Deorbit Maneuver Research Articles

De-Orbit Maneuver Demonstration Results of Micro-Satellite ALE-1 with a Separable Drag Sail

ALE-1, a micro-satellite created for the demonstration of artificial shooting stars, required orbital descent before mission execution due to safety aspects in orbit. ALE-1 utilized a drag sail called SDOM (Separable De-Orbit Mechanism) for a passive de-orbit maneuver, which was successfully completed, lowering the orbit from about 500 km down to about 400 km. This paper summarizes the detailed history of satellite operation and the results of the de-orbit maneuver demonstration during the past three years. Although the SDOM sail faced difficulty in keeping the desired deployed shape of the drag sail due to mechanical troubles, by letting the sail be a drag flag instead, it could still deliver a meaningful de-orbit performance to allow the satellite to successfully lower the orbit as planned. The de-orbit effect of the drag flag was evaluated using comparisons between orbit propagation simulations and the actual orbit transition flight data provided in the form of TLE (Two-Line Element) sets. Through this study, it is demonstrated that the SDOM can provide orbit transfer capabilities for satellites. Furthermore, the de-orbit performance of the drag flag can be evaluated, which could be an important reference for the future implementation of de-orbit devices to solve space debris problems.

Optimization of aeroassisted rendezvous and interception trajectories between non-coplanar elliptical orbits

Aeroassisted orbital maneuver can potentially reduce fuel consumption, compared with the traditional impulse maneuver. However, studies on aeroassisted orbital rendezvous and interception are limited to maneuvering between circular orbits. In this study, the optimal aeroassisted rendezvous and interception with elliptic initial and target orbits are investigated using the hp-adaptive pseudospectral method. Cases with minimum fuel and minimum time are considered. The initial state constraint at the deorbit position for an elliptical initial orbit is rederived first under the premise that the waiting time before the deorbit maneuver is an optimization variable. In addition, other interior constraints, including the constraints at the atmospheric entry and rendezvous/interception position, are derived. Considering the heating-rate and load factor constraints, numerical results demonstrate that the aeroassisted rendezvous and interception problems can be processed and easy to converge for optimization when elliptical initial and target orbits are considered. Moreover, comparative results with the traditional pure-impulse indicate that aeroassisted rendezvous and interception can reduce fuel consumption and adjust the rendezvous/interception time and phase with certain constraints.

Design and Operational Elements of the Robotic Subsystem for the e.deorbit Debris Removal Mission.

This paper presents a robotic capture concept that was developed as part of the e.deorbit study by ESA. The defective and tumbling satellite ENVISAT was chosen as a potential target to be captured, stabilized, and subsequently de-orbited in a controlled manner. A robotic capture concept was developed that is based on a chaser satellite equipped with a seven degrees-of-freedom dexterous robotic manipulator, holding a dedicated linear two-bracket gripper. The satellite is also equipped with a clamping mechanism for achieving a stiff fixation with the grasped target, following their combined satellite-stack de-tumbling and prior to the execution of the de-orbit maneuver. Driving elements of the robotic design, operations and control are described and analyzed. These include pre and post-capture operations, the task-specific kinematics of the manipulator, the intrinsic mechanical arm flexibility and its effect on the arm's positioning accuracy, visual tracking, as well as the interaction between the manipulator controller and that of the chaser satellite. The kinematics analysis yielded robust reachability of the grasp point. The effects of intrinsic arm flexibility turned out to be noticeable but also effectively scalable through robot joint speed adaption throughout the maneuvers. During most of the critical robot arm operations, the internal robot joint torques are shown to be within the design limits. These limits are only reached for a limiting scenario of tumbling motion of ENVISAT, consisting of an initial pure spin of 5 deg/s about its unstable intermediate axis of inertia. The computer vision performance was found to be satisfactory with respect to positioning accuracy requirements. Further developments are necessary and are being pursued to meet the stringent mission-related robustness requirements. Overall, the analyses conducted in this study showed that the capture and de-orbiting of ENVISAT using the proposed robotic concept is feasible with respect to relevant mission requirements and for most of the operational scenarios considered. Future work aims at developing a combined chaser-robot system controller. This will include a visual servo to minimize the positioning errors during the contact phases of the mission (grasping and clamping). Further validation of the visual tracking in orbital lighting conditions will be pursued.

