This work presents a design for a stand-alone Mars exploration mission utilizing a 12U CubeSat, showcasing the integration of DLR's in-house technologies. Focused on demonstrating the capabilities of miniaturized satellites in deep space, this innovative 4-year mission will travel independently to Mars propelled by a low-thrust propulsion system. Upon reaching Mars, the spacecraft will insert itself into a highly elliptical orbit, transitioning to a Primary Science Orbit (PSO) at 250 km altitude through aerobraking. This orbit is strategically chosen for its sun-synchronous and near-circular properties, optimizing scientific operations aimed at studying Mars' lower atmosphere and gravity field. The CubeSat features a 20.8 kg wet mass, a 6.3 km/s maneuvering capability, and can generate up to 90 W of power. It is equipped with a 2U infrared spectrometer, a 1U gravimeter, and a 12 Mpx CMOS camera for scientific data collection. Utilizing DLR's technology, including an integrated avionics stack combining communications, power, and onboard computer subsystems, the mission seeks to advance the CubeSat platform for interplanetary use, significantly reducing costs and fostering future exploration opportunities.

Mars is expected to become a focal point of exploration (human and robotic) in the coming century, with a very likely need for a robust space infrastructure. Be it communication and navigation satellite constellations or scientific missions in low Mars orbits (LMO) and Areosynchronous orbits (ASO), every individual satellite will have a definitive period of operation after which it becomes derelict. At the end-of-life (EOL) the satellite shall be proactively dealt with in a sustainable manner to protect our access to the space environment of Mars and opportunities to use this. Clearly, impacting Mars or escaping Mars’ gravity are no viable options. This paper aims at identifying graveyard orbit solutions in circummartian space for future Mars space debris. Orbital stability for a period of 200 years is studied for Martian orbits using the symplectic integration technique. Extensive validations are performed and propagation and integration settings are tuned to suit a variety of configurations. A plethora of candidate graveyard orbit solutions are found and presented for orbits in the ASO and LMO regimes. For example, it is found that transferring an ASO satellite to 400 km below the nominal orbit altitude would ensure a stability margin of ±25 km for at least 200 years. Multiple orbital geometry characteristics (combinations of semi-major axis, inclination, right ascension of ascending node), satellite geometries (various values of area-to-mass ratio) and uncertainties are studied to produce a comprehensive analysis of long-term stability of potential graveyard orbits around Mars, making them attractive for such purposes. The protected zones are found to be safe from debris even for an uncertainty in initial eccentricity of 0.01 and variations in cross-sectional area due to uncontrolled tumbling. The overall objective of this paper is to make designers of future missions to Mars aware of the EOL aspects and include this in their mission design proposals at an early stage already.

Due to ever increasing accessibility, recent years have seen a fast growing number of launches to space, especially to Sun-synchronous orbit. The spent rocket parts, and eventually non-functioning payloads of these launches remain in orbit. It is well established that this accumulation of space debris over time is quickly making this the most severe threat to future spaceflight operations. To address this, international guidelines have been established including a maximum of 25-year remaining orbital lifetime after end of operational life. This paper evaluates if Sun-synchronous satellites adhere to this guideline. To determine the compliance, the operational status of satellites with orbital control capabilities is established using a maneuver detection algorithm. For satellites without this capability, a model is created based on mass and design lifetime. The remaining orbital lifetime is determined using semi-analytic propagation. The results reveal that compliance was poor in the past, with 20 to 40% prior to 2014, but has increased to 95% in 2018. Satellites with a mass lower than 10 kg have a compliance of 86% compared to 35% for heavier satellites. Analysis shows that compliance is mostly a result of choosing an operational orbit with a sufficient natural decay, and less due to altitude lowering maneuvers near end-of-life. The relative popularity of SSO may demand re-evaluation of current guidelines to sustain future operations in these valuable orbits.

