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DESIGN AND LAUNCH OF THE HYBRID ROCKET DEMONSTRATOR COMPASS
M. Oechsle, J. Dobusch, M. Gritzka, P. Jochum, F. Merz, R. R¨
odinger
University of Stuttgart, Hybrid Engine Development - HyEnD e.V.
ABSTRACT
Hybrid Engine Development (HyEnD) is the student rock-
etry team of the University of Stuttgart. Since the end of
2019, HyEnD is participating in the educational program
STERN II (Studentische Experimentalraketen) of the Ger-
man Aerospace Center DLR. This program supports students
at German universities to design, build and launch an experi-
mental rocket within a project term of three years. HyEnD is
developing the hybrid sounding rocket N2ORTH, which will
launch from Esrange Space Center in Sweden in November
2022. The goal is to educate students in aerospace related
engineering topics and to surpass the altitude record of the
HEROS rocket (32.3 km) which was launched by HyEnD in
2016. Under the given constraints, the design is optimized
for maximum performance, resulting in a simple compo-
nent setup, an efficient propulsion system and an overall
lightweight design. In a first design study, the decision was
made to use liquid nitrous oxide (N2O) as oxidizer and a
custom developed, HTPB-based fuel in blow down operation.
New technologies and concepts were developed and tested
on the subscale demonstrator rocket Compass, which was
launched in June 2021 to an altitude of 3220 m. Compass
features a newly developed Type V CFRP pressure vessel
with an ETFE surface coating on the inside to ensure chem-
ical compatibility. Opening the compact and lightweight
pyrotechnic slider valve initiates the launch. The HyFIVE
hybrid rocket engine has a nominal thrust of 800 N and fea-
tures a CFRP casing. During the test campaign, the engine
was optimized in 41 hot fire tests, which were conducted at
the DLR Institute of Space Propulsion in Lampoldshausen.
A combined test of oxidizer tank, valves, ground support
equipment and engine was performed as well. The two-stage
recovery system, consisting of a drogue parachute and a main
parachute, was successfully tested in drop tests. During the
Compass launch campaign, HyEnD was able to demonstrate
the operating principle of all components. Although the de-
ployment of the main parachute was not fully successful, a
safe launch, stable flight and data acquisition was achieved.
All newly developed technologies were assessed and, if nec-
essary, optimized for their implementation in the N2ORTH
rocket.
Index Terms—hybrid, rocket, engine, propulsion, ni-
trous oxide, HTPB, launch, type v, tank, pyrotechnic valve
1. INTRODCUTION
The STERN (Studentische Experimentalraketen, engl.: stu-
dent experimental rockets) project of the German Aerospace
Center (DLR) provides funding for student groups of German
universities in order to develop and launch a rocket within
three years. The rockets are launched from the Esrange
launch site near Kiruna, Sweden and shall reach an altitude
of at least 3 km and the speed of sound [1]. However, several
student groups exceeded these minimum requirements [2].
In 2019, HyEnD applied together with the Institute of
Space Systems (IRS) of the University of Stuttgart for the
second time for the DLR STERN project. The application
included a proposal for the development of a successor of
the HEROS sounding rocket, which has reached an altitude
of 32.3 km in November 2016 and was developed within the
first STERN participation of HyEnD [3].
After the application was accepted, the team started work-
ing on the project in October 2019. The goal of HyEnD’s
second STERN participation is to launch a hybrid sounding
rocket called N2ORTH and to recover it safely. Furthermore,
N2ORTH shall surpass the altitude record currently hold by
the HEROS rocket.
The design of the N2ORTH rocket is based on the expe-
rience gained with the HEROS rockets. Several new tech-
nologies are developed and tested within the project. In or-
der to evaluate these technologies and to gain experience, it
was decided to incorporate these technologies into a subscale
demonstrator rocket called Compass. Compass is designed
for a total impulse of up to 8000 N s and shares the work-
ing principle of most components with the larger N2ORTH
rocket. Both rockets are compared in Figure 1.
