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To design a free-return orbit for manned lunar mission based on low earth orbit (LEO) rendezvous, an adaptive LEO-phase free-return orbit design method based on high-precision dynamics model is proposed. First, the radius of perilune and the absolute value of perilune velocity are decoupled using a coordinate system rotation, which is derived from moon-centric local vertical and local horizontal instantaneous coordinate system at the time of perilune. The two Euler rotation angles and the absolute value of perilune velocity are used as independent design variables because their initial values are easy to guess. Next, a two-segment numerical integration strategy is proposed to calculate orbital elements at the moments of trans-lunar injection and free-return vacuum perigee. Subsequently, an optimization algorithm software package for solving large-scale nonlinear sequential quadratic programming problems (SQP_snopt) is employed to search the objective free-return orbit with a fixed trans-lunar injection inclination and the other two constraints on radiuses of perigee at the times of trans-lunar injection and vacuum perigee. After that, an iteration algorithm is devised to adjust trans-lunar injection window for adaptive LEO-phase. Finally, numerical results show a fast and accurate performance of the direct optimization method, which can provide valuable references to manned lunar missions based on LEO rendezvous.

In view of the characteristics of the guidance, navigation and control (GNC) system of the lunar orbit rendezvous and docking (RVD), we design an auxiliary safety prediction system based on the human–machine collaboration framework. The system contains two parts, including the construction of the rendezvous and docking safety rule knowledge base by the use of machine learning methods, and the prediction of safety by the use of the base. First, in the ground semi-physical simulation test environment, feature extraction and matching are performed on the images taken by the navigation surveillance camera. Then, the matched features and the rendezvous and docking deviation are used to form training sample pairs, which are further used to construct the safety rule knowledge base by using the decision tree method. Finally, the safety rule knowledge base is used to predict the safety of the subsequent process of the rendezvous and docking based on the current images taken by the surveillance camera, and the probability of success is obtained. Semi-physical experiments on the ground show that the system can improve the level of intelligence in the flight control process and effectively assist ground flight controllers in data monitoring and mission decision-making.

In recent years, great experience has been accumulated in manned flight astronautics for rendezvous in near-Earth orbit. During flights of Apollo spacecraft with crews that landed on the surface of the Moon, the problem of docking a landing module launched from the Moon’s surface with the Apollo spacecraft’s command module in a circumlunar orbit was successfully solved. A return to the Moon declared by leading space agencies requires a scheme for rendezvous of a spacecraft launched from an earth-based cosmodromee with a lunar orbital station. This paper considers some ballistic schemes making it possible to solve this problem with minimum fuel expenditures.

With the rapid development of satellite astronomy technique, In order to ensure the effective operation of the spacecraft in orbit, the spacecraft requires full ground physical simulation tests. So the construction of the ground simulation test system is an important guarantee for the development of space technology [1]. In the system of full physical simulation test for lunar orbit rendezvous and docking, the large supportive stage with super smooth surface is one of the most important large-scale precision test equipment, this stage can provide a high precision horizontal support surface in the range of 40m*30m for the test load, Each test load is usually supported by three circular air cushion, whose diameter is 200mm~400mm. High pressure gas is stored in the load equipment, and is released into the air through the air holes on the air cushion, A thin layer of air film is formed between the air cushion and the stage, the thickness of gas film is 10μm~15μm, The load equipment can freely move on the surface of a large stage with the gas film. There are strict requirements for the deformation of the system, Since the floating height is 10 μm and the test load is 3T. So the altitude difference of neighboring platforms is in range of 10μm. In order to guarantee the experiment effects, the paper analyses mechanical properties of the stage, distribute index of 10μm, and provide theoretical support for subsystem [2], [3].

Vision Based Navigation consists in utilizing cameras as precision sensors for GNC after extracting information from images. To enable the adoption of machine learning for space applications, one of obstacles is the demonstration that available training datasets are adequate to validate the algorithms. The objective of the study is to generate datasets of images and metadata suitable for training machine learning algorithms. Two use cases were selected and a robust methodology was developed to validate the datasets including the ground truth. The first use case is in-orbit rendezvous with a man-made object: a mockup of satellite ENVISAT. The second use case is a Lunar landing scenario. Datasets were produced from archival datasets (Chang'e 3), from the laboratory at DLR TRON facility and at Airbus Robotic laboratory, from SurRender software high fidelity image simulator using Model Capture and from Generative Adversarial Networks. The use case definition included the selection of algorithms as benchmark: an AI-based pose estimation algorithm and a dense optical flow algorithm were selected. Eventually it is demonstrated that datasets produced with SurRender and selected laboratory facilities are adequate to train machine learning algorithms.

We propose a Nonlinear Model Predictive Control approach to spacecraft rendezvous in non-Keplerian Lunar orbits. The approach is based on the Pontryagin Minimum Principle and allows the accomplishment of minimum-propellant maneuvers. The relative motion between the chaser and the target is described by the nonlinear and unstable dynamics of the circular restricted three body-problem. In the proposed formulation, we design a minimum-propellant controller, which leads to a bang-bang behavior of the control signal. Under suitable assumptions, simplified dynamics is employed as prediction model, in order to reduce the complexity of the controller algorithm but, at the same time, without penalizing the controller tracking performance. The proposed approach's effectiveness is validated by a simulation example.

