Minimize LRO

LRO (Lunar Reconnaissance Orbiter) + LCROSS

LRO is a NASA mission to the moon within the Lunar Precursor and Robotic Program (LPRP) in preparation for future manned missions to the moon and beyond (Mars). LRO is the first mission of NASA's `New Vision for Space Exploration', which President Bush announced on January 14, 2004, in sending more robot and human explorers beyond Earth orbit. The LRO requirements call for a mission life of one year in lunar orbit. The objectives of LRO are to: 1) 2) 3) 4) 5)

• Identify potential lunar resources

• Gather detailed maps of the lunar surface

• Collect data on the moon's radiation levels

• Study the moons polar regions for resources that could be used in future manned missions or robotic sample return missions

• Provide measurements to characterize future robotic explorers, human lunar landing sites and to derive measurements that can be used directly in support of future Lunar Human Exploration Systems.

The orbiter project is managed by NASA/GSFC while NASA/ARC manages the LRO payload. The CDR (Critical Design Review) of LRO was completed in Nov. 2006.


Figure 1: Artist's view of LRO spacecraft (image credit: NASA)


The spacecraft is being built and integrated at NASA/GSFC (inhouse development), Greenbelt, MD. The spacecraft architecture emphasizes modularity through the use of standard interfaces. LRO is a 3-axis stabilized, nadir pointed spacecraft designed to operate continuously during the primary mission.

The ACS (Attitude Control Subsystem) consists of the following components: 10 CSS (Coarse Sun Sensors), 4 RW (Reaction Wheels), 2 A-STR (Autonomous Star Trackers), and a RLG (Ring Laser Gyroscope) of Honeywell, referred to as MIMU (Miniature Inertial Measurement Unit). MIMU provides attitude rate information up to 18 º/s and attitude rate polarity from 18º/s up to 60 rpm. The reaction wheels are specifically designed to provide very quiet, smooth changes in pointing of the spacecraft. Once in observing mode, the reaction wheels keep the boresight of the instruments pointing continuously at the surface of the moon. The 2 A-STR, built by SELEX Galileo, provide a spacecraft attitude quaternion in the J2000 ECI (Earth Centered Inertial) reference frame.

The ACS hardware is controlled by ACS flight software (FSW) resident on the SBC (Single Board Computer). This software also includes some FDC (Failure Detection and Correction) algorithms used in safing. Part of the ACS FSW function is to provide commands to the SA (Solar Array) and the HGA (High Gain Antenna). Attitude and momentum control functions are performed in ACS control modes that process sensor data and generate appropriate actuator commands. 6) 7)


Figure 2: Photo of the MIMU device (image credit: NASA)

EPS (Electric Power Subsystem): The EPS is comprised of an articulated solar array (1 wing, 2-axis tracking), a Li-ion battery of ABSL (UK), and a DET system (21-35 V). Battery mass of 35 kg, and a capacity of 126 Ah.

The C&DH (Command and Data Handling) subsystem comprises a radiation hardened SBC (Single Board Computer) for flight software, telemetry and command handling functions, system clock, and interfaces to all instruments. Data storage is provided by four DSB (Data Storage Board) devices. The onboard system architecture uses the SpaceWire bus in support of high-speed interfaces (LROC, Mini-RF, HK, UART, and LAMP), while the MIL-STD-1553B low-speed bus is used for LEND, DLRE, CRaTER, LOLA, ACS (Attitude Control Subsystem), PSE (Power System Electronics), and the propulsion subsystem.

Mass budget

1846 kg (total), dry mass of 949 kg,, fuel mass of 897 kg

Articulated solar array

Tri-panel solar array assembly of size 10.7 m2

Orbit power, power storage

824 W (average), 1.5 kW (peak)

Power storage

Li-ion battery, 84 parallel strings of 8 cells each, capacity of > 80 Ah (BOL)

Data volume, max downlink data rate

572 Gbit/day, 100 Mbit/s

Spacecraft pointing accuracy, knowledge

60 arcsec, 30 arcsec

Flight computer

RAD-750 processor executing at 133 MHz

C&DH (Command and Data Handling) subsystem

MIL-STD-1553, RS 422, & High Speed Serial Service

Onboard data storage

Two 100 GByte recorders for science data playback to Earth


4 insertion thrusters deliver a total force of ~350 N

Table 1: Overview of LRO spacecraft parameters

A one year primary mission is planned in ~50 km polar orbit, possible extended mission in communication relay/south pole observing, low-maintenance orbit. 8)


Figure 3: Block diagram of the LRO spacecraft (image credit: NASA/GSFC)

The spacecraft payload includes seven instruments, two of which are connected to the Command and Data Handling (C&DH) unit developed by NASA Goddard Space Flight Center via the SpaceWire network, as shown in Figure 4. The Mini-RF instrument is connected to the SpaceWire network through the SpaceWire ASIC. of BAE Systems (Manassas, VA). Within the C&DH unit, the RAD750 flight computer communicates with the instruments and other boards via three interfaces: a four-port SpaceWire router, a 32 bit, 33 MHz PCI bus, and a redundant MIL-STD-1553 bus. 9) 10) 11)

The SpaceWire router is implemented in the SpaceWire ASIC that is in turn connected to the RAD750 microprocessor via the PCI bus and the second generation enhanced Power PCI bridge ASIC. Both the Ka-band and S-band communications boards include SpaceWire interfaces with routers, implemented in Actel FPGAs. The LROC instrument is connected directly to one of the processor board’s SpaceWire links, while the Mini-RF connects to the Housekeeping and Input/Output (HK/IO) board that also implements a SpaceWire router using an Actel FPGA and is then routed to the processor board across the SpaceWire bus via the FPGA that implements another SpaceWire router. The 400 Gb LRO mass memory is implemented in synchronous DRAM that is interfaced to the RAD750 computer via the PCI bus on a custom C&DH backplane.


Figure 4: Block diagram of the C&DH subsystem and interfaces (image credit: NASA, BAE Systems)

The RAD750 SBC (Single Board Computer) is a Compact PCI 6U-220 card with two printed wiring boards (PWBs). The RAD750 microprocessor operates at 132 MHz with a 66 MHz bus to I/O and memory, both of which are accessed through the enhanced Power PCI bridge ASIC. A total of 36 MB of radiation hardened SRAM is available to the RAD750, along with 4 MB of EEPROM and 64 KB of Start-up ROM, all provided with additional bits for error correction code (ECC).

The SpaceWire ASIC, shown in Figure 3 is based on BAE Systems' reusable core architecture. Its primary function is to perform routing of data using the SpaceWire protocol via a router with four external links and two internal connections to the SoC (System-on-a-Chip) bus, the standard cross-bar switch connection medium of the reusable core architecture. The SpaceWire ASIC software is included in the RAD750 Board Support Package (BSP). The BSP is designed for operation with the VxWorks (versions 5.4, 5.5, and 6.2) RTOS (Real Time Operating System).

Two 16 kB blocks of on-chip scratchpad memory are provided, as well as a 32-bit RISC processor called the EMC (Embedded Microcontroller). The EMC performs housekeeping functions as well as providing support for the SpaceWire router. A PLL (Phase Locked Loop) is provided for the SpaceWire link interface, which is capable of 280 MHz operation.


Figure 5: Illustration of the SpaceWire ASIC (image credit: BAE Systems)

Software support for the CCSDS (Consultative Committee for Space Data Systems) File Delivery Protocol (CDFP) is split between the RAD750 CPU and the EMC within the SpaceWire ASIC. The function of the software executing within the SpaceWire ASIC is to assist in CDFP download to maximize downlink throughput by sending batches of CDFP data packets, known as Protocol Data Units (PDU), over the SpaceWire interface to the Ka-band communications link.


Figure 6: Design concept of the LRO spacecraft (image credit: NASA)


Figure 7: Instrument locations of deployed spacecraft (image credit: NASA)


Launch: The LRO and the companion LCROSS spacecraft were launched on June 18, 2009 on an Atlas V 401 launch vehicle from the Air Force Station at Cape Canaveral, FLA. LRO safely separated from LCROSS 45 minutes after launch. 12) 13)

LCROSS then was powered-up, and the mission operations team at NASA's Ames Research Center at Moffett Field, CA, performed system checks that confirmed the spacecraft is fully functional. LCROSS and its attached Centaur upper stage rocket separately impacted on the moon on Oct. 9, 2009, creating a pair of debris plumes that will be analyzed for the presence of water ice or water vapor, hydrocarbons and hydrated materials. The spacecraft and Centaur are tentatively targeted to impact the moon's south pole near the Cabeus region. The exact target crater will be identified 30 days before impact, after considering information collected by LRO, other spacecraft orbiting the moon, and observatories on Earth.

Orbit: Direct insertion orbit of LRO to the moon. 14)

• Minimum energy lunar transfer orbit (~ 4 days). The launch vehicle will inject LRO into a cis-lunar transfer orbit.

• Lunar orbit insertion sequence (4 maneuvers, 2-4 days, use of onboard propulsion system)

• Commissioning phase in lunar orbit: altitude of 30 km x 216 km, quasi-frozen orbit, up to 60 days

• Polar mapping phase for a duration of at least 1 year. The LRO orbit is nominally 50 km circular and polar, with a period of ~ 113 minutes. The orbital velocity is 1.6 km/s. LRO stays on near side of moon ~ 1 hour out of every two.


Figure 8: Illustration of lunar insertion orbit (image credit: NASA)

Viewing conditions of LRO from Earth for operations support:

• Communication/ranging (SLR) with the LRO spacecraft is possible during the near-side orbital phase of the moon

• Twice a month, LRO's orbit will be in full view of the Earth for roughly 2 days

• Twice a month, LRO will perform a momentum management maneuver while the ground has complete coverage

• Once a month, LRO will perform a station-keeping maneuver while the ground has complete coverage

• Twice a year, LRO's orbit will be in full view of the sun for roughly one month

• During the eclipse season, LRO will have a maximum lunar occultation of 48 minutes

• Twice a year, LRO will perform a 180º yaw maneuver

• Twice a year, the moon will pass through the Earth's shadow (lunar eclipse).

RF-communications: 15)

The S-band is used for TT&C (Telemetry, Tracking & Command) data. Proximity relay is planned to enable mission cross-support at S-band.

- Frequency: Transmit: 2271.2 MHz ±2.5 MHz; Receive: 2091.3967 MHz ±2.5 MHz

- Modulated RF (at the transponder output): 39.1±0.3 dBm (7.58 -8.71 W)

- Acquisition threshold: -121 dBm (receiver ON at all times)

- Modulation: BPSK

• Ranging:

- Coherent downlink ranging generation

- Compatible with STDN and DSN ranging modes

- 1.7 MHz downlink subcarrier; 16 kHz uplink subcarrier

• Data rates in S-band: Uplink:4 kbps uplink capability; Downlink: 0.125, 2, 16, 32, 64, 128, 256 kbit/s (with/without ranging) and 1.093 Mbit/s (direct modulation)

• Control and status interface: UART (Universal Asynchronous Receive Transmit) serial port

• DC Power: ≤ 45 W (full mode): ≤ 10 W (receive only)

The Ka-band is used for the downlink of instrument data (40 W transmitter and high-gain antenna).

- Frequency: 25.65 GHz

- Bandwidth: 300 MHz (±150 MHz)

- Modulation: OQPSK

- DC power: ≤ 30 W

- I/Q channel data inputs: LVDS interface, I and Q staggered by half of a symbol bit

• Symbol rate inputs (after rate ½, K=7 Convolutional and R/S encoding done by C&DH):

- 228.7 Msps (Mega samples per second, normal operations)

- 114.3 Msps or 57.2 Msps (contingency operations)


Figure 9: Schematic view of the projected LRO mossion timelime (image credit: NASA) 16)



LRO mission status:

• January 2014: With precise timing, the camera aboard NASA's Lunar Reconnaissance Orbiter (LRO) was able to take a picture of NASA's LADEE (Lunar Atmosphere and Dust Environment Explorer) spacecraft as it orbited our nearest celestial neighbor. The LROC (Lunar Reconnaissance Orbiter Camera) operations team worked with its LADEE and LRO operations counterparts to make the imaging possible. LADEE is in an equatorial orbit (east-to-west) while LRO is in a polar orbit (south-to-north). The two spacecraft are occasionally very close and on Jan. 15, 2014, the two came within 9 km of each other. Since LROC is a pushbroom imager, it builds up an image one line at a time, thus catching a target as small and fast as LADEE is tricky! Both spacecraft are orbiting the Moon with velocities near 1600 m/s, so timing and pointing of LRO needs to be nearly perfect to capture LADEE in an LROC image. 17) 18)


Figure 10: LRO imaged LADEE as it passed ~9 km beneath it, on Jan. 15, 2014 (image credit: NASA/GSFC, Arizona State University)

• In November 2013, new findings of the CRaTER (Cosmic Ray Telescope for the Effects of Radiation) instrument project at UNH (University of New Hampshire) were published in the journal Space Weather, documenting the different effects and instrument responses with some of the best long-term measurements ever made of radiation in deep space. 19) 20)

Radiation in deep space comes from cosmic rays, from the solar wind and from solar energetic particles emanated during a solar storm. Particles from these sources rocket through space. Many can pass right through matter, such as our bodies. So-called ionizing radiation knocks electrons off of atoms within our bodies, creating highly reactive ions. Within Earth's protective atmosphere and magnetic field, we receive low doses of background radiation every day. The radiation hazards astronauts face are serious, yet manageable thanks to research endeavors such as the CRaTER instrument.

