Minimize Lunar IceCube

Lunar IceCube

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Lunar IceCube is a selected NASA nanosatellite mission to prospect, locate, and estimate size and composition of water ice deposits on the Moon for future exploitation by robots or humans.

Lunar IceCube is one of several public-private partnerships chosen under NASA's NextSTEP (Next Space Technologies for Exploration Partnerships) Broad Agency Announcement for the development of advanced exploration systems. Among the first small satellites to explore deep space, Lunar IceCube will help lay a foundation for future small-scale planetary missions, mission scientists said. In addition to providing useful scientific data, Lunar IceCube will help inform NASA's strategy for sending humans farther into the solar system. The ability to search for useful assets can potentially enable astronauts to manufacture fuel and other provisions needed to sustain a crew for a journey to Mars, reducing the amount of fuel and weight that NASA would need to transport from Earth. 1) 2) 3) 4)

MSU (Morehead State University in Kentucky) is leading the 6U CubeSat mission, with significant involvement from scientists and engineers at NASA/GSFC (Goddard Space Flight Center) in Greenbelt, Maryland, and the Massachusetts-based Busek Company.

Under the university-led partnership, MSU's Space Science Center will build the 6U nanosatellite and provide communications and tracking support via its 21 m ground station antenna. Busek will provide the state-of-the-art electric propulsion system and Goddard will construct IceCube's only miniaturized instrument, the Broadband InfraRed Compact High Resolution Explorer Spectrometer (BIRCHES). The instrument will prospect for water in ice, liquid, and vapor forms from a highly inclined elliptical lunar orbit. Goddard also will model a low-thrust trajectory taking the pint-size satellite to lunar orbit with very little propellant.

"Goddard scientists and engineers have deep experience in areas that are critical to interplanetary exploration," said mission Morehead State University Principal Investigator Benjamin Malphrus, explaining why the university teamed with Goddard. "The significant expertise at Goddard, combined with Morehead State's experience in smallsats and Busek's in innovative electric-propulsion systems, create a strong team."

The Lunar IceCube mission science focus, led by the JPL Science PI (Pamela E. Clark), is on:

1) enabling broadband spectral determination of composition and distribution of volatiles in regoliths of the Moon and analogous bodies as a function of time of day, latitude, regolith age and composition;

2) providing geological context by way of spectral determination of major minerals; and

3) thus enabling understanding of current dynamics of volatile sources, sinks, and processes, with implications for evolutionary origin of volatiles. The mission is designed to address NASA HEOMD (Human Exploration and Operations Mission Directorate) Strategic Knowledge Gaps related to lunar volatile distribution (abundance, location, and transportation physics of water ice).

Lunar IceCube amplifies and extends the findings of previous missions. While the M3 (Moon Mineralogy Mapper) on the Chandrayaan-1 mission of ISRO provided a ‘snapshot' mosaic of the lunar nearside, indicating surface coating of OH/H2O near the poles, Lunar Ice Cube will provide coverage of the same swaths as a function of latitude at several times of day. Lunar Ice Cube will extend evidence for diurnal variation in OH absorption provided by Deep Impact and other C-, H-, O-, and S-bearing volatiles provided by LCROSS (Lunar CRater Observation and Sensing Satellite) through geospatial linkage. The Lunar Ice Cube measurements will completely encompass the broad 3 µm band resulting from absorption by several forms of water instead of cutting off at 3 µm as previous NIR spectrometers have done.




Cassini VIMS, Deep Impact

Surface water detection, variable hydration, with noonpeak absorption

Water & other volatiles, fully characterize
3 µm region as function of several times of
day for same swaths over range of latitudes
w/ context of regolith mineralogy and
maturity, radiation and particle exposure,
for correlation w/ previous data.

