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SP2016_3124774
HOT FIRING OF A N2O/C2H4 PREMIXED GREEN PROPELLANT: FIRST COMBUSTION
TESTS AND RESULTS
Lukas Werling(1), Nikolaos Perakis(2), Steffen Müller(3), Andreas Hauk(4), Helmut Ciezki(5), Stefan
Schlechtriem(6)
(1) Institute of Space Propulsion, German Aerospace Center (DLR), 74239 Hardthausen, Germany, Email:
Lukas.Werling@dlr.de
(2) Institute for Flight Propulsion, Technical University Munich, 85748 Garching, Germany,
Email: nikolaos.perakis@tum.de
(3) Institute of Space Propulsion, German Aerospace Center (DLR), 74239 Hardthausen, Germany, Email:
Steffen.Mueller@dlr.de
(4) Institute of Space Propulsion, German Aerospace Center (DLR), 74239 Hardthausen, Germany, Email:
Andreas.Hauk@dlr.de
(5) Institute of Space Propulsion, German Aerospace Center (DLR), 74239 Hardthausen, Germany, Email:
Helmut.Ciezki@dlr.de
(6) Institute of Space Propulsion, German Aerospace Center (DLR), 74239 Hardthausen, Germany, Email:
Stefan Schlechtriem@dlr.de
KEYWORDS: Green propellants, nitrous oxide
fuel blends, premixed monopropellant, nitrous
oxide, ethene, N2O, C2H4, characteristic velocity
c*, combustion efficiency
ABSTRACT:
Today hydrazine is the commonly used
monopropellant for attitude- and orbit control of
satellites and for powering probes or landers. Due
to changing political and economic framework,
(e.g. the REACH regulation in Europe) different
propellants for replacing hydrazine are currently
under development or qualification. The Institute
of Space Propulsion of the German Aerospace
Center (DLR) at Lampoldshausen is focusing on
two different propellants to replace hydrazine:
ADN-based monopropellants and mixtures of
nitrous oxide with hydrocarbons. The latter are so
called premixed monopropellants: oxidizer and
fuel are stored in a premixed state in one tank.
Thus the simplified propulsion system of a
monopropellant can be combined with the high ISP
of a bipropellant. To gain experience with a
propellant mixture consisting of nitrous oxide
(N2O) and ethene (C2H4), DLR is conducting hot
gas combustion tests with an experimental
combustor. The paper summarizes the results of
combustion tests conducted with the premixed
propellant injected in gaseous state. Calculated
and measured performance (c* and ),
depending on mixture ratio and chamber pressure
is shown and discussed.
1. INTRODUCTION
Since the early days of spaceflight hydrazine is
used as a monopropellant to power rockets,
satellites or probes. During the 50s and 60s of the
20th century a large number of different
propellants were tested to be used as a
monopropellant [1]. Among those, hydrazine
offered a good performance, long term storability,
handling without the danger of explosions and
relatively low costs. Space flight and the operation
of satellites are a business which is strongly
focused on reliability, thus development and
qualifications of new propellants and thrusters
consume lots of time and money. These constrain
make hydrazine the most established
monopropellant to this day.
During the last decade several things changed.
The high toxicity of hydrazine became a point of
concern. In the EU the so-called REACH-
Regulation [2] came into effect, here hydrazine
was set on the list of substances of very high
concern (SVHC). Thus it becomes more and more
likely that the use of hydrazine will be limited or
prohibited in future, even though exceptions for
the space industry might be given [3]. To
compensate a possible prohibition of hydrazine,
across the globe several so-called green
propellants are under development or being
qualified. Among those alternatives, the ADN-
based monopropellant LMP-103S seems to be a
promising candidate. The propellant is currently
under qualification by ESA, two 1N thrusters were
already tested in space on the PRISMA satellite
[4]. Nevertheless other propellants or propellant
mixtures may offer significant advantages and are
therefore under thorough investigations.
The following chapter will give a short overview of
some alternatives. The propellant mixture focused
on in this paper is a mixture of nitrous oxide and
ethen.
