47th AIAA Aerospace Sciences Meeting Including The New Horizons Forum and Aerospace Exposition
5 - 8 January 2009, Orlando, Florida
AIAA 2009-452
Numerical Bipropellant Thruster Simulation with Hydrazine
and NTO Reduced Kinetic Reaction Model
Kaori Ohminami1
Intelligent Modeling Laboratory, the University of Tokyo, Tokyo, 113-8656, Japan
Hiroyuki Ogawa2
Institute of Space and Astronautical Science / Japan Aerospace Exploration Agency (ISAS/JAXA), Sagamihara, 2298510, Kanagawa, Japan
and
Kuninori T. Uesugi3
Japan Aerospace Exploration Agency (JAXA), Sagamihara, 229-8510, Kanagawa, Japan
A combustion flow inside a film-cooled bipropellant thruster was simulated numerically.
The film-cooled bipropellant thruster model included a reduced detail kinetic reaction model
of hydrazine fuel and di-nitrogen tetroxide (NTO) oxidizer composed of 61 elementary
reactions. This kinetic reaction model had been originally composed of 245 elementary
reactions and was developed to be available for CFD simulations. Also bipropellant thruster
model included droplet motion and evaporation mechanism so that it could simulate the gas
generation with turbulent flow field. The flow simulation results showed the flame structure
and film-cooled mechanism in detail and could explain the film-cooling dry out mechanism
in accordance with experimental results.
I. Introduction
M
ANY kinds of bipropellant thrusters have been used in space-craft systems for orbital and attitude control.
The bipropellant thrusters have been adopted hypergolic fuels. MMH (monomethyl hydrazine, CH3NHNH2)
is a popular fuel and many bipropellant thrusters have adopted the mixture of MMH, and NTO (di-nitrogen
tetroxide, N2O4) or NTO derivatives as fuel and oxidizer respectively, because its mixture has high ignition
performance and stable combustion.
However a kind of scientific mission, e.g. sampling from the targets requires a no-carbon fuel and oxidizer
mixture: the exhausted chemical products shall not pollute the mission’s targets. Thus hydrazine (N2H4) and NTO
mixture has been required. Hydrazine fuel is more difficult to operate than MMH because the hydrazine generates
higher energy so that the combustion gas temperature is higher than that of MMH. Hence, missions using hydrazine
and NTO mixture have been not many.
ISAS/JAXA has developed a 20N bipropellant thruster (ISAS-20N bipropellant thruster) for attitude and orbital
controls of HAYABUSA (launched in 2003) using hydrazine fuel and NTO oxidizer. The ISAS-20N bipropellant
thruster thermal design is more difficult than that of the MMH/NTO thruster, because higher energy is released. To
avoid overheating of the chamber wall by combustion gases, the ISAS-20N bipropellant thruster has a film-cooling
device. Reducing heat flux to the chamber wall is necessary from the thermal point of view, while efficient
combustion with higher temperature improves the thruster’s performance. And the combustion flow structure in a
thruster is complex and very sensitive to thermal design. However the development of the thruster has been based on
a ‘make and test’ experimental iterative approach up till now and not constructed an analytical model. An effective
approach using CFD (Computation Fluid Dynamics) that is capable of providing more complex knowledge of
1
PhD, Intelligent Modeling Laboratory, the University of Tokyo, Tokyo, 113-8656, Japan, AIAA Member .
Associate Professor, Institute of Space and Astronautical Science (ISAS), AIAA Member.
3
Professor, Japan Aerospace Exploration Agency (JAXA), AIAA Member.
1
American Institute of Aeronautics and Astronautics
2
092407
Copyright © 2009 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.
combustion characteristics is required for next development of a 500N bipropellant ceramic thruster (ISAS-500N
ceramic thruster) in Planet-C mission.
