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SSOLVE is a heterodyne spectrometer that is under development as a technology demonstration for future lunar lander missions. The system contains a heliostat to target the Sun, optics, and submillimeter-wave, and Terahertz (THz) receivers to measure the column abundance of H 2 O, OH, and HDO.

In this work, problem of pinpoint soft landing of the “Resource Prospector Lander” (RPL) on the moon's surface is presented using the sliding mode control (SMC) Aero toolbox. The use of SMC Aero toolbox for the RPL control design simplifies the implementation of such a complicated control paradigm as relative degree-based SMC. The SMC Aero toolbox that include first order sliding mode control (1-SMC), second order sliding mode control (2-SMC), higher order sliding mode control (HOSMC), adaptive 1-SMC/2-SMC/HOSMC as well as higher order sliding mode differentiators are employed for the attitude and descent off-pulse controller design to achieve RPL soft landing on the moon's surface in the presence of the model perturbations caused by center of mass misalignments and actuator/thruster malfunctions. Adaptive SMC implementation of the RPL pulse-off strategy is facilitated using SMC Lunar Lander library of the SMC Aero toolbox. The efficacy of RPL control design using SMC Aero toolbox is illustrated via numerical simulations.

As NASA strives to return to the lunar surface with the Artemis program, there is a need to understand if an expendable or reusable architecture should be used. To address this question, this work presents a multidisciplinary model that explores the lunar lander design space to understand how reusability affects the chosen objectives of IMLEO, lifecycle cost, and safety. For the Artemis program, we found an expendable design to be more cost-effective than implementing reusability, as only five surface missions are currently planned. However, if additional missions are added, a hybrid implementation of reusability, which entails reusing the transfer and ascent elements multiple times by changing out line-replacement units after each mission, becomes the most cost effective. A fully reusable design, where the transfer and ascent elements only require consumables to be replenished between missions, was also investigated, although it was less cost effective than a hybrid design for the range of values studied.

Growing interest in autonomous landing on the lunar surface has motivated in-depth studies of the navigation architectures to support such missions. The research presented in this paper extends linear covariance models of terrain-relative navigation architectures to incorporate both image-based measurements to known landmarks and laser range measurements to the lunar surface. The benefit of the range measurements is shown to be significant in reducing relative position and velocity errors in the altitude channel, as compared to an image-only navigation architecture. The range measurements are also shown to reduce the sensitivity of final relative navigation errors to the turn-on time of both image and range measurements, resulting in a more robust system design.

So far, by the majority, most of the components used in space-related applications are specially designed for the very harsh space environment. However, these types of space-grade or military-grade Hi-reliability electrical components lack design flexibility due to limited options and cost a lot of money and time to acquire. Many private venture companies are entering the space industry with limited capital and human resources. It is very critical for them to build high-performance space systems with limited resources. This paper gives a brief overview of challenges faced during the use of Commercially Off-The-Shelf (COTS) semiconductor components for space application.

In response to the 2018 White House Space Policy Directive-1, to lead an innovative and sustainable lunar exploration, and to the Vice President's March 2019 direction to do so by 2024, NASA is working to establish humanity's presence on and around the Moon by: 1) sending payloads to its surface, 2) assembling the Gateway outpost in orbit and 3) demonstrating the first human lunar landings since 1972. NASA's Artemis program is implementing a multi-faceted and coordinated agency-wide approach with a focus on the lunar South Pole. The Artemis missions will demonstrate new technologies, capabilities and business approaches needed for future exploration, including Mars. Assessing options to accelerate development of required systems, NASA is utilizing public-private engagements through the Human Exploration and Operations (HEO) Mission Directorate's NextSTEP Broad Agency Announcements. The design, development and demonstration of the Human Landing System (HLS) is expected to be led by commercial partners. Utilizing efforts across mission directorates, the Artemis effort will benefit from programs from the Science Mission Directorate (SMD) and Space Technology Mission Directorate (STMD). SMD's Commercial Lunar Payload Services (CLPS) initiative will procure commercial robotic lunar delivery services and the development of science instruments and technology demonstration payloads. The Space Technology Mission Directorate (STMD) portfolio of technology advancements relative to HLS include lunar lander components and technologies for pointing, navigation and tracking, fuel storage and transfer, autonomy and mobility, communications, propulsion and power. In addition to describing the objectives and requirements of the 2024 Artemis mission, this paper will present NASA's approach to accessing the lunar surface with an affordable human-rated landing system, current status and the role of U.S. industry in 2024 and for a sustainable lunar presence.

Spacecraft equipped with liquid propulsion systems are susceptible to propellant sloshing during orbit correction maneuvers. During a maneuver, the liquid propellant is pushed towards the bottom of the tanks due to the effective gravity of the thrusters, resulting in lateral liquid motion and a sloshing phenomenon. The sloshing effect, if not mitigated by the attitude control system, may cause dynamic instabilities and reduce maneuver performance. It is difficult to synthesize fluid mechanics models in a control design process, and spacecraft dynamics under sloshing remains a great challenge in terms of attitude control design. Mechanical analogies, such as pendulums or spring-mass systems, have been found to capture the main effects of liquid sloshing and can serve as a control design model; however, such mechanical analogies are inherently uncertain and a robust control methodology is required for assuring the spacecraft stability and performance during a maneuver. In this work, the quantitative frequency design approach is applied for the first time to provide robust controllers for spacecraft with multi propellant tanks and an attitude dynamics model incorporating an uncertain sloshing mechanical analogy. This methodology has been used to design the attitude control of SpaceIL's lunar lander Beresheet and was proved successful in 10 orbit correction maneuvers. A case study of one of these maneuvers is presented as well as telemetry results from the mission.

