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Autonomous control of lunar landers is essential for successful space missions, where precision and efficiency are crucial. This study presents a novel control strategy that leverages proportional, integral, and derivative (PID) controllers to manage the altitude, attitude, and position of a lunar lander, considering time-varying mass and sloshing behavior. Additionally, neural network models are developed, to approximate the lander's mass properties as they change during descent. The challenge lies in the significant mass variations due to fuel, oxidizer, and pressurant consumption, which affect the lander's inertia and sloshing behavior and complicate control efforts. We have developed a control-oriented model incorporating these mass dynamics, employing multiple PID controllers to linearize the system and enhance control precision. Altitude is maintained by one PID controller, while two others adjust the thrust vector control (TVC) gimbal angles to manage pitch and roll, with a fourth controller governing yaw via a reaction control system (RCS). A cascade PD controller further manages position by feeding commands to the attitude controllers, ensuring the lander reaches its target location. The lander's TVC mechanism, equipped with a spherical gimbal, provides thrust in the desired direction, with control angles α and β regulated by the PID controllers. To improve the model's accuracy, we have introduced time delays caused by fluid dynamics and actuator response, modeled via computational fluid dynamics (CFD). Fluid sloshing effects are also simulated as external forces acting on the lander. The neural networks are trained using data derived from computer-aided design (CAD) simulations of the lander vehicle, specifically the inertia tensor and the center of mass (COM) based on the varying mass levels in the tanks. The trained neural networks (NNs) can then use lander tank levels and orientation to inform and accurately predict the lander's COM and inertia tensor in real time during the mission. The implications of this research are significant for future lunar missions, offering enhanced safety and efficiency in vehicle descent and landing operations. Our approach allows for real-time estimation of the lander's state and for precise execution of maneuvers, verified through complex numerical simulations of the descent, hover, and landing phases. [ABSTRACT FROM AUTHOR]

Autonomous control of lunar landers is essential for successful space missions, where precision and efficiency are crucial. This study presents a novel control strategy that leverages proportional, integral, and derivative (PID) controllers to manage the altitude, attitude, and position of a lunar lander, considering time-varying mass and sloshing behavior. Additionally, neural network models are developed, to approximate the lander’s mass properties as they change during descent. The challenge lies in the significant mass variations due to fuel, oxidizer, and pressurant consumption, which affect the lander’s inertia and sloshing behavior and complicate control efforts. We have developed a control-oriented model incorporating these mass dynamics, employing multiple PID controllers to linearize the system and enhance control precision. Altitude is maintained by one PID controller, while two others adjust the thrust vector control (TVC) gimbal angles to manage pitch and roll, with a fourth controller governing yaw via a reaction control system (RCS). A cascade PD controller further manages position by feeding commands to the attitude controllers, ensuring the lander reaches its target location. The lander’s TVC mechanism, equipped with a spherical gimbal, provides thrust in the desired direction, with control angles α and β regulated by the PID controllers. To improve the model’s accuracy, we have introduced time delays caused by fluid dynamics and actuator response, modeled via computational fluid dynamics (CFD). Fluid sloshing effects are also simulated as external forces acting on the lander. The neural networks are trained using data derived from computer-aided design (CAD) simulations of the lander vehicle, specifically the inertia tensor and the center of mass (COM) based on the varying mass levels in the tanks. The trained neural networks (NNs) can then use lander tank levels and orientation to inform and accurately predict the lander’s COM and inertia tensor in real time during the mission. The implications of this research are significant for future lunar missions, offering enhanced safety and efficiency in vehicle descent and landing operations. Our approach allows for real-time estimation of the lander’s state and for precise execution of maneuvers, verified through complex numerical simulations of the descent, hover, and landing phases.

