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NASA NTRSDate not supplied

Trends in Small Satellite Presentations from 2017-2019

As the small satellite community expands and new technologies emerge, it is important to understand and interpret these changes over time. In doing so, we observe progress and aid future development within the field. During the summer of 2020 we compiled for assessment purposes, archived presentation data for years 2017-2019 from three primary sources: the CubeSat Developers Workshop, Interplanetary Small Satellite Conference, and the Small Satellite Conference. A few examples of the information we recorded and compiled were presenter names and affiliations, presentation topic, and mission progress status. Ultimately, we reviewed roughly 600 presentations between the three conferences and the data obtained therein form the basis of our trend assessment. The paper focuses on trends interpreted through the assessment of key elements available in the content of each presentation to include: SmallSat mission developers, subsystem developments and the expanded scope for small satellite destinations. Data was generalized using various forms of analysis depending on the type of information being assessed. In the end, we achieved our goal of reducing the content to its key points and major takeaways from the conference proceedings. In this presentation, the observed trends in data and our findings will be discussed. First, we will cover changes in presentation topics by categorizing them as either: science, technology, science/technology, or other. Our next topic will be about the agencies and organizations at the forefront of small satellite research and development. The following section will explain trends in subsystem developments for telecommunications, propulsion, power, and thermal management. Further discussion will highlight how the scope of interplanetary spaceflight has expanded since 2017 as new small spacecraft missions venture beyond low-Earth orbit into deep space exploration. And finally, we will observe what information was lacking in archived presentations which included but not limited to: mission’s funding source, presentation focus area, and software used.

Small Satellite↗
NASA NTRSDate not supplied

Trends in Small Satellite Presentations from 2017 - 2019

As the small satellite community expands and new technologies emerge, it is important to understand and interpret these changes over time. In doing so, we observe progress and aid future development within the field. During the summer of 2020 we compiled for assessment purposes, archived presentation data for years 2017-2019 from three primary sources: the CubeSat Developers Workshop, Interplanetary Small Satellite Conference, and the Small Satellite Conference. A few examples of the information we recorded and compiled were presenter names and affiliations, presentation topic, and mission progress status. Ultimately, we reviewed roughly 600 presentations between the three conferences and the data obtained therein form the basis of our trend assessment. The paper focuses on trends interpreted through the assessment of key elements available in the content of each presentation to include: SmallSat mission developers, subsystem developments and the expanded scope for small satellite destinations. Data was generalized using various forms of analysis depending on the type of information being assessed. In the end, we achieved our goal of reducing the content to its key points and major takeaways from the conference proceedings. In this presentation, the observed trends in data and our findings will be discussed. First, we will cover changes in presentation topics by categorizing them as either: science, technology, science/technology, or other. Our next topic will be about the agencies and organizations at the forefront of small satellite research and development. The following section will explain trends in subsystem developments for telecommunications, propulsion, power, and thermal management. Further discussion will highlight how the scope of interplanetary spaceflight has expanded since 2017 as new small spacecraft missions venture beyond low-Earth orbit into deep space exploration. And finally, we will observe what information was lacking in archived presentations which included but not limited to: mission’s funding source, presentation focus area, and software used.