CRaTER measures realistic human radiation doses at the moon using a unique material called TEP (Tissue-Equivalent Plastic). Two pieces of this plastic, roughly 5 cm and 2.5 cm thick, respectively, are separated by silicon radiation detectors. The TEP-detector combo measures how much radiation may actually reach human organs, which may be less than the amount that reaches the spacecraft.

The LRO spacecraft launched as an exploration mission, a forerunner for humanity's return to the moon. But after completing its primary mission in 2010, LRO has become a powerful instrument for lunar and planetary science. CRaTER is an active participant in this scientific study, discovering a previously unmeasured source of hazardous radiation emanating from the moon itself.

This radiation comes from the partial reflection, also called an albedo, of galactic cosmic rays off the moon's surface. Galactic cosmic ray protons penetrate as much as a meter into the lunar surface, bombarding the material within and creating a spray of secondary radiation and a mix of high-energy particles that flies back out into space. This galactic cosmic ray albedo, which may interact differently with various chemical structures, could provide another method to remotely map the minerals present at the moon's surface.

According to the study, CRaTER directly measured the proton component of the moon's radiation albedo for the first time. The TEP radiation detector measures various components of radiation separately, which enables CRaTER to unfold the energy spectrum of the radiation albedo.

• Sept. 25, 2013: Repeat imaging of anthropogenic (human-made) targets on the Moon remains a Lunar Reconnaissance Orbiter Camera (LROC) priority as the LRO Extended Science Mission continues. These continuing observations of historic hardware and impact craters are not just interesting from a historical standpoint - each image adds to our knowledge of lunar science and engineering, particularly cartography, geology, and photometry. 21)


Figure 11: Luna 17, the Soviet Union spacecraft that carried the Lunokhod 1 rover to the surface. One can make out the rover tracks around the lander. LROC NAC image M175502049RE (image credit: NASA/GSFC, Arizona State University)

• On June 18, 2013, NASA's LRO (Lunar Reconnaissance Orbiter) was 4 years on orbit. Not only has LRO delivered all the information that is needed for future human and robotic explorers, but it has also revealed that the moon is a more complex and dynamic world than was expected. 22)

• January 2013: As part of the first demonstration of laser communication with a satellite orbiting the moon, scientists with NASA's LRO beamed an image of the Mona Lisa to the spacecraft from Earth. 23)

The image was transmitted in digital form from the NGSLR (Next Generation Satellite Laser Ranging) station at NASA/GSFC in Greenbelt, MD, to the LOLA (Lunar Orbiter Laser Altimeter) instrument on the spacecraft. By transmitting the image piggyback on laser pulses that are routinely sent to track LOLA's position, the team achieved simultaneous laser communication and tracking. This is the first time anyone has achieved one-way laser communication at planetary distances.

Precise timing was the key to transmitting the image. The Mona Lisa image was divided into an array of 152 x 200 pixels. Every pixel was converted into a shade of gray, represented by a number between zero and 4,095. Each pixel was transmitted by a laser pulse, with the pulse being fired in one of 4,096 possible time slots during a brief time window allotted for laser tracking. The complete image was transmitted at a data rate of about 300 bit/s (use of Reed-Solomon coding). LOLA reconstructed the image in the order the pixels were transmitted. The image was then sent back to Earth using radio waves.

This pathfinding achievement sets the stage for the LLCD (Lunar Laser Communications Demonstration), a high data rate laser-communication demonstrations that will be a central feature of NASA's next moon mission, the LADEE (Lunar Atmosphere and Dust Environment Explorer), a launch of LADEE is scheduled for the fall 2013.

The science objectives for the LRO Science Mission addressed five specific themes.

1) The bombardment history of the Moon.
Developed an improved understanding of the ancient impactor populations that affected all the planets in the inner Solar System, through analysis of global high-resolution topography. Improved the age dating of landforms by using crater counts from the new high-resolution images with Sun angles and illumination geometry optimized for morphology.

2) Lunar geologic processes.
Discovered the global population of small-scale, relatively young compressional structures that show the Moon is in a general state of relatively recent contraction. Characterized volcanic complexes, such as Ina, which appear to result from inflated lava flows.

3) The processes that have shaped the global lunar regolith.
Determined the global distribution of regolith surface temperature and rock abundance. Discovered that impact melt occurs as pools in lunar craters as small as 170 m, and that rough subsurface melt may extend beyond the visible surface melt.

4) Characterization of the volatiles on the Moon with emphasis on the polar regions.
Discovered significant subsurface hydrogen deposits in sunlit areas as well as in some, but not all, permanently shadowed regions.
Measured surprising amounts of several volatiles (e.g. CO, H2, and Hg) in the gaseous cloud released from Cabeus by the LCROSS impact.

5) The Moon’s interaction with its external space environment.
Measured galactic cosmic ray interactions with the Moon during a period with the largest space age cosmic ray intensities. Created the first cosmic ray albedo proton map of the Moon.

Table 2: Summary pf LRO science objectives and accomplishments as of the end of 2012 24)

• On Dec. 17, 2012, the LRO spacecraft witnessed the impact of the two GRAIL (Gravity Recovery and Interior Laboratory) spacecraft of NASA when they were intentionally crashed into a mountain near the moon's north pole. With just three weeks notice, the LRO team scrambled to get LRO in the right place at the right time to witness GRAIL's fiery finale. 25)

• August 2012: Scientists using the LAMP (Lyman Alpha Mapping Project) spectrometer on LRO, have made the first spectroscopic observations of the noble gas helium in the very rare atmosphere surrounding the moon. LAMP uses a novel method to peer into the perpetual darkness of the moon's so-called permanently shadowed regions. LAMP "sees" the lunar surface using the ultraviolet light from nearby space and stars, which bathes all bodies in space in a soft glow of ultraviolet light. -These remote-sensing observations complement in situ measurements taken in 1972 by the LACE (Lunar Atmosphere Composition Experiment) instrument deployed by Apollo 17. 26)

Although designed to map the lunar surface, the LAMP team expanded its science investigation to examine the far ultraviolet emissions visible in the tenuous atmosphere above the lunar surface, detecting helium over a campaign spanning more than 50 orbits. Because helium also resides in the interplanetary background, several techniques were applied to remove signal contributions from the background helium and determine the amount of helium native to the moon.

LRO was launched on June 18, 2009

- Spacecraft and instruments commissioned in a 30 x 200 km elliptical orbit

Exploration Mission: 9/16/2009 - 9/16/2010

- a one-year mapping of the Moon to search for resources, identify safe landing sites, and measure the radiation environment

- quasi-circular polar orbit (50 ± 15km)

Science Mission: 9/17/2010 - 9/16/2012

- more flexible operations for Planetary Science objectives

- quasi-circular orbit (50 ± 15km) until December, 2012

- in the summer of 2012 in a 30 x 200 km orbit

- through June 2012, LRO has delivered 325 TB of data into the PDS (Planetary Data System)

Extended Science Mission: 9/17/2012 - 9/16/2014 (proposed)

Table 3: Overview of mission phases: LRO flexible mission operations enabled new discoveries 27)


Figure 12: LRO orbit during various mission phase (image credit: NASA, Ref. 27)

• June 2012: According to data from NASA's LRO mission, ice may make up as much as 22% of the surface material in the Shackleton crater located near the Moon's south pole. The huge crater, named after the Antarctic explorer Ernest Shackleton, is ~ 4 km deep and 21.5 km in diameter. The small tilt of the lunar spin axis means Shackleton's interior is permanently dark and very cold. Researchers have long thought that ice might collect there. 28) 29) 30)

When a team of NASA and university scientists used LOLA (Lunar Orbiter Laser Altimeter) data to examine the floor of the Shackleton crater, they found it to be brighter than the floors of other nearby craters around the South Pole. This is consistent with the presence of small amounts of reflective ice preserved by cold and darkness.


Figure 13: Elevation (left) and shaded relief (right) image of the Shackleton Crater (image credit: NASA, MIT)

Legend to Figure 13: The structure of the crater's interior was revealed by a digital elevation model constructed from over 5 million elevation measurements from LOLA.

The LRO team was able to map the crater’s elevations and brightness in extreme detail, thanks in part to LRO’s path: The spacecraft orbits the moon from pole to pole as the moon rotates underneath. With each orbit, LOLA maps a different slice of the moon, with each slice containing measurements of both poles. The upshot is that any terrain at the poles — Shackleton crater in particular — is densely recorded. Zuber and her colleagues took advantage of the spacecraft’s orbit to obtain more than 5 million measurements of the polar crater from more than 5,000 orbital tracks. 31) 32)


Figure 14: High-resolution map showing the topography in the interior of the Shackleton crater as observed by LOLA (image credit: MIT, NASA)

Legend to Figure 14: The contours of elevation are plotted every 5 m. The colors show relative elevation with purple lowest and yellow highest. The crater is 4.1 km deep. The spatial resolution of the topography is 10 m and the radial accuracy is < 1 m.

• The LRO spacecraft and its payload are operating nominally in 2012. The mission has been successfully transitioned from NASA's ESMD (Exploration Systems Mission Directorate), it will continue to perform science and measurements throughout 2012.

- New images acquired by the LRO spacecraft show that the Moon’s crust is pulling apart – at least in some small areas. The high-resolution images from LROC (Lunar Reconnaissance Orbiter Camera) provide evidence that the Moon has experienced relatively recent geologic activity (Figure 15). A team of researchers discovered small, narrow trenches typically only hundreds to a few thousand meters long and tens to hundreds of meters wide, indicating the lunar crust is being pulled apart at these locations. These linear troughs or valleys, known as “graben”, are formed when crust is stretched, breaks and drops down along two bounding faults. A handful of these graben systems have been found across the lunar surface. The team proposes that the geologic activity that created the graben occurred less than 50 million years ago (very recent compared to the Moon’s current age of over 4.5 billion years). 33) 34)

- On March 14, 2012, LRO was 1000 days in lunar orbit. 35)


Figure 15: Newly detected series of narrow linear troughs are known as graben, and they formed in highland materials on the lunar farside (image credit: NASA/GSFC, Arizona State University, Smithsonian Institution)

Legend to Figure 15: The graben are located on a topographic rise with several hundred meters of relief revealed in topography derived from LROC stereo images.

• In November 2011, the science team of NASA's LRO mission released the highest resolution near-global topographic map of the moon's farside ever created. 36) 37) 38)


Figure 16: LROC WAC color shaded relief of the lunar farside (image credit: NASA/GSFC, DLR, ASU)

Legend to Figure 16: This new topographic map, from Arizona State University in Tempe, shows the surface shape and features over nearly the entire moon with a pixel scale close to 100 m. Called the Global Lunar DTM 100 m topographic model (GLD100), this map was created based on data acquired by LRO's WAC (Wide Angle Camera), which is part of the LROC imaging system.

During the period Aug. 10 to Sept. 6, 2011, the LRO orbit was adjusted by making it more elliptical. Two station-keeping maneuvers on Aug. 10 lowered LRO from its nominal altitude of ~50 km to an altitude that dipped as low as nearly 21 km (periapsis) as it passed over the moon's surface.

The spacecraft remained in this elliptical orbit for 28 days, long enough for the moon to completely rotate. This allowed full coverage of the surface by LROC's WAC (Wide Angle Camera). At the end of the cycle (Sept. 6, 2011), the spacecraft returned to its nominal 50 ±15 km near-circular orbit with another set of station-keeping maneuvers.

The main goal of the low-orbit lunar month was to obtain WAC coverage of the nearside mare at an average pixel scale of 50 m (or better). At the lowest altitudes, the LROC WAC pixel size decreases to < 40 m from the usual 75 m. High resolution imagery was also obtained by the NAC (Narrow Angle Camera) of LROC.

The higher resolution imagery permitted also to examine three Apollo landing sites (Apollo 12, 14 and 17).

Nominally the LRO orbits the Moon in a 50 km altitude, near-circular, polar orbit. The orbit is “near”-circular, as LRO’s altitude can vary between its lowest altitude (periapsis) of 35 km and its highest altitude (apoapsis) of 65 km over a twenty eight day period.

Table 4: Low-altitude LRO orbit for higher resolution imagery 39) 40)


Figure 17: Image of the Apollo 17 landing site with LROC-NAC taken at periapsis (image credit: NASA, ASU)

Legend: ALSEP (Apollo Lunar Surface Experiments Package), a portable scientific lab.

• In June 2011, NASA has declared full mission success for LRO after operating the spacecraft and its instruments for a one-year mission phase. Now that the final data from the instruments have been added to the agency's Planetary Data System, the mission has completed the full success requirements. The rich new portrait rendered by LRO's seven instruments is the result of more than 192 TB of data, images and maps, the equivalent of nearly 41,000 typical DVDs. 41)

- The LRO mission is ongoing with near continuous acquisition of science data.

- While studying the Hermite crater near the moon's north pole, LRO's Diviner Lunar Radiometer Experiment (DLRE) found the coldest spot in the solar system, with a temperature of 25 K.

• In May 2011, the following LR (Laser Ranging) results were presented at the 17th International Workshop on Laser Ranging, Bad Kötzting, Germany: (Ref. 79)

- One-way (uplink only) laser transponders have now been proven to work operationally (currently going on 2 years of operations)

- Two-way asynchronous transponders have been successfully demonstrated at planetary distances

- LRO-LR has been very successful thanks to support of ILRS

- LRO will be moved from 50 km circular mission orbit to reduced maintenance elliptical orbit late in 2011. LR is expected to continue at least through FY12.