1.1. Green propellants overview
Ionic liquids: ADN based propellants:
ADN based monopropellants consist of at least an
energetic salt (ammonium dinitramide) a fuel
component and water. The most known
propellants are LMP-103S, invented by the
Swedish company ECAPS and FLP-106,
developed by FOI (Swedish Defense Research
Agency). LMP-103S offers a 6% higher specific
impulse (253 s) than hydrazine and a 30% higher
density impulse while being less toxic and not
carcinogenic. Storability was already tested for
more than 7 years and the propellant can easily
be ignited by using a preheated catalyst (approx.
350°C) [5]. FLP-106 offers an even slightly higher
specific impulse (about 260 s) coming along with
higher combustion temperatures [6]. The main
difference of FLP-106 in comparison to LMP-103S
is the usage of a less volatile fuel. Detailed
information about ADN based monopropellants
can be found in [6–8].
Ionic Liquids: HAN based propellants
HAN (hydroxyl ammonium nitrate) based
propellants are another class of ionic liquids.
Those propellants were intensively investigated by
the United States Air Force Research Laboratory
(AFRL) to be used as liquid gun propellant. HAN
has been studied since the 1960s, e.g. during the
1980s the liquid gun propellant LP1846 was
developed. Due to the increasing toxicity concerns
of hydrazine, the HAN based propellants were
also considered to be used as a monopropellant
in space applications. Thus the AF-315E
propellant was developed. This propellant was
selected for the Green Propellant infusion mission
(GPIM) [9]. AF-M315E may offer an Isp of 257 s
and a 45% higher density than hydrazine. The
thrusters can be ignited by a preheated catalyst.
According to the higher Isp the combustion
temperature of AF-M315E exceed the combustion
temperatures of conventional hydrazine [10].
Hydrogen Peroxide, H2O2
Hydrogen peroxide is another well studied green
propellant alternative. The main advantages of
H2O2 are a negligible toxicity, easy ignitability via
catalyst and relatively low decomposition
temperatures (up to 1275 K). The drawbacks of
H2O2 are a lower Isp than hydrazine (up to 196s,
depending on concentration) and the
incompatibility with several materials (e.g. copper,
iron, magnesium alloys, titanium) [11–14]. Due to
the availability of high concentrated hydrogen
peroxide, well studied catalysts and a lot of
experience in handling the substance, hydrogen
peroxide seems to be a very promising alternative
for several monopropellant applications; e.g. H2O2
was under consideration for the A5ME upper
stage attitude control system [15].
Water electrolysis propulsion
Another possible - definitely green - alternative is
water electrolysis propulsion. The idea is to have
a satellite equipped with a water tank, two gas
tanks and an electrolyser. By using the
electrolyser the water is decomposed to gaseous
hydrogen and oxygen which are stored in
separate tanks. The satellites thrusters are then
powered by gaseous H2 and O2 at a low mixture
ratio to assure modest combustion temperatures.
Remaining oxygen can be used e.g. in a cold gas
propulsion system. Ignition of the H2/O2 can be
achieved by using a platinum catalyst. Easy
handling of the purified water, no safety and
toxicity concerns and available technology
(electrolysers) are the main advantages.
Drawbacks are a quite complex propulsion system
with three tanks, necessary pressure regulators,
valves, bipropellant injectors and a corresponding
high weight of the whole system. Additional the
avoidance of H2 leakage in a later propulsion
system might be challenging. Airbus D&S is
recently studying this kind of propulsion system
[16, 17].
Mixtures of hydrocarbons with nitrous oxide/
nitrous oxide fuels blends
Another class of green propellants is the so called
nitrous oxide fuel blends, mixtures of
hydrocarbons with nitrous oxide. Those mixtures
are no single species monopropellants as e.g.
hydrazine or hydrogen peroxide. The oxidizer
(N2O) and a fuel (e.g. C2H2, C2H4 or C2H6) or a
fuel combination are stored premixed, i.e.
monopropellant-like in one tank. In comparison to
a classical bipropellant system, only one tank, one
feeding line and one injection system is needed.
