The idea of the inflated method is to change the ballistic coefficient of the structure so that the gravitational force on the system can easily be manipulated. In the end, the system’s potential energy will be decreased while the pulling force towards Earth will be increased. When the debris enters the atmosphere, the friction will handle the dissipation of the kinetic energy. The most well-known study that researched this topic is the Gossamer Orbit Lowering Device (GOLD) (Nock et al., 2010). The summary of the concept is that the inflated balloon mechanism is integrated into the target satellite before it is launched. When a de-orbiting operation is needed, the envelope is filled with gas. The envelope material is ultra-thin and lightweight. Therefore, even a tiny amount of gas is enough to inflate the balloon. The drawback of this concept is the high risk of the possible damage to the balloon by other space debris having dangerously high momentum (Robinson, 2003), or this balloon-like structure can damage other operational satellites in space since it has no advanced positioning control system. The research conducted for this method was kept in concept development and simulation phases. No inflated method-based mission has been organized yet, and the highest TRL ever reacted for this method is TRL-4.

Space tethers are essentially used for propulsion purposes in space. These tethers are made of electrical conductor materials that can create a voltage difference between the edges of the tether. There are two common tethering methods in the literature, electrostatic tethering and electrodynamic tethering, and the detailed descriptions of these methods are given in the following subsections. When the tether physically interacts with the charged particles of the ionosphere in LEO, drag force is generated on the tether. Therefore, this principle is used to de-orbit any space object (Hou et al., 2020). The space tethers can be integrated on a spacecraft even before it is launched as an onboard system or attached by a different spacecraft in space. When the tether is spooled out, the debris’s potential energy decreases while the pulling force towards Earth is increased. This pulling force turns the potential energy of the debris into kinetic energy. Space tether systems are varying in terms of many different aspects, such as open-loop or closed-loop control (Sun and Zhong, 2020; Zhu et al., 2021) and rigid or elastic (Zhao et al., 2020; Razzaghi et al., 2021; Shan and Shi, 2021; Hu et al., 2022; Zhang et al., 2022).

Space tethers are very famous and have a wide range of application areas. When firstly proposed, the actual purpose of the electrodynamic tether method was to change the orbit of satellites. The NASA’s Gemini-11 mission demonstrated passive attitude stabilization of the two spacecraft connected by a tether and generated artificial gravity by spinning the integrated spacecraft. In ASI (Italian Space Agency) TSS-1 and ASI TSS-1R missions, it was aimed to generate electrical energy based on the Faraday effect. The ESA’s YES2 mission was focused on testing the usage of tethers as a de-orbiting propulsion system (Messidoro et al., 2017). The tethering applications conducted so far use a de-orbiting kit integrated on the target satellite before it is launched. The tethering process is triggered when the life of the satellite ends. Therefore, the de-orbiting kit and the target satellite start to have interaction. The disadvantage of this approach is that it is not valid for already existing uncooperative targets.

• Electrostatic tether method: the idea behind this tethering method is the creation of electrical potential between the edges of the tether so that its interaction with the charged particles of the ionosphere (Griffith et al., 2009) generates a drag force that pulls the space debris towards Earth. An integrated battery produces the electrical potential, and unlike the electrodynamic tether method, this battery is not charged by the particles of the ionosphere. One of the most advanced works conducted so far is the Finnish Meteorological Institute’s Electrostatic Tether Plasma Brake (Janhunen et al., 2017) in which verification and validation were performed with Aalto-1 satellite (3-U CubeSat). The advantage of the electrostatic tether over the electrodynamic tether is that the open-loop control of the electrostatic tether brings a severe amount of simplicity. The electrodynamic tether is frequently charged by the ionosphere’s particles so that this charge needs to be integrated into the closed-loop structure of the control algorithm. On the contrary, there is no charging requirement for electrostatic tethers. The electrostatic tether is controlled via DC voltage of an integrated battery (Tarabini Castellani et al., 2020). Electrostatic tether systems can be used not only for removing space debris but also for many different purposes, such as planetary exploration (Quadrelli et al., 2017) and space elevators (Bolonkin, 2005). The highest TRL ever reached for this technology is TRL-6.

