During ECSSMET 2023, the CREAM HDRM was introduced to the space community [1]. The major advantage of this mechanism is its novel approach to lock the release bolt within the mechanism. The locking mechanism is based on friction, which is generated by a collet-cone design. The major advantages of this strategy are simplicity in the design, low-shock characteristics and self-resetability. A new scaled version of the CREAM HDRM was developed within the last year due to the emerging need for larger HDRM mechanisms. The system just finished final qualification to TRL6. The principle of the locking mechanism remained the same. However, due to the significantly larger preload, upscaling of the design elements became necessary. The paper starts with a brief description of the design principle. A detailed characteristic of the push-pull device is shown, as this element required most design changes to allow high release forces. The achieved specifications of the CREAM 4.5 kN HDRM, including the shock emission at release, are discussed afterwards. The rest of the paper focuses on the details of the qualification campaign. This campaign includes mechanical (vibration, shock), thermal-vacuum and life testing. The paper closes with the overview of the commercialization process into the space market through an industrial partner.

The paper outlines the results of the development and testing of a 100 W deployable solar array for SmallSats. DEAR is a joint project between DLR, Astronika, AZUR SPACE Solar Power and ESA and aims at the development and ground demonstration of a membrane-based, deployable 100 W solar array that can be stowed in and deployed out of a 1 U CubeSat volume. In the paper the main requirements for the development will be addressed which will lead to specific design constraints. Therefore, the design is mainly driven by the required 100 W power End-of- Life, its operation in a low Earth orbit environment for at least 5 years and its tight volume and mass restrictions. Especially in terms of mass and volume constraints these requirements demand that flexible or at least semi-flexible solar arrays need to be addressed for the problem solution. The paper will give an overview on the overall system design, the functionality of the deployment process and will present the results of the environmental and functional engineering model test campaign. An overview for the further planned in-orbit demonstration missions will be presented.

This is a report about the simulation study of a lightweight rover being separated from a larger spacecraft in vicinity of a small Solar system body. Monte Carlo simulation and statistical analysis estimate the rate of requirement violation depending on uncertain rover and separation mechanism parameters. During the rover development phase, the full Monte Carlo analysis is only done with a simplified model. The first iterations show relatively high probability of non-compliance with requirements. Model refinement and more precise hardware values lead to the final assessment that separation meets the requirements within a ∼ 2 σ region. Another key result is the identification of relevant and irrelevant effects on the spacecraft rover separation. A more detailed model serves for analysis of selected samples during the development phase. For this model, a full Monte Carlo analysis is only done close to the end once the parameter ranges are fixed, mostly for time reasons. This more detailed analysis in general confirms the ∼ 2 σ requirement meeting. The main difference with the simplified model is the much larger number of outliers, a few of which also violate the allocated volume for the rover in proximity of the mothership.

The paper outlines the results of the intermediate development and testing of an 100 Wdeployable solar array for SmallSats. The activity is funded and carried out under ESA contract No. 4000133890/21/NL/MM/ra within the ARTES Advanced Technology programme. DEAR is a joint project between DLR, Astronika, AZUR SPACE Solar Power and ESA and aims at the development and ground demonstration of a membrane-based, deployable 100 W solar array that can be stowed in and deployed out of a 1U CubeSat volume. In the paper the main requirements for the development will be addressed which will lead to specific design constraints. Therefore, the design is mainly driven by the required 100 W power End-of-Life, its operation in a low Earth orbit environment for at least 5 years and its tight volume and mass restrictions. Especially in terms of mass and volume constraints these requirements demand that flexible or at least semi-flexible solar arrays need to be addressed for the problem solution. The paper will give an overview on the overall accommodation, the functionality of the deployment process and will present test results on breadboard level. An overview for the further ground demonstration of the development by means of an engineering model test campaign is presented.

