REXUS-4 - Vehicle and Experiments, Outlook on the REXUS/BEXUS Student Programme

Abstract

On the 22nd of October 2008, EuroLaunch launched the REXUS-4 rocket at Esrange in Northern Sweden. EuroLaunch is a joint venture of the DLR Mobile Rocket Base and the SSC Esrange Space Center. REXUS-4 was a two-stage unguided solid propellant sounding rocket. The vehicle consisted of a Nike motor as 1st stage, an Improved Orion motor as 2nd stage, a motor adapter, a recovery system, a service system, two experiment modules, and a nosecone. The REXUS-4 payload was comprised of five technological experiments from German and Swedish Universities. The rocket was spin-stabilized during the ascent. After the burn-out of the 2nd stage a yoyo system de-spun the rocket to a rate of only a few degrees per second. At an altitude of 71 km the nosecone was jettisoned. The payload reached its apogee at 175 km.
The REXUS-4 mission was also the maiden flight of a newly developed rocket service system. After this successful demonstration, it has been implemented into the REXUS/BEXUS programme. This German-Swedish student programme offers annual flights for student experiments on sounding rockets and stratospheric balloons.
This paper gives a short overview on the development of the REXUS service system and points out the advantages of using standard interfaces for student experiments. Furthermore it contains a description of the REXUS-4 vehicle, the mission, the campaign and the experiments. Some experiments are described in more detail. During the ballistic flight the MIRIAM experiment of the University of Armed Forces in München and the Mars Society Germany was separated from the main payload to test a balloon system that will be used for the entry of a probe in the Martian atmosphere in the future. Several cameras on the REXUS-4 payload as well as cameras and telemetry on the MIRIAM flight system monitored the separation and inflation during the ballistic flight phase. The VERTICAL experiment from the Technical University München verified the start-up procedures of the CubeSat MOVE and its solar panel deployment under real spaceflight conditions. The paper also gives an overview on the REXUS/BEXUS programme and its chances for students.

Full text

IAC-09-E1.1.2
REXUS-4 - VEHICLE AND EXPERIMENTS, OUTLOOK ON THE REXUS/BEXUS
STUDENT PROGRAMME
Andreas Stamminger
Deutsches Zentrum für Luft- und Raumfahrt (DLR), Institute of Space Systems, Robert-Hooke-Str. 7, 28359 Bremen,
Germany, Tel.: +49-421-24420-124, Email: andreas.stamminger@dlr.de
Manuel Czech
(2) Technische Universität München, Institute of Astronautics, Boltzmannstr. 15, 85748 Garching, Germany,
Tel.: +49-89-289-16019, Email: m.czech@lrt.mw.tum.de
Hannes Griebel
(3) Universität der Bundeswehr München, Institute of Space Technology, Werner-Heisenberg-Weg 39, 85579
Neubiberg, Germany, Tel.: +49-89-6004-2127, Email: hannes.griebel@unibw-muenchen.de
Marcus Hörschgen
(4) Deutsches Zentrum für Luft- und Raumfahrt (DLR), Mobile Rocket Base, Oberpfaffenhofen, 82234 Wessling,
Germany, Tel.: +49-8153-28-2172, Email: marcus.hoerschgen@dlr.de
Olle Persson
(5) Swedish Space Cooperation SSC, Esrange, P.O. Box 802, 98128 Kiruna, Sweden,
Tel.: +46-980-7-2205, Email: olle.persson@esrange.ssc.se
Markus Pinzer
(6) Deutsches Zentrum für Luft- und Raumfahrt (DLR), Mobile Rocket Base, Oberpfaffenhofen, 82234 Wessling,
Germany, Tel.: +49-8153-28-3028, Email: markus.pinzer@dlr.de
Jens Rießelmann
(7) Technische Universität Berlin, Institut für Luft- und Raumfahrttechnik, Marchstr. 12, 10587 Berlin, Germany, Tel.:
+49-30-314 - 24438, Email: jens.riesselmann@ilr.tu-berlin.de
ABSTRACT
On the 22nd of October 2008, EuroLaunch launched the REXUS-4 rocket at Esrange in Northern Sweden. EuroLaunch
is a joint venture of the DLR Mobile Rocket Base and the SSC Esrange Space Center. REXUS-4 was a two-stage
unguided solid propellant sounding rocket. The vehicle consisted of a Nike motor as 1st stage, an Improved Orion motor
as 2nd stage, a motor adapter, a recovery system, a service system, two experiment modules, and a nosecone. The
REXUS-4 payload was comprised of five technological experiments from German and Swedish Universities.
