SEARCH · Engineering Papers

Results for “Mars thermal control”

Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

SourceResearch connectionFrom yearThrough year

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 325 records · Page 18

All sources NASA NTRS DOE OSTI

NASA NTRS1989

Mars Observer spacecraft

The technical aspects of the spacecraft for the Mars Observer mission are discussed. The spacecraft development focuses on using existing flight subsystem designs and production techniques to offer a low-cost, reliable, production-type spacecraft. The scientific objectives of the mission and the scientific payloads of the spacecraft are considered. The spacecraft system and its performance are discussed. The subsystems are described in detail, including attitude and articulation control, electrical power supply, propulsion, structure, thermal control, command and data handling, telecommunications, mechanics, and flight software.

Potts, Dennis L.↗

NASA NTRSDate not supplied

Integrated Design Results for the MSR DAC-0.0 Mars Ascent Vehicle

The NASA Mars Sample Return (MSR) Campaign endeavors to return Martian regolith, rock, and atmospheric samples to Earth for scientific study. One of many significant challenges to overcome in the return of these samples lies in transporting them from the Martian surface to space. In order to surmount this challenge, the Campaign has conceptualized the need for a Mars Ascent Vehicle (MAV) to perform this function and deliver Martian samples to orbit. There, the samples will be ejected and captured by a separate spacecraft for return to Earth. Many concepts for a MAV have existed in the past, but it has not been until now that an integrated, detailed design solution has been developed and analyzed. Preliminary assessments of the initial architecture examined multiple methods of propulsion. The team ultimately determined that a Two Stage to Orbit (TSTO) solid propulsion vehicle would provide the most effective performance and be the most technologically ready to support this mission. Following the decision to adopt a TSTO solid propelled vehicle, the first official Design Analysis Cycle, DAC-0.0, was performed in Spring 2020 to formally advance the fidelity of the vehicle to a maturity level acceptable for NASA Key Decision Point A (KDP-A). This paper describes the resultant MAV design concept developed as part of the DAC-0.0 study by the NASA Marshall Space Flight Center (MSFC), in association with the NASA Jet Propulsion Laboratory (JPL). The TSTO vehicle features two solid rocket motors, one powering each stage. Their thrust vectors are controlled with Thrust Vector Control (TVC) systems consisting of independent electromechanical actuators acting on gimballed nozzles. The vehicle is designed to deliver up to 0.47kg of Martian samples to a Mars circular orbit of 343km at 27° inclination. Due to the unique environmental conditions that this vehicle is required to operate in, the subsystem design teams were compelled to develop creative and unorthodox designs to ensure a successful mission. The detailed design and analysis of these subsystems are discussed in this paper and include topics on the MAV Guidance, Navigation, and Control (GNC); structures and mechanisms; integrated vehicle thermal; avionics and flight software; a hydrazine-based Reaction Control System (RCS); aerosciences; and vehicle assembly, integration, and test (AI&T) considerations, among others. Following the conclusion of the MAV DAC-0.0, additional alternative architecture concepts were also studied to further reduce the mass of the overall system. The results of these studies will also be examined in this paper.

Darius Yaghoubi↗

NASA NTRS2003

CO2 Acquisition Membrane (CAM)