Use of high-power solar arrays, at power levels ranging from approximately 500 KW to several megawatts, has been proposed for a solar-electric propulsion (SEP) demonstration mission, using a photovoltaic array to provide energy to a high-power xenon-fueled engine. One of the proposed applications of the high-power SEP technology is a mission to rendezvous with an asteroid and move it into lunar orbit for human exploration, the Asteroid Retrieval mission. The Solar Electric Propulsion project is dedicated to developing critical technologies to enable trips to further away destinations such as Mars or asteroids. NASA needs to reduce the cost of these ambitious exploration missions. High power and high efficiency SEP systems will require much less propellant to meet those requirements.

It's July 20 as I start this column, exactly fifty years since Apollo 11 became the first space mission to land Earthlings on the moon. After the landing, Neil Armstrong and Buzz Aldrin spent a few hours collecting dust and rocks from the lunar surface before catching a bit of sleep and departing to rendezvous with Michael Collins in the command module orbiting above. The trio successfully splashed down in the ocean three days later, completing the first manned mission to the moon, which had been promised by President John F. Kennedy eight years earlier in 1961. Humans have had a fascination with the moon for millennia. An estimated 600 million people, or one-sixth of the global population at the time, watched the moon landing live.

A life support system concept has been developed for a new NASA lunar lander concept. The ground rules and assumptions driving the design of this vehicle are different from the Constellation Altair vehicle, and have led to a different design solution. For example, this concept assumes that the lander vehicle arrives in lunar orbit independently of the crew. It loiters in lunar orbit for months before rendezvousing with the Orion Multi-Purpose Crew Vehicle (MPCV), resulting in the use of solar power for this new lander, rather than fuel cells that provided product water to the life support system in the Altair vehicle. Without the need to perform a single Lunar Orbit Insertion burn for both the lander and the MPCV, the modules do not have to be centered in the same way, so the new lander has a smaller ascent module than Altair and a large habitat rather than a small airlock. This new lander utilizes suitport technology to perform EVAs from the habitat, which leads to significantly different requirements for the pressure control system. This paper describes the major trades and resulting concept design for the life support system of a new lunar lander concept. I

Mars Sample Return was presented as the highest priority planetary science mission of the next decade [1]. Lemke et al. [2] present a Mars Sample Return mission concept in which the sample is returned directly from the surface of Mars to an Earth orbit. The sample is recovered in Earth Orbit instead of being transferred between spacecraft in Mars Orbit. This paper provides the details of this sample recovery in Earth orbit and presents as such a sub-element of the overall Mars sample return concept given in [2]. We start from the assumption that a Mars Ascent Vehicle (MAV), initially landed on Mars using a modified SpaceX Dragon capsule, has successfully delivered the sample, already contained within an Earth Return Vehicle (ERV), to a parking orbit around Mars. From the parking orbit, the ERV imparts sufficient Delta-V to inject itself into an earthbound trajectory and to be captured into an Earth orbit eventually. We take into account launch window and Delta-V considerations as well as the additional constraint of increased safety margins imposed by planetary protection regulations. We focus on how to overcome two distinct challenges of the sample return that are driven by the issues of planetary protection: (1) the design of an ERV trajectory meeting all the requirements including the need to avoid contamination of Earth's atmosphere; (2) the concept of operations for retrieving the Martian samples in Earth orbit in a safe way. We present an approach to retrieve the samples through a rendezvous between the ERV and a second SpaceX Dragon capsule. The ERV executes a trajectory that brings it from low Mars orbit (LMO) to a Moon-trailing Earth orbit at high inclination with respect to the Earth-Moon plane. After a first burn at Trans-Earth Injection (TEI), the trajectory uses a second burn at perigee during an Earth flyby maneuver to capture the ERV in Earth orbit. The ERV then uses a non-propulsive Moon flyby to come to a near-circular Moon-trailing orbit. To perform the Earth Orbit Rendezvous (EOR), a second Dragon capsule is then launched from Earth and a similar lunar flyby is performed to rendezvous with the ERV. The requirements for rendezvous, close proximity operations and capture of the sample canister are described. A concept of operations for sample retrieval is presented along with design specifications of the ERV, the required modifications to the Dragon capsule, as well as the hardware, software, sensors, actuators, and capture mechanisms used. In our concept, a container is mounted to the front hatch of Dragon, capable of accommodating the sample canister and sealing it from the rest of the capsule. The sample canister is captured using a robotic arm with a magnetic grappling mechanism. Dragon then performs a propulsive maneuver to return to Earth for a controlled re-entry while the ERV (sans sample container) is left in the Moon trailing orbit. Contingency cases and related mitigation strategies are also discussed, including the advantages and disadvantages of performing the ERV rendezvous with a crew.