The objective of the present study is to develop an analytical thruster model and to explain the film-cooling
effect and flame structure for the ISAS-500N ceramic thruster. The thruster model is based on the knowledge of
fluid dynamics with chemical kinetics, and is simulated numerically using CFD. In the film-cooled bipropellant
thruster combustion chamber, oxidizer rich areas, fuel rich areas or well mixed areas are unevenly distributed so that
the combustion mechanism should be different locally. Hence the flow analysis should simulate combustion with
ignition delay, gas temperature or gas composition depending on the local condition of chemical gas compositions.
And to cover the variety of chemical composition combustion phenomenon, a detail kinetic reaction mechanism is
required.
In the present paper, to deal with those phenomena in the thruster simulation, we have incorporated the reduced
detail kinetic reaction model into the ISAS-500N ceramic thruster model. The reduced detail kinetic reaction model
had been developed for CFD simulations in previous our work [1]. Also fuel and oxidizer droplets motions and
evaporations are considered so as to take into account the fuel and oxidizer gas generation. The wall temperature is
obtained by the numerical thruster model simulation, which can be compared with the chamber wall temperature of
the development tests. The developed film-cooled bipropellant thruster model is used to investigate the effects of
flow parameters; e.g. mass flow rate, mixture ratio O/F (Oxidizer/Fuel mass flow ratio), flow injection method, on
physical processes such as combustion. Knowledge of these effects can be used to optimize design parameters and
improve the performance of the thruster.
II. General Description
The ISAS-500N ceramic thruster under development is shown in Fig.1. The thruster utilizes liquid hydrazine
(N2H4) as the fuel and liquid NTO as the oxidizer.
The thruster injector has three kinds of inlets. The oxidizer inlets are arranged close to the center of axis and the
fuel inlets surround the oxidizer inlets. The liquid fuel and oxidizer come from these inlets are impacted with each
other and contributes to direct firing of the reactant. Another kind of inlets for film-cooling are arranged outside of
those two. A key feature of the film-cooling is injecting cool fuel towards the chamber wall to protect it from the hot
combustion gas, as shown in Fig.2.
Figure 1. ISAS-500N film-cooled bipropellant ceramic thruster
2
American Institute of Aeronautics and Astronautics
092407
Film-cooling fuel inlet
Film-cooling
Chamber wall
Chamber wall
Injector
Injector
Combustion
Gas
Mixing oxidizer inlet
Mixing fuel inlet
Figure 2. Schematic of film-cooling bipropellant thruster injectors.
III. Modeling and Simulation
The combustion flow structure in the film-cooled bipropellant thruster is fairly complex, so that we model it as
follows:
1) Liquid jets come from each inlets
Fuel and oxidizer injection
impinges with each other or on the
N2H4 liquid jet
combustion chamber wall,
2) The impinged jets form liquid film around
N2O4 liquid jet
the impinging point,
3) The each kinds of fuel and oxidizer
Liquid film formation
droplets are generated from the liquid films,
N2H4 / N2O4
4) The droplets move interacting with the
liquid film
turbulent gas flow, collapsing, and
Secondary break up
evaporating in the hot gas,
5) The hydrazine and NTO gases generated
by each droplets start combusting,
Doplets generation
6) The combusted flow goes downstream with
N2H4 doroplet
turbulence while developing the gas phase
Droplet evaporation
reaction.
N2H4 droplet
N2H4 gas
This model schematic is shown in Fig. 3. To
N2O4 droplet
describe those phenomena, we take into account the
N2O4droplet
physics of:
N2O4 gas
I) Liquid film and, fuel and oxidizer droplet
Gas
phase
reaction
formation,
II) Droplet interaction with turbulent gas flow,
Combustion gas
N2H4 gas
III) Droplet evaporation,
N2 H2O OH
N2O4 gas
IV) Film-cooling droplet interaction with
NO NH3
combustion chamber wall,
V) Gas phase reaction of hydrazine and NTO,
VI) Turbulent flow model.
