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# RAAD: LIGHT-1 CubeSat’s Payload for the Detection of Terrestrial Gamma-Ray
Flashes

A. Di Giovanni 11footnotetext: Corresponding Author A. Di Giovanni:
<EMAIL_ADDRESS>F. Arneodo22footnotetext: Corresponding Author F.
Arneodo<EMAIL_ADDRESS>A. Al Qasim  H. Alblooshi  F. AlKhouri  L.
Alkindi  A. AlMannei  M. L. Benabderrahmane  G. Bruno  V. Conicella  O. Fawwaz
G. Franchi  S. Kalos  P. Oikonomou  L. Perillo  C. Pittori  M. S. Roberts  R.
Torres

###### Abstract

The Rapid Acquisition Atmospheric Detector (RAAD), onboard the LIGHT-1 3U
CubeSat, detects photons between hard X-rays and soft gamma-rays, in order to
identify and characterize Terrestrial Gamma Ray Flashes (TGFs). Three detector
configurations are tested, making use of Cerium Bromide and Lanthanum
BromoChloride scintillating crystals coupled to photomultiplier tubes or
Multi-Pixel Photon Counters, in order to identify the optimal combination for
TGF detection. High timing resolution, a short trigger window, and the short
decay time of its electronics allow RAAD to perform accurate measurements of
prompt, transient events. Here we describe the overview of the detection
concept, the development of the front-end acquisition electronics, as well as
the ground testing and simulation the payload underwent prior to its launch on
December 21st, 2021. We further present an analysis of the detector’s in-orbit
system behavior and some preliminary results.

##  1 Introduction

Terrestrial Gamma Ray Flashes (TGFs) are upward directed, highly luminous
bursts of photons, with durations of less than 1 ms and energies from 10 keV
reaching up to several tens of MeV[1, 2, 3, 4, 5, 6]. They are produced in the
high electric fields naturally occurring within thunderstorms, at altitudes of
10 - 15 km, when electrons are accelerated to relativistic speeds and produce
photons through bremsstrahlung [7, 8, 9, 5]. It is estimated that a sum of
more than 400,000 TGFs are produced annually [1], primarily concentrated
around the equator [3].

TGFs were first discovered by Fishman et al. using data obtained from the
Burst and Transient Source Experiment (BATSE) on NASA’s Compton Gamma Ray
Observatory (CGRO) in 1994 [3], and were found to be associated with lightning
activity. BATSE was built to detect gamma-rays from celestial sources and thus
was sensitive to energies of $\geq$20 keV. However, BATSE’s sampling rate of
$64\,$ms was longer than the typical sub-millisecond duration of a TGF,
resulting in very few detections [3, 5].

NASA’s Reuven Ramaty High Energy Solar Spectroscopic Imager (RHESSI),
originally built for studying solar flares, was able to detect TGFs as well.
However, RHESSI’s Germanium detectors were sensitive to photons with energies
up to only 18 MeV. [10]. Regardless, RHESSI’s detections proved that TGFs can
be detected by satellites operating il LEO (Low Earth Orbit) [11, 12]. The
“Astro-Rivelatore Gamma a Immagini Leggero” (AGILE) and the Gamma-Ray Burst
Monitor (GBM) onboard NASA’s Fermi Gamma-ray Space Telescope were able to
detect the full energy spectrum of TGFs [13, 4, 14, 15, 16, 17, 18]. As the
aforementioned missions were optimized to detect fainter phenomena of longer
duration, their detectors suffered from pileup and saturation when observing
high luminosity and energetic TGFs.

Thus far, TGFs have been primarily studied as by-products of missions designed
to detect other phenomena, thus creating an opportunity for specialized
missions. One such mission is European Space Agency’s Atmosphere-Space
Interactions Monitor (ASIM), installed in 2018 on the ISS, which managed to
confirm the hypothesized TGF models originating from previous studies by
combining observations between $50\,$keV - $30\,$MeV with optical data [19,
20].

Even though TGFs’ bright and brief nature made them hard to characterise with
the non-specialised equipment of the initial missions, it also made their
detection possible using detectors small enough to fit within CubeSats (i.e.,
pico-satellites developed at reduced costs). Thus we proposed the Rapid
Acquisition Atmospheric Detector (RAAD), an instrument specifically designed
for TGF detection, as the payload of the LIGHT-1 CubeSat mission. Table 1
shows the performance characteristics of LIGHT-1 compared to previous
missions. The RAAD acronym was also chosen for its similarity to the Arabic
word "Ra’ad" ( رعد> ) which means thunder.

