LunaH-Map (Lunar Polar Hydrogen Mapper) CubeSat Mission
LunaH-Map is a secondary 6U CubeSat payload of Arizona State University (ASU), selected by the NASA Science Mission Directorate as part of the SIMPLEx (Small, Innovative Missions for Planetary Exploration) program. LunaH-Map will fly as a secondary payload on the Artemis-1 mission, formerly the EM-1 ( Exploration Mission) of the SLS (Space Launch System), with a planned launch in 2020. This mission is led by a team of researchers, graduate students, and undergraduates at Arizona State University in collaboration with NASA centers, JPL, universities, and commercial space businesses. 1) 2) 3) 4) 5)
There is renewed interested in lunar exploration and, in particular, in a more detailed characterization of hydrogen-rich regions at the lunar south pole that have been identified by previous NASA planetary science missions to the Moon. The LunaH-Map mission, selected by NASA’s Science Mission Directorate in late-2015, will make neutron measurements from a low altitude perilune orbit to help place important constraints on our understanding of the distribution of lunar polar volatiles within permanently shadowed regions (PSRs). The LunaH-Map spacecraft is designed to fit within a 6U+ CubeSat form factor carrying one science instrument, a new type of neutron spectrometer. A new type of detector material is used, as this was required to achieve sufficient efficiencies for neutron detection in such a small volume. In order to improve upon the spatial resolution achieved by previous neutron instruments at the Moon, LunaH-Map will achieve perilune altitudes between 10 – 15 km above the lunar surface.
Science: Since the detection of enhanced hydrogen at the lunar poles by the Lunar Prospector Neutron Spectrometer (LPNS) in 1998, there has been significant and continued interest in the nature and distribution of volatiles at the Moon and in other solar system bodies. The detection of reduced epithermal neutron flux at the poles by neutron experiments on board the Lunar Prospector (LP) spacecraft and the LRO (Lunar Reconnaissance Orbiter) provided positive identification of increased hydrogen concentrations in PSRs, likely in the form of water-ice.
The LP and LRO missions used neutron spectrometers from orbit to map bulk hydrogen distributions at the lunar poles. Neutron spectrometers measure hydrogen (which is bound in water ice and other hydrated mineral phases) within the top meter of the lunar surface by detecting the neutrons created by interactions with high energy galactic cosmic rays that interact with lunar regolith and subsequently leak out of the surface. The energy distribution of neutrons leaking out of lunar surface will be a function of the bulk geochemistry, hydrogen abundance, depth distribution of hydrogen and a variety of other factors related to the regolith and surface properties.
Neutrons in the energy range of ~0.4 eV to ~10 keV will be most sensitive to hydrogen abundance and have been used to map quantities of hydrogen down to ~50 ppm. From orbit, however, neutron detectors typically have poor spatial resolution, with an effective field of view of about one and a half times the orbital altitude. Previous missions have orbited at altitudes of tens to hundreds of kilometers, resulting in coarse (many kilometer-scale) maps of neutron counts. These maps are sufficient to reveal the regional hydrogen distribution at the poles but are not high enough resolution to reveal distributions of hydrogen enrichments within the PSRs (Permanently Shadowed Regions). Higher spatial resolution neutron maps of these regions (at scales of 10 – 20 km) may reveal enrichments of water ice within the PSRs. These maps would place important constraints on lunar volatile emplacement and lunar polar wander as well as help future mission planning (landing site selections) at the lunar south pole.
The maps of hydrogen enrichments that will be produced by LunaH-Map are a function of the science orbit, the altitude above terrain, the neutron detector response, the surrounding lunar topography, and the hydrogen content of the lunar surface that passes within the detector’s sensing area each orbit. Preliminary analysis of LunaH-Map’s ability to resolve lunar PSRs is shown in Figures 1 and 2. Figure 1 shows a subset of the LunaH-Map ground tracks where even small (<10 km) PSRs are nearly resolved (black arrow). Figure 2 shows that the PSRs are not fully resolved (fully resolved PSRs have a contribution of 1); however, some fine scale features are prominent near periapsis.
