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This PowerPoint presentation will discuss a new small spacecraft architecture which takes advantage of ESPA (Evolved Expendable Launch Vehicle (EELV) Secondary Payload Adapter) Class rideshare opportunities.
This PowerPoint presentation will discuss efforts to increase rideshare initiatives: current technical challenges and initiatives underway.
As part of NASA's new deep space exploration system, the Space Launch System (SLS) will provide the United States with guaranteed access to deep space and an unparalleled capability for launching primary and co-manifested payloads beyond Earth's orbit. Planned missions for the new SLS family of vehicles include launching the Orion spacecraft and elements of the new Gateway astronaut-tended outpost to lunar orbit and sending robotic probes deep into the solar system, such as to Jupiter's moon Europa. If mission parameters allow, secondary payloads in 6U, 12U or larger sizes will also have rideshare opportunities, providing CubeSats with access to deep space. The SLS vehicle will evolve into progressively more powerful variants with fairings in several sizes available to meet an array of mission needs. Superior mass, volume and characteristic energy (C3) enable sending larger, heavier payloads to a variety of destinations. Several elements of the Block 1 vehicle for the first mission, Exploration Mission-1 (EM-1) are complete and have been delivered to the Exploration Ground Systems (EGS) Program at Kennedy Space Center (KSC), which has responsibility for integrating and launching the vehicle. Contractors are already at work manufacturing the second Block 1 vehicle and incorporating numerous lessons learned in manufacturing America's first super heavy-lift deep space rocket since the Apollo Program's Saturn V enabled humankind to take a giant leap forward.
Propulsion systems on SmallSats and CubeSats are becoming more common as launch vehicles like SLS begin providing rideshares to destinations beyond low earth orbit. While these spacecraft are a solution to achieve advanced science objectives at lower cost, they are still subject to the same rigorous standards as other larger spacecraft, such as NASA STD-5019, due to the high pressures within their propulsion systems at launch. A Post Launch Pressurization System would allow a spacecraft to be launched at low pressure, then be pressurized through the controlled chemical decomposition of a solid reactant after the spacecraft is jettisoned from the launch vehicle. This increases the safety of the secondary spacecraft, and effectively mitigates risks to the launch vehicle and other payloads.
Earth's magnetosphere is a large magnetic cavity formed through the interaction of the solar wind and Earth's intrinsic magnetic field. Solar wind energy enters this cavity through a boundary - the magnetopause - separating Earth's field from the solar wind. This energy leads to many forms of "space weather", including the aurora, geomagnetic storms, and energization of the Van Allen radiation belts. Despite decades of research, we still do not understand the extent of dayside reconnection sites, nor do we have a quantifiable understanding of how much energy enters the magnetosphere during different solar wind conditions - necessary for space weather prediction. On the nightside, impulsive flows at various spatial and temporal scales occur frequently during storms and substorms, and couple to the ionosphere through still unresolved physical mechanisms. Because the magnetosphere is so large, it has been understood since the dawn of the space age that a full understanding of this complex region could only be achieved with a large fleet of in situ spacecraft. NASA has studied one such constellation, the socalled "Magnetospheric Constellation" (MagCon), since the1990's, but it is deemed too expensive to implement using traditional approaches. The CubeSat/ SmallSat revolution represents a fundamental disruption to traditional mission architectures, and in this paper I will discuss how, by leveraging innovation in spacecraft subystems, advanced manufacturing, and access to space, we can finally realize this long-term vision of exploration and discovery. The proposed modular approach, utilizing rideshare and propulsive ESPAs, would also enable worldwide participation in the mission, and is applicable to any constellation mission, including Earth Science missions.
This PowerPoint presentation will discuss a new small spacecraft architecture which takes advantage of ESPA Class rideshare opportunities.