In this paper, we provide a detailed description of the payload electronics,
ground tests and simulations, as well as some preliminary in-orbit data. A
prototype of RAAD was built and tested at NYUAD, providing proof of concept.
Details on the prototype can be found in [21]. Details on the CubeSat’s bus
and subsystems can be found in [22].

Mission |  Trigger Window (ms) |  Resolution ($\mu$s) |  Dead Time Per Event ($\mu$s) |  Sensitivity (MeV)
---|---|---|---|---
BATSE |  64 |  2 |  6 |  0.02 $\div$ 2
RHESSI |  None |  0.95 |  8 |  0.003 $\div$ 17
AGILE |  0.293 |  1 |  65 |  0.3 $\div$ 100
FERMI |  16 |  2 |  2.6 |  0.001 $\div$ 40
LIGHT-1 |  0.5 |  0.5 |  0.04 |  0.02 $\div$ 3
Table 1: Characteristics of LIGHT-1 compared with previous missions: BATSE
[3], RHESSI [23], AGILE [24, 13], and FERMI [16].

##  2 The LIGHT-1 Mission

###  2.1 Mission Concept

Absorption by the atmosphere can suppress TGFs, therefore, solely measuring
the gamma-ray emission would not provide complete information about their
characteristics [25]. As a result, a combined measurement is, in general,
preferred over a single detection channel. Combining atmospheric gamma-ray
detection with X and gamma-ray surveys or lightning radar measurements could
provide insightful information on the production mechanisms of TGFs [26, 27].
Under this premise, a LEO operated gamma-ray detector, together with
correlated ground and space observations and measurements, would be the ideal
probe to pursue a TGF science program.

LIGHT-1 is a 3U CubeSat built primarily to study TGFs. Yet, LIGHT-1’s mission
extends to measuring orbital radiation and space-qualifying different
technologies for the detection of prompt, highly energetic, and intense
emissions typical of transient events. For size reference, one U of a CubeSat
consists of a cube of about
$\mathrm{10~{}cm~{}\times~{}10~{}cm~{}\times~{}10~{}cm}$. A little less than
two units of LIGHT-1 is dedicated to RAAD, while the rest is dedicated to the
subsystems of the CubeSat (reaction wheels, onboard computer, attitude control
system, electrical power system, etc.). A review on operating gamma-ray
detectors onboard CubeSats is available in [28].

RAAD utilizes three types of detectors, by coupling two types of photosensors
to two types of scintillating crystals. In fact, the LIGHT-1 mission is also
meant to conduct a direct comparison between different detector configurations
to find which is best suited for TGF detection. LIGHT-1 is the first mission
to conduct a direct comparison between Cerium Bromide and Lanthanum
BromoChloride scintillating crystals, coupled to either photomultiplier tubes
or silicon photomultipliers.

###  2.2 Mission Schedule

Figure 1: The life of the LIGHT-1 mission. Apogee and perigee altitudes of the
satellite are shown in red and blue respectively. The dotted black line
starting on the SpaceX Launch date (2021-12-21) until the deployment date
(2022-02-03) is the altitude of the International Space Station where LIGHT-1
was stowed prior to the start of the mission. Different colors denote the
different operation regimes: Launch and Early Orbit Phase (LEOP), Payload
Commissioning, Science Run, and de-orbit.

The LIGHT-1 CubeSat was launched on the 21st of December, 2021, onboard a
Space-X Falcon9/Dragon from the Kennedy Space Center, directed at the
International Space Station (ISS). The handling of LIGHT-1 for the launch and
subsequent deployment was taken care by the Japan Aerospace Exploration Agency
(JAXA). Once on the ISS, it was deployed via the Japanese Experiment Module
(JEM) on the 3rd of February, 2022, in a 51.6°orbit at an initial altitude of
$\mathrm{408~{}km}$.

Figure 1 shows the operation regimes of the LIGHT-1 mission as a function of
time. During the Launch and Early Orbit Phase (LEOP), the satellite bus
developed by NanoAvionics was fully tested and optimized for the science
program operations. The payload was powered on for the completion of one full
orbit on the 16th of March, 2022 to check its vital parameters. The
commissioning of the payload began on the 6th of April, 2022. On the 25th of
May, 2022 LIGHT-1 entered the Science Run mode. Communications with LIGHT-1
stopped on January 18th, 2023, marking the start of the de-orbiting phase and
the end of the mission.