Thus these propellants are sometimes called
“premixed monopropellants”, offering a
monopropellant like system while having a
bipropellant performance (Isp approx. 320 s).
Mainly the components are cooled down (< 253 K)
and mixed. The high vapor pressure of the
components can offer a self-pressurizing
propulsion system without any external pressure
supply. Beside the mentioned advantages, nitrous
oxide fuel blends provide some non-minor
challenges: Very high combustion temperatures (>
3000K) require an active cooling of the nozzle and
combustion chamber. Furthermore the propulsion
systems needs proper flashback arrestors as well
as newly designed ignition and injection systems.
The most known nitrous oxide fuel blend is
NOFBX from Firestar [18, 19]. Recently several
explosions occurred during DARPA/Boeing’s work
on a nitrous oxide acetylene propellant mixture
called NA-7 [20]. The German Aerospace Center
(DLR) in Lampoldshausen is working on a
dinitrogen monoxide/ethene mixture [21–24].
1.2. N2O & C2H4 propellant mixture (“HyNOx”)
The German Aerospace Center (DLR) in
Lampoldshausen has chosen a mixture consisting
of dinitrogen monoxide and ethene to be used as
a premixed monopropellant. The mixture was
named “HyNOx” (Hydrocarbons mixed with
nitrous oxide). Ethene was chosen as fuel due to
its not to different vapor pressure compared to
nitrous oxide (vapor pressure at 273 K: C2H4: 41
bar; N2O: 31.2 bar [25]). The similarity concerning
the vapor pressures should assure good
miscibility and simultaneous evaporation in a
propellant tank. Furthermore ethene is quite safe
to handle, so compared to acetylene self-
decomposition hazards can be avoided. Though
the theoretical vacuum specific impulse is lower in
comparison to a mixture of nitrous oxide and
acetylene (approx. 319 s for N2O&C2H4 to 330 s
for N2O&C2H2 [13]).
At the beginning of DLR’s research activities a
cooling and liquefaction setup for the N2O & C2H4
mixture was assembled [26]. After some
preliminary tests it was demounted. In 2016 an
improved system will be set up.
To gain experience with the propellant mixture
and to conduct the first tests, DLR chose to mix
the oxidizer and fuel in their gaseous state
upstream the injector. This offers several
advantages:
a) The mixture ratio can be adjusted easily via
exchange of orifices and/or adjusting the feeding
pressure;
b) Common gas bottles can directly be connected
to the setup;
c) Easy comparability to CFD simulations is
possible; no evaporation effects of the liquefied
propellant have to be considered;
d) The general performance of the propellant for
different mixture ratios can be adjusted easily;
e) If a hard ignition or a flashback across the
injector occurs, due to lower density and
propellant mass in the feeding lines less damage
will be caused;
f) Flashback across the injector is a very critical
issue, the gaseous mixture is easier to ignite, and
thus flashback is more likely. So by using a
gaseous mixture a “worst case” approach is
achieved. If the flashback can be avoided with
gaseous mixtures, it seems to be very likely that it
can also be avoided by using the liquid, cooled
Figure 1: Simplified P&ID of the test setup
propellant;
g) During the later use in a propulsion system 2-
phase or gaseous flow of the propellant cannot be
avoided for all situations. Especially during ignition
or shut down of a thruster, gaseous propellant will
flow through the injector. By using gases, the
occurring pressure drop as well as the ignition and
shutdown behavior at this operation points can be
studied in advance.
All described combustion tests were performed
with gaseous propellants changing the mixture
ratio by using different feeding pressures and a
calibrated set of orifices.
2. TEST SETUP AND COMBUSTOR DESIGN
The combustion tests were conducted at DLRs
M11 test complex. A green propellant test bench
was assembled at the test bench M11.5 [21, 27].
2.1. Test setup
A simplified sketch of the test bench’s fluid system
can be seen in Figure 1.