The objective of CAM is to develop, test, and analyze thin film membrane materials for separation and purification of carbon dioxide (CO2) from mixtures of gases, such as those found in the Martian atmosphere. The membranes are targeted toward In Situ Resource Utilization (ISRU) applications that will operate in extraterrestrial environments and support future unmanned and human space missions. A primary application is the Sabatier Electrolysis process that uses Mars atmosphere CO2 as raw material for producing water, oxygen, and methane for rocket fuel and habitat support. Other applications include use as an inlet filter to collect and concentrate Mars atmospheric argon and nitrogen gases for habitat pressurization, and to remove CO2 from breathing gases in Closed Environment Life Support Systems (CELSS). CAM membrane materials include crystalline faujasite (FAU) zeolite and rubbery polymers such as silicone rubber (PDMS) that have been shown in the literature and via molecular simulation to favor adsorption and permeation of CO2 over nitrogen and argon. Pure gas permeation tests using commercial PDMS membranes have shown that both CO2 permeance and the separation factor relative to other gases increase as the temperature decreases, and low (Delta)P(Sub CO2) favors higher separation factors. The ideal CO2/N2 separation factor increases from 7.5 to 17.5 as temperature decreases from 22 C to -30 C. For gas mixtures containing CO2, N2, and Ar, plasticization decreased the separation factors from 4.5 to 6 over the same temperature range. We currently synthesize and test our own Na(+) FAU zeolite membranes using standard formulations and secondary growth methods on porous alumina. Preliminary tests with a Na(+) FAU membrane at 22 C show a He/SF6 ideal separation factor of 62, exceeding the Knudsen diffusion selectivity by an order of magnitude. This shows that the membrane is relatively free from large defects and associated non-selective (viscous flow) transport mechanisms. The Membrane Test Facility (MTF) has been developed to measure membrane permeance over a wide range of temperature and pressure. The facility uses two volume compartments separated by the membrane that are instrumented to measure temperature, delta pressure across the membrane, and gas composition. A thermal shroud supports and encloses the membrane, and provides temperature control. Methods were developed to determine membrane permeance using the first order decay of the pressure difference between the sealed compartments, using the total pressure for pure gases, and partial pressure of each species in gas mixtures. The technique provides an end-to-end measurement of gas permeance that includes concentration polarization effects. Experiments have shown that in addition to membrane permeance properties, the geometry and design of associated structures play an important role in how membrane systems will function on Mars.

Mason, Larry W.↗

NASA NTRS1981

Surface penetrators for planetary exploration: Science rationale and development program

Work on penetrators for planetary exploration is summarized. In particular, potential missions, including those to Mars, Mercury, the Galilean satellites, comets, and asteroids are described. A baseline penetrator design for the Mars mission is included, as well as potential instruments and their status in development. Penetration tests in soft soil and basalt to study material eroded from the penetrator; changes in the structure, composition, and physical properties of the impacted soil; seismic coupling; and penetrator deflection caused by impacting rocks, are described. Results of subsystem studies and tests are given for design of entry decelerators, high-g components, thermal control, data acquisition, and umbilical cable deployment.

Murphy, J. P.↗

NASA NTRSDate not supplied

Development of a Microlith® Catalytic Oxidizer for Exploration Trace Contaminant Control

The state-of-the-art in long mission duration spacecraft trace contaminant control processes rely on high-temperature catalytic oxidation for light contaminant removal. Target compounds for oxidation include metabolic methane, carbon monoxide, and low molecular weight volatile organics such as formaldehyde and methanol. Precision Combustion, Inc. and NASA Marshall Space Flight Center have been developing and testing prototype high temperature catalytic oxidizers based on PCI’s patented Microlith® technology to meet the requirements of future spaceflight exploration missions. To this end, our latest generation Microlith oxidizer, featuring an integrated heat recuperator, was subjected to endurance testing to simulate catalyst ageing over a Mars transit mission duration. Periodic reactor health testing indicates our approach results in a robust contaminant control solution for exploration missions beyond low earth orbit. Based on the demonstrated performance of this unit, a next-generation prototype was designed to meet exploration contaminant load control demands while upgrading the prototype form and fit to match flight-compatible interfaces. Prototype design considerations intended to reduce power consumption, impacting process thermal and hydraulic performance, are discussed herein.