Small Satellites↗
NASA NTRS2015

Propulsion System Testing for the Iodine Satellite (iSAT) Demonstration Mission

CUBESATS are relatively new spacecraft platforms that are typically deployed from a launch vehicle as a secondary payload, providing low-cost access to space for a wide range of end-users. These satellites are comprised of building blocks having dimensions of 10x10x10 cm cu and a mass of 1.33 kg (a 1-U size). While providing low-cost access to space, a major operational limitation is the lack of a propulsion system that can fit within a CubeSat and is capable of executing high delta v maneuvers. This makes it difficult to use CubeSats on missions requiring certain types of maneuvers (i.e. formation flying, spacecraft rendezvous). Recently, work has been performed investigating the use of iodine as a propellant for Hall-effect thrusters (HETs) 2 that could subsequently be used to provide a high specific impulse path to CubeSat propulsion. 3, 4 Iodine stores as a dense solid at very low pressures, making it acceptable as a propellant on a secondary payload. It has exceptionally high ρIsp (density times specific impulse), making it an enabling technology for small satellite near-term applications and providing the potential for systems-level advantages over mid-term high power electric propulsion options. Iodine flow can also be thermally regulated, subliming at relatively low temperature (less than 100 C) to yield I2 vapor at or below 50 torr. At low power, the measured performance of an iodine-fed HET is very similar to that of a state-of-the-art xenon-fed thruster. Just as importantly, the current-voltage discharge characteristics of low power iodine-fed and xenon-fed thrusters are remarkably similar, potentially reducing development and qualifications costs by making it possible to use an already-qualified xenon-HET PPU in an iodine-fed system. Finally, a cold surface can be installed in a vacuum test chamber on which expended iodine propellant can deposit. In addition, the temperature doesn't have to be extremely cold to maintain a low vapor pressure in the vacuum chamber (it is under 10(exp -6) torr at −75 C), making it possible to 'cryopump' the propellant with lower-cost recirculating refrigerant-based systems as opposed to using liquid nitrogen or low temperature gaseous helium cryopanels. An iodine-based system is not without its challenges. The primary challenge is that the entire feed system must be maintained at an elevated temperature to prevent the iodine from depositing (transitioning from the gas phase directly back into the solid phase), which will block the propellant feed lines. Furthermore, deposition will occur unless the temperature in the lines is not greater than the temperature of the propellant reservoir. The flow rate can be controlled by adjusting the heating applied to the reservoir, but as with any thermal control there is a relatively slow response to changes in the heating rate. In the present paper, we describe the propulsion and propellant feed system for the iodine satellite (iSAT) flight demonstration mission. The system is based around the Busek BHT-200 Hall thruster, which has been modified for chemical compatibility with iodine vapor. While the gross propellant flow rate is maintained by the heated propellant reservoir, the flow to the anode and cathode are adjusted using two heated Vacco proportional flow control valves (PFCV), which provide very fast response on the flow rate adjustment. The flight mission design layout will be presented, showing how the system will be packaged into the overall 12-U spacecraft and the techniques being employed to protect the remaining spacecraft hardware from the propulsion system (e.g., plasma impingement, iodine deposition, thermal loads). In addition to the flight system design, results of testing the thruster and cathode with both operating on iodine propellant are presented. The tests are conducted on a thrust stand (see Fig. 1) in a large vacuum chamber containing a beam dump chilled to below -100 C to 'cryopump' the propellant. The thruster performance during these tests is presented, with these data used to evaluate the feed system and guide further refinements. Results of relatively long duration testing are presented to demonstrate the capability to operate for the length of the iSAT mission and to perform a number of re-starts as will be required by the mission concept of operations.

Polzin, Kurt A.↗
NASA NTRS2016

Engineering Model Propellant Feed System Development for an Iodine Hall Thruster Demonstration Mission

