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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.

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NASA NTRS1999

Thermal and Mechanical Microspacecraft Technologies for Deep Space Systems Program X2000 Future Deliveries

Thermal and mechanical technologies are an important part of the Deep Space Systems Technology (DSST) Program X2000 Future Deliveries (FD) microspacecraft. A wide range of future space missions are expected to utilize the technologies and the architecture developed by DSST FD. These technologies, besides being small in physical size, make the tiny spacecraft robust and flexible. The DSST FD architecture is designed to be highly reliable and suitable for a wide range of missions such as planetary landers/orbiters/flybys, earth orbiters, cometary flybys/landers/sample returns, etc. Two of the key ideas used in the development of thermal and mechanical technologies and architectures are: 1) to include several of the thermal and mechanical functions in any given single spacecraft element and 2) the architecture be modular so that it can easily be adapted to any of the future missions. One of the thermal architectures being explored for the DSST FD microspacecraft is the integrated thermal energy management of the complete spacecraft using a fluid loop. The robustness and the simplicity of the loop and the flexibility with which it can be integrated in the spacecraft have made it attractive for applications to DSST FD. Some of the thermal technologies to be developed as a part of this architecture are passive and active cooling loops, electrically variable emittance surfaces, miniature thermal switches, and specific high density electronic cooling technologies. In the mechanical area, multifunction architecture for the structural elements will be developed. The multifunction aspect is expected to substantially reduce the mass and volume of the spacecraft. Some of the technologies that will be developed are composite material panels incorporating electronics, cabling, and thermal elements in them. The paper describes the current state of the technologies and progress to be made in the thermal and mechanical technologies and approaches for the DSST Future Deliveries microspacecraft.

Birur, Gajanana C.↗
NASA NTRS2015

Microelectrospray Thrusters

Propulsion technology is often a critical enabling technology for space missions. NASA is investing in technologies to enable high value missions with very small spacecraft, even CubeSats. However, these nanosatellites currently lack any appreciable propulsion capability. CubeSats are typically deployed and tumble or drift without any ability to transfer to higher value orbits, perform orbit maintenance, or perform de-orbit. Larger spacecraft can also benefit from high precision attitude control systems. Existing practices include reaction wheels with lifetime concerns and system level complexity. Microelectrospray thrusters will provide new propulsion capabilities to address these mission needs. Electric propulsion is an approach to accelerate propellant to very high exhaust velocities through the use of electrical power. Typical propulsion systems are limited to the combustion energy available in the chemical bonds of the fuel and then acceleration through a converging diverging nozzle. However, electric propulsion can accelerate propellant to ten times higher velocities and therefore increase momentum transfer efficiency, or essentially, increase the fuel economy. Fuel efficiency of thrusters is proportional to the exhaust velocity and referred to as specific impulse (Isp). The state-of-the-art (SOA) for CubeSats is cold gas propulsion with an Isp of 50-80 s. The Space Shuttle main engine demonstrated a specific impulse of 450 s. The target Isp for the Mars Exploration Program (MEP) systems is >1,500 s. This propellant efficiency can enable a 1-kg, 10-cm cube to transfer from low-Earth orbit to interplanetary space with only 200 g of propellant. In September 2013, NASA's Game Changing Development program competitively awarded three teams with contracts to develop MEP systems from Technology Readiness Level-3 (TRL-3), experimental concept, to TRL-5, system validation in a relevant environment. The project is planned for 18 months of system development. Due to the ambitious project goals, NASA has awarded contracts to mature three unique methods to achieve the desired goals. Some of the MEP concepts have been developed for more than a decade at the component level, but are now ready for system maturation. The three concepts include the high aspect ratio porous surface (HARPS) microthruster system, the scalable ion electrospray propulsion system (S-iEPS), and an indium microfluidic electrospray propulsion system. The HARPS system is under development by Busek Co. The HARPS thruster is an electrospray thruster that relies on surface emission of a porous metal with a passive capillary wicking system for propellant management. The HARPS thruster is expected to provide a simple, high ΔV and low-cost solution. The HARPS thruster concept is shown in figure 1. Figure 1 includes the thruster, integrated power processing unit, and propellant reservoir.