Fig. 1. Compass and N2ORTH
2nd International Conference on Flight Vehicles, Aerothermodynamics and Re-entry Missions & Engineering (FAR)
19 - 23 June 2022. Heilbronn, Germany
2. DESIGN STUDY RESULTS
A design study for N2ORTH was conducted in order to iden-
tify the relevant technologies that are required to provide a
performance increase when compared to the HEROS rocket.
Multiple propellant combinations and propulsion system con-
figurations were compared. As Compass shares the same pro-
pellant combination and has the same working principle of
components, the results of the design study have a major im-
pact on the overall design of Compass.
2.1. Propulsion System
The DLR STERN project allows the use of rocket engines of
any type, including solid, liquid and hybrid propellant engines
[1]. Liquid rocket engines usually offer high performance and
can be throttled. However, the feeding and injection systems
require a comparably high development effort [4].
Solid propellant engines require no feeding system as the
propellants are already stored in the combustion chamber [4].
However, they offer lower performance than most liquid pro-
pellants. Additionally, the development of solid rocket motors
is subject to legal limitations and restrictions. In Germany,
certain licenses are required for production and handling [5],
making solid propellants unappealing for student groups.
Hybrid rocket engines can offer comparable performance
to liquid propellant engines, depending on the chosen pro-
pellants [3]. Due to the solid state of one of the components
(usually the fuel), the safety of all steps during production and
operation is increased [6]. Because of the less complex de-
sign, pressure-fed hybrid rocket engines are commonly used
by students and amateurs worldwide [7, 8, 9]. As HyEnD has
already successfully developed and tested several hybrid mo-
tors, the decision was made to develop a hybrid rocket within
the project.
Typical oxidizers used in hybrid rockets are liquid oxygen
(LOx, O2), nitrous oxide (N2O) and concentrated hydrogen
peroxide (HTP, H2O2) [4]. A comparison of relevant param-
eters can be found in Table 1.
Table 1. Considered Oxidizers
ρ[10] OFopt*ISP,SL*
O21285 kg m−3(at −214 ◦C) 2.2 260 s
N2O 785 kg m−3(at 20 ◦C) 6.7 231 s
H2O21452 kg m−3(at 20 ◦C) 5.7 246 s
*Calculated with NASA CEA [11]. Fuel: HTPB.
Flow frozen at throat, expansion from 30 to 1 atm.
While liquid oxygen delivers the highest specific impulse
(ISP,S L = 260 s) of the considered oxidizers, hydrogen per-
oxide has the highest density (1452 kg m−3). However, liquid
oxygen is cryogenic, requires expensive equipment and la-
borious handling [4]. H2O2is challenging to handle due to
the risk of decomposition and also the most expensive option.
Nitrous oxide offers the lowest performance (ISP,SL = 231 s)
and density (785 kg m−3) of the considered options, but is
affordable from an overall system perspective and often used
in student-built engines [3, 7, 12]. While the handling does
not require cryogenic temperatures, material compatibility
still has to be considered. Self-decomposition can result from
handling at elevated temperatures, pressures and impurities
[13]. Due to the high vapor pressure (pvap ≈51 bar at 20 ◦C
[10]), self-pressurized propulsion systems without a pressure
regulator and additional pressurant tank can be used.
Depending on the size of the rocket, the resulting weight
reduction of a self-pressurized system can compensate the
performance losses due to the use of nitrous oxide. Addi-
tionally, the high optimum mixture ratio (OFopt = 6.7with
HTPB) results in a larger rocket tank. While the density of
nitrous oxide is lower than the density of the considered fuels
(see Table 2), the oxidizer can usually be stored more effi-
ciently in a pressure vessel than the solid fuel can be stored
in the combustion chamber. The low volumetric loading frac-
tion is inherently caused by the design of a hybrid combus-
tion chamber, as the fuel grain requires one or multiple ports.
