Since having set a new standard for small, low cost, technology demonstrating satellites, the Delft University of Technology continues development on its PocketQube satellite class to make sure everyone can access space through miniaturised technology. Through its students, the design, testing and integration of various satellite subsystems can be achieved, such as the reaction wheel or magnetorquer. A design of a momentum bias wheel fit for a PocketQube pico-satellite has not yet been achieved, for which this thesis is dedicated. Finding out whether designing and testing a momentum bias wheel using commercial off the shelf or self-manufactured parts is feasible and whether this concept at PocketQube scale is competitive with other attitude actuators. By realising that a reaction wheel delivers a pointing accuracy below 1 degrees and a magnetorquer delivers one above 5 degrees, a perfect range appears in which the momentum bias wheel can operate to fill this pointing accuracy gap. The momentum bias wheel is thus designed to be able to limit the effect of disturbances; internal and external; to a pointing accuracy of 1 to 5 degrees in a nadir-pointing attitude scheme. By setting design requirements based on a statistically determined maximum aerodynamic disturbance torque over one orbit endured during the mission, deriving requirements from the
PocketQube standard, and using the other attitude actuators to derive competitive requirements, a momentum bias wheel design can be created that can be competitive to the other actuators, perform under required environments and achieve the set out pointing accuracy. The momentum bias wheel must be able to attain and maintain an angular momentum of 6.05 𝑚𝑁 𝑚𝑠 due to the aerodynamics disturbance torque estimated at 83.3 𝜇𝑁 𝑚 and must maintain a stability throughout its life of 0.180% as to stay within the desired pointing accuracy, while maximising itself to a power draw of 180 𝑚𝑊 and weighing no more than 41.4 grams. The design is ultimately derived through three trade-off scenarios; for the motor, the flywheel and mounting of the components to each other. As power usage, size and mass are preliminary set requirement due to the fact of the undone system engineering for the next PocketQube mission, these criteria will be approached and are treated as loose requirements. Having settled on a design utilising a Faulhaber vacuum proof 1509B motor and SC1801P speed controller, together with a bronze C67500 alloy lathed flywheel, mounted together utilising two PQ9 PCB’s and an additively manufactured motor mount and flywheel reinforcement, which unfortunately did not meet the size and mass requirement through prototyping, testing the built prototype through operational, micro-vibration, shaker and vacuum tests were the final steps. The operational modes and analysis thereof for only the motor and the complete momentum bias wheel system confirmed that the momentum bias wheel was able to achieve the angular momentum, stability, the pointing accuracy and nadir-pointing requirement, but for more power than maximally allowed. The motor alone was able to perform the required angular velocity at 256 𝑚𝑊 but the complete system was able to perform it at 534 𝑚𝑊. The imbalanced flywheel causes extra torque to be spun at high speeds, which is why the current draw is much high than with only the motor. Through micro-vibration and shaker testing was revealed that the designed mount, although made of plastic for the prototyping phase, was able to survive all its own induced vibrational forces and the launch induces stresses. A small amount of damage to the motor was detected through the decrease in performance post shaker test, meaning the flywheel reinforcement has to be revised. The vacuum test revealed that the steep power increase found during operational testing was due to the increased friction caused by the utilisation of a vacuum lubricant in atmospheric conditions. Conclusively, a COTS momentum bias wheel for a PocketQube is not feasible, as meeting the technical budget requirements of a PocketQube mission is not feasible given the currently available COTS motors. Furthermore, the availability of high velocity miniaturised vacuum rated electric motors is central in the feasibility and competitiveness of wheeled attitude actuators. The competitiveness of this prototype to the proposed PocketQube reaction wheel systems comes down to the motor utilised. Designing and creating a wheeled attitude actuators borders on what is feasibly possible with commercially available parts.

Deflecting an asteroid with a kinetic impactor is the only planetary defense method for asteroid impact avoidance tested on a full scale. This thesis investigates the initial design of a kinetic impactor and its deployment architecture. During hypervelocity impact, momentum enhancement due to ejecta plays a crucial role. The impact efficiency factor, accounting for impact obliquity, ejecta asymmetry, and gravity-induced trajectory bending, is modelled. Influential parameters, including spacecraft momentum, are considered. Seven realistic design options are evaluated, with an emphasis on departure strategy from Earth, interplanetary transfer and high-level spacecraft architecture. The contribution to the research area is a parametric framework to evaluate a kinetic impactor, including a model of the momentum enhancement and a low-thrust transfer trajectory. The assembled impactor is selected based on a performance trade-off, and preliminary requirements are identified.