The foaming method is based on spraying foam onto the target debris and manipulating its ballistic coefficient. The most important advantage of this method is that there is no need for a docking mechanism or de-tumbling the target so that this method is applicable for non-cooperative targets having high angular velocities. While the chaser satellite has proximity rendezvous with space debris, the foaming process is started as follows: the foam is sprayed towards the target satellite (Pergola et al., 2011), the foam ball that is created at the end increases the surface area of the target, the target slows down since its surface area interacts with the atmospheric particles more intensively, the rising area-to-mass ratio of the foam-covered debris generates an atmospheric drag force, the target’s potential energy is initially dissipated because of the enhancement of the friction and pressure of the atmospheric particles that hit towards the increased surface, and after the target is fully covered by the foam and becomes a foam ball, it starts getting into the atmosphere and burns. The contradictory point is that the foam should be stiff enough so that the foam ball will not be destroyed by small pieces of other space debris existing in the same orbit, and new space debris will not be created (Andrenucci et al., 2011). If the foam ball is overly soft, it will not last long. In conclusion, the chemical composition of the foam needs to have a balance between the stiffness and softness spectrum. This issue constitutes the disadvantage of using this method: determining these coefficients for varying space debris types and topology. The foam technology has also different usage areas in space, such as space debris shielding (Rostilov and Ziborov, 2021). No foam-based mission has been organized yet, and the concept has low TRL (Opiela, 2009). The highest TRL ever reached is TRL-4.

The idea behind the solar radiation force method is based on using the solar radiation pressure to create drag force and decrease the altitude, potential energy in other words, of the satellite, as long as the orientation of the surface of the solar sail wings is well aligned, as orthogonal, with the solar array.

Some of the advantages of this method are that solar sails are scalable, which means that it can be applied to a varying range of weight and geometry. Moreover, because of the topology of the principle, the method is easy to model (Borja and Tun, 2006; Johnson et al., 2012; Serfontein et al., 2021), easy to deploy (Lücking et al., 2011; Stiles et al., 2013), and passive (Fernandez et al., 2014; Seefeldt, 2017; Seefeldt et al., 2017). Passiveness of the method brings certain advantages in terms of energy consumption [extra˙73] [extra˙86].

Solar radiation pressure applied onto the wings of the solar sail, which have orthogonality with the solar array, generates a drag force that pushes the satellite towards Earth. Active control of the solar wings is crucially needed to ensure orthogonality throughout the process. While the potential energy of the debris is being dissipated using the solar sail wings, the distance of the debris to Earth is decreasing, which means that the debris will get into a more dense atmosphere. As a secondary output, the debris’s kinetic energy is dissipated too since it encounters atmospheric particles more frequently. The more the potential energy of the debris dissipated, the more the kinetic energy dissipated; this causality creates an infinite loop between the potential and kinetic energy dissipations. Another type of drag force is the atmospheric drag. The atmospheric drag is strictly dependent on the atmosphere’s density, which changes considerably throughout the orbit. Moreover, the atmospheric drag force is also dependent on the geomagnetic effect of the solar activity, which changes along the year.

A solar sail can generate both. However, a solar sail is generally used in higher orbits where there is no or neglectable atmospheric drag, and the atmospheric drag occurs closer to Earth, such as in LEO. The drawback of this method is that it highly depends on the Sun’s orientation. The method is not applicable for the altitude below 750 km because the atmosphere absorbs radiation. One of the most known examples of the solar radiation force method is the RemoveDEBRIS-Drag Sail mission. The method is promising, and advanced research has been conducted since it satisfies many debris removal aspects, such as minimal energy consumption via passive solar sails, compactness, and scalability for different satellite models planned to be launched in future. For space debris removal applications, the highest TRL ever reached with this method is TRL-6, achieved via the RemoveDEBRIS mission. The details of this mission are given in Section 4.