Although a pressure-fed combustion chamber has to handle
lower pressures than the tank, additional thermal protection is
required, thereby increasing the dry mass.
It was concluded that the use of nitrous oxide in blow-
down configuration is most suitable to achieve the highest
rocket performance within the project constraints. Helium
was chosen as an additional pressurant stored in the ullage
of the oxidizer tank. Supercharging with inert gases like he-
lium increases the lift-off thrust, ensures a sufficient pressure
difference over the injector and reduces effects of cavitation
within the feeding system and injector [12]. Additionally,
the risk of flame flashbacks during the gas phase blow down
is reduced.
Table 2. Considered Fuels
ρ OFopt*ISP,SL*
Paraffin 900 kg m−3[14] 7.4 234 s
HTPB/MDI 930 kg m−3[15] 6.7 231 s
Sorbitol 1490 kg m−3[14] 2.7 219 s
HDPE 947 kg m−3[16] 7.3 234 s
*Calculated with NASA CEA [11]. Oxidizer: N2O.
Flow frozen at throat, expansion from 30 to 1 atm.
As shown in Table 2, all considered fuels deliver compara-
ble specific impulse. Oxygenated fuels like sorbitol lead to
slightly reduced performance and to a lower optimum mix-
ture ratio. Due to the minor differences in performance, other
criteria are used for the selection of the fuel. Paraffin has
higher regression rates compared to the other fuels [17]. This
is important for hybrid rocket engines with high thrust lev-
2nd International Conference on Flight Vehicles, Aerothermodynamics and Re-entry Missions & Engineering (FAR)
19 - 23 June 2022. Heilbronn, Germany
els as they are considered for first stage applications and or-
bital vehicles. However, a high regression rate results in a
higher required combustion chamber diameter. It was con-
cluded that a rocket diameter higher than 250 mm is not feasi-
ble for N2ORTH due to the limited availability of appropriate
equipment and tools. Thus, the maximum possible fuel mass
at optimum mixture ratio is limited by the engine diameter.
Fuels with lower regression rates generally lead to longer fuel
grains with smaller diameter, thereby allowing a longer burn
time at a given fuel thickness. Fuel grains from HDPE either
have to be machined from a semi-finished product or brought
into the appropriate shape by plastic injection molding. Ad-
ditionally, HDPE leads to comparatively high L/D ratios of
the engine due to its low regression rate [18]. Therefore, it
was not regarded suitable for the project. The use of a casting
process with an HTPB-based fuel allows a less constrained
shaping of the fuel grain. While HTPB offers a performance
increase of up to 4.5 % when compared to pure sorbitol, one
common disadvantage is the use of toxic curing agents (MDI,
TDI) [19] during the fuel grain production process. In-house
development and testing of a HTPB-based fuel using a non-
toxic curing mechanism started prior to the STERN project.
As first results were promising, it was decided to fully develop
this new fuel within the STERN project.
2.2. Key Technologies
In order to improve the performance of N2ORTH when com-
pared to HEROS, the most relevant parameter is the structural
mass ratio. The key technologies to be developed and tested
with Compass were identified as:
• A type V CFRP liner-less pressure vessel that is inte-
grated into the outer hull structure
• A hybrid engine with a lightweight CFRP combustion
chamber hull
• Development of a non-toxic HTPB-based fuel, castable
at room temperature
• In-house developed pyrotechnical valves for oxidizer
supply and release
3. COMPASS ROCKET
The Compass rocket has a length of 2450 mm and an outer di-
ameter of 122 mm (see Figure 2). The oxidizer tank can store
up to 5 L of nitrous oxide and is filled via a quick disconnect
coupling. Helium is used to pressurize the oxidizer up to a
pressure of 65 bar, which is set by the relief valve. The rocket
also features an avionics section, a two-stage recovery system
and the pyrotechnically actuated main and release valve. It is
propelled by the HyFIVE hybrid rocket engine with a nominal
thrust of 800 N. The dry mass of Compass is 10 kg.