The autonomous control of landing procedures can provide the efficiency and precision that are vital for the successful, safe completion of space operations missions. Controlling a lander with this precision is challenging because the propellants, which will be expended during the operations, represent a significant fraction of the lander's mass. The mass variation of each tank profoundly influences the inertia and mass characteristics as thrust is generated and complicates the precise control of the lander state. This factor is a crucial consideration in our research and methodology. The dynamics model for our lander was developed where the mass, inertia, and center of mass (COM) vary with time. A feed-forward neural network (NN) is incorporated into the dynamics to capture the time-varying inertia tensor and COM. Moreover, the propellant takes time to travel through the feed lines from the storage tanks to the engine; also, the solenoid valves require time to open and close. Therefore, there are time delays between the actuator and the engine response. To take into account these sources of variations, a combined time delay is also included in the control loop to evaluate the effect of delays by fluid and mechanisms on the performance of the controller. The time delay is estimated numerically by a Computational Fluid Dynamics (CFD) model. As part of the lander's control mechanism, a thrust vector control (TVC) with two rotational gimbals and a reaction control system (RCS) are incorporated into the dynamics. Simple proportional, integral, and derivative (PID) controllers are designed to control the thrust, the gimbal angles of the TVC, and the torque required by the RCS to manipulate the lander's rotation and altitude. A complex mission with several numerical examples is presented to verify the hover and rotational motion control. [ABSTRACT FROM AUTHOR]

This research article presents a comparison between two mainstream Deep Reinforcement Learning (DRL) algorithms, Asynchronous Advantage Actor-Critic (A3C) and Proximal Policy Optimization (PPO), in the context of two diverse environments: CartPole and Lunar Lander. DRL algorithms are widely known for their effectiveness in training agents to navigate complex environments and achieve optimal policies. Nevertheless, a methodical assessment of their effectiveness in various settings is crucial for comprehending their advantages and disadvantages. In this study, we conduct experiments on the CartPole and Lunar Lander environments using both A3C and PPO algorithms. We compare their performance in terms of convergence speed and stability. Our results indicate that A3C typically achieves quicker training times, but exhibits greater instability in reward values. Conversely, PPO demonstrates a more stable training process at the expense of longer execution times. An evaluation of the environment is needed in terms of algorithm selection, based on specific application needs, balancing between training time and stability. A3C is ideal for applications requiring rapid training, while PPO is better suited for those prioritizing training stability.

Theoretical and experimental studies were performed with the goal of determining the best kind of information presented on a predictive display to provide the highest accuracy in space mission execution, taking into account the requirement for reduced propellant consumption. The following space missions were studied: docking performed in manual and teleoperator control modes and lunar landing. By using a theoretical technique, algorithms for the flight path marker in manual and teleoperator control modes were developed. For the landing mission, a technique for determining the program trajectory and its calculation in real time was proposed. The experiments performed using a ground-based simulator demonstrated the high effectiveness of a predictive display and allowed to select the best variant of information to be presented on the display. • The predictive display is a means of improving of flight safety in different space missions. • The proposed algorithm for time delay suppression improves precise tracking in a docking task performed in teleoperator mode. • The proposed law for optimizing the Lunar Lander trajectory is performed in real time and provides the economy of propellant. [ABSTRACT FROM AUTHOR]

The dynamic prediction of large spacecraft is a time-consuming process. The impulse-based substructuring (IBS) method is an efficient and accurate method for the dynamic analysis of large-scale structures. The ordinary IBS method couples the substructure impulse response functions (IRFs) with coincident interface nodes. However, the interface nodes in a practical structure are usually mismatched. This paper extends the ordinary IBS method and presents a method for coupling multi-domain substructures, including IBS, FEM and modal substructures, with non-matching interfaces. Both displacement compatibility and force equilibrium are satisfied on the non-matching interfaces of substructures by introducing the normalized interpolation technique into the framework of IBS. To evaluate the performance of the proposed method, numerical models and a practical application with a lunar lander and a rover are presented. The main bodies of the lander and the rover are coupled by non-matching interfaces, and the coupled system is excited by a buffer load. With the employment of this proposed method, the computing time of the landing simulation of the lunar lander is dramatically reduced, and the results show strong agreement with references. It is demonstrated that the proposed method is accurate and efficient in the transient dynamic analysis of engineering structures. [ABSTRACT FROM AUTHOR]

The autonomous control of landing procedures can provide the efficiency and precision that are vital for the successful, safe completion of space operations missions. Controlling a lander with this precision is challenging because the propellants, which will be expended during the operations, represent a significant fraction of the lander’s mass. The mass variation of each tank profoundly influences the inertia and mass characteristics as thrust is generated and complicates the precise control of the lander state. This factor is a crucial consideration in our research and methodology. The dynamics model for our lander was developed where the mass, inertia, and center of mass (COM) vary with time. A feed-forward neural network (NN) is incorporated into the dynamics to capture the time-varying inertia tensor and COM. Moreover, the propellant takes time to travel through the feed lines from the storage tanks to the engine; also, the solenoid valves require time to open and close. Therefore, there are time delays between the actuator and the engine response. To take into account these sources of variations, a combined time delay is also included in the control loop to evaluate the effect of delays by fluid and mechanisms on the performance of the controller. The time delay is estimated numerically by a Computational Fluid Dynamics (CFD) model. As part of the lander’s control mechanism, a thrust vector control (TVC) with two rotational gimbals and a reaction control system (RCS) are incorporated into the dynamics. Simple proportional, integral, and derivative (PID) controllers are designed to control the thrust, the gimbal angles of the TVC, and the torque required by the RCS to manipulate the lander’s rotation and altitude. A complex mission with several numerical examples is presented to verify the hover and rotational motion control.

