# REX: X-ray experiment on the Water Recovery Rocket

Martin Urban<EMAIL_ADDRESS>Ondřej Nentvich
<EMAIL_ADDRESS>Tomáš Báča<EMAIL_ADDRESS>Ivo Veřtát
<EMAIL_ADDRESS>Veronika Maršíková<EMAIL_ADDRESS>Daniela
Doubravová<EMAIL_ADDRESS>Vladimír Dániel<EMAIL_ADDRESS>Adolf Inneman  Ladislav Pína  Randall L. McEntaffer<EMAIL_ADDRESS>Ted B.
Schultz  Drew M. Miles  James H. Tutt  Czech Technical University in Prague,
Czech Republic  University of West Bohemia, Pilsen, Czech Republic  Rigaku
Innovative Technologies Europe, s.r.o., Prague, Czech Republic  Advacam,
s.r.o., Prague, Czech Republic  Czech Aerospace Research Centre a.s., Prague,
Czech Republic  Pennsylvania State University, State College, PA, USA

###### Abstract

This paper presents Rocket Experiment (REX) that was part of a dual-payload
rocket campaign for NASA’s sounding rocket Black Brant IX with water recovery
technology. This mission was a suborbital sounding rocket flight that was
launched and recovered on April 4, 2018 and targeted the Vela supernova
remnant. The purpose of REX was to classify the Technology Readiness Level of
onboard devices designed for space applications. The devices were two wide-
field X-ray telescopes consisting of a combination of Lobster-Eye (LE) optics
with an uncooled Timepix detector (256 $\times$ 256 $\mathrm{px}$ @
$55\text{\,}\mathrm{\SIUnitSymbolMicro m}$), and additional sensors. The first
telescope uses a two-dimensional combination of LE modules with a focal length
of $1\text{\,}\mathrm{m}$ and a Field of View (FOV) of 1.0 $\times$ 1.2
$\mathrm{\SIUnitSymbolDegree}$ and operates in the energy range of 3 – 60
$\mathrm{keV}$. The second telescope was a one-dimensional LE with a focal
length of $250\text{\,}\mathrm{mm}$ and a FOV of 2.7 $\times$ 8.0
$\mathrm{\SIUnitSymbolDegree}$ for the energy range 3 – 40 $\mathrm{keV}$. The
X-ray telescopes were supplemented by a camera in the visible spectrum with
1,280 $\times$ 1,024 $\mathrm{px}$ resolution, which was used to obtain images
of the observed sources and to verify the resulting pointing of the rocket
carrier. Other devices also include infrared array sensors and inertial
measurement units tested for future small satellite missions. The data handler
and communication system were built using the Robot Operating System, and both
the system and the electronics were deployed and operated in-flight. The
hardware was successfully recovered after the launch and the data were
extracted.

###### keywords:

Sounding rocket , X-ray , Lobster-Eye , Radiation imaging

††journal: Acta Astronautica

##  1 Introduction

Sounding rockets, a type of Suborbital rockets, provide a low-cost observing
platform for carrying instruments for scientific and educational purposes, as
well as for rapid technology validation. This is very advantageous for testing
the Technology Readiness Level (TRL) of space instrumentation without the need
for expensive boosters and advanced telemetry systems. Therefore, the price of
the launch is lower than for satellite missions. The disadvantage of using
suborbital rockets is in the limited experiment time, which is usually 5 – 20
$\mathrm{min}$. Nevertheless, they are still useful for verification of
technologies in space which do not require a long experimental time. A major
consideration for the TRL of devices used in space is not only the capability
of working in vacuum, extreme temperatures, temperature changes, and
radiation, but also the ability to survive the violent conditions during
launch and booster separation [1].

Aside from TRL verification, sounding rockets can also be used for scientific
purposes. One example of a scientific mission was the CHAMPS mission launched
in October, 2011. Two rockets were used to detect and measure the density and
size distribution of meteoritic smoke particles in the upper atmosphere [2].
Another mission was focused on the examination of the Sun’s atmosphere and was
launched in 2009 by NASA’s sounding rocket. Its goal was the research of the
helium amount and how the solar wind originated and accelerated [3]. The last
mentioned example is from the _RockOn!_ series of sounding rocket missions
carrying experiments developed by the students during NASA supported
workshops111https://www.nasa.gov/wallops/2019/feature/nasa-launches-student-
experiments-to-space-on-a-suborbital-
rocket222https://spacegrant.colorado.edu/national-programs/rockon-home.
Rockets were launched from Wallops Flight Facility in Virginia regularly since
2008 [4, 5].

