Overview of LE-X Research and Development
By Hideo SUNAKAWA, Akihide KUROSU, Keiichiro NODA1),
Takashi TAMURA, Akira OGAWARA2)
Tsutomu MIZUNO, and Satoshi KOBAYASHI3)
1)
Space Transportation Directorate, Japan Aerospace Exploration Agency, Tsukuba, Japan
2)
Mitsubishi Heavy Industries, Nagoya, Japan
3)
IHI Corporation, Tokyo, Japan
JAXA has begun to study the next generation launch vehicle, which aims for significantly reduced cost and higher
reliability comparable to the human transportation mission. LE-X is the booster liquid rocket engine for the next generation
launch vehicle. The LE-X applies the open expander cycle, which is suitable for the next generation launch vehicle and
human transportation system. The design of the LE-X engine system and component, that is, thrust chamber and FTP, is
progressing and ready for production of the full scale component test. Simulation tools are prepared to mitigate liquid
rocket engine major technical issues, such as combustion stability estimation, regenerative cooling estimation, combustion
chamber life time prediction, and hazard simulation. The full scale thrust chamber and FTP tests are planned to be
conducted in 2012-2013. This full scale tests results will be the first demonstration of the open expander booster engine in
the world.
Key Words:
LE-X, Open expander cycle
transportation system. In our future plan, this vehicle is not only
for the post H-IIA/B, but also will evolve as a launch system for
human spaceflight. LE-X is the booster stage engine of the
vehicle. The LE-X demonstration program is ongoing in JAXA,
and the demonstration test of the full scale component (the thrust
chamber and the fuel turbopump) is planned in 2012-2013. In
this paper, the status of the LE-X program is reported.
1. Introduction
JAXA has initiated research for the Japanese next generation
launch vehicle with a new concept which is reliable, capable,
flexible and low-cost to support our nation's access to space. The
vehicle configurations, such as number of stages and engines,
are under trade study. Figure 1 shows one of the launch vehicle
reference. Figure 2 is JAXA’s future plan for the space
Light
SSO 2.5t
Medium
SSO 6t
Heavy
GTO 4t
Heavy+
GTO 6t
Heavy++
LEO 23t
Human
rated
Common Upper Stage
Fig. 1 Japanese next generation launch vehicle familiy (reference image)
1
2010
11
12
13
14
15
16
17
18
19
20
21
22
23
24
25
26
27
Next Generation Space Transportation System
LNG
R&D Program Propulsion
LE-X Demonstration
H-IIA/B upgrade
Demo Flight
Expanding capability / man rated LV development
Next Generation Launch Vehicle development /operation
Liquid Rocket
H-IIA/B upgrade operation
H-IIA/B upgrade
H-IIA/B opeation
Interplanetary Transportation System Development
TF1
Solid Rocket
Epsilon Rocket
Development
Evolving Epsilon Rocket
Fig. 2 Future plan for space transportation system
Regeneratively
heated fuel
GG
Engine Cycle
Preburner
Gas Generator
Turbine drive gas
Staged Combusition
Closed Expander
Combustion gas
Open Expander
Regeneratively heated fuel gas
Maxmum Thrust
◎
Suitable for booster
stage
◎
Suitable for booster stage
△
Limited
Suitable for upper stage
○
Limited
Capable for booster stage
Isp
△
Low due to discarded
propellant
◎
Highest
○
High
△
Low due to discarded fuel
○
Inherently safe
◎
Inherently safe
Robust to failure event
Safety
Start Sequence
Complexity
△
Contolled safety (Controlling preburner/GG
combustion temperatue is necessary)
○
GG starter is necessary
△
Difficult
◎
Easy
○
Many pipelines and
components
△
Many (high temp.)
pipelines and
components
◎
Simple
Fig. 3
Engine cycle comparison
2. Open expander cycle
(1) Maximum thrust
Compared with closed expander cycle, open expander cycle
can achieve higher combustion pressure and larger thrust
capable for booster stage engine. Presently, there is no expander
cycle booster engine, and the LE-X will become the first booster
engine with expander cycle.