• In April 2011, the PI of the Mini-RF instrument, Ben Bussey, is reporting that the anomaly of the instrument transmitter could not be fixed. Hence, the project was not able in collecting science data with the instrument anymore since the beginning of 2011.

- However, the rest of the Mini-RF is functioning nominally, and the project is investigating opportunities to conduct interesting science using Mini-RF in a “receive only” mode. One possibility is to conduct bistatic experiments, using the Arecibo Radio Telescope in Puerto Rico as the transmitter.

- Since entering lunar orbit in June 2009, Mini-RF has imaged over two thirds of the lunar surface, including more than 98% of both polar regions. 42)

• On March 15, 2011, the LRO team released the final set of data from the mission's exploration phase along with the first measurements from its new life as a science satellite. - Note: the science mission started after the exploration mission on Sept. 16, 2010 and is projected for two years. 43) 44)

- Among the latest products is a global map with a resolution of 100 m/pixel from the LROC (Lunar Reconnaissance Orbiter Camera). To enhance the topography of the moon, this map was made from images collected when the sun angle was low on the horizon. Armchair astronauts can zoom in to full resolution with any of the mosaics—quite a feat considering that each is 34,748 pixels by 34,748 pixels, or approximately 1.1 GByte.

- The complete data set contains the raw information and high-level products such as mosaic images and maps. The data set also includes more than 300,000 calibrated data records released by LROC. All of the final records from the exploration phase, which lasted from Sept. 15, 2009 through Sept. 15, 2010, are available through several of the Planetary Data System nodes and the LROC website.

• On Jan. 4, 2011, the Mini-RF instrument team for LRO found that the Mini-RF radar had suffered an anomaly and is not currently producing useful science data. Preliminary analysis indicates a possible fault in the Mini-RF radar transmitter. Mini-RF has suspended normal operations until analysis of the situation is completed. 45)

• In December 2010, NASA presented a most precise and complete map to date of the moon's complex, heavily cratered landscape. The digital elevation map was compiled with a dataset of LOLA (Lunar Orbiter Laser Altimeter). This dataset is being used to make digital elevation and terrain maps that will be a fundamental reference for future scientific and human exploration missions to the moon. While the current maps contain ~ 3 billion data points of LOLA so far, the project expects to continue these measurements for the next two years of the mission's science phase and beyond. 46)

The positional errors of image mosaics of the lunar far side, where direct spacecraft tracking (the most accurate) is unavailable, have been 1-10 km in the past. The LRO project is reducing these dimensions to a level of ≤ 30 m in the horizontal and to ≤ 1 m in the vertical plane. At the poles, where illumination rarely provides more than a glimpse of the topography below the crater peaks, the LRO project found systematic horizontal errors of hundreds of meters as well.


Figure 18: LOLA topographic map of the moon's southern hemisphere. The false colors indicate elevation: red areas are highest and blue lowest (image credit: NASA/GSFC)

• The first 2-year extended mission phase of LRO started in mid-September 2010 for additional lunar science measurements supported by NASA’s Science Mission Directorate (SMD).

• The one-year “exploration phase” of the LRO mission was completed on Sept. 16, 2010, meeting all objectives. It produced a comprehensive map of the lunar surface in unprecedented detail; searched for resources and safe landing sites for potential future missions to the moon; and measured lunar temperatures and radiation levels.

The mission is now turning its attention from exploration objectives to scientific research.

• In the summer months of 2010, the Mini-RF instrument of LRO is about half way through its first high-resolution polar-mapping campaign. It is imaging within 20º latitude of both poles using its S-zoom mode. Recently, Mini-RF imaged a potentially ice-rich crater near the north pole of the moon. Located at 84ºN, 157ºW, this permanently shadowed crater, about 8 km in diameter, lies on the floor of the larger, more degraded crater Rozhdestvensky (177 km in diameter). With no sunlight to warm the crater floor and walls, ice brought to the moon by comets or formed through interactions with the solar wind could potentially collect here. 47) 48)

The crater was first identified as a region of interest with Mini-SAR, a NASA instrument flown on the Chandrayaan-1 mission of ISRO in 2009, when it was seen to exhibit unusual radar properties consistent with the presence of ice. But with a Mini-RF resolution 10 times better than the radar (Mini-SAR) aboard the Chandrayaan-1 spacecraft, Mini-RF allows the project to see details of the crater’s interior. In particular, the CPR (Circular Polarization Ratio) measures the polarization characteristics of the radar echoes, which give clues to the nature of the surface materials. The inset image in Figure 19 shows a "same-sense" radar image of the crater (left) next to a colorized CPR image of the crater. Red pixels have CPR values greater than 1.2. The CPR values inside the crater are almost all greater than 1, whereas the CPR values outside the crater are generally low (much less than 1). Regions with CPR greater than 1 are relatively rare in nature, but are commonly seen in regions with thick deposits of ice (such as the Martian polar caps, or the icy Galilean satellites). They are also seen in rough, blocky ejecta around fresh, young craters, but in that occurrence, scientists also observe high CPR outside the crater rim. This feature has high CPR inside its rim, but low CPR outside. The Mini-RF team plans to examine data from the other LRO instruments, particularly temperature and topographic measurements, to better characterize the environment and setting of these unusual features near the poles of the moon.


Figure 19: The SAR instrument Mini-RF returns first high-resolution view of an unusual crater near the moon’s north pole (image credit: NASA)

Using data from the NASA Mini-SAR instrument on the Chandrayaan-1 spacecraft of ISRO, scientists have detected ice deposits near the moon's north pole. Mini-SAR found more than 40 small craters with water ice (Figure 20). The craters range in size from 2 - 15 km in diameter. Although the total amount of ice depends on its thickness in each crater, it is estimated there could be at least 600 million metric tons of water ice. 49)


Figure 20: Mini-SAR map of the moon's north pole region CPR distribution (image credit: NASA)

• On June 23, 2010, LRO has been one full year in lunar orbit. In this timeframe of the mission, LRO has gathered more digital information than any previous planetary mission in history. To celebrate one year in orbit, NASA provided a list of 10 cool things already observed by LRO. Among these items - the Diviner instrument of LRO measured a temperature of -248º C (or 35 K) in the floor of the moon's Hermite Crater. This represents the coldest place measured anywhere in the solar system. 50)


Figure 21: The lunar far side topography observed by the LOLA instrument with the highest peaks of 6000 m (red) and the lowest areas of -6000 m (blue), image credit: NASA/GSFC

• The LRO spacecraft and its payload are operating nominally in 2010 in lunar orbit. LRO will have approximately 210 m/s of ΔV remaining after the 1-year nominal mapping mission is completed in mid-September of 2010. These reserves will be available for extended mission operations (Ref. 14).

• Rediscovery of the Russian Lunokhod-1 and -2 retroreflectors locations on the lunar surface (Luna 17 landed on the moon on Nov.17, 1970 releasing Lunokhod-1): Using LRO's mapping data, researchers at the UCSD (University of California San Diego) successfully pinpointed the location of a long lost light reflector on the lunar surface by bouncing laser signals from Earth to the Russian Lunokhod 1 retroreflector. The initial imaging of the two Russian rover locations, Lunokhod-1 and -2 were made in early 2010 by the LROC (Lunar Reconnaissance Orbiter Camera) team, led by Mark Robinson from Arizona State University in Tempe, AZ. 51) 52)

On April 22, 2010, Tom Murphy from UCSD and his team sent pulses of laser light from the 3.5 m telescope at the Apache Point Observatory in New Mexico, zeroing in on the target coordinates provided by the LROC images and altitudes provided by the LOLA (Lunar Orbiter Laser Altimeter). The new locations of Lunokhod-1 and -2 were quickly verified by the signal response from the retroreflectors.

Figure 22: Illustration of the Lunokhod-1 retroreflector (image credit: NASA)

• At the end of 2009, LROC (Lunar Reconnaissance Orbiter Camera) has mapped in high resolution all the Apollo landing sites and 50 sites that were identified by NASA's Constellation Program to be representative of the wide range of terrains present on the moon (Figure 24). 53)

• The LRO mission played a major role in the support of the Oct. 9, 2009 LCROSS impact experiment. The role of LRO is to make detailed observations before, during, and after the LCROSS impact in a permanently shadowed crater near the lunar South Pole. The impact will occur within the highlands impact crater Cabeus on October 9, 2009. LRO has made detailed observations to support the LCROSS selection of this crater for the impact. The LRO team used laser altimetry data from both the LRO LOLA and JAXA Kaguya laser altimeters to determine regions of permanent shadows that would be the most likely regions to harbor frozen volatiles, if any are preserved in significant concentrations. The topography measurements also identified regions within the polar impact craters where the local slope is sufficiently small to maximize the transfer of kinetic energy from LCROSS impact into the lunar regolith target. LRO’s LEND neutron spectrometer provided maps of enhanced hydrogen concentrations that could indicate water ice embedded in the upper meter of the lunar regolith. LRO Diviner temperature measurements of the South Polar region were used to reveal the extremely low temperatures of cold traps that can potentially preserve volatiles in ice form. Mini-RF dual-frequency radar polarization imaging provided information that indicates the blockiness of the impact site and which can be used to further test for the presence of significant water ice on the basis of anomalous scattering behavior. These combinations of LRO measurements as well as other factors led the LCROSS team to choose the south polar crater Cabeus for its impact target. Before the impact occurs the Cabeus target area will be exhaustively observed by the LRO LROC, LOLA, and Mini-RF instruments in detail to characterize the pre-impact geology. 54)

• NASA has successfully completed its testing and calibration phase and entered its mapping orbit of the moon on Sept. 17, 2009. The spacecraft already has made significant progress toward creating the most detailed atlas of the moon's south pole to date. Scientists released preliminary images and data from LRO's seven instruments. 55) 56)


Figure 23: This image shows the daytime and nighttime lunar temperatures recorded by DIVINER (image credit: NASA/UCLA) 57)

• On Sept. 15, 2009 the LRO orbiter spacecraft was moved into a polar-inclination of 89.7º circular 50 km mean altitude orbit for a planned one-year duration to execute its baseline ESMD (Exploration Systems Mission Directorate) mission phase, the so-called “Exploration Mission”. The orbital period is typically 113 minutes. This polar orbit allows repetitive measurements at high latitudes, producing a dense net of observations. Since the Moon rotates once each sidereal month, successive groundtracks are separated by about 31 km at the equator. Observations accumulated during the one-year Exploration Mission result in complete global coverage.

LRO has already proved its keen eyes, imaging fine details of the Apollo landing sites in August with the LROC (Lunar Reconnaissance Orbiter Camera) imager. - During the nominal mission phase, the Maneuver Team designed maneuvers to allow for the successful viewing of the LCROSS impact on October 9, 2009. Maneuvering LRO for the LCROSS impact viewing included many iterations and re-plans to adapt to changing requirements for viewing the impact (Ref. 14).


Figure 24: LROC image of the Apollo-12 landing site taken in Aug. 2009 (image credit: NASA)

Legend: Figure 24 of LROC (Lunar Reconnaissance Orbiter Camera) shows the spacecraft's first look at the Apollo 12 landing site (Apollo-12 was launched in Nov. 1969). The Intrepid lunar module descent stage, the experiment package ALSEP (Apollo Lunar Surface Experiment Package), and Surveyor 3 spacecraft are all visible. Astronaut footpaths are marked with unlabeled arrows. This image is 824 m in width. 58)

• During the commissioning phase, it was determined that LRO could perform coordinated observations with the ISRO (Indian Science Research Organization’) spacecraft Chandrayaan-1 (launched on October 22, 2008). As part of its instrument suite, Chandrayaan-1 carried the MiniSAR instrument – a synthetic aperture radar and sister instrument to LRO’s Mini-RF instrument. The goal for the coordinated observation was to perform a bistatic SAR experiment whereby Chandrayaan-1 would transmit from MiniSAR into a lunar South Pole crater and both Chandrayaan-1 and LRO would attempt to receive the return signal with their sister instruments. The experiment would attempt to find water ice in Erlanger Crater (Longitude: 29.16º, Latitude: -87.01º), one of the permanently shadowed craters at the lunar South Pole.

The experiment took place on August 20, 2009 with the corresponding instruments’ sensor footprints overlapping over the Erlanger crater for roughly 35 seconds. The close approach between the satellites was approximately 22.5 km, a majority of the difference being in the radial direction. After analyzing the encounter data, ISRO determined that, due to deteriorating spacecraft hardware, Chandrayaan-1 was not pointing at the Erlanger Crater during the experiment time. - A second attempt using a different crater was being investigated when communications were lost with Chandrayaan-1 on August 29, 2009 and the Chandrayaan-1 mission was terminated (Ref. 14).

• The first two successful SLR passes between a terrestrial ground station and a spacecraft orbiting the moon were obtained on July 1, and July 2, 2009 between the NGSLR station at Greenbelt, Maryland, USA, and the LRO (Lunar Reconnaissance Orbiter).