As previously described, the propellant is mixed in
its gaseous state upstream the injector of the
combustor. Nitrous oxide and ethene are stored in
50l pressure tanks outside of the test container.
The pressure, mixture ratio and the resulting mass
flow are adjusted via pressure regulators and
calibrated orifices. Additionally the test bench is
equipped with H2 and O2 feeding lines to supply
the torch igniter. For flushing of the combustor, to
realize redline shutdown sequences and for
operation of the pneumatic valves nitrogen supply
is needed. Each feeding line is equipped with at
least one pressure transducer upstream the main
valve. At the N2O and C2H4 lines upstream each
orifice a pressure sensor and a thermocouple are
mounted, corresponding measurement data are
used for mass flow calculation. Downstream the
tube junction where the nitrous oxide and the
ethene are mixing, another pressure transducer is
situated. This sensor is delivering the inlet
pressure to the corresponding injector.
2.2. Combustor design
A sectional view of the combustor can be seen in
Figure 2. On the left hand side upstream the
combustor a Plexiglas tube is mounted. Thus a
flashback during a test run can be observed
optically via camera.
The Plexiglas tube is connected to the injection
element of the combustor. These elements are
designed to be exchanged easily, so different
injector types or injector geometries can be
tested. All the tests described in this paper were
conducted with a showerhead injector. The
injector consisted of 17 boreholes with a diameter
of 0.65 mm. Figure 3 shows a photo of the
injector.
Figure 2: Combustor design
Two of the originally foreseen boreholes were
blocked due to a manufacturing failure, so only 17
of 19 injector holes were completely drilled.
Upstream the injector a pressure transducer (P-
INJ) is mounted, thus the pressure drop across
the injector can be derived.
Figure 3: Showerhead Injector
The igniter is a H2/O2 torch igniter commonly used
for different research activities at various test
benches. The igniter is equipped with hydrogen
and oxygen feeding lines in which calibrated
orifices are mounted. The orifices assure an
oxygen/hydrogen mixture ratio of about 1.5 in
case of a sonic flow. Due to the design of the
igniter during all operation modes sonic speed
was reached in the orifices. By using a big excess
of hydrogen relatively low temperatures (1277 K
[13]) in the igniter were achieved. The overall
igniter mass flow at its nominal operation point is
about 2.6 g/s. In the frame of the later described
tests, the igniter’s mass flow was reduced to
about 1.3 g/s. The excess of hydrogen and the
torch’s flame significantly influence the
combustion process of the N2O/C2H4 propellant
mixture during the first 1.5 s of a test run. At the
igniter a pressure sensor (P-ZÜND) and a
thermocouple (T-ZÜND) are installed. During the
ignition sequence the proper function of the ignitor
can be supervised by those sensors.
The HyNOx combustion chamber itself consists of
Elbrodur (CuCr1Zr) segments with different axial
length. The recent configuration’s overall length is
110 mm with a combustion chamber diameter of
24 mm; the first segment is 50 mm, the other two
30 mm long. At the center of each segment three
thermocouples (T-BK-X-Y) and a pressure sensor
(P-BK-X) are installed. The thermocouples are
placed in 3 mm, 8 mm and 12 mm radial distance
from the inner combustion chamber wall.
The combustion chamber is completed by a
nozzle segment. Here CuCr1Zr nozzles with
different throat diameters and expansion ratios
can be used. During the described tests a
truncated nozzle (ε = 1) with a throat diameter of
5mm was used.
3. EXPERIMENTAL PROCEDURE AND TEST
RESULTS
All combustion tests were conducted by using the
gaseous premixed propellant. Nitrous oxide and
ethene were mixed about 0.3 m upstream the
injector at a tube junction. The mixture ratio and
the mass flow rates were controlled by using
calibrated orifices and adjusting the feeding
pressure of the corresponding gases.