This method is based on applying a plasma ion beam onto the space debris and lowering its altitude (lower potential energy, in other words) over time. The electromechanical model of this method relies on the dynamics of ion actuation (Bombardelli and Pelaez, 2011). Outputs of this method with different ion actuation models are given in the work of Merino et al., 2011, and Merino et al., 2013. Fundamentally, the ion beam acts as a physical link between the chaser and target satellites, and when properly controlled, it works as a stable, spring-like mechanism to transfer the necessary forces on the debris for de-orbiting purposes. During the ion-beam actuation, the relative position stability between the chaser satellite and the debris can be ensured using the H ∞ optimal control framework (Alpatov et al., 2018).

This concept can be applied to non-cooperative and cooperative targets on both LEO and GEO orbits (Bombardelli et al., 2013; Kitamura et al., 2014). Cooperative targets can provide their location to be de-orbited by the chaser satellite, yet this method was developed to find a solution primarily for uncooperative targets. The velocity of ions is exceptionally high, around 38 km/s for xenon ions at 1-kV beam, and a higher power energy source can increase this value. The method is highly space debris topology dependent, and energy consumption requirement is hard to overcome. No ion-beam shepherd mission has been organized yet, and the concept has low TRL. The highest TRL ever reached for this method is TRL-4.

In this method, the chaser satellite deploys a harpoon mechanism to penetrate space debris (Dudziak et al., 2015) (Figure 4A). After the successful shot, the chaser satellite, harpoon, and target become integrated, the satellite and target are connected via an elastic tether, and the chaser satellite will pull the debris to re-enter the atmosphere and burn together or fly to a graveyard orbit. One disadvantage of this method is that it is tough to shoot an object tumbling at a high angular velocity in space. To achieve this, the chaser satellite needs to be supported by an advanced GNC infrastructure to get aligned and close-proximity rendezvous with the debris. The harpoon method is among the promising debris capturing methods, so there is ongoing research on this topic, both in simulation and experiment levels. An uncertain part of the simulation studies conducted for this method is the difficulty of proving the results since the simulation results may not be fully compatible with the experimental results. Furthermore, verification and validation processes of the product in space are needed (Forshaw et al., 2017). Some simulation studies conducted in this field include harpoon modelling (Forshaw et al., 2016), piercing geometry modelling (Pražák, 2021), pulling mechanism modelling, and GNC infrastructure support for harpoon systems (Surrey, 2022). The disadvantage of this method is that an unsuccessful shot may create more space debris depending on the debris material and targeted point, and there is only one chance to shoot (von Kries, 2002). The concept missions using harpoon technology are given in Review Technologies and Missions.

The net capturing method is based on shooting a flexible net onto the target debris. The net is shot from a chaser satellite by releasing corner masses outward in the target debris’s orientation. Therefore the net wraps the debris. This method has several advantages; for instance, the net capturing method enables shooting from a relatively long distance between chaser satellite and target debris such that close rendezvous and docking operations with the debris are not obligatory, and there is not necessary to de-tumble the target satellite before capturing (Botta et al., 2020). In other words, this method does not require a chaser satellite that has to achieve proximity rendezvous with precision. Moreover, the method is independent of the debris’s geometry, which means that the net can catch space debris in different scales (Bischof et al., 2003; Medina et al., 2017). The most common materials used for net design are Kevlar, Dyneema, Vectran, and Zylon (Zinner et al., 2011; Lavagna et al., 2012; Benvenuto and Carta, 2013). The disadvantage of this method is determining the elasticity of the net’s material and geometry of the net’s topology since these parameters will dramatically affect the first interaction dynamics (Bessonneau and Marichal, 1998; Tibert and Gardsback, 2005; Gao et al., 2012). Another disadvantage of this method is that the net can be shot only once.

In conclusion, net-based capturing methods in the literature are very similar since the main principle is to use a net to capture space debris. The only difference in the proposed studies is the nets’ geometric and material properties. Several concept missions have been developed for net capturing methods, and the research for the applicability of this method is kept active by different institutions. More detailed information about this method and related research are given in Review of Technologies and Missions. The highest TRL ever reached for this method is TRL-6.