##  3 Scientific Payload Specifications

Parameter |  Design value |  Actual Value
---|---|---
Mass [kg] |  2.05 $\pm$ 0.07 |  1.981 $\pm$ 0.001
Average Power Consumption [W] |  $<$ 5.9 W |  $<$ 4.8 W
Data Downlink [MB/24 h] |  40 |  $\sim$ 40
Duty Cycle [%] |  50 |  $\sim$ 50
Life Time [months] |  $\sim$ 6 |  $\sim$ 10
Temperature operative range [${}^{\circ}\mathrm{C}$] |  0 $\div$ 45 |  10$\div$ 40
Table 2: Payload specification design in comparison with the values measured
on the flight model.

RAAD is designed to resolve events hundreds of ns apart, measure the energy
deposited, and assign a timestamp for comparison with TGF and lightning
catalogs generated by other experiments.

###  3.1 Detector Structure

RAAD consists of two detectors, different in size, fitting 1 U and 0.7 U of
the spacecraft, respectively. The detector structure is shown in Figure 2.

Figure 2: An explosion view of the MPPC CAD model. The PMT payload structure
is identical except for the size of the photosensors.

The smaller one, the MPPC payload, is equipped with four S13361-6050AE-04
Multi-Pixel Photon Counters (MPPCs) manufactured by Hamamatsu Photonics and
coupled to a 4-channel Low Background Cerium Bromide (CeBr3(LB)) crystal
array, manufactured by Scionix and shown in Figure 3. The larger detector, the
PMT payload, is equipped with four R11265-200 photomultiplier tubes (by
Hamamatsu Photonics) and coupled to one array of four scintillating crystals
organized in one pair of CeBr3(LB) and one pair of Lanthanum BromoChloride
(LBC), both manufactured by Scionix. We refer to each photosensor array
coupled with the corresponding scintillating crystals as the detection target
of RAAD. The differences between the two photosensors are shown in Table 3.

Parameter |  R11265-200 |  S13361-6050AE-04
---|---|---
Type of photosensor |  PMT |  MPPC
Dimensions ($L\cdot D\cdot H$) [$\mathrm{mm^{3}}$] |  $26\cdot 26\cdot 19$ |  $25\cdot 25\cdot 1.4$
Mass [g] |  24 |  2
Max Sensitivity [nm] |  400 |  450
Quantum Efficiency - Photon Detection Efficiency [%] |  43 |  40
Operating Voltage [V] |  900 |  55
Gain at working point |  $\sim 10^{6}\div 10^{7}$ |  $\sim 10^{6}\div 10^{7}$
Dark Counting Rate at working point, 300 K [Hz] |  < 1 |  $>10^{7}$
Operating Temperature [K] |  240 - 320 |  250 - 330
Table 3: Characteristics of the photosensors used in RAAD

The performances of the crystals at different energies, shown in Table 4, make
them complementary. The intrinsic activity of LBC, while on one hand an
obvious nuisance, on the other hand, may provide an embedded calibration tool.
The high dark counting rate of the MPPC does not represent a problem when
triggering at $\mathrm{\gtrsim 20\,keV}$ (see 4.2).

Each scintillating crystal unit is
$\mathrm{23~{}mm~{}\times~{}23~{}mm~{}\times~{}45~{}mm}$ in dimensions, with
the whole scintillating crystal arrays of
$\mathrm{60~{}mm~{}\times~{}60~{}mm\times~{}48~{}mm}$ in dimensions and
$\mathrm{615~{}g}$ and $\mathrm{595~{}g}$ in mass, for the
$\mathrm{CeBr_{3}(LB)}$ and LBC respectively.

Figure 3: Top view of the CeBr3(LB) crystal array used in the MPPC payload,
embedded in the VETO system and its readout electronics. The view of the PMT
payload is identical.

The four scintillating crystals are separated by Polytetrafluoroethylene
(PTFE) to avoid optical cross-talk. In order to prevent the effects of water
vapor contamination during the assembling phase, the hygroscopic crystal array
is housed in an airtight container. Five sides are covered with aluminum,
while on the sixth side, a 2 mm thick layer of fused silica is used to
optically connect the photosensor array to the crystals.