This paper introduces a flexible framework to enable high-fidelity simulations of the fuel slosh dynamics for verification of the Guidance, Navigation and Control (GNC) algorithms. The sloshing phenomenon must be tackled during the control system design phase and the GNC software run and validated within appropriate simulation frameworks. Equivalent mechanical models can well approximate sloshing under specific assumptions, allowing the derivation of the multibody vehicle's equations of motion; however, while suitable for analysis and control synthesis purposes, this approach presents several limitations for implementation as a simulation model. In this work, the multibody plant embedding slosh dynamics is modeled by means of the DLR's Vertical Landing Vehicles Library (VLVLib) written using the object-oriented Modelica modeling language. The advantages of using Modelica are explored throughout the paper. Given that control synthesis, analyses and simulation campaigns are usually performed in the Matlab/Simulink environment, a core point of the proposed solution is to achieve a good synergy within the latter and the Modelica plant model with the least readaptation burden for the developer. The strategies to achieve a full integration and verification tool chain are illustrated. The potential of this approach is demonstrated via two different test cases: a lunar landing scenario of the ALINA spacecraft developed by PTS, and CALLISTO reusable rocket demonstrator. • A flexible framework to enable high-fidelity simulations of the fuel slosh dynamics for verification of the GNC algorithms is introduced. • The spacecraft model embedding the slosh dynamics is implemented exploiting Modelica language. • A GNC testing framework is obtained with the integration of the causal Simulink environment with the compiled Modelica model. • Two test cases are presented: the ALINA spacecraft lunar landing scenario and CALLISTO reusable Vertical-Takeoff–Vertical-Landing demonstrative mission. [ABSTRACT FROM AUTHOR]

This paper proposed a multi-objective, multi-disciplinary design optimization and multi-attribute evaluation method for the manned lunar lander descent stage. A system design model is established considering multiple disciplines such as propulsion, structure, trajectory and cost. With the goal of minimizing mass, minimizing cost and maximizing velocity increment, the overall scheme for hybrid rocket motors was optimized by altering various grain shapes and feed systems, acting as an alternative propulsion scheme for the manned lunar lander. Under the same design conditions, the optimal scheme of hybrid rocket motors considering continuous and discrete attributes was studied and compared with that of liquid rocket engines to elucidate the characteristics of the hybrid rocket motors for deep space exploration. The results showed that the evaluation results obtained by considering only continuous attributes were different from those by considering both continuous and discrete attributes. The liquid propulsion scheme with liquid oxygen/kerosene is superior to the hybrid propulsion schemes due to its excellent cost and specific impulse performance when only continuous attributes are considered. However, hybrid rocket motors have shown good performance in terms of operability, manufacturability, safety and environmental protection. So, after introducing discrete attributes, the hybrid propulsion schemes show greater potential. In brief, based on the multi-attribute evaluation method considering comprehensive attributes, the hybrid rocket motor provided with tube grain and gas pressure feed system was considered as the optimal propulsion system for the lunar lander. Furthermore, the parametric analysis showed that the fuel grain outside diameter, the initial design thrust and the initial oxygen-to-fuel ratio had a significant influence on the performance of the hybrid rocket motor for the lunar lander, and in particular, the effect of the fuel grain outside diameter was more than 50%. [ABSTRACT FROM AUTHOR]

A new method is developed to determine a number of operating parameters of the dashboard thermal control system (DTCS) at the initial design stage. An application example of the method is presented using a thermal mathematical model of the DTCS as part of the lunar lander. Particular attention is paid to the aspects of reproducing external thermal effects and the adequate representation of the DTCS functional diagram based on heat pipes. DTCS modification options are discussed. The DTCS cooling capacity and temperature states as part of the lunar lander are calculated. [ABSTRACT FROM AUTHOR]