Trace Contaminant Control↗

NASA NTRS2004

Miniature Loop Heat Pipe (MLHP) Thermal Management System

The MLHP Thermal Management System consists of a loop heat pipe (LHP) with multiple evaporators and condensers, thermal electrical coolers, and deployable radiators coated with variable emittance coatings (VECs). All components are miniaturized. It retains all the performance characteristics of state-of-the-art LHPs and offers additional advantages to enhance the functionality, versatility, and reliability of the system, including flexible locations of instruments and radiators, a single interface temperature for multiple instruments, cooling the on instruments and warming the off instruments simultaneously, improving. start-up success, maintaining a constant LHP operating temperature over a wide range of instrument powers, effecting automatic thermal switching and thermal diode actions, and reducing supplemental heater powers. It can fully achieve low mass, low power and compactness necessary for future small spacecraft. Potential applications of the MLHP thermal technology for future missions include: 1) Magnetospheric Constellation; 2) Solar Sentinels; 3) Mars Science Laboratory; 4) Mars Scouts; 5) Mars Telecom Orbiter; 6) Space Interferometry Mission; 7) Laser Interferometer Space Antenna; 8) Jupiter Icy Moon Orbiter; 9) Terrestrial Planet Finder; 10) Single Aperture Far-Infrared Observatory, and 11) Exploration Missions. The MLHP Thermal Management System combines the operating features of a variable conductance heat pipe, a thermal switch, a thermal diode, and a state-of-the-art LHP into a single integrated thermal system. It offers many advantages over conventional thermal control techniques, and can be a technology enabler for future space missions. Successful flight validation will bring the benefits of MLHP technology to the small satellite arena and will have cross-cutting applications to both Space Science and Earth Science Enterprises.

Ku, Jentung↗

NASA NTRS1990

Technologies for aerobraking

Aerobraking is one of the largest contributors to making both lunar and Mars missions affordable. The use of aerobraking/aeroassist over all-propulsive approaches saves as much as 60 percent of the initial mass required in LEO, thus reducing the number and size of earth-to-orbit launch vehicles. Lunar transfer vehicles, which will be used to transport personnel and materials from LEO to lunar outposts, will aerobrake into earth's atmosphere at aproximately 11 km/sec on return from the lunar surface. Current plans for both manned and robotic missions to Mars use aerocapture during arrival at Mars and at earth return. At Mars, the entry velocities will range from about 6 to 9.5 km/sec, and at earth the return velocity will be about 12.5 to 14 km/sec. In order to successfully design aerobrakes to withstand the aerodynamic forces and heating associated with these entry velocities, as well as to make them efficient, several critical technologies must be developed. These are vehicle concepts and configurations, aerothermodynamics, thermal protection system materials, and guidance, navigation, and control systems. This paper describes the status of each of these technologies and outlines what must be accomplished in each area to meet the requirements of the Space Exploration Initiative.

Cooper, David M.↗

NASA NTRS2003

Thermal Analysis of Acicular Shaped Magnetite

We are in the process of developing a database on the thermal properties of well-characterized Martian analog materials in support of future Mars surface missions. The database contains the thermal behaviors of these analog materials under reduced and Earth ambient pressures. Magnetite in planetary materials has received considerable attention in recent years since the identification of fine grain single-domain magnetite in Martian meteorite ALH84001 and their possible link to past life on Mars (i.e., possible biominerals of magnetotactic bacteria). Because of its possible importance to Mars science, we report here the thermal properties of magnetite particles with acicular morphology, i.e., needle-shaped magnetite. Acicular shaped magnetite can be commercially produced from goethite (FeOOH) as the starting material via a H2 reduction process. However, instead of using this process or procedure, we report here on the thermal characterization of acicular magnetite formed under reducing conditions from well-characterized needle-shaped goethite at low temperature in controlled CO-CO2 1-bar atmosphere gas mixing furnaces.