The rocket was spin-stabilized during the ascent. After the burn-out of the 2nd stage a yoyo system de-spun the rocket
to a rate of only a few degrees per second. At an altitude of 71 km the nosecone was jettisoned. The payload reached its
apogee at 175 km.
The REXUS-4 mission was also the maiden flight of a newly developed rocket service system. After this successful
demonstration, it has been implemented into the REXUS/BEXUS programme. This German-Swedish student
programme offers annual flights for student experiments on sounding rockets and stratospheric balloons.
This paper gives a short overview on the development of the REXUS service system and points out the advantages of
using standard interfaces for student experiments. Furthermore it contains a description of the REXUS-4 vehicle, the
mission, the campaign and the experiments. Some experiments are described in more detail. During the ballistic flight
the MIRIAM experiment of the University of Armed Forces in München and the Mars Society Germany was separated
from the main payload to test a balloon system that will be used for the entry of a probe in the Martian atmosphere in
the future. Several cameras on the REXUS-4 payload as well as cameras and telemetry on the MIRIAM flight system
monitored the separation and inflation during the ballistic flight phase. The VERTICAL experiment from the Technical
University München verified the start-up procedures of the CubeSat MOVE and its solar panel deployment under real
spaceflight conditions. The paper also gives an overview on the REXUS/BEXUS programme and its chances for
students.

1. INTRODUCTION
The development of REXUS-4 started with the Kick-Off
Meeting in December 2006, held by EuroLaunch at DLR
Mobile Rocket Base in Oberpfaffenhofen. Two major
drivers have been influencing the time plan: The
development of the new service system and the main
scientific payload, the MIRIAM experiment. The
MIRIAM experiment is a probe, located under the
nosecone during the ascent phase and ejected exo-
atmospherically. This re-entry experiment needed a
higher altitude than a standard single-stage REXUS
comprising an Improved Orion rocket motor can achieve.
This was the reason to use a two-staged motor
combination.
Figure 1: REXUS-4 Vehicle on Launcher
Even though the launch had to be postponed several
times, the major requirement concerning the REXUS
programme was to have a successful flight of the new
service system with a large enough time gap before the
REXUS-5 flight in March 2009. This was achieved by the
launch in October 2008. Table 1 shows the critical time
frame in overview.
2. VEHICLE DESIGN AND SUBSYSTEMS
REXUS-4 was a two-stage unguided solid propellant
sounding rocket. The vehicle consisted of a Nike motor as
1st stage, an Improved Orion as 2nd stage, a motor adapter,
a recovery system, a service system, two experiment
modules, a nosecone adapter ring with the MIRIAM
electronics and an ejectable ogive nosecone.
Milestones Date
Kick-Off Meeting 2006-12
REXUS SM Assembly 2007-06
MIRIAM Life Cycle Test 2008-08-22
REXUS SM Vibration Test 2008-09-03
MIRIAM Vibration Test 2008-09-12
Experiment Acceptance Tests 2008-09-15
Flight Simulation Test at DLR 2008-09-18
Transport of EGSE to Esrange 2008-09-29
Transport of PL to Stockholm 2008-10-01
Balance Test at Packforsk, Stockholm 2008-10-07
Beginning of Campaign at Esrange 2008-10-13
Practise Countdown 2008-10-22
Launch of REXUS-4 2008-10-22
Table 1: The REXUS-4 Schedule
2.1. Lift-Off Configuration
The total lift-off mass of REXUS-4 was 1175 kg with the
motor mass contributing more than 1000 kg. Including
both motors the REXUS-4 rocket had a length of 9.5 m.
The payload mass was 152 kg including a scientific
payload mass of 75 kg. The polar moment of inertia of the
payload was measured to be 2.65 kgm².
