Results for “lander” — Page 12

NASA NTRS2024

Design, Development, and Use of a Lunar Lander Simulation for NASA's Artemis Program

This paper describes the design, development, and initial use of a generalized and configurable lunar lander simulation to support NASA’s Artemis Program. This simulation is being developed for the Crew Compartment Office (CrewCo) in the Human Landing Systems (HLS) program and is called the HLS CrewCo Lander Simulation (HCLS). The HCLS provides insight into the challenges associated with returning humans to the Moon and the particular difficulties of operating at the Lunar South Pole. A generalized lunar landing spacecraft based on a government reference design has been modeled but the simulation can be modified and adapted to model vendor designs as well. The simulation architecture and toolsets provide a flexible framework that allows for quickly prototyping and evaluating various aspects of a piloted lunar landing system. This includes the modeling of all principal human controlled flight phases: rendezvous and docking with crew transfer systems; docked orbital outpost operations; undocking and lunar transfer; lunar orbital operations; lunar deorbit, descent, and landing (DDL); lunar surface operations; lunar ascent; and return to the orbital outpost. The flexibility of the simulation allows for the integrated evaluation of potential vehicle subsystems, crew displays, guidance and control modes, and trajectory designs. The purpose of the HCLS is not to design the ideal lunar lander, but rather to understand the strengths and weaknesses of vehicle design choices.

Edwin Z Crues↗

NASA NTRS2024

Design, Development, and Use of a Lunar Lander Simulation for NASA’s Artemis Program

This paper describes the design, development, and initial use of a generalized and configurable lunar lander simulation to support NASA’s Artemis Program. This simulation is being developed for the Crew Compartment Office (CrewCo) in the Human Landing Systems (HLS) program and is called the HLS CrewCo Lander Simulation (HCLS). The HCLS provides insight into the challenges associated with returning humans to the Moon and the particular difficulties of operating at the Lunar South Pole. A generalized lunar landing spacecraft based on a government reference design has been modeled but the simulation can be modified and adapted to model vendor designs as well. The simulation architecture and toolsets provide a flexible framework that allows for quickly prototyping and evaluating various aspects of a piloted lunar landing system. This includes the modeling of all principal human controlled flight phases: rendezvous and docking with crew transfer systems; docked orbital outpost operations; undocking and lunar transfer; lunar orbital operations; lunar deorbit, descent, and landing (DDL); lunar surface operations; lunar ascent; and return to the orbital outpost. The flexibility of the simulation allows for the integrated evaluation of potential vehicle subsystems, crew displays, guidance and control modes, and trajectory designs. The purpose of the HCLS is not to design the ideal lunar lander, but rather to understand the strengths and weaknesses of vehicle design choices.

James Gentile↗

NASA NTRS2006

Dust Devil Tracks and Wind Streaks in the North Polar Region of Mars: A Study of the 2007 Phoenix Mars Lander Sites

The 65-72 latitude band of the North Polar Region of Mars, where the 2007 Phoenix Mars Lander will land, was studied using satellite images from the Mars Global Surveyor (MGS) Mars Orbiter Camera Narrow-Angle (MOC-NA) camera. Dust devil tracks (DDT) and wind streaks (WS) were observed and recorded as surface evidence for winds. No active dust devils (DDs) were observed. 162 MOC-NA images, 10.3% of total images, contained DDT/WS. Phoenix landing Region C (295-315W) had the highest concentration of images containing DDT/WS per number of available images (20.9%); Region D (130-150W) had the lowest (3.5%). DDT and WS direction were recorded for Phoenix landing regions A (110-130W), B (240-260W), and C to infer local wind direction. Region A showed dominant northwest-southeast DDT/WS, Region B showed dominant north-south, east-west and northeast-southwest DDT/WS, and region C showed dominant west/northwest - east/southeast DDT/ WS. Results indicate the 2007 Phoenix Lander has the highest probability of landing near DDT/WS in landing Region C. Based on DDT/WS linearity, we infer Phoenix would likely encounter directionally consistent background wind in any of the three regions.