Fig. 2. Compass (CAD Model)
3.1. Oxidizer Tank
The oxidizer tank of Compass is a composite pressure vessel,
which features a thin ETFE fluoropolymer surface coating on
the inside. This design was chosen primarily due to the po-
tential weight savings for the N2ORTH rocket compared to
HEROS. The production process as well as the application in
the rocket was validated with Compass.
In the production process, two aluminum tank domes are
coated with an ETFE fluoropolymer on the inner surface. The
cylindrical part of the pressure vessel uses a thin aluminum
tube as a winding mandrel, which is coated with an ETFE
layer and an adhesion promoter on the outside. The ETFE
coating ensures leak tightness as well as nitrous oxide com-
patibility [13]. The adhesion promoter improves the bond-
ing to the CFRP overwrap. Additional ETFE is used to weld
together the tank domes and the ETFE-coated mandrel. Af-
ter the CFRP overwrap is applied, the cylindrical aluminum
winding mandrel is dissolved with a sodium hydroxide solu-
tion. During this process, the sodium hydroxide solution must
be cooled continuously due to the exothermic reaction. The
laminate is heat cured at a temperature of 65 ◦Cfor 24 h. An
overview of the tank design is shown in Figure 3.
The structural calculation was conducted using classical
laminate theory in combination with Puck’s action plane fail-
ure criterion [20]. The tank was designed for an operating
pressure of 65 bar with a safety factor of two. The design
was validated by destructive testing with tank prototypes. The
obtained failure pressures were in good accordance with the
calculated values with a deviation of less than 4 %. An empty
mass of 1050 gwas achieved for the flight version of the 5 L
oxidizer tank.
2nd International Conference on Flight Vehicles, Aerothermodynamics and Re-entry Missions & Engineering (FAR)
19 - 23 June 2022. Heilbronn, Germany
Fig. 3. Compass Oxidizer Tank
3.2. Engine
The HyFIVE engine was developed for the use in the Com-
pass rocket, but also to determine the fuel’s regression rate
and general properties. In total, 41 hot fire tests with differ-
ent configurations were conducted. The modular test engine
HyFIVE-1 was designed to have a short turnaround time to
enable a high test cadence required for evaluation of differ-
ent fuel compositions. With HyFIVE-2 the focus was shifted
towards optimization of the engine regarding combustion ef-
ficiency and insulation component design. The thrust was in-
creased from the 500 N range (depending on the fuel and in-
jector configuration) to 800 N. HyFIVE-3 was designed with
a CFRP casing, reducing the overall engine mass.
3.2.1. Fuel Development
A non-toxic HTPB-based fuel was developed. The curing
mechanism is based on the reaction between hydroxyl termi-
nated polybutadiene (HTPB) and maleic anhydride grafted
polybutadiene, thereby forming an ester linkage and a car-
boxylic acid group (half ester). Higher anhydride grafting
concentrations in the polybutadiene backbone may lead to
incompatibilities between the HTPB and the maleinized
polybutadiene, mainly caused by the increased polarity of
the latter. This issue is solved by introducing an epoxidized
HTPB as a third component, whose epoxy groups may react
with the carboxylic acid groups formed in the first occuring
anhydride-hydroxyl reaction. The fuel can be stored at 0◦C
for several weeks after mixing while remaining castable. Cur-
ing is initiated by heating the mixture to 65-75 ◦Cfor at least
12 h. The fuel has a density of 930 kg m−3. The hardness can
be varied by the molecular weight of the HTPB and the graft-
ing grade of the maleinized polybutadiene. The regression
rate of the fuel was evaluated in a test campaign by measuring
the oxidizer mass flow and the mass of the fuel grain before
and after the test. The following correlation between the fuel
regression rate (in mm s−1) and the oxidizer mass flux GOx
(mass flow per port cross section) was obtained:
˙r= 0.085 ·GOx
0.5(1)
valid for 125 kg
s m2< GOx <350 kg
s m2
Problems with insufficient adhesion to phenolic liners were
solved by coating of the latter by an epoxy formulation with
excess amine fraction. The formulation bonds physically
to the sanded phenolic liner surface, whereby the unre-
acted amine groups react with the anhydride groups of the
maleinized polybutadiene, ensuring a chemical bond between
fuel grain and liner. This method of adhesion promotion is
also feasible for isocyanate cured HTPB standard formula-
tions.