The rigid capturing method can be evaluated under the impact energy dissipation category since as soon as the mechanical interaction is ensured between the chaser and target satellites, impact energy occurs on the chaser satellite. This operation is needed to achieve robust capturing because both objects are rigid, and any misalignment of the contact between the chaser and debris can push the debris far away from the chaser’s rigid capturing mechanism, especially in case where the object is tumbling at high angular velocity (Wu et al., 2018). Therefore, the rigid capturing method is more applicable for cooperative targets that have a docking port/mechanism (Somov et al., 2019). After completing the capturing phase, the chaser satellite can carry the target satellite anywhere in space, to the atmosphere or an outer orbit. Several studies vary from the simulation to experiment level, conducted to research the applicability of rigid robotic capturing systems for space debris mitigation problems. Any rigid capturing mechanism needs to be lightweight, easy to control, and compatible with the varying topology of space debris (Nishida et al., 2018). In the literature, there are several rigid robotic system designs from single arm to multiple arms (Yan et al., 2020), in which the control strategy of multiple arms to catch the debris is collaborative. Multiple arms are controlled by more sophisticated control system approaches, such as sliding mode control and adaptive control (Wang et al., 2021a), compared with control system approaches for single-arm robotic capture. For single-arm rigid capturing methods, the mostly well-known PID control approach is enough to achieve position and velocity control of the end-effector (Chen et al., 2020b). In some cases of multiple rigid robotic arms usage, one of the arms is used to catch the debris, whereas the other arm is used for manoeuvring purposes and creating rotational and translational movements in space (Yang et al., 2019). On another point, there is ongoing research for autonomy and teleoperation of rigid capturing mechanisms in space (Meng et al., 2018). However, most of these works cannot go beyond the theoretical level since it is tough to realize sophisticated rigid capturing scenarios in frictionless space conditions.

A tentacle-like structure was proposed by the ROGER project of the ESA (Liu et al., 2014). One of the scopes of the ROGER project was to design a rigid capturing system with a tentacle design to get the first interaction with uncooperative space debris in the geostationary orbit. ROGER aimed to create a basic research infrastructure for this capture mechanism. The geostationary orbit is critical, since many commercial and military telecommunication satellites exist in GEO. The goal was to carry the debris to the outer orbit after a successful catch. However, high TRL never reached the project since it was concept research. TRL-4 is achieved at the end of the project.

One problem of rigid capturing methods is to get a robust mechanical interaction with a fully rigid structure in space since rigidness creates a lack of appropriate impact energy dissipation in a friction-free environment. In other words, it is tough to catch a rigid object with a rigid structure in an environment that has no friction (Sallaberger et al., 1997; Rembala and Ower, 2009).

The flexible capturing method can be evaluated under the impact energy dissipation category since the capturing mechanisms of this method decrease the impact energy of the debris at the first contact. Several flexible capturing mechanisms fit with the scope of this method, such as shape memory alloy capturing mechanisms or pneumatic capturing mechanisms (Sinatra et al., 2019). Verification and validation processes of pneumatic capturing mechanisms have given satisfying results for on-ground conditions, yet the performance is still questionable for the zero-gravity condition of space. Moreover, the air consumption in space is critical and needs to be limited due to the necessary air consumption of the propulsion system. Therefore, the industry is inclined to use shape memory alloy-based capturing mechanisms as a flexible method in space. The actuation of this flexible capturing system is realized by the electrical voltage difference applied to the shape memory alloy material. The capturing mechanism made of shape memory alloy material fully complies with the debris during the capturing phase and reduces its impact energy. During this coherence, the chaser satellite absorbs the impact energy of the debris. After the full capture is ensured, the debris can be moved into the atmosphere or outer space. The recent advancement of shape memory alloys is becoming a hot topic in the material science discipline. The actuation of the mechanical structures made of this kind of material does not require a considerable high-power consumption. In other words, energy consumption is a critical matter for every space application, so these materials are excellent candidates for space debris capturing mechanisms since the structures that are made of shape memory alloys can easily be re-shaped by temperature or electrical potential gradient (Viscuso et al., 2021). The most advanced study conducted for this topic is MEDUSA. MEDUSA has arms that can grasp any object with flexible structures actuated by an electrical signal. When a simple electrical signal triggers the circuits consisting of nitinol wires of MEDUSA, the arms of MEDUSA start to change their topology and grasp the space debris (Feng et al., 2019). Several experiments were conducted in ground conditions for MEDUSA, and the results were satisfying in terms of robustness of the mechanical interaction. Another advantage of using shape memory alloy material for space applications is that since it can be controlled by either open loop or closed loop very effectively, the computational complexity of the control system is manageable (Lu et al., 2019; Wang et al., 2021b). No flexible capturing method-based mission has been organized yet, and the concept has low TRL, which is TRL-4. However, the method is promising since there is much ongoing research for the flexible capturing method, as given in the abovementioned references.