• On June 30, 2009 the LROC NAC and WAC cameras were activated. The cameras are working well and have returned first images of a region a few kilometers east of Hell E crater in the lunar highlands south of Mare Nubium. 59)

• On June 23, 2009, 4 1/2 days after launch, LRO has successfully entered orbit around the moon. During transit to the moon, engineers performed a mid-course correction to get the spacecraft in the proper position to reach its lunar destination. 60)



Sensor complement: (CRaTER, DLRE, LAMP, LOLA, LROC, LEND, Mini-RF)

The spacecraft payload consists of six instruments and one technology demonstration to perform investigations specifically targeted for preparing for future human exploration. The instruments are provided by various organizations in the United States, one is from Russia. 61)


Mass allocation (kg)

Power orbit average allocation (W)




Table 5: Summary of LRO instrument mass and power allocations

Instrument name

Instrument type

Characteristic range

Characteristic resolution

Spatial resolution from 50 km orbit

Spatial coverage

Data rate (Gbit/day)


Laser altimeter

Range window 20-70 km

10 cm vertical

Five 5 m laser spots, 25 m spacing

Polar grid 0.001º latitude, 0.04º longitude



High resolution camera

Broadband centered at 550 nm

±150 nm

50 cm/pixel

Targeted >10% lunar surface, 100% > 85.5º lat.



Multispectral camera

315-680 nm

Spectral filters centered at 315 nm 360 nm, 415 nm, 560 nm, 600 nm, 640 nm, 680 nm

100 m/pixel VIS
400 m/pixel UV

Full lunar surface at each wavelength and various lighting angles



Neutron detector

Thermal to 15 MeV

Four bands
Thermal <0.4 eV, Epithermal: 0.4 eV-10 keV, Fast: 10 keV-1 MeV, Energetic: 1 MeV-15 MeV

Epithermal 10 km
FWHM (see test for
other bands)

Full lunar surface and deep space




30-400 K

5 K

400 m

Full lunar surface day/night temperatures



UV imaging spectrograph

52 to 187 nm

3.5 nm

260 m

Full lunar surface



Primary and albedo cosmic ray sensor

LET spectra 0.2 keV/µm to 7 MeV/µm


77 km

Full lunar surface and deep space

7.8 (peak)


X- and S-band SAR

4 cm (X-band)
12 cm (S-band)

Sensitivity: -30 dB (S), -25 dB (X)

75 m/pixel, 7.5 m/pixel (zoom)

Limited during the nominal mission

7.7 (for 4 min observations)

Table 6: Overview of the LRO instrument complement


CRaTER (Cosmic Ray Telescope for the Effects of Radiation):

CRaTER PI: Harlan E. Spence, UNH (University of New Hampshire), Durham, NH. The primary goal is to characterize the global lunar radiation environment and its biological impacts. The instrument consists of a single, integrated sensor and electronics box with simple electronic and mechanical interfaces to the spacecraft. The CRaTER sensor frontend design is based on standard stacked-detector, cosmic ray telescope systems. 62)

The objective of CRaTER is to measure LET (Linear Energy Transfer) spectra produced by incident galactic cosmic rays (GCRs) and solar energetic protons (SEPs). GCRs and SEPs with energies >10 MeV have sufficient energy to penetrate even moderate shielding. CRaTER is designed to return the following required data products:

• Measure and characterize that aspect of the deep space radiation environment, LET spectra of galactic and solar cosmic rays (particularly above 10 MeV), most critically important to the engineering and modeling communities to assure safe, long-term, human presence in space.

• Investigate the effects of shielding by measuring LET spectra behind different amounts and types of areal density, including tissue-equivalent plastic.

The CRaTER telescope consists of five ion-implanted silicon detectors (red areas in Figure 27), mounted on four detector boards (green areas), and separated by three pieces of tissue-equivalent plastic, hereinafter referred to as TEP (tan areas). All five of the silicon detectors are 2 cm in diameter. 63) 64) 65)

Low LET (Light Emitting Transistor) detectors

9.6 cm2 circular, 1000 µm thick. 0.2 MeV threshold

High LET detectors

9.6 cm2 circular, 140 µm thick. 2 MeV threshold

TEP (Tissue-Equivalent Plastic) absorber 1

5.4 cm cylinder

TEP absorber 2

2.7 cm cylinder

Zenith FOV (Field of View)

35º, 6 detector coincidence

Nadir FOV

75º, for D3D4D5D6 coincidence

Geometric factor

0.1 cm2 sr (D1D2 events)

LET range

0.2 - 7 MeV/µm (Si)

Incident particle energy range

> 20 MeV (H), > 87 MeV/nucleon (Fe)

Table 7: Parameters of CRaTER


Figure 25: Detailed view of the CRaTER telescope (image credit: BU)


Figure 26: Illustration of detector location in CRaTER (image credit: University of Tennessee) 66)

Legend to Figure 26: The detectors (D1-D6) are made of silicon, the TEPs are composed of hydrogen, carbon, nitrogen, oxygen, fluorine, and calcium, in a tissue-equivalent mixture (A-150 plastic). The end caps are made of aluminum.

CRaTER is composed of three sets of detectors. The first set of detectors consists of thin silicon (140 µm thick) followed by a second, thicker detector (1000 µm thick). Thin detectors primarily detect particles with a high LET while thick detectors primarily detect low LET particles. Sandwiched between each of the three pairs of detectors is a slab of A-150 tissue-equivalent plastic (TEP). The first silicon detector pair D1 and D2 is on the zenith end, which faces away from the lunar surface out into deep space.

Then there is a 5.4 cm long section of TEP, followed by another detector pair D3 and D4, followed by 2.7 cm long section of TEP, and the final detector pair D5 and D6.


Figure 27: Photo of the CRaTER instrument (image credit: BU)


DLRE (Diviner Lunar Radiometer Experiment):

DLRE PI: D. Paige, UCLA. The overall objective is to measure the lunar surface thermal environment (temperatures) at scales that provide essential information for future surface operations and exploration (resolution 300 m). DLRE is a a multi-channel (9 channels) solar reflectance and infrared filter radiometer utilizing uncooled thermopile detector arrays. DLRE's spectral channels are distributed between two identical, boresighted telescopes, and an articulated elevation/azimuth mount allows the telescopes to view the lunar surface, space, and calibration targets. The IFOV response of each channel is defined by a linear, 21-element, thermopile detector array at the telescope focal plane, and its spectral response is defined by a focal plane bandpass filter.

The DLRE structure consists of an instrument optics bench assembly (OBA), an elevation/azimuth yoke, and an instrument mount. The OBA contains all of the instrument optical subassemblies, and is suspended from the yoke. Elevation and azimuth motors mounted on the yoke drive instrument articulation. The OBA is temperature controlled, and internal temperature gradients are minimized by design. Radiometric calibration is provided by views of blackbody and solar targets mounted on the yoke. The electronics subassemblies control signal processing, instrument operation and articulation, command processing, and data processing and are distributed between the OBA and the yoke. 67)


Figure 28: Illustration of the DLRE device (image credit: NASA)

The operation of DLRE is continuously in nadir pushbroom mapping mode using 21 detectors cross-track for each of its nine spectral channels. The FOV of each detector is 3.6 mrad cross track, yielding a resolution of 180 m on the lunar surface at an orbital altitude of 50 km. To facilitate spatial registration of DLRE's surface footprints in multiple spectral bands, and to reduce along-track smear, the integration period will be 0.128 seconds. The mapped data products will generally be at a resolution of ~500 m/pixel to increase the SNR (Signal-to-Noise Ratio), and to allow for anticipated errors in the reconstruction of the position and pointing of the LRO spacecraft.


Channel No

Minimum wavelength (µm)

Maximum wavelength (µm)

Purpose of observation

Minimum detectable signal





Solar reflectance in permanently shadowed regions
Solar reflectance in sunlit regions
Thermal emission near Christiansen feature
Thermal emission near Christiansen feature
Thermal emission near Christiansen feature
Thermal mapping

40 K*
55 K*
160 K
160 K
160 K
75 K





Thermal mapping
Thermal mapping
Thermal mapping

45 K
32 K

Table 8: Spectral channel parameters of the DLRE instrument

Note: * is the intensity of reflected radiation from an isotropic reflector with broadband solar albedo of 0.1 in thermal equilibrium at the quoted temperature.


LAMP (Lyman-Alpha Mapping Project):

LAMP PI: A. Stern, SwRI (Southwest Research Institute), San Antonio, TX. The LAMP instrument is an imaging UV spectrometer. The objectives of LAMP are: 68)

• LAMP will be used to identify and localize exposed water frost in PSRs (Permanently Shadowed Regions)

• Provision of landform mapping (using Lyman-α albedos) in and around the PSRs of the lunar surface

• Demonstrate the feasibility of using starlight and UV sky-glow for future night time and PSR surface mission applications

• Assay the lunar atmosphere and its variability.

Viewing in the nadir direction from LRO, LAMP measures the signal reflected from the nightside lunar surface and PSR (Permanently Shadowed Regions) using Lyman-α skyglow and UV starlight as a light source. The LAMP data are taken entirely in pixel list (i.e., time tagged) mode, allowing mapping at a variety of resolutions. The reflectance data yield albedo maps of PSRs, the spectra of PSRs yield exposed water frost abundances, and the atmospheric spectra yield species abundances and variability.

Instrument mass, power

5.0 kg, 4.3 W (each with reserves)


Pluto-ALICE UV Spectrograph, no new technologies


2 year (required), 5 year (goal)


1200-1800 Å

Effective area

0.4 cm2 @ 1216 Å (Lyman-α)


0.2º x 6º

Spectral resolution

< 20 Å FWHM (Full Width Half Maximum) across passband

Spatial resolution

<1º (Nyquist sampled,PSF)

Filled slit spectral resolution

< 40 Å FWHM average across passband

Stray light

< 10-5 at 7º off-axis

Maximum count rate

> 15 kHz (~50% deadtime loss)

Dark count rate

< 50 counts/s (total array)

Detector output

Continuous pixel list

Table 9: Summary of the LAMP instrument parameters

The LAMP instrument is of ALICE heritage flown on the Rosetta mission of ESA and the New Horizon mission of NASA. LAMP is comprised of a telescope and Rowland-circle spectrograph. LAMP has a single 40×40 mm2 entrance aperture that feeds light to the telescope section of the instrument. Entering light is collected and focused by an f/3 off-axis paraboloidal (OAP) primary mirror at the back end of the telescope section onto the instrument's entrance slit. After passing through the entrance slit, the light falls onto a toroidal holographic diffraction grating, which disperses the light onto a double-delay line (DDL) microchannel plate (MCP) detector. The 2D pixel format detector (1024 x 32) is coated by a CsI solar-blind photocathode and has a cylindrically curved MCP stack that matches the Rowland-circle. LAMP is controlled by an Intel 8052 compatible microcontroller, and utilizes lightweight, compact, surface mount electronics to support the science detector, as well as the instrument support and interface electronics.


Figure 29: Schematic view of the LAMP instrument (image credit: SwRI)


Figure 30: The LAMP design as seen from above (left) and below (right), image credit: SwRI


LEND (Lunar Exploration Neutron Detector):

LEND is a contributed instrument of the Federal Space Agency of Russia (Roskosmos). In addition, there are many collaborators in the project from inside and outside of Russia. The LEND instrument PI is Igor Mitrofanov of IKI (Space Research Institute), Moscow. The LEND investigations are based on the detection of the moon's neutron albedo. Specific objectives are to provide: 69) 70) 71) 72)

• High resolution hydrogen distribution maps with sensitivity of about 100 ppm of hydrogen weight and a horizontal spatial resolution of 5 km

• Characterization of surface distribution and column density of possible near-surface water ice deposits in the moon's polar cold traps

• Creation of a global model of neutron component of space radiation at altitude of 30-50 km above moon's surface with spatial resolution of 20-50 km at the spectral range from thermal energies up to 15 MeV.

LEND is capable of providing high spatial resolution mapping of epithermal neutrons with collimated epithermal neutron detectors. LEND is able to detect a hydrogen-rich spot at one of the lunar poles with as little as 100 ppm of hydrogen and a spatial resolution of 10 km (pixel diameter), and to produce global measurements of the hydrogen content with a resolution of 5–20 km. If the hydrogen is associated with water, a detection limit of 100 ppm hydrogen corresponds to ~ 0.1% weight water ice in the regolith. High energy neutron data from another LEND sensor could help to distinguish between areas in which hydrogen was implanted by solar wind and potential water ice deposits.

LEND features a full set of sensors for thermal (STN 1-3), epithermal (SETN) and high energy neutrons (SHEN) to provide data for neutron components of radiation environment in the broad range of more than 9 decades of energy. The LEND instrument design is based on the Russian HEND (High Energy Neutron Detector), which continues to perform well in its fifth year of science measurements onboard NASA's Mars Odyssey mission.


Figure 31: Conceptual view of the cosmic ray induced neutron flux on the lunar surface (image credit: NASA) 73)

LEND's primary sensor type is the 3He counter, used for the detectors CSETN 1-4, STN 1-3, and SETN. The 3He counter produces an electrical pulse proportional to the number of ions formed. The major difference between LEND and HEND, is the collimation of neutron flux before detection. Collimating modules around the 3He counters of CSETN 1-4 effectively absorb neutrons that have large angles with respect to the normal on the moon's surface, leading to spatial resolution of 10 km full width at half maximum signal from the nominal 50 km orbit.