The suborbital mission described in this manuscript was a two-payload campaign
carried by the NASA’s sounding rocket with water recovery technology for
science payloads which was launched from the Kwajalein Atoll in the Marshall
Islands on April 4, 2018. Figure 1 shows photos from the launch date.

(a) At the launch site.

(b) During the launch.

(c) Recovery at sea.

Figure 1: Rocket carrier and its launch from Kwajalein Atoll in the Marshall
Islands.

The observation of the Vela Supernova Remnant (SNR) was the primary scientific
astrophysical target of this sub-orbital experiment. The Vela SNR is a shell-
type remnant located in the southern constellation Vela at a distance of
approx. $250\text{\,}\mathrm{pc}$ [6]. This source is Type II and it exploded
approximately 11,000 – 12,300 years ago. The centre of the Vela SNR is formed
by a neutron star with a strong magnetic field, with a mass close to our Sun
and a diameter of approx. $20\text{\,}\mathrm{km}$ [7] (the apparent diameter
of $\approx$ $8\text{\,}\mathrm{\SIUnitSymbolDegree}$ [8]). Due to the
apparent size of the object and high surface brightness, this SNR is a
frequent source of observation and testing of equipment for both sounding
rockets [9, 10, 11] and space telescopes [12, 13]. In addition to scientific
observation, the mission aims at an important technological goal, namely the
suborbital flight verification of scientific payloads such as the Multi-Foil
Optics (MFO) type of LE with a Timepix pixel detector [14], and X-ray
reflection gratings with a hybrid Complementary Metal–Oxide–Semiconductor
(CMOS) detector.

The sounding rocket was prepared by the Pennsylvania State University (PSU)
together with the main measuring instrument — a soft X-ray grating
spectrometer (bandpass 0.25 – 0.75 $\mathrm{keV}$) named the Water Recovery
X-ray Rocket (WRXR) payload. The spectrometer consisted of a mechanical
collimator, X-ray reflection gratings, grazing incidence mirrors, and a hybrid
CMOS detector. In-depth description of the mission goals and the WRXR payload
can be found in [8, 15, 16, 17].

WRXR payloadREX payloadStar tracker mount point (a) The sensor and optic
compartment of the rocket payload.

2D telescope1D telescopeXimea cameraIR sensor (b) Apertures at the front of
the REX instrument.

Figure 2: The final arrangement of the rocket vacuum segment with REX payload
(1D and 2D LE telescopes) and WRXR payload (soft X-ray spectroscope).

The second payload — the REX — was prepared by the Czech team and contained
two sets of MFO in combination with the Timepix detector, which formed two
separate X-ray telescopes. The first of the REX telescopes consists of two LE
modules in the Schmid’s arrangement [18, 19] which covered a FOV of 1.0
$\times$ 1.22 deg in the energy range of 3 – 60 $\mathrm{keV}$. The second
telescope has a single one-dimensional (1D) LE module [20, 21] with a field of
view of 2.75 $\times$ 8.0 $\mathrm{\SIUnitSymbolDegree}$ for the energy range
3 – 40 keV. Moreover, the payload was equipped with additional sensors:
_Ximea_ camera for visible spectrum with the resolution of 1,280 $\times$
1,024 $\mathrm{px}$, an infrared (IR) sensor array and an Inertial Measurement
Unit (IMU). A photo of the payload is shown in Figure 2 with the WRXR on top
and the REX on the bottom.

##  2 REX instrument description

The REX payload consisted of two main parts: the optics modules including
sensors, and a hermetically sealed electronics box. An airtight bulkhead
separated both parts and they were electrically connected via a vacuum
feedthrough. The payload was also connected to a rocket interface panel via
another vacuum feedthrough.

The optics and sensors section contained all the sensors and detectors except
for the IMU, which was placed in the hermetic box. The sensor section was
designed as hermetically sealed in order to hold an artificial low-pressure
atmosphere to prevent liquid condensation during launch. The visible-light
camera was used alongside the X-ray telescopes as well as the IR sensor array.
A carbon-fiber baffle covered the optical path between the optic modules and
Timepix detectors to avoid side rays. A schematic of the payload is shown in
Figure 3.

Figure 3: 3D rendition of the vacuum part of the X-ray optical system REX.