Parameter |  CeBr3(LB) |  LBC
---|---|---
Density [$\mathrm{g\cdot cm^{-3}}$] |  5.1 |  4.9
Hygroscopic |  Yes |  Yes
Light Emission Peak [nm] |  370 |  380
Energy Resolution at 122 keV [%] |  10 |  7
Energy Resolution at 662 keV [%] |  4 |  3
Decay Time [ns] |  20 |  35
Intrinsic Activity [$\mathrm{Bq\cdot cm^{-3}}$] |  $<0.01$ |  $\sim$ 1
Table 4: Characteristics of the scintillating crystals used in RAAD.

The front-end electronics are directly coupled to the back of the photosensor
array as seen in Figure 4.

Figure 4: The two photosensor arrays from the payload of the LIGHT-1 mission.
(left) The four R11265-200 photomultiplier tubes used in the PMT payload.
(right) The four S13361-6050AE-04 Multi-Pixel Photon Counters used in the MPPC
payload. Both sensors have been manufactured by Hamamatsu Photonics.

A detailed study and characterization of the detection concept is reported in
[21].

A VETO system surrounds the detection target on four sides. Its purpose is to
partially identify and suppress the charged particle induced background. It
consists of eight independent units, each composed of one 5 mm thick plastic
scintillator tile with an embedded wavelength shifter fiber and read out at
one edge by a SMD Silicon photomultiplier (SiPM) manufactured by AdvanSiD srl
(ASD-NUV1C-P-40). The CAD model of the VETO system is shown in Figure 2.

The detection target along with the photosensor array and related electronics
are placed inside an aluminum enclosure designed to be mechanically coupled to
the CubeSat spacecraft. The final assembly of the PMT payload is shown in
Figure 5.

Figure 5: The fully assembled PMT payload. The view of the MPPC payload is
identical.

The effect of the aluminum walls has been evaluated by a Geant4 simulation,
fully presented in Section 4.2, indicating a hardware detection threshold for
gamma rays of about 20 keV.

An aerospace-grade silicon resin (Momentive RTV615) was used for filling the
voids in the assembly in order to mitigate the effect of vibration on the
structure and to provide effective electrical insulation for the high voltage
($\mathrm{-800~{}V\div-700~{}V}$) required to operate the payload. Overall,
the multiple enclosing layers comply with the safety requirements of the
mission, mitigating the risk of fragmentation.

###  3.2 Payload Readout and Control Electronics

RAAD electronics consist of three main components.

  1. 1.

VETO readout electronics

  2. 2.

Photosensor power supply

  3. 3.

Front-end and Controller boards

Figure 6: Block diagram of the RAAD electronics

A block diagram for the electronics is shown in Figure 6 and its main
characteristics are summarized in Table 5.

Parameter |  Design value
---|---
Average Power Consumption [W] |  $<$ 5.9
Mass [g] |  58 (MPPC), 66 (PMT)
ADC resolution range [bits] |  10 $\div$ 16
Components |  SMD, all COTS, Automotive grade
Operating modes |  noise, default, science, custom, safe
Table 5: The main design characteristics of the RAAD electronics

The power supply electronics of the MPPC and PMT payloads are shown in Figure
7.

Figure 7: The custom made power supply electronics of the LIGHT-1 PMT payload
which is used to bias the photomultiplier tubes at variable gains, depending
on the science case.

With the single exception of the photosensor sockets, no connectors have been
used in the LIGHT-1 payload to maximize compactness and reliability.

###  3.3 Firmware Overview

The onboard firmware is designed in order to assign all data collected to
specific categories (i.e. gamma and charged particle-induced events, TGF
events, and orbit monitoring data).

The detection strategy can be summarised in a series of steps listed below (
see Figure 8).

  1. 1.

Particle interaction and signal generation in the detector;

  2. 2.

the charge amplifier integrates the signal. Its output is sampled by an ADC
(16 bits, 10 MHz) and passed through a matched filter [28] to extract the peak
charge value;

  3. 3.

if the charge value is larger than the detection threshold, an event timestamp
is assigned and the event is sent to the spacecraft bus.

Figure 8: A schematic description of the detection pipeline of LIGHT-1. A
particle interaction triggers a signal on the photosensor that is then
amplified and digitized. If the signal is above a software set threshold it is
recorded and sorted to the corresponding buffer.