This research investigates the performance of a space-based laser system to remove debris objects with size smaller than 10 cm. The laser system is placed in a 800 km Sun Synchronous Orbit and consists of a 20 kW laser that shoots 300 J energy pulses with a repetition frequency of 66.66 Hz. The system is able to detect and track debris objects in situ using a 2.0 m mirror from 800 km distance. From a distance of about 500 km, the laser fluence on the targets is sufficiently high to trigger ablation, which decelerates the debris objects and reduces their lifetime. The feasibility of the concept is tested in scenarios where the laser system targets the debris objects from a different orbiting altitude and from varying azimuth angles. For many geometries, the laser is capable of significantly reducing the lifetime of the debris object. Extrapolating to longer periods of operation, the laser can be expected to provide a significant reduction of the population of small debris objects in LEO.

Orbital mission analysis is an iterative procedure in which several solutions are tested until the one that better fulfils mission objectives is selected. With new missions in the Circular Restricted Three-Body Problem (CR3BP) becoming more frequent, the mission analysis process for trajectories within this model must be improved. So far, most of the research efforts have been devoted to finding the most accurate optimization algorithm for a particular problem, mainly related to the Gateway mission. However, this study presents a novel approach whose aim is to improve the initial phase of mission design by providing a multitude of optimal solutions that cover most of the trajectory possibilities. Moreover, this should be performed without an initial guess input from the user, nor any kind of previously known solution, so that it can be applied to all kinds of problems. To do so, the proposed research performs a multi-objective optimization on direct, manifold and flyby transfers and places the non-dominated solutions in a Pareto front. In this way, the user can easily choose which solution better meets their requirements in terms of ΔV and time of flight. Once the solution is chosen, it can used as a first guess for a further optimization with the Analysis, Simulation and Trajectory Optimization Software (ASTOS), which increases the accuracy and reliability of the results, by being verified in a higher fidelity model. To achieve this purpose, the tool has been carefully designed, selecting the design variables that with the least amount of information, completely define the full trajectories. Then, a mixed integer distributed ant colony optimizer (MIDACO) is used to find the best solutions. The results obtained include two sample cases. First a LEO to L2 southern Halo orbit and then a GTO to 9:2 resonant L2 NRHO, all in the Earth-Moon system. For all cases, a populated Pareto front with dozens of optimal solutions, was obtained and a single trajectory solution was selected for representation purposes.

The ability to reduce the cost of space missions beyond Earth orbit by leveraging innovative concepts is of great interest in the field of spaceflight. In this paper, the trajectory design for a mission from an inclined geosynchronous transfer orbit (GTO) to an Earth-Moon L2 Halo orbit is presented. The mission scenario for this investigation involves the use of a small satellite launched in a rideshare configuration and the use of a low-energy transfer to reach the target orbit. As a consequence of choosing a rideshare launch, the mission scenario entails critical uncertainties on the time of launch and injection parameters of the spacecraft, which could complicate the insertion into a low-energy transfer. Thus, the goal of the project is to develop a robust design methodology to deal with the launch uncertainties and assess launch readiness at any time of the year.

Inspired by the Keplerian Map and the Flyby Map, a Gravity Assist Mapping using Gaussian Process Regression for the fully spatial Circular Restricted Three-Body Problem is developed. A mapping function for quantifying the flyby effects over one orbital period is defined. The Gaussian Process Regression model is established by proper mean and covariance functions. The model learns the dynamics of flyby's from training samples, which are generated by numerical propagation. To improve the efficiency of this method, a new criterion is proposed to determine the optimal size of the training dataset. We discuss its robustness to show the quality of practical usage. The influence of different input elements on the flyby effects is studied. The accuracy and efficiency of the proposed model have been investigated for different energy levels, ranging from representative high- to low-energy cases. It shows improvements over the Kick Map, an independent semi-analytical method available in literature. The accuracy and efficiency of predicting the variation of the semi-major axis are improved by factors of 3.3, and 1.27×10⁴, respectively.