TugSat: Removing Space Debris from Geostationary Orbits Using Solar Sails

Space debris accumulation is on the rise as commercial space applications become more affordable and existing satellites near the end of their operational lifetimes. For geosynchronous equatorial orbits, solar radiation pressure provides a unique solution for the removal of larger debris such as inoperable satellites. Given the current state of the art in solar sailing technology, a CubeSat can be tasked to deorbit satellites on the order of 1000 kg using a high-performance solar sail for propulsion. This CubeSat, called “TugSat,” is simulated in this study, virtually deorbiting a satellite from geostationary orbit without the use of standard propulsion systems. This same TugSat can be reused indefinitely, between the geosynchronous equatorial orbit belt and the retirement orbit, to continuously remove debris from valuable geostationary orbit slots. The entire deorbit maneuver will demonstrate control of a satellite’s semimajor axis, eccentricity, inclination, and geosynchronous equatorial orbit belt longitude: all using techniques to optimize time rates of change in the satellite’s orbital elements. The TugSat concept provides a low-cost reusable means for geosynchronous equatorial orbit belt cleanup or open-ended geosynchronous equatorial orbit belt mission applications.

Environmental effect of space debris repositioning

One of the proposed measures to limit the number of near-Earth orbiting fragments to a sustainable level is to actively remove large derelict objects from crowded orbital regions. The two main removal procedures considered so far are (1) a direct targeted reentry maneuver or (2) a deorbit maneuver resulting in a predicted 25-year lifetime for the target object. We study here the viability of a third option, which consists of repositioning the target to an optimally chosen altitude according to a selected benefit/cost objective function. The objective function accounts for both the maneuver cost and the reduction of environmental criticality of the object. Numerical simulations are conducted to determine the optimal sequence of repositioning maneuvers for a given available deorbiting propellant. Results show that an optimal repositioning campaign tends to displace ton-class objects from around 900–1000km altitude down to around 750–800km altitude and to redistribute debris mass from altitudes around 1500km across lower density nearby altitudes. Comparisons with a 25-year lifetime deorbiting suggest a significant performance improvement.

Simulation of a tethered microgravity robot pair and validation on a planar air bearing

A software model has been developed to simulate the on-orbit dynamics of a dual-mass tethered system where one or both of the tethered spacecraft are able to produce propulsive thrust. The software simulates translations and rotations of both spacecraft, with the visco-elastic tether being simulated as a lumped-mass model. Thanks to this last feature, tether longitudinal and lateral modes of vibration and tether tension can be accurately assessed. Also, the way the spacecraft motion responds to sudden tether tension spikes can be studied in detail. The code enables the simulation of different scenarios, including space tug missions for deorbit maneuvers in a debris mitigation context and general-purpose tethered formation flight missions. This study aims to validate the software through a representative test campaign performed with the MIT Synchronized Position Hold Engage and Reorient Experimental Satellites (SPHERES) planar air bearing system. Results obtained with the numerical simulator are compared with data from direct measurements in different testing setups. The studied cases take into account different initial conditions of the spacecraft velocities and relative attitudes, and thrust forces. Data analysis is presented comparing the results of the simulations with direct measurements of acceleration and Azimuth rate of the two bodies in the planar air bearing test facility using a Nylon tether. Plans for conducting a microgravity test campaign using the SPHERES satellites aboard the International Space Station are also being scheduled in the near future in order to further validate the simulation using data from the relevant operational environment of extended microgravity with full six degree of freedom (per body) motion.

Optimal Design and Deorbiting Performance of Thermionic Tethers in Geostationary Transfer Orbits

The application of a recent concept, a thermionic bare tether (that is, a long conductor coated with a thermionic material), to a practical engineering problem (deorbit space debris from geostationary transfer orbit) is presented. Lorentz drag on a thermionic bare tether, during each pass through an arc close to the perigee, lowers the apogee progressively and produces the object reentry. The performance of a spacecraft equipped with a thermionic bare tether is studied at two different levels, using models that couple thermal and electrical effects. In first place, a simple formula for the eccentricity decrement produced during each perigee pass as a function of TBT properties is derived and used to select TBT optimal dimensions. For a given tether mass, the formula shows that long tethers with small cross-section areas, but large enough to accomplish mechanical constraints, yield the best performance. Second, full numerical simulations of the deorbit maneuver including Lorentz force, air drag, and perturbations are carried out. A spinning thermionic bare tether with a mass of about 16 kg, as well as a length, width, and thickness equal to 6 km, 2 cm, and , respectively, passively deorbits a half-ton spacecraft (with a natural deorbit time of about 50 years) in less than six months, without using propellant, expellant, or power supply. The important roles played by the eclipses and the Earth’s oblateness on thermionic bare tether performance are highlighted.