Flexible Printed Circuit Boards (Flex PCB) are commonly used for the electrical routing on spacecrafts (S/C). The implementation of Flex PCB as structural material and harness for the Gossamer Solar Array (GoSolAr) is currently studied by the German Aerospace Center (DLR). The Flex PCB as well as the other parts of the array are folded and stowed into a small volume for transport into space and deployed once the spacecraft starts operating. The folding areas, also referred to as hinge, are subject of this paper. The hinges and their design have major impact on the required deployment force and geometry. The aim is to get a better understanding of the relevant parameters, which are the deployment force, the angles between photovoltaic generators and the deployed distance. Deployment experiments were carried out on four different design options, which were then compared to each other with respect to their deployment behaviour. Important aspects for this evaluation were the controllability during the deployment and the force required for tensioning the array. Main differences of the evaluated designs were the amount of elastic deformation energy stored in the different configurations (stowed and deployed). The experiments revealed that a design, which has its minimum of elastic deformation energy in stowed configuration is beneficial to be used for GoSolAr. Furthermore, a numerical model of the hinge based on the FEM is introduced. The experiments were used to correlate the model. This model can be further used for design studies and optimization. The design of Flex PCB as hinges adds a mechanical function to its otherwise only electrical purpose and therefore opens a whole new field of applications, especially for thin film and lightweight structures.

Highly efficient low-thrust propulsion is increasingly applied beyond commercial use, also in mainstream and flagship science missions, in combination with gravity assist propulsion. Another recent development is the growth of small spacecraft solutions, not in size but in numbers and individual capabilities. Just over ten years ago, the DLR-ESTEC GOSSAMER Roadmap to Solar Sailing was set up to guide technology developments towards a propellant-less and highly efficient class of spacecraft for solar system exploration and applications missions: small spacecraft solar sails designed for carefree handling and equipped with carried application modules. Soon, in three dedicated GOSSAMER Roadmap Science Working Groups it initiated studies of missions uniquely feasible with solar sails such as Displaced L-1 (DL1) space weather advance warning and monitoring, Solar Polar Orbiter (SPO) delivery to very high inclination heliocentric orbit, and multiple Near-Earth Asteroid (NEA) rendezvous (MNR). Together, they demonstrate the capability of near-term solar sails to achieve at least in the inner solar system almost any kind of heliocentric orbit within 10 years, from the Earth-co-orbital to the extremely inclined, eccentric and even retrograde. Noted as part of the MNR study, sail-propelled head-on retrograde kinetic impactors (RKI) go to this extreme to achieve the highest possible specific kinetic energy for the deflection of hazardous asteroids. At DLR, the experience gained in the development of deployable membrane structures leading up to the successful ground deployment test of a (20 m)(2), i.e., 20 m by 20 m square solar sail at DLR Cologne in 1999 was revitalized and directed towards a 3-step small spacecraft development line from as-soon-as-possible sail deployment demonstration (GOSSAMER-1) via in-flight evaluation of sail attitude control actuators (GOSSAMER-2) to an envisaged proving-the-principle flight in the Earth-Moon system (GOSSAMER-3). First, it turned the concept of solar sail deployment on its head by introducing four separable Boom Sail Deployment Units (BSDU) to be discarded after deployment, enabling lightweight 3-axis stabilized sailcraft. By 2015, this effort culminated in the ground-qualified technology of the DLR GOSSAMER-1 deployment demonstrator Engineering Qualification Model (EQM). For mission types using separable payloads, such as SPO, MNR and RKI, design concepts can be derived from the BSDU characteristic of DLR GOSSAMER solar sail technology which share elements with the separation systems of asteroid nanolanders like MASCOT. These nano-spacecraft are an ideal match for solar sails in micro-spacecraft format whose launch configurations are compatible with ESPA and ASAP secondary payload platforms. Like any roadmap, this one contained much more than the planned route from departure to destination and the much shorter distance actually travelled. It is full of lanes, narrow and wide, detours and shortcuts, options and decision branches. Some became the path taken on which we previously reported. More were explored along the originally planned path or as new sidings in search of better options when circumstance changed and the project had to take another turn. But none were dead ends, they just faced the inevitable changes when roadmaps face realities and they were no longer part of the road ahead. To us, they were valuable lessons learned or options up our sleeves. But for future sailors they may be on their road ahead. (C) 2021 COSPAR. Published by Elsevier B.V. All rights reserved.