The rapidly increasing utilization of small satellites has led to a growing demand for improved solar array technologies. Traditional solar arrays are limited by their mass and volume constraints, and thus are struggling to meet the increased demands. This feasibility study explores the potential of roll-out solar arrays as a solution to enhance the power generation capabilities of small satellites. These innovative systems are characterised by their unfurling method of deployment, and have demonstrated improved performance compared to traditional solar arrays. By analysing the feasibility of designing a roll-out solar array for small satellites, it is hoped that the operational capabilities of small satellites can be enhanced.

A comprehensive literature study revealed a significant performance gap in high power-to-mass ratios for satellites with a power range of 300 − 1000 W , leading to the development of a roll-out solar array design specifically for this power range. The study combined an analytical model with the non- dominated sorting genetic algorithm (NSGA-II), an efficient multi-objective optimisation algorithm, to generate a set of optimal designs by maximising the power-to-mass ratio and stowed power density, with the latter parameter being another key performance indicator for roll-out solar arrays. The analytical model incorporated the core design elements of a roll-out solar array like deployment mode, deployment mechanism, solar cell type, and stiffening method. A multitude of constraints ensured the feasibility of the designs, while several verification and validation tests were conducted to confirm the accuracy of the analytical model.

The optimisation results revealed that a variety of design combinations exist that could outperform the state of the art solar array capabilities. Optimal designs demonstrated power-to-mass ratios exceeding 145 W /kg and stowed power densities over 55 kW /m3. These designs would deliver over double the power-to-mass ratio of traditional solar arrays in the 300 − 1000 W range, while remaining comparable to the 120 − 140 W /kg state of the art performance of arrays outside of this range. Furthermore, optimal designs offer up to 40% increases in stowed power density compared to the 40 kW /m3 state of the art benchmark. Despite the promising gains, the study reveals some key limitations to roll-out solar arrays. These are the increased development costs, the higher risks, and their bulky geometry being potentially inefficient for smallsats. Ultimately, the study provides for a strong foundation for the development of a roll-out solar array. The results confirm the potential of such a system, and future research is recommended to solidify the feasibility of this innovative technology.

The Deployable Space Telescope (DST) project aims to reduce the stowed volume compared to other Earth observation satellites with a similar resolution. The DST demonstrator functions in the thermal infrared (TIR) domain, which imposes stringent thermal requirements on the detector. The detector must be cooled to 150 K, and the instrument box to 200 K. This thesis documents the lowering of the detector and instrument box temperatures by minimizing the heat input on the detector and the radiators. The addition of a (deployable) Earth shade door minimizes the external heat input on the radiator surfaces. The impact of the satellite's orbit on the thermal environment and the minimum and maximum worst-case scenarios were investigated, which determined that constellation flight was possible. The proposed thermal control system design is fully passive and results in stable detector and instrument box temperatures that meet the set requirements.

ADCS magnetometers on board of satellites suffer from the disruptive magnetic field produced by the satellite itself. For CubeSats and PocketQubes this magnetic field is usually not quantified. If the largest magnetic field sources are known and if there is freedom in the magnetometer placement, the magnetometer is simply placed as far from the largest sources as possible. The magnetometer is not calibrated for the remaining error due to the satellite magnetic field. The research objective of this thesis is, ’to develop a measurement system which is capable of characterizing the satellite magnetic fields of CubeSats and PocketQubes close to the satellite body itself’. The measurement system is specifically designed for PocketQubes and is designed using the steps of the engineering design process. The design and manufacturing process was conducted separately for three subsystems; the magnetometers, the supporting structure and the data acquisition and processing. The measurement system features four magnetometers, which can be placed in various locations around the satellite with an accuracy of 0.6 mm. In the functional validation several magnetometer configurations were tested using the Delfi-PQ PocketQube as a test case. The functional validation showed that the magnetometers measure the magnetic field produced by a PocketQube with an uncertainty of 0.2 µT for each axis. The magnetic field of the test case, Delfi-PQ, ranges from approximately 5 µT close to the battery pack to 0.5 µT close to the bottom of the satellite. The measurement system can detect dynamic components of the field up to a limiting frequency of 18 Hz, this can in turn be used to trace magnetic noise contributions back to their source.
The uncertainty of 0.2 µT for each axis allows for calibration of on board ADCS magnetometers up to this uncertainty. For an ADCS magnetometer placed at the bottom of Delfi-PQ, this means a reduction in measurement uncertainty from 0.5 µT to 0.2 µT. This reduction would reduce the pointing error introduced by the satellite magnetic field disturbance and EMF model uncertainty from 0.6 to 0.3 degrees for the magnetic inclination and 1.5 to 0.7 degrees for the magnetic declination. The measurement system can also be used to select the optimal location for placing ADCS magnetometers, based on where the satellite magnetic field is the lowest and least dynamic.
The measurement system is not designed to conduct measurements on Delfi-PQ with the antennae deployed, this is a missed input requirement and a recommendation for future work. Additionally, the calibration of the magnetometers features a residual linearity error, the reduction of which is recommended as the subject of future work.