Mars↗

NASA NTRS2005

Exploring Europa with a Surface Lander Powered by a Small Radioisotope Power System (RPS)

This paper describes a conceptual landed mission to the Jovian satellite Europa using a small RPS powered lander that would ride piggyback on the proposed Jupiter Icy Moons Orbiter (JIMO). This mission study was performed to assess the feasibility of landing a realistic science driven payload using a conceptual small radioisotope power system (US) to provide electrical and thermal power during the extended duration cruise phase (up to 13 years) and the nominal 30 day surface science mission. This paper includes individual sections that describe the key science goals, the mission architecture, and the conceptual design of the Europa Lander Mission (ELM) spacecraft.

lander↗

NASA NTRS2013

Jovian Tour Design for Orbiter and Lander Missions to Europa

Europa is one of the most interesting targets for solar system exploration, as its ocean of liquid water could harbor life. Following the recommendation of the Planetary Decadal Survey, NASA commissioned a study for a flyby mission, an orbiter mission, and a lander mission. This paper presents the moon tours for the lander and orbiter concepts. The total delta v and radiation dose would be reduced by exploiting multi-body dynamics and avoiding phasing loops in the Ganymede-to- Europa transfer. Tour 11-O3, 12-L1 and 12-L4 are presented in details and their performaces compared to other tours from previous Europa mission studies.

lander↗

NASA NTRS2020

Long-duration Venus lander for seismic and atmospheric science

An exciting and novel science mission concept called Seismic and Atmospheric Exploration of Venus (SAEVe) has been developed which uses high-temperature electronics to enable a three-order magnitude increase in expected surface life (120 Earth days) over what has been achieved to date. This enables study of long-term, variable phenomena such as the seismicity of Venus and near surface weather, near surface energy balance and atmospheric chemical composition. SAEVe also serves as a critical pathfinder for more sophisticated landers in the future. For example, first order seismic measurements by SAEVe will allow future missions to deliver better seismometers and systems to support the yet unknown frequency and magnitude of Venus events. SAEVe is focused on science that can be realized with low data volume instruments and will most benefit for temporal operations. The entire mission architecture and operations maximize science while minimizing energy usage and physical size and mass. The entire SAEVe system including its protective entry system is estimated to be around 45 kg and approximately 0.6 m diameter. These features allow SAEVe to be relatively cost effective and be easily integrated onto a Venus orbiter mission. The technologies needed to implement SAEVe are currently in development by several funded activities. Component and system level work is ongoing under NASA’s HOTTech program and by the Long Lived Insitu Solar System Explorer (LLISSE) project. The SAEVe long duration Venus lander promises groundbreaking science and is an ideal complimentary element to many future Venus orbiter missions being proposed or planned today.

Venus↗

NASA NTRS1999

Navigation Strategy for the Mars 2001 Lander Mission

The MSP 2001 project will send an orbiter, a lander, and a rover to Mars in the 2001 opportunity. The lander will demonstrate precision landing at Mars by utilizing improved approach navigation and hypersonic aeromaneuvering.

Navigation↗

NASA NTRS1996

Mars Pathfinder Lander Deployment Mechanisms

The Mars Pathfinder Lander will land, right itself, and open so that none of its airbag covers its solar cells. This requires two different types of mechanisms, the Airbag Retraction Actuators, and the Lander Petal Actuators which are designed for a high torque, low temperature, dirty environment, and limited life application. The development of these actuators investigated low temperature lubrication, Electrical Discharge Machining (EDM) to cut gears, and gear design for limited life use.