We develop a Gravity Assist Mapping to quantify the effects of a flyby in a two-dimensional circular restricted three-body situation based on Gaussian Process Regression (GPR). This work is inspired by the Keplerian Map and Flyby Map. The flyby is allowed to occur anywhere above 300 km altitude at the Earth in the system of Sun-(Earth+Moon)-spacecraft, whereas the Keplerian map is typically restricted to the cases outside the Hill sphere only. The performance of the GPR model and the influence of training samples (number and distribution) on the quality of the prediction of post-flyby orbital states are investigated. The information provided by this training set is used to optimize the hyper-parameters in the GPR model. The trained model can make predictions of the post-flyby state of an object with an arbitrary initial condition and is demonstrated to be efficient and accurate when evaluated against the results of numerical integration. The method can be attractive for space mission design.

Verified interval orbit propagation provides mathematically guaranteed solutions of satellite position and velocity over time. These verified solutions are useful for conjunction analysis and other space-situational-awareness activities. Unfortunately, verified methods suffer from overestimation and explosive interval growth, limiting the possible propagation time and thus their applicability. Different orbital-element state models have been shown to increase the maximum propagation time to a degree, but at the expense of significant overestimation introduced by the state transformations. This paper proposes the Dromo state model for verified integration. Dromo is a regularized variation-of-parameter formulation of the perturbed two-body equations of motion. Taylor models are implemented for both integration and transformation. Moreover, a technique is developed for dealing with time uncertainty resulting from verified regularized propagation. Dromo significantly prolongs the maximum forecasting window, producing verified trajectories of days up to weeks in duration for the low Earth orbit regime. A sensitivity analysis of the integrator settings identifies combinations that produce stable and computationally efficient solutions. A sensitivity study of the orbital parameters shows that the method is applicable to a large orbital regime, especially for low Earth orbit regions that contain high densities of space debris.

This paper presents the theory for linear cotangential transfers and safe orbits for elliptic orbit rendezvous. Expressions for the transfer angle and the required ΔV’s are derived. Singularities in the algorithm can occur if the two orbits intersect. Alternative maneuvers for such singular cases are developed. The linear cotangential transfer algorithm is compared with the nonlinear cotangential transfer and the algorithm is found to be very similar. The development of the linear cotangential transfer leads to a new set of relative orbital elements that are well suited for defining safe trajectories. The characteristics of safe trajectories are discussed and a linear safety checking algorithm is developed. Finally, the combination of the cotangential transfers and safe orbits is used to define safe rendezvous trajectories for elliptical orbit rendezvous.

The Lunar Meteoroid Impact Observer (LUMIO) is one of the four projects selected within ESA’s SysNova competition to develop a small satellite or scientific and technology demonstration purposes to be deployed by a mothership around the Moon. Themission utilizes a 12U form-factor CubeSat which carries the LUMIO-Cam, an optical instrument capable of detecting light flashes in the visible spectrum to continuously monitor and process the meteoroids impacts. In this chapter, we will describe the mission concept and focus on the performance of a novel navigation concept using Moon images taken as byproduct of the LUMIOCam operations. This new approach will considerably limit the operations burden on ground, aiming at autonomous orbit-attitude navigation and control. Furthermore, an efficient and autonomous strategy for collection, processing, categorization, and storage of payload data is also described to cope with the limited contact time and downlink bandwidth. Since all communications have to go via a lunar orbiter, all commands and telemetry/data will have to be forwarded to/from the mothership. This will prevent quasi-real-time operations and will be the first time for CubeSats as they have never flown without a direct link to Earth. This chapter was derived from a paper the authors delivered at the SpaceOps 2018 conference.