There is great interest in missions to small bodies in our solar system such as asteroids, comets, Centaurs, dwarf planets, and moons. A key aspect of these bodies that makes them interesting is their varying morphologies over (and under) their surfaces. To characterize these bodies with in-situ exploration, a surface landing architecture is necessary that allows for the study of multiple diverse areas. Current stationary surface landers have only a short range, limited to capabilities of their on-board instruments and robotic arms. Surface rovers, while more mobile, are inherently more complex than landers and incur high costs both in development and operations. A surface lander capable of performing multiple hops could not only lower the cost of this mobility significantly, but also allow for the characterization of even more diverse areas than a rover. While a rover can cover tens of kilometers over the course of its lifetime, a hopper could potentially perform multiple hops of that same distance. While multiple mission concepts have been proposed (CHOPPER, Mars Geyser Hopper, GRUNT), and some have even flown (MINERVA-II, NEAR Shoemaker, Philae), the only mission to execute a planned hop on the surface of another body is the lunar lander, Surveyor 6. Performed in November of 1967, the results of this three-meter hop experiment contain lessons that will aid the development of future hopping mission concepts. This paper will first summarize previous hop attempts - both intentional and unintentional-and proposed concepts for future hopping spacecraft. A more in-depth study of the Surveyor 6 mission will be included, and lessons learned from the study will be applied to other small bodies. Analysis of ballistic, Surveyor 6-like hops, are performed for a diverse catalogue of small bodies including dwarf planet Ceres, comet 67P, and the moon Triton. Feasibilities of hops on each body are then evaluated based on distance and $\Delta-\mathbf{V}$ , among other considerations. The utility of another style of hop, the “jump-and-wait” method, is also examined for the same bodies and compared to the Surveyor style hop results. The paper will conclude with a summary of the applicability of various types of hops for a variety of small bodies.

This paper discusses the adoption of Simulation Based Design (SBD) approaches involving 3D Virtual Reality (VR) based digital mockup models to support the study of transportation and assembly alternatives in the context of Deep Space Missions as well as landing on the Moon. An initial set of designs have been proposed using immersive VR environments. The modeling, planning and simulation scope includes the transfer of equipment from the lunar lander to the moon surface, design of astronaut habitats and assembly of such habitats using robots. Using this SBD approach, the project team proposed, compared and analyzed transport and assembly alternatives for these design contexts. The simulation environments were built using the immersive Vive VR platform and Unity 3D software tools.

This paper focuses on the study of the scattering of radar waves due to rugged topography relief. It is a helpful assistance to identify features in GPR reflection profile to separate true subsurface reflectors and artifacts caused by surface scatterers. Two examples are presented here to illustrate the approach. The first case is for a numerical simulation of the edge diffraction from rugged topographic relief on Earth. The other is one from the numerical model of the 60-MHz radar wave propagation to look into the edge diffraction caused by the impact crater depressions based on the real topography and stratigraphy of the lunar surface at the landing site of the 2014 Chinese lunar lander Chang'e-3. The purpose of this study is to explore the possibility to explain the featured events in the lunar penetration radar (LPR) profile recorded by the Yutu lunar rover, which were originally identified as subsurface reflectors from stratigraphic interfaces. It is noteworthy to point out that the 60-MHz radar system on the Yutu lunar rover uses the unshielded rod antennas. Substantial portion of the radiated energy is in the air above the lunar surface. The simulation of lunar radar wave propagation suggests that it is likely that the later time events in the 60-MHz LPR profile might be associated with the focused surface scattering waves from the edge of the impact craters near the Chang'e-3 landing site.

As NASA focuses on Mars mission readiness, a crewed lunar lander concept may be timely, as there are both Mars exploration and lunar exploration motives for its development. A lunar lander can be used as a Mars lander technology demonstrator, including demonstrating technologies used in the final phase of Mars descent/landing, as well as technologies/techniques for habitation and ascent to orbit. Utilizing the cislunar Deep Space Gateway and Space Launch System (SLS) transportation capabilities, the addition of a lander would enable international and scientific communities to explore the lunar surface The Boeing lunar lander concept supports a crew of four on a two week (one lunar day) surface sortie and is extensible to longer duration missions with the addition of separately delivered surface assets such as a mobile habitat. The lander is optimized to take advantage of the cislunar Deep Space Gateway architecture; the lander elements are delivered using the SLS Block 2 cargo launch capability and the SLS Orion/cargo co-manifest capability. Overall mass of the ascent module was minimized by allocating habitation and crew surface access functions to the descent module. Recurring costs were reduced by designing a reusable ascent module. This paper will describe the lander design and operation concept, key trade studies and rationale for the selected options. Trade studies were conducted in areas such as spacecraft elements, engines, propellant, airlock, and life support subsystems. Analysis of the Deep Space Gateway orbit is discussed with respect to its compatibility with: a) surface landing operations, b) earth-to-orbit element/cargo delivery operations, and c) access to various lunar surface regions. The paper will show that use of the SLS, Orion and the Deep Space Gateway in support of a human lunar lander exploration architecture is feasible and practical and will explain how the design described can be extensible to future Mars lander designs.