Mars↗

NASA NTRS1996

(abstract) Through the Europan Ice: Advanced Lander Mission Options

A variety of options for missions to visit Europa and study it in depth are being discussed. A first reconaissance may be made with an orbiter only. However, detailed examination of any water-ice ocean, particularily to determine the possible existence of biochemical materials, will require landing on and perhaps melting down through the ice itself to directly explore the Europan ocean. A first mission to Europa will attempt to determine whether or not this liquid water exists. Assuming this first mission determined it likely that liquid water did indeed exist, a more sophisticated lander package would be appropriate. This paper will describe strawman designs of this advanced lander as well and will discuss the many opportunities and challences in creating such spacecraft. Work in progress at both Leicester and JPL to further determine cryobot feasibility through numerical modeling and mechanical prototyping will be described.

Europa↗

NASA NTRSDate not supplied

Moving Beyond Apollo: Vacuum Ground Testing to Reduce Plume-Surface Interaction Risks to Lunar Landers

NASA’s Artemis Program will return humans to the surface of the Moon for the first time since Apollo using the Human Landing System (HLS). Plume-surface interactions (PSI) pose a potential hazard to all propulsive landing vehicles and future nearby assets that will be part of a sustained lunar architecture. Risks due to uncertainty in PSI predictions have challenged lunar landers since the 1960s, and understanding these phenomena further remains critical to enabling NASA’s lunar exploration goals. To this end, the HLS Program has funded a risk reduction ground test to obtain data relevant for application to environments produced by large landing systems. New data are needed to understand PSI and effects with the potential to differ from those experienced by the Apollo landers. This presentation will discuss the test concept, facility, research goals, methods, and planned data products.

Moon↗

NASA NTRSDate not supplied

Aerothermal Analysis and Thermal Protection System Design of the Mars Sample Retrieval Lander [SRL].

Mars Sample Retrieval Lander, part of the Mars Sample Return (MSR) mission, is being designed to land the heaviest payload yet, to the surface of Mars. SRL is being designed to carry the Lander, Sample Transfer System, Mars Acent Vehicle, and two Sample Recovery Helicopters. Compared to MSL and Mars 2020, SRL has a significantly higher ballistic coefficient, and flies at a higher lift/drag configuration. While the SRL heatshield is very similar to that of MSL and M2020, the backshell is very different, so as to accommode the payload. SRL is shielded by the same TPS materials as MSL and Mars 2020, with changes to design reflecting the SRL configuration and ConOPS. The aerothermal analysis and TPS design methodology of SRL relies on the successes of MSL and Mars 2020, and the lessons learned from MEDLI and MEDLI2. However, the constraints on mass require us to revisit all of our prediction models and analysis assumptions, in an attempt to reduce conservatism and TPS mass. MSL and Mars 2020 reconstruction, and detailed comparisons against MEDLI/MEDLI2 data are being used to justify our analysis approach and refine uncertainties and margins.

Mars↗

NASA NTRSDate not supplied

A Hybrid CFD/Engineering Model Tool for Lunar Lander Surface Erosion Prediction

Plume-Surface Interaction (PSI) between lander engine plumes and landing area regolith poses risks to the landers and space exploration missions through view obscuration and high-energy ejecta impacts from eroded particle clouds and vehicle tilting from plume-induced craters. To address these risks, NASA MSFC Fluid Dynamics Branch (ER42) has developed a cascade of predictive simulation capabilities for PSI, including the recent development of a hybrid CFD/engineering model tool that allows for rapid PSI simulations with a descending/ascending vehicle and an eroding surface. The development of this fully-coupled, moving body/eroding surface simulation capability is detailed along with a demonstration of the tool’s capabilities on the Apollo 12 Lunar Module landing. Predictions of eroded mass flow rates as a function of time are verified against Apollo 12 flight data, from which the developed, viscous erosion model has been calibrated.