3.1. Preliminary tests
Prior to the combustion tests, several sets of
orifices were calibrated. The effective diameters of
the orifices were derived by using an experimental
setup equipped with a Coriolis mass flow meter,
pressure and temperature sensors and assuring
sonic flow through the orifice. Equation (1) and (2)
were used to determine the effective diameter of
the orifice for a chosen pressure. The density
upstream the orifice () and the heat capacity
ratio for a given pressure () and temperature
were taken from the REFPROP database [25].
For each combustion test the diameter was used
to calculate the mass flow rate (equation (3)).
=2
+ 1
+ 1
(1)
=4
2
(2)
=
4
2
(3)
3.2. Test preparation and sequence
In preparation of each combustion test the whole
setup was pressurized and leakage tested, the
pressure transducers and thermocouples were
checked and the test sequence was programmed.
At the beginning of the test sequence, the whole
setup was purged with nitrogen from -8 to -5 s to
assure identical start conditions (see Figure 4).
After nitrogen purging, all valves stayed closed for
additional 5 seconds. The H2/O2 igniter was
lighted at 0 s, N2O and C2H4 valves opened at 0.5
s and the propellant mixture was ignited. The H2
and O2 valves closed at 1.0 s, the propellant
continued to burn for 10 seconds until the main
valves were shut. After closing the N2O and C2H4
propellant lines, all valves stayed closed for
additional 10 seconds, then post test run nitrogen
purging started.
3.3. Measurement data
Figure 4 shows the pressure curves of a test run
with a mass flow of about 12 g/s and an oxidizer
to fuel ratio of about 10.
Figure 4: Pressure data during combustion test
V_1 212
In the diagram, P-C2H4-01 (blue) marks the
ethene pressure upstream the orifice, respectively
P-N2O-01 (magenta) names the nitrous oxide
pressure. P-ZÜND-01 (dark green) is the
pressure in the torch igniter, P-ZUL-01 (light
green) the pressure of the mixture in the
combined N2O/C2H4 feeding line, P-INJ-01
(yellow) the pressure directly upstream the
injector and P-BK-01 names the combustion
chamber pressure.
The diagram (Figure 4) shows the nitrogen
purging at the start of the test run, the ignition at
0s, the pressure rise due to the torch igniter and
the stable combustion regime of the N2O/C2H4
after turning off the igniter.
The propellant valves were closed 11 s after the
torch igniter lit up and 10 seconds after the H2/O2
valves were closed. 10 s after all valves were
closed, nitrogen purging started again (at 22s).
Figure 5: Ignition of N2O/C2H4 mixture, V_1 212
Detail
Figure 5 shows in detail the pressure
development during ignition. The igniter started
working at approx. 0.1 s, raising the pressure in
the igniter to 6 bar. At 0.5 s the N2O/C2H4 mixture
is injected. Due to the igniter and the
corresponding mass flow of the N2O/C2H4 mixture,
the chamber pressure rises to more than 10 bar.
Turning off the igniter is leading to a decrease of
the combustion chamber pressure to the value of
around 8.8 bar. The oscillations of the igniter’s
pressure (green line) are characteristic for the
used torch igniter. They are caused by the H2/O2
combustion in combination with the coaxial
injector of the igniter.
Figure 6: Screenshot of test video
3.4. Mass flow rates and mixture ratio
The mass flow rates, the average chamber
pressure and the oxidizer to fuel ratio of the
conducted tests are shown in Table 1. The test
sequence was conducted as described above.
The mixture ratio and mass flow were derived
from the pressure drop across the orifices. During
test V_1 201 to V_2 209 the N2O and C2H4 valves
stayed open for 5.5 seconds, for the tests V_3
209 to V_1 212 the opening time of the valves,
respectively the N2O/C2H4 combustion time was
increased to 10.5 s.
To ignite the mixture during each test the torch
igniter was used. The total mass flow of H2 and O2
was 2.6 g/s for the tests V_1 201 to V_3 206. For
the following runs (V_1 207 to V_3 207) the
igniter’s mass flow was reduced to 1.95 g/s,
followed by a reduction to 1.3 g/s during startup of
the runs V_1 208 to V_1 212. The H2/O2 mass
flow was reduced to avoid high thermal loads on
the combustion chamber and to keep the
influence of excess hydrogen during the ignition
process as low as possible.