TeamIndus lunar landing mission objectives include enabling lunar lander to soft land on Moon, traverse the rover on the surface of Moon and send high definition videos and images to Earth. The lander is controlled by a Lander On-Board Computer (LOBC) card consisting of a SPARC processor. The spacecraft is required to be commanded by ground station for multiple scenarios: execute commands immediately, execute commands at a particular time, execute sequence of commands, execute commands after certain event has expired and execute memory management commands. The limited availability of uplink bandwidth requires optimal design of telecommand packet structure. In order to provide ease of operation of spacecraft and reuse of already uplinked command, programmability and flexibility features are provided by telecommand processing (TCP) software. This paper describes telecommand processing software implementation on LOBC and packet structure design for Real time commands, Absolute time tagged commands, Differential time tagged commands, Patch commands, Configurable command blocks and Event based commands. The autonomy features of the designed TCP software has minimized ground intervention and hardcoding of sequence based modules. The general purpose TCP software does not change with modifications in the packet structure or commands in the mission command dictionary, making it reusable across various missions.

Space field has experienced vigorous advancement with respect to evolution of vision system, image storage and processing. Real time image processing has become one of the most important tools for navigation and landing for planetary and lunar missions. Information of horizontal velocity with high accuracy will be required to do accurate pin point landing. For the testability of such a system as well as to have the understanding of Lander dynamics, prior landing image sequences are required to initially testing the algorithm [1]. This paper deals with FPGA based processing on Image sequence to find the relative velocity and also implements Image Storage in a microSD card. Landing sequence is stored in SD card and post landing they are downloaded. These images provide a very useful information about lighting, lander dynamics to fine tune algorithm for future mission. Besides this image data serves the testability during various phases of testing during development.

This paper is on polynomial guidance law for future lunar lander in power descending phase. To achieve pin-point landing, the guidance trajectory needs to be accurate. Boundary conditions are required to solve the polynomial guidance law, but the guidance law is designed in bases where the boundary conditions are nominal. This paper proposes the method to obtain boundary conditions from approximation when initial error occurs and give optimal trajectory even with the initial error. This paper shows the series of different initial errors and discusses the effect of the approximate error to the trajectory and fuel consumption. The results give some ideas for future navigation technology for lunar landing, which has error at the start of their power descending phase.

The Moon-based Extreme Ultraviolet Camera (EUVC) installed on the lunar lander of Chinese Chang'e 3 mission can obtain global and instantaneous EUV images of the Earth's plasmasphere in meridian view (side view). The movies of the continuously observed time-series images by EUVC reveal the dynamic variations of the plasmasphere during substorms. The EUVC instrument, observations of plasmasphere, data products, and main scientific results are summarized in this paper.

In this study, we will present the novel application of Type-2 (T2) fuzzy logic to the popular video game called Lunar Lander. The proposed T2 fuzzy moon landing system structure is composed of the error signal generator and the T2 fuzzy logic control structure which give the opportunity to transform the moon landing problem of the spaceship as a multivariable tracking control problem. The landing problem of the game can be seen as one of the classical multivariable control problems including uncertainties due to the randomization process occurring the game environment. Thus, we will employ T2 fuzzy logic controllers since they are capable of handling a high level of uncertainties. Then, by optimizing the T2 fuzzy moon landing system via the particle swarm optimization, we will show that the resulting T2 fuzzy moon landing system resulted with an adequate control and game performance in the presence of the uncertainties, disturbances and nonlinear system dynamics in comparison with its type-1 and conventional counterparts. We believe that the results of this paper will be an important step for a wider deployment of T2 fuzzy logic in the research area of computer games.