Space is becoming increasingly congested as the number of on-orbit satellites and debris objects continues to grow. Space traffic management (STM) is critical for ensuring that the expanding orbital population operates safely and efficiently, avoiding collisions and radio-frequency interference while still facilitating widespread space operations. Recent events such as the FCC approval of SpaceX’s ~12,000 satellite constellation, the signing of Space Policy Directive 3 (which moves Space Situational Awareness responsibilities away from the Department of Defense and to a civil agency), and the growth in rideshare and small launch vehicles illustrate the rapidly changing nature of this domain. This paper will describe the concept of operations (ConOps) for a civilian STM research initiative, which has been developed from previous NASA work to enable safe operation of small unmanned aircraft systems. The STM ConOps proposes an architecture to enable efficient data sharing and coordination between participants to facilitate safe spaceflight operations. It is designed to utilize and promote the emerging field of commercial STM services, as a complement to existing government-provided STM services. The concept envisions a phased evolution that would gradually integrate additional capabilities, proposing a first phase architecture and tentative plans for a broader system. Work towards developing an STM research and prototyping platform is also discussed.
NASA is embarking on a new generation of missions to the Moon, known as the Artemis program. The Agency’s new super heavy-lift launch vehicle, the Space Launch System (SLS), is a critical enabling capability for these efforts, which will serve as a proving ground for future crewed expeditions to Mars. SLS is designed to return astronauts to the Moon in the Orion spacecraft and to launch more mass and provide more volume for critical payloads than commercially available vehicles, giving NASA a unique asset for deep space exploration. When performance margin and volume is available, as it is on the first flight, Artemis I (previously Exploration Mission-1), SLS can offer CubeSats access to deep space, beyond Earth’s orbit. The Artemis I flight has 13 6U (14 kg) CubeSats manifested, and the Program is currently accepting proposals for 6U and 12U payloads for the Artemis II flight through the agency’s CubeSat Launch Initiative (CSLI) program. Proposals from U.S.-based payload developers to fly on Artemis II will be accepted through CSLI until November 4, 2019. When ridesharing on SLS deep space missions, CubeSats must demonstrate they will not interfere with primary mission objectives. The 13 Artemis I CubeSats that hail from industry, academia, NASA and its international partners represent an array of exciting deep space science investigations and technology demonstrations that may help inform future Artemis missions. The initial SLS Block 1 vehicle for the Artemis I flight is fully manufactured; several elements are complete and have been delivered to the Exploration Ground Systems (EGS) Program at Kennedy Space Center (KSC), which has responsibility for integrating and launching the SLS and Orion stack. Completed elements of the Artemis I vehicle include the Orion Stage Adapter (OSA), which houses the 13 Artemis I CubeSats. With the Artemis I flight hardware and software nearing completion, work is in progress for the second Block 1 launcher, designated for the crewed Artemis II flight. Hardware is being manufactured for every element of the Artemis II vehicle. In this paper, the author will review the status of the Artemis I vehicle and the payloads manifested for the initial flight. Information on the Secondary Payload Deployment System (SPDS) that the SLS Program provides to payload developers will also be included.
There is considerable interest in utilizing Small Spacecraft beyond low Earth orbit. In November of 2018, successful data relay operations of the MarCO CubeSats during the entry, descent, and landing (EDL) of the Mars InSight mission showed the viability of using CubeSats for interplanetary missions. Additional testament to the promise of Small Spacecraft class rideshare missions is the upcoming Artemis-1 flight test that will launch thirteen 6U CubeSats, as well as the establishment of NASA’s SIMPLEx program which will conduct stand-alone planetary science missions that launch with a primary payload. It is anticipated that continued innovations in Small Spacecraft capabilities combined with small EDL systems will expand the range of potential missions to allow for in situ investigations. Recently, NASA invested in the development of a new, efficient and capable ablative thermal protection system (TPS), utilizing 3-D Weaving. The new dual layer TPS, Heatshield for Extreme Entry Environment Technology (HEEET), is tailorable, scalable, robust, mass efficient and capable of supporting in situ missions across the solar system. Combining the HEEET entry system with innovative Small Spacecraft technology will substantially expand the range of Small Spacecraft mission applications by offering the capability for high speed entry or aerocapture at destinations with atmospheres. This paper will describe the HEEET aeroshell system and highlight various mission concepts including a dual technology demonstration mission that is under development and other concepts to deliver instruments for planetary science.