In addition to the previously mentioned concepts, some technologies are being used as part of the concepts classified as ET2, such as adhesive material technology or tentacle design architecture. For example, gecko-inspired adhesive pads or voltage-induced electrostatic adhesion props can be assembled on the end-effector of both rigid and flexible capturing mechanisms (DeLuca et al., 2013). The tentacle-like design may consist of multiple joint-link couples with rigid robotic arms or flexible robotic arms. Figure 5 shows examples of robotic grippers with adhesive technology.

One of the most troublesome tasks in space is to achieve a robust mechanical interaction between the chaser and target satellites. Grasping tasks using robotic arms/grippers are not easy, and the infrastructure needs to be supported by dedicated GNC hardware and software components. In this sense, adhesive materials have been proposed to hit and stick on the surface of the debris and slow down its velocity (Crane, 2017; Zhang et al., 2021). The first adhesive technology for ADR was proposed by Astroscale with the idea of using the technology to reduce the complexity of the mechanical interaction or to de-tumble the debris before the capturing processes are initiated (Astroscale, 2022).

The JPL of the NASA proposed another use of the adhesive technology. They developed a gecko-like adhesive material for uncooperative space debris (Modabberifar and Spenko, 2018; Onrobot, 2022). The tests of this product were conducted in parabolic flights and gave satisfactory results. However, despite the promising results, no mission that uses adhesive material has been organized yet. The technology has low TRL (TRL-4). Nevertheless, the horizon for this technology is very promising since the research of this type of material is ongoing for many different space applications. Moreover, any adhesive material type can be combined and integrated with different space-related material and structure types to synthesize hybrid concepts with higher TRL. For instance, an adhesive material can support shape memory alloys or tentacle-like structures.

When the magnetic coupling is generated between a ferromagnetic object and an electromagnetic coil, virtual spring and damping elements occur between two objects. By controlling the electromagnet using specific control algorithms, the coefficients of the damping and spring elements can be manipulated iteratively, which means that the transferred impact energy can be controlled iteratively (Liu et al., 2018; Yalçın et al., 2018; Yalçın and Erkan, 2021) since the damping is the dissipative energy element. In contrast, the spring is the energy storage element in control theory. The chaser can both push and pull the target, which means that by adjusting the sequence of the push/pull forces at high frequency, the impact of debris onto the chaser satellite will be minimized. One can say that if the chaser pushes the target, the target pushes the chaser as well due to Newton’s well-known third law (when two bodies interact, they apply forces to one another that are equal in magnitude and opposite in direction). However, this opposite reaction force can be compensated by the propulsion system of the chaser satellite. As the name of the category refers to neutral energy balance, the kinetic energy of both chaser and target satellites is balanced to minimize the impact energy. In other words, the electromagnetic coils play a perfect compliance mechanism role, even though it is not a physical compliance mechanism.

The magnetic force is only used for soft capture/soft docking of the debris when the close-proximity operation is achieved. The advantage of this method is the low possibility of crash and fracture since the virtual damping and spring elements generated by the magnetic field create a virtual cushion between two objects. However, the disadvantage of this method is that either debris needs to be cooperative with a ferromagnetic docking port or uncooperative with a ferromagnetic body.

In recent years, magnetic systems have been extensively used in many applications, from clean-room design to robotic capturing. The magnetic system of ELSA-d mission, given in Figure 6, uses the magnetic capturing method since the chaser satellite uses a unique magnetic coil to soft capture/soft docking the debris. ELSA-d consists of two satellites stacked together, a chaser designed to safely remove debris from the orbit and a target satellite that serves as a piece of mock-up debris, launched by GK Launch Services into a 550 km orbit on a Soyuz rocket from the Baikonur Cosmodrome in Kazakhstan on Monday, 22 March 2021, at 6:07 a.m. (UTC) (Astroscale, 2022).

