Lunar Landing Mission Research Articles

A fast calculation method for three-impulse contingency return trajectory during the near-moon phase based on differential algebra

Aimed at the demand of contingency return at any time during the near-moon phase in the manned lunar landing missions, a fast calculation method for three-impulse contingency return trajectories is proposed. Firstly, a three-impulse contingency return trajectory scheme is presented by combining the Lambert transfer and maneuver at the special point. Secondly, a calculation model of three-impulse contingency return trajectories is established. Then, fast calculation methods are proposed by adopting the high-order Taylor expansion of differential algebra in the two-body trajectory dynamics model and perturbed trajectory dynamics model. Finally, the performance of the proposed methods is verified by numerical simulation. The results indicate that the fast calculation method of two-body trajectory has higher calculation efficiency compared to the semi-analytical calculation method under a certain accuracy condition. Due to its high efficiency, the characteristics of the three-impulse contingency return trajectories under different contingency scenarios are further analyzed expeditiously. These findings can be used for the design of contingency return trajectories in future manned lunar landing missions.

MEDEA: A New Model for Emulating Radio Antenna Beam Patterns for 21 cm Cosmology and Antenna Design Studies

In 21 cm experimental cosmology, accurate characterization of a radio telescope’s antenna beam response is essential to measure the 21 cm signal. Computational electromagnetic (CEM) simulations estimate the antenna beam pattern and frequency response by subjecting the EM model to different dependencies, or beam hyperparameters, such as soil dielectric constant or orientation with the environment. However, it is computationally expensive to search all possible parameter spaces to optimize the antenna design or accurately represent the beam to the level required for use as a systematic model in 21 cm cosmology. We therefore present the Model for Emulating Directivities and Electric fields of Antennas (MEDEA), an emulator that rapidly and accurately generates far-field radiation patterns over a large hyperparameter space. MEDEA takes a subset of beams simulated by CEM software, spatially decomposes them into coefficients on a complete, linear basis, and then interpolates them to form new beams at arbitrary hyperparameters. We test MEDEA on an analytical dipole and two numerical beams motivated by upcoming lunar lander missions, and then employ MEDEA as a model to fit mock radio spectrometer data to extract covariances on the input beam hyperparameters. We find that the interpolated beams have rms relative errors of at most 10−2 using 20 input beams or less, and that fits to mock data are able to recover the input beam hyperparameters when the model and mock are derived from the same set of beams. When a systematic bias is introduced into the mock data, extracted beam hyperparameters exhibit bias, as expected. We propose several extensions to MEDEA to potentially account for such bias.

Organ dose equivalents of albedo protons and neutrons under exposure to large solar particle events during lunar human landing missions

Astronauts participating in lunar landing missions will encounter exposure to albedo particles emitted from the lunar surface as well as primary high-energy particles in the spectra of galactic cosmic rays (GCRs) and solar particle events (SPEs). While existing studies have examined particle energy spectra and absorbed doses in limited radiation exposure scenarios on and near the Moon, comprehensive research encompassing various shielding amounts and large SPEs on the lunar surface remains lacking. Additionally, detailed organ dose equivalents of albedo particles in a human model on the lunar surface have yet to be investigated. This work assesses the organ dose equivalents of albedo neutrons and albedo protons during historically large SPEs in August 1972 and September 1989 utilizing realistic computational anthropomorphic human phantom for the first time. Dosimetric quantities within human organs have been evaluated based on the PHITS Monte Carlo simulation results and quality factors of the state-of-the-art NASA Space Cancer Risk (NSCR) model, as well as ICRP publications. The results with the NSCR model indicate that the albedo contribution to organ dose equivalent is less than 3 % for 1 g/cm2 aluminum shielding, while it increases to more than 30 % in some organs for 50 g/cm2 aluminum shielding during exposure to low-energy-proton-rich SPEs.

Tripping on the Moon

Two recent lunar lander missions have toppled over on touchdown. We analyse a simplified model to determine the lateral speed of an object that results in tipping, as a practical example of an inelastic collision that conserves angular momentum. The solution employs rotational dynamics and conservation laws covered in an introductory mechanics course. We suggest further investigations, including simple experiments that can be performed in the teaching laboratory.