This summary is about the highlights of the final design of the LAMP (Low Altitude Modular Platform). This report follows the project plan, baseline report and midterm report. This report presents the market analysis for the platform followed by the detailed design of the platform. The design of each subsystem is treated on its own after which the integration, manufacturing and operations of all subsystems are discussed. The low-altitude modular platform is a versatile satellite platform with a wide range of capabilities. It bridges the gap between small CubeSats and high-end Earth observational satellites, while also flying at 300 Km, enabling higher resolutions in a small form factor. While the market share of CubeSats has grown a lot in recent years, their capabilities are still limited. Due to practical constraints of miniaturisation, the spacecraft bus platform typically occupies approximately 50% to 80% of the total satellite internal volume. This problem is however remedied with the use of larger satellites, which is the market gap LAMP tries to occupy. It has both the advantages of standardisation, ease of production, and low cost of CubeSats, while also possessing a large payload fraction and the bus capabilities to accommodate a high-end earth observation payload. LAMP is also innovative in its communication capabilities: It is planned to be the first satellite platform to use the SpaceX Starlink constellation. This gives LAMP unparalleled communication capabilities for an earth observational satellite in its class. LAMP is capable of sending all the information of its design payload (the DST) in livelink. In certain orbits, it is even capable of streaming 1080p 60fps video live to Earth. This opens it for a great number of new applications related to civil, law enforcement, and military surveillance...

The miniaturization of satellites throughout the decades is mostly contributed by the advancement in electronics and technology. Volume and mass of a satellite are factors of the launch costs. The Deployable Space Telescope (DST) program
aims to reduce the stowed volume of the telescope and subsequently reduce the costs.
Telescopes have an inherit design problem with reducing their size. The focal length is a property of how far a telescope can see. This length correlates with the physical length that the optical elements are separated. This is necessary for a telescope to function. The distance between the primary mirror and the
secondary mirror largely dictates the focal length. The deployment of the DST is the change in length between these optical elements after launch to when it is in orbit. This thesis is describing the development of the secondary mirror deployment mechanism for a thermal infrared space telescope.

In the DST project, a deployable space telescope is designed for the thermal infrared spectrum. For this project, a new Systems Engineering (SE) approach was needed, since the SE was outdated and the project worked in silos. Important goals for the SE approach were that it was easy to use and minimizes the workload by automating as much as possible. It was found that an automated version of classical SE would be the most suitable. To implement this, a set of tools was created. In this SE Toolset the team members do their own SE part while the program handles the overall SE. The SE Toolset showed that it can provide a solution for the gap between classical SE and Model Based Systems Engineering (MBSE), without getting as complicated as MBSE. To improve the overall SE in the DST project, a top level trade-off was done to investigate the effect of different deployment options on the stowed volume. This trade-off was used to find a suitable concept for the DST TIR 30cm demonstrator mission. The DST concepts were also compared to reference satellites to determine if the concepts were state of the art and can compete with fixed telescopes. It was found that the DST concepts perform very well and outperform current existing telescopes.