The hermetic box contained NiCd batteries and switching power supply
electronics with DC/DC converters for powering the devices during the entire
duration of the experiment (approx. 2 hours). Two on-board Odroid-XU4 ARM
computers managed the sensors and miscellaneous electronics provided an
interface with the rocket. Each of the computers controlled a single Timepix
detector (one per telescope), and the rest of the sensors were evenly
distributed between the computers. Mission software using Robot Operating
System (ROS)333http://ros.org was developed to operate and control the entire
experiment. ROS is a widely adopted open-source middleware for the integration
of sensors and data processing algorithms for autonomous systems.
Simultaneously, an operating interface for Timepix detectors in ROS was
developed under the name _Rospix_ 444http://github.com/rospix/rospix [22].

###  2.1 X-ray telescope based on Lobster-Eye optics and Timepix detector

The REX experiment consisted of two LE telescopes designed by Czech Technical
University in Prague (CTU), manufactured by Rigaku Innovative Technologies
(RITE), a two-dimensional (2D) in Schmidt’s arrangement and a 1D Lobster-Eye
optics. Table 1 presents all the parameters of the optics: the coating of the
mirror layers, the focal length or FOV and the physical position of the module
in the experiment. The 1D LE optics has the same parameters as an identical
copy used in our previous successful mission on the nanosatellite VZLUSAT-1
[7, 23].

The position of both 1D and 2D telescopes are visualised in Figure 3. The 2D
telescope contained two 1D optical modules placed in series. The modules were
mutually rotated by $90\text{\,}\mathrm{\SIUnitSymbolDegree}$ along their
optical axes to form a 2D optical system. Figure 4 shows an illustration of
the 2D system configuration. A Timepix detector was placed
$1,438\text{\,}\mathrm{mm}$ from the front aperture in the focal plane of the
2D system. The 1D telescope was placed alongside the 2D, with another Timepix
detector in the focal plane. The focal distance of the 1D was
$280\text{\,}\mathrm{mm}$ and the system contains only one optical module and
thus the values for vertical dimension are void.

Figure 4: An illustration of the arrangement of the 2D optical system. Two 1D modules were combined in series with $90\text{\,}\mathrm{\SIUnitSymbolDegree}$ mutual rotation along their optical axes. Table 1: Parameters of the 1D and 2D Lobster-Eye optics for rocket experiment. | 1D | 2D
---|---|---
Module centre position (horizontal) | $250\text{\,}\mathrm{mm}$ | $1,363\text{\,}\mathrm{mm}$
Module centre position (vertical) | - | $1,097\text{\,}\mathrm{mm}$
Focal length (horizontal) | $243\text{\,}\mathrm{mm}$ | $1,345\text{\,}\mathrm{mm}$
Focal length (vertical) | - | $1,079\text{\,}\mathrm{mm}$
Field of view (horizontal) | $2.75\text{\,}\mathrm{\SIUnitSymbolDegree}$ | $1.04\text{\,}\mathrm{\SIUnitSymbolDegree}$
Field of view (vertical) |  $\approx$ $8\text{\,}\mathrm{\SIUnitSymbolDegree}$ | $1.29\text{\,}\mathrm{\SIUnitSymbolDegree}$
(limited by housing) |  |
Mirror thickness | $0.145\text{\,}\mathrm{mm}$ | $0.35\text{\,}\mathrm{mm}$
Mirror spacing | $0.3\text{\,}\mathrm{mm}$ | $0.75\text{\,}\mathrm{mm}$
Mirror dimension | 60 $\times$ 25 ${\mathrm{mm}}^{2}$ | 150 $\times$ 75 ${\mathrm{mm}}^{2}$
Number of mirrors | 56 | 2 $\times$ 47
Reflective surface | Au | Au
Angular resolution (horizontal) | $6.12\text{\,}\mathrm{arcmin}$ | $1.40\text{\,}\mathrm{arcmin}$
Angular resolution (vertical) | - | $1.59\text{\,}\mathrm{arcmin}$

####  2.1.1 Simulating the optical systems

The Vela X-1 pulsar produces a flux of $3.0\cdot 10^{-11}$
$\mathrm{erg}\text{\,}{\mathrm{cm}}^{-2}\text{\,}{\mathrm{s}}^{-1}$ in the
energy range of 4 – 25 $\mathrm{keV}$ [24] to the input aperture of each
optic. For example, the Point Spread Function (PSF) can be used with a
combination of two 1D Lobster-Eye optics for higher efficiency of the
telescope, which leads to higher photon flux on the detector and thus in
combination with a coded mask leads to a faster identification of events in
the sky as is proposed in [25]. Figures 5 and 6 show simulated results of
detector images where the optics have 100$\%$ reflectivity and the detector
100$\%$ quantum efficiency.

(a) Simulating high flux.