The LEND collimators of neutrons provide very high spatial resolution maps of neutron emission at the lunar surface. No other neutron instrument with this imaging capability has ever flown in space. LEND has a total of nine neutron sensors:

+ A - 4 3He collimated counters CSETN1-4 for epithermal neutrons >0.4 eV
+ B - 2 3He counters of Doppler filter STN1-2 for thermal neutrons
+ C - 3He counter STN-3 for thermal neutrons
+ D - 3He counter SETN for epithermal neutrons
+ E - stylbene scintillation sensor SHEN with 16 energy channels from 0.3 up to > 15 MeV (the stylbene scintillation spectrometer is being used for the detection of fast neutrons).

- Four of the 3He counters are collimated with a combination of polyethylene and 10B powder.

- Collimated detectors are also surrounded by Cd shields to filter out thermal neutrons with energies below ~0.4 eV so they are primarily sensitive to epithermal neutrons.

These detectors are also surrounded by Cd shields to filter out thermal neutrons with energies below ~0.4 eV so they are primarily sensitive to epithermal neutrons. The epithermal neutron flux is very sensitive to the presence of hydrogen in the lunar regolith and the collimated LEND 3He counters will provide detection of hydrogen near the poles to levels of 100 ppm or better with spatial resolution of 5 km (Half Width Half Maximum). If the hydrogen is associated with water, a detection limit of 100 ppm of hydrogen corresponds to ~0.1 wt% of water ice homogeneously distributed in the regolith. Over the course of the one-year LRO mission, LEND will be able to produce global maps of hydrogen content with resolutions of 5-20 km.


Figure 32: The LEND instrument with four collimated sensors of epithermal neutrons CSETN 1-4 (image credit: IKI)

A numerical simulation of LEND performance showed that the instrument, with the optimal shaping of the collimators of sensors CSETN 1-4, may provide a detection limit (3σ) of hydrogen of about 82 ppm for a polar spot with a diameter of 10 km (FWHM), given a baseline 1 year mapping mission from a 50 km polar orbit. This detection sensitivity increases for larger spots, and decreases for locations more distant from the pole.


Figure 33: Schematic view of LEND collimator detection concept (image credit: IKI)

Instrument mass, power

23.7 kg, 13 W

Instrument size

372 mm x 344 mm x 327 mm

Instrument type

Collimated neutron spectrometer

Energy range

From thermal energies up to 15 MeV

Time resolution

Variable, > 1 s

Spatial resolution on the lunar surface

5 km from the 50 km orbit

Spatial resolution inside the subsurface

1-2 m

Working range of temperature

From -20ºC up to +50ºC

Telemetry data volume

250 MB per day

Instrument design life

5 years

Table 10: Overview of some performance parameters of the LEND instrument


Figure 34: The LEND instrument with 4 detectors shown - the other 5 sensors are inside the collimator module (image credit: Roskosmos, IKI)


LOLA (Lunar Orbiter Laser Altimeter):

LOLA PI: D. E. Smith, NASA/GSFC. The objectives are to provide a precise global lunar topographic model and geodetic grid that will serve as the foundation of this essential understanding. Topography at scales from local to global is necessary for landing safely and, in addition; it preserves the record of the evolution of the surface which contributes to decisions as to where to explore.

LOLA is a “geodetic tool” to derive a precise positioning of observed features with a framework (grid) for all LRO measurements: 74) 75) 76) 77)

- Topography of the moon to an accuracy ±1 m and 0.1 m precision

- Surface slopes in 2 directions to better than 0.5º on a 50 m scale

- Surface roughness to 0.3 m

- Surface reflectance of the moon at 1064 nm to ~ 5%

- Establish a global lunar “geodetic” coordinate system

- Improve knowledge of the lunar gravity field.

• LOLA is a 70 m wide swath altimeter (includes field of view of detectors) providing 5 profiles

- Along-track sampling in latitude of 25 m

- Cross-track sampling in longitude 0.04º (~25 m above latitude 85º and ~1.2 km at the equator), after 1 year of operation.

• LOLA characterizes the swath in elevation, slope and surface roughness, and brightness

• Knowledge of pixel locations determines map resolution.

Instrumentation: The LOLA instrument pulses a single laser through a DOE (Diffractive Optical Element) device to produce five beams that illuminate the lunar surface. For each beam, LOLA measures time of flight (range), pulse spreading (surface roughness), and transmit/return energy (surface reflectance). With its two-dimensional spot pattern, LOLA unambiguously determines slopes along and across the orbit track.

The LOLA instrument design is of MOLA (Mars Orbiter Laser Altimeter) and of MLA (Mercury Laser Altimeter) heritage. However, LOLA has five laser beams and five receiver channels. LOLA's laser transmitter consists of a single stage diode-pumped and Q-switched Nd:YAG laser with a 1064 nm wavelength, a 2.7 mJ pulse energy, a 6 ns pulse, a 28 Hz pulse rate, and a 100 µrad beam divergence angle. A diffractive optics element made of fused silica with an etched-in diffraction pattern is used to split the single incident laser beam into five off-pointed beams, creating the 50 m diameter 5-spot cross-pattern on the lunar surface. The reflected signal is collected by a 14 cm diameter telescope with a 5-optical-fiber array at the focal plane. Each of the five optical fibers collects the reflected signal from one of the five laser spots on the lunar surface, and delivers it to one of the five avalanche photodiodes.


Figure 35: Photo of the LOLA optical fiber array (image credit: NASA)

The transmitted laser pulse and the five received laser pulses are time stamped with respect to the spacecraft mission elapsed time using a set of time-to-digital converters at < 0.5 ns precision. In addition, LOLA measures the transmitted and received pulse by integrating the pulse waveforms. The on-board science algorithm, running on an embedded microprocessor, autonomously adjusts the receiver detection threshold levels and detector gain to keep the range window tracking the lunar surface returns.


Pulse energy
Pulse width and rate
Beam splitting
Beam divergence (per beam)
Beam separation

2.7 ± 0.3 mJ
6 ns FWHM, 28 Hz
1064.30 ±0.1 nm
5-way, > 13% total per beam
100 µrad
500 µrad

Receiver optics

Receiver aperture
FOV (Field of View)
Optics transmission
Optical bandwidth

Atel = 0.015 m2 (Ø = 0.14 m)
400 µrad
> 70 %
0.8 nm


Detector type
Detector active area
Detector quantum efficiency
Noise equivalent power
Electrical bandwidth

5 Si-APD
0.7 mm
0.05 pW/Hz1/2
100 MHz

Timing electronics

Timing resolution
Clock frequency uncertainty
Laser pulse epoch time accuracy

< 0.5 ns
< 1 e-7

< 3 ms

LOLA instrument mass, power

12.6 kg, 34 W


35 cm x 35 cm x 29 cm

Data rate

28 kbit/s

Table 11: Parameter specification of the LOLA instrument

LOLA is a pulse detection time-off-light altimeter that incorporates a five-spot pattern that measures the precise distance to the lunar surface at 5 spots simultaneously, thus providing 5 profiles across the lunar surface. Each spot within the five-spot pattern has a diameter of 5 m; the spots are 25 m apart, and form a cross pattern (Figure 36). The 5-spot pattern enables the surface slope to be derived in the along-track and cross-track directions; the pattern is rotated approximately 26º to provide five adjacent profiles, 10 to 12 m apart over a 50 to 60 m swath, with combined measurements in the along track direction every 10 to 12 m.


Figure 36: Schematic illustration of the LOLA observation pattern (image credit: NASA)

Since LOLA provides global observations, the LOLA altimetry data can be used to improve the spacecraft orbit, and the knowledge of far side lunar gravity - which is currently extremely poorly known but is required for precise landing and low-altitude navigation.


Figure 37: Functional block diagram of the LOLA instrument (image credit: NASA)


Figure 38: Illustration of the LOLA instrument - two views (image credit: NASA)

LOLA (and other LRO instruments) require accurate orbits of LRO

- high quality tracking

- improvement in the lunar gravity field

• Baseline tracking of LRO is S-band Doppler at 1 mm/s at 5 second rate from White Sands (NM), and 8 mm/s from other S-band systems enabling 24 hours/day, 7 days/week coverage (when LRO is visible)

• Simulations of the LRO mission show S-band tracking will not provide enough information to precisely determine the lunar gravity field.


Figure 39: LR (Lunar Ranging) flight system components (image credit: NASA)


Figure 40: LR (Lunar Ranging) operations overview (image credit: NASA)


Figure 41: Simplified LOLA-LR block diagram (image credit: NASA)

Resulting products:

6) Relative range measurements to LRO spacecraft at <10 cm precision at 1 Hz

7) Gravity model with sufficient accuracy to calculate knowledge of spacecraft position to within 50 m along-track, 50 m cross-track, and 1 m radial.


Figure 42: Illustration of LOLA-measured topography in the vicinity of the Apollo 11 landing site (image credit: NASA, Ref. 44)


LRO-LR (Lunar Reconnaissance Orbiter-Laser Ranging)

LRO-LR is the first mission for the ILRS (International Laser Ranging Service) whose primary laser tracking method is transponder ranging. LRO-LR is a one-way (uplink only) ranging technique where the Earth laser station measures the fire times of its outgoing laser pulses and LOLA (Lunar Orbiter Laser Altimeter) measures the receive times. The range gate for the Earth received pulses in LOLA’s detector #1 is called the Earth Window. During this window the detector is gated on to receive Earth events. LOLA performs signal processing on the received Earth events and provides the signal processing results to the ground via its housekeeping packets.

These housekeeping packets are routed through the LRO Mission Operations Center (MOC) to the LOLA Science Operations Center (SOC) where the relevant data is extracted and put into a real-time website which is displayed from the Crustal Dynamics Data Information System (CDDIS) server. This real-time LRO-LR website provides feedback to all participating stations while they are ranging to LRO. Unlike two-way ranging where the laser light returns to the station and provides the feedback, this website is the only feedback that the stations have while they are ranging to LRO. The latency of the website is nominally 10 to 20 seconds, but has been observed to be as long as several minutes. 78) 79)

There are ten stations supporting laser ranging to LRO. These stations are shown in Table 12 along with their first successful ranges and their system characteristics.

Four of the participating stations are NASA MOBLAS systems. These systems were modified to permit ranging to LRO. A new Windows computer with a Guidetech timing card (model GT658) was added to each system to provide the precision needed for the fires, and the systems were all modified to fire their laser at 10 Hz.



Sync to LOLA

Fire rate (Hz)

Max #/s in LOLA window

Expected energy at LRO (FJ/cm2)

frequency source

Date of first successful ranging to LRO

LR status


Maryland, USA




1 to 5

Maser (18-Oct-2010)




Texas, USA



2 to 4

1 to 10




MOBLAS-7, Greenbelt

Maryland, USA



2 to 4

1 to 3

Maser (18-Oct-2010)




Great Britain




1 to 3

Maser (13 May- 2010)








1 to 3

Oversized crystal oscillator








1 to 10




MOBLAS-6, Hartebeesthoek

South Africa



2 to 4

1 to 3








2 to 4

1 to 3

Maser (14-May-2010)



MOBLAS-4, Monument Peak

California, USA



2 to 4

1 to 3

GPS steered Rubidium







2 to 4

1 to 10




Table 12: Participating ILRS stations and their characteristics

Some of the participating ground stations control their laser fires to ensure the pulses arrive when the LOLA Earth Window is open. These systems are referred to as synchronous ranging stations. NGSLR, Herstmonceux and Zimmerwald are all synchronous stations. All other systems are asynchronous. The MOBLAS systems and Grasse all fire at 10 Hz. MLRS fires at approximately 10 Hz. Systems that fire at 10 Hz get two pulses per second into the LOLA Earth Window most of the time, and occasionally they will get four pulses per second into the Window. Wettzell fires at 7 Hz and they tune their fire frequency to match the range-rate.

Figure 43: Overview of ILRS station locations (image credit: NASA)

LR results: As of May 2011, there were over 1000 hours of Laser Ranging data collected from all of the stations. NGSLR has over 45% of the global data collected since launch, with Yarragadee at 18%, Monument Peak at 15%, and MLRS at 13%. The global data rate appears to be increasing as shown in the plot of Figure 44.

In the early months after launch only a single station was scheduled to range to LRO at any time. This was to give the stations some experience in using the real-time LRO-LR website for feedback. Simultaneous ranging to LRO by two or more stations allows comparison of station ranging and biases. Three-way simultaneous ranging can potentially provide a geometric solution of the spacecraft location. Simultaneous ranging opportunities are now scheduled for all NGSLR, MLRS, MOBLAS-7 (Greenbelt), and MOBLAS-4 (Monument Peak) passes. In addition Grasse and Zimmerwald are also always scheduled for simultaneous ranging opportunities. More stations will be simultaneously scheduled in the near future.


Figure 44: Plot of LR data from launch (June 2009) to May 2011 (image credit: NASA)

With LRO-LR entering year 3 of successful operations, one-way (uplink only) laser transponders have now been proven to work operationally. Thanks to the support of the ILRS and the participating stations, over 1000 hours of LR data has been collected and used to determine spacecraft time to UTC, and will be used to provide more precise orbits. In addition time transfer between ground stations using LRO will be attempted later in 2011, initially between Wettzell and NGSLR.