Lauer, H. V., Jr.↗

NASA NTRS2006

NASA Tech Briefs, March 2006

Topics covered include: Medical Signal-Conditioning and Data-Interface System; Instruments for Reading Direct-Marked Data-Matrix Symbols; Processing EOS MLS Level-2 Data; Ground Processing of Data From the Mars Exploration Rovers; Estimating Total Electron Content Using 1,000+ GPS Receivers; NASA Solar Array Demonstrates Commercial Potential; Improved Control of Charging Voltage for Li-Ion Battery; Programmable Pulse-Position-Modulation Encoder; Wavelength-Agile External-Cavity Diode Laser for DWDM; Pattern-Recognition Processor Using Holographic Photopolymer; Submicrosecond Power-Switching Test Circuit; Three-Function Logic Gate Controlled by Analog Voltage; Integrated System for Autonomous Science; Montage Version 3.0; Utilizing AI in Temporal, Spatial, and Resource Scheduling; Satellite Image Mosaic Engine; Architecture for Control of the K9 Rover; HFGMC Enhancement of MAC/GMC; Automated Activation and Deactivation of a System Under Test; Cleaning Carbon Nanotubes by Use of Mild Oxygen Plasmas; Generating Aromatics From CO2 on Mars or Natural Gas on Earth; Attaching Thermocouples by Peening or Crimping; Heat Treatment of Friction-Stir-Welded 7050 Aluminum Plates; Generating Breathable Air Through Dissociation of N2O; High-Performance Scanning Acousto-Ultrasonic System; Correction for Thermal EMFs in Thermocouple Feedthroughs; Using Quasiparticle Poisoning To Detect Photons; Estimating Resolution Lengths of Hybrid Turbulence Models; Education and Training Module in Alertness Management; Cargo-Positioning System for Next-Generation Spacecraft; Micro-Imagers for Spaceborne Cell-Growth Experiments; Holographic Solar Photon Thrusters; Plasma-Based Detector of Outer-Space Dust Particles; and Generation of Data-Rate Profiles of Ka-Band Deep-Space Links.

Source record↗

NASA NTRS1988

Mars rover technology development requirements

The technology development requirements for various Mars rover range capabilities are discussed, focusing on local navigation of the rover. The capabilities of two methods are compared. In one method, operators on the earth view stereo pictures sent by the rover and determine short traverse paths which the rover follows. The other method achieves more autonomous capability by using computer vision from orbital imagery with approximate long routes commanded from earth. The locomotion, navigation, ground operations, computation, power, thermal control, communications, sample acquisition, and analysis and preservation requirements are examined.

Bedard, Roger↗

NASA NTRS2015

Investigations of Methane Production in Hypersaline Environments

The recent reports of methane in the atmosphere of Mars, as well as the findings of hypersaline paleo-environments on that planet, have underscored the need to evaluate the importance of biological (as opposed to geological) trace gas production and consumption. Methane in the atmosphere of Mars may be an indication of life but might also be a consequence of geologic activity and/or the thermal alteration of ancient organic matter. Hypersaline environments have now been reported to be extremely likely in several locations in our solar system, including: Mars, Europa, and Enceladus. Modern hypersaline microbial mat communities, (thought to be analogous to those present on the early Earth at a period of time when Mars was experiencing very similar environmental conditions), have been shown to produce methane. However, very little is known about the physical and/or biological controls imposed upon the rates at which methane, and other important trace gases, are produced and consumed in these environments. We describe here the results of our investigations of methane production in hypersaline environments, including field sites in Chile, Baja California Mexico, California, USA and the United Arab Emirates. We have measured high concentrations of methane in bubbles of gas produced both in the sediments underlying microbial mats, as well as in areas not colonized by microbial mats in the Guerrero Negro hypersaline ecosystem, Baja California Mexico, in Chile, and in salt ponds on the San Francisco Bay. The carbon isotopic (δ13C) composition of the methane in the bubbles exhibited an extremely wide range of values, (ca. -75 per mille ca. -25 per mille). The hydrogen isotopic composition of the methane (δ2H) ranged from -60 to -30per mille and -450 to -350per mille. These isotopic values are outside of the range of values normally considered to be biogenic, however incubations of the sediments in contact with these gas bubbles reveals that the methane is indeed being produced by these sediments. Substrate limitation of methanogenesis in these environments, and not methane oxidation, would explain the isotopic values of the methane in these environments. Incubations with both isotopically labeled and unlabeled putative substrates for methanogenesis have shown that the substrates most important for methanogenesis in these environments are the so-called non-competitive substrates, e.g., methylamines, dimethylsulfide, and methanol. Acetate and bicarbonate appear not to be important substrates for methanogens in these environments. Extraction of DNA and analysis of a gene used for methane production (mcrA) has revealed that the community composition of methanogens is consistent with organisms known to use non-competitive substrates. Our work has shown that hypersaline environments have the potential to both produce and preserve methane for analysis, e.g., by capable rovers. Our work expends the range of methane isotopic values now known to be produced by active methanogenesis