(b) Simulating expected flux.

Figure 5: (a) Simulated image with 100$\%$ mirror reflectivity for an energy
of $8\text{\,}\mathrm{keV}$ and a uniform distribution of the photons on the
input aperture and (b) calculated photons during the observation time
($282.5\text{\,}\mathrm{s}$). Both (a) and (b) images use 1D Lobster-Eye
optics with parameters listed in Table 1.

(a) Simulating high flux.

(b) Simulating expected flux.

Figure 6: (a) Simulated image with 100$\%$ mirror reflectivity for an energy
of $8\text{\,}\mathrm{keV}$ and a uniform distribution of the photons on the
input aperture and (b) calculated photons during the observation time
($282.5\text{\,}\mathrm{s}$). Both (a) and (b) images use 2D Lobster-Eye
optics with parameters listed in Table 1.

####  2.1.2 Timepix detector

The Timepix detector [14] was used as an X-ray detector for the REX
experiment. This hybrid silicon pixel detector developed at CERN consisted of
a high-density matrix of 256 $\times$ 256 sensitive pixels with a pixel pitch
of $55\text{\,}\mathrm{\SIUnitSymbolMicro m}$, and the total sensitive area of
the detector was 14.1 $\times$ 14.1 ${\mathrm{mm}}^{2}$. The Timepix sensor
(see Figure 7) consists of two parts: the first part is the semiconductor
detection layer ($300\text{\,}\mathrm{\SIUnitSymbolMicro m}$ silicon in this
case used for the energy range of 3 – 60 $\mathrm{keV}$), which is bump bonded
to the second part, an Application-Specific Integrated Circuit (ASIC) readout
chip containing preamplifier, Analog-to-Digital Converter (ADC), and a counter
for each pixel. Data from the sensor can be read out as fast as 100 images per
second (depending on the accompanying hardware interface [26, 27, 28, 29]),
which makes it ideal for particle tracking applications. On the other hand,
long acquisition times on the order of minutes are possible, thanks to the
sensor’s noise-less output. The matrix of all 65,535 pixels can be equalized
for a given energy threshold which makes the sensor filter out events with
lower energy than the threshold. Thus it is possible to record even single-
photon events while measuring their energy.

Timepix has already been extensively tested in space [30], e.g., onboard the
International Space Station [31, 32], the Proba-V satellite [33], CubeSat
VZLUSAT-1 [34, 35, 23], TechDemoSat-1 [36, 37], and RISESAT [38]. Two spare
detectors from the RISESAT project were used for the REX payload.

(a) Timepix sensor with Si detector.

(b) An image of particle tracks.

Figure 7: Timepix (a) sensor consists of a
$300\text{\,}\mathrm{\SIUnitSymbolMicro m}$ Si detector and an ASIC read-out
chip. The sensor produces images with 256 $\times$ 256 $\mathrm{px}$
resolution. (b) Ionizing particles leave characteristic marks (yellow —
electrons/muons, red — ions, blue — gamma), which can be processed by machine
learning algorithms [22] to extract the particle type and its energy.

###  2.2 Camera for the visible spectrum

The camera system for the visible spectrum was employed to verify the REX’s
target pointing in addition to the primary Attitude Determination and Control
System (ADCS) of Black Brant IX rocket. Verification of the small, cheap, and
uncooled industrial class CMOS camera was the secondary purpose of this
system, as it is considered for the future CubeSat missions. The camera system
was based on the Ximea MQ013CG-E2 module with a small size (26 $\times$ 26
$\times$ 26 ${\mathrm{mm}}^{3}$) and low weight ($26\text{\,}\mathrm{g}$). The
camera has a 1/1.8” CMOS sensor with $1.3\text{\,}\mathrm{Mpx}$ native
resolution. The camera module was equipped with ThorLabs MVL50M23 lens, with a
focal length of $50\text{\,}\mathrm{mm}$, offering a field of view
approximately $9.8\text{\,}\mathrm{\SIUnitSymbolDegree}$. Figure 8(b) shows
the camera module and the used lens. The low power consumption and the
aluminum camera case attached to the massive rocket mechanical structure
mitigated the possibility of the camera overheating in vacuum. An absence of
plastic materials in the camera body avoided the pollution of the telescopes
by material outgassing. The USB3 interface and the availability of drivers for
the ROS made it easier to implement the camera control software in the Odroid
on-board computer. The camera outputs uncompressed RAW images which allow
post-processing of captured images after the rocket’s recovery.

(a) Ximea camera.