Figure 1: LunaH-Map science phase ground tracks (in blue) superimposed on Moon LRO LOLA Hillshade 237 m (v4, LOLA Science Team, retrieved from USGS web site). PSRs are labeled in red (image credit: LunaH-Map Team)
Trajectory: An innovative trajectory and mission profile has been designed in order to safely deliver the LunaH-Map spacecraft into a final lunar orbit that will fulfill the mission objectives. 6) The trajectory is partitioned into three primary phases; an Earth-Moon transfer phase, lunar orbit transition phase, and science phase, each of which presents unique challenges.
The Earth-Moon transfer phase begins with the spacecraft’s deployment from the SLS EM-1 Interim Cryogenic Propulsion Stage (ICPS) host vehicle at a distance from Earth of ~70,000 km. This insertion state is favorable as it occurs after the ICPS has exited the Van Allen Radiation Belts, but early enough to allow time for telemetry, tracking, and propulsive maneuvers prior to the first lunar flyby, labeled Periselene-1.
Figure 2: LunaH-Map altitude above terrain vs latitude along each blue ground-track shown in Figure 1. Bottom) Contribution to the observed neutron count rate by PSRs for the ground-tracks shown in Figure 1 (image credit: LunaH-Map Team)
Periselene-1 is used to contain the spacecraft within the Earth-Moon system and target the Sun-Earth WSB (Weak Stability Boundary) which allows for a low-energy transfer from Periselene-1 to Lunar orbit insertion. This nominal transfer trajectory includes one loop around the Earth with perigee > 100,000 km altitude to avoid the Van Allen Belts and concludes with a set of maneuvers to targets weak capture and orbit insertion at the second lunar encounter, Periselene-2.
After weak capture, the lunar orbit transition phase begins with a set of maneuvers that ensures the spacecraft is captured in a stable lunar orbit. After “strong capture,” a long cadence of maneuvers spanning several months are executed to reduce the orbital energy and eventually enter into the final science orbit. These maneuvers will occur contiguously when the spacecraft is not in eclipse nor communicating with ground stations and burn updates will occur at regular intervals to ensure the transition profile is maintained. If necessary, the spacecraft may enter a stable circular orbit during its transition to avoid Earth eclipses before targeting the final science orbit. After maneuvering for several months, the spacecraft’s orbital parameters will eventually match those of the final desired science orbit. An illustrated view of the transition from weak lunar capture to elliptical science orbit is shown in Figure 3.
Figure 3: LunaH-Map transition phase. Earth-centered view of LunaH-Map transition from weak lunar capture to elliptical science orbit (image credit: LunaH-Map Team)
An elliptical science orbit with periselene above the lunar south pole will then be maintained with deterministic orbit adjustment maneuvers occurring at apolune. The periselene altitude of each pass above the south pole will be between 5-25 km, enabling productive scientific return. This science orbit will be maintained for at least 282 lunar orbits (46 days), with the possibility of extending the operations and orbital maintenance if desired. The trajectory design was constrained such that maneuvers are aligned in the velocity or anti-velocity direction and may not occur on consecutive orbits, due to operations constraints. This results in approximately 60 orbits requiring deterministic maneuvers to maintain the desired elliptical orbit. The elliptical science orbit is illustrated in Figure 4.
LunaH-Map is manifested on the SLS (Space Launch System) Artemis-1 mission. Upon deployment from SLS, LunaH-Map will use its iodine-ion-thruster to position itself into highly elliptical orbit with the closest point to the Moon over the south pole at a height of between 8 and 20 km. It will take up to 70 days for the spacecraft to be weakly captured by the Moon’s gravity. Then, within about one year, LunaH-Map will achieve an elliptical orbit. At its farthest point in this orbit, the spacecraft will be 3,150 km from the lunar surface (above the north pole), while its nearest point will be 8 km (above the south pole). Once this has been achieved, there will be a minimum two-month science phase when neutron measurements will be made as the spacecraft flies by the south pole during each orbit. The orbital period is about 4 hours and communication with the spacecraft is achieved using NASA's DSN (Deep Space Network). Radio contact and maneuver planning will occur every 3 to 5 days, with science data downlinked back to Earth on each contact. Finally, after the science data have been transmitted back to Earth and most of the fuel reserves consumed, LunaH-Map will crash into the lunar South Pole, perhaps within one of the permanently shadowed regions or near the pole. 7) Teri Crain, Ernest Cisneros, Nathan Cluff,8)
Figure 5: Illustration of the deployed LunaH-Map spacecraft (image credit: Arizona State University)
Table 1: LunaH-Map flight system details
LunaH-Map will use a solid iodine ion propulsion system, X-band radio communications through the NASA DSN (Deep Space Network), star tracker, C&DH and EPS systems from Blue Canyon Technologies of Boulder, CO, solar arrays from MMA Design LLC, of Louisville CO, mission design and navigation by KinetX Inc. of Tempe, Arizona. Spacecraft systems design, integration, qualification, test and mission operations are performed by Arizona State University.