1st Stage: 598.5 kg
2nd Stage 424.5 kg
Payload 152.3 kg
Motor adapter + Balancing Mass 1
Recovery System
IGAS Module
Service System
Extension Ring
Experiment Module 1
Exp. Mod. 2 + Ballast + Bal. Mass 2
MIRIAM Camera Module
MIRIAM + Pod
Nosecone
11.9 kg
19.9 kg
10.7 kg
18.9 kg
2.3 kg
10.7 kg
23.3 kg
8.9 kg
31.7 kg
13.9 kg
Total 1175.3 kg
Table 2: The REXUS-4 Mass Budget [17]
Two requirements on the payload configuration are
important to ensure a safe flight of a sounding rocket. The
first requirement is the center of gravity during lift-off
that has to be ahead of the center of pressure.

The center of gravity for the lift-off configuration
referenced to the aft plane of the motor adapter was
measured at 1315 mm, 40.6 % of the payload length [12].
Figure 2: REXUS-4 two-stage unguided solid propellant
sounding rocket [17]
2.2. Re-entry Configuration
For the re-entry a center of gravity close to 50 % of the
payload length is necessary to avoid a stable attitude of
the vehicle during the descent and prevent the recovery
system from too much heating. Since the main experiment
MIRIAM was ejected during the ballistic flight it was
necessary to integrate 12 kg of ballast mass into the
nosecone adapter.
The center of gravity for the re-entry configuration was
measured at 780 mm (50.1 %) which ensured near perfect
re-entry conditions. The re-entry payload mass was 95 kg.
3. THE NEW SERVICE SYSTEM
The newly developed REXUS service system has the
capability to provide five experiments with 1 ampere at
28 volts and a serial data interface for up- and downlink.
The service system provides GPS position and velocity
data, 3-axis acceleration data and rotation rates from 3-
axis rate gyros. A standard TV channel can be used by
one experiment. More information on the REXUS service
system can be found in [5].
4. EXPERIMENTS
The REXUS-4 payload comprised of five university
experiments and one DLR Mobile Rocket Base
experiment in an extra module.
Figure 3: REXUS-4 experiment configuration [17]
During the ballistic flight the MIRIAM experiment of the
Universität der Bundeswehr München and the Mars
Society Germany was separated from the main payload to
test a balloon system that will be used for the entry of a
probe in the Martian atmosphere in the future. Several
cameras on the REXUS-4 payload as well as cameras and
telemetry on the MIRIAM flight system monitored the
separation and inflation during the ballistic flight phase.
The VERTICAL experiment from the Technische
Universität München verified the startup procedures of
the CubeSat MOVE and its solar panel deployment under
real spaceflight conditions.
The HISPICO experiment of the Technische Universität
Berlin tested a high- integrated S-Band transmitter. The
REWICAS of the Technical University of Luleå consisted
of three cameras. The EMSADA experiment from the
same university is a multiple sensor and data acquisition
unit [17].
4.1. MIRIAM
The MIRIAM Experiment (Main Inflated Reentry Into the
Atmosphere Mission test for ARCHIMEDES) is placed
under the nosecone. It was the spaceflight test for the
ARCHIMEDES project. ARCHIMEDES is an effort to
probe the atmosphere of planet Mars by means of a
hypersonic drag balloon, a device known as a “ballute”
[10], [14] (a term coined by the Goodyear Corporation in
1959 by combining the terms “balloon” and “parachute”)
[9]. The project is currently under study, proposed and
supported by the Mars Society Germany, the Universität
der Bundeswehr München, the AMSAT-DL e.V.
organization and several other research institutions and
industrial companies [6].
Figure 4: MIRIAM flight system stack
4.1.1. MIRIAM Experiment Design
The most important step in the development of
ARCHIMEDES so far was the spaceflight test MIRIAM.
MIRIAM was to test the deployment mechanism, the

inflation process, the ballute behaviour during inflation
and the high speed entry into the atmosphere on a ballistic
trajectory [8]. This is important to validate trajectory- and
CFD analyses [7].
The spaceflight system consisted of 3 major elements:
- The MIRIAM ballute spacecraft which comprises an
instrumented pod and the helium-filled hypersonic
drag balloon (ballute).
As stated above, MRIAM’s pod instrumentation
already closely resembled that of ARCHIMEDES, but
purely for flight analysis purposes. The FMI-provided
ATMOS-B pressure sensor was installed inside the
balloon. The magnetometer for MIRIAM (MiriMag)
was contributed by the IGEP institute and MAGSON.