The chaser is equipped with proximity rendezvous technologies and a magnetic docking mechanism, while the target, mimicking a defunct satellite, has a magnetic plate that allows docking. The chaser will repeatedly release and dock with the target in a series of technical demonstrations, proving the capability to find and dock with debris. Demonstrations include target search, target inspection, target rendezvous, and both non-tumbling and tumbling docking. The highest TRL ever reached is TRL-6.

This method aims at destroying small debris targets in case where the laser beam has enough energy (Esmiller et al., 2014). Examples of this method are given in Figure 7. For chaser satellites, one of the main problems in space is how to carry a battery that can produce enough energy to create a laser beam since the weight is a severe constraint for any space application. If the laser beam does not contain enough energy, or the battery is not capable of doing so, then to thoroughly burn the debris, the chaser satellite will change the debris’s kinetic or potential energy depending on the orientation of the laser beam. Therefore, the residual mass will keep on existing in the orbit. However, the industry is inclined to thoroughly burn the debris in space using a laser system for future applications. The State Key Laboratory of Laser Propulsion from China has been developing a space-based laser system, as assembled on a chaser satellite, that is capable of targeting debris whose diameter is up to 20 cm and consisting of mostly aluminium-, carbon-, and iron-based materials (Shen et al., 2014). They claim that, for an aluminium debris particle with a 10 cm diameter, the required laser pulse energy to destroy the debris is 1 kJ.

On the other hand, the current de-orbiting applications of laser focus on changing the orbiting trajectory of the debris by sending a laser beam from an earth-based station. An example of this concept is the Project ORION (Orbital Debris Removal Using Ground-Based Sensors and Lasers) (Phipps et al., 1996; Phipps et al., 1997). In ORION, a ground-based system to shoot the debris is proposed. According to the team, a laser destruction system combined with a space debris detection system can work for a varying distance of space debris, between 1 cm and 500 km.

Ground-based laser methods contain many negative aspects, such as the possible interference of other objects between the laser beam and debris. Additionally, the laser’s energy is not enough to melt the debris if the target is too big. In this case, the debris may be divided into several pieces or change its trajectory in an uncontrolled way. The Australian space company Electro Optic Systems is developing a pair of ground-based lasers to tackle small debris orbiting Earth (Securing Space, 2022). Generally speaking, there are several difficulties to overcome for this method, such as debris topology and material dependency, high energy consumption, or a possible collision/interfere risk with some other object in space. All these are crucial parameters for the chaser satellites orbiting in space. Currently, the technology has low TRL, which is TRL-4, and no laser system method-based mission has been organized yet (Phipps et al., 2019; Peltoniemi et al., 2021).

RemoveDEBRIS was a satellite research and development project to demonstrate different ADR technologies, such as the net, the harpoon, and the drag sail concepts (Aglietti et al., 2020; Forshaw et al., 2020). The Surrey Space Centre led the mission, and the satellite was manufactured by Surrey Satellite Technology Ltd. (SSTL). The platform was equipped with a net, a harpoon, a laser ranging instrument, a drag sail, and two CubeSats. Unique mechanisms were adapted to the platform to validate different ADR technologies. For example, to test the harpoon ADR concept, a target was assembled to the chaser satellite with a link; thus, it did not have a tumbling state, and no close-proximity operation was needed.

After end-to-end testing, the spacecraft was placed in an ISS cargo transfer bag and assembled with the CRS-14 SpaceX Dragon Spacecraft. The Dragon resupply mission with RemoveDEBRIS onboard was launched on April 2, 2018, arriving at the ISS on April 4, 2018. The deployment of the satellite took place on 20 June 2018. The satellite deployed a large sail, and the drag sail brought RemoveDEBRIS from the low orbital altitude of the space station into Earth’s atmosphere to burn.