A new method is developed to determine a number of operating parameters of the dashboard thermal control system (DTCS) at the initial design stage. An application example of the method is presented using a thermal mathematical model of the DTCS as part of the lunar lander. Particular attention is paid to the aspects of reproducing external thermal effects and the adequate representation of the DTCS functional diagram based on heat pipes. DTCS modification options are discussed. The DTCS cooling capacity and temperature states as part of the lunar lander are calculated. [ABSTRACT FROM AUTHOR]

In this study, ground tests of a lab-scale hybrid rocket motor were conducted to verify the feasibility of the hybrid propulsion system for lunar lander application. The primary goal is to assess the realizability of hybrid rocket by testing its throttleability and soft landing capability with a scale-down lunar module. A design thrust of 200 N was achieved by clustering four identical 50 N-class gaseous oxygen (GOX)/high-density polyethylene (HDPE) hybrid rocket motors with multi-port solid fuels. Ground tests were carried out via two main experiments: static test and drop test. Static test was focused on the overall performance of the clustering module such as cold injection, uniform oxidizer distribution, throttleability and simultaneous ignition of the four motors, while the drop test was performed to investigate the planned throttle behavior using a 1-D vertical drop test stand. The clustering module was controlled in an open-loop setup with a simple ballistic flight simulation input. The landing velocity of 1.01 m/s was achievable, confirming the possibility of soft landing missions on lunar surface using hybrid rocket motors.

This work proposes a systematic methodology for designing an active leveling system (ALS) actuator for lunar landing application. The ALS actuator is integrated into an inverted tripod leg layout, exploiting a honeycomb crushable damper as a shock absorber. The proposed ALS actuator is fitted within the leg's primary strut and features a custom permanent-magnet synchronous machine rigidly coupled with a lead screw. The actuator aims to both provide proper leg deployment functioning and compensate for the different shock absorber deformations during landing. The leg dynamic behavior is simulated through a parameterized multi-body model to investigate different landing scenarios. First, a parametric sensitivity approach is used to optimize the transmission system and the electric machine characteristics. Then, the electric motor model is numerically validated and optimized through electromagnetic finite element analysis. To validate the proposed ALS design methodology, a virtual test bench is used to assess the ALS performances under different load scenarios. It is found that the proposed methodology is able to yield a compact, well-sized actuator which is numerically validated with the EL3 platform as a case study. [ABSTRACT FROM AUTHOR]

The dynamic prediction of large spacecraft is a time-consuming process. The impulse-based substructuring (IBS) method is an efficient and accurate method for the dynamic analysis of large-scale structures. The ordinary IBS method couples the substructure impulse response functions (IRFs) with coincident interface nodes. However, the interface nodes in a practical structure are usually mismatched. This paper extends the ordinary IBS method and presents a method for coupling multi-domain substructures, including IBS, FEM and modal substructures, with non-matching interfaces. Both displacement compatibility and force equilibrium are satisfied on the non-matching interfaces of substructures by introducing the normalized interpolation technique into the framework of IBS. To evaluate the performance of the proposed method, numerical models and a practical application with a lunar lander and a rover are presented. The main bodies of the lander and the rover are coupled by non-matching interfaces, and the coupled system is excited by a buffer load. With the employment of this proposed method, the computing time of the landing simulation of the lunar lander is dramatically reduced, and the results show strong agreement with references. It is demonstrated that the proposed method is accurate and efficient in the transient dynamic analysis of engineering structures.

This paper proposed a multi-objective, multi-disciplinary design optimization and multi-attribute evaluation method for the manned lunar lander descent stage. A system design model is established considering multiple disciplines such as propulsion, structure, trajectory and cost. With the goal of minimizing mass, minimizing cost and maximizing velocity increment, the overall scheme for hybrid rocket motors was optimized by altering various grain shapes and feed systems, acting as an alternative propulsion scheme for the manned lunar lander. Under the same design conditions, the optimal scheme of hybrid rocket motors considering continuous and discrete attributes was studied and compared with that of liquid rocket engines to elucidate the characteristics of the hybrid rocket motors for deep space exploration. The results showed that the evaluation results obtained by considering only continuous attributes were different from those by considering both continuous and discrete attributes. The liquid propulsion scheme with liquid oxygen/kerosene is superior to the hybrid propulsion schemes due to its excellent cost and specific impulse performance when only continuous attributes are considered. However, hybrid rocket motors have shown good performance in terms of operability, manufacturability, safety and environmental protection. So, after introducing discrete attributes, the hybrid propulsion schemes show greater potential. In brief, based on the multi-attribute evaluation method considering comprehensive attributes, the hybrid rocket motor provided with tube grain and gas pressure feed system was considered as the optimal propulsion system for the lunar lander. Furthermore, the parametric analysis showed that the fuel grain outside diameter, the initial design thrust and the initial oxygen-to-fuel ratio had a significant influence on the performance of the hybrid rocket motor for the lunar lander, and in particular, the effect of the fuel grain outside diameter was more than 50%.