In this paper we present a simulation tool that has been used to develop an architecture for a low-cost Position, Navigation and Time (PNT) system to provide PNT services on and around the Moon using non-dedicated low-cost orbital and ground assets. The simulation tool has been developed to be flexible and is capable of modeling and analyzing the many different capabilities and configurations that the non-dedicated assets could support. The tool models the creation of an ad hoc swarm, the localization of this swarm and the subsequent provision of PNT services from this swarm. We present results from several studies of select configurations chosen to reflect existing and future real-world needs and capabilities. Over the next few decades there is expected to be a substantial increase in Lunar missions supporting and inspired by NASA’s Artemis Program. It is expected that a large fraction of these missions will be low cost, utilizing rideshares and CubeSats, just as has seen in Earth orbit over the past decade. Many of these missions will need navigation capabilities but may be unable to support the large power, mass and weight that a weak GNSS or DSN based navigation solution would entail. As Lunar PNT service demand is not likely to be needed over the entire lunar surface 24/7, the creation of a dedicated Lunar GNSS constellation cannot be justified. Hence, our proposed architecture envisions utilizing existing Lunar science and exploration assets to create ad hoc and on demand Lunar PNT swarms capable of providing PNT services to low cost missions. The simulation reflects the two distinct parts of the architecture: the creation of a Lunar PNT swarm using non-dedicated existing assets, and the quantitative modeling of the quality of the PNT services that this swarm provides. To support the former, the simulation supports various swarm localization techniques, including centralized and distributed EKFs both of which support pluggable dynamics models. In modeling the PNT service performance the simulation adopts standard techniques from the GNSS community, including providing degree of precision (DOP) estimates for theoretical end users. Crucially, all asset capabilities, including clock accuracy, independent location self-knowledge and timing measurement precision can be set independently for each asset, reflecting the key concept of utilizing non-dedicated assets. The simulation is predominantly implemented in MATLAB, with GMAT being used for the propagation of orbital assets. The paper will present results from the simulation reflecting the tool’s flexibility and focusing on scenarios that match real-world proposed missions, including scenarios designed to provide PNT support to lunar surface missions similar to NASA Ames' forthcoming VIPER mission. The performance of centralized and decentralized swarm configuration and localization techniques will be compared. Finally, performance of the PNT service provided by an autonomous Lunar PNT swarm will be compared to existing radiometric and weak GNSS methods.
Spiral escape trajectories enable a spacecraft to escape Earth’s gravity and travel into inter-planetary space under its own power. This analysis demonstrates how to design spiral escape trajectories that leverage single and double Lunar gravity assists and compares the effectiveness of each trajectory type. The trajectories are constructed by combining backwards propagated Q-Law and a perturbed Sims-Flanagan transcription to design the spiral and interplanetary phases, respectively. We demonstrate that these mission types result in significant propellant savings when compared to a conventional, no Lunar swingby spiral escape, and are highly beneficial for the interplanetary rideshare mission concept.
There is considerable interest in utilizing Small Spacecraft beyond low Earth orbit. In November of 2018, successful data relay operations of the MarCO CubeSats during the entry, descent, and landing (EDL) of the Mars InSight mission showed the viability of using CubeSats for interplanetary missions. Additional testament to the promise of Small Spacecraft class rideshare missions is the upcoming Artemis-1 flight test that will launch thirteen 6U CubeSats, as well as the establishment of NASA’s SIMPLEx program which will conduct stand-alone planetary science missions that launch with a primary payload. It is anticipated that continued innovations in Small Spacecraft capabilities combined with small EDL systems will expand the range of potential missions to allow for in situ investigations. Recently, NASA invested in the development of a new, efficient and capable ablative thermal protection system (TPS), utilizing 3-D Weaving. The new dual layer TPS, Heatshield for Extreme Entry Environment Technology (HEEET), is tailorable, scalable, robust, mass efficient and capable of supporting in situ missions across the solar system. Combining the HEEET entry system with innovative Small Spacecraft technology will substantially expand the range of Small Spacecraft mission applications by offering the capability for high speed entry or aerocapture at destinations with atmospheres. This paper will describe the HEEET aeroshell system and highlight various mission concepts including a dual technology demonstration mission that is under development and other concepts to deliver instruments for planetary science.