In June 2021, an ADEO-N solar sail was mounted on the 220 kg “ION Platform” from D-Orbit (Dorbit, 2022), which was launched on a Falcon9 by SpaceX. The ION Platform is a space vehicle that can transport many satellites in an orbit and release them into distinct orbital points, reducing the time from launch. After 6 months of launch, the ADEO-N will open its 3.5 square metre drag sail, and the spacecraft will start to de-orbit. An in-orbit verification and validation of a bigger ADEO-L drag sail is planned for a future EU mission.

The NanoSail-D mission was launched onboard a Falcon-1 rocket in August 2008. The overall system contained a sail subsystem integrated into a CubeSat-2U for a main CubeSat 1U body. Due to budget limitations, no onboard camera or sensory infrastructure was installed to image the deployed sail or measure the attitude dynamics. Unfortunately, Falcon-1 rocket malfunctioned during the launch, and the NanoSail-D could not be deployed (Johnson et al., 2011). NanoSail-D was not dedicated to showing debris removal properties. Instead, it was planned to test the navigation capabilities of the sail.

The LightSail programme organized by the Planetary Society includes two missions, LightSail-1 and LightSail-2. LightSail-1 was launched on 20 May 2015, and the mission objectives were to test preliminary and straightforward CubeSat functionalities about the solar sail deployment process. The avionics part was assembled into a 1U volume, whereas the solar sail part was integrated into a 2U volume. After the success of the LightSail-1 mission, the LightSail-2 mission was planned to be organized to test more advanced scenarios, such as active control of the solar sail. The active control was achieved using a single-axis momentum wheel and magnetic torque tools. Compared with the NanoSail-D mission, LightSail missions have sophisticated sensory and actuation infrastructures, such as onboard cameras, momentum, and magnetic actuation parts. The Planetary Society announced that LightSail-2 mission success on 31 July 2019 (Spencer et al., 2021). LightSail-2 was not dedicated to showing debris removal properties. Instead, it was planned to test the navigation capabilities of the sail.

The InflateSail mission was a preliminary part of the RemoveDEBRIS-Drag Sail mission. The InflateSail mission was organized by the Surrey Space Centre and the Von Karman Institute. The total weight of the 3U CubeSat was 3.2 kg and equipped with a 1 m-long inflatable mast and a 10 m ² drag sail. InflateSail successfully demonstrated the de-orbiting process of a CubeSat in LEO for the first time using European drag sail technologies (Underwood et al., 2019). It was launched at 505 km altitude on 23 June 2017 and de-orbited 72 days after the deployment of the drag sail.

The DEOS (Deutsche Orbitale Servicing) mission can be evaluated under the impact energy dissipation category since the idea was to use a rigid robotic arm to achieve interaction with the target satellite and manipulate it in the orbit (Posch et al., 2012). This operation requires intensive impact energy dissipation at the first contact. DEOS was organized by DLR (German Space Agency). It was a mission to practice conducting tasks, such as maintenance or debris removal. DEOS consisted of two satellites, a client and a servicer. The client acts as the satellite requiring maintenance or removal. The servicer carries out the duty on the client and uses a robotic manipulator controlled remotely (Rank et al., 2011). The two satellites were to be launched together and brought into the orbit at 550 km height. DLR started the DEOS mission in 2012, and three phases were completed: phase 0- concept phase, exploring different systems and mission concepts, phase A- feasibility phase, where the technical aspects and uncertainty risks were researched, and phase B- preliminary design definition phase, where technical requirements of the system were defined. DEOS was expected to be launched in 2018, but the project was cancelled after phase B because of technical difficulties.

RemoveDEBRIS was deployed in June 2018 and focused on demonstrating different ADR technologies.

In 2012, the ESA launched the Clean Space Initiative to tackle the environmental impact of the space debris problem (Cieslak et al., 2019; Henry et al., 2019). Since the beginning, a central focus of Clean Space has been decreasing space debris through “design for destroying.” As time goes by, the ESA Clean Space Initiative has focused on both in-orbit servicing and ADR topics and started to work on the e. Deorbit mission. In the e. Deorbit mission, a 1,600 kg spacecraft was to be launched into an orbit at an altitude of 800 + 1,000 km (LEO). Once in the orbit, the spacecraft would rendezvous with the defunct satellite Envisat which is in an unknown condition, uncooperative, and probably tumbling at high angular velocity.