Figure 3. Modeling schematic of phenomena inside filmThe droplet formation mechanism, the first one, was
cooling bipropellant thruster injectors.
originally proposed by Inamura et al. [2], and was
improved by Yamanishi and Amemiya [3] to adjust
to the ISAS-500N ceramic thruster model. We take it into the ISAS-500N ceramic thruster model to set the initial
condition of the CFD simulations. The items of II to VI are solved in CFD simulation together. A gas phase kinetic
reaction mechanism, the fifth one, was constructed by another scheme in our previous work [1], and is incorporated
into this simulation model. The gas flow (the continuous phase) and the droplets (the discrete phase) motion are
coupled in CFD simulation. To simulate multiphase, the Euler-Lagrange approach is adopted. The coupled
simulation of the continuous phase calculation with the kinetic reaction mechanism and the discrete phase
calculation is concurrently performed by commercial CFD code, Fluent 6.3 [4].
3
American Institute of Aeronautics and Astronautics
092407
A. Governing Equations for the Continuous Gas Phase
For three-dimensional flow calculation, the continuous gas phase is governed by Navier-Stokes equation. The
governing equation is analyzed numerically using a commercial CFD code, Fluent 6.3 [4].In the Fluent Solver, the
pressure-based coupled solver was adopted and realizable k- turbulent mode is used. The simulation was performed
by third-order MUSCL (Monotone Upstream-centered Schemes for Conservation Laws) convection scheme. The
governing equations for the flow can be expressed as
Mass Conservation:
t
+
(
xi
ui ) = S m ,
(1)
Momentum Conservation:
( ui u j )
( ui )
+
=
t
xj
p
xi
ij
xj
+ S M .i ,
(2)
Energy Conservation:
( ui h j )
( h)
=
+
t
xi
eff
xi
T
+
xi
J ki hk + u j
ij
+ Sh ,
(3)
k
Species Transport Equation for kth species:
t
where
ij
( Yk ) +
( u i Yk )
=
xi
xi
J ki + S Yk , k =1, 2,
, K,
(4)
is the viscous stress tensor. The transport equations for all species are solved at each finite volume with
each time step, and Yk is the mass fraction of kth species. The transport properties; the viscosity µ , the effective
thermal conductivity eff and the diffusion flux of kth species J k , are defined by kinetic theory. The gases are
assumed ideal gas. The density is obtained from
=
pM
,
RT
(5)
where M is the mixture molecular weight defined by local gas compositions.
Source terms in Eq. 2 to 4 come from the combustion or discrete phase calculation. The mass source term Sm is
generated by droplets evaporation. The momentum source SM,i is equal to the total droplets momentum loss in a cell.
Sh includes energy generation by combustion and exchange with the droplets evaporating calculation. SYk comes
from chemical species generation or consumption by the combusting calculation, and gas generation of hydrazine
and NTO by the droplets evaporating calculation.
B. Discrete Phase Model
For multiphase treatment of liquid propellant and gaseous combusting flow, we adopt the Euler-Lagrange
approach. The gas phase is treated as a continuum fluid by solving Navier-Stokes equations shown in the previous
section, while the liquid phase is composed of a discrete number of droplets that are traced individually in the
continuous gaseous flow. The discrete phase is solved by tracking a large number of droplets in a Lagragian frame
of reference through the calculated continuous flow field with exchanging momentum, mass and energy.
4
American Institute of Aeronautics and Astronautics
092407
Only two kinds of droplets, hydrazine and NTO are considered. The evaporation of these droplets generates each
gas. We include the hydrazine and NTO liquid as the discrete phases to simulate the gas flame construction source
and the wall cooling effect. The hydrazine and NTO boiling point is under 500K while the flame temperature is over
3000K. The droplets may evaporate quickly, then the gas phase mostly occupies in the combustion chamber. Hence
we assume that the droplets-droplets interactions and the effects of the droplets volume fraction are negligible. Also
collision of the droplets is not considered.