Dankanich, John↗
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 NTRS2017

Ground Demonstration on the Autonomous Docking of Two 3U CubeSats Using a Novel Permanent-Magnet Docking Mechanism

Small spacecraft autonomous rendezvous and docking is an essential technology for future space structure assembly missions. A novel magnetic capture and latching mechanism is analyzed that allows for docking of two CubeSats without precise sensors and actuators. The proposed magnetic docking hardware not only provides the means to latch the CubeSats but it also significantly increases the likelihood of successful docking in the presence of relative attitude and position errors. The simplicity of the design allows it to be implemented on many CubeSat rendezvous missions. A CubeSat 3-DOF ground demonstration effort is on-going at NASA Langley Research Center that enables hardware-in-the loop testing of the autonomous approach and docking of a follower CubeSat to an identical leader CubeSat. The test setup consists of a 3 meter by 4 meter granite table and two nearly frictionless air bearing systems that support the two CubeSats. Four cold-gas on-off thrusters are used to translate the follower towards the leader, while a single reaction wheel is used to control the attitude of each CubeSat. An innovative modified pseudo inverse control allocation scheme was developed to address interactions between control effectors. The docking procedure requires relatively high actuator precision, a novel minimal impulse bit mitigation algorithm was developed to minimize the undesirable deadzone effects of the thrusters. Simulation of the ground demonstration shows that the Guidance, Navigation, and Control system along with the docking subsystem leads to successful docking under 3-sigma dispersions for all key system parameters. Extensive simulation and ground testing will provide sufficient confidence that the proposed docking mechanism along with the choosen suite of sensors and actuators will perform successful docking in the space environment.

Pei, Jing↗
NASA NTRS1985

An overview of integrated flywheel technology for aerospace application

Space missions ranging from small scientific satellites to large manned spacecraft have, for many years, utilized systems of spinning flywheels to maintain vehicle attitude. These systems have included momentum and reaction wheels as well as control moment gyros. Extension of that technology to satisfy the additional tasks associated with energy storage has also been pursued. The combining of control and energy storage features into one system has been examined by NASA for space applications and demonstrated in the laboratory. The impact of technology advances in such areas as composite material rotors, magnetic suspensions, motor/generators, and electronics have prompted a re-evaluation of the viability of the flywheel storage system concept for aerospace applications. This paper summarizes the results of this re-examination and identifies shortfalls in the various technology areas.

Keckler, C. R.↗
NASA NTRS1994

Advanced Mico-Gn&C Technology for Low-Cost Planetary

The new NASA paradigm calls for more frequent, low-cost, small spacecraft missions capable of returning high-value planetary science. This challenge also extends to the rapid insertion of advanced technologies across all spacecraft subsystems as an enabling tool for building these highly capable miniaturized science platforms in the spirit of.

planetary spacecraft tracking gravity-field mappin↗
NASA NTRS2015

NASA Ames UV-LED Poster Overview

UV-LED is a small satellite technology demonstration payload being flown on the Saudisat-4 spacecraft that is demonstrating non-contacting charge control of an isolated or floating mass using new solid-state ultra-violet light emitting diodes (UV-LEDs). Integrated to the rest of the spacecraft and launched on a Dnepr in June 19, 2014, the project is a collaboration between the NASA Ames Research Center (ARC), Stanford University, and King Abdulaziz City for Science and Technology (KACST). Beginning with its commissioning in December, 2015, the data collected by UV-LED have validated a novel method of charge control that will improve the performance of drag-free spacecraft allowing for concurrent science collection during charge management operations as well as reduce the mass, power and volume required while increasing lifetime and reliability of a charge management subsystem. UV-LED continues to operate, exploring new concepts in non-contacting charge control and collecting data crucial to understanding the lifetime of ultra-violet light emitting diodes in space. These improvements are crucial to the success of ground breaking missions such as LISA and BBO, and demonstrates the ability of low cost small satellite missions to provide technological advances that far exceed mission costs.

KACST↗
NASA NTRS2010

Solar Sailing is not Science Fiction Anymore

Over 400 years ago Johannes Kepler envisioned the use of sunlight to propel a spacecraft. Just this year, a solar sail was deployed in orbit for the first time and proved that a spacecraft could effectively use a solar sail for propulsion. NASA's first nano-class solar sail satellite, NanoSail-D was designed and developed in only four months. Although the first unit was lost during the Falcon 1 rocket failure in 2008, the second flight unit has been refurbished and is waiting to be launched later this year. NanoSail-D will further the research into solar sail enabled spacecraft. It will be the first of several more sail enabled spacecraft to be launch in the next few years. FeatherSail is the next generation nano-class sail spacecraft being designed with the goal to prove low earth orbit operational capabilities. Future solar sail spacecraft will require novel ideas and innovative research for the continued development of space systems. One such pioneering idea is the Small Multipurpose Advanced Reconfigurable Technology (SMART) project. The SMART technology has the potential to revolutionize spacecraft avionics. Even though solar sailing is currently in its infancy, the next decade will provide great opportunities for research into sailing in outer space.