Propulsion System: LunaH-Map is equipped with an EP (Electric Propulsion) system for orbit transfer and station keeping, developed by Busek Co. Inc. of Natick, MA. The BIT-3 (Busek Ion Thruster – 3 cm grid) propulsion system uses solid-storable iodine as propellant, a pioneering technology that can enable a wide variety of deep-space CubeSat missions. The ability to use iodine as propellant is a game-changer for CubeSat propulsion, because iodine can be stored as a dense solid (4.9 g/cc), and its torr-level storage and operating vapor pressure is safe to launch while allowing for very lightweight and conformal tanks. In contrast, legacy EP propellant xenon has to be stored as highly compressed gas (>2,000 psi) and requires bulky, spherical-shaped pressure vessels that are unfavorable for a CubeSat’s unique form factors. Busek’s BIT-3 ion thruster and its complementary cathode neutralizer also hold the distinction of being the world’s first EP device ever to fire on iodine propellant, and its performance has been verified on the ground via direct thrust measurement.
The BIT-3 uses an inductively-coupled plasma (ICP) discharge to eliminate the need for an internal hot cathode and increase overall lifetime. Thruster life is dominated by grid erosion, which by simulation exceeds 20,000 hours. The most unique feature of BIT-3 is its compatibility with iodine propellant, a demonstrated drop-in replacement for xenon in terms of thrust and Isp performance. Iodine stores as a dense solid (>2x storage density than xenon) and eliminates the need for high-pressure tanks. 9) 10)
Figure 6: BIT-3 system layout, gimballed (image credit: Busek)
Table 2: Specifications of the BIT-3 CubeSat flight system
Figure 7: Actual performance of the iodine BIT-3 flight system (image credit: Busek)
Solar array: To meet or exceed the requirements of the LunaH-Map mission, MMA designed an eHaWK™ (Enhanced High Watts Per Kilogram) solar array which leveraged high TRL-8 HaWK design components to produce an estimated 96 W at BOL (Beginning of Life). MMA’s solution can be stowed efficiently in a 2U x 3U volume. The LunaH-Map eHaWK™ incorporates proven cell performance with the XTJ (NeXt Triple Junction) prime cells manufactured by Spectrolab, Inc. These cells have an average efficiency of 30.7%, with an active area of 27.22 cm2. They are protected by a cerium-doped coverglass (127 µm thick) with an anti-reflective coating. Spectrolab measures the electrical performance of each CIC (Cell-Interconnect-Coverglass) using a steady-state solar simulator and this data is used by MMA Design to design each string assembly and provide similar performance across all strings. The LunaH-Map eHaWK™ integrates a discrete bypass diode at the CIC level for reverse bias protection. Each string of 7 cells in series also includes a single blocking diode for string protection.