hypersaline paleo-environments↗

DOE OSTI2021

Resistive wall mode stability and resonant field amplification in MAST high beta plasma

The n=1 (n is the toroidal mode number) resistive wall mode (RWM) instability and the resonant field amplification (RFA) due to a stable RWM response are numerically investigated for a MAST high-pressure plasma scenario, utilizing the MARS-F (Liu Y.Q. et al 2000 Phys. Plasmas 7 3681) and MARS-K (Liu Y.Q. et al 2008 Phys. Plasmas 15 112503) codes, for the purpose of understanding the potential RWM behavior in MAST-U plasmas. Ideal kink stability analysis identifies a target plasma, with parameters similar to that of the reference equilibrium reconstructed from a MAST high-pressure discharge, that accesses the RWM regime. The unstable n=1 RWM for the target equilibrium is subject to strong damping by the plasma toroidal flow and/or the drift kinetic effects from thermal particles. As a result, the mode is found to be stable under the experimental flow conditions, or even without flow stabilization if drift kinetic stabilization is included. The stability prediction is robust against variation of the assumed resistive wall minor radius. Active magneto-hydrodynamic (MHD) spectroscopy modeling, using the magnetic coils designed for controlling the edge localized modes (ELMs) in MAST-U as the antenna, shows strong resonant field amplification due to a stable RWM response in the target plasma. Maximal amplification, of 6-7 times larger than the applied vacuum field within the plasma, is obtained assuming a coil phasing of 120-180 degrees between the upper and lower rows of the ELM control coils. The MHD-kinetic hybrid model predicts generally higher RFA than the fluid model, in particular near the low-field and high-field sides of the plasma boundary surface. Furthermore, the MARS-F/K modeling thus shows that the RWM in MAST-U is subject to strong flow and kinetic stabilization, but can nevertheless be detected by active MHD spectroscopy.

70 PLASMA PHYSICS AND FUSION TECHNOLOGY↗

NASA NTRS2019

BioBot: Innovative Offloading of Astronauts for More Effective Exploration

The BioBot concept consists of a robotic rover which is capable of traversing the same terrain as a spacesuited human. It carries the primary life support system for the astronaut, including consumables, atmosphere revitalization systems (e.g., CO2 scrubbing, humidity and temperature management, ventilation fan), power system (e.g., battery, power management and distribution),and thermal control system (e.g., water sublimator, cooling water pump), along with umbilical lines to connect to the supported astronaut. Although not technically part of life support, it would be logical for the BioBot to also provide long-range communications, video monitoring, tool and sample transport, and other functions to enable and enhance EVA productivity in planetary surface exploration.The design reference scenario for this concept is that astronauts involved in future lunar or Mars exploration will be on the surface for weeks or months rather than days, and will be involved in regular EVA operations. It is not unreasonable to think of geologists spending several days inEVA exploration each week over a prolonged mission duration, with far more ambitious operational objectives than were typical of Apollo. In this scenario, each astronaut will be accompanied by a "BioBot", which will transport their life support system and consumables, an extended umbilical and umbilical reel, and robotic systems capable of controlling the position and motion of the umbilical. The astronaut will be connected to the robot via the umbilical, carrying only a small emergency open-loop life support system similar to those contained in every PLSS. The robotic mobility base will be designed to be capable of traveling anywhere the astronaut can walk, and will also be useful as a transport for the EVA tools, science instrumentation, and collected samples. In addition, the BioBot can potentially carry the astronaut on traverses as well. Such a system will also be a significant enhancement to public engagement in these future exploration missions, as the robotic vehicles can also support high-resolution cameras and high bandwidth communications gear to providehigh-definition video coverage of each crew throughout each EVA sortie.