Droplets are initially generated around the film disks that are modeled by Inamura’s scheme [2]. We set the
droplets’ diameter, position, velocity and temperature as initial conditions. Then the droplets move in continuous
fluid phase. The droplets trajectories are solved three-dimensionally by the force balance on the droplet using the
continuous phase cell condition. In this simulation, the force effect on the droplets is the drag force, and drag
coefficient includes dynamic drag model considering the droplets distortion. We also take into account the droplets
secondary breakup effect by TAB (the Taylor analogy breakup) model. And the dispersion of droplets due to
turbulence in the fluid are predicted using the stochastic tracking model (random walk model). While the droplets
lose the momentum by drag force effect, the momentum exchange appears as a momentum sink in the continuous
phase momentum balance in Eq.2 as SM.
The droplets’ heating, evaporation and boiling effect are considered in this simulation. The generated mass,
heat and chemical species of hydrazine (N2H4) and NTO (N2O4) by droplets are appear in the continuous phase
source terms as in Eq.1, 3 and 4.
The discrete phase calculations are performed individually at specified intervals during the fluid phase
calculation. We adjust the discrete phase calculation intervals in the continuous phase iterations.
C. Computational Geometry and Boundary Conditions
The computational model geometry is shown
= +18°
in Fig. 4. The model is three dimensional and has
= -18°
cut nozzle in accordance with the experimental
36°
test model. The model computational domain is
1/10 of full size with angles of
-18° to +18°,
includes 550000 cells, and is composed of solid
domains of a injector and a ceramics chamber
wall, and fluid domains shown in Fig.5.
Thermal conductance of the solid domains is
z=0.077[m]; Th r oa t posit ion
considered. The external combustion chamber
x
wall boundary set as the radiative wall: radiative
z=0[m ]; In ter n a l wa ll bou n da r y of in jector
heat transfers to/from the external environment
z
(environmental temperature is 293.15 K) with an
y
emissivity of 0.88 are considered. The internal
Figure 4. Computational model geometry.
combustion chamber wall boundary set as the
Non-Slip wall for fluid, and the tangential
coefficient of restitution is only considered (the
normal composition is zero) for droplets. The
internal injector wall boundary set as the Non-Slip,
and the elastic collision wall for the droplets. The
external injector wall boundary that is contact
area with the test facilities has constant
temperature with 298.15 K in accordance with the
experimental data.
Droplet sources are arranged around the liquid
film and calculated using Inamura model [2] with
adjustment of hydrazine and NTO unlike
impinging or hydrazine impinging to wall
phenomenon by Yamanishi and Amemiya[3].
Total oxidizer and fuel ratio O/F is 0.77 and filmFigure 5. Computational domains and droplet sources
cooling fuel ratio is 30%.
5
American Institute of Aeronautics and Astronautics
092407
D. Hydrazine and NTO Kinetic Reaction Model
A gas-phase finite-rate chemical reaction model is considered in the flow model. In present study, we adopt the
eddy-dissipation-concept (EDC) model to include the detailed chemical mechanism in the turbulent flow. It assumes
that reaction occurs in small turbulent structures, called the fine scale with fine time scale *. A system of a chemical
reaction is expressed as:
K
k =1
K
' kr X k
k =1
"kr X k
r =1, 2,
, R,
(6)
where 'kr and ' 'kr are forward and reverse stoichiometric coefficients for the kth species in the rth reaction, and Xk is
the molecular concentration of the kth species. The molar production/consumption rate of kth species, k , is expressed
as:
K
&k =
R
r =1
(
kr
kr )k Fr
K
[Xk ]
' kr
k =1
1 Patm
K p , r RT
(
kr
kr
)
k =1
K
[Xk ]
" kr
,
(7)
k =1
where Kp,r is the equilibrium constant for the rth reaction, Patm denotes atmospheric pressure, and kFr is forward rate
constant for the rth reaction using Arrhenius expression:
k Fr = ArT
r
exp[ Ea , r / RT ] .