CUBESATS are relatively new spacecraft platforms that are typically deployed from a launch vehicle as a secondary payload, providing low-cost access to space for a wide range of end-users. These satellites are comprised of building blocks having dimensions of 10x10x10 cu cm and a mass of 1.33 kg (a 1-U size). While providing low-cost access to space, a major operational limitation is the lack of a propulsion system that can fit within a CubeSat and is capable of executing high (Delta)v maneuvers. This makes it difficult to use CubeSats on missions requiring certain types of maneuvers (i.e. formation flying, spacecraft rendezvous). Recently, work has been performed investigating the use of iodine as a propellant for Hall-effect thrusters (HETs) 2 that could subsequently be used to provide a high specific impulse path to CubeSat propulsion. 3, 4 Iodine stores as a dense solid at very low pressures, making it acceptable as a propellant on a secondary payload. It has exceptionally high ρIsp (density times specific impulse), making it an enabling technology for small satellite near-term applications and providing the potential for systems-level advantages over mid-term high power electric propulsion options. Iodine flow can also be thermally regulated, subliming at relatively low temperature (< 100 C) to yield I2 vapor at or below 50 torr. At low power, the measured performance of an iodine-fed HET is very similar to that of a state-of-the-art xenon-fed thruster. Just as importantly, the current-voltage discharge characteristics of low power iodine-fed and xenon-fed thrusters are remarkably similar, potentially reducing development and qualifications costs by making it possible to use an already-qualified xenon-HET PPU in an iodine-fed system. Finally, a cold surface can be installed in a vacuum test chamber on which expended iodine propellant can deposit. In addition, the temperature doesn't have to be extremely cold to maintain a low vapor pressure in the vacuum chamber (it is under 10(exp −6) torr at −75 C), making it possible to 'cryopump' the propellant with lower-cost recirculating refrigerant-based systems as opposed to using liquid nitrogen or low temperature gaseous helium cryopanels. In the present paper, we describe the design and testing of the engineering model propellant feed system for iSAT (see Fig. 1). The feed system is based around an iodine propellant reservoir and two proportional control valves (PFCVs) that meter the iodine flow to the cathode and anode. The flow is split upstream of the PFCVs to both components can be fed from a common reservoir. Testing of the reservoir is reported to demonstrate that the design is capable of delivering the required propellant flow rates to operate the thruster. The tubing and reservoir are fabricated from hastelloy to resist corrosion by the heated gaseous iodine propellant. The reservoir, tubing, and PFCVs are heated to ensure the sublimed propellant will not re-deposit within the feed system. Heating is accomplished through a number of individual zones to control the overall power expended on heating the system and insulation is employed to minimize the amount of power used to heat the system prior to thruster operation.

Polzin, Kurt A.↗
NASA NTRSDate not supplied

Lunar Node 1 and Beyond

Lunar Node 1 (LN-1) is an S-band Navigation beacon for lunar applications that was recently designed and built at MSFC. As part of NASA's Commercial Lunar Payload Services initiative, this beacon will be delivered to the moon's surface on Intuitive Machine's NOVA-C lunar lander in November 2021. During this mission, LN-1’s goal will be to demonstrate navigation technologies that can support local surface and orbital operations around the moon, enabling autonomy which would decrease dependency on heavily utilized Earth based assets like the Deep Space Network. To do this, LN-1’s design leverages Cubesat components as well as the Multi-spacecraft Autonomous Positioning System (MAPS) algorithms, which enable the autonomous spacecraft positioning through communication-integrated navigation measurements. In addition to demonstrating the MAPS payload, the radio will also be used in standard tone-based non-coherent ranging and Doppler tracking to provide an alternate approaches and comparisons for navigation performance. LN-1 will represent a single node in a potential greater MAPS network of assets. While LN-1 awaits launch to demonstrate this initial use of one-way navigation, designs are already under way for proposed subsequent missions like LN-2 to enable expanded capability. The next steps in these missions seek to include a receive capability to allow two-way ranging support with other operational spacecraft in the lunar vicinity, and to provide some component upgrades related to the radio, power, and thermal systems that will improve the longevity of the payload in the harsh lunar environment. The LN-1 design details, status, and potential forward work with mission like LN-2 will be outlined in this presentation.

Evan John Anzalone↗
NASA NTRS2020

PAssive Thermal Coating Observatory Operating in Low Earth Orbit (PATCOOL)

The PATCOOL is a NASA sponsored, University of Florida developed 3U CubeSat meant to investigate the feasibility of using a cryogenic selective surface coating as a new, more efficient way of passively cooling components in space. Initial tests on the ground demonstrate that this coating should provide a much higher reflectance of the Sun's irradiant power than any existing coating, while still providing far-infrared power emission. The ultimate validation of this technology requires on-orbit testing. PATCOOL hosts a 4-sample housing, with the samples shaped as thin cylinders (coin-like). Two samples are coated with state-of-the-art material, while the other pair uses the new coating to be evaluated. The temperatures of all samples during the mission (minimum 72 hours of data collection) are measured via thermistors. The samples are connected via thin Kevlar strings to the housing, to minimize heat transfer. The housing is designed to shield the samples from Earth's thermal radiation, and the CubeSat is attitude stabilized and controlled via a gravity gradient boom, magnetorquers and a reaction wheel set. Thermal Desktop simulations show PATCOOL's ability to thermally isolate the samples from heat exchanges other than with Sun and deep space, thanks to its thermal design and the chosen attitude profile