Figure 8: MMA design LunaH-map solar array. Existing 3-panel wing design (top). Wing and launch restraint scaled for stowed requirements (bottom left). Also showing solar array drive assembly (bottom right). The LunaH-Map eHaWK™ incorporates MMA’s patented CubeSat SADA (Solar Array Drive Assembly). This facilitates higher average orbital power and enables peak power tracking. The SADA features ±180º of actuation, up to 16 signal/power feed-through conductors per wing, and actuation speeds up to 0.188 revolutions per minute (image credit: ASU)
Avionics, Control, Command and Data Handling: BTC (Blue Canyon Technologies) is providing the spacecraft avionics for the LunaH-Map mission. The BCT provided avionics include the LunaH-Map GN&C (Guidance, Navigation, & Control Subsystem), the C&DH (Command & Data Handling Subsystem) and the EPS (Electrical Power Subsystem). BCT is utilizing the heritage XB1-50 avionics suite to provide the GN&C and C&DH subsystems with a heritage Power Subsystem including a battery and power switching board. The BCT XB1-50 avionics provide a complete bus solution in a highly integrated, precision spacecraft platform. The LunaH-Map avionics provides a unified, tested package that leverages BCT’s GN&C expertise and experience to provide solutions to the unique requirements of the LunaH-Map mission.
Figure 9: Exploded View of BCT XB1-50 Avionics Module for LunaH-Map (image credit: BCT)
RF communications: This subsystem for LunaH-Map is based on previously developed JPL hardware for other interplanetary CubeSat missions such as INSPIRE and MarCO. 11) The key component of the telecommunication system is the Iris radio which provides support for telemetry, command and navigation functions. It operates in the Near-Earth X-band frequency range (7145-7190 MHz for uplink, 8450-8500 MHz for downlink), and provides coherent transmission with an 880/749 turn around ratio to support ranging. The uplink modulation is PCM/PSK/PM, NRZ with BCH encoding and data rates ranging from 62.5 bit/s to 8 kbit/s. Specifically, LunaH-Map selected uplink data rates are 62.5 bit/s (safe mode) and 1 kbit/s (nominal mode). The downlink modulation is BPSK with options for RS (255,223), convolutional (K=7, r=1/2) or turbo encoding (1/2,1/3,1/6 with 1784 or 8920 bit frames) and data rates ranging from 62.5 to 256 kbit/s.
LunaH-Map will use 62.5 bit/s (safe mode) and a set of nominal mode data rates ranging from 1 kbit/s to 128 kbit/s. The choice of downlink data rates will depend on the phase of the mission (deployment vs. cruise, vs. science phase), the ground station used (either the 34 m dish at DSN, or the 21 m dish at Morehead State University), and the specific pointing capabilities of the spacecraft during the different mission phases. The encoding solutions implemented will mostly be alternating between Turbo 1/2 and Turbo 1/6, depending again on mission phase and ground station used.
The Iris radio is connected in the receiving path to the LNA (Low Noise Amplifier), and the LNA is then connected to two 6 dBi patch antennas placed on opposite side of the spacecraft for coverage maximization. In the transmitting path, the Iris radio is connected to the SSPA (Solid State Power Amplifier) which provides 2 W of amplification. The 2 W radio frequency signal is then transmitted through other two 6 dBi patch antennas which are placed also on opposite sides of the spacecraft.
In terms of ground support, LunaH-Map will use the services of the DSN (Deep Space Network) and in particular the 34 m antennas at Goldstone, Canberra, Madrid, and the new 21 m antenna at Morehead State University which is now DSS number 17 and part of the NASA DSN complex. Link and coverage analysis have been performed to assess the data rate capabilities at different points of the mission. The data rate capabilities and range variation, when using DSN and when using the 21 m antenna at Morehead State University are shown in Figure 10.
Operations: Mission operations for the LunaH-Map program, including spacecraft command and control, telemetry processing and analysis, science data retrieval, archiving and product generation, and mission planning and scheduling, is performed on-campus at ASU. ASU’s state-of-the-art Mission Operations Center is a multi-mission facility that will also house operations for components of the Mars 2020 rover and Psyche asteroid missions. For downlink operations, telemetry from the spacecraft is transmitted over X-Band via the Iris transponder, and received by the DSN. The DSN forwards the telemetry to ASU over a VPN (Virtual Private Network) connection where it is processed by the JPL-developed AIT (AMMOS Instrument Toolkit) Ground Data System. Throughout each satellite contact, real-time data from the spacecraft is downlinked for graphical trending and analysis for state of health monitoring purposes. Additional stored flight software and spacecraft state of health telemetry is available for immediate analysis by the engineering and science teams, while the science data is labeled and archived following PDS (Planetary Data System) standards. Following the satellite contact, custom data products are generated for the science, engineering, and KinetX mission planning teams. KinetX receives attitude prediction and history files, as well as thruster firing history and a predicted events file; these are used to generate spacecraft ephemeris reconstruction and prediction files, navigation tracking requests, maneuver interface files, and light time files, which are all incorporated in to the mission operations planning cycle. Similarly, the engineering and science teams use their data products to generate calibration or instrument requests to submit to the planning cycle, as necessary.