(8)
In the calculation, combustion at the fine scale is assumed to occur as a constant pressure reactor, with initial
condition taken as the current species and temperature in the cell. Reactions proceed over the fine time scale *,
governed by Eq. 7, and are integrated numerically. Then the kth species total amount of mass fraction change & k M k ,
Mk is the kth species molecular weight, appears in the combustion source term in Eq. 4, and enthalpy change
& k M k hk appears in the combustion source term in Eq. 5.
A kinetic gas phase reaction mechanism as the hydrazine and NTO combustion model is considered. The model
was suggested by Ohminami et al. [1] which includes 61 elementary reactions in N/H/O system.
A bipropellant thruster model using CFD has capability to simulate the complicated flow inside the thruster
combustion chamber; the flow analysis should simulate combustion depending on local O/F with ignition delay,
gas temperature or chemical species compositions. There have been some bipropellant thruster simulations before,
but those numerical models have included only global chemical reaction models or the eddy-dissipation model,
because of no detail hydrazine and NTO combustion model. We had tried to incorporated 2-steps global reaction
model by Sawyer and Glassman [5] into CFD thruster simulations before, but the combustion gas temperature had
become higher than the adiabatic flame temperature: The global reaction model has been found to be not enough.
To construct a reasonable combustion model and to simulate the combustion flame mechanism including filmcooling effects, a detail kinetic model of the fuel/oxidizer combustion is required. Hence we had constructed a
hydrazine and NTO combustion model that includes detail a kinetic mechanism and is useful for bipropellant
thruster simulation.
In previous our work, first we formulated a kinetic reaction model from the available literature. The
experimental results and studies of hydrazine and NTO reactions have not been much reported. No reasonable
combustion model of hydrazine and NTO has been developed, although an N/H/O kinetic reaction model has been
investigated. Thus we constructed the hydrazine and NTO reaction model from the elementary reactions in the
N/H/O system. A total elementary reaction was 245 for 31 chemical species [6].
6
American Institute of Aeronautics and Astronautics
092407
Temperature, K
Second, to incorporate reaction
3500
model in the thruster CFD model, we
extracted the efficient elementary
3000
reactions by sensitivity analysis by
SENKIN included in CHEMKIN-II
2500
package. Then we could construct a
Pc=0.581Mpa, O/F=1.84
hydrazine and NTO kinetic reaction
2000
Pc=0.654Mpa, O/F=1.57
model including 61 elementary reactions
Pc=0.715Mpa, O/F=1.36
for 28 chemical species [1]. This is the
1500
hydrazine and NTO reduced detail
Pc=0.764Mpa, O/F=1.21
kinetic reaction model. In order to
Pc=0.72Mpa, O/F=1.09
validate the model, zero-dimensional
1000
Pc=0.671Mpa, O/F=0.96
combustion simulations performed. In
Pc=0.643Mpa, O/F=0.82
this model, the chemical species
500
0
0.002
0.004
0.006
0.008
0.01
0.012
composition changes in time towards the
Time, sec
equilibrium sate: i.e. the gas temperature
Figure 6. Temperature profiles of zero-dimensional simulation
and the chemical composition after
by the reduced detail kinetic reaction mechanism.
enough long time agree with those in
equilibrium state. The equilibrium gas
temperature
(adiabatic
flame
temperature) is obtained theoretically by
the method based on equilibrium
constants. Therefore we compared the
gas temperature composition obtained by the equilibrium calculation with those after enough long time calculated by
the combustion model with the finite rate constants. Also we calculated ignition delay by the zero-dimensional
simulation using the reduced model. It was coincident with that by original 245 reaction model and experimental
value. The temperature profiles for O/F equals to 0.82 to 1.84 are shown in Fig.6, and the mole fraction profiles are
shown in Fig.7. In the thruster simulation, oxidizer and fuel ratio O/F is 0.77 totally and 1.09 except film-cooling
fuel in accordance with the core flame structure ratio.
