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All sourcesNASA NTRSDOE OSTI
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 100 kW 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 12.5 kW 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 100 hours 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 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 NTRS2019

Aeroassist Technologies for Small Satellite Missions

Orbit insertion operations that require large V maneuvers using conventional propulsive technologies are mass inefficient and challenging to package within SmallSat form factors such as the popular CubeSat. Aeroassist technologies offer an alternative approach for V maneuvers and could revolutionize the use of SmallSats for exploration missions and increase the science return while reducing costs for orbital or entry missions to Mars, Venus and return to Earth. Aeroassist refers to the use of an atmosphere to accomplish a transportation system function using techniques such as aerobraking, aerocapture, aeroentry, and aerogravity assist. Aeroassist technologies are power efficient and tolerant to the radiation and thermal environment encountered in deep space, and can be integrated around or within SmallSat geometries. This presentation will discuss various Aeroassist technologies including conventional rigid aeroshells, inflatable decelerators, mechanically deployable decelerators and other drag devices and control methods that should be considered by Small Satellite mission design teams.

Cassell, Alan M.↗
NASA NTRS1992

The effects of thermal gradients on the Mars Observer Camera primary mirror

The paper discusses the effect of thermal gradients on the optical performance of the primary mirror of Mars Observer Camera (MOC), which will be launched on the Mars Observer spacecraft in September 1992. It was found that mild temperature gradients can have a large effect on the mirror surface figure, even for relatively low coefficient-of-thermal-expansion materials. However, in the case of the MOC primary mirror, it was found that the radius of curvature (ROC) of the reflective surface of the mirror changed in a nearly linear fashion with the radial temperature gradient, with little additional aberration. A solid-state ROC controller using the thermal gradient effect was implemented and verified.

Applewhite, Roger W.↗
NASA NTRS2011

Performance of the Micropower Voltage Reference ADR3430 Under Extreme Temperatures

Electronic systems designed for use in space exploration systems are expected to be exposed to harsh temperatures. For example, operation at cryogenic temperatures is anticipated in space missions such as polar craters of the moon (-223 C), James Webb Space Telescope (-236 C), Mars (-140 C), Europa (-223 C), Titan (-178 C), and other deep space probes away from the sun. Similarly, rovers and landers on the lunar surface, and deep space probes intended for the exploration of Venus are expected to encounter high temperature extremes. Electronics capable of operation under extreme temperatures would not only meet the requirements of future spacebased systems, but would also contribute to enhancing efficiency and improving reliability of these systems through the elimination of the thermal control elements that present electronics need for proper operation under the harsh environment of space. In this work, the performance of a micropower, high accuracy voltage reference was evaluated over a wide temperature range. The Analog Devices ADR3430 chip uses a patented voltage reference architecture to achieve high accuracy, low temperature coefficient, and low noise in a CMOS process [1]. The device combines two voltages of opposite temperature coefficients to create an output voltage that is almost independent of ambient temperature. It is rated for the industrial temperature range of -40 C to +125 C, and is ideal for use in low power precision data acquisition systems and in battery-powered devices. Table 1 shows some of the manufacturer s device specifications.

Patterson, Richard L.↗
NASA NTRS2019

Mars Ascent Vehicle Hybrid Propulsion Development

Hybrid propulsion is being investigated as a propulsion method for a possible Mars Ascent Vehicle (MAV) application. MAV is part of a proposed larger Mars Sample Return (MSR) campaign plan to bring samples from Mars to earth for examination. The Mars Ascent Vehicle would launch Mars surface samples found and packaged by the Mars 2020 mission to orbit around Mars. This version of hybrid propulsion is based on a wax based solid fuel, called SP7A, and a Mixed Oxides of Nitrogen oxidizer, MON-25. SP7 is a new fuel formulation developed by Space Propulsion Group and was modified for this application to be resistant to Mars temperature extremes and modified again to lower the regression rate to become SP7A. MON-25 was chosen for its low freezing temperature. Due to cost constraints, MON-3 was the oxidizer used during testing through 2018. In 2019, full scale hybrid testing with MON-25 commenced in Mojave, CA by Whittinghill Aerospace. One flight motor will be subjected to thermal cycling in a vacuum and later fired in a vacuum to demonstrate the proposed Liquid Injection Thrust Vector Control system performance at White Sands Test Facility (WSTF). In addition, there will be MON-25 characterization work done at Purdue University and WSTF. Additional testing of subscale and full scale motors will be conducted with MON-3 with fuel grain stress, fuel grain support and case design test objectives by Space Propulsion Group Inc. of Butte, MT. This paper documents some of the testing, issues and accomplishments with the MAV hybrid propulsion option that is being considered (along with a two-stage solid propulsion option).

