system weight can be reduced by approximately 14% andpower consumption can be reduced by approximately 33% by applying a greenpropellant RCS with SHP, compared with the equivalent hydrazine RCS.
・ For safety verification, a detonability test was conducted using steel tubes. From the test
results, propellant compositions with non detonability and a higher specific impulse andhigher density than hydrazine were found.
・ A firing test was conducted by use of a thruster with SHP and iridium-based catalysts forevaluation of the lifetime of the system, aiming at 2,000 s cumulative operating time. As
a result, 2,002 s operating time and 200 s continuous operating time were achieved toobtain the outlook of applying the green propellant RCSs to small satellites for which
operating time is comparatively short as several thousands seconds.
In the future, we will accelerate the development of green propellant RCSs with SHP toestablish the technology through the development of more thermostable catalysts and theimprovement of the heat resistance of catalyst support structures, aiming for higher performance,
longer life time, and the development of a series of green propellant RCSs.

References1.

Tetsuya Matsuo， Katsumi Furukawa， Takashi Nakamura， Keiichi Hori， Syujiro Sawai， Nobuyuki
Azuma and Hidefumi Shibamoto，Safety Evaluation of HAN-based Liquid Propellant and its Advantages
Applied to a Spacecraft， IAA 50th Anniversary Celebration Symposium in Nagoya – Climate Change /Green Systems (2010)