(b) ThorLabs lens. Figure 8: Small CMOS camera module from Ximea with ThorLabs
lens, used in the REX experiment as star tracker.

###  2.3 Additional sensors

The infrared array sensor _Grid-EYE_ AMG88 (8 $\times$ 8 $\mathrm{px}$
resolution) was added to the REX payload as a technology demonstrator of Earth
horizon detection. It could be used as one of the inputs to ADCS algorithms of
a small CubeSat satellites. Furthermore, we employed an off-the-shelf IMU, the
MPU6000. It is a miniature low cost 3-axis Micro-electro-mechanical System
(MEMS) accelerometer and gyroscope, widely used in commercial electronics and
robotics. As this class of electronics is not dedicated for space application,
the main reason for using it in this mission was to test the capability of the
sensor to survive the hard conditions of the rocket launch. The possibility of
its future usage for recognition of flight phases and experiments triggering
as well as future usage in attitude determination and control systems of small
satellites will be examined.

##  3 Experimental results

###  3.1 Launch event and payload recovery

The Water Recovery X-ray Rocket was launched on April 4, 2018 from the
Kwajalein Atoll in the Marshall Islands. Figure 1 shows the rocket before,
during and after the launch. The total experiment time was
$919\text{\,}\mathrm{s}$ with $282.5\text{\,}\mathrm{s}$ while pointing on
target. Both telescopes survived the launch, observation and impact when the
rocket landed on the water. A damage to the rocket fuselage was caused during
descend by uneven heat dissipation. However, despite a breach of the
electronics compartment has caused a water leakage on our hermetic electronics
box, the electronics inside of the box survived and data were recovered.

Analysis of the data shown that the internal pressure in the optics
compartment has temporarily spiked during the launch [17]. This was presumably
caused by a water vapor outgassing. The optics and electronics section of the
payload were supposed to be isolated according to the original flight plan and
the optics compartment was planned to be pumped out to high vacuum. Instead,
the both sections were connected due to a last-moment failure of the ion pump
for the optics section. It is suspected that water from the payload skins
caused a rise in the compartment pressure [17] and could later condensed and
formed ice on the instrumentation.

The goal of this mission was to verify the TRL of the devices which
successfully worked during the REX mission.

###  3.2 Inertial Measurement Unit

The recorded data show consistent values from the beginning of the mission
until the rocket parachute landed into the ocean and stayed floating in the
water. All of these facts indicate that the sensor survived the mission. The
main phases of the flight are labeled in Figure 9, and it shows the thrust
phase of the first and the second rocket engine stage, fast rocket detumbling
before observation of stars, zero gravity observing phase, reentry into the
atmosphere, parachute descending, impact onto the water surface and free
floating in water. The MPU6000 sensor was partly in saturation during the full
engine thrust due to its higher sensitivity setting, but it survived an
acceleration of almost $13.7\text{\,}\mathrm{g}$ and an angular velocity of
more than $3,000\text{\,}\mathrm{d}\mathrm{e}\mathrm{g}\mathrm{/}\mathrm{s}$.
The achieved results are promising and, therefore, the MPU6000 will be used in
the future mission of PilsenCUBE II satellite.

Figure 9: Data from the onboard 3-axis MEMS accelerometer with recognized and
marked key phases of the suborbital mission.

###  3.3 Visible camera and payload pointing

Figure 10: Logged temperature profile of camera system during the whole
mission.

Reliable recognition of low brightness stars in the image required long
exposure time and high gain of the image sensor due to its small pixel pitch.
However, these parameters also increase the dark current and thermal noise
which lead to the degradation of images, especially when the image sensor has
a higher temperature. Since many aspects of the suborbital mission were not
known during the payload preparation, we did not have simulations of the
expected camera heating. We proposed a special camera capture mode with
changing exposure settings in the loop and with idle intervals between the
individual batches of image frames to overcome this issue of thermal
uncertainty. The capture mode started recording a batch of image frames as
soon as the rocket electronics were switched on shortly before the launch.
Each batch of image frames consists of several different combinations of
exposure time settings (from $100\text{\,}\mathrm{ms}$ up to
$900\text{\,}\mathrm{ms}$) and sensor gain settings (from
$0\text{\,}\mathrm{dB}$ up to $18\text{\,}\mathrm{dB}$). The low power mode
was inserted for a short time between two batches to reduce heating of the
camera. Several different exposure settings minimized the risk of
inappropriate camera settings due to unforeseen temperature conditions and
could also be used for the combination of images during the post-processing.