Ojeda, Carlos↗
NASA NTRS2020

CubeSat Platform Enables an Inexpensive Space Telescope

Space telescopes have historically been expensive to plan, build, and launch. The Astrophysics Division of NASA’s Science Mission Directorate is now leveraging the development of small spacecraft known as CubeSats by universities and industry to enable capable and reliable platforms with well-focused goals, rapid development times, and affordable costs

technology↗
NASA NTRS2016

Green Monopropellant Status at Marshall Space Flight Center

NASA Marshall Space Flight Center is continuing investigations into the use of green monopropellants as a replacement for hydrazine in spacecraft propulsion systems. Work to date has been to push technology development through multiple activities designed to understand the capabilities of these technologies. Future work will begin to transition to mission pull as these technologies are mature while still keeping a solid goal of pushing technology development as opportunities become available. The AF-M315E activities began with hot-fire demonstration testing of a 1N monopropellant thruster in FY 14 and FY15. Following successful completion of the preliminary campaign, changes to the test stand to accommodate propellant conditioning capability and better control of propellant operations was incorporated to make testing more streamlined. The goal is to conduct hot-fire testing with warm and cold propellants using the existing feed system and original thruster design. Following the 1N testing, a NASA owned 100 mN thruster will be hot-fire tested in the same facility to show feasibility of scaling to smaller thrusters for cubesat applications. The end goal is to conduct a hot-fire test of an integrated cubesat propulsion system using an SLM printed propellant tank, an MSFC designed propulsion system electronic controller and the 100 mN thruster. In addition to the AF-M315E testing, MSFC is pursuing hot-fire testing with LMP-103S. Following our successful hot-fire testing of the 22N thruster in April 2015, a test campaign was proposed for a 440N LMP-103S thruster with Orbital ATK and Plasma Processes. This activity was funded through the Space Technology Mission Directorate (STMD) ACO funding call in the last quarter of CY15. Under the same funding source a test activity with Busek and Glenn Research Center for testing of 5N AF-M315E thrusters was proposed and awarded. Both activities are in-work with expected completion of hot-fire testing by the end of FY17. MSFC is continuing to coordinate with the AF and academia on understanding the chemical reactions that occur in AF-M315E. An on-going investigation of the catalyst bed species using Raman Spectroscopy through the NASA Technology Research Fellowship Program (NSTRF) is looking for ways to minimize the amount of computation required by understanding the intermediate species created in the catalyst bed. The MSFC team is also working with commercial partners through Cooperative Agreement Notices (CAN's). Partnerships with commercial and academia include work in non-catalytic ignition of AF-M315, spark ignition of hybrid cubesat systems, printed SLM tanks, and dual-mode (electric and chemical) propulsion systems is continuing.

Burnside, Christopher G.↗
NASA NTRS2013

NASA Facts: Nanosatellite Launch Adapter System (NLAS)

The Nanosatellite Launch Adapter System (NLAS) was developed to increase access to space while simplifying the integration process of miniature satellites, called nanosats or cubesats, onto launch vehicles. A standard cubesat measures about 4inches (10 cm) long, 4 inches wide,and 4 inches high, and is called a one-unit (1U) cubesat. A single NLAS provides the capability to deploy 24U of cubesats. The system is designed to accommodate satellites measuring 1U, 1.5U, 2U, 3U and 6U sizes for deployment into orbit. The NLAS may be configured for use on different launch vehicles. The system also enables flight demonstrations of new technologies in the space environment.