Chandrayaan M3

H2O and OH (<3 µm) in mineralogical context nearside snapshot at one lunation


Ice, other volatile presence and profile from impact in polar crater


H+ in first meter (LP, LEND) & at surface (LAMP) inferred as ice
abundance via correlation with temperature (DIVINER), PSR and
PFS (LROC, LOLA), H exosphere (LADEE)

Table 1: Lunar IceCube versus Previous Missions (Ref. 3)




Development of the Lunar IceCube deep space CubeSat bus leverages MSU's CubeSat mission experience, utilizes systems with significant flight heritage (in LEO), and incorporates new NanoSat technologies to develop an evolved, radiation-tolerant 6U CubeSat that can support interplanetary investigator science. The 6U Lunar IceCube is based on MSU's bus heritage and incorporates high power generation (120 W of continuous power in cislunar space adapted by Pumpkin Inc. from their lower power Supernova design), a radiation-hardened flight computer (Proton 400K made by Space Micro, Inc.) , a highly-capable micronized GNC (Guidance, Navigation and Control) system and BCT's (Blue Canyon Technologies) XACT attitude determination and control subsystem. Several options were considered for communications including COTS systems and a high throughput X-band communication system designed by JPL for lunar CubeSat missions. The JPL X-band radio, known as Iris, was selected. Iris 2.1 is a full transponder that is capable of high data throughput and has Doppler ranging capabilities.

The 6U Lunar IceCube bus was essentially derived from MSU's successful 2U CXBN (Cosmic X-ray Background NanoSatellite) mission and 1U KySat-2 missions. Additionally, several of the subsystems have successfully flown on numerous NanoSat and MicroSat missions and most of the COTS subsystems will have flight heritage prior to the development of a complete Lunar IceCube flight unit.

The spacecraft design is at a CDR level of maturity; an up-to-date specification overview is provided in Table 1, with Figure 2 and Figure 3 showing the subsystem layout. The combination of flight qualified hardware and innovative solutions to difficult engineering challenges provides for a robust spacecraft bus solution for the Lunar IceCube program.

ADCS (Attitude Determination and Control Subsystem): Attitude control will be provided by the BCT( Blue Canyon Technology) XACT which is an integrated ADCS. This fully integrated system includes star trackers, IMU, and RWAs flight heritage on the MinXSS mission, and can interface with thrusters. Several of the NASA EM-1 CubeSats utilize the BCT XACT.

C&DH (Command and Data Handling): The C&DH selected for Lunar IceCube is the Space Micro Inc. Proton P400K-SGMII-2-PCI104S-SD Space Computer. The Proton400K TM computing platform is a high performance, low power radiation hardened processing solution that meets the challenge of the space environments. This product utilizes the Freescale advanced 45 nm dual-core microprocessor and combined with Space Micro's patent pending radiation mitigation technologies.

RF communications: Communications with the Lunar IceCube are provided by by the X-band JPL Iris Radio and dual patch antennas. MSU has a 21 m dish antenna that is becoming part of the DSN. Anticipated data rate~128 kbit/s with the 34 m DSN antennas and 64 kbit/s with the MSU 21 m Ground Station.

Launch mass

~14 kg

Payload mass capability, volume

3.0 kg, 2.0 U

Pointing accuracy

±.007º (1σ)

Orbit knowledge

10 m, 0.15 m/s

Maneuver rate


Payload power capability

17.78 W

Prime power generated

120 W continuous

Voltages available

28 V, 12 V, 5 V, 3.3 V

Performance of Ion Propulsion System

Nominal power: 70 W (max 80 W)
Nominal thrust: 1.0 mN
Nominal Isp (including neutralizer): 2130 s
Maximum ΔV capability: 2.9 km/s (at max power)
Total impulse capability: 38,800 Ns

Table 2: Performance parameters of the 6U Lunar IceCube nanosatellite


Figure 1: Layout of the Lunar IceCube with sample science payload (image credit: MSU)

Propulsion: The Busek Company of Natik, MA, with sponsorship from NASA, developed a low-thrust electric propulsion system named BIT-3 (Busek Ion Thruster-3 cm grid). The BIT-3 system is capable of delivering variable Isp and thrust of 2,130 s and 1.0 mN, respectively, at the designed power of 70 W. The BIT-3 RF ion thruster is regarded as the world's first gridded ion thruster ever to operate on iodine propellant. 5) 6)


Figure 2: Left: BIT-3 RF ion thruster flight assembly; Right: prototype demonstration firing with iodine propellant (image credit: Busek)

The concept of utilizing iodine (I2) as an EP propellant was identified over a decade ago and patented by the U.S. Air Force. Busek pioneered the practical application of such propellant in 2010 with the development of an iodine-fueled 200 W HET (Hall Effect Thruster) known as BHT-200. This development has led to a patent involving the iodine feed system that was the departure point of all Busek's iodine EP technologies. Since then, Busek has successfully demonstrated iodine-fueled HETs up to 10 kW power and verified that I2 has very similar performance characteristics compared to the legacy EP propellant xenon (Xe).