Paired with an optical still image camera it provided
attitude information. The still image camera though is
a commercially available low resolution unit which can
be integrated cheaply and easily and is sufficient for an
occasional attitude fix in combination with the other
sensors. A suite of two different sets of accelerometers
built by the ARCHIMEDES team and universities in
Iasi and Pitesti, Romania gave deceleration and roll
rate information.
- The Miriam Ballute was 4m in diameter and made of
UPILEX 25 RN. The ballute had 32 segments which
were bonded with a high-temperature high strength
UPILEX-RN tape specially manufactured to MIRIAM
mission specifications by Lohmann Tapes of Neuwied.
- The Service Module (SM, see Figure 5) contained the
inflation system, structural box, release mechanism, a
telemetry and a live television subsystem. It also
contained a set of cold gas thrusters. These thrusters
were used to pull the Service Module away from
MIRIAM after inflation.
- The Camera Module which remained attached to the
rocket. It documented the release and operation of the
SM/Miriam system, as well as providing the structural
interface between the MIRIAM flight system stack and
the REXUS payload section. The camera module also
contained the release mechanism for the Service
Module / Miriam combined system.
All three elements combined formed the MIRIAM Flight
System Stack (see Figure 4).
Due to strong cost requirements in the development of
MIRIAM’s structure, special emphasis had to be put on
commercially available and reliable components which
are suitable for use in a zero-g and vacuum environment.
All components have to also withstand the loads
occurring during the launch and the ascent on top of a
solid fuel rocket.
Figure 5: MIRIAM Service Module
The Helium gas was stored in three CFRP pressure tanks
at 200 bars (see Figure 5). These tanks are carbon-fibre
wound low cost tanks normally used for paintball games.
It was found that these components fulfil the MIRIAM
requirements very well. To reduce the piping between the
tanks and the valves, most of the piping was integrated
into the base plate of the inflation systems deck. This
plate is divided horizontally into two parts and the
plumbing is milled directly into the aluminium plates.
Sealing is done by conventional O-rings. Tests have
shown that leakage is not a problem. The valves are
connected directly to the channels in the base plate via
special adapters with integrated gas channels. The only
pipes necessary are those for connecting the channels with
the thrust nozzles. One more is used to connect the main
inflation control valve (ICV) to the central expansion
chamber assembly.
Figure 6: MIRIAM’s Instrument Pod Layout and
Instrument Positions. The Green Area Labelled
PodCon represents the instruments pod’s flight
sequence control computer

Due to weight and volume limitations, no conventional
pressure regulator could be used. Two chambers of
expansion volumes together with specially designed
throttles and an inflation control valve (ICV) were used
instead. The ICV was a high pressure injector valve and is
controlled dynamically at a high frequency, thereby
controlling the inflation hose pressure through pulse
width modulation [3].
The Instrument Pod was directly attached to the ballute. It
consisted of a hexagonal shaped thin-walled container,
completely milled out of aluminium. It contained all the
sensors, camera, transmitter and computer as well as the
batteries of the ballute spacecraft. It was located mainly
inside the balloon. The ballute envelope thin film was
clamped between the container and the circular cover.
This cover had the same curvature as the ballute itself to
give the spacecraft a perfectly spherical shape. All
communication between the Instrument Pod and the
Camera Module (CM) was done via an infra-red link
through the camera window.
Figure 7: MIRIAM Mission Sequence and Elements: The
Deployment (left) and the inflation of the Miriam
spacecraft (right). Note size ratio of balloon / service
module, which dictates the necessary packaging
efficiency
The folded ballute spacecraft was held down inside the
spring-loaded deployment container by a clamp ring.
4.1.2. MIRIAM Mission Results
MIRIAM unfortunately did not meet all of its mission
goals after one of the three main interlock bolts failed to
properly separate the spacecraft from the rocket. As a
result the MIRIAM ballute spacecraft was not fully
deployed. However, the deployment and inflation control
systems functioned as planned, as well as the observation
platform, validating the method and yielding important
data and experience based upon which an improved
system can be designed.
The next step in the development of ARCHIMEDES
would therefore be the flight test of an improved version
of MIRIAM. Based on the architecture of MIRIAM, its
successor would feature not only a different separation
mechanism, but also an improved ballute, an improved
spacecraft bus and an improved observation system.
However, funding for this mission remains to be raised.