A design for a 4 m long, ultra-light, high-gain, helical Antenna made from fiber-composite material will be presented. The antenna was designed for the DLR NanoSatellite Mission AISat to receive signals from the Automatic Identification System (AIS) of maritime applications. A description of the antenna deployment strategy including release mechanisms will be given. The proof of concept will be presented based on experimental results gained during the 15. DLR parabolic flight campaign (PFC) in March 2010 and several development tests. Finally, in-orbit demonstration was performed during the two years of operation of the AISat after the successful launch on June 30, 2014 from Sriharikota (India). The AISat satellite was developed at the DLR Institute of Space Systems aiming at the worldwide receiving of AIS signals. These signals can usually be received along coast lines or from ship to ship in range of sight. They provide identity, position, velocity and heading and are therefore used for ship tracking. A number of AIS satellites already exist but especially in areas with high ship traffic density identification problems arose due to the high signal density. Therefore AISat has a distinctive ultra-light, high-gain, helical antenna which allows to focus on comparably small areas on the Earth surface. IT thus shall enable the receiving of Class A and B and SART signals especially in high traffic density zones. The antenna is a 4 m long and 0.57 m in diameter deployable helix antenna made from fiber composite material, which can be stowed in a very flat volume of merely 100 mm height. The wire of the antenna is made from carbon fiber material with a diameter of 8 mm. It is covered with a copper cord for high electrical conductivity. Based on its design with 8 windings the total length of the wire itself is approx. 16 m. Through the dedicated usage of fiber composite materials this wire weighs less than 1 kg including the copper cord. In stowed configuration, in which it is held down by 3 release mechanisms, the antenna has stored elastic energy like in a spiral spring. After release the structure deploys autonomously in orbit to a length of 4 m. When deployed, the antenna is still pre-stressed using control cords in order to increase its bending stiffness.

In recent years, the German Aerospace Center (DLR) developed Gossamer deployment systems in different projects. As power requirements of spacecraft are getting more and more demanding, DLR recently focused on the development of new deployable photovoltaic (PV) technologies that are suitable for generating 10’s of kW per array. Possible space applications that may also require high power supply are missions using electric propulsion such as interplanetary missions, placing of geostationary (GEO) satellites in their orbit or even more future oriented as space tugs or lightweight power generation on extra-terrestrial infrastructures. The paper gives an overview about a feasibility study for flexible solar arrays based on new thin-film photovoltaics. It is expected that the combination of new thin-film PV technologies, e.g., copper indium gallium selenide (CIGS) cells or gallium–arsenide (GaAs) cells, together with Gossamer deployment technologies, could significantly increase the power availability for spacecraft. Based on a requirement, analysis system concepts were evaluated. A focus is on the potential of CIGS PV combined with a two-dimensional deployment of the array and DLR’s coilable carbon fibre-reinforced plastic (CFRP) booms. Therefore, a concept based on crossed booms with a foldable PV membrane is considered as baseline for further developments. The array consists of rectangular PV generators that are interconnected by flexible printed circuit board (PCB) harness. By a double-folding technique, these generators are laid on top of each other in such that the membrane can be extracted from its stowing box during the deployment in a controlled manner. Considering constantly increasing efficiencies of the CIGS PV combined with Gossamer structures, there is clear potential of reaching a very high specific power value exceeding that of conventional PV systems. Furthermore, the CIGS PV appears to be more radiation resistant and has already reached more than 21% efficiency in laboratories. Such efficiencies are expected to be achieved in the near future in a standard manufacturing process. However, flexible, thin-film GaAs cells are also subject of consideration within GoSolAr. With this prospect, DLR’s research has the goal to develop a Gossamer Solar Array (GoSolAr) to exploit the described potential.

The "Heat Flow Property Package Instrument" (HP3) is part of NASA's current Mars mission "InSight", which was launched in 2018 and currently operates on the surface of Mars. The instrument needs to remain at its initial position and orientation during operation. Although the landing site can have significant tilt and can be covered with low cohesion soil, any mechanical excitation might make the instrument slip. Therefore, the instrument is using a tailored feet design, which can withstand lateral loads. Future instruments might require higher resistance against slip. This can be due to stronger tilted landing sites or due to higher shocks emitted from stronger penetration probes. This paper introduces a novel design for those instruments based on the idea of the "spaced-link track" of Bekker to further minimize slippage. This design concept is originally used on tracks of heavy machinery. It is presented how the major design feature can be incorporated into the current design. A newly developed analytical-numerical model is utilized to estimate the track force of the new design. The paper closes with a design study at which the new design and the current design are compared to each other for different sized feet.