Plume Surface Interaction↗

NASA NTRS2023

Hybrid CFD Engineering Model of Plume Induced Erosion and Crater Formation During Descent of Lunar Landers

With rapidly increased worldwide interest in landing on the moon, the issue of Plume Surface Interactions (PSI) is gaining attention. Hazards posed by lander plume induced dust and debris, as well as landing site deformation can be mitigated when better understood through predictive simulations. As simulation enabling computational power continues to increase, hybrid Computational Fluid Dynamics (CFD)/Engineering models provide the immediate ability to conduct parametric/trade studies driving design decisions for landers. Reduced order erosion engineering models apply correlations of the surface erosion rate to the plume induced surface forces with the correlations anchored to flight observations from Apollo LM landings. The MSFC propulsion fluid dynamics branch has developed such a hybrid model for predicting the plume induced viscous erosive regression of the Lunar surface beneath a landing vehicle. The (Descent Interpolated Gas Granular Erosion Model) DIGGEM was originally implemented as a post processing tool to calculate induced erosion rates through vehicle descent using CFD solutions of the vehicle plume at several fixed altitudes. This capability has since been advanced in the Loci/Chem-DIGGEM model to allow transient, moving vehicle, fully coupled viscous erosion modeling of vehicle descent PSI. This model has recently been used to make preflight predictions of the erosive regression of the ground beneath a Commercial Lunar Payload Services (CLPS) vehicle in support of measurements to be made by the Stereo Cameras for Lunar Plume Surface Studies (SCALPSS) instrument.

Plume Surface Interaction↗

NASA NTRS2023

Hybrid CFD Engineering Model of Plume Induced Erosion and Crater Formation During Descent of Lunar Landers

With rapidly increased worldwide interest in landing on the moon, the issue of Plume Surface Interactions (PSI) is gaining attention. Hazards posed by lander plume induced dust and debris, as well as landing site deformation can be mitigated when better understood through predictive simulations. As simulation enabling computational power continues to increase, hybrid Computational Fluid Dynamics (CFD)/Engineering models provide the immediate ability to conduct parametric/trade studies driving design decisions for landers. Reduced order erosion engineering models apply correlations of the surface erosion rate to the plume induced surface forces with the correlations anchored to flight observations from Apollo LM landings. The MSFC propulsion fluid dynamics branch has developed such a hybrid model for predicting the plume induced viscous erosive regression of the Lunar surface beneath a landing vehicle. The (Descent Interpolated Gas Granular Erosion Model) DIGGEM was originally implemented as a post processing tool to calculate induced erosion rates through vehicle descent using CFD solutions of the vehicle plume at several fixed altitudes. This capability has since been advanced in the Loci/Chem-DIGGEM model to allow transient, moving vehicle, fully coupled viscous erosion modeling of vehicle descent PSI. This model has recently been used to make preflight predictions of the erosive regression of the ground beneath a Commercial Lunar Payload Services (CLPS) vehicle in support of measurements to be made by the Stereo Cameras for Lunar Plume Surface Studies (SCALPSS) instrument.

Plume Surface Interaction↗

NASA NTRSDate not supplied

Capabilities for Long-Duration Landers in Extreme Environments

Missions to the surface of planets that experience high temperatures, like Venus or Mercury, have had limited consideration and development in recent decades partially because of the extreme temperatures and environments the planets experience. In the case of Mercury this is up to 430C for nearly 30 days and for Venus almost 60 days at 460C. Several landers have been sent to Venus many decades ago but 127 minutes was the longest any operated on the surface. Venus, and Mercury, hold many mysteries and successful surface missions will result in compelling new science that will have significant bearing on us here on Earth. To enable this compelling new science, NASA has been developing capabilities for a small lander that is designed to operate for months in the extreme temperatures found on Venus and Mercury. The capabilities promise to enable new missions not yet considered. This work summarizes technical advances that are preparing us for long-duration (months) operations in extreme environments on other planets.