The need for Earth observation telescopes with high spatial and temporal resolution is constantly increasing, however, current state-of-the-art telescopes are costly. The Deployable Space Telescope project at TU Delft aims to provide a solution with an Earth observation telescope that has the same optical performance as today's best ones, but at a reduced cost. A baffle is required to surround the telescope to provide stable thermal environment, limit stray-light, and protect the optical components from debris. A deployable baffle consisting of pantographic arms has been designed that has only one degree of freedom, therefore the whole structure follows the configuration change if one angle is changed in it, and the diameter and height change happen synchronously, successfully reducing the required number of actuators. The proposed thermal solution decreases the thermal gradients and temperature extremes within the baffle considerably, and successfully shifts the overall temperature of the telescope towards colder regions.

The Earth magnetic field is used in many applications such as navigation. Magnetic field sensors are mainly used on satellites for attitude control and in limited missions as scientific instruments. These magnetic field instruments are: specifically designed for the mission, expensive and/or require knowledge of the magnetic noise sources present in the complete satellite. Therefore a proposal was made of a magnetic field instrument consisting of multiple COTS sensors that doesn't require knowledge of the satellite and could be placed inside the satellite. Two measurement processing methods were used to test the effectiveness of such a system. Both methods were tested simultaneously testing if an ideal sensor placement exists and the effect of the number of sensors was considered.

The requirements for a sensor system to be used for attitude control and/or scientific purposes are as follows. Firstly for both kind of systems (attitude or scientific) a measuring frequency of at least 10 Hz is preferred. Secondly the sensitivity for an attitude control instrument needs to be 50 nT or less and for a scientific instrument 0.5 nT or less. Lastly the accuracy for an attitude control instrument needs to be 50 nT and/or 3 degrees or less and for a scientific instrument 0.5 nT or less.

The two measurement processing methods used in this research are the averaging method which averages the measurements of multiple sensors and the SOE method which uses a system of equations to determine the location, direction and magnetic strength of the noise sources. It was found that the SOE method can maintain an accuracy of less than 50 nT if a determined or overdetermined system of equation could be realized. This can only be the case if the number of sensors is $3\frac{1}{3}$ times as much as the noise sources which is not feasible for a satellite. Another limitation of the SOE method is the computational time which is at least 7 order of magnitude bigger than the computational time of the averaging method. The averaging method can produce better accuracies than 50 nT and/or 3 degrees but is very dependent on the sensor placing and less dependent on the number of sensors. An increase in accuracy can only be gained if all sensors are placed far from the biggest noise sources and there should be a minimum distance from small noise sources. If one of the sensors doesn't adhere to this then there is a decrease in accuracy and a single sensor could perform better. The sensitivity of a sensor system improves when multiple sensors are used.

This is the final report in a project which aims to design a mega-constellation to survey the Earth for environmental, safety and security reasons: monitoring gases, humans, vehicles, and potential other aspects such as forest fires...

The Deployable Space Telescope (DST) is a telescope that uses foldable optical elements to reach state-of-the-art ground resolutions while minimising the launch volume and total mass of the system. One of the promising concepts that is part of the DST design is the Compliant Rolling-Element (CORE) Hinge. This hinge uses compliant links to let two semi-circular cams roll over each other with minimal friction. For this thesis, a thermo-mechanical analysis is conducted of the DST CORE hinges to determine whether the current design complies with the stability and in-orbit drift requirements of the DST. From this analysis, it can be concluded that this is currently not the case, but that a simple solution is at hand. By elongating the baffle of the DST, which protects the DST from solar irradiation, more solar radiation is blocked and the stability and in-orbit drift are kept within limits.

The need for higher spatio-temporal resolution Earth observation increases rapidly. To fill this need, the Deployable Space Telescope (DST) project aims to make a light-weight, low-volume deployable telescope. In doing so, the achievable ground resolution is high while the cost per telescope stays low. This allows for the DST to be used in constellations, thus effectively achieving a high spatio-temporal resolution. One of the issues of this design are the relative translation and rotation of the secondary mirror due to temperature fluctuations. This thesis work focused on first finding these movements by identifying the temperature variations of the secondary mirror support structure using ESATAN TMS simulations, and subsequently designing a system to keep these movements within the allowed budgets. The end-result is a novel design in which all displacements are measured by means of 4 Displacement Measuring Interferometers and corrected by means of 4 linear piezo actuators.