Figure 11 represents a sample of original images captured during the
suborbital flight, which were strongly affected by fixed pattern noise and
thermal noise. Only a few of the brightest stars can be recognized, but not
clearly due to strong noise artifacts in the images.

Camera temperature was almost constant near
$34.7\text{\,}\mathrm{\SIUnitSymbolCelsius}$ during the observation part of
the mission, as shown in Figure 10. Small periodical increases and decreases
of camera temperature were caused by the low power mode being switched on
between batches of images to avoid overheating, but the main temperature
envelope was given by the temperature of mechanical structures in the rocket
and their changes during ascending, observation, descending and the landing
phase of the mission.

Most of the fixed pattern noise was removed by the dark image subtraction,
because it is almost constant for given exposure settings and camera
temperature. Dark images were captured for all exposure settings when the
rocket was in the ascending and descending phase of the mission in a closed
instrumentation section. Results of dark image compensation are shown in
Figure 12. Nonlinear amplification and thresholding were also applied to
increase the image clarity.

(a)

(b)

(c)

Figure 11: Cut-out of the original captured images with the exposure time
$500\text{\,}\mathrm{ms}$ and gains (a) $18\text{\,}\mathrm{dB}$, (b)
$6\text{\,}\mathrm{dB}$ and (c) $0\text{\,}\mathrm{dB}$.

(a)

(b)

(c)

Figure 12: Cut-out of the images (same as in Figure 11) after fixed pattern
noise suppression by the subtraction of the dark image for the three different
exposures with constant time $500\text{\,}\mathrm{ms}$ and gains (a)
$18\text{\,}\mathrm{dB}$, (b) $6\text{\,}\mathrm{dB}$ and (c)
$0\text{\,}\mathrm{dB}$.

(a)

(b)

(c)

Figure 13: Cut-out of the images (same as in Figure 11 and Figure 12) after
spatial and temporal average filtering for the three different exposures with
constant time $500\text{\,}\mathrm{ms}$ and gains (a)
$18\text{\,}\mathrm{dB}$, (b) $6\text{\,}\mathrm{dB}$ and (c)
$0\text{\,}\mathrm{dB}$.

(a)

(b)

Figure 14: Comparison of star map in _Stellarium_ (a) and the processed images
(b). The pink circle in the top-right part of (a) and in the middle of (b) was
the target of X-ray telescopes.

Bright stars can be well recognized after removing the fixed pattern noise
artifacts, however the high amount of thermal noise is still present in the
images, covering and masking the weak stars. The thermal noise was
additionally suppressed by applying a spatial and temporal averaging filter on
the series of images captured during the attitude stabilized phase of star
observation. The results of averaging eight images in series and spatial
averaging with 3 $\times$ 3 neighboring pixels are shown in Figure 13.

Most of the thermal noise was suppressed by averaging filters and this enabled
less bright stars to be recognized (apparent magnitude 8). Some residual fixed
pattern and thermal noise is still present in the images, but stars can be
successfully identified. A comparison of the processed image with the
_Stellarium_ star map shows a good match in Figure 14.

The small and cheap camera successfully fulfilled the task in the suborbital
mission and confirmed the targeting of the X-ray telescopes. Use of these
cameras in small satellites like CubeSat requires capturing a dark image and
having a long exposure time (which requires attitude stabilization).

###  3.4 Infrared sensor array

Unfortunately, the timing of the rocket stabilization and door opening and
closing in the scientific section did not show the Earth to the FOV of the
Grid-EYE sensor, but valuable data were nevertheless obtained. As this and
similar IR arrays contain IR lenses, sensors should not have to experience
high acceleration and the conditions during rocket launch are far behind the
recommended operation limits. However, the thermal image had sharp edges
during the observation part of the mission so it indicates that the IR lenses
survived the launch conditions. During the ascending and descending phase of
the mission, the sensor’s FOV was the door and the mechanical structure of the
rocket. The thermal data from the sensor (Figure 15) indicated the heating of
the inner rocket structure from the initial 25 – 30
$\mathrm{\SIUnitSymbolCelsius}$ up to
$80\text{\,}\mathrm{\SIUnitSymbolCelsius}$, due to friction of the atmosphere
during the ascending phase, followed by a fast decrease of temperature down to
$-150\text{\,}\mathrm{\SIUnitSymbolCelsius}$ (post-processed value), after the
door opening for the phase of star observation. Closing the science section
door and reentry into the atmosphere caused a short increase of temperature of
the inner mechanical structure, followed by cooling during its slow parachute
descent and landing into the ocean.