One of the features of the LE-X engine is its engine cycle. The
open expander cycle (expander bleed cycle) will be applied to
the LE-X engine. In open expander cycle, hydrogen pumped by
the fuel turbopump is partly directed to the main combustion
chamber cooling channels and then used to drive the turbines.
The turbine drive hydrogen is injected into the main combustion
flow at nozzle extension. Each pump fed engine cycle is
compared by 5 factors in Fig. 3.
(2) Specific impulse
In the open engine cycle, that is, gas generator cycle and open
expander cycle, specific impulse is less than the closed engine
2
cycle, such as staged combustion cycle adopted in the LE-7A. In
the LE-X and the Japanese next generation launch vehicle
concept study, this disadvantage is recovered by increasing
thrust.
(3) Safety
Gas generator cycle and staged combustion cycle use
combustion energy for turbine drive. Mixture ratio in gas
generator or preburner is the main factor of the turbine thermal
condition. The mixture ratio trouble directly leads to catastrophic
failure and must be controlled. Therefore, the safety of gas
generator cycle and staged combustion cycle is so called
“controlled safety”.
On the other hand, expander cycle uses regenerative energy of
main combustion chamber, and it has “inherent safety” potential.
The comparison of transient simulation in case of the breakage
in turbine upstream duct is shown in Fig. 4. Gas temperature
doesn’t change drastically in open expander cycle. In sight of
human transportation system, open expander cycle engine can
shut down safely and shift to the abort phase.
Fig. 5 LE-5B combusiton chamber in H-II F8
(5) Complexity
There is the sub combustion chamber in gas generator cycle
and staged combustion cycle, therefore, they have more
components and pipeline made of special high temperature
resistant material with special cooling mechanism due to the gas
generator / preburner combusiton gas.
Expander cycle engine has the advantage of less components
and pipelines and relatively low temperature condition. This
leads to low cost and high reliability engine. In addition, since
moisture doesn’t exist at turbine drive hot gases, the expander
cycle can realize easier maintenance by minimizing a hot purge
operation.
LE-X：Expander Bleed Cycle
LE-7A:Staged Combustion Cycle
110 ton EXP BLEED - LEAK AREA 8cm2
LE-7A - LEAK AREA 8cm2
3000
2000
1500
1000
500
Leak
Mixture ratio of Preburner increases
suddenly
 Gas temperature exceeds over material
melting point
Shut Down
0
10
Chamber
TEMPERATURE wall
[K]
temperature
1200
10.5
11
11.5
TIME [sec]
LE-7A - LEAK AREA 8cm2
12
12.5
2500
2000
Gas temperature doesn’t change
1500
 Shut down transition is steady.
Therefore, the LE-X engine cycle is suitable for the Japanese
next generation launch vehicle and human spaceflight. Figure 6
shows the LE-X engine cycle.
1000
500
T_T1F
Leak Shut Down
0
19.5
20
20.5
21
21.5
22
TIME [sec]
110 ton EXP BLEED - LEAK AREA 8cm2
22.5
1200
TEMPERATUREwall
[K]
Chamber
temperature
9.5
TEMPERATURE [K]
TEMPERATURE [K]
Turbine gas
temperature
2500
Turbine gas
temperature
3000
T_T1F
Material Melting
Point
1000
1000
800
800
600
600
400
400
Chamber Wall Temp
200
200
Leak Shut Down
0
9.5
10
10.5
11
11.5
TIME　[sec]
12
12.5
0
19.5
Chamber Wall Temp
Leak Shut Down
20
20.5
21
21.5
TIME [sec]
22
22.5
Fig. 4 Transient simulation result
(4) Start sequence
In case of gas generator cycle and staged combustion cycle,
the start sequence is sensitive to the start up failure. For example,
FTP suction trouble causes the abnormal mixture ratio in gas
generator or preburner, which results in melting the turbine. Gas
generator cycle also needs to be started pyrotechnically or by
pneumatic gas, which may decrease its reliability.
Compared to these, expander cycle has robustness to the start
sequence. In the launch of H-II F8, upper stage engine is open
expander cycle, LE-5B. In the H-II F8, first stage engine was in
trouble, and the LE-5B had to start in severe condition, such as
inadequate chill down. Nonetheless, the LE-5B engine reached
steady state with slightly slower start-up rate. (Fig. 5) This fact
shows the robustness of this cycle.