3.2.2. Combustion Chamber Design
A cutaway view of the flight version of the engine is shown in
Figure 4. It uses a 320 mm long fuel grain with an initial port
diameter of 35 mm. A doublet impingement injector made
off brass is fitted between the aluminum bulkhead and the
connection element to the rocket’s structure. All insulation
components are made from paper phenolic composites, with
the only exception being the graphite throat. A vortex ele-
ment made from silica phenolic is placed in the post chamber
to increase the reaction efficiency. Relevant parameters of the
flight engine are listed in Table 3. The operation of the engine
is started by igniting an asymmetrically mounted commercial
of the shelf (COTS) pyrotechnical flare charge 1.5 s before
opening the oxidizer supply.
Fig. 4. HyFIVE-3 Engine Design (CAD Model)
Table 3. HyFIVE-3 Flight Engine Parameters
Parameter Value
Nominal thrust 800 N
Nominal chamber pressure 25 bar
Operation time up to 10 s
Nozzle expansion ratio 1:4
Mixing ratio (OF ) 6-7
Fuel mass 685 g
Dry mass 1488 g
2nd International Conference on Flight Vehicles, Aerothermodynamics and Re-entry Missions & Engineering (FAR)
19 - 23 June 2022. Heilbronn, Germany
In March 2021, the flight version of the engine was tested at
the test bench M11.5 of the DLR Institute of Space Propulsion
in Lampoldshausen [21] for five seconds. The temperature,
pressure and mass flow of the oxidizer as well as the com-
bustion chamber pressure and thrust was measured. Based on
this data, a specific impulse of ISP = 211 s and combustion
efficiency of ηc∗= 97% was obtained. Figure 5 shows the cut
engine after the test. All insulation components, the fuel grain
and the vortex element were observed to be in good condition.
Fig. 5. HyFIVE-3 Engine Opened after Hot Fire Test
3.2.3. Blow-Down Operation
The engine was tested together with the rocket’s oxidizer
tank, valves and ground support equipment in a blow-down
test in May 2021. Figure 6 shows the overall setup with the
oxidizer tank, fluid system and engine. The tank was filled
with 2.1 kg of nitrous oxide at a temperature of 18 ◦Cand
super charged with helium to a pressure of 67 bar.
(a) Fluid System Set-Up (b) Engine Hot Fire
Fig. 6. Compass Blow Down Test
Figure 7 shows the measured tank and combustion chamber
(CC) pressure. The injector design was identical to the design
used for the previous static tests. Due to the blow down op-
eration and resulting decrease of oxidizer tank pressure, the
average thrust is reduced to 500 Nand the nominal operating
point (800 N) is reached shortly after ignition. In theory, the
thrust profile can be adjusted by a variation of super charging
pressure, injector inlet area and the oxidizer temperature.
0 2 4 6 8 10 12 14 16
0
20
40
60
Time / s
Pressure / bar
Tank
CC
Fig. 7. Pressure Data of Blow Down Test
3.3. Valves
A schematic of the fluid system of Compass is shown in Fig-
ure 8. It is designed for a pressure of 65 bar with a safety
factor of two. The three main design drivers are weight, com-
pactness and minimizing total pressure loss of the flow. The
main and release valves were developed in house. For the
quick disconnect, the check valve and the relief valve, ade-
quate COTS components are available.