Chartres, James↗
NASA NTRS2015

Nanosatellite Launch Adapter System (NLAS)

The Nanosatellite Launch Adapter System (NLAS) was developed to increase access to space while simplifying the integration process of miniature satellites, called nanosats or CubeSats, onto launch vehicles. A standard CubeSat measures about 10 cm square, and is referred to as a 1-unit (1U) CubeSat. A single NLAS provides the capability to deploy 24U of CubeSats. The system is designed to accommodate satellites measuring 1U, 1.5U, 2U, 3U and 6U sizes for deployment into orbit. The NLAS may be configured for use on different launch vehicles. The system also enables flight demonstrations of new technologies in the space environment.

Nanosatellite Launch Adapter System↗
NASA NTRS2017

Advanced Deployable Shell-Based Composite Booms for Small Satellite Structural Applications Including Solar Sails

State of the art deployable structures are mainly being designed for medium to large size satellites. The lack of reliable deployable structural systems for low cost, small volume, rideshare-class spacecraft severely constrains the potential for using small satellite platforms for affordable deep space science and exploration precursor missions that could be realized with solar sails. There is thus a need for reliable, lightweight, high packaging efficiency deployable booms that can serve as the supporting structure for a wide range of small satellite systems including solar sails for propulsion. The National Air and Space Administration (NASA) is currently investing in the development of a new class of advanced deployable shell-based composite booms to support future deep space small satellite missions using solar sails. The concepts are being designed to: meet the unique requirements of small satellites, maximize ground testability, permit the use of low-cost manufacturing processes that will benefit scalability, be scalable for use as elements of hierarchical structures (e.g. trusses), allow long duration storage, have high deployment reliability, and have controlled deployment behavior and predictable deployed dynamics. This paper will present the various rollable boom concepts that are being developed for 5-20 m class size deployable structures that include solar sails with the so-called High Strain Composites (HSC) materials. The deployable composite booms to be presented are being developed to expand the portfolio of available rollable booms for small satellites and maximize their length for a given packaged volume. Given that solar sails are a great example of volume and mass optimization, the booms were designed to comply with nominal solar sail system requirements for 6U CubeSats, which are a good compromise between those of smaller form factors (1U, 2U and 3U CubeSats) and larger ones (12 U and 27 U future CubeSats, and ESPA-class microsatellites). Solar sail missions for such composite boom systems are already under consideration and development at NASA, as well as mission studies that will benefit from planned scaled-up versions of the composite boom technologies to be introduced. The paper presents ongoing research and development of thin-shell rollable composite booms designed under the particular stringent and challenging system requirements of relatively large solar sails housed on small satellites. These requirements will be derived and listed. Several new boom concepts are proposed and other existing ones are improved upon using thin-ply composite materials to yield unprecedented compact deployable structures. Some of these booms are shown in Fig. 1. For every boom to be introduced the scalable fabrication process developed to keep the overall boom system cost down will be shown. Finally, the initial results of purposely designed boom structural characterization test methods with gravity off-loading will be presented to compare their structural performance under expected and general load cases.

Fernandez, Juan M.↗
NASA NTRS2014

Life Cycle Analysis of Dedicated Nano-Launch Technologies

Recent technology advancements have enabled the development of small cheap satellites that can perform useful functions in the space environment. Currently, the only low cost option for getting these payloads into orbit is through ride share programs. As a result, these launch opportunities await primary payload launches and a backlog exists. An alternative option would be dedicated nano-launch systems built and operated to provide more flexible launch services, higher availability, and affordable prices. The potential customer base that would drive requirements or support a business case includes commercial, academia, civil government and defense. Further, NASA technology investments could enable these alternative game changing options.With this context, in 2013 the Game Changing Development (GCD) program funded a NASA team to investigate the feasibility of dedicated nano-satellite launch systems with a recurring cost of less than $2 million per launch for a 5 kg payload to low Earth orbit. The team products would include potential concepts, technologies and factors for enabling the ambitious cost goal, exploring the nature of the goal itself, and informing the GCD program technology investment decision making process. This paper provides an overview of the life cycle analysis effort that was conducted in 2013 by an inter-center NASA team. This effort included the development of reference nano-launch system concepts, developing analysis processes and models, establishing a basis for cost estimates (development, manufacturing and launch) suitable to the scale of the systems, and especially, understanding the relationship of potential game changing technologies to life cycle costs, as well as other factors, such as flights per year.