Story, G. T.↗
NASA NTRS2011

The Development of a CO2 Test Capability in the NASA JSC ARCJet for Mars Reentry Simulation

The Atmospheric Reentry Materials and Structures Evaluation Facility (ARMSEF) located at NASA Johnson Space Center is used for simulating the extreme environment experienced upon reentry for the development and certification of thermal protection systems (TPS). The facility supports a large variety of programs and was heavily leveraged for the certification and operational support of the TPS for the Orbiter and, more recently, the development of the heat shield for CEV. This paper will provide more detail into the heritage of the facility. Unique attributes of the facility include a modular aerodynamically stabilized arc heater and independently controlled O2 and N2 for the test gases. When combining the O2 and N2 in a 23:77 mass ratio respectively the Earth s atmosphere is accurately simulated and via modification of this ratio the investigation of the effects of atomic oxygen on a material s response is possible. In the summer of 2010 a development effort was started to add CO2 as a third independently controlled test gas such that, when combined with N2, opens up the possibility of accurately simulating a Martian reentry environment. This paper will discuss the test facility, especially the arc heater, in more detail. Initial testing involved relatively low concentrations of CO2 combined with N2 for the primary purpose of gathering data to answer two pressing safety concerns. The first being the rate of production of carbon monoxide (CO) within the ejector vacuum system. The main concern was that CO can be flammable and possibly explosive at high enough concentrations and pressures. The hazard control during the development phase involved the use of injecting N2 inside the test chamber diffuser to dilute and reduce the concentration of any and all CO present. A residual gas analyzer (RGA) was used to determine the relative amount of CO in the exhaust gas and provide a conversion rate of CO2 to CO. This paper will discuss in more detail the results of the RGA data and the calculated conversion rate. The second safety concern addressed is the possible formation of hydrogen cyanide (HCN) and cyanide (CN). HCN would primarily be present in the cooling water while the CN would most probably condense onto the interior surfaces of the test chamber. Water samples and wipes of the test chamber surfaces were analyzed by an industrial hygienist for the presence of HCN and CN. His paper will discuss these results in more detail. Throughout this development effort measurements of the CO2:N2 flowfield were made with heat flux and pressure probes and with laser induced fluorescence (LIF) of the atomic oxygen. This paper will discuss these results

DelPapa, Steven V.↗
NASA NTRS1983

New approaches to planetary exploration - Spacecraft and information systems design

Approaches are recommended for use by the NASA Solar System Exploration Committee (SSEC) in lowering the costs of planetary missions. The inclusion of off-the-shelf hardware, i.e., configurations currently in use for earth orbits and constructed on a nearly assembly-line basis, is suggested. Alterations would be necessary for the thermal control, power supply, telecommunications equipment, and attitude sensing in order to be serviceable as a planetary observer spacecraft. New technology can be developed only when cost reduction for the entire mission would be realized. The employment of lower-cost boost motors, or even integrated boost motors, for the transfer out of earth orbit is indicated, as is the development of instruments that do not redundantly gather the same data as previous planetary missions. Missions under consideration include a Mars geoscience climatology Orbiter, a lunar geoscience Orbiter, a near-earth asteroid rendezvous, a Mars aeronomy Orbiter, and a Venus atmospheric probe.