(a)

(b)

(c)

(d)

(e)

(f)

$-150$$-125$$-100$$-75$$-50$$-25$$0$$25$$50$$75$$100$Temperature (∘C)

Figure 15: Obtained thermal images from the AMG88 IR sensor; (a) before
launch, (b) during ascending, (c) after science section door opening, (d)
after science section door closing, (e) during descending, and (f) after
landing.

(a) Pixel (2,5).

(b) Pixel (6,5).

(c) Pixel (8,3).

Figure 16: Examples of acquired temperature measurements affected by a
software overflow in numerical expression of temperatures; (a) in pixel
position (2,5), (b) in pixel position (6,5), (c) in pixel position (8,3).

The most interesting data were expected after opening the science section and
subsequent exposure of the sensor to the absence of heat from space, and to
the heat of the rocket structure, both concurrently in the FOV of the sensor,
where the sensor is exposed to a high dynamic range of measured temperatures.
However, the sensor suffered from a software overflow or sensitivity switching
in this situation and the sensor’s pixels shift numeric representation of
temperatures approximately one half of the numeric range, which leads to a
false interpretation of the measured temperatures from two complement binary
representation. This can be seen in Figure 16, where the readout temperature
of the pixels alternates between high temperature and low temperature.

The overflow behavior was compensated in post-processing based on the numeric
range of the temperature oscillations and knowing the expected temperature
values for specific locations of the rocket structure in the sensor FOV.
Compensated temperatures with the opened scientific section are shown in
Figure 17 as the mean value of four pixels in the middle of the AMG88 Grid-EYE
sensor.

Figure 17: Temperature in the Field of View after compensation of numeric
representation overflow — plotted as mean value of four central pixels.

The tested IR array sensor AMG88 is usable for simple and non-critical
application in similar missions, even with the instability of numeric
representation of temperatures. The temperature changes in the FOV are
detectable, only the absolute value of temperature is unreliable. The sensor
can be used, for example, as a simple detector for the release of the
deployable parts, covering the sensor in the stowed state and opening the
sensor FOV to space in the deployed state. Using this sensor as an Earth
horizon detector in the ADCS of small satellites is not suitable with this
particular sensor due to its instability of numeric representation of
temperatures.

###  3.5 X-ray telescopes

Both X-ray detectors were autonomously operated by a custom-built software
[22] in ROS. Data from both detectors were acquired using as a continuous
stream of $1\text{\,}\mathrm{s}$ acquisitions that started before launch.

(a) 1D LE integral image.

(b) 2D LE integral image.

Figure 18: Integral image during the entire duration rocket’s flight; (a) 1D
LE telescope, (b) 2D LE telescope.

(a) Integral image.

(b) Row sums.

(c) Column sums.

(d) Cluster size histogram.

Figure 19: Data obtained from the 1D telescope taken while pointing at the
Vela pulsar: (a) integral images with particle tracks labeled according to
their size (b) profile of summation of image rows (1 – 4$\mathrm{px}$) (c)
profile of the summation of image columns (1 – 4 $\mathrm{px}$) (d) bar graph
representing the frequency of events according to their size.

(a) Integral image.

(b) Row sums.

(c) Column sums.

(d) Cluster size histogram.

Figure 20: Data obtained from the 2D telescope taken during the pointing time:
(a) integral images with tracks labeled according to their size (b) profile of
row summation of the image (1 – 4 $\mathrm{px}$) (c) profile of the column
summation of the image (1 – 4 $\mathrm{px}$) (d) bar graph representing the
frequency of events according to their size.

Both X-ray telescopes (1D and 2D) took images during the entire duration of
the rocket’s flight (Figure 18). Figures 19(a), and 20(a) represent filtered
integral images from both telescopes while the rocket pointing at the Vela
pulsar. Thanks to the data sensing system, it was possible to recognize and
filter out individual events based on the size of their tracks or their
dissipated energy in the detection layer. The results of track size
classification are shown in color. The main areas of interest were gamma
particles (photons with a range of energies between $3\text{\,}\mathrm{keV}$
to $10\text{\,}\mathrm{keV}$), which in the case of the Si detector, leave
mostly single-pixel events on the detector [39]. However, photons can strike
multiple pixels due to charge-sharing among the neighboring pixels. During the
evaluation of the measurement, we considered 1 – 4 $\mathrm{px}$ events as
gamma photons.