Structural samples removed from the Surveyor 3 probe and returned to Earth by the Apollo 12 astronauts have shown that close proximity lunar landing operations could damage previously landed lunar hardware. During the final landing approach the jet engines of the Apollo 12 lunar lander stirred up lunar soil. Parts of the dust flow impacted on the previously landed Surveyor 3 and “sandblasted” the surface of the probe. Similar scenarios might occur during future lunar missions as part of resupply or sample return scenarios. The presented experimental work focuses on the identification and quantitative characterization of possible damages to technical surfaces caused by lunar dust impacts for different impact angles and velocities. It can be assumed that, during lunar landing operations, the majority of the stirred-up dust particles will not impact perpendicular to the target surface. Therefore, impact experiments using particles of the lunar analogue material JSC-1A were conducted at different impact angles at impact velocities of approximately 200 m/s for grain sizes between 450 and 560 microns. In the current study 10 mg of the dust particles were accelerated with an electromagnetic eddy current accelerator. The accelerated lunar dust simulant particles were impacted on three different technical materials, Al6061-T6, silica glass, and Makrolon® (polycarbonate). Microscopic images of the impact area on the specimens and the mass change of the specimens resulting from the impacts were used to identify whether the major interaction between surface and particles is erosion or particle adhesion. The two ductile materials, aluminum and polycarbonate, are mainly affected by particle adhesion or the embedding of particles, while chipping is the major surface degradation effect for silica glass specimens. However material removal could only be measured for an impact angle of 15°. Macroscopic degradation phenomena were quantified by measuring the alteration of surface properties of the specimens. Change in surface properties were measured as the change of visible light transmission for both optical media, while the surface roughness (Ra) was measured to quantify the change of surface properties for the aluminum specimens. The results show a significant correlation between impact angle and the change of surface properties regardless of the investigated material. It was found, that the main degradation effects for surface roughness as well as for the transmission occur between impact angles normal to the surface and 15°, while the surface properties are less affected by dust impacts towards grazing impact angles. A comparison of these results with previously conducted experiments, which investigated the impact wear of JSC-1A particles, was done. The previous study impacted 10 mg of JSC-1A with grain sizes between 250 and 350 microns at different impact angles at impact velocities of approximately 360 m/s under the same experimental conditions. The comparison showed that the surface properties are more affected by the 360 m/s-impacts, although the single particle has less kinetic impact energy. The reason is that more particle impacts occur per unit area for the faster but smaller particles.

Lunar dust grains threaten the safety of lunar lander and astronauts, because of adhesion. In order to reduce their threat, electrostatic adhesion mechanism and adhesive mechanics model of lunar dust are studied by using Discrete Element Method, Finite Element Method and Image Method in this paper. An electrostatic adhesive mechanics model is derived with electrostatic field theory and verified based on the adhesive force plug-in which is developed by applying the API function on the software platform of EDEM. The model and corresponding simulated adhesive force values can provide a theoretical support for attempts on reducing lunar dust contamination.

Real-time performance metrics were developed to quantify workload, situational awareness, and manual task performance for use as visual feedback to pilots of aerospace vehicles. Results from prior lunar lander experiments were replicated and extended to provide insights for the development of real-time metrics. Automated Speech Recognition (ASR) was employed to detect verbal callouts and measure subjects' situational awareness. ASR achieved an F-measure of 0.8, despite large variances within subjects. A one-dimensional manual tracking task and simple instructor-model visual feedback scheme was developed. Experiments showed that for this low-complexity task, visual feedback did not change subject performance, but did increase the subjects' measured workload.

In this paper, to reduce batch process TRN (Terrain Relative Navigation) correlation problem caused by large position errors and for the improvement of computational effort, MRT (Mean Removal technique) and two step search methods are applied for lunar lander. MRT is the method subtracting the mean of both the measurement and DEM profile is applied. And two step search is the method using overall coarse and partially fine DEM. To meet position accuracy for precise lunar landing, a lunar navigation is operated by combination of INS (Inertial Navigation System) and TRN using altimeter, LIDAR etc. Until altimeter and LIDAR are operated, kilometers of cumulative errors caused by INS with time occur. Since these position errors induce bad correlation with measured terrain height profile and DEM (Digital Elevation Map) profile and cause wrong position fix, TRN considering large position error should be studied. By using MRT and two step search method, improvement of TRN position correction is shown in simulation.

This paper presents how to estimate two parameters: slope and roughness representing lunar terrain quality for HDA (Hazard Detection and Avoidance) of lunar lander. Robust plane fitting technique based on LMS (Least Median Square) is applied to standardized DEM (Digital Elevation Model) from LIDAR (Light Detection and Ranging) sensor measurement in order to estimate the parameters. Since a number of hazards exist on the lunar surface such as craters, rocks and cliff, which are able to damage lunar lander, it is important for HDA system to precisely detect hazards and select safe landing site by meticulously characterizing terrain condition. For this reason, out research has been conducted using various sensors such as camera, infrared sensor, LIDAR sensors for HDA algorithm. Because sensor measurement can be processed in high speed relative to optical image and obtained at any light condition, it is an effective way to perform hazard detection and avoidance landing based on the LIDAR measurement. Therefore, it is confirmed that safe landing site can be extracted based on the parameters through simulation in this paper.

To design the optimal lunar landing trajectory, the periapsis of de-orbit burn phase is usually used as a starting point of the powered descent phase. And the optimal problem is constructed based on the initial states at this point. For this optimal problem, some kinds of result trajectories have the increase in their own altitude at the early stage to earn the enough time to reduce their huge horizontal velocity for the low thrust, and this phenomenon increase the consumption of the propellant mass. For this reason, it is possible to exist another phase angle to reduce this phenomenon instead of zero for periapsis of de-orbit burn phase because the vertical velocity is not zero at this point. To find out this optimal phase angle, in this paper, the initial free state optimal problem method is applied without any additional assumption and change of hardware for the traditional optimal lunar landing problem. Using this approach, it shows that the optimal phase angle does not always equal to the periapsis for some cases, and it is possible to reduce the fuel consumption of the lunar lander.