I2 has many advantages over Xe from the perspective of storage and handling requirement, launch safety, and cost. Table 1 compares the physical properties between these two propellants. I2 appears naturally as solid crystal in diatomic form, with relatively-low dissociation energy at around 1.54 eV. Atomic iodine (I) is only slightly lighter than Xe, and its electron impact ionization cross section is larger and first ionization potential is lower. These characteristics make I2-based plasma just as easy to create and efficient to use for propulsion purpose as with xenon. Scientists used to have reservation regarding I2's dissociation energy cost may reduce an EP device's overall thrust efficiency; this turned out to be a non-issue as reported by previous HET test results. It is true that in a plasma generator some of the I2 molecules may become ionic dimers (I2+) without dissociating, but as James Szabo and Mike Robin point out, these species are relatively low in percentage (<10%) and most of the propellant exit as singly-charged monomer I+. 7)

Figure 3 shows a conceptual flow diagram of an iodine BIT-3 propulsion unit, with both the thruster and the neutralizer operating on a single iodine storage and feed system. During operation, the propellant reservoir is heated to produce iodine vapor via sublimation, which is then fed to the BIT-3 thruster and the BRFC-1 neutralizer (an iodine-compatible RF cathode). The feed lines and flow control valves are heated to above the tank temperature in order to prevent condensation. The vapor flow rate is controlled by the tank heater, with feedback from the tank's temperature sensors (primary), measured ion beam current (secondary), and pressure transducer (tertiary backup). The control scheme works best for long-duration, steady-state maneuvers as the response time between applied heater power and actual I2 flow rate change can be slow.


Figure 3: Iodine BIT-3 system flow diagram (image credit: Busek)

Figure 4 shows the configuration of the BIT-3 flight system currently in development. The system envelope is 180 x 88 x 102 mm (~1.6U) with a wet mass of 3 kg. In addition to the thruster and the neutralizer, other notable subsystem components include a 1.5 kg-capacity iodine storage tank, a custom 2-axis gimbal integrated to the thruster body, a highly miniaturized and radiation-tolerant PPU (Power Processing Unit), and COTS (Commercial-of-the-Shelves) pressure transducer and iodine-compatible control valves. The 2-axis gimbal utilizes flight-grade stepper motors and is capable of ±10º actuation and 0.1º accuracy. The gimbal enables the single primary thruster to de-saturate the RWA (Reaction Wheel Assembly) while in deep space. It is also capable to balance out any C.G. shift in-flight as the propellant is consumed.

The expected performance of the iodine BIT-3 system is currently based on actual thrust and Isp data obtained with I2 propellant, with some estimation on the heater power consumption and FM PPU efficiency. The flight BIT-3 system will be throttleable between 56 and 80 W at the PPU input and capable of accepting 28-37 V unregulated voltage and RS-485 command. Expected thrust and Isp within this power range are 0.66-1.24 mN and 1,400-2,640 s, respectively. The 80 W max power is not outrageous for 6U CubeSats, as current SOA (State-of-the-Art) solar array for 6U can generate up to 120 W prime power at 1 AU distance from the Sun. When given sufficient power, the iodine BIT-3 system can provide a 6U/14 kg CubeSat an unprecedented ΔV capability of 2.9 km/s. The BIT-3 system has also been recognized as having the highest volumetric impulse (total impulse per unit volume) among all SOA CubeSat propulsion options.

In summary, the iodine BIT-3 propulsion system offers extremely high Isp (>2,000 s) heretofore unavailable to nanosatellites. This type of propulsion technology would enable a variety of deep space CubeSat missions including lunar reconnaissance, asteroid rendezvous, inner planet flyby, as well as low flying Earth observation.