Table 3: XB1-50 LunaH-map avionics performance
Sensor complement (Mini-NS)
Mini-NS (Miniature Neutron Spectrometer)
Mini-MS uses a set of CLYC scintillators to detect neutrons and has been designed with a gadolinium shield to provide sensitivity primarily to neutrons above 0.5 eV. The Mini-NS consists of two detectors that are comprised of four modules (or sensors), shown in Figure 11 right. The modules consist of a hermetically sealed CLYC (Cs2LiYCl6) scintillator in an aluminum can, and the design provides 200 cm2 of the scintillator’s area facing the Moon. The sensors are encased with a 0.5 mm layer of gadolinium. The readout electronics are separated into two sections, where the analog components are placed near the detector and the digital electronics are further away.
Figure 11: Left: The flight Mini-NS with ruler for scale. Right: A single Mini-NS detector module without PMT (Photomultiplier Tube) mounted. Up to four detector modules can be operated by one electronics board assembly. The Mini-NS consists of 8 detector modules, which operate as two independent 2x2 detector arrays (image credit: ASU)
Figure 12: a) Mechanical structure of the Mini-NS. The near-side panel is drawn as transparent to see into the instrument. b) Outer mechanics are removed showing the eight detector modules and inner readout electronics. c) Flight Mini-NS with ruler for scale (image credit: ASU, Ref. 8)
The digital readout electronics are mounted to the exterior support frame of the instrument, where the heat generated in these electronics is directly dissipated to the spacecraft chassis. The instrument interfaces with the spacecraft via RS422, providing a per second heartbeat to monitor the instrument health. Data products are stored locally on the instrument, as the real time is issued by the spacecraft to co-register the data to the spacecraft trajectory. Event data from each PMT is tagged and individual module values (gain, bias voltage, pulse shape analysis parameters) can be adjusted in flight. Each electronics board can operate up to 4 Mini-NS modules and future missions can customize the number of Mini-NS modules based on the science mission requirements and configure them based on the spacecraft mass, power and volume constraints.
The mass of the flight Mini-NS instrument (8 modules, 2 electronics board assemblies, mechanical structure) is 3.4 kg with an estimated power draw of 9.6 W. In stand-by mode, the spacecraft will interface with the instrument to transfer data or to initiate data acquisition, where the power draw in this state is estimated to be 3.6 W.
Figure 13: Photo of the fully assembled Mini-NS flight unit (image credit: ASU)
Figure 14: Orbit ground track shown for entire 60 (Earth) day science phase: 141 passes over target area initially (and periodically) centered on the Shackleton Crater (-89.9º latitude), with a close-approach of 7 km at each perilune crossing. The yellow circle denotes LunaH-Map altitude of 8 km; the green circle denotes the LunaH-Map altitude of 12 km (image credit: ASU)
Hydrogen Detection Using Neutron Spectroscopy (Ref. 12)
Every planet in our solar system is constantly bombarded by cosmic ray protons. The Moon doesn't have an atmosphere, so these protons constantly bombard the lunar surface, interacting with the regolith (e.g. rocks and soils) to producing neutrons. As these neutrons “leak” out from the surface, they lose energy and are slowed by collisions. Materials that are more enriched in hydrogen (i.e. lunar regolith within permanently shadowed regions) are more effective at moderating neutrons than unenriched regolith. A (over)simplified analogy for thinking about planetary neutron spectroscopy uses colliding balls. Think of a neutron as a ping pong ball. It will lose about half its energy when it hits another ping pong ball. Every time a neutron (ping pong ball) hits a hydrogen atom (another ping pong ball) it loses about half its energy. Collisions between neutrons and every other element in the periodic table are more like a ping pong ball hitting a bowling ball. This results in increased numbers of low energy (thermal) neutrons and decreased numbers of moderate energy (epithermal) neutrons when more hydrogen is present in the lunar regolith.