Diaz, A. V.↗
NASA NTRS2020

Mars Ascent Vehicle Solid Propulsion Configuration

As part of a Mars Sample Return (MSR) campaign, two Mars Ascent Vehicle (MAV) configurations have been designed in parallel. Each ascent vehicle configuration has a different propulsion system, which ultimately leads to two unique vehicle designs. As part of a Preliminary Architecture Assessment (PAA), these vehicle designs were developed to the same level of maturity in order to inform the selection of one of the vehicles as the point of departure design for the campaign. The selection will be made in November 2019. Although the initial MSR architecture called for a hybrid-based propulsion MAV featuring solid wax fuel with liquid oxidizer, a configuration using more traditional solid propulsion was developed as an additional risk mitigation option. Though lacking in the single stage to orbit (SSTO) and throttle flexibility of a hybrid configuration, a solid configuration vehicle allows a simpler design with significantly longer flight heritage and higher Technology Readiness Level (TRL). This paper describes the design of the solid propulsion configuration. An additional paper will be published describing the design of the hybrid propulsion configuration. The solid propulsion configuration MAV was developed in 2019 by NASA Marshall Space Flight Center (MSFC) in association with NASA Jet Propulsion Laboratory (JPL). It features two stages with a modified STAR-17 motor for the second stage and a traditional electromechanical actuator Thrust Vector Controller (TVC). The vehicle was designed to deliver approximately 0.47kg of Martian geological samples to a circular orbit at Mars of 343km at a 25° inclination. Although solid motor designs in general are at a relatively high TRL, the integrated vehicle subsystems that operate in conjunction with these propulsion elements do not typically operate in a Martian environment, which in this application can get as cold as -40°C. The PAA advanced the maturity of these subsystems by performing detailed design and analysis on the vehicle with respect to structures and mechanisms, Guidance/Navigation/Control (GNC) systems, avionics, Reaction Control System (RCS), TVC, thermal environments, and advanced Computational Fluid Dynamics (CFD). This paper will summarize the results of these studies.

Yaghoubi, Darius↗
NASA NTRS2020

Mars Ascent Vehicle Solid Propulsion Configuration

As part of a Mars Sample Return (MSR) campaign, two Mars Ascent Vehicle (MAV) configurations have been designed in parallel. Each ascent vehicle configuration has a different propulsion system, which ultimately leads to two unique vehicle designs. As part of a Preliminary Architecture Assessment (PAA), these vehicle designs were developed to the same level of maturity in order to inform the selection of one of the vehicles as the point of departure design for the campaign. The selection will be made in November 2019. Although the initial MSR architecture called for a hybrid-based propulsion MAV featuring solid wax fuel with liquid oxidizer, a configuration using more traditional solid propulsion was developed as an additional risk mitigation option. Though lacking in the single stage to orbit (SSTO) and throttle flexibility of a hybrid configuration, a solid configuration vehicle allows a simpler design with significantly longer flight heritage and higher Technology Readiness Level (TRL). This paper describes the design of the solid propulsion configuration. An additional paper will be published describing the design of the hybrid propulsion configuration. The solid propulsion configuration MAV was developed in 2019 by NASA Marshall Space Flight Center (MSFC) in association with NASA Jet Propulsion Laboratory (JPL). It features two stages with a modified STAR-17 motor for the second stage and a traditional electromechanical actuator Thrust Vector Controller (TVC). The vehicle was designed to deliver approximately 0.47kg of Martian geological samples to a circular orbit at Mars of 343km at a 25° inclination. Although solid motor designs in general are at a relatively high TRL, the integrated vehicle subsystems that operate in conjunction with these propulsion elements do not typically operate in a Martian environment, which in this application can get as cold as -40°C. The PAA advanced the maturity of these subsystems by performing detailed design and analysis on the vehicle with respect to structures and mechanisms, Guidance/Navigation/Control (GNC) systems, avionics, Reaction Control System (RCS), TVC, thermal environments, and advanced Computational Fluid Dynamics (CFD). This paper will summarize the results of these studies.

Yaghoubi, Darius↗
NASA NTRS2021

Low Cost Access to Mars Surface using a Small Impact Lander

In order to reduce the cost of landing small payloads on Mars, a new technology is being developed: the Small High Impact Energy Landing Device concept (SHIELD). The purpose of SHIELD is to provide a low-cost option to deliver up to 6 kg of science payload to the surface of Mars. SHIELD could be launched as a hosted payload, rideshare on an EELV secondary payload adapter (ESPA), or launching from a dedicated small launch vehicle using a kick stage. The SHIELD concept can achieve low-cost access to Mars by taking advantage of a low ballistic coefficient design to decelerate safely instead of using the typically required parachutes and/or propulsion to decelerate and stabilize itself during entry, decent, and landing (EDL). Current finite element simulations suggest that SHIELD would experience an impact acceleration pulse ranging from 1000 to 2000 g’s with a duration of 8 milliseconds. The SHIELD concept payload subsystem includes a “ruggedized” small warm electronics box (WEB). The WEB houses the telecommunications, command and data handling, thermal control, electrical power, and payload subsystems while maintaining an internal operating temperature ranging between -20° to 20° C. The system is designed to survive the Martian night by utilizing electric heaters powered with solar cells during the day and secondary batteries. The WEB is designed to be impact resistant capable of surviving an impact acceleration pulse equal to or less than 2000 g’s.