In view of the landing site positioning of the lunar probe, this paper proposes a positioning reduction method based on VLBI tracking data, and applies VLBI delay-observation data from more than three baselines to complete positioning reduction with the distance between the lunar lander and the center of the earth as a constraint. Simulation results show that the positioning accuracy of landing point of the lunar lander can reach 100 m if the standard deviation of the VLBI delay-observation noise is less than 1 ns (0.3 m). This method also supports positioning reduction of other deep space probes' precise landing.

A human-in-the-loop experiment was conducted at the NASA Ames Research Center Vertical Motion Simulator, where instrument-rated pilots completed a simulated terminal descent phase of a lunar landing. Ten pilots participated in a 2 x 2 mixed design experiment, with level of automation as the within-subjects factor and failure frequency as the between-subjects factor. The two evaluated levels of automation were high (fully automated landing) and low (manual controlled landing). During test trials, participants were exposed to either a high number of failures (75% failure frequency) or low number of failures (25% failure frequency). In order to investigate the pilots' sensitivity to changes in levels of automation and failure frequency, the dependent measure selected for this experiment was accuracy of failure diagnosis, from which D Prime and Decision Criterion were derived. For each of the dependent measures, no significant difference was found for level of automation and no significant interaction was detected between level of automation and failure frequency. A significant effect was identified for failure frequency suggesting failure frequency has a significant effect on pilots' sensitivity to failure detection and diagnosis. Participants were more likely to correctly identify and diagnose failures if they experienced the higher levels of failures, regardless of level of automation.

This paper presents a safety estimation methodology for soft landing explorations on the Moon. In our application we use digital elevation models (DEM) of given lunar terrain extracted from on-board lidars. Currently most landing safety analysis methods are based on topographical analysis which needs quite a lot of computation. Here we adopt a Monte-Carlo simulation method which doesn't need to identify complex lunar topographical characteristics and needs less computation. After developing lunar lander's collision detection model, lander gradient model, and lander aspect model, Monte-Carlo simulations of landing on given DEM are conducted iteratively. Safety estimation is implemented by analyzing corresponding simulation results using statistical analysis method. Simulation errors of this algorithm are also deduced and analyzed in this paper. Computer simulation results show that this method can provide valuable information to estimate landing safety of given terrain for lunar soft landing missions.

This paper investigates the relationship between the difficulty and the interestingness of individual problem candidates from within a class of related problems, using Lunar Lander as a case study. In this class of problems, a 2D spaceship must be controlled by a simple set of macro-actions, including both linear and angular impulses, such that it fulfils a set of weighted criteria relating to landing on a jagged landscape with flat landing pads. It is demonstrated that a very simple measure based on standard deviations of improvement can be used to guide evolution to develop interesting problems in this class of problems, which in turn can be solved using evolution strategies to get a high level of improvement based on initial random performance. We examine the impact of the measure used on the evolution of the problems, and also what aspects of this problem class affect the difficulty and interestingness the most.

One of the major tasks in the design and optimization process of a new landing gear system for a lunar lander is to accurately determine the loads and energy absorption capability during the landing event. The works of this paper describes a new approach of landing impact dynamic analysis using nonlinear finite element method. Abaqus/Explicit, which is part of Abaqus suite of finite element analysis software, is selected to simulate the landing event for its excellent nonlinear, transient dynamics capabilities. The aluminum honeycomb shock absorber is modeled by plastic crushable foam material model, while the lunar soil is modeled by Drucker-Prager/Cap material model. Simulation results, including the load at connector between structure and landing gear, acceleration response of structure and dissipated energy by the shock absorber and lunar soil, are given and discussed. It shows that the performance of landing gear meets the design requirements.

This project focuses on identification of craters in terms of its characteristics and detection of these visual features of the moon to determine a safe landing site for a lunar Lander. Cheng et al. proposed using craters as landmarks for navigation purposes because this geometric model grants a robust detection under different lighting conditions. Moreover, craters appear in enough density on most planetary system bodies of interest and also known as a fairly stable in appearance or shapes over time or under different conditions and environments. These special features make them an appropriate type of landmark to observe. Currently, there are a lot of ongoing researches mainly on craters detection and optical navigation systems for the moon but still using a complex and similar approach such as detection using the Hough transform method. To part from this limitation, the author decides to build a simple algorithm for detecting craters on a moon surface which will detect the craters based on two important measurements that are the distance and angle measurements. The advantages of using this approach are (1) its uncomplicatedness (2) fast detection (3) can be used further in ellipse reconstruction algorithm to determine the position and orientation of the crater. This paper will discuss the method of employing MATLAB and image processing tool on an optical image as well as the morphological image detection fundamentals. In addition, some geometrical projection analysis in reconstructing an ellipse as a disc will be evaluated in order to obtain the orientation of the disc (crater) for autonomous optical navigation system.

It is necessary to investigate the lunar lander soft-landing by considering the flexible body and legs because the impact of deformation is significant. There is a new method for research on landing impact performance in this paper. Firstly a dynamical model of 1/4 lunar lander with four suspension legs is build and simplified. After getting oscillation equation of the model, founds two simulation models respectively by multi-rigid-body dynamics method and nonlinear finite element method. The simulation of soft-landing selects MSC.ADMAS and MSC.DYTRAN. Main analysis objects are buffer distance of primary strut in lunar lander and vertical acceleration of box's centroid, then contrasts two simulation results based on the objects. Research results show that the influence of flexible body reduces the maximum of primary strut's buffer distance, but amplifies the vertical acceleration of centroid with periodic oscillation.