NASA’s Mars Exploration Program is studying a potential Mars Sample Return (MSR) campaign consisting of a series of missions over the next decade that would return samples collected at Mars for analyses in terrestrial laboratories. It is anticipated that during such a campaign, many in the Mars science community would seek to continue high-priority science investigations in parallel to those provided by geological and astrobiological sample return investigations. To respond to this anticipated desire of the science community, JPL is performing a study of small spacecraft mission concepts to Mars that could bridge the gap between MSR and other desired science investigations at Mars. The goal of the study is to utilize smaller, affordable missions in performing highpriority science investigations as defined in the National Academy of Sciences Decadal Survey, Mars Exploration Program Analysis Group (MEPAG) goals, and Human Exploration and Operations (HEO) Strategic Knowledge Gaps. The study targets the use of small spacecraft with greater science capability than currently achievable with CubeSats. The target spacecraft wet mass is approximately 100 to 350 kilograms. Methods of access to Mars considered in this study include a self-propelled transit from Earth geosynchronous transfer orbit (GTO) to Mars as a secondary payload in a rideshare configuration. The study investigates mission concepts, science objectives, mission designs, concept of operations, enhancing technologies, and mission costs, along with launch vehicle interfaces. The cost estimates of the mission concepts studied range from below $100 million to less than $300 million for development through launch. This paper concludes with an outline of several examples of small spacecraft mission concepts to Mars that demonstrate significant scientific capability, are technically feasible, and fit within the desired cost range.
The Venusian atmosphere creates inhospitable temperature and pressure conditions for the surface of Venus, Earth’s twin planet, making in-situ measurements of any appreciable length difficult, expensive, and risky to obtain. Yet, because of the apparent youthfulness of Venus’ surface features, long-duration seismic observations are in high demand in order to determine and understand the dynamic processes taking place in lieu of plate tectonics. The Venus Airglow Measurements and Orbiter for Seismicity (VAMOS) mission concept would make use of the dense Venusian atmosphere as a medium to conduct seismic vibrations from the surface to the ionosphere. Here, the resulting atmospheric gravity waves and acoustic waves can be observed in the form of perturbations in airglow emissions, the basic principles for which have been demonstrated at Earth following a tsunami and at Venus with the European Venus Express’s Visible and Infrared Thermal Imaging Spectrometer (VIRTIS) instrument. In addition, these observations would enable VAMOS to determine the crustal structure and ionospheric variability of Venus without approaching the surface or atmosphere themselves. Equipped with an instrument of modest size and mass, the baseline VAMOS spacecraft is designed to fit within a SmallSat form factor and travel to Venus predominantly under its own power. VAMOS would enter into an orbit uniquely suited for the longduration, full-disk staring observations required for seismic readings. VAMOS’ journey would be enabled by modern solar electric propulsion technology and SmallSat avionics, which allow the spacecraft to reach Venus and autonomously filter observation data on board to detect Venus-quake events. Currently, trade studies are being conducted to determine mission architecture robustness to launch and rideshare opportunities. Key spacecraft challenges for VAMOS, just as with many SmallSat-based mission concepts, include thermal and power management, onboard processing capabilities, telecommunications throughput, and propulsion technology. The VAMOS mission concept is being studied at JPL as part of the NASA Planetary Science Deep Space SmallSat Studies (PSDS3) program, which will not only produce a viable and exciting mission concept for a Venus SmallSat, but will have the opportunity to examine many issues facing the development of SmallSats for planetary exploration. These include SmallSat solar electric propulsion, autonomy, telecommunications, and resource management that can be applied to various inner solar system mission architectures.