Fig. 8. Schematic of the Compass Fluid System
The main valve (see Figure 9) consists of a slider with a
trough hole, pushed into the opened position by two pistons.
To ensure a simultaneous actuation of both pistons, a single
nitrocellulose pyrocharge is used. The pistons allow for a
complete separation of the pyrocharge and the oxidizer. In
total, the Compass main valve has a mass of 230 g. For
N2ORTH, the slider results in a compact system and elimi-
nates the need for additional piping. This advantage is less
significant for Compass due to other design constraints.
Fig. 9. Main Valve
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19 - 23 June 2022. Heilbronn, Germany
For the release valve, a simple and light-weight design was
developed (see Figure 10). The valve is screwed into a
manifold with a membrane sealing the oxidizer inlet. Upon
activation, the membrane is punctured by a pyrotechnically
actuated hollow nail, releasing the oxidizer through the out-
lets. The reaction products of the pyrocharge are separated
from the oxidizer side by two O-rings. The design results in
a weight of 35 g.
Fig. 10. Release Valve (CAD Model)
3.4. Recovery
The recovery system is positioned below the nose cone sec-
tion. The parachutes are ejected laterally to the rocket’s body
axis (see Figure 11). The two-stage system includes a mortar-
like ejection mechanism of the drogue parachute. The main
parachute (packed in a deployment bag), a three-ring release
system and the parachute’s harness are stored in a CFRP tub
underneath the mortar. At the bottom end of the section, a
ring-shaped mount is used to connect the harness. To reduce
the parachute’s opening forces, a polyamid shock absorber is
used. The ejection and release systems are pyrotechnically
triggered.
Fig. 11. Components of the Two-Stage Recovery System
The Compass recovery system resembles a smaller version of
the system designed for N2ORTH, resulting in a similar se-
quence of the descent phase (shown in Figure 12). At apogee
the drogue parachute is ejected, reducing the descent veloc-
ity to 30 m s−1. At an altitude of 400 m the three-ring release
system is triggered, causing the drogue parachute to pull the
main parachute out of the recovery tub. The main parachute
decelerates the rocket to a terminal velocity of 5 m s−1to en-
sure a safe and damage-free landing.
Fig. 12. Sequence of the Descent Phase
3.5. Avionics
The avionics are split up into modules enabling a parallel de-
velopment of multiple functions. The individual nodes are
connected by a CAN bus, allowing the measurements (see Ta-
ble 4) that are acquired on different nodes to be merged into
one telemetry downlink. Furthermore, the CAN bus is used
to initiate state transfers and monitor the current state of the
nodes.
Table 4. Compass Data Acquisition
Measurement Range Sample Rate
a±12 g100 Hz
a±24 g400 Hz
a±64 g800 Hz
ω±1000 ◦s−1100 Hz
ω±2000 ◦s−11000 Hz
T(4x) −50 −250 ◦C 10 Hz
pCC 0−50 bar 1000 Hz
pT ank 0−100 bar 1000 Hz
pAmbient 10 −1200 mbar 100 Hz
GPS - ≈5 Hz
Despite all required measurements for an apogee detection
being available, the development of a parachute deployment
avionics was deemed unfeasible due to the lack of adequate
testing possibilities. Therefore, a flight proven COTS product
in form of two Altus Metrum Telemega v4 was chosen.
The power management structure is displayed in Figure 13.
The power control and distribution unit (PCDU) consists of
two converter boards (CV) and one switch board (SW). The
CV implements a radio silence switch (RSS) which can be
controlled by the ground support equipment (GSE) to disable
2nd International Conference on Flight Vehicles, Aerothermodynamics and Re-entry Missions & Engineering (FAR)
19 - 23 June 2022. Heilbronn, Germany
the entire avionics. All required voltages are provided by the
CV. In order to avoid a single point of failure each Telemega
is directly connected to one CV. The SW combines the out-
puts of both CVs to supply the remaining avionics nodes. In
case of a battery or converter failure the SW will automati-
cally switch to the remaining functioning CV. In total, three
power sources are used: The GSE, which is only present prior
to launch, a four cell lithium ion battery acting as main power
source as well as a three cell lithium polymer battery used for
the pyrotechnical igniters.