Multi-Objective Robust Trajectory Design for Powered Descent and Landing

A robust trajectory optimization approach for guidance algorithm gain and target vector selection for powered descent and landing is developed. A genetic algorithm is used to determine optimized guidance algorithm parameters to minimize the impact of initial condition, environment, navigation, and vehicle property uncertainty on flight performance for a given sensor suite. Vehicle state uncertainties are computed efficiently using linear covariance analysis techniques. When implemented in the guidance algorithm, the optimized gains and target vectors shape a trajectory that has more favorable conditions for a given navigation sensor suite, resulting in improved flight performance. As a demonstration of this method, the optimized guidance parameters are found for a multiphase trajectory from powered descent initiation to touchdown for a robotic lunar landing mission. Single-objective optimization results demonstrate a reduction in uncertainty and an improvement in nominal performance. Multi-objective optimization results showing the tradeoff between terminal position uncertainty and total propellant usage are presented for multiple sensor suite compositions. Further, guidance parameters selected using the developed robust trajectory design approach may enable acceptable flight performance with fewer and/or lower-quality sensors. The resulting Pareto fronts present the optimal trade space early in the mission design process to enable informed decision-making.

High frequency Lunar Penetrating Radar quality control, editing and processing of Chang’E-4 lunar mission

Chinese lunar landing mission Chang’E-4 reached the far side of the Moon in January 2019 and has been providing unprecedented Lunar Penetrating Radar data able to explore the lunar subsurface down to more than 40 m (with its more resolutive high frequency band). Data are periodically released to the scientific community in raw PDS4 format. Here we provide different versions of the radar dataset after editing (i.e. pre-processing), partial, and full processing in order to provide a complete ready-to-use dataset to end-users (data collected since 4th January 2019 until 27th March 2023) which can be directly exploited for analysis, interpretation, inversion, as well as integration with imagery or other information. In particular, we implemented an efficient and objective way to remove duplicated traces representing more than 90% of original data, as well as a processing flow able to retain all the original data information, while avoiding redundancies. The provided datasets can be implemented with future data releases and straightforwardly exploited for any future analysis.

Abort Guidance During Lunar Powered Descent

During powered descent of a crewed lunar landing mission, the ability to abort and ascend into a clear pericynthion orbit is critical in case any contingency renders the landing infeasible or excessively risky. In such a case, it is incumbent upon the guidance system to determine autonomously a safe abort trajectory and the associated guidance command. To accomplish this task reliably, a novel two-phase abort guidance strategy is proposed in this paper where a pull-up maneuver redirects the velocity vector to help mitigate any ground collision risk and achieve an appropriate initial ascent condition, followed by a fuel-optimal ascent guidance algorithm to insert the spacecraft into a predefined safe orbit around the moon. Development of the pull-up guidance laws and a description of the fuel-optimal ascent guidance based on the indirect method of optimal control are presented. In-depth investigation is conducted to provide a full understanding of the validity of the abort solutions. The solutions and fuel efficiency of the proposed guidance strategy are independently verified by using a direct method of optimal control. Monte Carlo closed-loop simulations demonstrate the effectiveness and robustness of this method throughout the entire powered descent.

Design and implementation of regional integrated environmental measurement for China space station

The monitoring of cabin air pressure, temperature, humidity, microgravity and other environmental parameters is very important for the long-term and reliable operation of large manned spacecraft. Based on the characteristics of 1553B bus, a design method of regional integrated environmental measurement for the space station was proposed, which realized the high precision and reliable measurement of environmental parameters such as pressure, temperature, humidity and microgravity of china space station. The measurement data can accurately reflect the environmental changes of astronauts’ in-orbit living, working and extravehicular activities, which verifies the effectiveness of the design. It provides the necessary new principle application verification for higher precision pressure measurement in lunar landing missions.

DESIGN AND PERFORMANCE OF A 110 N ATTITUDE CONTROL THRUSTER FOR LUNAR MISSIONS

The design and development of an attitude control thruster, designated A110, developed for broad attitude control applications, including lunar landers, is described. The A110 thruster produces 110 N (25 lbf) of nominal thrust and in-excess of 305.5 s of specific impulse (I<sub>sp</sub>) while operating on a fuel blend of 20% monomethylhydrazine-80% hydrazine (M20) with mixed oxides of nitrogen containing 3% nitric oxide (MON3) as the oxidizer. The injector and combustion chamber are additively manufactured with cobalt-chrome and niobium C103, respectively, to enable the rapid development required of the industry. The thruster is capable of response times of under 10 ms and currently produces a minimum impulse bit (MIB) of 0.436 N.s and has demonstrated operation in pulse-mode with a variety of duty cycles during development testing. These capabilities allow for the generation of precise pulses for accurate control of a lander. The testing of this thruster is used to evaluate steady-state and pulse-mode specific impulse, impulse repeatability, and combustion stability. The measured performance (i.e., specific impulse and characteristic velocity (C*), are compared with theoretical values. The capabilities demonstrated by the A110 thruster are applicable to a wide range of lunar lander missions and can be qualified for minor changes in thruster requirements for additional missions without the need for major design changes. This article is published with the permission of the authors granted to 3AF - Association Aéronautique et Astronautique de France (www.3AF.fr) organizer of the Space Propulsion International Conference.