To ensure the lunar lander landing safely, and execute the expected scientific experiments smoothly, a new terrain simulation method is proposed in this paper, which can build a terrain model of Sinus Iridum with a high resolution of better than 1 meter. The terrain surface trend is fitted from a low-resolution terrain with genetic programming (GP) and genetic algorithm (GA). A high resolution terrain with a resolution up to 1 cm in a 10m×10m region is built by applying the distribution laws and the models of craters and rocks in Sinus Iridum region. This method can be used to analyze the landing safety of lunar lander and decide a safe landing site by simulating precisely the real terrain characters.

With the goal of developing a preliminary hardware and software design of a Guidance, Navigation and Control Subsystem of a robotic lunar lander for various potential missions including the International Lunar Network (ILN), navigation sensor technologies are researched and their maturity is assessed. 12The results of the broad technology assessment are compared with requirements generated from known or estimated design constraints, modeling and simulation results, and subject-matter expert consultations to validate the assembly design. The results identify the current baseline and some alternative components for the lander. The nominal sensor configuration baselines assume that precision landing is desired for the mission even though some of the mission concepts, including the ILN mission, could be achieved without precision landing. In these cases, the specific sensors used for precision terrain relative position measurements are removed from the configuration.

Landing dynamics analysis is necessary for the success of the lunar lander soft-landing. Recent advances in computational speed have made lunar lander soft-landing modeling and simulation using a nonlinear, transient dynamics, finite element analysis code more feasible. A new approach to study the soft-landing of lunar lander has been described in this paper. The simulation of lunar lander soft-landing selected MSC. DYTRAN. By taking a certain type of lunar lander as the research object, the flexible models of the lander structure and the lunar regolith are established. The lander has four landing legs, and has corresponding shock absorbers using Al-honeycomb. Two specific nonlinear characteristics of the honeycomb and the lunar regolith of simulation model are preset, and the nonlinear frictions are considered. The results are then compared to and validated by test data. Simulation results are presented to investigate the performance of energy absorbing system in soft-landing. The time history of internal energy of the lander structure and the lunar regolith during the entire soft-landing event are obtained.

The influence of flexible body and legs of lander, plastic deformation of lunar regolith on landing impact performance of lander is rarely researched, and literatures about using transient dynamics to analysis soft landing impact performance of lunar lander are also less. In this paper, a new idea and method for research on landing impact performance is presented. Flexible models and transient dynamics are combined to research soft landing impact performance of lunar lander. Firstly flexible models of lunar lander and lunar regolith were built using MSC.PATRAN, and the transient dynamics simulation software MSC.DYTRAN was used to simulate and research the landing impact performance of lunar lander soft landing along uphill. The research results show that the actual landing conditions of lunar lander are simulated more realistic by using flexible models and transient dynamics method together. When lunar lander's four legs land along uphill terrain non-symmetrically, landing impact load of the lunar lander is the least and energy absorption efficiency of the lander's buffering is the highest.

Considering the requirements for the lunar rover positioning in the future lunar reconnaissance, several methods of lunar rover positioning are compared and analyzed firstly, and a new method is proposed in this paper according to the characters of the lunar rover activities, which uses the asynchronous transmission incoherent spread-spectrum communication/ranging system, and dual-UHF-antenna short-baseline carrier coherent angle detection method, to achieve high accuracy of ranging and azimuth angle results, and can figure out the rectangular coordinates position and vector of the lunar rover. The autonomous and real time precise positioning of the lunar rover over the whole activities on the moon can be realized with the use of the new method proposed here.

In this paper, the key influences of lunar lander landing model are illuminated systematic. Then the touchdown dynamic model of lunar module is built. And the simulation model for lunar landing described on ADAMS is discussed. As an example, one kind of lunar module model with four cantilever truss style landing gear is built by ADAMS and the simulations are focused on the influences on landing by the characteristics of buffers in primary and secondary struts of the landing gear. The conclusions can be used for the design of lunar lander.

The goal of this paper is to describe a first-order performance analysis of a lunar hopper 1,2. A hopper is a vehicle that has both landing and surface mobility capabilities on a single platform. Unlike rovers, which traverse the lunar surface while in contact with the ground, hopping reuses the landing propulsion system to lift back off again and “hop” over the lunar terrain. Hopping, as a form of surface mobility, is a novel concept. As such, analysis must be performed to assess how it would fit with an overall lunar landing system architecture. Two trajectory categories are investigated to perform this assessment: the ballistic hop, where the vehicle launches itself into a ballistic trajectory toward the destination, and the hover hop, in which the vehicle ascends and maintains a constant altitude as it travels toward its desired location. Initially, parametric studies of the ballistic and hover hop are carried out in order to make observations about the performance of each hop. Using this data, it is possible to investigate the fuel-optimal hop trajectory. The delta-V costs for the ballistic and hover hops are compared for hop distances between 500 meters and 5000 meters, and in this range it is found that the ballistic hop and hover traverse have comparable delta-V costs. For the entire hop maneuver, however, the hover hop will always be the more delta-V expensive option due to the ascent and descent phases. Nevertheless, this does not rule out the hover hop as a feasible option due to its operational advantages over the ballistic hop.