Figure 8: MIRIAM Experiment after Seperation [Source:
MarsSociety]
4.2. VERTICAL
VERTICAL (VERification and Test of the Initiation of
CubeSats After Launch) is a mission of the Institute of
Astronautics (LRT) of the TU-München for verification
of critical components for the pico-satellite First-MOVE
(Munich Orbital Verification Experiment) [4], which will
be launched in late 2009.
Figure 9. The VERTICAL experiment consists of the
Solar Panel Verification Assembly (SPVA), the
Deploymen Switch Verification Assembly (DSVA)
and the Electronics Module (EM) [15]
4.2.1. VERTICAL Mission Objectives
As the purpose of VERTICAL is to verify components
which are used very closely after separation of a satellite
from a launch vehicle, sounding rockets are a particularly

suitable option for verification of these operations, as
mission characteristics are very similar to those of an
orbital mission in such an early phase. A flight on a
sounding rocket closely simulates the following
conditions to those on a CubeSat orbital launcher:
- Load characteristics during ascent phase
- Decrease of pressure to vacuum during ascent phase
- Vacuum environment in high altitudes beyond Earth
atmosphere during experiment operations
- Reduced gravity environment during experiment
operations
Verification Items (VI) have been phrased which shall
define how those critical systems can be efficiently
verified using the flight opportunity on Rexus-4.
- Deployment Switches (DS): Launch service providers
require CubeSats to be powered off during launch until
deployment. Normally micro switches are used for
deployment detection, which detect the loss of contact
to the launch vehicle. All following operations,
including power up, require feedback from these
switches to be initiated. On VERTICAL, a total of 16
commercial switches are verified tocover a broad
spectrum of manufacturers, the CubeSat standard only
requires oneswitch on-board.
- VI2 Solar Array Deployment Mechanism (SADM):
MOVE’s solar panels have to be stowed during launch
and cannot be deployed before or during the ejection of
the CubeSat from the launch vehicle. The SADM,
consisting of a set of springs and a Hold Down Release
Mechanism (HDRM) was newly developed at TUM
and has no space heritage. The HDRM is activated by
melting a nylon wire and the solar panel is deployed by
a set of springs.
Figure 10: The VERTICAL Experiment
The function of the DSVA is to verify the deployment
switches. A total of 16 micro switches are mounted on
two brackets. During the launch the switches have to be
depressed. This is accomplished by a so-called retraction
plate which is actively pushed towards the switches by an
actuator. After launch the switches have to be triggered,
then the actuator is commanded to unlock the plate, which
is pushed away from the switches by four springs. The
retraction plate retracts with 0.2 m/s, which is slower than
an actual CubeSat deployment [11].
The SPVA achieves two objectives, it detects the
deployment of the solar panel, confirming the HDRM’s
function and it also measures the panel’s motion during
deployment. To accomplish this, photo diodes and LEDs
are mounted opposite each other on two plates located on
both sides of the solar panel. When the panel deploys, the
sensors are shaded for a very short time.
Figure 11: VERTICAL Schematic and Hardware
4.2.2. VERTICAL Mission Results
Out of the 16 deployment switches, 15 worked as
expected. All working switches remained open during the
ascent. When the retraction plate was commanded to
retract, all switches closed within 73 ms of each other.
This was consistent with ground test results and is due to
the mechanical configuration of the individual switches.
The HDRM released the solar panel 1.48 s after the
melting wire was activated. This was the same duration as
observed during ground tests. The maximum panel
velocity observed was 1.66 m/s, this is also consistent
with ground test results. The EM performed as expected
and did not show any anomalous behavior. More detailed
results can be found in a paper concerning the
performance of the experiment [15].

4.3. HISPICO
One of the main focuses of the technology research at the
Chair of Astronautics of the TU Berlin is the development
of pico-satellites and their subsystems by involving
students. A High Integrated S-Band Transmitter for Pico-
Satellites (HISPICO) was developed by TU Berlin and the
company IQ wireless GmbH.
HISPICO achieves 1 Mbps data rate with a transmission
power of 27 dBm and power consumption of 5 W. These
technical parameters can only be reached by the use of
modern channel coding algorithms – especially
Turbocode. Turbocode is typically used in deep-space and
large satellite missions. The innovation of HISPICO lies
in adoption of turbo-coding for pico-satellite applications.
The communication of nearly all launched pico-satellites
is UHF based [1].