The apparent youthfulness of Venus’ surface features, given a lack of plate tectonics, is very intriguing; however, longduration seismic observations are essentially impossible given the inhospitable surface of Venus. The Venus Airglow Measurements and Orbiter for Seismicity (VAMOS) mission concept uses the fact that the dense Venusian atmosphere conducts seismic vibrations from the surface to the airglow layer of the ionosphere, as observed on Earth. Similarly, atmospheric gravity waves have been observed by the European Venus Express’s Visible and Infrared Thermal Imaging Spectrometer (VIRTIS) instrument. Such observations would enable VAMOS to determine the crustal structure and ionospheric variability of Venus without approaching the surface or atmosphere. Equipped with an instrument of modest size and mass, the baseline VAMOS spacecraft is designed to fit within an ESPA Grande form factor and travel to Venus predominantly under its own power. Trade studies have been conducted to determine mission architecture robustness to launch and rideshare opportunities. The VAMOS mission concept was studied at JPL as part of the NASA Planetary Science Deep Space SmallSat Studies (PSDS3) program, which has not only produced a viable and exciting mission concept for a Venus SmallSat, but has also examined many issues facing the development of SmallSats for planetary exploration, such as SmallSat solar electric propulsion, autonomy, telecommunications, and resource management that can be applied to various inner solar system mission architectures.
This paper describes the evolution of the NASA Double Asteroid Redirection Test (DART) mission design and navigation. Specifically, the mission has been conceived as (1) a hydrazine bus on a ballistic trajectory, (2) a low-thrust bus launching from a geostationary transfer orbit and spiraling to escape, and (3) a lowthrust bus that launches with a small positive escape energy. This paper discusses the rationale in favor of the third concept, low energy escape, and describes the key mission design and navigation studies. In an effort to be compatible with an unknown co-manifest partner, the trajectory design must account for a large range of launch energies, orientations, and dates. The navigation approach must account for sensitive regions in the trajectory and plan for both low-thrust and chemical phases of flight. These findings are relevant to other missions pursuing low-cost interplanetary rideshare concepts.
Selected by NASA as an ESPA-class rideshare technology demonstration mission to launch with the Interstellar Mapping and Acceleration Probe (IMAP) mission in 2025, the NASA MSFC Solar Cruiser mission will mature solar sail technology for use in future Heliophysics missions, as well as missions of interest across a broad user community including NOAA, Earth, and Planetary Sciences. Solar sails, which use reflected sunlight to derive thrust, can be used to create artificial equilibria and near-indefinite station-keeping at locations sunward of L1 along the Sun Earth Line (SEL), or at any desired offset from the SEL leading or trailing the Earth in its orbit. They can change the heliocentric inclination of a spacecraft from the ecliptic to as high as solar polar, stopping and remaining at any intermediate inclination orbit in between or can be used around a range of solar system bodies. The Solar Cruiser mission will fly a small spacecraft (~100 kg) with a large (>1600 square meter) solar sail containing embedded reflectivity control devices (RCDs) and photovoltaic cells, attaining a characteristic acceleration of >0.12 mm/s2. The mission concept includes successful deployment of the solar sail, validation of all sail subsystems, controlled station-keeping inside of the Sun-Earth L1 point, demonstration of pointing performance for science imaging, and finally an increase in heliocentric inclination (out of the ecliptic plane). Solar Cruiser would serve as a pathfinder for missions that observe the solar environment from unique vantage points such as a high inclination solar mission, opening a fundamentally new range of observational capabilities for the Heliophysics Program and for space weather monitoring. Observations away from the Sun-Earth line (SEL) present unique opportunities for answering the outstanding science questions of Heliophysics, for improving space-weather monitoring and prediction, and for revealing new discoveries about our Sun and solar system. High solar inclinations are particularly compelling. Investment in, and demonstration of, the technology needed to enable polar missions is essential to making this unique vantage point a reality in the next decade.
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.