Akin, David L.↗

NASA NTRS2004

Electronic packaging for extended Mars surface missions

Extended Mars missions will require vehicles to survive a large number of extended temperature cycles. To address this issue for electronics, previous strategies have been to locate critical electronics in a "warm electronics box" where thermal management is more easily maintained. However, that strategy limits the number and location of electronics. An alternative strategy allows electronics to be remotely located on actuator and wheel arms with no heating which has the advantage of distributed control.

Mojarradi, Mohammad M.↗

NASA NTRS2013

Mars Sample Return Landed with Red Dragon

A Mars Sample Return (MSR) mission is the highest priority science mission for the next decade as recommended by the recent Decadal Survey of Planetary Science. However, an affordable program to carry this out has not been defined. This paper describes a study that examined use of emerging commercial capabilities to land the sample return elements, with the goal of reducing mission cost. A team at NASA Ames examined the feasibility of the following scenario for MSR: A Falcon Heavy launcher injects a SpaceX Dragon crew capsule and trunk onto a Trans Mars Injection trajectory. The capsule is modified to carry all the hardware needed to return samples collected on Mars including a Mars Ascent Vehicle (MAV), an Earth Return Vehicle (ERV) and Sample Collection and Storage hardware. The Dragon descends to land on the surface of Mars using SuperSonic Retro Propulsion (SSRP) as described by Braun and Manning [IEEEAC paper 0076, 2005]. Samples are acquired and deliverd to the MAV by a prelanded asset, possibly the proposed 2020 rover. After samples are obtained and stored in the ERV, the MAV launches the sample-containing ERV from the surface of Mars. We examined cases where the ERV is delivered to either low Mars orbit (LMO), C3 = 0 (Mars escape), or an intermediate energy state. The ERV then provides the rest of the energy (delta V) required to perform trans-Earth injection (TEI), cruise, and insertion into a Moon-trailing Earth Orbit (MTEO). A later mission, possibly a crewed Dragon launched by a Falcon Heavy (not part of the current study) retrieves the sample container, packages the sample, and performs a controlled Earth re-entry to prevent Mars materials from accidentally contaminating Earth. The key analysis methods used in the study employed a set of parametric mass estimating relationships (MERs) and standard aerospace analysis software codes modified for the MAV class of launch vehicle to determine the range of performance parameters that produced converged spacecraft designs capable of meeting mission requirements. Subsystems modeled in this study included structures, power system, propulsion system, nose fairing, thermal insulation, actuation devices, and GN&C. Best practice application of loads and design margins for all resources were used. Both storable and cryogenic propellant systems were examined. The landed mass and lander capsule size provide boundary conditions for the MAV design and packaging. We estimated the maximum mass the Dragon capsule is capable of landing. This and the volume capability to store the MAV was deduced from publically available data from SpaceX as well as our own engineering and aerodynamic estimates. Minimum gross-liftoff mass (GLOM) for the MAV were obtained for configurations that used pump-fed storable bi-propellant rocket engines for both the MAV and the ERV stage. The GLOM required fits within our internal estimate of the mass that Dragon can land at low elevation/optimal seasons on Mars. Based on the analysis, we show that a single Mars launch sample return mission is feasible using current commercial capabilities to deliver the return spacecraft assets.