The HP 3 instrument measures the thermal flux through the Martian crust using a penetration probe. Launched on the InSight mission in 2018, HP 3 was deployed for penetration activities in the beginning of 2019. This paper shortly introduces the HP 3 instrument and shows the design and analysis process of the feet pattern and of each foot. The pattern of the feet is driven by the accommodation of the instrument on the lander and the need to ensure stability of the instrument on the surface. A four-feet design was chosen as it can be fitted very good into the available volume. During operation, the instrument is vulnerable to tipping as the landing site can be tilted and partially covered with rocks and depressions. Therefore it is shown analytically, that the instrument is stable, although it might be partially located on rocks or depression in combination with a tilted landing site. The new design of the feet is introduced and compared to a flat-plate foot. The capability of each design to produce slippage resistance is estimated analytically. The analytical estimations show, that the chosen flight-design generates more slippage resistance than the flat-plate approach. Even though the analytical estimations do not cover all effects generating slippage resistance.

On the background of the highly successful flight of the first interplanetary solar sail, JAXA’s IKAROS, and with NASA’s upcoming NEASCOUT nanospacecraft solar sail, there is a rising interest in large lightweight structures in space. Deployable membrane or ‘gossamer’ structures can provide very large functional area units for innovative space applications which can be stowed into the limited volumes of launch vehicle fairings as well as ESPA or ASAP secondary payload launch slots, depending on the scale of the mission. Large area structures such as solar sails which have been studied for many decades require a technology that allows controlled and safe deployment. Scientific as well as commercial mission require demonstrated reliability, i.e., TRL 6 or higher. A reliable technology that enables controlled deployment was developed in the GOSSAMER-1 solar sail deployment demonstrator project of the German Aerospace Center, DLR. Its functionality was verified in various laboratory tests to qualify the hardware for a first demonstration in low Earth orbit. We provide an overview of the GOSSAMER-1 hardware development and qualification campaign. On its engineering models, all aspects of the deployment were tested at ambient conditions. The key components and mechanisms of this technology were subjected to environmental qualification testing, enabling future flight use of an innovative stowing and deployment strategy for a controlled deployment. The stowing and deployment strategy was verified by tests with an engineering qualification model using one complete of the four Boom Sail Deployment Units (BSDU) of the DLR GOSSAMER concept, two membrane sections, and two simulated BSDUs. This efficient approach was possible due to the inherent symmetry of the design. Test-asyou-fly procedures included vibration tests, venting, thermal-vacuum tests, and ambient deployment. In these tests the deployment strategy proved to be suitable for a controlled deployment of gossamer spacecraft, and deployment on system level was demonstrated to be robust and controllable. The GOSSAMER-1 solar sail membranes were also equipped with small thin-film photovoltaic arrays to supply the core spacecraft. Beyond solar sailing, which is a uniquely performant method of propulsion when it comes to asteroid rendezvous missions, we explore other use cases for thin membrane-based structures in planetary defense and related missions. In these we consider the small spacecraft based solutions for mass and volume efficient designs as developed in the DLR-ESTEC GOSSAMER Roadmap and ASTEROIDFINDER/SSB studies, and implemented in our small spacecraft missions such as AISAT-1 and MASCOT. **************************************