Woolley, Ryan↗
NASA NTRS2024

Lessons Learned in Designing A Proposed Ultraviolet Sterilization System for Space

This paper presents a number of lessons learned while designing a proposed sterilization system for Mars Sample Return. This sterilization system is needed to inactivate any potentially hazardous Mars material on the exterior surface of the vessel containing sealed sample tubes filled with Mars rock cores, regolith and atmosphere. These returned samples would provide information on the geologic history of Mars, the evolution of its climate and the potential for ancient life. Mars Sample Return is categorized at Planetary Protection Category V Restricted Earth Return, so it is required to protect the Earth–Moon system from the biological impact of returning samples from Mars to Earth. This article reviews lessons learned in the development of a particular engineering implementation to support the protection of the Earth–Moon biosphere: the use of in situ ultraviolet LED illumination. The details of the biological efficacy of this approach or the policy-related impacts are outside of the scope of this manuscript. The lessons learned presented here include establishing design requirements for the system, the selection of a light source, optical design options, contamination control and approaches to thermal and power management.

Planetary Protection↗
NASA NTRS2013

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 a 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. The major element required for the MSR mission are described and include an integration of the emerging commercial capabilities with small spacecraft design techniques; new utilizations of traditional aerospace technologies; and recent technological developments. 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 Supersonic Retro Propulsion (SRP). After previously collected samples are transferred to the ERV, the single-stage MAV launches the ERV from the surface of Mars to a Mars phasing orbit. The MAV uses a storable liquid, pump fed bi-propellant propulsion system. After a brief phasing period, the ERV, which also uses a storable bi-propellant system, performs a Trans Earth Injection (TEI) burn. Once near Earth the ERV performs Earth and lunar swing-bys and is placed into a Lunar Trailing Orbit (LTO0 - an Earth orbit, at lunar distance. A later mission, using a 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. Other capsule type vehicles and associated launchers may be applicable. 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 entry conditions on the thermal protection system (TPS), currently in use for Dragon missions, has been made. And shows no significant stressors. A useful mass of 2.0 mt is provided and includes mass growth allowances for the MAV, the ERV, and mission unique equipment. We also report on alternate propellant options for the MAV and options for the 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. This work shows that emerging commercial capabilities can be used to effectively integrated into a mission to achieve an important planetary science objective.

Mars Sample Return↗
NASA NTRS2014

Independent Verification of Mars-GRAM 2010 with Mars Climate Sounder Data

The Mars Global Reference Atmospheric Model (Mars-GRAM) is an engineering-level atmospheric model widely used for diverse mission and engineering applications. Applications of Mars-GRAM include systems design, performance analysis, and operations planning for aerobraking, entry, descent and landing, and aerocapture. Atmospheric influences on landing site selection and long-term mission conceptualization and development can also be addressed utilizing Mars-GRAM. Mars-GRAM's perturbation modeling capability is commonly used, in a Monte Carlo mode, to perform high-fidelity engineering end-to-end simulations for entry, descent, and landing. Mars-GRAM is an evolving software package resulting in improved accuracy and additional features. Mars-GRAM 2005 has been validated against Radio Science data, and both nadir and limb data from the Thermal Emission Spectrometer (TES). From the surface to 80 km altitude, Mars-GRAM is based on the NASA Ames Mars General Circulation Model (MGCM). Above 80 km, Mars-GRAM is based on the University of Michigan Mars Thermospheric General Circulation Model (MTGCM). The most recent release of Mars-GRAM 2010 includes an update to Fortran 90/95 and the addition of adjustment factors. These adjustment factors are applied to the input data from the MGCM and the MTGCM for the mapping year 0 user-controlled dust case. The adjustment factors are expressed as a function of height (z), latitude and areocentric solar longitude (Ls).

Justh, Hilary L.↗
NASA NTRS2011

Monitoring Mars for Electrostatic Disturbances

The DSN radio telescope DSS-13 was used to monitor Mars for electrostatic discharges from 17 February to 11 April, 2010, and from 19 April to 4 May, 2011, over a total of 72 sessions. Of these sessions, few showed noteworthy results and no outstanding electrostatic disturbances were observed on Mars from analyzing the kurtosis of radio emission from Mars. Electrostatic discharges on mars were originally detected in June of 2006 by Ruf et al. using DSS-13. he kurtosis (normalized fourth moment of the electrical field strength) is sensitive to non-thermal radiation. Two frequencies bands, either 2.4 and 8.4 GHz or 8.4 and 32 GHz were used. The non-thermal radiation spectrum should have peaks at the lowest three modes of the theoretical Schumann Resonances of Mars. The telescope was pointed away from Mars every 5 minutes for 45 seconds to confirm if Mars was indeed the sources of any events. It was shown that by including a down-link signal in one channel and by observing when the kurtosis changed as the telescope was pointed away from the source that the procedure can monitor Mars without the need of extra equipment monitoring a control source.