Figure 4: The 1.6U volume, 3 kg wet, iodine BIT-3 flight system configuration (image credit: Busek)


Figure 5: Artist's rendition of the Lunar IceCube spacecraft in lunar orbit (image credit: Morehead State University, NASA)


Launch: Lunar IceCube is a selected NASA nanosatellite mission to prospect, locate, and estimate size and composition of water ice deposits on the Moon for future exploitation by robots or humans. It will fly as a secondary payload mission on the first flight of the SLS (Space Launch System), EM-1 (Exploration Mission-1) scheduled to launch in 2018.

The satellites selected for EM-1 will be installed inside the adapter, which connects Orion to the upper stage of NASA's newest rocket — the SLS launch vehicle designed to ferry humans and gear around the moon and beyond. Once the rocket reaches a certain position on its way to the moon, ground controllers will send a command to release the payloads, which will follow their own trajectories to their final destinations in and around the moon.

Busek's RF Ion BIT-3 thruster, along with a carefully designed trajectory modeled by Goddard's state-of-the-art trajectory-design software, will get IceCube to its destination in about three months, according to Dave Folta, the Goddard orbital engineer who has developed advanced tools for modeling lunar orbits for spacecraft equipped with both chemical and low-thrust propulsion systems.

The journey of the Lunar IceCube will begin after deployment— and will be another challenge given the miniscule real estate set aside for propellant. Ground controllers will fire Busek's miniaturized electric thrusters — the world's only propulsion system powered with an iodine propellant — driving the spacecraft along a complex path that uses the gravity of the sun, Earth and moon, looping around Earth a couple times and then to its destination. Because the thrusters operate electrically using small amounts of propellant, an orbital path that takes advantage of gravitational acceleration from the Earth and moon is vital. This approach utilizes the interplanetary superhighway- a low energy trajectory using gravitational manifolds.


Figure 6: Getting to the moon will require that the Lunar IceCube take a circuitous route that uses the gravity of the Sun, Earth and Moon (image credit: NASA, Dave Folta)



Sensor complement (BIRCHES)

BIRCHES (Broadband InfraRed Compact, High-resolution Exploration Spectrometer, a miniaturized version of OVIRS on OSIRIS-REx)

BIRCHES is a compact (1.5U, 2.5 kg, 10-15 W including cryocooler) point spectrometer with a compact cryocooled HgCdTe focal plane array for broadband (1 to 4 µm) measurements, achieving sufficient SNR (>400) and spectral resolution (10 nm) through the use of a Linear Variable Filter to characterize and distinguish important volatiles (water, H2S, NH3, CO2, CH4, OH, organics) and mineral bands. The instrument has built-in flexibility, using an adjustable 4-sided iris, to maintain the same spot size regardless of variations in altitude (by up to a factor of 5) or to vary spot size at a given altitude, as the application requires. Compact instrument electronics are also being developed which can be easily reconfigured to support the instrument in ‘imager' mode, once the communication downlink bandwidth becomes available, and the H1RG family of focal plane arrays.

Thermal design is critical for the instrument. The compact and efficient Ricor cryocooler is designed to maintain the detector temperature below 120 K. In order to maintain the optical system below 220 K, a special radiator is dedicated to optics alone, in addition to a smaller radiator to maintain a nominal environment for spacecraft electronics.

The BIRCHES instrument has a size of 10 x 10 x 15 mm. The power subsystem requirement of 10-15 W includes the 3 W detector electronics, a 1.5 W AFS controller, and the 5-10 W cryocooler.

BIRCHES observation requirements (Ref. 3):

• A footprint of 10 km from a lunar altitude of 100 km. Footprint of 10 km in along-track direction regardless of altitude, larger in cross-track direction above 250 km

• Nyquist sampling of the lunar surface

• FOV of the instrument is 100 mrad (6º)

• An AFS (Adjustable Field Stop) shall maintain the FOV to 100 km in size

• Based on spacecraft velocity, exposes shall be taken at intervals of 2.7 s (TBC).



Wavelength (µm)


Water vapor



OH stretch
OH stretch

Liquid water


H-OH fundamental
H-OH fundamental



OH stretch overtone
HOH bend overtone



M3 Feature
Total H2O

Hydroxyl ion


OH stretch (mineral)
OH (surface or structural) stretches



Cation-OH bend
Structural OH

Bound H2O


Houck et al. (Mars)
H2O of hydration
H2O stretch (Mars)



Feature w/2.95

Adsorbed H2O


R. Clark



Band depth-layer correlated
Strong feature



Pieters et al.