Figure 12. HISPICO system mounted on bulkhead (from
top left to bottom right): HISPICO with power
splitter, power control and data handling unit
(PCDH), camera
4.3.1. HISPICO Experiment Design
The goal of the experiment was to test the RF connection
between HISPICO and the ground station.
The main components are the S-band transmitter
HISPICO and the patch antennas. Because of the angle of
beam spread of the antennas of about 90° four S-band
patch antennas are mounted axially symmetrical at the
rocket. The Radio Frequency signal was divided by a 4-
way splitter to the antennas. To control the transmitter and
to provide power and data a Payload Control and Data
Handling unit (PCDH) was used. In addition to this a
camera delivers payload data. The PCDH and the camera
represent an electrical and data interface of a “real”
satellite [2].
During the REXUS-4 flight the HISPICO system were
mounted on a bulkhead as shown in Figure 12.
4.3.2. HISPICO Mission Results
After 55 seconds simultaneously to YoYo despin of the
rocket at an altitude of 66km HISPICO was started to
transmit data at a frequency of 2228.5MHz. Ten minutes
after the beginning of experiment rocket power is
switched off and the test ended.
Periodical fluctuations of the receiving signal could be
meassured. They were caused by the interferences of the
four antennas in their overlapping area of beams. Also
after the YoYo despin the rocket was slightly rolling. At a
realistic satellite communication link these interferences
do not appear.Due to these effects of a sounding rocket
environment the overall bit error rate was not better then
10E-3. Filtered of the periodically errors caused by the
antenna interferences the calculated bit error rate was
better then 10E-5 [2].
The successful data link demonstrated the resistance of
HISPICO against launch loads and high relative motion,
caused by the rocket.
4.4. EMSADA
The EMSADA Experiment (Experimental Multiple
Sensors And Data Aquisition) of the Luleå University of
Technology was mounted in the experiment module. It
measured acceleration, pressure and temperature in the
experiment.
4.5. REWICAS
The final version of the REWICAS system of the
Technical University Luleå consisted of 3 cameras, all
placed inside their own box containing individual
controlling electronics and a 9V battery. These 3 camera
boxes were then connected to the main box where the
video was to be stored. All the other components of
REWICAS (such as a video receiver, channel chooser,
analog-digital recorder, IR-controller, microcontroller and
a switching regulator power supply) were mounted inside
the main box.
4.5.1. REWICAS Mission Objectives
REWICAS’s primary objective was to visually present a
soundings rocket's flight path from 3 different angles.
Starting from lift-off, the smooth 25 pictures per second
video would contain rocket separations, parachute release
and later on touch-down. The secondary objective was to
test and promote the use of COTS equipment on sounding
rockets.

4.5.2. REWICAS Mission Results
Due to an electrical leakage from the batteries of the
camera boxes. No video was unfortunaly captured from
space. This leakage was known and some electrical
adjustments were made pre launch. Even so, a slow
degradation of the batteries power supply took place
during the 24 hour rocket assembly and lift-off
preparation, together with the night in storage waiting for
launch. Thus, already by lift-off, the batteries were
depleted and could not provide any power to the cameras.
All other functions operated nominally.
4.6. IGAS (Intelligent GPS Antenna System)
The IGAS experiment was developed at the DLR Mobile
Rocket Base and consisted of an intelligent antenna and a
standard antenna system with the appropriate GPS
receivers. Additionally, a new type of command antenna
for P-Band (450 MHz) was qualified. More information
can be found in [13].
5. FLIGHT RESULTS
REXUS-4 was launched from Esrange, Sweden on the
22nd of October 2008 at 12:30 UTC. The maximum
acceleration of 20.3 g was reached after 2.5 s and the
burn-out of the 1st stage was 3.4 s. The rocket was spin-
stabilized during the ascent. After the burn-out of the 2nd
stage a yoyo system de-spun the rocket at an altitude of
66 km to a rate of only a few degrees per second.
Figure 13: Roll-Rate during Ascent Phase [17]
At an altitude of 71 km the nosecone was jettisoned. The
MIRIAM release was initiated at 76 s but a fully
separation of the experiment occurred 13 s later at an
altitude of 109 km. The 2nd stage motor was separated
from the payload at 112 s in an altitude of 132 km.
Because of the late MIRIAM release and impulse of the
separating motor on the main payload it approached
towards the MIRIAM probe and collided at an altitude of
135 km. Both, the main payload and MIRIAM reached its
apogee in 175 km.