The payload data generated are organized into four buffers, each designed to
store different types of events as shown in Table 6.

Data Buffer |  Description
---|---
NonVeto |  All the events in which there is no VETO flag. The ADC resolution is lowered to 10 bits, timestamp at 1 ms resolution
Veto |  All the events in which there is at least one (out of eight) VETO flag active. The ADC resolution is lowered to 14 bits, timestamp at 100 $\mu s$ resolution
TGF |  All the events falling into a coincidence window of 500 $\mu s$ in which there is no VETO flag. The ADC resolution is nominal (16 bits), 500 ns timestamp resolution
Orbit |  Payload housekeeping events collected once every 20 s including the payload temperature, the particle rate of each channel and the VETO, the detector operating voltage, the payload operating configuration (e.g. noise, default, flight mode), and all the working parameters required to run diagnostic checks
Table 6: Data buffer protocol used for the scientific data of the LIGHT-1
mission.

The science data is transmitted to the ground using an S-Band system during
daily passes above ground stations dedicated to the mission [22]. Each contact
(up 6 times per day) can last up to 10 minutes. Spacecraft telemetry, which
includes housekeeping data, operational commands, and payload scripts, is
transmitted using a UHF transceiver.

##  4 Pre-Flight Simulations and Tests

In order to space qualify RAAD, all the payload components and the payload as
a whole have been subjected to several environmental and functionality tests,
in order to assess the surviving capability to the various regimes typical of
the mission (launch phase, thermal stress, radiation, etc …).

###  4.1 Stress Analysis

Prior to conducting vibrations tests on the physical assembly of the payload,
a stress analysis was conducted to study the effect of the launch conditions
on the structure. In particular, a Finite Element Analysis was conducted in
order to calculate the frequency and peak amplitude of the first 20
vibrational modes.

The large computational requirement for such analyses forced a geometric
simplification of the 3D models used. Yet the simplification was within a
reasonable range so as to not contribute more than 10% of the payload’s total
mass.

The general requirement for space qualifying such equipment to survive launch
is for the natural modes to be above 100 Hz (see [29]). Through the
simulation, we verified that our design fulfills this requirement as all
normal modes are between 1200 Hz and 2000 Hz.

###  4.2 Geant4 Simulations

Using the Geant4 framework a C++ particle physics simulation was developed to
evaluate the efficiency of the detectors, as well as the background energy
deposition on the crystals and the veto caused by trapped charged particles in
the earth’s magnetic field. The detector geometry was simplified to increase
performance. The simulation code and analysis can be found in [30].

We were able to estimate the hardware threshold of the detector by sending
photons with energies ranging from $1\ keV$ to $1\ MeV$ and recording the
energy deposited on each component of the detector. Fig. 9 shows the ratio of
the energy deposited to the crystals over the incoming energy as a function of
the incoming photon energy. We see that we start detecting photons at roughly
$20\ keV$.

Figure 9: Geant4 hardware threshold estimation. Single photons were sent to
the detector with equiprobable energy ranging from $1\ keV$ to $1\ MeV$. A
histogram of the ratio of energy deposited on the crystals to incoming
particle energy is plotted as a function of incoming energy. Brighter colors
correspond to higher probability of observing the ratio. For example, most
particles with energy 100 keV deposit around 90% of their energy in the
crystals, while almost none of them deposit between 30% and 60%. We can
conclude that particles start depositing energy at $20\ keV$ and that they
deposit roughly 100% of their energy in the crystals.

Using the European Space Agency’s SPace ENVironment Information System
(SPENVIS) [31] the flux and energy spectrum for trapped protons and electrons
along the orbit of LIGHT-1 were estimated, providing the input for a
Geant4-based simulation of the detectors’ background spectrum. The energy
deposition on the crystals vs the original particle energy spectrum is shown
in Fig. 10 while a more detailed comparison between the energy deposited on
the veto vs the crystals is shown in Fig. 11.

Figure 10: Simulated trapped charged particles energy spectra along the orbit
of LIGHT-1. In pink and blue the incoming energy spectrum of electrons and
protons is shown. In lighter colors, the energy spectrum deposited on the
detector channels is shown. Figure 11: Energy of trapped charged particles
deposited on the detector and the veto as a function of the incoming particle
energy.