electrostatic discharges↗
NASA NTRS1990

Spectral albedo and emissivity of CO2 in Martian polar caps - Model results

In this paper, a snow albedo model previously developed for terrestrial snow is extended to the case of CO2 snow on Mars. Pure CO2 snow is calculated to have high albedo at visible wavelengths but not as high as that of water snow. At any given wavelength, the primary variable controlling albedo and emissivity is the snow grain size, with albedo decreasing and emissivity increasing as grain size increases. Observations that red albedo is much higher than blue albedo in the Martian south polar cap indicates that the snow or the atmosphere is contaminated with red dust. The absorption coefficient of CO2 ice in the thermal infrared is two to three orders of magnitude smaller than that measured for H2O ice. CO2 snow emissivity is therefore much lower than H2O snow, varying substantially with wavelength and quite sensitive to grain size and emission angle. Factors tending to increase emissivity are large grain size, small emission angle, and large concentrations of dust or water.

Warren, Stephen G.↗
NASA NTRS1994

A model for the evolution of CO2 on Mars

We have constructed a model that predicts the evolution of CO2 on Mars from the end of the heavy bombardment period to the present. The model draws on published estimates of the main processes believed to affect the fate of CO2 during this period: chemical weathering, regolith uptake, polar cap formation, and atmospheric escape. Except for escape, the rate at which these processes act is controlled by surface temperatures which we calculate using a modified version of the Gierasch and Toon energy balance model. The modifications account for the change in solar luminosity with time, the greenhouse effect, and a polar and equatorial energy budget. Using published estimates for the main parameters, we find no evolutionary scenario in which CO2 is capable of producing a warm (global mean temperatures greater than 250 K) and wet (surface pressures greater than 30 mbar) early climate, and then evolves to present conditions with approximately 7 mbar in the atmosphere, less than 300 mbar in the regolith, and less than 5 mbar in the caps. Such scenarios would only exist if the early sun were brighter than standard solar models suggest, if greenhouse gases other than CO2 were present in the early atmosphere, or if the polar albedo were significantly lower than 0.75. However, these scenarios generally require the storage of large amounts of CO2 (greater than 1 bar) in the carbonate reservoir. If the warm and wet early Mars constraint is relaxed, then we find best overall agreement with present day reservoirs for initial CO2 inventories of 0.5-1.0 bar. We also find that the polar caps can have a profound effect on how the system evolves. If the initial amount of CO2 is less than some critical value, then there is not enough heating of the poles to prevent permanent caps from forming. Once formed, these caps control how the system evolves, because they set the surface pressure and, hence, the thermal environment. If the initial amount of CO2 is greater than this critical value, then caps do not form initially, but can form later on, when weathering and escape lower the surface pressure to a point at which polar heating is no longer sufficient to prevent cap formation and the collapse of the climate system. Our modeling suggests this critical initial amount of CO2 is between 1 and 2 bar, but its true value will depend on all factors affecting the polar heat budget.

Haberle, Robert M.↗
NASA NTRS2008

25th Space Simulation Conference. Environmental Testing: The Earth-Space Connection