Extreme Environment Lander↗

NASA NTRSDate not supplied

A Hybrid CFD/Engineering Model Tool for Lunar Lander Surface Erosion Prediction

Plume-Surface Interaction (PSI) between lander engine plumes and landing area regolith poses risks to the landers and space exploration missions through view obscuration and high-energy ejecta impacts from eroded particle clouds and vehicle tilting from plume-induced craters. To address these risks, NASA MSFC Fluid Dynamics Branch (ER42) has developed a cascade of predictive simulation capabilities for PSI, including the recent development of a hybrid CFD/engineering model tool that allows for rapid PSI simulations with a descending/ascending vehicle and an eroding surface. The development of this fully-coupled, moving body/eroding surface simulation capability is detailed along with a demonstration of the tool’s capabilities on the Apollo 12 Lunar Module landing. Predictions of eroded mass flow rates as a function of time are verified against Apollo 12 flight data, from which the developed, viscous erosion model has been calibrated.

Plume Surface Interaction↗

NASA NTRSDate not supplied

Results from the Radio Frequency Mass Gauge Technology Demonstration on the Intuitive Machines Nova-C Lunar Lander

A cryogenic propellant mass gauge known as the Radio Frequency Mass Gauge (RFMG) was integrated into the Intuitive Machines (IM) Nova-C lunar lander and provided an estimate of the liquid oxygen and liquid methane mass in the lander propellant tanks throughout the IM-1 mission, including during microgravity coast phases. An RFMG electronics controller was used to measure and record the spectrum of the RF signal reflected from an antenna sensor in each tank over the frequency range 100 to 1,300 MHz. The RF spectrum of each of the tanks is unique and is sensitive to the index of refraction of the propellants and the spatial distribution of the liquid within the tanks. Electromagnetic simulation software was used to simulate the antenna response spectra for a given tank geometry, fluid properties, and liquid–vapor configurations within the tank. Over 10,000 antenna response simulations were completed for each propellant tank prior to the IM-1 mission and represented various volumetric fill levels and fluid configurations. The simulated spectra served as a database against which measured tank spectra were compared. For analysis, a spectral matching algorithm was used to find the best match between measured and simulated spectra, and the gauged mass was calculated from the most highly correlated fluid mass simulations. RFMG measurements were recorded during tank loading on the launch pad and during translunar coast, lunar orbit insertion, low lunar orbit, powered descent to the lunar surface, and postlanding on the Moon. This paper describes the RF and fluid simulations, the RFMG measurements and analysis of spectral data, the RFMG instrument, and the gauged results throughout all phases of the IM-1 mission.

lunar lander↗

NASA NTRSDate not supplied

Results from the Radio Frequency Mass Gauge Technology Demonstration on the Intuitive Machines Nova-C Lunar Lander

A cryogenic propellant mass gauge known as the Radio Frequency Mass Gauge (RFMG) was integrated into the Intuitive Machines (IM) Nova-C lunar lander and provided an estimate of the liquid oxygen and liquid methane mass in the lander propellant tanks throughout the IM-1 mission, including during microgravity coast phases. An RFMG electronics controller was used to measure and record the spectrum of the RF signal reflected from an antenna sensor in each tank over the frequency range 100 to 1,300 MHz. The RF spectrum of each of the tanks is unique and is sensitive to the index of refraction of the propellants and the spatial distribution of the liquid within the tanks. Electromagnetic simulation software was used to simulate the antenna response spectra for a given tank geometry, fluid properties, and liquid–vapor configurations within the tank. Over 10,000 antenna response simulations were completed for each propellant tank prior to the IM-1 mission and represented various volumetric fill levels and fluid configurations. The simulated spectra served as a database against which measured tank spectra were compared. For analysis, a spectral matching algorithm was used to find the best match between measured and simulated spectra, and the gauged mass was calculated from the most highly correlated fluid mass simulations. RFMG measurements were recorded during tank loading on the launch pad and during translunar coast, lunar orbit insertion, low lunar orbit, powered descent to the lunar surface, and postlanding on the Moon. This paper describes the RF and fluid simulations, the RFMG measurements and analysis of spectral data, the RFMG instrument, and the gauged results throughout all phases of the IM-1 mission.

lunar lander↗