An innovative approach for thermal analysis and design of small satellites consisting in the study of its thermal behavior and properties from a global perspective is investigated in this research project. Spacecraft analysis and design is usually carried out in a tailored manner, based on the particular characteristics of each mission. The increasing interest in PocketQubes as space platforms, which share multiple design features, opens the possibility to develop general thermal control procedures applicable to multiple satellites independently of their payloads, configurations and orbits, within certain ranges. An exploration of the design and environmental parameters that influence the thermal behavior of picosatellites is carried out along with a sensitivity analysis in order to better understand their influence on temperatures. The Delfi-PQ satellite has been chosen as a case study for which a finite element model has been developed using ESATAN. As well, a Matlab tool has been developed for processing the data generated. Based on the results produced from these analyses, generalized conclusions on how thermal control could be achieved for satellites such as the Delfi-PQ and similar PocketQubes are derived. This study aims to set the basis for approaching thermal control of highly standardized spacecraft from a global perspective, opening new possibilities of lowering the costs and increase its reliability and performance. The outcomes and lessons learned could be later applied to other categories of similar satellites such as CubeSats.

Due to their low mass pocketqubes are relatively inexpensive to launch. This makes them ideal for the creation of large constellations in low Earth orbit and a sufficiently large constellation would be able to perform measurements over the same area multiple times per day. Equipping these pocketqubes with cameras would increase the chance of cloud free images and would allow scientists to study dynamic processes which happen on sub-daily timescales. A camera in a pocketqube would have a spatial resolution of about 40 meters which would be sufficient for, for example, ocean color measurements or measuring land related variables, such as vegetation extent or land use, which allows biologists
to better study conservation efforts. The economics of pocketqubes requires that the development cost and production cost per pocketqube must be comparable to the launch cost. For this reason it is preferable to use commercial off the shelf over purpose built cameras. to test whether or not those commercial off the shelf cameras would survive in space, environmental testing is required. The launch environment, consisting of the accelerations and vibrations experienced during launch, and the thermal-vacuum environment potentially have the largest immediate impact on a camera and should be tested for. Other environmental effects in low Earth orbit, such as atomic oxygen, UV radiation, ionizing radiation, et cetera, are damaging over time but might only require analysis rather than extensive testing. Due to limited resources only a thermal-vacuum test was conducted. For this test two cameras, the See3CAM_CU30 and a generic ELP H264 720p usb camera, were selected. These cameras were subjected to a thermal-vacuum test by placing the cameras inside a vacuum-oven in which the cameras were exposed to a vacuum while being heated to a temperature of 50∘퐶. Because the vacuum-oven was not suitable to test the cameras at temperature below room temperature a thermal-ambient test was also conducted in which the cameras experienced four cycles of heating up and cooling down to the expected in-orbit temperature extremes of 50∘퐶 and −5∘퐶 respectively. Before, in between, and after the thermal-ambient and thermal-vacuum tests a performance test was conducted to identify changes to the cameras. These performance tests measured the modulation transfer function, change in color representation, change in full well capacity, change in average dark signal, change in chromatic aberrations, and change in image distortion.
Both cameras still functioned after the thermal-ambient and thermal-vacuum tests and showed no significant outgassing. Both cameras also showed no change in the distortion and chromatic aberration after the conducted tests indicating that the lenses survived the thermal-vacuum environment without
detectable changes. After the thermal-ambient test, conducted prior to the thermal-vacuum test, only the See3CAM_CU30 showed some minor changes in its color representation. The ELP camera was unaffected. After the thermal-vacuum test the See3CAM_CU30 only showed again some minor changes in the color representation. The ELP camera became unusable. It experienced a large change in color representation and showed significant performance deterioration in the full well capacity and average dark signal. This indicates the camera overexposing all its images. However, a test measuring the effective integration time at various camera settings before and after the experiments showed that
the camera integration time is unlikely to have been affected.
Both cameras still functioned after the thermal-vacuum experiments. However, the fact that the ELP camera, and to a lesser extent the See3CAM_CU30, experienced performance degradation shows the necessity for good and rigorous performance testing of commercial off the shelf cameras before
using them in pocketqubes. Although the See3CAM_CU30 showed potential it is too early to conclude whether or not it can survive in space for any length of time. In order to determine this at least a test is required demonstrating its ability to survive the mechanical environment experienced during launch.