The High Performance Processor (HPP) Task of the Advanced Avionics and Processor Systems (AAPS) Project, part of the Exploration Technology Development Program (ETDP), was to evaluate several high performance multicore processor architectures with respect to their ability to provide real time hazard detection and avoidance for the Constellation Program's Altair Lunar Lander. 12In this paper we review the Tilera Tile64 processor, the hazard detection and avoidance algorithm, strategies for parallelizing these algorithms, and preliminary performance study results. We were presented with the requirements of 30 Hz LIDAR frame processing rate and 10 second processing time for ALHAT HDA processing and were able to meet that requirement with the Tile64. We then project the performance of these algorithms on the OPERA MAESTRO Processor, a radiation tolerant version of the Tile 64 being developed by the Boeing Company.

This paper1,2 uses linear covariance analysis to investigate the expected navigation performance and corresponding parameter sensitivities for lunar landing and surface mobility (via “hopping”). The Draper Linear Covariance (LinCov) tool is a statistical simulation that approximates the state error covariance results of a six-degree-of-freedom Monte Carlo simulation in a fraction of the time. This ability to quickly analyze navigation performance for a wide variety of sensor suite configurations given any landing or hopping trajectory allows for increased vehicle and mission design flexibility, as well as insight into sensor and vehicle cost, mass, and power sizing.

As demand for highly reliable complex systems increases, engineers are being forced to consider the risk implications of design decisions earlier in the conceptual phase of projects and with greater accuracy. In highly mass constrained systems, in order to achieve increased reliability, buying down risk with mass cannot be considered from a ‘stove-piped’ stand point, but rather it must be approached from a global system perspective. The Altair lunar lander design team was able to implement a process of Risk Informed Design utilizing the Valador Reliability Tool [1] (VRT) in order to achieve high reliability. This tool is able to quickly and accurately produce estimates of the risk of Loss of Mission (LOM) and Loss of Crew (LOC) and provide insight to the designers as to how their decisions will impact overall mission success.

Eight pilots participated as test subjects in a fixed-based simulation experiment to evaluate advanced vision display technologies such as Enhanced Vision (EV) and Synthetic Vision (SV) for providing terrain imagery on flight displays in a Lunar Lander Vehicle. Subjects were asked to fly 20 approaches to the Apollo 15 lunar landing site with four different display concepts — Baseline (symbology only with no terrain imagery), EV only (terrain imagery from Forward Looking InfraRed, or FLIR, and LIght Detection and Ranging, or LIDAR, sensors), SV only (terrain imagery from onboard database), and Fused EV and SV concepts. As expected, manual landing performance was excellent (within a meter of landing site center) and not affected by the inclusion of EV or SV terrain imagery on the Lunar Lander flight displays. Subjective ratings revealed significant situation awareness improvements with the concepts employing EV and/or SV terrain imagery compared to the Baseline condition that had no terrain imagery. In addition, display concepts employing EV imagery (compared to the SV and Baseline concepts which had none) were significantly better for pilot detection of intentional but unannounced navigation failures since this imagery provided an intuitive and obvious visual methodology to monitor the validity of the navigation solution.

The requirement for controlling lunar lander in the powered descending phase is analyzed in the paper. With suboptimal minimum fuel considered first, the expected non-iterative calculation formulas of the attitude angles are accquired. The explicit guidance project in relevance with the state variables and terminal restrict are deduced. According to deflection between the real attitude angles and the desired attitude angles, the attitude tracking system is designed by adopting the method of exponent reaching law and chattering reducing of boundary layers based on variable structure control with sliding mode. The simulation results show that the control system can complete flying mission in the powered descending phase, and has good robustness and adaptability as well.

The aim of Terrain Relative Navigation systems (TRN) is to augment inertial navigation by providing position estimates relative to known lunar surfaces. Also the purpose of TRN systems is to assist a lunar landing spacecraft with precise and safe landing. Such systems collect the height values from the lunar surface with the help of active range sensors which are then matched within a terrain Digital Elevation Map (DEM) database. Although lunar terrain elevation maps are rare and have low resolution quality, after Lunar Reconnaissance Orbit (LRO) mission is completed by 2010, maps of 20–25 m resolution will be gained. In this proposed work, a model of the lunar surface is developed with terrain generation algorithms. All the possible profiles (height values) and the slopes of the profiles that the sensor may collect during a position estimation phase are determined in advance. The DEM is pre-processed and reorganized into a Digital Profile Attributes Database (DPAD). DPAD records are organized in a way to optimize the lander's position estimation process. DPAD data is available as a function of lunar latitudes and longitudes. The DPAD over which the lunar lander is flying is loaded into the lander's memory. A novel TRN algorithm has been developed to estimate the precise location of the lander. This algorithm matches the determined slope values within the DPAD. Where only one match is found, the position of the lander is determined. When more-than-one matches arise, the system iterates until only one position solution is reached. In order to achieve accurate continuous navigation, the sensor should make measurements at specific intervals. The difference between the lander's estimate and the actual height that are measured by laser altimeter, can be used to calculate the position errors, thus providing the lander a continuous navigation solution as in TERCOM [7].