Fig. 6 LE-X engine cycle
3. Engine System Design
In the expander bleed cycle engine, a main technical challenge
is how to extract turbine power. Turbine drive gas of the
expander bleed cycle is regenerative heated hydrogen, therefore,
the energy of turbine drive gas depends on combustion chamber
heat load, while they are independent in the staged combustion
cycle and the gas generator cycle which use combustion energy.
In order to increase the turbine energy output, the combustion
chamber of the LE-X will be longer than that of the LE-7A.
However, a longer chamber brings heavier engine weight.
Increasing the ratio of turbine drive gas flow rate, which
3
decreases the engine performance, is one way to solve the
problem. A smaller pressure loss though the regenerative coolant
channel is preferred from the view of engine performance, which
might result in excessive high temperature at combustion
chamber wall. As above, in order to determine properly the
system and component specification, it is important to evaluate
the whole engine system.
Therefore, the parametric design method shown in Fig. 8 is
applied to the LE-X engine system and the design parameters of
each component are optimized along with the requirement from
the launch vehicle system.
design parameters
system requirement
baseline configuration
evaluation functions
(Performance/Cost/Reliability)
Assign to orthogonal array
Fig. 7 LE-X engine 3D layout model
Orthogonal array
production
variation
4. Engine component design
Calculating each engine cycle
RSM of each evaluation functions
4.1 Combustion Chamber
The LE-X combustion chamber consists of large size copper
alloy inner liner and steel based super alloy outer shell. To
increase the energy of the turbine drive gas, the combustion
chamber (MCC) is divided into two parts; upper chamber, which
includes chamber throat and injector interface, and lower
chamber. (Fig.9) Pratt & Whitney Rocketdyne (PWR)’s hot
isostatic pressure (HIP) brazing technology will be applied to
fabricate the combustion chamber cooling channel to lower the
production cost significantly. The forming process of inner liner
copper alloy is the flow forming process and shear forming
process. The production demonstrations of these forming
processes are completed.(Fig. 10) After the production
demonstration of the upper chamber HIP brazing, the production
for the LE-X thrust chamber test will be started.
(Performance/Cost/Reliability)
analysis error
development risks
margin visualization
factor effect analysis
Configuration optimization
new engine configuration
Fig. 8 Engine system design process
The principal specification of the LE-X baseline configuration
is shown in Table 1, and the 3D layout model of the LE-X
engine is shown in Fig. 7. The 100% vacuum thrust of the LE-X
engine is set to 1,448kN to maximize the payload transportation
capability. And single stage impeller with 2 stage inducer is
applied to the FTP to achieve high pump head with low cost.
Table 1. LE-X baseline configuration
100% thrust
60% thrust
Combustion time
140 sec
60 sec
(for SSO mission)
Engine thrust (vacuum)
1,448 kN
868 kN
Engine thrust (sea level)
1,217 kN
638 kN
Chamber O/F
6.9
6.7
ISP(vacuum)
432 sec
435 sec
FTP
OTP
2 stage inducer
single stage inducer
single stage impeller
single stage impeller
2 stage impulse turbine
2 stage impulse turbine
Upper
chamber
Lower
chamber
Fig. 9 Main combustion chamber
Fig. 10 MCC inner liner forming demonstration
4
4.2 Injector
Figure 11 is the 3D model of the LE-X injector. Eliminating
the LOX inlet manifold of the injector will reduce the production
cost. The effect of the LOX manifold and main oxidizer valve
was estimated by both flow test and CFD simulation. (Fig. 12)
Fig. 11 Injector
Fig. 14 Fuel turbopump
Fig. 12 Flow test set up and simulation result
The injector design affects combustion efficiency and
combustion stability, which is one of the biggest technical issues
for the LE-X. Especially, combustion instability should be
avoided at the baseline design phase. The combustion instability
estimation is approached by the injector element test and some
instability model using computational analysis. (5.(1)) Single
element firing tests were conducted in Kakuda Space Center (Fig.
13), and some element shapes are selected for the candidate of
the LE-X injector elements. Multi element firing tests will
determine the LE-X injector element and the injector will shift to
production.