In 2013, orbit determination for LRO is accomplished by utilizing Doppler and altimetric crossover data with an uncertainty in spacecraft positioning of only ~12 m. By incorporating the high-precision LR measurements into this solution, the positioning accuracy is expected to improve, enabling more accurate orbital mapping products. To facilitate the incorporation of the LR data into the orbit solutions, the project is now focusing on the precise referencing of MET (Mission Elapsed Time) to TDB (Barycentric Dynamical Time) times. This shall be achieved by developing an improved clock model that includes effects of relativity as well as other effects. The project expects that the clock calibration and referencing (MET to TDB) is more accurate than the standard conversions provided by the SCLK (Spacecraft Clock Kernel). The data may enable improvements in orbit determination and gravity field modeling. While excellent gravity field data are available from the GRAIL mission, it will be interesting to analyze the benefits of the LR data. The accurate GRAIL fields provide a basis for quantitatively evaluating improvements in –pre-GRAIL models. 80)


LROC (Lunar Reconnaissance Orbiter Camera)

LROC PI: M. Robinson, of ASU (Arizona State University), Tempe, AZ, USA. The LROC Science Team includes participants from Brown University, Washington University, and the University of Arizona. The objectives of LROC are to address two requirements: 1) landing site certification and 2) polar illumination. Specific mission goals are: 81) 82) 83) 84)

- Landing site identification and certification, with unambiguous identification of meter-scale hazards

- Unambiguous mapping of permanent shadows and sunlit regions

- Meter-scale mapping of polar regions with continuous illumination

- Overlapping observations to enable derivation of meter-scale topography

- Global multispectral imaging to map ilmenite and other minerals

- Determine current impact hazard by reimaging 1-2 m/pixel Apollo images

- Global morphology base map

- Characterize regolith properties.

Instrumentation: LROC is a modified version of CTX (ConTeXt Camera) and MARCI (MARs Color Imager) flown on the MRO (Mars Reconnaissance Orbiter) mission. The LROC is comprised of two NACs (Narrow-Angle Cameras), a WAC (Wide-Angle Camera), and the SCS (Sequence and Compressor System). The total mass of LROC is 16 kg. The instrument is being developed by MSSS (Malin Space Science Systems) in San Diego, CA.

• Each NAC uses a Ritchey-Chretien telescope with a focal length of 700 mm that images onto a 5000 pixel CCD line array, providing a cross-track FOV of 2.86º. The NAC readout noise is better than 100 e-, and the data are sampled at 12 bit, then compressed to 8 bit, square root encoded values prior to downlink. The NAC internal buffer holds 256 MB of uncompressed data, enough for a full swath image 25 km long or a 2 x 2 binned image 100 km long.

• The WAC has two short focal length lenses imaging onto the same 1000 x 1000 pixel, electronically shuttered CCD area array, one imaging in the visible/near infrared (EFL = 6.0 mm), and the other in the UV range (EFL = 4.5 mm). The optical systems have a cross-track FOV of 90º and 60º respectively. From the nominal 50 km orbit, the WAC will provide a nadir, ground sample distance of 100 m/pixel in the visible, and a swath width of ~100 km. The seven band color capability of the WAC is provided by a color filter array mounted directly over the detector, providing different sections of the CCD with different filters acquiring data in the seven channels in a “pushframe” mode. Continuous coverage in any one color is provided by repeated imaging at a rate such that each of the narrow framelets of each color band overlap.

Spatial resolution, IFOV (Instantaneous Field of View)

0.5 m, 10 mrad


2.86º (0.05 radian) per NAC

Maximum image size

2.5 km x 25 km


f/3.59 Cassegrain (Ritchey-Chretien)

Effective focal length

700 mm

Primary mirror diameter

195 mm

MTF (Modulation Transfer Function) at Nyquist frequency

> 0.20

Structure + baffle

graphite-cyanide composite


Kodak KLI-5001G

Pixel format

1 x 5,000


100 e-

A/D converter

Honeywell ADC9225

FPGA (Floating Point Gate Array)

Actel RT54SX32-S

Volume, power

70 cm x 26 cm diameter, 10 W (peak), 6 W (average)

Spectral range

400-750 nm

Instrument mass

5.4 kg (each device)

Table 13: Specification of the NAC devices


Figure 45: View of the LROC NAC device (image credit: ASU)


Figure 46: View of the LROC WAC device (image credit: ASU)


Figure 47: NAC optics cutaway (left) and NAC optics and electronics (image credit: ASU)

Imager type

Pushbroom multispectral camera

Image format

1024 x 16 pixels monochrome
704 x 16 pixels 7-filter color

Spectral bands (center frequency)

No 1: 315 nm
No 2: 360 nm
No 3: 415 nm
No 4, 560 nm
No 5: 600 nm
No 6: 640 nm
No 7: 680 nm

Spatial resolution

1.5 mrad, 75 m/pixel at nadir (visible)
2.0 mrad, 400 m/pixel at nadir (UV, 4x binned)

Swath width (from a 50 km lunar orbit)

105 km (visible monochrome)
57 km (visible color)
57 km (UV)

FOV (Field of View)

90º (visible)
60º (UV)


f/5.1 (visible)
f/5.3 (UV)

Effective focal length

6.0 mm (visible), 4.6 mm (UV)

Entrance pupil diameter

1.19 mm (visible), 0.85 mm (UV)

System MTF (Nyquist)

> 0.2


Kodak KLI-1001

Pixel format

1024 x 1024


50 e-

Instrument volume, power

14.5 cm x 9.2 cm x 7.6 cm, 4 W

Instrument mass

0.6 kg

Table 14: Parameters of the WAC instrument


Figure 48: Photo of the WAC device (image credit: Arizona State University)

The NACs and WAC interface with the SCS (Sequence and Compressor System), the third element of the LROC. The SCS commands individual image acquisition by the NACs and WAC from a stored sequence, and applies lossless compression to the NAC and WAC data as they are read out and passed to the spacecraft data system. The SCS provides a single command and data interface between the LROC and the LRO spacecraft data system.


Xilinx XQR2V3000


SpaceWire, LVDS

Power supply

Interpoint SMSA2805S

Power consumption

~3 W (idle), ~4 W (peak)


11.4 cm x 16.5 cm x 3.8 cm

Instrument mass

0.6 kg

Table 15: Parameters of SCS (Sequence and Compressor System)


Mini-RF (Miniature Radio Frequency) instrument - a technology demonstration

Mini-RF was developed by an JHU/APL (Johns Hopkins University/Applied Physics Laboratory) and Navy team (PI: Benjamin J. Bussey). Mini-RF represents a significant step forward in spaceborne RF technology and architecture. It combines SAR (Synthetic Aperture Radar) at two wavelengths (S-band and X-band) and two resolutions (150 m and 30 m) with interferometric and communications functionality in one lightweight (14 kg) package. 85) 86)

Previous radar observations (Earth-based, and one bistatic data set from Clementine) of the permanently shadowed regions of the lunar poles seem to indicate areas of high CPR (Circular Polarization Ratio) consistent with volume scattering from volatile deposits (e.g. water ice) buried at shallow (0.1–1 m) depth, but only at unfavorable viewing geometries, and with inconclusive results. The LRO Mini-RF utilizes new wideband hybrid polarization architecture to measure the Stokes parameters of the reflected signal. These data will help to differentiate “true” volumetric ice reflections from “false” returns due to angular surface regolith. Additional lunar science investigations (e.g. pyroclastic deposit characterization) will also be attempted during the LRO extended mission.

The objectives of the Mini-RF instrument are:

1) Flight verification of an advanced lightweight RF technology for future NASA and DoD (Department of Defense) communications applications

2) Demonstration of a hybrid-polarity SAR (Synthetic Aperture Radar) architecture

3) Obtaining measurements of the lunar surface as a function of radar band (S and X) and resolution (150 m, 30 m) which could identify water ice deposits in the permanently shadowed polar regions

4) Production of topographic data using interferometry (S-band) and SAR stereo techniques

5) Mapping of areas of interest identified by the Chandrayaan-1 forerunner Mini-SAR experiment and other lunar instruments. Coordinated, bistatic imaging in S-band, to be compatible with the Chandrayaan-1 and the LRO spacecraft, can unambiguously resolve ice deposits on the moon.

Background: The Mini-RF payload will address key science questions during the LRO primary and extended missions. These include exploring the permanently shadowed polar regions and probing the lunar regolith in other areas of scientific interest (e.g. pyroclastic deposits). The nature and distribution of the permanently shadowed polar terrain of the moon has been the subject of considerable controversy.

The Mini-RF hardware is based on DoD communications technology and methodology. Precursor Mini-RF technology was flight-tested by NRL (Naval Research Laboratory) in the low Earth orbit on the USAF MightySat-2 and XSS-10 missions as a Space Ground Link System (SGLS).

In 2004, the DoD and NASA initiated the Mini-RF program to develop and flight-test advanced lightweight radar and communication systems and NASA elected to apply the technology to lunar exploration by building two payloads. The first, “Forerunner” Mini-SAR (Miniature SAR) instrument, was developed and integrated into the ISRO (Indian Space Research Organization) Chandrayaan-1 mission to the moon (launch Oct. 22, 2008) as a NASA guest payload and the second, on the LRO spacecraft as a technology demonstration. The Mini-SAR assembly had to operate in the lunar thermal and radiation environment, yet was simpler in design and operation, providing significant experience and reduction of risk for the more advanced LRO Mini-RF system.

In May 2006, ISRO and NASA signed a MOU in Bangalore on the inclusion of two US instruments, namely Mini-SAR and M3 (Moon Mineralogy Mapper), to be flown on the Chandrayaan-1 mission.

The LRO Mini-RF affords NASA and the DoD an opportunity to flight-qualify lightweight technology for a range of applications, including deep space communications. The flexibility, reconfigurability, and capability of Mini-RF will be demonstrated by a communications and radar mode utilizing the same hardware. The constraints of a lunar mission (range, limited duty cycle over the poles) and the low mass of advanced lightweight RF technology allows a technology demonstration which met the payload constraints of both the Chandrayaan-1 and LRO spacecraft, and provided an opportunity to collect unique and valuable lunar science data. The new technologies being qualified on LRO Mini-RF include: MPM (Microwave Power Module) based multi-frequency transmitter, wideband dual-frequency panel antenna, all digital receiver and waveform synthesizer incorporating FPGA (Field Programmable Gate Array) and analog-to-digital conversion at 1 GHz sampling.

The Mini-RF parts qualification program, which included commercial technology, allowed innovative components to gain space qualification. A comparison of the Mini-RF radar and communications performance with existing and previously flown technology, illustrating mass and performance improvements, is shown in Table 16.


SAR instrument mass

DC power input

RF power

Frequency band

(dry mass)

Orbital altitude

Launch date




1 kW


2300 kg

800 km

June 27, 1978

Magellan (Venus Radar Mapper)

154 kg

1 kW

325 W


1035 kg

249 km x 8543 km

May 4, 1989


11,000 kg

3-9 kW


L, C, X


225 km

Apr. 9, 1994
Sept. 30, 1994


8.1 kg

< 100 W

15 W


525 kg

100 km

Oct. 22, 2008


13.9 kg

~150 W

25 W


< 1000 kg

50 km

June 18, 2009

Table 16: Performance comparison of various SAR missions


Mini-RF instrument investigation and description:

The Mini-RF instrument features a new hybrid-polarity architecture, a dual-polarized system with a linearly-polarized antenna - leading to a simpler yet more capable radar. The essential feature of the hybrid-polarity architecture is: transmit circular polarization (by driving the orthogonal linear feeds simultaneously by two identical waveforms, 90º out of phase), and receive H and V linear polarizations, coherently (measurement of the 2 x 2 covariance matrix of the backscattered field). Once calibrated, the H and V single-look complex amplitude data are sufficient to form all four Stokes parameters, from which the circular-polarization ratio may be found, along with several other quantitative characterizations in the image domain. 87) 88) 89) 90) 91)


Figure 49: Schematic diagram of the generic hybrid-polarity SAR instrument (image credit: JHU/APL)

As the Mini-RF system probes the lunar regolith at two frequencies (S-band and X-band) it will provide additional information on the physical properties of the upper 1-2 m of the lunar surface. Under the proposed observational constraints, Mini-RF can identify areas of high CPR (~1), which could be caused by ice deposits. Areas that do show high CPR can be analyzed with greater sensitivity through their backscattering features. It is hypothesized that “ice” and “regolith” will have differentiable characteristics as seen through their respective Stokes parameters at two wavelengths. When supported by Chandrayaan-1 and other LRO data (e.g. neutron spectroscopy, shadow and lighting, roughness and surface texture, thermal environment), the LRO Mini-RF measurements should provide more conclusive evidence as to the likelihood that ice deposits occur in permanently shadowed areas.