Figure 14: REXUS-4 Payload and Motor after the Ascent
Phase [Source: Mars Society]
The vehicle reached the lower atmosphere at altitudes of
<100 km and by definition started the re-entry at 335 s.
The maximum deceleration during the descent phase
occurred 53 seconds later at 26 km altitude.
The heat shield, stab-chute and beacon of the recovery
system were activated at 4.5 km altitude. The stab-chute
was de-reefed 5 seconds later and the main chute released
at 3.1 km altitude. The payload landed safely on ground
and was recovered by the helicopter and brought back to
Esrange one hour later in excellent condition.
No
. Time [s] Alt [km] Event
0 T-600.0 0.332 Experiments Power On
(except HISPICO)
1 T-360.0 0.332 MIRIAM Wake Up in
Flight Standby
2 T-180.0 0.332 MIRIAM F-SCET Sync to -
180 and main interlock
actuators hold lock engage
3 T-120.0 0.332 HISPICO Power On
4 T- 90.0 0.332 Internal Power

5 T- 60.0 0.332 REWICAS Start of
Experiment
6 T- 60.0 0.332 VERTICAL Start of Data
Storage
7 T+ 0.0 0.332 Lift-Off
8 T+ 0.0 0.332 MIRIAM Lift-Off, F-SCET
Reset to 0
9 T+ 0.0 0.332 VERTICAL Lift-Off Signal
10 T+ 0.0 0.332 HISPICO Activation
11 T+ 0.0 0.332 EMSADA Activation of the
acceleration sensor
12 T+ 0.0 0.332 REWICAS Lift-Off Signal
13 T+ 0.0 0.332 IGAS Lift-Off Signal
14 T+ 2.5 0.808 Maximum Acceleration
(20.3 g)
15 T+ 3.4 1.307 Burnout 1st Stage (Nike)
16 T+ 9.2 4.207 Ignition 2nd Stage (Imp.
Orion)
17 T+ 35.7 36.978 Burnout 2nd Stage (Imp.
Orion)
18 T+ 54.8 66.057 YoYo Despin
19 T+ 55.0 66.344 HISPICO Signal (together
with Yo-Yo Despin)
20 T+ 58.9 71.739 Nosecone Ejection
21 T+ 62.0 76.158 EMSADA Signal
22 T+ 64.0 78.869 MIRIAM P4MS
Subsequence (Prepare for
MIRIAM Separation)
23 T+ 76.0 94.375 MIRIAM SM Release
(Separation)
24 T+ 89.0 109.617 MIRIAM Balloon
Deployment (Clamp Ring
Release)
25 T+ 89.0 109.617 MIRIM Separation (as
occurred)
26 T+111.6 132.306 Motor Separation
27 T+115.0 135.304 P/L – MIRIAM Collision
28 T+170.0 168.764 VERTICAL Activation of
mechanism
29 T+170.0 168.764 HISPICO Signal (together
with VERTICAL Signal)
30 T+207.6 175.412 Apogee
31 T+224.0 174.161 MIRIAM Balloon Release
from MIRIAM SM
32 T+335.0 99.791 Begin of Atmospheric
Reentry + Camera Switch
33 T+388.7 26.745 Maximum Deceleration (8.9
g)
34 T+400.0 ~ 16.000 Begin of subsonic flight
35 T+501.0 4.489 Heatshield, Stab Chute
Activation + Beacon
Activation
36 T+507.3 3.983 Stab Chute De-Reefing
37 T+525.7 3.141 Main Chute Activation
38 T+532.5 2.977 Main Chute De-Reefing
39 T+600.0 2.535 Power Off for Experiments
(except IGAS +
VERTICAL)
40 T+800.0 ~1.0 km Power Off TM/TV
Table 3: REXUS-4 Time Events [17]
6. THE REXUS/BEXUS PROGRAM
The REXUS-4 mission was the maiden flight of a newly
developed rocket service system. After this successful
demonstration, it has been implemented into the
REXUS/BEXUS programme. This German-Swedish
programme allows students from universities across
Europe to carry out scientific and technological
experiments on research rockets and balloons. Each year,
two rockets and two balloons are launched, carrying up to
20 experiments designed and built by student teams.