The filtered images, formed only by impacts of gamma particles, were evaluated
as horizontal and vertical projections. Sums of individual rows and columns
were plotted as a line graphs, which effectively produces integral images of
256 line-detectors. This technique makes it easier to detect the position the
focus line (1D) and cross (2D), since the focus is always aligned with the
detector rows and columns. Figure 19 shows the processed and filter data from
the 1D telescope, Figure 20 depicts the data from the 2D telescope.

As it can be seen in the images, a small number of particles was collected on
both detectors. Despite the approx. 43 (1D) respectively 35 (2D) X-ray gamma
events taken during the observation time, no apparent focal point/line or
preferential event positions was observed. The total supposed X-ray counts
during the whole observation were 5 photons for the 1D optics and 27 photons
for the 2D optics. The results show no statistical tendency to form a focus
line by the 1D telescope or focus point by the 2D telescope. In fact, majority
of the measured signal originates from ambient background radiation, which was
recorded equally by both detectors.

Another reasons for such a small photon count yield might be a water
condensation on the optics surfaces, as suggested by [17]. Water vapor
condensing on the surfaces of the LE optics foils would diminish the
intensity, especially for soft X-rays. However, the real reason is unclear as
not enough data supports this hypothesis for the REX payload. Importantly for
the TRL evaluation, both the optics and the electronics have survived the
experiment and all the subsystems were fully operational.

##  4 Conclusion

The Rocket Experiment was one of two payloads on the suborbital rocket
campaign launched from the Marshall Islands on April 4, 2018. The mission goal
was to demonstrate the space performance of X-ray multifoil telescopes in
combination with an uncooled pixel detector Timepix and the verification of
the Technology Readiness Level. The payload contained a number of other
sensors and devices, for observation in the visible, infrared, and X-ray
spectra. The payload system was operated by the custom-made automatic software
Rospix, developed specifically for this experiment. Due to the low intensity
of the source, the measurements did not yield sufficient results for X-ray
observations of the Vela Supernova Remnant, however, we can report that all
instruments in the proposed system worked successfully and without problems
during the experiment. Based on the obtained results, the functionality and
the possibility of using the proposed system for similar upcoming experiments
was confirmed.

## Acknowledgements

The team of authors of this article would like to thank the whole team of
colleagues who participated in this project, not only: Ondřej Petr (Rigaku
Innovative Technologies s.r.o.), Robert Filgas (Czech Technical University in
Prague), Richard Linhart (University of West Bohemia) and Pavel Fiala
(University of West Bohemia) and many others without which this project did
not arise.

This project was performed in collaboration with several universities,
companies and scientific institutions in the Czech Republic and abroad.
Namely: Czech Technical University in Prague (CZE) - Institute of Experimental
and Applied Physics and Faculty of Electrical Engineering, University of West
Bohemia (CZE), Rigaku Innovative Technologies s.r.o. (CZE), Czech Aerospace
Research Centre a.s. (CZE), PANTER facility of the Max Planck Institute for
Extraterrestrial Physics (DEU), Pennsylvania State University (USA),
University of Colorado at Boulder (USA).

Many thanks for the willingness, help and pieces of advice belonging to
namely: D. Burrows with his team (PSU), W. Cash with teams (CU Boulder), V.
Burwitz with the team (PANTER).

The project was supported by the Technology Agency of the Czech Republic under
grants no: TA03011329, TA04011295, the Grant Agency of the Czech Republic
under grant no. 18-10088Y, the NASA Astrophysics Research and Analysis grant
NNX17AC88G, the Grant Agency of the Czech Technical University in Prague no.
SGS18/186/OHK3/3T/13, and by the Ministry of Education and Youth of the Czech
Republic, project no. LTAUSA18094. This work was done on behalf of Medipix2
collaboration.

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  *[REX]: Rocket Experiment
  *[LE]: Lobster-Eye
  *[FOV]: Field of View
  *[TRL]: Technology Readiness Level
  *[SNR]: Supernova Remnant
  *[MFO]: Multi-Foil Optics
  *[CMOS]: Complementary Metal–Oxide–Semiconductor
  *[PSU]: Pennsylvania State University
  *[WRXR]: Water Recovery X-ray Rocket
  *[1D]: one-dimensional
  *[2D]: two-dimensional
  *[IR]: infrared
  *[IMU]: Inertial Measurement Unit
  *[ROS]: Robot Operating System
  *[CTU]: Czech Technical University in Prague
  *[RITE]: Rigaku Innovative Technologies
  *[PSF]: Point Spread Function
  *[ASIC]: Application-Specific Integrated Circuit
  *[ADC]: Analog-to-Digital Converter
  *[ADCS]: Attitude Determination and Control System
  *[MEMS]: Micro-electro-mechanical System