With the further exploration of the Moon, lunar landers need to land in complex and unexplored areas. However, current lunar landers are confined to basic landing conditions (low landing velocity and plain terrains). This research proposes a novel lunar lander with magnetorheological fluid (MRF) dampers that can adapt its damping forces to various complex landing conditions. The design process and coupled electromagnetic hydrodynamic model analysis of the lunar lander were systematically elucidated. According to the derived theory model of the damping force, a fuzzy control law is proposed. The controllable damping forces under different control currents are simulated according to the proposed fuzzy control law. Additionally, landing dynamic simulations of landers with different dampers are carried in MSC. Adams to compare with their landing performances. MRF dampers under fuzzy control can decrease the largest acceleration of the lander cabin by 63.6%. Finally, prototype drop simulations and experiments are performed, which results match each other well. By increasing the control current from 1 A to 2 A, the damping forces of the MRF damper increase approximately 20% in both the simulations and experiments. Simulations and experiments prove that MRF dampers can be used for energy impact absorption as variable dampers. [ABSTRACT FROM AUTHOR]

Multiple nations are planning activity on the Moon's surface, and to deconflict lunar operations we must understand the sandblasting damage from rocket exhaust blowing soil. Prior research disagreed over the scaling of the erosion rate, which determines the magnitude of the damage. Reduced gravity experiments and two other lines of evidence now indicate that the erosion rate scales with the kinetic energy flux at the bottom of the laminar sublayer of the gas. Because the rocket exhaust is so fast, eroded particles lifted higher in the boundary layer do not impact the surface for kilometers (if at all; some leave the Moon entirely), so there is no saltation in the vicinity of the gas. As a result, there is little transport of gas kinetic energy from higher in the boundary layer down to the surface, so the emission of soil into the gas is a surprisingly low energy process. In low lunar gravity, a dominant source of resistance to this small energy flux turns out to be the cohesive energy density of the lunar soil, which arises primarily from particles in the 0.3 to 3 μ m size range. These particles constitute only a tiny fraction of the mass of lunar soil and have been largely ignored in most studies, so they are poorly characterized. • Reduced gravity erosion experiments produced insight into the physics. • Roberts' theory of soil erosion under landing rockets is incorrect. • Erosion under a rocket plume is a surprisingly low energy phenomenon. • Erosion rate scales with energy flux in the laminar sublayer, not shear stress. [ABSTRACT FROM AUTHOR]

In the companion paper ("Erosion rate of lunar soil under a landing rocket, part 1: identifying the rate-limiting physics", this issue) an equation was developed for the rate that lunar soil erodes under the exhaust of a landing rocket. That equation has only one parameter that is not calibrated from first principles, so here it is calibrated by the blowing soil's optical density curve during an Apollo landing. An excellent fit is obtained, helping validate the equation. However, when extrapolating the erosion rate all the way to touchdown on the lunar surface, a soil model is needed to handle the increased resistance to erosion as the deeper, more compacted soil is exposed. Relying on models derived from Apollo measurements and from Lunar Reconnaissance Orbiter (LRO) Diviner thermal inertia measurements, only one additional soil parameter is unknown: the scale of increasing cohesive energy with soil compaction. Treating this as an additional fitting parameter results in some degeneracy in the solutions, but the depth of erosion scour in the post-landing imagery provides an additional constraint on the solution. The results show that about 4 to 10 times more soil was blown in each Apollo landing than previously believed, so the potential for sandblasting damage is worse than prior estimates. This also shows that, with further development, instruments to measure the soil erosion during lunar landings can constrain the soil column's density profile complementary to the thermal inertia measurements, providing insight into the landing site's geology. • A new erosion rate theory is compared to dust blown in the Apollo 16 landing. • The theory correctly predicts the blowing dust's optical density curve. • A soil model is needed to predict erosion after the loose layer is removed. • Eleven to 26 tons of soil were blown, 4 to 10 times the prior estimates. • Soil blown by lunar landings is a worse problem than previously believed. [ABSTRACT FROM AUTHOR]