Following the highly successful flight of the first interplanetary solar sail, JAXA's IKAROS, with missions to come such as NASA's NEAscout nanospacecraft solar sail and JAXA's Solar Power Sail (a solar-electric propelled mission to a Jupiter Trojan asteroid), and on the back-ground of the ever increasing power demand of GEO satellites now including all-electric spacecraft, there is renewed interest in large lightweight structures in space. Among these, deployable membrane or `gossamer' structures can provide very large functional areas for innovative space applications which can be stowed into limited volumes of launch vehicle fairings as well as secondary payload launch slots, depending on the scale of the mission. Large area structures such as solar sails or high-power photovoltaic generators require a technology that allows their controlled and thereby safe deployment. Before using such technology for a dedicated science or commercial mission, it is necessary to demonstrate its reliability, i.e., TRL 6 or higher. A reliable technology that enables controlled deployment was developed in the Gossamer-1 solar sail project of the German Aerospace Center. It included the verification of its functionality with various laboratory tests to qualify the hardware for a first demonstration in low Earth orbit. We provide an overview of the Gossamer-1 hardware development and qualification campaign. The design is based on a crossed boom configuration with triangular sail segments. Using engineering models, all aspects of the deployment were tested under ambient environment. Several components were also subjected to environmental qualification testing. An innovative stowing and deployment strategy for a controlled deployment and the required mechanisms has been worked out. The tests conducted provide allow a mechanical characterization of this process, in particular the measurement of the deployment forces. The stowing and deployment strategy was verified by tests with an engineering qualification model of one (out of four) Gossamer-1 deployment units. According to a test-as-you-fly approach the tests included vibration testing, venting, thermal-vacuum testing and ambient deployment testing. In these tests the deployment strategy proved to be suitable as a controlled deployment for gossamer spacecraft. Deployments on system level were demonstrated to be robust and controllable. The Gossamer-1 solar sail membranes were also equipped with small thin-film photovoltaic arrays intended to supply the core spacecraft. In our follow-on project GoSolAr, the focus is now entirely on deployment systems for huge thin-film photovoltaic arrays. Based on the Gossamer-1 experience, deployment technology and qualification strategies, new technologies for the integration of thin-film photovoltaics are being developed and qualified for a first in-orbit technology demonstration within five years. Main objective is the further development of a deployment technology for a 25 m 2 gossamer solar power generator and a flexible photovoltaic membrane. GoSolAr enables a wider range of deployment concepts beyond solar sail optimized methods. It uses the S2TEP bus system developed at the Institute of Space Systems as part of the DLR satellite roadmap.

In recent years the German Aerospace Center (DLR) developed Gossamer deployment systems in different projects. DLR recently focused on the development of new deployable photovoltaic (PV) technologies that are suitable for generating 10's of kW per array as power requirements of spacecraft are getting more and more demanding. Possible space applications that may also require high power supply are missions using electric propulsion like interplanetary missions, placing of geostationary (GEO) satellites in their orbit or even more future oriented as space tugs or lightweight power generation on extraterrestrial infrastructures. The paper gives an overview about a feasibility study for a flexible solar array based on new thin-film photovoltaics. It is expected that the combination of new thin-film PV technologies, e.g. Copper Indium Gallium Selenide (CIGS) cells, together with Gossamer deployment technologies, could significantly increase the power availability for spacecraft. Based on a requirement analysis system concepts were evaluated. A focus is on the potential of CIGS PV combined with a two-dimensional deployment of the array and DLR's coilable Carbon Fiber Reinforced Plastic (CFRP) booms. Therefore, a concept based on crossed booms with a foldable PV membrane is considered as baseline for further developments. The array consists of rectangular PV generators that are interconnected by flexible Printed Circuit Board (PCB) harness. By a double folding technique these generators are laid on top of each other in such that the membrane can be extracted from its stowing box during the deployment in a controlled manner. Considering constantly increasing efficiencies of the CIGS PV combined with Gossamer structures, there is clear potential of reaching a very high specific power value exceeding that of conventional PV systems. Furthermore, the CIGS PV appears to be more radiation resistant and has already reached more than 20% efficiency in laboratories. Such efficiencies are expected to be achieved in the near future in a standard manufacturing process. DLR's research has the goal to develop a Gossamer Solar Array (GoSolAr) in order to investigate and exploit the described potential.