Fig. 15 FTP inducer water tunnel test
5. Evolving design analysis
In the LE-X demonstration program, we are trying to estimate
the quantitative reliability of the liquid rocket engine based on
the failure mode abstraction in the early design phase and the
high fidelity computational simulation. Simulation tools are
prepared for the liquid rocket engine major technical issues, such
as combustion stability estimation, regenerative cooling
estimation, combustion chamber life time prediction, and hazard
simulation.
(1) Combustion stability estimation
Our approaches to the combustion instability prediction are
two types. One is the stability analysis of the transfer function
model (Fig. 16), and the other is the pressure oscillations energy
estimation. Both approaches utilize CFD and acoustic analysis.
Subscale combustion tests are planned to verify these
approaches, which will estimate the LE-X full scale test.
Fig. 13 Injector element test at Kakuda Space Center
LOX injection
TF：Ao
mo’
Combustion(o)
TF：exp(-sτo)
mB’Inj
4.3 Fuel turbopump (FTP)
Figure 14 shows the 3D model of the FTP. The LE-X is a
open cycle engine, therefore, the higher turbopump efficiency is
necessary to achieve high engine specific impulse. The inducer
design and test is completed. (Fig. 15) The impeller and the
turbine for the rig test are designed. The turbopoup efficiency
also affects the fuel temperature of injector inlet, hence, the rig
test result will refrect not only to the FTP but also to the injector
design.
Fuel injection
TF：Af
mf’
Combustion(f)
TF：exp(-sτf)
Pc’
Injection-coupled instability
mB’
Combusiton
TF：AB
= n(1-exp(-sτ))
Pc’
Chamber
Acoustic
TF：Rc
Pc’
Burning-coupled instability
Fig. 16 Combusiton instability transfer function
(2) Regenerative cooling analysis
Figure 17 shows the image of regenerative cooling analysis.
Precise estimation of combustion heat flux and cooling channel
heat absorption is important especially in case of the LE-X
5
engine. Subscale combustion tests and the cooling channel tests
are utilized to improve the regenerative cooling analysis.
Fig. 17 Regenerative cooling analysis
P[MPa]
20.0
(3) Combustion chamber life time prediction analysis
Combustion chamber is the most critical component for the
LE-X engine life time. And the life time prediction of the
combustion chamber is important to realize the high reliability
rocket engine. Our life time prediction tool models the coupling
between combustion gas, coolant flow and chamber structure.
The tool are verified by the chamber deformation case we have
experienced in the LE-5B development program. (Fig. 18) This
tool will be applied to the LE-X life time estimation.
T[K]
600.0
0.0
Pressure Dist.
20.0
Temperature Dist.
Fig. 19 Entire engine simulation
7. Future works
Test @ HATS
After the production of the full scale thrust chamber and FTP,
the demonstration tests are planned to be conducted in
2012-2013. Figure 20 is the overview of the thrust chamber test
at Tashiro. This full scale tests will be the first demonstration of
the open expander booster engine in the world.
Test @ Ground
161.0
Throat Diameter [mm]
160.0
159.0
× Measurement
158.0
Analysis
157.0
156.0
Start
Stop
Start
Stop
155.0
0
1000
2000
3000
Time[sec]
4000
5000
6000
Fig. 18 Combustion chamber deformation simulation
(4) End to end engine hazard simulation
Human spaceflight launch system is often discussed about its
safety. PRA (Probabilistic Risk Assessment) is one method to
estimate the risks each failure mode has. However, it is difficult
to develop the failure scenario which leads to the end state, such
as LOC (loss of crew) or LOM (loss of mission). To improve the
risk estimation, we have started numerical hazard simulation of
the LE-X. Numerical simulation in most cases are applied to
each engine component individually. However, the risks cannot
be estimated completely by the individual component analysis
since interaction between components cannot be considered.
One possible way is to perform high fidelity, three-dimensional
CFD analyses of the the “entire” engine system.
As a first step, feasibility simulation study of the full engine
system is carried out. (Fig. 19) This large scale simulation has
been accomplished using the JSS supercomputer of JAXA. The
results of each component test and simulation will be compared
to evaluate the validity and accuracy of this entire engine system
simulation.
Fig. 20 Chamber test overview at Tashiro
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