0.8
mole fraction
0.7
N2H4
0.6
0.5
0.4
N2
NO2
H2O
0.3
H2
0.2
0.1
0
0.E+00
N2O
H
NO
NH3
1.E-04
2.E-04
3.E-04
4.E-04
OH
5.E-04
time, sec
6.E-04
7.E-04
8.E-04
9.E-04
1.E-03
N2
NO
NO2
NO3
N2O
N2O4
NH
NH2
NH3
NNH
N2H2
N2H3
N2H4
HNO
HONO
HNO3
H
H2
O2
OH
HO2
H2O
H2O2
Figure 7a. Mole fraction profiles of zerodimensional simulation by the reduced detail
kinetic
reaction
mechanism
(Pc=0.72MPa,
O/F=0.77).
1
0.9
0.8
0.7
mole faraction
1
0.9
0.6
N2H4
0.5
NO2
N2
0.4
H2O
0.3
0.2
H2
N2O
0.1
0
0.E+00
H
NH3
1.E-04
2.E-04
3.E-04
4.E-04
5.E-04
time, sec
6.E-04
7.E-04
NO
8.E-04
OH
9.E-04
1.E-03
N2
NO
NO2
NO3
N2O
N2O4
NH
NH2
NH3
NNH
N2H2
N2H3
N2H4
HNO
HONO
HNO3
H
H2
O2
OH
HO2
H2O
H2O2
Figure 7b. Mole fraction profiles of zerodimensional simulation by the reduced detail
kinetic
reaction
mechanism
(Pc=0.72MPa,
O/F=1.09).
IV. Results and Discussions
The gas temperature and chemical species mass fraction distribution are shown in each Fig. 8, and the gas
temperature contours in the x-y cross section (normal direction to the z-axis) are shown in Fig. 9. In Fig. 9, the
combustion gas temperature becomes homogeneous and expanded through the nozzle smoothly. The homogeneous
flame structure about temperature can be explained by the combustion process. In fig. 8, intermediate chemical
species are distributed over the chamber. It is thought if the hydrazine and NTO droplets generate a little of the
hydrazine and NTO gas, the hydrazine and NTO gas decomposes into the intermediate species and they are
distributed over the combustion chamber before ignition. Then the combustion starts with its distributed and mixed
7
American Institute of Aeronautics and Astronautics
092407
intermediate species. It can be said that the intermediate species work as the cause of ignition. Since these ignition
sources, the intermediate species, are mixed well and distributed before the ignition, the homogeneous temperature
flame is formed. In the adopted reduced detail kinetic reaction mechanism, the reaction paths from the reactant
hydrazine and NTO to the intermediate chemical species are constructed before ignition. Then the reaction network
is activated at the ignition delay time and expands by constructing variety of reaction paths within 10 microseconds.
During that ignition, the system gas temperature goes up to over 3000K immediately as shown in Fig.6.
N
N22OO44 dr oplet s
Z, m
0.005
N 2O 4
0.0
0.0
0.2
0.0
0.1
0.2
0.4
2.0
N 2H 4
0.1
NH3
N 2 H 4Ndr2 Hoplet
s
4
0.03
N O2
N 2O
0.06
0.0
0.35
NO
0.0
0.2
0.1
0.7
NH 2
0.0
0.2
NH
0.1
0.0
0.002
0.004
0.07
0.08
0.0
300
0.105
1600
Temper a t u r e [K
[K]]
2900
1e-5
2e-5
Ign itin g
Figure 8. Simulation results of chemical species mass
fraction and temperature distribution.
Figure 9. Temperature contour of x-y
cross-section.