During the Apollo landings, onboard astronauts, along with analysis and instructions from mission control, performed the majority of complex tasks beyond automated Guidance, Navigation, and Control (GN&C). The crew played a significant role in the landings and were critical to navigating to the landing site, selecting a safe landing aim point, and commanding the spacecraft via a hand controller. Thus the requirements and constraints for site selection, launch dates, and GN&C design were driven in large part by human capabilities. Each of the major tasks performed by the crew will be described to provide an understanding of the functions that must be performed by either automation or people (on or off-board) for the next generation lunar lander. Additionally, several of the Apollo missions faced significant issues in identifying the landing site and assessing the safety of that landing site. These missions serve as case studies for future landing challenges that must be overcome.

Space exploration and mission success are dependent upon assessing mission viability. When a new mission and vehicle are defined, a reasonable facsimile must be built to conduct trade studies. In the case of the Altair lunar lander, we have developed a simulation lab with a specific hardware and software architecture to facilitate trade studies. Our starting point is to design and implement simplified guidance and navigation and control algorithms for the lunar landing trajectory. This trajectory gets integrated into an analysis of the communications system. We are using a software and hardware platform that allows for multiple teams and companies to integrate their simulation models into a single evaluation model. We discuss the software and hardware architecture which includes autocode generation targeted for a single board computer, and Ethernet connections used for integrating multiple labs. A discussion of how to incorporate flight hardware into the simulation environment is also presented.

The methods and experiences from the Apollo Program are fundamental building blocks for the development of lunar landing strategies for the Constellation Program. Each of the six lunar landing Apollo missions landed under near ideal lighting conditions. The astronauts visually performed terrain relative navigation while looking out of windows, and were greatly aided by external communication and well lit scenes. As the LM approached the landing site, the astronauts performed visual hazard detection and avoidance, also under near-ideal lighting conditions. The astronauts were looking out of the windows trying to the best of their ability to avoid rocks, slopes, and craters and find a safe landing location. NASA has expressed a desire for global lunar access for both crewed and robotic sortie lunar exploration missions [2] [3]. Early NASA architecture studies have identified the lunar poles as desirable locations for early lunar missions. These polar missions provide less than ideal lighting conditions that will significantly affect the way a crewed vehicle is to land at such locales. Consequently, a variety of hazard identification methods should be considered for use by the crew to ensure a high degree of safety. This paper discusses such identification methods applicable to the poorly lit polar lunar environment, better ensuring global access for the soon to be designed Lunar Lander Vehicle (LLV).

This paper describes and discusses the Constellation point of departure lunar capability architecture. Constellation is NASA's program to implement the human exploration of the moon and Mars. This paper focuses on the overall design and operational strategy for the Constellation lunar transportation system architecture including the Ares V heavy lift launch vehicle, Altair lunar lander and early concepts for lunar surface systems including habitats and mobile vehicles. The architecture builds on the initial capability vehicles of Ares I and Orion as they will initially fly to the international space station. The architecture is based on extensive trade studies of surface system objectives, lunar lander design options and Ares V design options. In addition to the description of the architecture this paper will briefly provide status of the initial capability building blocks of Ares I and Orion.

All Apollo landings were performed by the crew, manually commanding the lunar module (LM) attitude and rate-of-descent (ROD). In future missions, the astronauts will again need manual control of the flight path and attitude. Crew interaction mechanisms have been proposed to re-designate a landing aimpoint during the approach phase. However, manual control modes, crew control capability, and resulting vehicle performance during terminal descent have not been thoroughly investigated. A rate-control attitude hold (RCAH) mode was ultimately used in the LM for lateral flight and descent rate was controlled incrementally (P66), although other modes were considered and evaluated. These modes, as well as others proposed during Apollo, are reviewed and discussed in terms of their applicability to Altair. The ALHAT guidance, navigation, and control (GNC) algorithms for autonomous precision lunar landing were modified to include two manual flight control modes: RCAH with incremental ROD (A66) and incremental lateral velocity control with incremental ROD (A68) using Altair LDAC1-Delta vehicle parameters. Crew interactions with the ALHAT GNC system are described throughout the mission phases from lunar orbit to touchdown, focusing on manual control of flight path and attitude during terminal descent. These ALHAT manual attitude control modes are described and vehicle performance is discussed in terms of its estimated impact on handling quality ratings.

Execution of NASA's Vision for Space Exploration through Project Constellation involves the integration of multiple mission architecture and vehicle design teams. One of these is the Altair Lunar Lander Design Team, which had chosen to approach the vehicle development through a process of risk-informed design. Following the “de minimus” concept, a single string minimally functional vehicle that can perform all the mission requirements was created as a baseline design. This exercise, performed by the Altair Design Team, ensured the minimal mass vehicle would fit within the design constraints and established what mass margin would be available for enhancement of the design.