Soft Landing Parameter Measurements for Candidate Navigation Trajectories Using Deep Learning and AI-Enabled Planetary Descent

Smart instruments, sensors, and AI technologies are playing an important role in many fields such as medical science, Earth science, astronomy physics, and space study. This article attempts to study the role of sensors, instruments, and AI (artificial intelligence) based smart technologies in lunar missions during navigation of trajectories. Lunar landing missions usually divide the power descent phase into three to four sub-phases. Each sub-phase has its own set of initial and final constraints for the desired system state. The landing systems depend on human competencies for making the most crucial landing decisions. Trajectory planning and designing are very significant in lunar missions, and it requires inputs with precision. The manual systems may be prone to errors. In contrast, AI and smart sensor-based measurements give an accurate idea about the trajectory paths and make appropriate decisions where manual systems may turn into disasters. The manual systems are either pre-fed or have manual controls to guide the trajectory. For autonomous landing problems, trajectory design is a very crucial task. The automated trajectories play a vital role in the measurement and prediction of landing state parameters of the space rocket. Nowadays, sensors, intelligent instruments, and the latest technologies go hand in hand to devise measurement methods for accurate calculations and make appropriate decisions during landing space rockets at the designated destination. Space missions are very expensive and require huge efforts to design smart systems for navigation trajectories. This paper attempts to design all possible candidates of reference navigation trajectories for autonomous lunar descent by employing 3D non-linear system dynamics with randomly chosen initial state conditions. The generated candidates do not rely on multiple hops and thus exhibit an ability to serve autonomous missions. This research work makes use of smart sensors and AI federated techniques for smartly training the system to serve the ultimate purpose. The trajectories are simulated in an automated simulating environment to perform exhaustive analyses. The results accurately approximate the trajectories analogous to their numerical counterparts and converge to their measured final state estimates. The generation rate of feasible trajectories measures the accuracy of the algorithm. The algorithm’s accuracy is near 0.87 for 100 sec flight time, which is reasonable.

Positioning and Velocity Performance Levels for a Lunar Lander using a Dedicated Lunar Communication and Navigation System

<h3>Abstract</h3> The interest in Moon exploration has grown substantially in the last few years, appointing the Moon as the first step toward deep space exploration. However, the current state-of-the-art approach for lunar landing does not always reach the required performance levels. This contribution presents a potential implementation of a dedicated lunar communication and navigation service (LCNS) and the performance levels achievable by a representative lunar lander mission that uses the LCNS. The expected positioning precision during the final descent and at the landing site is demonstrated here with a variance-covariance analysis starting from reasonable assumptions about the capabilities of a potential dedicated LCNS system. The performance in positioning and navigation achievable during a generic moon-landing phase significantly outperforms existing ground-based baseline solutions, enabling the stringent requirements from the International Space Exploration Coordination Group (ISECG) to be met.

Landing Trajectory Generation and Energy Optimization for Unmanned Lunar Mission

The moon is recognized as an important destination for space science and exploration. To find a satisfactory answer for the mystery of the universe and to make use of the lunar resources for the welfare of human beings, several space agencies are planning manned and unmanned missions on the moon. As a result, the concept of lunar vehicles has begun with an advanced descent scheme to execute a precise and safe landing on the surface of the moon. On the contrary, the energy budget is an important issue for any space mission. To reduce the cost of a space mission, it is necessary to design the vehicle trajectory based on optimized energy resources. Fuel is the main energy in a space mission. Therefore, a fuel-optimized energy generation technique is focused on this research. The design of an algorithm that generates a real-time trajectory for the descent and landing of a lunar probe is critical to ensuring a successful lunar landing mission. A scheme of dual-step trajectory generation for lunar descent is also investigated in this paper. In the algorithm developing process, the thrust-to-mass ratio is considered as a principle variable. Algorithm design along with mathematical modeling and simulation results are described in detail. In addition, the proposed method for generating reference trajectory profiles is also analyzed for fuel consumption and robustness.