The REXUS/BEXUS programme is realised under a
Bilateral Agency Agreement between the German
Aerospace Center (DLR) and the Swedish National Space
Board (SNSB). The Swedish share of the payload has
been made available to students from other European
countries through a collaboration with the European
Space Agency (ESA).
EuroLaunch is responsible for the campaign management
and launch vehicle operations. Experts from SSC, DLR
and ESA provide technical support to the student teams
throughout the project.
Figure 15: REXUS/BEXUS Programme Logo

Since the first “Call for Proposals” in Autumn 2007
twenty student experiments in the scientific fields of
atmospheric physics, Earth magnetic field, milligravity,
space biology, communications, balloon or rocket
technology, space technology and reentry technology
have flown on four BEXUS and two REXUS missions
[16].
The “Call for Proposals” for current round of the student
programme is open until November 2009. For further
information visit www.rexusbexus.net.
7. REFERENCES
1. Alavi R. / Brieß K. / Jäckel K. / Podolski H., S-Band-
Sender für Nano- und Pico-Satelliten, Deutscher
Luft- und Raumfahrtkongress 2008, Darmstadt, Sep.
2008.
2. Alavi R. et al., In Space Verification of the Pico-
Satellite S-Band Transmitter “HISPICO” on a
Sounding Rocket, 60th International Astronautical
Congress 2009, Daejeon, Republic of Korea, Oct.
2009
3. Barth A., Durchführung von Aufblassystemtests für
den Raumflugversuch MIRIAM, Diplomarbeit,
Universität der Bundeswehr München, 2008.
4. Czech M., First-Move in Satellite Development at the
TU-München, 7th IAA Symposium on Small
Satellites For Earth Observation, 2009.
5. Ettl J. / Pinzer M., Principle Design of the Service
Module of REXUS, 19th ESA PAC Symposium, Bad
Reichenhall, 2009.
6. Griebel H. / Häusler B. / Mundt C. / Rapp H., Project
ARCHIMEDES – A Novel Approach to Balloon
Deployment on Mars, Paper IAC-04-Q.P.02, 55th
International Astronautical Congress, Vancouver,
2004
7. Griebel H., MIRIAM Spaceflight Test Mission
Definition, Version 2R1, The Mars Society
Deutschland e.V., March 2007.
8. Grieger B. et al, The ARCHIMEDES Mission Science
Proposal, Max-Planck-Institute for Solar System
Research, Katlenburg-Lindau, 2005
9. Jaremenko I. M., Ballute Characteristics in the 0.1 to
10 Mach Number Speed Regime, Goodyear
Aerospace Corporation, Journal of Spacecraft and
Rockets, vol.4, No.8, 1967, pp. 1058-1063
10. Katzlowski, M. / Griebel H., The Atmospheric Entry
of ARCHIMEDES Balloon, 3rd International
Symposium on Atmospheric Reentry Vehicles and
Systems, Arcachon, March 2003
11. Lan W, Poly Picosatellite Orbital Deployer Mk III
IC, California Polytechnic State University, 2007.
12. Löfgen O., Inertia and balancing report REXUS-4,
SRI10-S2, SSC Esrange, 2008-10-13.
13. Markgraf M. / Ettl. J. / Hassenpflug F. / Turner P.,
IGAS – A Novel GPS Antenna Concept for Spin-
stabelized Sounding Rockets, 19th ESA PAC
Symposium, Bad Reichenhall, 2009.
14. Mundt Ch. / Griebel H. / Welch Ch., Studies of
atmospheric entry of vehicles with very low ballistic
coefficient, 13th AIAA/CIRA International Space
Planes and Hypersonic Systems and Technologies
Conference, Capua, Italy, May 2005
15. Olthoff C. / Purschke R. / Winklmeier R., Testing of
Critical Pico-Satellite Systems on the Sounding
Rocket REXUS-4, 7th IAA Symposium on Small
Satellites for Earth Observation, 2009.
16. Roth M. / Magnusson P. / Page H., The First Two
Years Of The REXUS/BEXUS Student Programme,
19th ESA PAC Symposium, Bad Reichenhall, 2009
17. Stamminger A., REXUS-4 Campaign Report,
EuroLaunch, 2009
18. Stamminger A. / Czech M. / Griebel H. / Hörschgen
M. / Persson O. / Pinzer M. / Riesselmann J. /
Shahsavar A., REXUS-4 – Vehicle and Subsystem
Design, Flight Performance and Experiments