Satellite reentry predictions are used to determine the time and location of impacts of decaying objects. These predictions are complicatedby uncertainties in the initial state and environment models, and the complex evolution of the attitude. Typically, the aerodynamic and error propagation are done in a simplistic fashion. Full six-degrees-of-freedom modeling and attitude control is proposed for studying the historic reentry case of the Gravity Field and Steady-State Ocean Circulation Explorer satellite. Improved error modeling and estimation of the initial state and atmospheric density are introduced for both Global Positioning System and two-line elements states. A sensitivity analysis is performed to identify the driving parameters for several models and epochs. The predictions are compared against Tracking And Impact Predictions, and predictions by the European Space Agency Space Debris Office. The performed predictions are consistently closer to the true decay epoch for several starting epochs, while providing narrower windows than other predictions with higher confidence.

At present, tracking data for planetary missions largely consists of radio observables: range-rate (Doppler), range and angular position (VLBI/Δ DOR). Future planetary missions may use Interplanetary Laser Ranging (ILR) as a tracking observable. Two-way ILR will provide range data that are about 2 orders of magnitude more accurate than radio-based range data. ILR does not produce Doppler data, however. In this article, we compare the relative strength of radio Doppler and laser range data for the retrieval of parameters of interest in planetary missions, to clarify and quantify the science case of ILR, with a focus on geodetic observables. We first provide an overview of the near-term attainable quality of ILR, in terms of both the realization of the observable and the models used to process the measurements. Subsequently, we analyse the sensitivity of radio Doppler and laser range measurements in representative mission scenarios for parameters of interest. We use both an analytical approximation and numerical analyses of the relative sensitivity of ILR and radio Doppler observables for more general cases. We show that mm-precise range normal points are feasible for ILR, but mm-level accuracy and stability in the full analysis chain are unlikely to be attained, due to a combination of instrumental and model errors. We find that ILR has the potential for superior performance in observing signatures in the data with a characteristic period of greater than 0.33–1.65 hours (assuming 2–10 mm uncertainty for range and 10 μ m/s at 60 s for Doppler). This indicates that Doppler tracking will typically remain the method of choice for gravity field determination and spacecraft orbit determination in planetary missions. ILR data will be able to supplement the orbiter tracking data used for the estimation of parameters with a once-per-orbit signal. Laser ranging data, however, are shown to have a significant advantage for the retrieval of rotational and tidal characteristics from landers. Similarly, laser ranging data will be superior for the construction of planetary ephemerides and the improvement of solar system tests of gravitation, both for orbiter and for lander missions.

Space debris is a growing problem for modern spaceflight. This holds in particular for spacecraft in Low Earth Orbit (LEO) and in Geostationary Orbit (GEO). In order to control or even reduce the problem in GEO, the Inter-Agency Space Debris Coordination Committee has proposed that when at their end of life, GEO satellites should be moved to a safe graveyard orbit which is several hundreds of kilometres above the nominal GEO altitude. This rule is internationally acknowledged, and accepted by the United Nations as a guideline. This study investigates the rate of compliance to this guideline. In particular, the satellite ephemeris data provided by the US Space Surveillance Network is used to analyse the lifetime orbital behaviour of all relevant objects, and to assess the state of compliance for any of the spacecraft. This is done by inspecting the semi-major axis, eccentricity, inclination and longitude position of individual satellites over the time frame from 1977 until 2016. In total, almost 1000 spacecraft have been reviewed. Although the developed algorithm with which the status of individual objects is established is open to minor flaws occasionally, it can be concluded that a statistical analysis only as function of the orbital elements of GEO satellites can be done and relevant conclusions can be drawn about their state of compliance with the mitigation regulations. The overall conclusion is that the compliance rate is considerably low for satellites launched in the late 1970s (below 40%), but that this number has been on the rise ever since and is approaching 70% for spacecraft launched in 1992. The compliance rates for later launches cannot be assessed since (a large) part of the missions is still active. However, in view of the observed trend the compliance rate can be expected to reach 100% for spacecraft that have been launched recently.