Frequency bands

S-band (12.6 cm wavelength) and
X-band (4.2 cm wavelength)

2380 MHz (±10 MHz) and
7140 MHz (±10 MHz)

Instrument mass

≤ 16 kg

13.9 kg total; 8.5 kg electronics; 5.4 kg antenna

Radiation tolerance

20 kRad, SEU/SEL 75 MeV

Selective wavers

Thermal environment

-20ºC, –30ºC


DC current

< 4.8 A transmitter; < 4.7 A remainder

5.2 A transmitter; 1.7 A at 27 V
±0.25 dBm, ±0.3 dBm

RF transmitter power


42.6 dBm S-band; 41.1 dBm X-band

Operating time (duty cycle)

3 min on, 50 min off;
3 min on

10 min on, 20 min off,
10 min on (limit of test)

Transmit polarization

Circular polarization

Circular polarization

Polarization isolation

1.7 dB axial ratio

≤ 2.0 dB axial ratio, (to be verified in flight)

Receive polarization

H and V

≥ 40 dB isolation

Channel to channel power

±1 dB “knowledge”

±0.2 dB S-band


-0.1 dB X-band

±0.2 dB X-band

S-band baseline resolution

150 m azimuth x range

150 m x 150 m

S-band Noise Equivalent

-30 dB at 50 km altitude

-33.6 dB

Radar cross section



Number of looks



Range of swaths

6 km in S-band, 4 km in X-band

6 km in S-band, 4 km in X-band

S-band zoom capability

15 m x 30 m

15 m x 30 m

X-band baseline resolution

150 m

150 m x 150 m

X-band Noise Equivalent

-24 dB at 50 km altitude

-26.3 dB

Radar cross section



Number of looks



PRF (Pulse repetition Frequency)

-18 dB

-18 dB

X-band zoom capability

30 m

15 x 30 m

Radiometric stability

±0.5 dB

±0.5 dB

Communications demonstration


2380 MHz


Half duplex

Half duplex


Modulated signal

BPSK, Manchester, Variable data rate and modulation index

Received signal

Digitize signal

Continuous 500 kbit/s

Table 17: Mini-RF instrument requirements and performance (Ref. 85)

Technology demonstrations: The Mini-RF observations are made possible within the mass and power constraints imposed by LRO via application of a number of technologies. Two key technologies are a wideband MPM (Microwave Power Module) based transmitter and a lightweight broadband antenna and polarization design. The Mini-RF also has an S-band-only interferometric mode with 3.5 km wide strips with ±15 m mapping accuracy to measure lunar topography. This will be the first demonstration of interferometric SAR techniques in a planetary mission. The Mini-RF antenna architecture is shown in Figure 49. The H and V right circular polarization components are transmitted coherently, which are then reconstructed as Stokes’ parameters during the data processing step.

Both the communications and the radar astronomical objectives impose a requirement for circular polarization on the transmitted field. Conventional radar that would measure CPR (Circular Polarization Ratio) then would have to be dual-circularly polarized on receiver. The hybrid-polarity approach provides weight savings by eliminating circulator elements in the receiver paths, which reduces mass, increases RF efficiency, and minimizes cross-talk and other self-noise aspects of the received data. The H and V signals are passed directly to the ground-based processor. It is well known that the Stokes parameters comprise a full characterization of the backscattered field.

The values of the four Stokes parameters do not depend on the choice of receiver polarization, so this architecture minimizes hardware while maintaining full science value. The result provides significant advantages over the conventional “CPR-measuring” dual-circular-polarized approach, yet the radar is simpler. The use of possible Stokes parameter-based products (e.g. CPR, degree-of-depolarization, degree-of-linear-polarization, phase “double bounce”) have a number of significant advantages over traditional radar systems: less hardware is needed, resulting in fewer losses and a “cleaner,” simpler flight instrument. The signal levels are comparable (within ~2 dB) in both channels allowing relatively relaxed specifications on channel-to-channel cross-talk and more robust phase and amplitude calibration. The processor has a direct view through the entire receiver chain; including the antenna receives patterns and other radar parameters (e.g., gain and phase). These parameters are applied in processing “Levels” (Level 0, 1) which correspond to successive data processing stages, as shown in Figure 49.

The design allows selective Doppler weighting to maximize channel–channel coherence (e.g., reduce the H & V beam mismatch). As CPR is less sensitive to channel imbalance by at least a factor of 2 with respect to explicit RCP/LCP, Stokes parameter-based backscatter decomposition strategies can help distinguish “false” from “true” high CPR areas (e.g., analysis of “m-δ” feature space (Ref. 87).

Mini-RF instrument: The Mini-RF Instrument is comprised of the following elements: (1) antenna, (2) transmitter, (3) digital receiver/quadrature detector waveform synthesizer, (4) analog RF receiver, (5) Control Processor, (6) interconnection module, and (7) supporting harness, RF cabling, and structures. The functional block diagram is shown in Figure 50 while its layout is shown in Figure 51.


Figure 50: Functional block diagram of the Mini-SAR instrument (image credit: JHU/APL)


Figure 51: Mechanical configuration of the Mini-RF instrument (image credit: JHU/APL)

Antenna: An “egg crate” antenna (Figure 52) allows a broadband, dual-frequency design with a single antenna panel, without any deployable mechanisms (e.g. feeds) while also meeting stringent weight and volume constraints. The elements are sized to allow a 3:1 frequency range. Each element incorporates radiators and physical phasing combines their power. The thermal design, materials selection, manufacturing, and test qualification heritage of the single-frequency Chandrayaan-1 Mini-SAR antenna was applied to the dual frequency LRO Mini-RF unit. Because of this heritage, the Mini-RF antenna is robust and lightweight (4 kg) while satisfying all technical requirements.


Figure 52: Illustration of the Mini-RF antenna design (image credit: JHU/APL)

Transmitter: The LRO Mini-RF transmitter (Figures 53, 54) takes full advantage of the capabilities of the wideband antenna. The transmitter is the first implementation of the MPM (Microwave Power Module) technology on a long-duration spaceflight, which affords a significant breakthrough in available bandwidth and power efficiency with reduced mass as compared to previous TWT (Traveling Wave Tube) systems. The MPM combines a solid state RF driver/preamplifier with a traveling wave tube amplifier, a hybrid approach combining the advantages of both solid state and vacuum electronic technology. Flight-testing the MPM technology is a major goal of the Mini-RF demonstration. The MPM is enabling in giving Mini-RF its dual-band capability within the challenging mass, power, and volume constraints of the LRO spacecraft (Ref. 85).


Figure 53: Functional block diagram of the Mini-RF transmitter (image credit: JHU/APL)


Figure 54: Illustration of the MPM/TWT (Microwave Power Module/Traveling Wave Tube), image credit: JHU/APL

IM (Interconnect Module): The IM combines and splits the RF energy and serves as the interface between the transmitter, receiver, and antenna. Its design specifically handles issues such as multipaction using selected materials and geometry.

Mini-RF calibration: Laboratory calibration data was acquired during spacecraft integration and test. The overarching goal of these activities was to insure production of a calibrated instrument. All waveforms in the waveform table were tested on brassboard hardware while selected waveforms were tested on flight hardware. This waveform testing is inherent in the overall Mini-RF integration and test program. Additional waveform testing was conducted on the flight instrument during thermal vacuum temperature ramp cycles. Internal calibration data are acquired every time that Mini-RF takes a data collect; a chirp, noise, and tone calibration is done both immediately before and immediately after a data collect. No end-to-end range tests were possible during integration and test, which necessitated the use of in-flight external calibration.

External calibration is planned in-flight by in conjunction with ground-based assets at the Greenbank and the Arecibo Radio Telescopes. These measurements will include polarization purity or axial ratio and antenna pattern. A transmitted signal from the LRO Mini-RF is received by Greenbank while the antenna pattern is scanned over a range of angles. Specifically, the scan will be ±12º from boresight in both elevation and azimuth, sampling at 0.5º increments. At each orientation, mini-RF will transmit for approximately 40 ms. Subsequently, each axis (azimuth or elevation) of the antenna will be parallel to the Earth’s equator, with the boresight pointed towards Greenbank. The antenna will then be scanned parallel to the Earth’s equator, at 0.4º/s 12º in one direction, then back to boresight, then 12º in the other direction, then back to boresight. During scanning, Mini-RF will transmit for 40 ms every 1.25 seconds, corresponding to an angle change between transmits of 0.5º. The scan should take approximately two minutes to complete. An S-band received calibration will also be conducted using signals transmitted from Arecibo following the same geometry as the transmit calibration.


Figure 55: High-level block diagram of the Mini-RF calibration methodology (image credit: JHU/APL)



Figure 56: Artist's rendition of Mini-RF imaging (image credit: NASA, JHU/APL)

Changing operational requirements for Mini-RF during the mission:

The Mini-RF instrument data has proved to be extremely useful to the project - resulting in requests for considerably more operational time during the mission than was originally planned. In fact, the total data volume for the primary LRO mission increased three hundred-fold from 200 GB to 60 TB - causing in particular operational problems for the POC (Payload Operations Center) at JHU/APL to scale up to the increased demand for science data. The success of the operations team has enabled Mini-RF to support LCROSS targeting and to continue mapping large portions of the Moon at both the polar and equatorial regions. More than 40 TB of Mini-RF science data have been delivered to the PDS (Planetary Data System) as of December 2010.

The Mini-RF team is currently (2011) supporting the LRO two-year extended mission that began on September 15, 2010, throughout which it will continue to gather data, publish results, and support the Lunar Reconnaissance Orbiter.The Mini-RF project anticipates future PDS archive deliveries to be at the same approximate rate while LRO continues to operate in its elliptical orbit. Mini-RF will continue taking non-polar data until the LRO solar arrays are parked due to the high Sun angle. During those semi-annual periods, Mini-RF will renew intense campaigns to collect polar data on every orbit, and add non-polar data when downlink capacity permits. 92)

The two high-data-rate instruments onboard LRO are the LROC (LRO Camera) and Mini-RF. Since Mini-RF was baselined as a technology demonstration with minimal observing time, the spacecraft design connected both instruments to the same SpaceWire bus to send data to the SSDR (Solid State Data Recorder). The design of the light software that managed data on the bus precludes both instruments from operating simultaneously. This was not anticipated to be an issue due to the limited operational time of Mini-RF in the mission baseline plan.

LROC acquires less data during the nighttime portion of the orbits, while Mini-RF radar can operate in the dark. This situation, along with the fact that the high data rate downlink capacity exceeds that required by LROC and the other instruments, allows Mini-RF to collect data at night. This is only done during orbits that include a high data rate downlink using LRO’s White Sands ground station Ka-band capability.


Figure 57: North (top) and South (bottom) polar mosaics show radar brightness (left) and the circular polarization ratio (right) taken by Mini-RF on LRO (image credit: NASA)

Figure 57 shows some of the images collected by Mini-RF from the LRO spacecraft. These mosaics cover from 70º to the pole for both the north (top) and south (bottom) polar regions. The left-hand images show radar brightness while the right-hand color images also show the circular polarization ratio. Over 500 images collected during the June 2010 portion of the polar campaign are combined into the mosaics shown.



LCROSS (Lunar CRater Observation and Sensing Satellite)

LCROSS is a separate secondary payload spacecraft of NASA/ARC which will be launched on the same Atlas-Centaur rocket (Atlas V 4001) as LRO. After the orbiter (LRO) separates from the Atlas V launch vehicle for its own mission, the LCROSS system will use the spent Centaur upper stage of the rocket as a 2,300 kg lunar impactor, targeting a permanently shadowed crater near the lunar South Pole.

The LCROSS concept was selected for flight by NASA in April 2006 (critical design review in Feb. 2007). The objective of LCROSS is to advance the Vision for Space Exploration (VSE) by identifying, with a high probability of success, the presence of water ice at the moon's south pole. LCROSS carries a 2,300 kg Kinetic Impactor that creates nearly a 1000 metric ton plume of lunar ejecta on impact. This powerful impact is achieved by steering the entire launch vehicle's spent Earth Departure Upper Stage (EDUS) into a crater at the lunar south pole. According to estimates, the Centaur's collision with the moon will excavate about 220 tons of material from the lunar surface. 93) 94) 95) 96)

The scientific basis for the LCROSS mission had roots in the Clementine (1994) and Lunar Prospector (1998) missions which performed complementary forms of resource mapping. This mapping led the lunar science community to conclude that there might be water-ice trapped in permanently-shadowed craters on the moon.

If successful, the LCROSS mission would conduct the first in-situ study of a pristine, permanently shadowed lunar crater and would:

• Confirm the presence of water ice in a permanently shadowed region

• Determine the nature of hydrogen signatures detected at the lunar poles on the previous lunar missions, Clementine and Lunar Prospector

• Determine the amount of water, if present, in the lunar regolith or soil

• Determine the composition of the lunar regolith.


Figure 58: Accommodation of the LRO and LCROSS spacecraft in the launch fairing (image credit: NASA)


LCROSS is a bare-bones spacecraft designed to use cameras and spectrometers to watch its 2200 kg upper stage slam into hydrogen-rich Shackleton Crater. The LCROSS Probe, is referred to as S-S/C (Shepherding Spacecraft) with a mass of about 700 kg. On approach to the moon, the Shepherding Spacecraft will position the upper stage for a precision impact, then separate and perform a braking maneuver in order to observe the upper stage's impact into the moon. Sensors will observe and monitor the debris plume, searching for water ice or vapor. Shortly after the Centaur impact, the Shepherding Spacecraft will also impact the moon, creating a second smaller plume. 97) 98) 99)


Figure 59: Illustration of the shepherding spacecraft (image credit: NASA/ARC)

LCROSS is a rapid response mission (26 months to delivery), the NASA/ARC contract for the spacecraft was awarded to NGC (Northrop Grumman Corporation) in 2006. Since LCROSS is a secondary payload to LRO, an ESPA (EELV Secondary Payload Adapter) ring is being used as the interface to the EELV upper stage and the primary payload, LRO. In effect, the ESPA ring serves as the LCROSS spacecraft structure (Figures 59 and 61). The ESPA ring functions as a multifunctional integrating element which supports the LRO adapter; contains an independent set of avionics; a small 344 kg capacity monopropellant-propulsion system, a single-panel body mounted solar array and battery; and mounts for the impact observation instruments, two S-band omni antennas, and 2 medium-gain horns.