Soft landers are a common systems for planetary exploration and have been successfully landed on Moon, Mars and the comet Churyumov–Gerasimenko. In this study numerical simulations of the touchdown process of soft landers on the moon surface are presented. The focus is on the soil footpad interaction and the resulting loading on the primary strut of the landing gear. Main influence parameters like slope inclination, touchdown velocities and soil properties are varied to show each individual influence on the structural loading of the landing gear. The numerical simulations are done using the CEL-Method allowing for significant soil deformation at touchdown. [ABSTRACT FROM AUTHOR]

This work proposes a systematic methodology for designing an active leveling system (ALS) actuator for lunar landing application. The ALS actuator is integrated into an inverted tripod leg layout, exploiting a honeycomb crushable damper as a shock absorber. The proposed ALS actuator is fitted within the leg’s primary strut and features a custom permanent-magnet synchronous machine rigidly coupled with a lead screw. The actuator aims to both provide proper leg deployment functioning and compensate for the different shock absorber deformations during landing. The leg dynamic behavior is simulated through a parameterized multi-body model to investigate different landing scenarios. First, a parametric sensitivity approach is used to optimize the transmission system and the electric machine characteristics. Then, the electric motor model is numerically validated and optimized through electromagnetic finite element analysis. To validate the proposed ALS design methodology, a virtual test bench is used to assess the ALS performances under different load scenarios. It is found that the proposed methodology is able to yield a compact, well-sized actuator which is numerically validated with the EL3 platform as a case study.

• A DEM-FEM-MBD coupling algorithm is established to simulate the landing process of the lander, and is verified with the full-scale experiment of lunar lander on earth. • The influence of landing speed, lander mass and landing attitude on safe landing is analyzed based on DEM-FEM-MBD numerical modelling. • Two different inclined landing modes are compared with vertical landing to obtain the safe landing mode. • The energy consumption mechanism is revealed considering the buffer energy dissipation of lunar soil during the landing process. 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. [ABSTRACT FROM AUTHOR]

• Uncertainty of dynamic response of lunar lander landing procedure is considered. • The uncertain input parameters are modeled by random fields and interval fields. • The gradients in interval field are tuned towards gradients present in random field. • Impulse Based Substructure method is employed to improve the analysis efficiency. This paper presents an objective comparison of random fields and interval fields to propagate spatial uncertainty, based on a finite element model of a lunar lander. The impulse based substructuring method is used to improve the analysis efficiency. The spatially uncertain input parameters are modeled by both random fields and interval fields. The objective of this work is to compare the applicability of both approaches in an early design stage under scarce information regarding the occurring spatial parameter variability. Focus is on the definition of the input side of the problem under this scarce knowledge, as well as the interpretation of the analysis outcome. To obtain an objective comparison between both approaches, the gradients in the interval field are tuned towards the gradients present in the random field. The result shows a very similar dependence and correlation structure between the local properties for both approaches. Furthermore, through the transient dynamic estimation, it is shown that the response ranges that are predicted by the interval field and random field are very close to each other. [ABSTRACT FROM AUTHOR]

NASA is planning to launch robotic landers to the Moon as part of the Artemis lunar program. We have proposed sending a greenhouse housed in a 1U CubeSat as part of one of these robotic missions. A major issue with these small landers is the limited power resources that do not allow for a narrow temperature range that we had on previous spaceflight missions with plants. Thus, the goal of this project was to extend this temperature range, allowing for greater flexibility in terms of hardware development for growing plants on the Moon. Our working hypothesis was that a mixture of ecotypes of Arabidopsis thaliana from colder and warmer climates would allow us to have successful growth of seedlings. However, our results did not support this hypothesis as a single genotype, Columbia (Col-0), had the best seed germination, growth, and development at the widest temperature range (11-25 oC). Based on results to date, we plan on using the Columbia ecotype, which will allow engineers greater flexibility in designing a thermal system. We plan to establish the parameters of growing plants in the lunar environment, and this goal is important for using plants in a bioregenerative life support system needed for human exploration on the Moon. [ABSTRACT FROM AUTHOR]