modeling↗
NASA NTRS2016

State-of-the-Art for Small Satellite Propulsion Systems

SmallSats are a low cost access to space with an increasing need for propulsion systems. NASA, and other organizations, will be using SmallSats that require propulsion systems to: a) Conduct high quality near and far reaching on-orbit research and b) Perform technology demonstrations. Increasing call for high reliability and high performing for SmallSat components. Many SmallSat propulsion technologies are currently under development: a) Systems at various levels of maturity and b) Wide variety of systems for many mission applications.

Small Satellite↗
NASA NTRS2014

Iodine Hall Thruster Propellant Feed System for a CubeSat

There has been significant work recently in the development of iodine-fed Hall thrusters for in-space propulsion applications.1 The use of iodine as a propellant provides many advantages over present xenon-gas-fed Hall thruster systems. Iodine is a solid at ambient temperature (no pressurization required) and has no special handling requirements, making it safe for secondary flight opportunities. It has exceptionally high ρI sp (density times specific impulse), making it an enabling technology for small satellite near-term applications and providing system level advantages over mid-term high power electric propulsion options. Iodine provides thrust and efficiency that are comparable to xenonfed Hall thrusters while operating in the same discharge current and voltage regime, making it possible to leverage the development of flight-qualified xenon Hall thruster power processing units for the iodine application. Work at MSFC is presently aimed at designing, integrating, and demonstrating a flight-like iodine feed system suitable for the Hall thruster application. This effort represents a significant advancement in state-of-the-art. Though Iodine thrusters have demonstrated high performance with mission enabling potential, a flight-like feed system has never been demonstrated and iodine compatible components do not yet exist. Presented in this paper is the end-to-end integrated feed system demonstration. The system includes a propellant tank with active feedback-control heating, fill and drain interfaces, latching and proportional flow control valves (PFCV), flow resistors, and flight-like CubeSat power and control electronics. Hardware is integrated into a CubeSat-sized structure, calibrated and tested under vacuum conditions, and operated under under hot-fire conditions using a Busek BHT-200 thruster designed for iodine. Performance of the system is evaluated thorugh accurate measurement of thrust and a calibrated of mass flow rate measurement, which is a function of reservoir temperature/pressure, the flow resistors, and the setting of the PFCV. The calibration is performed using independent flow control monitoring techniques, providing an in situ measure of the flowrate as a function of controllable parameters. The reservoir temperature controls the iodine sublimation rate, providing propellant to ths thruster by pressurizing the propellant feed system to approx.1-2 psi. Control of the temperature and the PFCV are used to maintain reservoir pressure and keep the thruster discharge current constant.

Polzin, Kurt A.↗
NASA NTRS2017

SmallSat Navigation via the Deep Space Network: Lunar Transport

Spacecraft component miniaturization, the standardized CubeSat form factor, and a corresponding increase in launch ride-share availability have led to a renewed interest in smaller, more agile spacecraft missions. Naturally, interest in SmallSat capabilities extends outside of low Earth orbit for a variety of science and technology demonstration applications. As with larger missions, near-term deep-space SmallSats will more than likely rely on telecommunications and tracking via NASA's Deep Space Network or similar facilities. Given the predicted growth in the number of deep space missions, e ective use of telecommunication resources will be more critical than ever, even with Multiple Spacecraft Per Aperture capability. In particular, SmallSat missions will likely face stricter limitations on two-way contact with the tracking stations, thus making other options like one-way data types more attractive. While all missions must eventually develop their own detailed navigation plans, a common set of references is needed to support SmallSat missions, especially in the early development phases. Our investigation provides this initial survey of expected navigation performance for standard radiometric data types, from traditional two-way Doppler and ranging capabilities to one-way equivalents, including delta-di erential one-way range. In this investigation, we examine transport within the Earth-Moon region, the rst step outward from Earth into deep space exploration.