Other volatiles


1.65, 2, 2.2

N-H stretch


2, 2.7

C-O vibration and overtones





1.2, 1.7, 2.3, 3.3

C-H stretch fundamental and overtones

Mineral Bands



crystal field effects, charge transfer


1, 2, 2.9

crystal field effects



crystal field effects

Iron oxides


crystal field effects


2.35, 2.5

overtone bands



conduction bands

Hydrated silicates


vibrational processes

Anticipate wavelength of peak for water absorption band to be structural<bound<adsorbed<ice

Yellow = water-related features in the 3 µm region

Table 3: IceCube measurements will encompass the broad 3 µm band to distinguish overlapping OH, water, and ice features. Will have near 10 nm resolution in this band.


Figure 7: Block diagram of the BIRCHES instrument (image credit: NASA)

BIRCHES utilizes a compact Teledyne H1RG HgCdTe FPA (Focal Plane Array) and a JDSU (formerly JDS Uniphase Corporation) of Milpitas, CA, LVF (Linear Variable Filter) detector assembly leveraging OSIRIS-REx OVIRS.


Figure 8: COTS AFRL developed AIM SX030 microcryocooler with cold finger to maintain detector at ≤ 115 K and IRIS controller (image credit: NASA)


Figure 9: Illustration of the BIRCHES instrument (image credit: NASA)


1) "Lunar IceCube to Take on Big Mission From Small Package," NASA, August 4, 2015 (updated on Jan. 29, 2016), URL:

2) Michael Tsay, John Frongillo, Kurt Hohman, Benjamin K. Malphrus, "Linar Cube: A Deep Space 6U CubeSat with Mission Enabling Ion Propulsion Technology," Proceedings of the 29th Annual AIAA/USU Conference on Small Satellites, Logan, Utah, USA, August 8-13, 2015, paper: SSC15-XI-1, URL:

3) P. E. Clark, Ben Malphrus, Kevin Brown, Dennis Reuter, Robert MacDowall, David Folta, Avi Mandell, Terry Hurford, Cliff Brambora, Deepak Patel, Stuart Banks, William Farrell, Noah Petro, Michael Tsay, V. Hruby, Carl Brandon, Peter Chapin, "Lunar Ice Cube Mission: Determining Lunar Water Dynamics with a First Generation Deep Space CubeSat," 47th Lunar and Planetary Science Conference (2016), The Woodlands, Texas, March 21-25, 2016, paper: 1043.pdf, URL:

4) Pamela E. Clark, Ben Malphrus, D. Reuter, T. Hurford, R. MacDowall, N. Petro, W. Farrell, C. Brambora, D. Patel, S. Banks, P. Coulter, D. Folta, P. Calhoun, B. Twiggs, Jeff Kruth, Kevin Brown, R. McNeill, M. Tsay, V. Hruby, Carl Brandon, Peter Chapin, "Broadband InfraRedCompact High-resolution Exploration Spectrometer: Lunar Volatile Dynamics for the Lunar Ice Cube Mission," Proceedings of the 30th Annual AIAA/USU SmallSat Conference, Logan UT, USA, August 6-11, 2016, URL:

5) Michael Tsay, John Frongillo, Kurt Hohman, "Iodine-Fueled Mini RF Ion Thruster for CubeSat Applications," Proceedings of the 34th International Electric Propulsion Conference (IEPC), Hyogo-Kobe, Japan, July 4-10, 2015, paper: IEPC-2015-273, URL:

6) Michael Tsay, John Frongillo, Joshua Model, Jurg Zwahlen,Lenny Paritsky, "Flight Development of Iodine BIT-3 RF Ion Propulsion System for SLS EM-1 CubeSats," Proceedings of the 30th Annual AIAA/USU SmallSat Conference, Logan UT, USA, August 6-11, 2016, URL:

7) James Szabo,Mike Robin, "Plasma Species Measurements in the Plume of an Iodine Fueled Hall Thruster," Journal of Propulsion and Power, Vol. 30, No. 5, Sept.2014, pp. 1357-1367,

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 (

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