Table 1: Test results, mass flow, mixture ratio and
chamber pressure
Test-No.
Average
mass
flow [g/s]
Mixture
Ratio
(O/F)
Average
chamber
pressure
[bar]
V_1 201
4,35
9,04
3,05
V_2 201
4,33
8,88
3,08
V_1 202
6,56
10,02
4,7
V_2 202
6,59
10,1
4,75
V_1 203
9,55
10,39
7,01
V_2 203
9,68
10,49
7,02
V_1 204
12,07
9,25
8,90
V_2 204
12,16
9,26
8,90
V_1 205
13,6
8,62
10,1
V_2 205
13,8
8,75
10,19
V_1 206
16,64
9,02
12,49
V_3 206
17,07
9,15
12,64
V_1 207
16,91
9,05
12,53
V_2 207
16,96
9,1
12,5
V_3 207
16,39
8,78
12,42
V_1 208
16,59
8,98
12,72
V_1 209
4,46
10,82
3,15
V_2 209
4,44
10,8
3,14
V_3 209
4,46
10,83
3,14
V_1 210
6,76
10,3
4,89
V_2 210
6,75
10,3
4,88
V_1 211
9,54
9,45
7,02
V_2 211
9,61
9,46
7,05
V_1 212
12.06
10,09
8,83
The prefix of the test number (V_1, V_2, and V_3)
indicates a repetition of the test under identical
conditions. Those tests were conducted directly
after each other, only separated by nitrogen
purging. Due to the short time intervals in between
the tests, the capacitively cooled combustion
chamber heats with each test run.
Figure 7:Temperatures during test run V_1 212
This results in high combustion chamber
temperatures at the start of the subsequent tests
(approx. 100°C). Figure 7 shows the temperature
course at several positions of the combustion
chamber and setup. T-BK-01-03 marks the
temperature at the first chamber segment in 3 mm
distance to the inner combustion chamber wall.
T-INJ-01 is the temperature in 3 mm distance
from the hot wall at the faceplate. The sudden
decay at a test time of 5 s is caused by a lost
contact of the thermocouple to the stainless steel
surface.
T-PORÖS marks a thermocouple in the N2O/C2H4
propellant feeding line, directly upstream the
injector. The temperatures of the gases upstream
the orifices are named T-N2O-01 and T-C2H4-01.
A detailed analysis of the occurring heat flux and
the temperature development during the
described tests can be found in [28].
3.5. Flashback challenges
In several tests flashback into the feeding line was
observed. The occurrence of flame propagation
across the injector was recorded via video
camera, focused on the Plexiglas tube [23]. For
those events, a distinct pressure and temperature
rise in the feeding line could be detected. For the
conducted tests immediate propagation of the
flame across the injector was found only if a hard
ignition of the propellant occurred. It is assumed
that a delayed ignition of the mixture leads to an
explosive decomposition. Detailed analysis of the
flashback and ignition events will take place in
future.
4. DISCUSSION
With the results for the chamber pressure, the
mass flow and the corresponding mixture ratio, c*
is derived. The experimental c* is compared to the
theoretical performance calculated with NASA
CEA [13]. Furthermore possible measurement
errors and deviations of the theoretical and
experimental performance will be described.
4.1. Theoretical and experimental c*
Figure 8 shows the absolute c* above the mixture
ratio for the described test runs. The experimental
c* was derived using equation (4). The black
squares mark the experimental c* for the
corresponding test number.
Figure 8: Experimental and theoretical c* for
different chamber pressures and mixture ratios
During the tests, the mixture ratio was varied in
between 8.5 and 11. Different results for c* at the
same mixture ratio are caused by different
combustion chamber pressures. The green
triangle above each experimental value (black
square) indicates the corresponding theoretical c*.
The theoretical c* was calculated via NASA CEA,
by using the reaction model “frozen at throat”.