Mars Ascent Vehicle (MAV)↗

NASA NTRS2018

Development of High Power Hall Thruster Systems to Enable the NASA Exploration Vision

The next phase of space exploration missions requires high power Solar Electric Propulsion (SEP) systems for large-scale science missions and cargo transportation. Development is underway at Aerojet Rocketdyne on Hall thruster systems that are intended to bracket the needs of future NASA SEP missions in support of space exploration. The Advanced Electric Propulsion System (AEPS) program is developing and qualifying a 13.3kW Hall thruster system to be demonstrated on the Power and Propulsion Element (PPE), which is intended to be the first element of a Lunar Outpost Platform - Gateway (LOP-G). The NextSTEP program is integrating a nested Hall thruster into a 100kW system and testing it for 100 hours. These two programs will provide a path to efficient in-space propulsion that will allow NASA to transfer the large amounts of cargo that is needed to support human missions - first to the moon and then on to Mars. The Advanced Electric Propulsion System (AEPS) program is completing development, qualification and delivery of five flight 13.3kW EP systems to NASA. The flight AEPS system includes a magnetically shielded long-life Hall thruster, Power Processing Unit (PPU) and a Xenon Flow Controller (XFC). The Hall thruster, developed and demonstrated by NASA, operates at input powers up to 12.5kW while providing a specific impulse over an estimated 2800s at an input voltage of 600V. The power processor is designed to accommodate an input voltage range of 95-140V, consistent with operation beyond the orbit of Mars. The integrated system input power is continuously throttleable between 3 and 13.3kW. Component level testing of the EP String has begun with prototype hardware. The NextSTEP program is developing a 100kW Electric Propulsion (EP) system using a nested Hall thruster designed for powers up to 250kW, a modular power processor and a modular mass flow controller. While the program objective is to operate the integrated EP system continuously at 100kW for 100hrs to demonstrate thermal stability and support the development of system life time models, it builds on decades of experience with long-life Hall thrusters and the design is evolvable to a capability of 250kW. Design upgrades that demonstrate the 100kW EP system have been completed and tested. Aerojet Rocketdyne (AR) is excited to support NASA as it extends human reach into deep space and believes that these programs will provide the propulsion to make such missions affordable and sustainable. These systems provide NASA with a range of options to power its deep space transport vehicles. This paper presents the mission requirements for supporting the NASA exploration vision, as well as the status for the high power Hall thruster systems in development.