###  4.3 Vibration Tests

Vibration tests are typically conducted on fully assembled CubeSats to test
their ability to survive launch. However, since our payload made use of quartz
windows and crystals, JAXA required additional vibration tests to be carried
out on these components in isolation, prior to their integration into the main
assembly.

The vibration tests were conducted at New York University Abu Dhabi (NYUAD). A
permanent magnet shaker was used to emulate the vibrational profile of the
launch as well as other parameters specified by JAXA. The testing procedure is
outlined below:

  1. 1.

A 20 Hz - 2 kHz sweep to identify the component’s normal modes.

  2. 2.

A 7s test using a frequency sweep designed to simulate the launch conditions
of a possible transport vehicle as provided by JAXA.

  3. 3.

A 20 Hz - 2 kHz sweep to compare the component’s normal modes after exposure
to launch conditions.

This test was carried out to simulate three spacecraft candidates, namely H-II
Transfer Vehicle (HTV), Cygnus NG, and SpaceX Dragon. The latter became the
chosen transport vehicle for LIGHT-1. Both crystal arrays and a PMT unit were
subject to these tests, all of which survived without mechanical failures,
ensuring that the RAAD components were launch-safe.

###  4.4 Thermal-Vacuum Tests

Figure 12: This figure shows the last four cycles of the thermal vacuum test
conducted in Yah-Sat Lab’s thermal vacuum chamber during one of the CeBr3
crystal array tests.

Placing four crystals in the same case was a solution to the limited space
problem that exists in CubeSat missions, but a concept that was not tested for
space missions. Thus, it was crucial to conduct thermal vacuum tests on the
crystal arrays to determine their survivability in the space environment. Two
types of thermal vacuum tests were carried out on the crystal arrays:

  1. 1.

One Day Tests: The arrays were left in the thermal vacuum chamber for
approximately 24 hours at a set temperature of 40 oC and a set pressure of 1
$\mu$Torr.

  2. 2.

Three Days Tests: The arrays were left for approximately 72 hours at a set
pressure of 1 $\mu$Torr with a thermal cycle where the temperature oscillates
every five hours between -40 oC and 50 oC, simulating the conditions faced in
space by satellites. An example of this cycle can be found in figure 12.

In both thermal vacuum tests the arrays were placed on top of an insulator
material, to prevent the lower side of the arrays from heating up faster than
the rest of the array due to their contact with the chamber’s inner metal
base. These tests helped in reevaluating the structure of the arrays so that
they could be optimized for space flight. In particular, a manufacturing error
was uncovered that led to condensation inside one of the arrays during the
thermal cycle, necessitating its repair by the manufacturer. These tests were
carried out both at YahSat Lab in Khalifa University and at NYUAD.

##  5 On-Flight Perfomance

The performance and health of the payload have been evaluated continuously
throughout the mission. In this section, we relate some of the experimental
data observed in orbit to the parameters we have used to evaluate the
operating condition of the LIGHT-1 payload.

###  5.1 Instrument Health

Throughout the mission, we have monitored the health of the hardware through
onboard sensors. In particular, the Texas Instruments LM71 temperature sensor
embedded in each of the readout electronics boards measured the payload
temperature over the lifespan of the mission. The resulting measurements from
the commissioning of the payload until de-orbit are shown in Fig. 13.

Figure 13: Temperature of the PMT and MPPC Payloads from the Payload
Commissioning (2022-04) until the de-orbit (2023-01) of LIGHT-1. The
temperatures of the PMT and MPPC detectors are shown in red and blue
respectively (after smoothing by a Savitzky-Golay with a window length of 1
day). The unfiltered temperature is overlaid with lighter colors. Using two
insets we highlight the temperature increase during two different duty cycles.
As can be seen from the two insets of randomly picked duty cycles, the
temperature variation is high and difficult to predict as it varies from orbit
to orbit.

The payload’s temperature data during orbit is important to correct for
detector gain drift [21], and for adjusting the operations duty cycle in
flight to avoid overheating. As can be seen in the insets of Fig. 13, which
represent the temperature recorded while the payload was operational, during a
single orbit the temperature rises significantly when the payload is
operating. This occurs due to the power dissipated by the electronics. Already
in its design phase, the payload mission concept has been tailored around a
reduced 50 $\%$ duty cycle in order to meet the mission requirements in terms
of the power budget. From the commissioning to the de-orbit phase, the
electronics have been subjected to 1861 power cycles. Under these conditions
and for the entire mission, no emergency procedure has been triggered due to
temperature failure.