Topics covered include: Methods of Helium Injection and Removal for Heat Transfer Augmentation; The ESA Large Space Simulator Mechanical Ground Support Equipment for Spacecraft Testing; Temperature Stability and Control Requirements for Thermal Vacuum/Thermal Balance Testing of the Aquarius Radiometer; The Liquid Nitrogen System for Chamber A: A Change from Original Forced Flow Design to a Natural Flow (Thermo Siphon) System; Return to Mercury: A Comparison of Solar Simulation and Flight Data for the MESSENGER Spacecraft; Floating Pressure Conversion and Equipment Upgrades of Two 3.5kw, 20k, Helium Refrigerators; Affect of Air Leakage into a Thermal-Vacuum Chamber on Helium Refrigeration Heat Load; Special ISO Class 6 Cleanroom for the Lunar Reconnaissance Orbiter (LRO) Project; A State-of-the-Art Contamination Effects Research and Test Facility Martian Dust Simulator; Cleanroom Design Practices and Their Influence on Particle Counts; Extra Terrestrial Environmental Chamber Design; Contamination Sources Effects Analysis (CSEA) - A Tool to Balance Cost/Schedule While Managing Facility Availability; SES and Acoustics at GSFC; HST Super Lightweight Interchangeable Carrier (SLIC) Static Test; Virtual Shaker Testing: Simulation Technology Improves Vibration Test Performance; Estimating Shock Spectra: Extensions beyond GEVS; Structural Dynamic Analysis of a Spacecraft Multi-DOF Shaker Table; Direct Field Acoustic Testing; Manufacture of Cryoshroud Surfaces for Space Simulation Chambers; The New LOTIS Test Facility; Thermal Vacuum Control Systems Options for Test Facilities; Extremely High Vacuum Chamber for Low Outgassing Processing at NASA Goddard; Precision Cleaning - Path to Premier; The New Anechoic Shielded Chambers Designed for Space and Commercial Applications at LIT; Extraction of Thermal Performance Values from Samples in the Lunar Dust Adhesion Bell Jar; Thermal (Silicon Diode) Data Acquisition System; Aquarius's Instrument Science Data System (ISDS) Automated to Acquire, Process, Trend Data and Produce Radiometric System Assessment Reports; Exhaustive Thresholds and Resistance Checkpoints; Reconfigurable HIL Testing of Earth Satellites; FPGA Control System for the Automated Test of MicroShutters; Ongoing Capabilities and Developments of Re-Entry Plasma Ground Tests at EADS-ASTRIUM; Operationally Responsive Space Standard Bus Battery Thermal Balance Testing and Heat Dissipation Analysis; Galileo - The Serial-Production AIT Challenge; The Space Systems Environmental Test Facility Database (SSETFD), Website Development Status; Simulated Reentry Heating by Torching; Micro-Vibration Measurements on Thermally Loaded Multi-Layer Insulation Samples in Vacuum; High Temperature Life Testing of 80Ni-20Cr Wire in a Simulated Mars Atmosphere for the Sample Analysis at Mars (SAM) Instrument Suit Gas Processing System (GPS) Carbon Dioxide Scrubber; The Planning and Implementation of Test Facility Improvements; and Development of a Silicon Carbide Molecular Beam Nozzle for Simulation Planetary Flybys and Low-Earth Orbit.

Packard, Edward↗
NASA NTRS2020

Mars Ascent Vehicle Hybrid Propulsion Configuration

As part of a Mars Sample Return (MSR) campaign, two Mars Ascent Vehicle (MAV) configurations have been designed in parallel. Each ascent vehicle configuration has a different propulsion system which ultimately leads to two unique vehicle designs. As part of a Preliminary Architecture Assessment (PAA), these vehicle designs were developed to the same level of maturity in order to inform the selection of one of the vehicles as the point of departure design for the campaign. The selection will be made in November 2019. The initial MSR architecture called for a hybrid-based propulsion MAV. This type of propulsion system calls for a solid wax motor that would utilize liquid MON-25 as an oxidizer. Hybrid rocket propulsion allows for more flexibility than traditional solid or liquid propulsion options, and typically benefits from the advantages of both. A hybrid motor can be throttled and shut down easily, and avoids significant risk in manufacturing and handling. On a theoretical level, hybrid motors perform at a higher specific impulse (Isp) than solid motors. The primary disadvantage of hybrid motors comes from additional complexity and significantly less flight heritage and low Technology Readiness Level (TRL). This paper describes the design of the hybrid propulsion configuration. An additional paper will be published describing the design of the solid propulsion configuration1. The hybrid propulsion configuration MAV was developed in 2019 by NASA Marshall Space Flight Center (MSFC) in association with NASA Jet Propulsion Laboratory (JPL). It features a Single Stage to Orbit (SSTO) design with an SP7A solid wax fuel and MON-25 liquid oxidizer. The liquid portion of the vehicle allows for a Liquid Injection Thrust Vector Controller (LITVC) as well as hypergolic propellant additives for ignition. The vehicle was designed to deliver approximately 0.31kg of Martian geological samples to a circular orbit at Mars of 343km at a 25o inclination. Although hybrid propulsion in general has been used on launch vehicles in the past, the integrated vehicle subsystems that operate in conjunction with these propulsion elements do not typically operate in a Martian environment, which in this application can get as cold as -40oC. The PAA advanced the maturity of these subsystems by performing detailed design and analysis on the vehicle with respect to structures and mechanisms, Guidance/Navigation/Control (GNC) systems, avionics, Reaction Control System (RCS), LITVC, thermal environments, and advanced Computational Fluid Dynamics (CFD). This paper will summarize the results of these studies.

Yaghoubi, Darius↗