As the new space movement develops, so comes the request for cheaper space solar panels with a shorter delivery time. This master thesis investigates the possibilities of using a Semi-flexible terrestrial solar panel in a low Earth orbit space environment. A Semi-flexible solar panel is a sandwich of polymer films with the solar cells in between, these layers are laminated to create the end product. This, for space, new process is applicable to all major solar cell types. It enables a wide range of possible designs and could result in a drastically lower price, shorter production time and a lighter and smaller end product. By using system engineering methodology, the risks of Semi-flexible solar panels in space are identified after which the major risks have been tested. The risks investigated in this report are: outgassing, temperature cycling, vacuum UV radiation, charged particle radiation and the stiffness in deployed and stowed position. The results show no red flags for a seven-year 600 km Earth orbit. A transmission degradation of 10% is observed, resulting in lower power output for the selected material. When entering higher orbits, the amount of radiation leads to delamination and potentially no power output. When comparing the Semi-flexible solar panel to a conventional solar panel, a price difference of factor 4 and a power output difference of 33% is expected. The vast price difference shows the potential for the concept, but further investigation is needed to find out if the transmission degradation could be mitigated and whether designs comply with stiffness and vibration requirements.

The PocketQube is an emerging satellite class, which pushes the miniaturization of space technology beyond the well-established CubeSats, promising rapid design-to-orbit cycles while lowering the cost of accessing space. A showstopper in the success story of nano- and picosatellites are their high mission failure rates. Environmental testing before flight is an effective means to identify design flaws and workmanship errors, and thus improve the chances of mission success. However, inefficiencies in the design process, low-budgets, and stringent schedule requirements often motivate small satellite developers to postpone environmental testing towards the end of the design lifecycle, where recovering for design flaws is inefficient. The project circumstance of nano- and picosatellite missions require, therefore, cost- and time-effective test strategies that facilitate early design evaluation. This research proposes and implements a thermal screening method for PocketQube subsystems to identify temperature hotspots and verify their compliance against operational hardware limits. Key elements of the test method are a thermal IR temperature scan at ambient conditions and an estimation of the worst-case flight temperature using experimentally derived graphs that describe the vacuum heating of thermal hotspots. The study of subsystem layout options complements the screening method by providing solutions to mitigate hotspot overheating. Moreover, the study proposes to lower the required pressure levels for thermal-vacuum testing of PocketQube subsystems. The analysis shows that, due to the small form factor, pressure levels by four orders of magnitude larger than those used in environmental test standards for larger satellites suffice to maintain the resulting temperature errors below 5K on the hot side of the temperature spectrum. Both the screening method and moderation in vacuum requirements contribute to the development of subsystem test methods that match the needs of small satellite developers.

There are several benefits of using autonomous sensors in spacecraft. Avoidance of wired connections reduces cost, mass, and increases the flexibility and reliability of the system. The impact of wire reduction can be significant, especially for small satellites with many sensors, like temperature and sun sensors. Previous research has already focused on wireless intra- spacecraft communications. This research tests the self-powering capabilities of a system based on a COTS thermoelectric generator connected to a Bluetooth Low energy communication system, with a built-in controller and temperature sensor, and a power management interface. The system will be considered as a candidate for an autonomous temperature sensor in a future PocketQube mission of the university.

Controlled temperature differences can be achieved in a test environment, allowing the measurement of the generator power capabilities. It is tested that the system requires, for operation, a minimum temperature difference of 2.31 degrees between the extremes of the thermoelectric generator. It generates a peak power of 234 μW for that difference. In addition, the voltage difference obtained of 35.5 mV exceeds the minimum voltage required by the power management subsystem to be used. The power management sub-system consists of an ultra-low power converter that provides an output voltage of 4.1 V and a measured power efficiency of 32 % Moreover, thanks to the management of the Bluetooth sleeping modes, with the built-in controller and several operational amplifier comparators, an average power consumption of 5 μW is required during operation. The case studied would allow measuring temperature and sending the data over a Bluetooth link to the on-board computer every 16.2 seconds

It is concluded that the technology, based on COTS components, can be implemented and considered as the first step for a fully autonomous sensor with thermoelectric power generation in small satellites. Its implementation may provide substantial advantages for remote or/and locations where wiring is difficult to integrate. The tested performance values provide the foundation to develop the technology further.

the aim of this thesis is to regard all aspects concerning mass optimization with respect to small satellite structures and deployers, specifically aimed at the DelfiPQ PocketQube mission.