LunaH-Map will use a miniature neutron spectrometer (Mini-NS) which uses a neutron sensitive material called a scintillator. Scintillators produce a small light flash when neutrons or other particles interact within their structure. Photomultiplier tubes (or other photosensitive, light-catching and amplifying, devices) then amplify the signal and translate the light flashes into neutron counts. The Mini-NS is designed to detect the suppression of only epithermal neutrons. The Mini-NS uses a gadolinium shield to effectively "screen" (absorb) the thermal neutrons, making the detector sensitive to epithermal neutrons which are related primarily to the abundance of hydrogen. When making measurements over the permanently shadowed regions, where hydrogen is enriched, the Mini-NS will detect lower numbers of epithermal neutrons. These "neutron suppressed regions" correspond to hydrogen enrichments, most likely in the form of water (H2O) or hydroxide (OH).
Figure 15: Illustration of lunar surface bombardment of cosmic particles (image credit: ASU)
Launch: The LunaH-Map 6U CubeSat will fly as a secondary payload on the Artemis-1 mission, originally known as the Orion EM-1 (Exploration Mission) of the SLS (Space Launch System), with a planned launch in the second half of 2021. 14)
Overview of secondary payloads on the Artemis-1 mission (formerly Orion/EM-1 mission)
The first flight of NASA’s new rocket, SLS ( Space Launch System), will carry 13 CubeSats/Nanosatellites to test innovative ideas along with an uncrewed Orion spacecraft in 2020. These small satellite secondary payloads will carry science and technology investigations to help pave the way for future human exploration in deep space, including the journey to Mars. SLS’ first flight, referred to as EM-1 (Exploration Mission-1), provides the rare opportunity for these small experiments to reach deep space destinations, as most launch opportunities for CubeSats are limited to low-Earth orbit. 15) 16)
The secondary payloads, 13 CubeSats, were selected through a series of announcements of flight opportunities, a NASA challenge and negotiations with NASA’s international partners.
All the CubeSats will ride to space inside the Orion Stage Adapter, which sits between the ICPS ( Interim Cryogenic Propulsion Stage) and Orion (Figure 16). The cubesats will be deployed following Orion separation from the upper stage and once Orion is a safe distance away.
The SPIE ( Spacecraft and Payload Integration and Evolution) office is located at NASA/MSFC (Marshall Space Flight Center) in Huntsville, Alabama, which handles integration of the secondary payloads.
These small satellites are designed to be efficient and versatile—at no heavier than 14 kg, they are each about the size of a boot box, and do not require any extra power from the rocket to function. The science and technology experiments enabled by these small satellites may enhance our understanding of the deep space environment, expand our knowledge of the moon, and demonstrate technology that could open up possibilities for future missions. 19)
A key requirement imposed on the EM-1 secondary payload developers is that the smallsats do not interfere with Orion, SLS or the primary mission objectives. To meet this requirement, payload developers must take part in a series of safety reviews with the SLS Program’s Spacecraft Payload Integration & Evolution (SPIE) organization, which is responsible for the Block 1 upper stage, adapters and payload integration. In addition to working with payload developers to ensure mission safety, the SLS Program also provides a secondary payload deployment system in the OSA (Orion Space Adapter). The deployment window for the CubeSats will be from the time ICPS disposal maneuver is complete (currently estimated to require about four hours post-launch) to up to 10 days after launch. 20)
Figure 16: The CubeSats will be deployed from the Orion Stage Adapter (image credit: NASA, Ref. 16)
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The information compiled and edited in this article was provided by Herbert J. Kramer from his documentation of: ”Observation of the Earth and Its Environment: Survey of Missions and Sensors” (Springer Verlag) as well as many other sources after the publication of the 4th edition in 2002. - Comments and corrections to this article are always welcome for further updates (firstname.lastname@example.org).