#### Pulse Per Second Signal Loss.

By design, a key feature of the RAAD electronics is the ability of assigning a
sub-microsecond time stamping to each acquired event. To be compliant with the
power constraints of the mission, since the design phase, the use of fast
electronics capable of coping with the typical time response of MPPCs and PMTs
(order of $\mathrm{\sim 1~{}GS/s}$) could not be considered.

The solution implemented on the LIGHT-1 timing architecture utilizes a
distributed disciplined clock, controlled by a Phase Locked Loop (PLL).

The PLL generates an high frequency output ($\mathrm{10~{}MHz}$) from a low
frequency reference ($\mathrm{1~{}Hz}$). The RAAD timing circuit concept is
based on a PLL using the $\mathrm{1~{}Hz}$ Pulse-Per-Second (PPS) hardware
signal obtained from the spacecraft GPS receiver coupled to a
$\mathrm{10~{}MHz}$ oscillator.

However, an unidentified failure made the PPS signal unavailable. As a
consequence, the capability to calculate and assign the timestamp to events on
the payload data buffers was compromised from the beginning of the LEOP Phase
until the Science Run began (see Fig. 1).

To overcome this issue, a software patch has been implemented during the
Payload Commissioning phase (see Fig. 1). The patch uses the onboard
computer’s signal to provide a, less accurate, PPS signal digitally. We used
further reconstruction techniques on the ground after obtaining the data in
order to retrieve the original accuracy.

#### Seasonal Temperature Variation.

Figure 14: Temperature of the PMT payload as a function of latitude during
1${}^{\text{st}}$ \- 23${}^{\text{rd}}$ of October 2022. The shaded region
shows the probability density of the temperature over the latitude, while the
red, dotted line is the maximum of this distribution. We observe the
characteristic hotter temperatures for the southern hemisphere during October.

We further verified the health of the onboard electronics by tracking the
temperature of the satellite during October as a function of latitude. In Fig.
14 the dotted line shows the most probable temperature measured by the
satellite as a function of latitude. In the plot, we observe the expected
seasonal variation, namely hot temperatures in the southern hemisphere due to
increased sunlight during October, and, conversely, colder temperatures in the
northern hemisphere. Such results were used to verify the integrity of our
reconstructed data.

##  6 Conclusions

In this paper, we describe RAAD, the scientific payload of the LIGHT-1 CubeSat
mission, designed to detect fast (< 1ms) X and gnmma-ray transients with two
detectors fitting in 1.7 CubeSat units. The payload underwent rigorous
physical testing and simulation in order to space-quality the instrument. In
particular, the thermal vacuum and vibration tests of the individual
components and the assembly certified RAAD according to JAXAs specifications
for surviving typical launch conditions for such instrumentation. Space
qualification was further validated by carrying out a modal analysis of the
apparatus, showing vibrational modes within the approved range for
withstanding launch conditions. The expected trapped charged particle flux
along the orbit of the detector was estimated using data from the European
Space Agency’s SPace ENVironment Information System (SPENVIS) [31] allowing us
to carry out an exhaustive Geant4 particle physics simulation to predict the
expected background and optimize the detection threshold accordingly as well
as the geometry and optical insulation of the instrument.

The LIGHT-1 satellite was launched from the Kennedy Space Center on December
21st, 2021, on a SpaceX rocket, which docked at the International Space
Station the day after. It was deployed on February 3rd, 2022. Contact was lost
on January 14, 2023. While the analysis of the data is ongoing, the
housekeeping data presented show that, with the exception of the failure that
led to the PPS loss, RAAD withstood the launch stress as expected, and the
detectors operated throughout the mission.

## Acknowledgments

We gratefully acknowledge the support of the UAE Space Agency through the 2018
MiniSat Competition, and the NYUAD Kawader program for supporting one of the
authors (L. AlKindi). Special thanks to Sebastien Celestin for providing TGF
models. We also thank Khalifa University and NSSA for funding their master’s
students to work on the CubeSat’s Bus system design. Finally, we express our
gratitude to the NYUAD Core Technology Platforms for their invaluable
assistance and particularly to the machine shop team for the realization of
the aluminum enclosures of the detection targets.

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