This results in an equilibrium reaction in the
combustion chamber. The dashed curves indicate
the theoretical c* for a variation of the chamber
pressure (5, 10 and 15 bar). C* rises with
increasing combustion pressure. The increase of
the experimental and theoretical c* with increasing
chamber pressure is caused by an increase of the
combustion temperature (), a decay of the
isentropic coefficient () and small changes in the
specific gas constant () due to changes in the
combustion products composition.
=
4 =
2
+ 1
(4)
Equation (4) shows the relationship of c* and the
mentioned thermodynamic variables. Here is
the combustion chamber pressure, the nozzle
throat diameter and the overall mass flow.
= .
.
(5)
The results for the characteristic velocity (c.)
were compared to the theoretical results. The ratio
c./c. indicates the combustion efficiency
for the chosen chamber, injector and nozzle
geometry (see Equation (5)).
Figure 9: Combustion efficiency for the conducted
test runs
The results for depending on the mixture ratio
are seen in Figure 9. For the test series the
combustion efficiency was in between 86% and
92%.When plotting the absolute c* above the
chamber pressure for the tests runs, the
described increase of c* with rising chamber
pressure can be seen in Figure 10.
Both, the theoretic and the measured c* increase
with rising chamber pressure. Here Figure 10
shows the theoretical characteristic velocity for
reactions frozen at the nozzle throat (green) and
for reactions frozen at the combustor’s end (blue)
assuming an infinite area combustor. The solid
lines are linear fits through the calculated and
measured data. The lines for the theoretical c*
(colored green and blue) show nearly identical
gradients, while the experimental value’s fit seem
to get closer to the theoretical values with
increasing chamber pressure. This effect can also
be seen in Figure 11.
Figure 10: Experimental an theoretical c*
depending on chamber pressure
The fit of the combustion efficiency increases from
about 88% to 91% with a rise of the chamber
pressure from 3 to 12.5 bar.
Figure 11: Combustion efficiency depending on
chamber pressure
The effect of rising combustion efficiency with
increasing combustion chamber pressure is
mentioned in the literature [29]. With higher
chamber pressure the reaction paths shift more to
the equilibrium conditions, which are calculated in
NASA CEA. Another effect additionally increases
the combustion efficiency: with rising chamber
pressure the released energy per volume
increases. The rise of additional energy exceeds
the energy losses via heat conduction at the
chamber walls. So the heat generation due to the
chemical reactions grows stronger than the rise of
heat conduction due to higher chamber pressure.
Thus more energy for acceleration of the exhaust
gases is available and the combustion efficiency
increases.
4.2. Measurement error
To calculate the measurement uncertainty and
derive the error bars, the deviation of each sensor
was calculated or compared to calibrated devices.
After deriving the sensor’s deviation, equation (6)
was used to calculated the uncertainty of the
experimental c*, the pressure and the mixture
ratio.
In equation (6) marks the overall deviation,
depending on the derivative of equation y (in our
case c*) to the single variables x1, x2, etc. (here
, , ).
The maximum error for the coriolis mass flow
sensor (Emerson CMF025) which was used to
calibrate the orifices is 2.4%. This deviation
occurs only when the mass flow is at the far end
of the measurement range. Due to the uncertainty
of the mass flow sensor, the mixture ratio’s
deviation was derived correspondingly.
Furthermore the pressure sensors of the
combustion chamber were pressurized parallel
with a calibration module (Beamex MC5). The
resulting data recorded by the measurement
system was compared to the measured values by
the calibration module. As result of this end to end
comparison a maximum sensor deviation of +/-
0.1 bar was found. The nozzle’s throat diameter
was measured under ambient temperatures
before each test run and in hot conditions after the
test runs. Due to heating of the nozzle segment
the throat diameter varied by 0.08 mm (4.92 mm
up to 5 mm). The calibration of the orifices (see
chapter 3.1) took place some time ago. Thus a
cross check of the mass flow calculation via the
earlier determined orifice diameter and an
additional measurement with the coriolis sensor
took place. This comparison showed deviations of
smaller than 0.5% in between the calculated and
the measured values, thus these errors were
neglected. Furthermore the contraction ratio of the
combustion chamber is quite high (23.41), this
results in low Mach numbers (approx. 0.025). With
this the deviation in between the static and
absolute pressure is approximately 0.3%. Thus for
the c* calculation the experimental measured
static pressures were used.