Within the preparation phase of the e. Deorbit mission, several concepts were analyzed and studied, 1) a robotic arm-based solution that catches the target and 2) a flexible link-based solution (capturing the target using a net or a harpoon). From this point of view, the e. Deorbit mission can be evaluated under the impact energy dissipation category since all methods include impact energy dissipation requirement at the first interaction. The launch was planned for 2025 onboard a Vega launch vehicle. However, the funding of the mission stopped in 2018. Instead, ESA member states now focus on the ClearSpace-1 mission, which is now under the development phase.

In 2019, the ClearSpace start-up was selected by the ESA to lead the first ADR mission by 2025 (Clear Space, 2022). In 2025, ClearSpace will launch the first ADR mission called ClearSpace-1. It will rendezvous, capture, and de-orbit to burn up in the atmosphere the upper part of a Vespa (Vega Secondary Payload Adapter), which was left in a “gradual disposal” orbit (altitude 801 km by 664 km). There will be no de-tumbling or separate docking mechanism. The rigid robotic capturing mechanism will include tentacles for the first mechanical interaction with the space debris. The tentacles will encircle the debris and align with it before dragging it into the lower levels of the atmosphere.

ClearSpace-1’s Vespa target is 112 kg (classified as a small satellite). Its compact shape and homogeneous geometry make it an appropriate first step before progressing to larger, more advanced captures by the subsequent missions. The future goal of the ClearSpace-1 mission is to achieve multiple capturing operations in a single mission. The ClearSpace-1 mission will be launched into a 500 km LEO for initial feature tests before moving to the target orbit for rendezvous and capture. In the end, the chaser combined with the Vespa target will de-orbit and burn in the atmosphere.

The mission will validate an innovative soft-capture/soft-docking mechanism, as well as the CONOPS (Concept of Operations) for capturing and removing non-tumbling and tumbling targets from the orbit. The target satellite has a DP (docking plate) that can be used to interact with the magnetic coil of the chaser satellite. Moreover, the chaser satellite has an autonomy mode, which means that the chaser satellite is able to operate either by itself or receiving commands from the ground. If needed, different algorithms can be uploaded from the ground station to the chaser satellite during the operation phase. The chaser minisatellite’s weight is 180 kg, and the target microsatellite’s weight is 20 kg.

The target satellite incorporates S-band communications, GPS positioning, a 3-axis control system, and a laser retro-reflector. The constellation platform is designed for 5–100 kg range operational missions. It also carries an HD camera and lighting to record the capture sequences during the eclipse in the absence of light. The ELSA-d mission was launched by GK Launch Services into a 550 km orbit on a Soyuz rocket from Kazakhstan on 22 March 2021, at 6:07 a.m. (UTC), and still is in operation phase. The data flow, such as odometry, energy consumption, and trajectory data flows, is being realized between the satellite and ground station. Nevertheless, the data recorded by the embedded devices will be investigated during the post-operation phase after the mission ends. The outputs of the mission have not been shared with the public yet.

In May 2021, Astroscale signed a £2.5 M agreement for research and development of space debris removal technology innovations with OneWeb (Astroscale, 2022). The communication infrastructure of the Elsa-m mission will be supported by OneWeb and will be based on 5G technology. Within ELSA-m, the technology to tackle multiple retired satellites in a single mission will be verified and validated, and it is claimed that the technology will be ready to use by 2024. This multi-client strategy will reduce launch and operation costs. Each constellation satellite will carry a passive magnetic docking port (DP) consisting of permanent physical magnets. The DP will be used to create a weak magnetic attraction force for the re-assembly of the chaser and target satellites, while the chaser satellite’s magnetic coil system generates a magnetic field. Sophisticated control algorithms can drive the magnetic coil of the chaser satellite to create the appropriate magnetic field patterns. These control algorithms can be uploaded to the chaser satellite from the ground station even when the operation is on. In this way, the chaser satellite will catch multiple satellites in the constellation.