Fig. 13. Compass Avionics Power Management
4. LAUNCH REPORT
Compass was launched on 25th of June, 2021 from the Trup-
penuebungsplatz (military training area) Heuberg (see Fig-
ure 15). Due to altitude and range restrictions on the launch
site, the rocket was filled remotely with only 1.85 kg nitrous
oxide (46 % of tank volume). The GSE ignited the engine
1.5 s before opening the main valve. Figure 14 depicts the
height and roll rate before drogue parachute deployment over
time. The roll rate increased to 1.6 Hz at the end of the thrust
phase (t+9.5 s). This could be attributed to fin misalignment.
Compass reached a height of 3220 m at t+24 s. The drogue
parachute was deployed and lead to an average descent rate
of 21.6 m s−1.
0 50 100 150
0
1000
2000
3000
4000
Height / m
Height
Roll Rate
0 50 100 150 0
0.5
1
1.5
2
Time / s
Roll Rate / Hz
Fig. 14. Height and Roll Rate of Compass
The main parachute was triggered but unable to inflate. Thus,
the rocket performed a drogue landing and was subsequently
damaged. Due to lack of adequate testing the self developed
avionics down link was not utilized resulting in only telemetry
data of the Telemegas being transmitted in flight. The avion-
ics node responsible for acquiring engine and tank pressure
failed to record the measurements. The cause of the main
parachute deployment failure was traced back to the retain-
ing straps that secure the bag in the recovery bay during the
descent phase. The straps likely failed after the hatch was
opened, causing the main parachute bag to get tangled around
the fuselage and in its own harness. Therefore, the main
parachute could not be pulled out of the bag by the drogue
parachute.
(a) Oxidizer Loading (b) Launch
Fig. 15. Compass Launch
5. LESSONS LEARNED FOR N2ORTH
In retrospect, the most important contribution of the Compass
demonstrator rocket to the project was the experience gain of
the team. This ranges from effective and constructive commu-
nication in demanding situations and project management to
practical experiences with the technical systems and working
fluids. However, the scalability of the technical challenges
is limited. The supersonic flight regimes expected for the
N2ORTH rocket can not be reached with a subscale demon-
strator. Moreover, the monetary, labor and time effort for
building of the Compass rocket were significant. Despite the
limited scalability, a number of important technical lessons
could be learned. The design flaw of the recovery systems
was identified and addressed in the design of the N2ORTH
rocket. The fin alignment procedure was changed, including
improved methods to quantify the misalignment. Important
experience was gained regarding the transferability of the in-
jection behavior from static to blow-down operation, and pre-
diction of the overall propulsion system performance.
2nd International Conference on Flight Vehicles, Aerothermodynamics and Re-entry Missions & Engineering (FAR)
19 - 23 June 2022. Heilbronn, Germany
6. ACKNOWLEDGMENTS
The DLR STERN project is funded by the DLR Space Ad-
ministration with funding from the Federal Ministry for Eco-
nomic Affairs and Climate Action. HyEnD’s participation is
administered by the Institute of Space Systems at the Uni-
versity of Stuttgart. All hot-fire tests were performed at the
test complex M11 of the DLR Institute of Space Propulsion
in Lampoldshausen. The launch took place at the military
training area (Truppenuebungsplatz) Heuberg of the German
Armed Forces. The support of the DLR department of satel-
lite and orbital propulsion and the personnel of Heuberg is
greatly acknowledged. While the authors of this paper were
responsible for the development of the presented subsystems
of Compass, the project would have not been possible without
the work of more than 50 active students.
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