Wood, Lincoln↗
NASA NTRS2020

Sample Transponder Implementations for PN DDOR

The design philosophy and recommended standards for PN DDOR transponders are proposed in RFM 20-04. The referenced document illustrates how to pick a chip rate, roll-off factor, and other PN parameters for a PN DDOR transmitter. The document also proposes various recommendations for these parameters. This document expands upon RFM 20-04 and presents two PN DDOR transmitters that were designed and tested at JPL. First the IRIS cubesat radio is presented with an 8 MHz spread spectrum DDOR signal. Next the Univeral Space Transponder (UST) transmitter is presented with a 32 MHz spread spectrum DDOR signal. These transmitter designs show that the proposed recommendations are reasonably implemented with current technology.

Volk, Christopher P↗
NASA NTRS2014

Life Cycle Analysis of Dedicated Nano-Launch Technologies

Recent technology advancements have enabled the development of small cheap satellites that can perform useful functions in the space environment. Currently, the only low cost option for getting these payloads into orbit is through ride share programs - small satellites awaiting the launch of a larger satellite, and then riding along on the same launcher. As a result, these small satellite customers await primary payload launches and a backlog exists. An alternative option would be dedicated nano-launch systems built and operated to provide more flexible launch services, higher availability, and affordable prices. The potential customer base that would drive requirements or support a business case includes commercial, academia, civil government and defense. Further, NASA technology investments could enable these alternative game changing options. With this context, in 2013 the Game Changing Development (GCD) program funded a NASA team to investigate the feasibility of dedicated nano-satellite launch systems with a recurring cost of less than $2 million per launch for a 5 kg payload to low Earth orbit. The team products would include potential concepts, technologies and factors for enabling the ambitious cost goal, exploring the nature of the goal itself, and informing the GCD program technology investment decision making process. This paper provides an overview of the life cycle analysis effort that was conducted in 2013 by an inter-center NASA team. This effort included the development of reference nano-launch system concepts, developing analysis processes and models, establishing a basis for cost estimates (development, manufacturing and launch) suitable to the scale of the systems, and especially, understanding the relationship of potential game changing technologies to life cycle costs, as well as other factors, such as flights per year.

nanolaunch↗
NASA NTRS2019

SmallSat missions enabled by paired low-thrust hybrid rocket and low-power long-life Hall thruster

The capabilities of a SmallSat-class spacecraft targeting the outer solar system and using a combined chemical and electric propulsion system are explored. The development of compact hybrid rockets has enabled high-thrust engines to be packaged tightly enough to fit on cubesat and SmallSat spacecraft. These hybrid rockets provide 10’s-100 N of thrust depending on the propellant load & >300 s of specific impulse and have been demonstrated in both ambient and vacuum environments. Advancements in low-power long-life Hall thruster technologies have provided the potential for significantly greater propellant throughputs, enabling their use as a primary propulsion element on interplanetary spacecraft. In a recent characterization test campaign, the MaSMi-DM Hall thruster demonstrated power throttling from 150 – 1000 W with >1500 s of specific impulse available at >500 W and >40% total thrust efficiency available at >300 W; peak values of 1940 s and 53% were observed. A notional low-mass spacecraft employing a combined hybrid rocket and low-power electric propulsion system was designed and used for mission concept analysis targeting the outer solar system. Using an imposed wet mass limit of 400 kg, mission trajectories to Saturn and Uranus were generated. Orbit capture with >40% of the launch mass was shown to be possible at either target, with mission transfer times of 7.5 years and 13.5 years for Saturn and Uranus, respectively. Significant follow-on mission activities near Saturn (e.g. to Titan & Enceladus) were also possible by carrying extra propellant mass while remaining under the total wet mass limit.

Karp, Ashley C.↗