Temperature, K
The prediction and observation results of the external chamber wall temperature along the center of thruster axis
Z are shown in Fig. 10. In the observation temperature curves, the point where the temperature immediately goes up
exists before the nozzle throat convergence.
2000
This point is called a film-cooling dry-out point.
Chamber wall geometry profile
1800
In the prediction curve, the film-cooling dry-out
1600
point appears clearly, too. The film-cooling
1400
Expermental data
dry-out
appearance
agrees
with
the
1200
Model prediction
experimental result well, however, there are the
1000
differences of the dry-out position and the
800
maximum temperature between the prediction
600
and observation.
400
200
Figure 11 shows the heat flux from the
0
gaseous flow and the droplets to the chamber
0
0.02
0.04
0.08
0.1
wall. Before throat area, the droplet phase
Z, m 0.06
absorbs heat from the heated chamber wall,
Figure 10. Chamber wall temperature profiles along zwhile the gas phase reduces heat flux from the
8
American Institute of Aeronautics and Astronautics
092407
7.E+06
heat flux; gas
3.E+06
Mass fraction
Wall
Wall
heat
heat
flux,
flux,
W/m2
W/m2
5.E+06
heat flux; droplet
1.E+06
-1.E+06
-3.E+06
-5.E+06
0
0.02
0.04
0.06
0.08
1
0.9
0.8
0.7
0.6
0.5
0.4
0.3
0.2
0.1
0
N2H4; hydrazine
NH3; ammonia
0
0.1
0.02
0.04
0.06
0.08
0.1
Z, m
Z, m
Figure 12. Mass fraction profiles of hydrazine
and ammonia along z-axis.
Figure 11. Wall heat flux profiles of gas and
droplets phase along z-axis.
hot combustion gas to the wall. From point of view this, it can be said that the film-cooling is composed of the gas
and droplets phase. To see the detail of the film-cooling gas layer, hydrazine and ammonia mass fractions are shown
in Fig. 12. Hydrazine and ammonia has large mass fractions in the film-cooling gas layer. Around the film-cooling
dry-out point: z=0.07, the hydrazine decomposes to ammonia and is disappeared immediately, and wall temperature
increases coincidentally. And the heat flux of the droplets shown in Fig. 11 turns to zero around z=0.07.
Accordingly, it is thought that the film-cooling is composed of both the hydrazine droplets and the hydrazine gas
layer. Disappearing both droplets and gas layer causes film-cooling dry-out suddenly.
It is successfully showed that the reduced detailed hydrazine and NTO combustion model is useful for the filmcooled bipropellant thruster CFD simulation to investigate the film-cooling effect and the combustion mechanism. In
future works, we will try to improve the combustion mechanism or film-cooling model to agree with experimental
data about the film-cooling dry-out point and the maximum wall temperature.
References
1
Ohminami, K., Ogawa, H., A. Hayashi, K. “Construction of Hydrazine and NTO Kinetic Reaction Model for Bipropellant Thruster
Simulation” in Japanese,
The Japan Society for Aeronautical Space Science, Vol. 7, 2008, pp. 1-10.
2
Inamura, T., Oguro, S., Kumakawa, A. and Tamura, H., Institute for Liquid Atomization and Spray Systems, Vol. 15, 2006,
pp. 147-152.
3
Yamanishi, N. and Amemiya, T., private contact.
4
Fluent, Software Package, Ver. 6.3, “User’s manual”.
5
Sawyer, R. F., I. Glassmann, “Gas-Phase Reaction of Hydrazine with Nitrogen Dioxide, Nitric Oxide and Oxygen”, 11th
Symp. Comb., 1967, pp. 861-869.
6
Ohminami, K. and Ogawa, H., “Survey of Reaction Rate Constants in the Reaction System of Hydrazine and NTO” in Japanese, The
Japan Society for Aeronautical Space Science, Vol. 6, 2007, pp. 55-60.
9
American Institute of Aeronautics and Astronautics
092407