Control of Space Debris Using an Elastic Tether and Wave-Based Control

One proposed solution to the space debris problem involves sending an active spacecraft (or “chaser”) to rendezvous with the target debris and to make a physical connection with it by means of a net or harpoon at the end of an elastic tether. The chaser then needs to control the debris motion via the elastic tether, for example, to achieve a controlled deorbit. Compared with robotic capture, the use of an elastic tether reduces the need for precise chaser–target coordination during capture, reduces the danger of collisions, and limits subsequent jerk with its potentially damaging consequences. However, target motion control via an elastic tether presents control challenges, for example, due to the radically different system dynamics, depending on whether the tether is taut or slack. Tether entanglement is also a danger, especially if the debris is tumbling. This research focuses on using wave-based control to have the chaser give the target an arbitrary change in the magnitude and direction of its velocity, while maintaining tether and debris control. It then considers a complete, controlled deorbit maneuver. To test the technique, ESA’s Envisat is used as a representative debris satellite.

Swing principle for deployment of a tether-assisted return mission of a re-entry capsule

Dynamics and control of a tether-assisted return mission of a re-entry capsule are considered. Efficiency of the braking process of the capsule depends on a deflection angle of the tether from a local vertical before separating the capsule from the tether. The aim is of this paper to find a tether length control law that allows to increase the deflection angle of the tether from the local vertical. The control law is based on a principle of a swing. This control law may be applied to the final phase to two possible options to perform a tether-assisted deorbit maneuver: static and dynamic release. An approximate analytical solution for the deflection angle from the local vertical is obtained for the control law. The numerical simulations have shown that the application of the control law allows a reduction in the required tether length of a tether-assisted deorbit maneuver. The proposed control law can be applied to develop new space tethered systems.

Multidisciplinary Design Optimization of a Deorbit Maneuver Considering Propulsion, TPS, and Trajectory

Unguided reentry capsules are usually involved in ballistic entry. The final states (such as altitude and velocity of parachute activity) depend on initials parameters. Reentry trajectory parameters differently affect the thermal protection system (TPS), required deorbit propellant, and structural load. The purpose of this paper is to optimal design of deorbit parameters to minimize the thermal protection system mass and deorbit propellant mass using multidisciplinary design optimization technique. Genetic Algorithm (GA) is used to simultaneously optimize TPS discipline, propulsion discipline and trajectory in present of the all trajectory and configuration constraints. To do this, every discipline is mathematically modeled. A suitable framework of Multidisciplinary design optimization (MDO) is developed. Then, the reentry mission is optimized according to the each discipline. The results show that simultaneous optimization is more efficient in comparison with single discipline optimization such as TPS optimization or deorbit propulsion system (propulsion and propellant) optimization.

Active debris removal of multiple priority targets

Today’s space debris environment shows major concentrations of objects within distinct orbital regions for nearly all size regimes. The most critical region is found at orbital altitudes near 800km with high declinations. Within this region many satellites are operated in so called sun-synchronous orbits (SSO). Among those, there are Earth observation, communication and weather satellites. Due to the orbital geometry in SSO, head-on encounters with relative velocities of about 15km/s are most probable and would thus result in highly energetic collisions, which are often referred to as catastrophic collisions, leading to the complete fragmentation of the participating objects. So called feedback collisions can then be triggered by the newly generated fragments, thus leading to a further population increase in the affected orbital region. This effect is known as the Kessler syndrome.Current studies show that catastrophic collisions are not a major problem today, but will become the main process for debris generation within the SSO region in the near future, even without any further launches. In order to avoid this effect, objects with a major impact on collisional cascading have to be actively removed from the critical region after their end of life. Not having the capability to perform an end-of-life maneuver in order to transfer to a graveyard orbit or to de-orbit, many satellites and rocket bodies would have to be de-orbited within a dedicated mission. In such a mission, a service satellite would perform a de-orbit maneuver, after having docked to a specific target.In this paper, chemical and electric propulsion systems were analysed with the main focus on removing multiple targets within one single mission. The targets were chosen from a previously defined priority list in order to enhance the mission efficiency. Total mission time, ΔV and system mass were identified as key parameters to allow for an evaluation of the different concepts. It was shown that it is possible to remove up to five high priority targets per year using a chemical propulsion system, however, missions may result in too high ΔV and/or mission duration depending on the orbital distribution of the targets. When using an electric propulsion system, the required fuel mass is significantly reduced when compared to the chemical propulsion system, but it was shown that mission duration strongly depends on the mass of the selected targets. More powerful engines as well as out-of-plane thrust are thus required to achieve the defined mission goals.