5. SUMMARY AND OUTLOOK
5.1. Summary
DLR’s institute of Space Propulsion set up a
green propellant test bench to analyze the
combustion and ignition behavior of a so called
premixed green propellant. The propellant mixture
consists of nitrous oxide (N2O) and ethene (C2H4)
and was called “HyNOx” – hydrocarbons with
nitrous oxide. First combustion tests with gaseous
N2O and C2H4 were conducted. During a test
campaign the mass flow and mixture ratio was
varied. The results of the conducted tests are:
a) Pressure and temperature data at
different positions of the combustor were
collected. A set of 24 hot runs with
different mass flow and mixture ratios was
conducted.
b) The combustor is equipped with
thermocouples at different axial and radial
chamber wall positions; the heat fluxes for
the test runs were derived (see [28]).
c) The mixture’s theoretical c* reaches up to
1650 m/s. During the test campaign a
maximum combustion efficiency of 92%
could be observed, reaching a c* of about
1480 m/s.
d) The combustion efficiency and the
absolute c* of the propellant mixture
increases with rising chamber pressure
e) The appropriate design of flashback
arrestors is essential for the use of the
propellant mixture
f) The setup will be modified to conduct
tests with a wider range of mixture and
mass flow ratios. Other injection and
ignition methods as well as flashback
elements will be tested.
5.2. Modifications of the setup
Due to test results, several modifications at the
setup will take place. A wider range of mixture
ratios and mass flow variations will be
investigated. Further new injectors and flashback
arrestors will be tested [30].
To increase the accuracy of the mass flow and
mixture ratio data, the feeding lines of N2O and
C2H4 will be equipped with a coriolis mass flow
meter each. One main benefit of this modification
is that the effective diameter of the orifices does
not have to be derived prior to the tests.
Additionally no sonic speed in the orifices is
needed to increase the accuracy of the measured
data.
Furthermore the test bench will be equipped with
an optimized thrust measurement. So in addition
=
+
+
(6)
to the c* values, the thrust of the combustor can
be derived. With the thrust measurement the Isp
under ambient conditions can be derived. With the
measured Isp the resulting vacuum Isp will be
predicted.
5.3. Testing Flashback arrestors
Critical for a safe use of the premixed propellant
mixture in a combustor is the avoidance of
flashback during start up, stationary operation and
shut down. To analyze and design appropriate
flashback arrestors, a measurement section for
flashback analyses will be set up. Here two
chambers with the propellant mixture in a defined
condition will be separated by a permeable wall of
a porous material. By igniting one of the
chambers, propagation of the flame to the other
chamber or its extinction will be investigated.
Different porous materials will be tested;
furthermore the occurring pressure drop of the
elements will be measured and calculated [31].
5.4. Liquefaction facility
For final tests with liquid propellants, a liquefaction
setup will be reconstructed. The setup will
produce small batches of liquefied N2O and C2H4
to conduct tests with the premixed propellant. The
mixture will be produced with different oxidizer to
fuel ratios. The resulting performance of the liquid
propellant will be compared to the performance of
the gaseous mixture.
6. ACKNOWLEDGEMENTS
The authors would like to thank the M11 test
bench team: Hagen Friedrich, Ingo Dörr, Jan
Buddenberg, and Konstantin Manassis for helping
in conduction and preparation of the tests.
Furthermore the authors would like to thank
Christoph Kirchberger, Mario Kobald and Andreas
Gernoth for reviewing the paper. This work is
funded by DLRs’ “Future Fuels” strategic project.
The support of the other institutes, especially
DLR’s Institute of Combustion Technology in
Stuttgart is greatly acknowledged.
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