Alhorn, Dean C.↗
NASA NTRS1984

U.S. planetary exploration program technology implications

As a consequence of the widespread acceptance of the recommendations of the Solar System Exploration Committee, the U.S. Program for exploring the planets has entered a new phase. The objectives to be pursued involve a reduction of costs, while maintaining a high level of scientific return. Plans for the activities to be conducted in this new phase are related to a Core Program and to 'augmentation missions'. One part of the Core Program is concerned with the utilization of the technology, developed for earth-orbiting spacecraft, in missions within the inner solar system to targets ranging from Venus to the inner portion of the asteroid belt. However, modified earth-orbiting buses are not suitable for missions outside the inner solar system. For the second part of the Core Program, which is concerned with the outer solar system and small bodies, a modularized spacecraft based on Viking, Voyager, and Galileo technology will be developed. 'Augmentation missions' will be conducted when possible or desirable.

Diaz, A. V.↗
NASA NTRS2006

A Low Cost Spacecraft Architecture for Robotic Lunar Exploration Projects

A program of frequent, capable, but affordable lunar robotic missions prior to return of humans to the moon can contribute to the Vision for Space Exploration (VSE) NASA is tasked to execute. The Lunar Reconnaissance Orbiter (LRO) and its secondary payload are scheduled to orbit the moon, and impact it, respectively, in 2008. It is expected that the sequence of missions occurring for approximately the decade after 2008 will place an increasing emphasis on soft landed payloads. These missions are requited to explore intrinsic characteristics of the moon, such as hydrogen distribution in the regolith, and levitated dust, to demonstrate the ability to access and process in-situ resources, and to demonstrate functions critical to supporting human presence, such as automated precision navigation and landing. Additional factors governing the design of spacecraft to accomplish this diverse set of objectives are: operating within a relatively modest funding profile, the need tb visit multiple sites (both polar and equatorial) repeatedly, and to use the current generation of launch vehicles. In the US, this implies use of the Evolved Expendable Launch Vehicles, or EELVs, although this design philosophy may be extended to launch vehicles of other nations, as well. Many of these factors are seemingly inconsistent with each other. For example, the cost of a spacecraft usually increases with mass; therefore the desire to fly frequent, modestly priced spacecraft seems to imply small spacecraft (< 1 Mt, injected mass). On the other hand, the smallest of the EELVs will inject approx. 3 Mt. on a Trans Lunar Injection (TLI) trajectory md would therefore be wasteful or launching a single, small spacecraft. Increasing the technical capability of a spacecraft (such as autonomous navigation and soft landing) also usually increases cost. A strategy for spacecraft design that meets these conflicting requirements is presented. Taken together, spacecraft structure and propulsion subsystems constitute the majority of spacecraft mass; saving development and integration cost on these elements is critical to controlling cost. Therefore, a low cost, modular design for spacecraft structure and propulsion subsystems is presented which may be easily scaled up or down for either insertion into lunar orbit or braking for landing on the lunar surface. In order to effectively use the approx.3 Mt mass-to-TLI of the EELV, two low cost spacecraft will be manifested on the same launch. One spacecraft will be located on top of the other for launch and the two will have to be released in sequence in order to achieve all mission objectives. The two spacecraft could both be landers, both orbiters, or one lander and one orbiter. In order to achieve mass efficiency, the body of the spacecraft will serve the dual purposes of carrying launch loads and providing attachment points for all the spacecraft subsystems. In order to avoid unaffordable technology development costs, small liquid propulsion components and autonomous, scene-matching navigation cameras may be adapted from military missile programs in order to execute precision soft landings.

Lemke, Lawrence G.↗
NASA NTRS2024

DiskSat: A Two-Dimensional, High Power, High Aperture, Maneuverable Spacecraft

NASA is funding designers of small spacecraft at The Aerospace Corporation in El Segundo, California, to develop a technology demonstration of an evolutionary alternative to the CubeSat standard that maintains the benefits of that platform while overcoming key limitations. DiskSat is a plate-shaped satellite 40 inches in diameter and an inch thick that could offer more power and surface area for instruments, providing more opportunities for NASA to expand upon target mission objectives for small spacecraft.

Julianna L Fishman↗
NASA NTRS1999

Parallel Stirling Converters Being Developed for Spacecraft Onboard Power

Stirling Technology Co., as part of a NASA Lewis Research Center Phase II Small Business Innovation Research contract, has successfully demonstrated paralleling two thermodynamically independent Stirling converters. A system of four Stirling converters is being developed by NASA and the Department of Energy as an alternative high-efficiency radioisotope power source for spacecraft onboard electric power for NASA deep space missions. The high Stirling efficiency, exceeding 20 percent for this application, will greatly reduce the necessary isotope inventory in comparison to the current radioisotope thermoelectric generators (RTG s), significantly reducing mission cost and risk. Stirling is the most developed converter option of the advanced power technologies under consideration.