In this paper, a method for predicting the landing stability of a lunar lander by a classification map of the landing stability is proposed, considering the soft soil characteristics and the slope angle of the lunar surface. First, the landing stability condition in terms of the safe (=stable), sliding (=unstable), and tip-over (=statically unstable) possibilities was checked by dropping a lunar lander onto flat lunar surfaces through finite-element (FE) simulation according to the slope angle, friction coefficient, and soft/rigid ground, while the vertical touchdown velocity was maintained at 3 m/s. All of the simulation results were classified by a classification map with the aid of logistic regression, a machine-learning classification algorithm. Finally, the landing stability status was efficiently predicted by Monte Carlo (MC) simulation by just referring to the classification map for 10,000 input datasets, consisting of the friction coefficient, slope angles, and rigid/soft ground. To demonstrate the performance, two virtual lunar surfaces were employed based on a 3D terrain map of the LRO mission. Then, the landing stability was validated through landing simulation of an FE model of a lunar lander requiring high computation cost. The prediction results showed excellent agreement with those of landing simulations with a negligible computational cost of around a few seconds.

Abstract The Moon is the nearest celestial body to the Earth and the starting point of the space exploration for the human being. The impact energy absorption and landing stability of lunar landers are extremely critical for lunar exploration. However, traditional landers cannot adapt to complex landing conditions and response to any unforeseen accidents during the landing process. Although some researchers have proposed the idea for the lander with MRF dampers, there are relatively few detailed structural designs and practical landing research studies. In this paper, we propose a detailed new type of lander with magnetorheological fluid dampers implemented as the primary struts, which can absorb the impact energy from landing. Its damping force is controlled by semi-active controlled currents based on the derived hydrodynamics of magnetorheological fluid dampers and can adjust to practical landing conditions. The simulation models of landers with the aluminum honeycomb and semi-active control are constructed in MSC Adams separately to compare their landing performances. Their simulation results show that this new type of lander can effectively reduce the largest acceleration by 26.18% under the largest acceleration response condition and the largest compression by 11.77% under the largest compression of primary struts condition. Its performance for more inclined and both smoother and rougher landing surfaces is better than that of traditional passive landers. It is believed that the proposed new type of lander is capable of landing on more complex and unexplored terrains for future lunar explorations. Highlights • A new type of lunar lander using magnetorheological fluid is proposed. • The damping forces of the new lander can be controlled by the currents. • The new lander has better landing performances under critical conditions. • The new lander is suitable for the more complex environment. [ABSTRACT FROM AUTHOR]

Abstract During the landing process of lunar landers, the lunar surface is eroded by the exhaust plume and a large amount of lunar dust entrained into a high-velocity spray. This high-speed lunar dust can have many adverse effects on the normal operation of both the lunar lander and any potential lunar surface facilities. It is necessary, therefore, to develop better ways to model this interaction between the plume and the lunar dust. This paper details the development of a macroscopic gas-particle two-way coupled method for simulating the interaction between the plume and the lunar dust, combining a dynamic method for simulating particle motion with the direct simulation Monte Carlo method for the rarefied plume flow field. This method considers macroscopic aerodynamic forces, convective heat transfer from the plume to the lunar dust, and the reaction of lunar particles to the plume, based on the principles of momentum and energy conservation and the interaction between the plume and its particles; it can meet the conservation conditions at each time step. A microscopic gas-particle two-way coupled method is also adopted to compare with the macroscopic method. This method considers collision behavior, momentum, and heat transfer to particles at the microscopic level, and realizes the effects of solid particles on gas by considering the behavior of gas molecules reflecting on the particle surfaces. The macroscopic simulation results show that particle size has a large influence on their spatial distribution. The results indicate that the particles decelerate as they interact with gas molecules; the particles also delay gas expansion. The results of the microscopic method differ from this method (e.g., slower particle velocities), due to its different theoretical basis. The results obtained by the macroscopic method appear to be more realistic, and this method can better demonstrate the serious impact of lunar dust on the lander and the surrounding environment. Highlights • The interaction between plume and lunar particles under vacuum condition is studied. • A macroscopic two-way coupled method is developed. • The results are obviously different between macroscopic and microscopic two-way coupled method. • Each time-step's momentum and energy conservation between two phases are guaranteed by macroscopic method. [ABSTRACT FROM AUTHOR]