Jackson, Jerry↗

NASA NTRS2014

Mars Sample Return Using Commercial Capabilities: Mission Architecture Overview

Mars Sample Return (MSR) is the highest priority science mission for the next decade as recommended by the recent Decadal Survey of Planetary Science. This paper presents an overview of a feasibility study for an MSR mission. The objective of the study was to determine whether emerging commercial capabilities can be used to reduce the number of mission systems and launches required to return the samples, with the goal of reducing mission cost. We report the feasibility of a complete and closed MSR mission design using the following scenario that covers three synodic launch opportunities, beginning with the 2022 opportunity: A Falcon Heavy injects a SpaceX Red Dragon capsule and trunk onto a Trans Mars Injection (TMI) trajectory. The capsule is modified to carry all the hardware needed to return samples collected on Mars including a Mars Ascent Vehicle (MAV), an Earth Return Vehicle (ERV), and hardware to transfer a sample collected in a previously landed rover mission to the ERV. The Red Dragon descends to land on the surface of Mars using Super Sonic Retro Propulsion (SSRP). After previously collected samples are transferred to the ERV, the single-stage MAV launches the ERV from the surface of Mars. The MAV uses a storable liquid bi-propellant propulsion system to deliver the ERV to a Mars phasing orbit. After a brief phasing period, the ERV, which also uses a storable bi-propellant system, performs a Trans Earth Injection (TEI) burn. Upon arrival at Earth, the ERV performs Earth and lunar swing-bys and is placed into a lunar trailing circular orbit - an Earth orbit, at lunar distance. A later mission, using Dragon and launched by a Falcon Heavy, performs a rendezvous with the ERV in the lunar trailing orbit, retrieves the sample container and breaks the chain of contact with Mars by transferring the sample into a sterile and secure container. With the sample contained, the retrieving spacecraft makes a controlled Earth re-entry preventing any unintended release of pristine martian materials into the Earth's biosphere. The analysis methods employed standard and specialized aerospace engineering tools. Mission system elements were analyzed with either direct techniques or by using parametric mass estimating relationships (MERs). The architecture was iterated until overall mission convergence was achieved on at least one path. Subsystems analyzed in this study include support structures, power system, nose fairing, thermal insulation, actuation devices, MAV exhaust venting, and GN&C. Best practice application of loads, mass growth contingencies, and resource margins were used. For Falcon Heavy capabilities and Dragon subsystems we utilized publically available data from SpaceX, published analyses from other sources, as well as our own engineering and aerodynamic estimates. Earth Launch mass is under 11 mt, which is within the estimated capability of a Falcon Heavy, with margin. Total entry masses between 7 and 10 mt were considered with closure occurring between 9 and 10 mt. Propellant mass fractions for each major phase of the EDL - Entry, Terminal Descent, and Hazard Avoidance - have been derived. An assessment of the effect of the entry conditions on the thermal protection system (TPS), currently in use for Dragon missions, shows no significant stressors. A useful payload mass of 2.0 mt is provided and includes mass growth allowances for the MAV, the ERV, and mission unique equipment. We also report options for the MAV and ERV, including propulsion systems, crewed versus robotic retrieval mission, as well as direct Earth entry. International planetary protection policies as well as verifiable means of compliance will have a large impact on any MSR mission design. We identify areas within our architecture where such impacts occur. We also describe preliminary compliance measures that will be the subject of future work. This work shows that emerging commercial capabilities as well as new methodologies can be used to efficiently support an important planetary science objective. The work also has applications for human exploration missions that use propulsive EDL techniques

Red Dragon↗

NASA NTRS2011

Status of NASA In-Space Propulsion Technologies and Their Infusion Potential

Since 2001, the In-Space Propulsion Technology (ISPT) program has been developing in-space propulsion technologies that will enable or enhance NASA robotic science missions. These in-space propulsion technologies have broad applicability to future competed Discovery and New Frontiers mission solicitations, and are potentially enabling for future NASA flagship and sample return missions currently being considered. This paper provides status of the technology development of several in-space propulsion technologies that are ready for infusion into future missions. The technologies that are ready for flight infusion are: 1) the high-temperature Advanced Material Bipropellant Rocket (AMBR) engine providing higher performance; 2) NASA s Evolutionary Xenon Thruster (NEXT) ion propulsion system, a 0.6-7 kW throttle-able gridded ion system; and 3) Aerocapture technology development with investments in a family of thermal protection system (TPS) materials and structures; guidance, navigation, and control (GN&C) models of blunt-body rigid aeroshells; and aerothermal effect models. Two component technologies that will be ready for flight infusion in FY12/13 are 1) Advanced Xenon Flow Control System, and 2) ultra-lightweight propellant tank technology advancements and their infusion potential will be also discussed. The paper will also describe the ISPT project s future focus on propulsion for sample return missions: 1) Mars Ascent Vehicles (MAV); 2) multi-mission technologies for Earth Entry Vehicles (MMEEV) needed for sample return missions from many different destinations; and 3) electric propulsion for sample return and low cost missions. These technologies are more vehicle-focused, and present a different set of technology infusion challenges. Systems/Mission Analysis focused on developing tools and assessing the application of propulsion technologies to a wide variety of mission concepts.

Anderson, David↗