The body-mounted solar array is structurally designed to be extremely high frequency and uses a large, 12.5 cm thick honeycomb, ESPA-ring mount. The solar array is sized to provide 650 W with the S-S/C and Centaur stack pointed in a ± 10º ACS (Attitude Control System) dead band to the sun. Standard 28% multi-junction solar cells are used in the array. With the instruments on and transmitting telemetry, the battery system (four 20 Ah batteries) provides nearly 2 hours of operation without charging from the solar array.

The ACS consists of a STA (Star Tracker Assembly), MIMU (Miniature Inertial Measurement Unit), a CSSA (Coarse Sun Sensor Assembly), and the PDE (Propulsion & Deployment Electronics). The ACS is based primarily on LRO hardware and software in the same single strung arrangement. Actuation was provided by a set of eight monopropellant 5 N thrusters. Two additional 22 N thrusters provided orbit maneuvering capability. The ACS featured twelve control mode/submode combinations, six tailored for specific operations while attached to the Centaur, and a second set for use after Centaur separation. The LCROSS propellant tank contained just over 305 kg of hydrazine for both attitude control and orbit maneuvering.

Use of a RAD750-based single-board computer, communications card, and power and thruster control electronics. Onboard communications employ mixed SpaceWire and MIL-STD 1553 buses. The LCROSS flight software is derived heavily from software on previous programs, including EO-1 and WMAP.


Figure 60: The LCROSS spacecraft employed a novel use of an ESPA ring (image credit: Northrop Grumman)


Figure 61: Alternate view of the shepherding spacecraft with its elements (image credit: NASA/ARC)


Figure 62: Configuration of the LCROSS system (image credit: NASA)


Figure 63: Artist's rendition of LCROSS (image credit: NASA)

RF communications: The LCROSS communication baseline system (S-band) is single-strung (two omnidirectional antennas and two medium-gain antennas) and can deliver 1.5 Mbit/s real-time data from the moon to the DSN (Deep Space Network) 70 m dish using one of the two medium gain horn antennas, or can deliver 40 kbit/s using one of the two omni antennas, pointed ±30º from Earth. The S-S/C uses the existing LRO transponder (along with other LRO passive microwave components). At least one of the three DSN sites has visibility to the spacecraft at all times.



Instruments of S-S/C:

The LCROSS science payload, developed at NASA ARC, combined processing and control electronics DHU (Data Handling Unit) with nine instruments to aid in water detection. The DHU accommodated all sensor interfaces, all digital video system functionality and all interfaces with the S-S/C avionics. The instrument package comprised 5 cameras (1 visible, 2 NIR, 2 Mid IR), 3 spectrometers (1 visible, 2 NIR) and one photometer with a total mass of 12.4 kg, a power consumption of 27 W, and a total data rate of 554.5 kbit/s.

Eight of nine instruments were co-aligned along the S-S/C longitudinal axis and provided nadir-pointed sensing during the Centaur impact event. One of two near-infrared spectrometers was side-pointed to provide spectra of sunlit material rising in the Centaur ejecta plume. A spring-loaded cover protected the nadir-looking instruments from direct sun exposure during launch and through the early part of the mission.

Visible and NIR (Near Infrared) cameras (3). The objective is to: 1) observe the impact of EDUS, and 2) observe ejecta cloud morphology and evolution. For the visible sensor, a high-end broadcast-quality CCD video camera is being used outputting PAL format (752 H x 582 V pixels). FOV= 6º, resolution < 0.5 km.

Mid-IR imagers (2). The objective of these two cameras (in 2 wavelengths: 7 and 12.3 µm) is to look down on the permanently shadowed lunar surface to map pre-impact terrain (warmer vs cooler = rocks vs regolith), thermal evolution of plume (dependent upon H2O vapor concentration in plume), ejecta blanket, and freshly exposed regolith. The baseline mid-IR sensors will be a flight-proven alpha-silicon uncooled micro-bolometer, most sensitive in the 7-14 µm spectral range, the data output is in PAL format (384 H x 288 V pixels).


Figure 64: View of the visible and IR imagers (image credit: NASA)

NIR spectrometers (2 COTS instruments). The objective is to monitor spectral bands (every second) associated with water vapor, ice, and hydrated minerals in NIR (1.35-2.45 µm, ~0.01 µm spectral resolution) covering the first overtones of H2O ice (band is free of interference, more brightly illuminated by sunlight than fundamentals near 3 µm).

The regions near 1.4 and 1.9 µm (usually obscured by Earth's atmosphere) also provide sensitive indication of water vapor from ice, shape of band may provide info regarding nature of ice crystals and mineral hydrate. Broad minima at 1.5 and 2.0 µm are indicative of water ice. Resolution: 1 km. The two identical NIR spectrometers are being coupled with fiber optics to telescopes, one focused along the impactor trajectory, the second aimed laterally through the plume towards the limb during the last ten seconds before S-S/C impact.


Figure 65: View of the NIR spectrometer (image credit: NASA)

Visible total luminance diode (1). Broadband from 0.4 - 0.9 µm, sample rate >100 Hz, power: < nW NEP @ 100 Hz. The goal is to observe the impact flash.

- Light flash due to thermal heating and vaporization

- Shape of the flash's light curve can be used to determine certain initial conditions of the impact

- Flash peak intensity depends on impact velocity angle, target & projectile types.

Sequence of events:

• After launch, the LCROSS spacecraft will arrive in the lunar vicinity independent of the LRO satellite. On the way to the moon, the LCROSS spacecraft's two main parts, the S-S/C and the Centaur Upper Stage, will remain coupled.

• As the spacecraft approaches the moon's south pole, the Centaur (EDUS) will separate, and then will impact a crater in the moon's polar region. The impact speed is estimated to be ~2.5 km/s, and a resulting moon crater of size 30 m in diameter and 4.8 m in depth is expected - tossing up about 200 tons of lunar debris.

• A plume from the Centaur crash will develop as the S-S/C heads in towards the moon. The S-S/C (mass of 700 kg) will fly through the plume, and six instruments (cameras and spectrometers) on the spacecraft will analyze the cloud to look for signs of water and other compounds.

• About 15 minutes after the upper stage booster's impact the S-S/C will also crash into the crater floor of the moon

• In addition, spaceborne and earth-based instruments will be pointed to the moon's south pole to study the huge plume, which scientists expect to be larger than 200 metric tons.



LCROSS mission status:

• LCROSS was cast as a Class D mission, which means it can accept more technical risk than other mission types in NASA. So why have “freeboard”, a.k.a. Extra technical margin? One of the ways that LCROSS kept its risk in check was by keeping complexity as low as possible while satisfying project requirements. To address the remaining complexity in the design, having technical performance measures which have a fair amount of margin could be invaluable. This extra margin is a commodity that can be used in many different and sometimes unplanned ways during the mission. Extra fuel, power, thermal, or RF link can provide operational degrees of freedom when anomalies are encountered. 100)

The key to capabilities-driven, cost capped missions like LCROSS is to keep it simple and to manage the risk. It is not about eliminating risk, which is very costly. It is about managing risk to a level commensurate with project programmatic constraints. LCROSS did this by making use of existing investments by the Agency, existing commercial hardware, and being sufficiently creative to see opportunities to buy-down risk.
Ultimately, LCROSS succeeded because the individuals and organizations in the LCROSS team walked and shared road on a mission to the moon and worked together to make it succeed. Each party on this team had both mutual and self-interests for why they wanted to participate. The Agency wanted to show there was an effective way to make use of excess launch capability and to work cheaply; NASA/ARC wanted to show it was able to run small, fast-paced, light-weight missions; NG (Northrop Grumman) wanted to show that it could be nimble and carve out a new market for itself; and the commercial sector found an onramp to space and lunar applications which could propel their businesses into a new market. One of the great successes of LCROSS was aligning each the team member’s needs into a common purpose which benefited everyone in a win-win-win scenario.

Table 18: LCROSS programmatic summary (Ref. 96)

• In Nov. 2009, preliminary data of LCROSS indicate that the mission successfully uncovered water during the Oct. 9, 2009 impacts into the permanently shadowed region of Cabeus crater near the moon’s south pole. 101)

• On October 9, 2009, the LCROSS spacecraft was slammed into a crater near the lunar south pole. No light flash was visible in the thermal images broadcast on NASA television, as the 2.3 ton rocket impacted the Cabeus crater at 11:31 UTC. A second shepherding spacecraft flew through the debris plume, collecting and relaying key data back to Earth before it too plowed into the lunar surface, according to NASA. The LCROSS mission is hoping to uncover whether there is water or ice below the moon's surface that could be used by astronauts on future space missions. 102) 103)

Even without big explosions or bright plumes of ejecta, for all intents and purposes it appears LCROSS's impact on the moon was a smashing success. While the mainstream media and the public seemed disappointed in the lack of visual data, mission managers said the mission has garnered plenty of spectroscopic data, and that's where the real science can be found.

• Centaur separation was performed successfully 9 hours 40 minutes prior to Centaur impact (Ref. 97). One minute following separation, the S-S/C flipped 180º to point the payload at the receding Centaur. The spacecraft payload was activated to transmit imagery of the Centaur for 15 minutes (via 70 m DSN antenna), to determine whether the separation had caused the Centaur to tumble. Forty minutes after separation, the S-S/C performed the Braking Burn, a ΔV maneuver used to induce a four-minute delay between the Centaur and S-S/C impact events (598 km range at Centaur impact).

• On Sept. 9, NASA selected the target crater for lunar impact. LCROSS is racing toward a double-impact on the moon at 7:30 am EDT on Oct. 9, 2009. The target crater is Cabeus A. It was selected after an extensive review of the best places to excavate frozen water at the lunar south pole. 104) 105)


Figure 66: Illustration of the impact crater region around the lunar south pole (image credit: NASA)

• On June 23, 2009, LCROSS successfully completed its most significant early mission milestone with a lunar swingby and calibration of its science instruments. With the assist of the moon's gravity, LCROSS and its attached Centaur booster rocket successfully entered into polar Earth orbit. The maneuver puts the spacecraft and Centaur on course for a pair of impacts near the moon's south pole on Oct. 9, 2009. 106)


Figure 67: Artist's view of LCROSS EDUS ready to separate from S-S/C (image credit: NASA)


Figure 68: LCROSS/EDUS heading-in with S-S/C in the foreground (image credit: NASA)



Figure 69: LCROSS plume developing with S-SC looking down and outward prior to its own impact (image credit: NASA)



LRO ground segment:

LRO’s routine support requirements include: 107) 108)

• 30 minutes of S-band tracking per 113 minute lunar orbit

- Range and range rate measurements

- Commanding

- Realtime housekeeping telemetry

• 600 Gbit per day of Ka-band downloads

- Recorded science data

- Recorded housekeeping telemetry

- CCSDS CFDP protocol with loop closed via S-band


The LRO ground segment is comprised of the following elements:

• Mission Operations Center & Flight Dynamics Facility at GSFC

• Primary Ground Station at White Sands (Ka-band and S-band)

• Global S-band TT&C provided by NASA GN & SN

• SOC (Science Operations Centers) at PI institutions

• S-band tracking augmented by laser ranging system to improve accuracy.

Routine Operations Network:

The WS1 (White Sands 1) station will provide the Ka-band download service as well as S-band coverage for all of the LRO orbits visible from White Sands (approximately 45% of all LRO orbits).

Note: In Nov. 2007, NASA/GSFC showcased the new operational 18 m near-Earth Ka-band antenna network (a three antenna network), the first in NASA history, during a ribbon cutting ceremony (Nov. 8 2007) at the White Sands Complex, N.M. White Sands was chosen as the location for the new antennas because of the existing infrastructure available there, making it a cost effective option. Two of the three antennas will be used to accommodate the continuous high volume data stream of SDO (Solar Dynamics Observatory). The third antenna will be used for LRO and will have the highest data volume stream ever received from a lunar spacecraft. 109)

A five station network (WS1, Dongara, Weilheim, Kiruna, and Hawaii) provides nearly complete S-band coverage above 5º elevation with 81% multistation coverage for scheduling flexibility.

• LRO S-band support consists of alternating 56 minute view / no view periods for TT&C functions.

• Ka-band support consists of at least four 56 minute views per day from WS1. Ka-band utilization is approximately 61% of capacity.


Figure 70: Overview of the LRO ground segment (image credit: NASA)

To make the LRO observations accessible to both human exploration planners as well as the science community, calibrated LRO data will be rapidly deposited into the PDS (Planetary Data System) by each instrument SOC (Science Operations Center). Figure 71 illustrates the locations where the various LRO observational data and data products will be located. Each instrument team is required to submit to the PDS several types of data products. Typically Level 0 data products are raw data in the form of counts accumulated during specific time intervals of measurements, with orbital information and engineering parameters. Higher level data products consist generally of count rates converted into physical units and projected onto the lunar surface as maps of geophysical quantities.

Each instrument SOC is required to deliver validated and calibrated data to the PDS every three months, starting six months after the beginning of the primary Exploration Mission (i.e., March 15, 2010). The initial delivery will include all measurements made from the time of launch through the first three months of the primary orbit (e.g., June 18 to December 15, 2009). Subsequent deliveries will be made every three months after the first delivery and will include all data that is no more than three months old. After the Exploration Mission ends, some investigations plan to reprocess data and develop higher level composite data products. These will be delivered to the PDS no later than six months after end of the Exploration Mission.


Figure 71: LRO SOCs (Science Operation Centers) and PDS Nodes and Data Nodes (NASA/GSFC, Ref. 24)

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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.