Design and implementation of GNC system of lander and ascender module of Chang&amp;rsquo;e-5 spacecraft

<p indent="0mm">The lander and ascender module of Chang’e-5 spacecraft achieved China’s first lunar landing and ascent mission. The guidance, navigation, and control (GNC) system is one of the most important subsystems. To meet the weight, safety, and intelligence requirement of the mission, light-weight integration designs are implemented in the GNC system, so that a single control system could meet the mission requirement of landing, ascent, and rendezvous. Moreover, a high-rate robust optical obstacle avoidance system is applied with a hierarchical controller, and landing security is enhanced while fuel consumption is minimized. Furthermore, autonomous attitude determination on the lunar surface and orbit insertion are achieved during the lunar ascent phase, and momentum wheels are relied on to obtain good attitude stability for the spacecraft during the rendezvous phase. Finally, the GNC system has fully functional redundancy and autonomous management, which significantly increases the fault detection and isolation capabilities. Here, we present key design specifications and implementations in the GNC system of the lander and ascender module.

DEM-FEM-MBD coupling analysis of landing process of lunar lander considering landing mode and buffering mechanism

The safe lander dynamics is an important part of the lunar landing mission. In this paper, a discrete element method (DEM) for lunar soil is set up, and the finite element method (FEM) for the lander is set up by shell elements and beam elements. The lander is regarded as a multibody system composed of a cabin, legs and footpads, and its motion characteristics are solved by multibody dynamics (MBD). A DEM-FEM-MBD coupling algorithm is developed to simulate the landing process of the lander considering landing mode and buffer mechanism, whose correctness is verified by comparison with a full-scale experiment involving lunar lander on earth. The effects of the mass, landing velocity and attitude of the lander on the safe landing are discussed. The buffering mechanism and influencing factors of lunar soil are analyzed. The results show that the impact force peak and impact depth gradually increase with the increase in the landing velocity and mass of the lander. Two kinds of inclined landing modes are defined and compared with vertical landing. It is found that the force on the landing leg that first contacts the lunar soil is significantly greater than that on the landing leg, that contacts later. The impact force peak on the lander under the two inclined landing modes is similar, but the impact depth of the 1-2-1 mode is significantly greater than that of the 2-2 mode. In the process of landing, lunar soil has the function of buffering dissipation. The energy dissipation rate is affected by the physical characteristics of lunar soil and the mechanical energy of the lander.

The science mission of SpaceIL’s Beresheet lander

SpaceIL’s lunar lander mission Beresheet was launched on February 22, 2019 and impacted its targeted landing site in Mare Serenitatis on the Moon on April 11, 2019. The spacecraft carried a package of scientific instruments including a fluxgate magnetometer and a retroreflector array for laser ranging, as well as a suite of cameras. Orbital measurements of the magnetic field from Kaguya and Lunar Prospector guided the selection of the landing site to a location west of Posidonius crater in the Serenitatis plains, where the magnitude of the modeled magnetic field reaches 8–10 ​nT ​at the surface. Data was collected by the SILMAG magnetometer from Earth orbit, lunar orbit, and during the descent maneuver, although its interpretation is hindered by the presence of a spacecraft field.

VISUAL-BASED INTEGRATED NAVIGATION SYSTEM APPLIED TO A SIMULATION OF LUNAR MODULE LANDING

Abstract. With the development of space technology, more and more lunar researches are performed by different countries. For the lunar landing mission success, the lunar landing module should equip with advanced Positioning and Orientation System (POS) for the navigation requirements. For the pinpoint landing mission formulated by NASA, a good POS with error less than 100 meters is needed in order to make the lunar module land safely at the exact destination on lunar surface. However, the existing technologies for lunar navigation, such as satellite positioning and star tracker, have poor performance for the navigation requirements. The visual-based positioning technology is an alternative way to make sure a lunar landing module reaches the destination. There are two types of visual-based positioning technology, absolute and relative navigation. The relative navigation system can provide the solution at a higher rate, but the error would accumulate over time. On the contrary, the absolute navigation could provide an initial position or updates of position and attitude for relative navigation. Thus, the integrated navigation system from those two methods can take advantage of both stand-alone systems. On the other hand, the Inertial Navigation System (INS) can help it overcome the disadvantage that the images much closer to the lunar surface are not available. This study shows an integrated navigation system that integrates a visual-based navigation system and an INS, which is implemented in a simulated lunar surface.