During the descent and ascent of the lunar lander, the gas plume exhausted from its engine imposes a significant thermal load on the lander when it approaches the lunar surface. This challenge becomes more pronounced for multi-engine landers due to the intricate interactions between their plumes, making it difficult to ensure adequate thermal protection for the lander. In this study, a thermal protection method utilizing a supersonic cold gas film is proposed to shield the lander bottom from high-temperature backflow from the lunar surface. Unlike conventional film cooling that relies on the boundary layer flow, our approach is specifically designed to accommodate irregular surfaces of the lander bottom. The applicability of this heat protection method in vacuum or low-pressure environments is first confirmed. Subsequently, the effects of three key annular injector parameters, including total pressure, expansion ratio, and injection angle are investigated. Our findings reveal that increased coolant total pressure results in better protection, while the effects of the expansion ratio vary with different coolant total pressures. Under our conditions, an annular injector angle of 45∘ is determined to be the optimal choice for enhancing cooling efficiency. [ABSTRACT FROM AUTHOR]

As a growing number of exploration missions have successfully landed on the Moon in recent decades, ground infrastructures, such as radio beacons, have attracted a great deal of attention in the design of navigation systems. None of the available studies regarding integrating beacon measurements for pinpoint landing have considered uncertain initial beacon locations, which are quite common in practice. In this paper, we propose a radio beacon/inertial measurement unit (IMU)/altimeter localization scheme that is sufficiently robust regarding uncertain initial beacon locations. This scheme was designed based on the sparse extended information filter (SEIF) to locate the lander and update the beacon configuration at the same time. Then, an adaptive iterated sparse extended hybrid filter (AISEHF) was devised by modifying the prediction and update stage of SEIF with a hybrid-form propagation and a damping iteration algorithm, respectively. The simulation results indicated that the proposed method effectively reduced the error in the position estimations caused by uncertain beacon locations and made an effective trade-off between the estimation accuracy and the computational efficiency. Thus, this method is a potential candidate for future lunar exploration activities.

Dynamic analysis of soft-landing is very important for the design of lunar lander. At present, the determined landing attitude and speed are considered while not considering the uncertainty of these parameters in the analysis of soft-landing dynamics. Based on Chebyshev interval analysis method, an analysis process of landing dynamic interval based on nonlinear finite-element model is proposed for the dynamic characteristics of landing process. The upper and lower bounds of dynamic response are calculated using Chebyshev method and compared with the simulated results of Monte Carlo method. Comparative result shows that the analyzed results of Chebyshev interval analysis method can fully cover those of Monte Carlo method, and the dynamic interval is not enlarged. The influence of truncation order on the analytic error of dynamic interval was analyzed. The analyzed result shows that the truncation order has little influence on analysis error. Chebyshev method has the advantage of high accuracy and efficiency. [ABSTRACT FROM AUTHOR]

Abstract Radio detection of Ultra High Energetic Cosmic Rays and Neutrinos (UHECRv) which hit the Moon has been investigated in recent years. In preparation for near-future lunar science missions, we discuss technical requirements for radio experiments onboard lunar orbiters or on a lunar lander. We also develop an analysis of UHECRv aperture by including UHECv events occurring in the sub-layers of lunar regolith. It is verified that even using a single antenna onboard lunar orbiters or a few meters above the Moon's surface, dozens of lunar UHECRv events are detectable for one-year of observation at energy levels of 10 18 – 10 23 eV. Furthermore, it is shown that an antenna 3 m above the Moon's surface could detect lower energy lunar UHECR events at the level of 10 15 – 10 18 eV which might not be detectable from lunar orbiters or ground-based observations. [ABSTRACT FROM AUTHOR]

In this paper, the soft-landing analysis of a lunar lander spacecraft under three loading case was carried out in ABAQUS, using the Explicit Finite Element method. To ensure the simulation result's accuracy and reliability, the energy and mass balance criteria of the model was presented along with the theory and calculation method, and the results were benchmarked with other software (LS-DYNA) to get a validated model. The results from three loading case showed that the energy and mass of the models were conserved during soft landing, which satisfies the energy and mass balance criteria. The overloading response, structure steady state, and the crushing stroke of this lunar lander all met the design requirements of the lunar lander. The buffer energy-absorbing properties in this model have a good energy-absorbing capability, in which up to 84% of the initial energy could be dissipated. The design parameters of the model could guide the design of future manned landers or larger lunar landers. [ABSTRACT FROM AUTHOR]