Optimal Deorbit Guidance

can be hundreds of seconds apart from that of the impulsive solution), duration of the burn, and thrust direction duringtheburn.Inthispaperthe deorbitguidanceproblem isformulatedasanoptimalcontrolproblemwith finitetime burn structure. Several sets of general targeting conditions are derived to be met at the entry interface. Numerical solutions are obtained by employing a fast multiburn shooting algorithm. The solution contains the coast times before and after a burn, duration of the burn, and thrust-direction profile during the burn. This study investigates both single- and two-burn deorbit maneuvers, and demonstrates how each works in different mission scenarios.

Concept and Technology of HTV-R: an Advanced Type of H-II Transfer Vehicle

The first and second H-II Transfer Vehicle (HTV) successfully completed their missions in Nov. 2009 and March 2011. JAXA got the capability to upload lots of logistics to support human activities on the International Space Station (ISS). Mid of these HTV missions, JAXA established ISS Cargo Return Vehicle Research and Development Office as a new section to study a new space vehicle concept. It is called and an improved type of HTV to enable JAXA to recover various samples of experiments conducted on the ISS. HTV-R equips a new module which is called HRV(HTV Return Vehicle) for the enhanced purpose. Most part of HTV-R mission is not changed from original HTV's, but once the HRV separates from the HTV-R after the de-orbit maneuver completion, the HRV will fly by itself and conduct lifting reentry to the atmosphere with controlling attitude and trajectory to the predefined splashdown point. Because HRV flight profile is similar to human returning flight, JAXA is considering that HTV-R is not only as a cargo recovery system but also as the pathfinder of a human transport system.

Operational use of guided reentry capsules – system design solutions and mission safety considerations

Future utilization of the experimental facilities on-board the International Space Station (ISS) exhibits a demand for frequent sample return opportunities that can be provided by small, unmanned but controllable re-entry capsules. Based on the mission and system requirements for payload retrieval, a conceptual vehicle design down to subsystem level is presented. The vehicle is of a sphere-cone shape with an oblate bottom surface and equipped with a body flap. This configuration was established by means of extensive numerical analyses. A tether-assisted deorbit maneuver is considered, which yields significant mass savings over a conventional propulsive approach if applied in an operational scenario. For reentry guidance, which is a key technology for this type of vehicle, an explicit near-optimal guidance law is presented and its performance evaluated. Furthermore, mission safety issues in case of GNC system failures are addressed. Six-degree-of-freedom analyses simulating the loss of the control capability during re-entry flight indicate that the vehicle is capable to drastically reduce landing dispersions and thus to improve safety on ground, if proper on-board emergency procedures are implemented.

Guidance and Control System Design of the Viking Planetary Lander

This paper describes the salient features and unique problem solutions for the Viking lander guidance and control system design. Because of the following design requirements imposed, the lander guidance and control function presented stringent design challenges. These requirements were as follows: 1) a completely autonomous operation was required although the onboard computer memory size was limited to solve the descent problem; 2) the lander deorbit maneuver had to be very accurate to achieve a satisfactory planet atmospheric entry and an accurate touchdown at the predicted landing site; 3) the landers had to descend through an unknown atmosphere with a large variation in predicted characteristics, and they had to land at an unknown terrain height; and 4) they had to be as fuel-efficient as possible, particularly in the terminal descent phase. This paper describes how a safe and accurate landing on the surface of Mars while meeting these design requirements was achieved by the guidance and control systems of both Viking landers.