First look by the Yutu-2 rover at the deep subsurface structure at the lunar farside

The unequal distribution of volcanic products between the Earth-facing lunar side and the farside is the result of a complex thermal history. To help unravel the dichotomy, for the first time a lunar landing mission (Chang’e-4, CE-4) has targeted the Moon’s farside landing on the floor of Von Kármán crater (VK) inside the South Pole-Aitken (SPA). We present the first deep subsurface stratigraphic structure based on data collected by the ground-penetrating radar (GPR) onboard the Yutu-2 rover during the initial nine months exploration phase. The radargram reveals several strata interfaces beneath the surveying path: buried ejecta is overlaid by at least four layers of distinct lava flows that probably occurred during the Imbrium Epoch, with thicknesses ranging from 12 m up to about 100 m, providing direct evidence of multiple lava-infilling events that occurred within the VK crater. The average loss tangent of mare basalts is estimated at 0.0040-0.0061.

Application of the particle trajectory imaging for modelling dusty plasma levitation on the Moon

The future robotic and human lunar landing missions to the Moon has many factors which turn them into difficult technology and science tasks. One of these factors is the influence of the dust. Upper layers of the regolith are an insulator. The regolith exposed to permanent bombardment by micrometeorites and acquire a charge due to solar wind fluxes and solar UV radiation. These factors create a charge distribution on the surface of the Moon: positive on the illuminated side and negative on the night side. On the day side of the Moon near the surface layer exists possibility of formation an electric field. Charged dust particles of micron and submicron sizes can take off and levitate over the surface. The aim of the work is to get visualization of the dynamic of dust particles under a charged surface with simultaneously estimating its parameters as accurately as possible. The experimental setup based on vacuum chamber for physical modelling of dusty plasma levitation is described. For visualization of the dust particles trajectory a stereo system of two cameras with a laser as source of illumination is used. Image processing techniques for estimating the particle trajectory in three dimensional coordinates and examples of processing results are presented.

Fixed-thrust Earth–Moon free return orbit design based on a hybrid multi-conic method of pseudo-perilune parameters

This study proposes a high-precision fixed-thrust free return orbit design method for a manned lunar mission centred on the rendezvous and docking with the lunar space station. Future manned lunar landing missions based on the lunar space station will require different spacecraft to achieve rendezvous and docking with it in different mission stages. Adopting the perilune parameters to design the Earth–Moon flight trajectory would be more convenient for decoupling in this kind of series missions. Thus, a new set of pseudo-parameters representing the perilune motion states is defined, and based on these, a hybrid multi-conic method is then proposed to design the free return orbit for manned spacecraft. The new method is more accurate than the traditional four-phase patched-conic method, and can produce initial values with a higher converging speed for a high-precision trajectory iteration design. Then, a fast and high-precision continuation method for the fixed-thrust trajectory is proposed based on the impulse-based solutions, and the gravity loss of the impulse-thrust transfer is analysed. Finally, the general characteristics and launch window of the free return orbit, as well as the reachable domain of the pseudo-perilune parameters, are analysed using the hybrid multi-conic method. The results will provide a theoretical basis for future missions planning.

High‐Resolution Terrain Analysis for Lander Safety Landing and Rover Path Planning Based on Lunar Reconnaissance Orbiter Narrow Angle Camera Images: A Case Study of China's Chang'e‐4 Probe

AbstractChina's Chang'e‐4 (CE‐4) probe will explore the South Pole‐Aitken basin in late 2018. Its preselected landing area is located on the southeastern floor of the Von Kármán crater. Landing experience of China's Chang'e‐3 probe may not be applied to CE‐4 mission directly because the topography of lunar farside is more rugged. Moreover, due to scale dependence and smooth effects, the previous topographic studies derived from digital elevation model data with lower resolution cannot represent the meter‐scale topographic features. Because of low time and space complexities in lunar images, image segmentation algorithm is especially suitable for recognizing lunar features. We divided the narrow angle camera (NAC) mosaic image into three parts: dark area, bright area, and flat area, based on a double‐threshold Otsu method. The first two parts corresponded to the undulating areas (positive and negative terrains). And then we calculated the flat area percentage (Fap) of the previous successful lunar landing missions (including Chang'e‐3, Apollo, Surveyor, and Luna series) and obtained the Fap threshold (&gt;0.6) for lunar safe landing. According to Fap map (~1.5 m per pixel) of CE‐4 preselected landing area, the divided square grids with a size of 0.01° can be classified into safe grids (Fap &gt; 0.6) and unsafe grids (Fap ≤ 0.6). The CE‐4 preselected landing areas can be greatly reduced to five potential landing areas. In the last, we proposed a route planning method, which took both the distance and security into account, for CE‐4 rover to dynamically generate a safe and short route between current grid and target grid.