20 years after the successful ground deployment test of a (20 m)(2) solar sail at DLR Cologne, and in the light of the upcoming U.S. NEAscout mission, we provide an overview of the progress made since in our mission and hardware design studies as well as the hardware built in the course of our solar sail technology development. We outline the most likely and most efficient routes to develop solar sails for useful missions in science and applications, based on our developed 'now-term' and near-term hardware as well as the many practical and managerial lessons learned from the DLR-ESTEC GOSSAMER Roadmap. Mission types directly applicable to planetary defense include single and Multiple NEA Rendezvous ((M)NR) for precursor, monitoring and follow-up scenarios as well as sail-propelled head-on retrograde kinetic impactors (RKI) for mitigation. Other mission types such as the Displaced L1 (DL1) space weather advance warning and monitoring or Solar Polar Orbiter (SPO) types demonstrate the capability of near-term solar sails to achieve asteroid rendezvous in any kind of orbit, from Earth-coorbital to extremely inclined and even retrograde orbits. Some of these mission types such as SPO, (M)NR and RKI include separable payloads. For one-way access to the asteroid surface, nanolanders like MASCOT are an ideal match for solar sails in micro-spacecraft format, i.e. in launch configurations compatible with ESPA and ASAP secondary payload platforms. Larger landers similar to the JAXA-DLR study of a Jupiter Trojan asteroid lander for the OKEANOS mission can shuttle from the sail to the asteroids visited and enable multiple NEA sample-return missions. The high impact velocities and re-try capability achieved by the RKI mission type on a final orbit identical to the target asteroid's but retrograde to its motion enables small spacecraft size impactors to carry sufficient kinetic energy for deflection.

On May 05, 2018 NASA JPL launched its mission to Mars called “InSight”. Main objective of this mission is to gain more knowledge of the evolution of terrestrial planets. Beside a number of different scientific instruments onboard the lander, there are two instruments that will perform measurements on the Martian ground. One of the instruments is the Heat Flow and Physical Properties Package (HP 3 ), which was developed by the German Aerospace Center (DLR) to measure the heat flow of the Martian outer crust. It uses a hammering mechanism equipped with heating foils on the outer hull for regolith thermal conductivity measurements and pulls a tether approx. 5 m into the soil. The tether is equipped with temperature elements for the determination of temperature gradients in the ground. There is the need of a separate system to be able to perform those activities on the surface. This system is called the “HP 3 Support System”. Its main task is to ensure a stable, nearly perpendicular position of the hammering mechanism relative to the soil on the Martian surface before initial penetration. It also houses the instrument for penetration depth measurement and serves as electrical connection to the lander. The paper gives an overview of the development and the qualification of the structure of the Support System. It will focus on the mechanical design, the analysis of the structural dynamics but in particular on the testing which includes standard environmental testing but also numerous development tests that are very mission specific. The mechanical design of the Support System is mainly driven by a unique set of requirements derived from the working environment on Mars, the deployment from the lander deck and the mechanically separated operation on the surface. The instrument design will be explained to show which design elements were implemented to ensure proper functionality. Various development tests were developed and performed to show compliance of the instrument design to the requirements. Such tests are: Separation Tests from the lander deck in cold environment under various tilt angles, Tether Deployment Tests, under various temperatures, foldings and routings, Feet Sliding Resistance Tests on sand with different slopes and the “Cold Drop and Dig test”, which covers the initial hammering phase of the Mole. The paper will give an overview on all tests necessary for the support system qualification and will describe test setups and the results.

The Martian Moons eXploration (MMX) is a mission by the Japan Aerospace Exploration Agency, JAXA,to the Martian moons Phobos and Deimos. It will primarily investigate the origin of this moon by bringing samples back from Phobos to Earth and deliver a small (about 25 kg) Rover to the surface. The Rover is a contribution by the Centre National d’Etudes Spatiales (CNES) and the German Aerospace Center (DLR). Its currently considered scientific payload consists of a thermal mapper (miniRAD), a Raman spectrometer (RAX) a stereo pair of cameras looking forward (NavCAM)and two cameras looking at the interface wheel-surface (WheelCAM) and consequent Phobos’ regolith mechanical properties.The cameras will serve for both, technological and scientific needs. The MMX rover will be delivered from an altitude of <100 m and start uprighting and deploying wheels and a solar generator after having come to rest on the surface. It is planned to operate for three months on Phobos and provide unprecedented science while moving for a few meters to hundreds of meters. MMX will be launched in September 2024 and inserted into Mars orbit in 2025, the Rover delivery and operations are planned for 2026-2027.