Thieme, Lanny G.↗
NASA NTRS2019

Using telemetry to navigate the MarCO cubesats to Mars

The two MarCO “cubesat” spacecraft were launched alongside NASA’s InSight in May 2018, operating primarily as a technology demonstrator for small satellites in deep space, with a nominal (but experimental) mission to provide relay support for the primary spacecraft during entry, descent, and landing at Mars. Due to their small size and experimental nature, extensive use of telemetry beyond that commonly used by deep space missions was necessary to complete adequate orbit determination. In particular, telemetry was valuable in two areas: use of wheel speeds during thruster calibrations to improve knowledge of individual thruster force levels, and the use of propellant temperature and pressure data to correctly model small thrusting events on board the vehicle.

Martin-Mur, Tomas↗
NASA NTRS2019

The Space Debris Sensor Experiment

The Space Debris Sensor (SDS) is a NASA Class 1E technology demonstration external payload aboard the International Space Station (ISS). With approximately one square meter of detection area, the SDS is attached to the European Space Agency Columbus module facing the ISS velocity vector with minimal obstruction from ISS hardware. The SDS is the first flight demonstration of the Debris Resistive/Acoustic Grid Orbital NASA-Navy Sensor (DRAGONS) technology developed and matured over 10 years by the NASA Orbital Debris Program Office (ODPO), in concert with the DRAGONS consortium, to provide information on the sub-millimeter scale orbital debris environment. The SDS demonstrated the capacity to read 4 resistive grids at 1 Hz, 40 acoustic sensors at 500 kHz, and record and downlink impact data to the ground. Observable and derived data from the SDS could provide information to models that are critical to understanding risks the small debris environment poses to spacecraft in low Earth orbit. The technology demonstrated by the SDS is a major step forward in monitoring and characterizing the space debris environment. This paper will address the technical performance of the SDS during its operational lifetime and its realization of technical and scientific goals. The SDS was intended to operate for 3 years; however, the payload incurred multiple anomalies during its operational life. Subsequently termed “Anomaly #1,” the first was the symptomatic loss of low data rate 1553 channel command and telemetry. The second, Anomaly #2, was loss of all low- and medium-data rate (Ethernet) telemetry. Anomaly #2 proved to be unrecoverable, leading to loss of the payload after approximately 26 days on-board the ISS. Therefore, this paper also addresses the anomalies that occurred during operation of the SDS, their attribution, and their resolution. Lessons learned are described when relevant to anomaly identification, attribution, and resolution.

Anz-Meador, P.↗
NASA NTRS2003

Evolutionary Design of an X-Band Antenna for NASA's Space Technology 5 Mission

We present an evolved X-band antenna design and flight prototype currently on schedule to be deployed on NASA s Space Technology 5 spacecraft in 2004. The mission consists of three small satellites that wall take science measurements in Earth s magnetosphere. The antenna was evolved to meet a challenging set of mission requirements, most notably the combination of wide beamwidth for a circularly-polarized wave and wide bandwidth. Two genetic algorithms were used: one allowed branching an the antenna arms and the other did not. The highest performance antennas from both algorithms were fabricated and tested. A handdesigned antenna was produced by the contractor responsible for the design and build of the mission antennas. The hand-designed antenna is a quadrifilar helix, and we present performance data for comparison to the evolved antennas. As of this writing, one of our evolved antenna prototypes is undergoing flight qualification testing. If successful, the resulting antenna would represent the first evolved hardware in space, and the first deployed evolved antenna.

Lohn, Jason D.↗
NASA NTRS2016

The Impact of New Trends in Satellite Launches on the Orbital Debris Environment

The main goal of this study is to examine the impact of new trends in satellite launch activities on the orbital debris environment and collision risk. As a foundation for the study, we developed a deployment scenario for satellites and associated rocket bodies based on publicly announced future missions. The upcoming orbital injection technologies, such as the new launch vehicles dedicated for small spacecraft and propulsive interstages, are also considered in this scenario. We then used a simulation tool developed in-house to propagate the objects within this scenario using variable-sized time-steps as small as one second to detect conjunctions between objects. The simulation makes it possible to follow the short- and long-term effects of a particular satellite or constellation in the space environment. Likewise, the effects of changes in the debris environment on a particular satellite or constellation can be evaluated. It is our hope that the results of this paper and further utilization of the developed simulation tool will assist in the investigation of more accurate deorbiting metrics to replace the generic 25-year disposal guidelines, as well as to guide future launches toward more sustainable and safe orbits.

conjunction analysis↗