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)

Spacecraft:

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 ST (Star Trackers), and a RLG (Ring Laser Gyroscope). 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⁾

The EPS (Electric Power Subsystem) is comprised of an articulated solar array (1 wing, 2-axis tracking), a Li-ion battery, and a DET system (21-35 V). 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.

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. ⁷⁾

Figure 2: 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 3. 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. ^{8) 9) 10)}

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 3: 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 4: 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 5: Design concept of the LRO spacecraft (image credit: NASA)

Figure 6: 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. ^{11) 12)}

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. ¹³⁾

• 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 2 hour period. The orbital velocity is 1.6 km/s. LRO stays on near side of moon ~ 1 hour out of every two.

Figure 7: 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: ¹⁴⁾

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 8: Schematic view of the projected LRO mossion timelime (image credit: NASA)

LRO mission status:

• 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. ^{15) 16)}

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 9 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 9: 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 10). 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. ¹⁷⁾

Figure 10: 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. ¹⁸⁾

Figure 11: 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. 13).

• 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. ^{19) 20)}

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 12: 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. ²¹⁾

• 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. ^{22) 23)}

Figure 13: This image shows the daytime and nighttime lunar temperatures recorded by DIVINER (image credit: NASA/UCLA) ²⁴⁾

• On Sept. 15, 2009, a final orbital maneuver put LRO into its science and mapping orbit of 50 km altitude above the lunar surface. 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. 13).

Figure 14: LROC image of the Apollo-12 landing site taken in Aug. 2009 (image credit: NASA)

Legend: Figure 14 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. ²⁵⁾

• 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. 13).

• 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. ²⁶⁾

• 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. ²⁷⁾

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. ²⁸⁾

Table 2: Summary of LRO instrument mass and power allocations

Table 3: Overview of the LRO instrument complement

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

CRaTER PI: H. E. Spence, Boston University (BU), Boston, MA. 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.

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 17), 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. ^{29) 30)}

Table 4: Parameters of CRaTER

Figure 15: Detailed view of the CRaTER telescope (image credit: BU)

Figure 16: Illustration of detector location in CRaTER (image credit: University of Tennessee) ³¹⁾

Legend to Figure 16: 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 17: Illustration 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. ³²⁾

Figure 18: 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.

Table 5: 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: ³³⁾

• 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.

Table 6: 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 19: Schematic view of the LAMP instrument (image credit: SwRI)

Figure 20: 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: ^{34) 35) 36) 37) 38)}

• 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 21: Conceptual view of the cosmic ray induced neutron flux on the lunar surface (image credit: NASA)

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 22: 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 23: Schematic view of LEND collimator detection concept (image credit: IKI)

Table 7: Overview of some performance parameters of the LEND instrument

Figure 24: 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: ^{39) 40) 41)}

- 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 25: 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.

Table 8: 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 27). The 5-spot pattern enables the surface slope to be derived in the along-track and across 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.

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

Figure 27: Schematic illustration of the LOLA instrument (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 28: LR (Lunar Ranging) flight system components (image credit: NASA)

Figure 29: LR (Lunar Ranging) operations overview (image credit: NASA)

Figure 30: Simplified LOLA/LR block diagram (image credit: NASA)

Resulting products:

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

2) 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.

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: ^{42) 43) 44)}

- 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 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.

Table 9: Specification of the NAC devices

Figure 31: View of the LROC NAC device (image credit: ASU)

Figure 32: View of the LROC WAC device (image credit: ASU)

Figure 33: NAC optics cutaway (left) and NAC optics and electronics (image credit: ASU)

Table 10: Parameters of the WAC instrument

Figure 34: Detailed view 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.

Table 11: 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: Ben 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. ^{45) 46)}

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 12.

Table 12: 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. ^{47) 48) 49) 50) 51)}

Figure 35: 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.

Table 13: Mini-RF instrument requirements and performance (Ref. 45)

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 35. 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 35.

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. 47).

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 36 while its layout is shown in Figure 37.

Figure 36: Functional block diagram of the Mini-SAR instrument (image credit: JHU/APL)

Figure 37: Mechanical configuration of the Mini-RF instrument (image credit: JHU/APL)

Antenna: An “egg crate” antenna (Figure 38) 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 38: Illustration of the Mini-RF antenna design (image credit: JHU/APL)

Transmitter: The LRO Mini-RF transmitter (Figures 39, 40) 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. 45).

Figure 39: Functional block diagram of the Mini-RF transmitter (image credit: JHU/APL)

Figure 40: 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 41: High-level block diagram of the Mini-RF calibration methodology (image credit: JHU/APL)

Figure 42: Artist's rendition of Mini-RF imaging (image credit: NASA, JHU/APL)

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. ^{52) 53) 54) 55) 56)}

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

Spacecraft:

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. ^{57) 58)}

Figure 44: 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 44 and 45). 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 45: Alternate view of the shepherding spacecraft with its elements (image credit: NASA/ARC)

Figure 46: Configuration of the LCROSS system (image credit: NASA)

Figure 47: 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 48: 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 49: 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:

Table 14: LCROSS programmatic summary (Ref. 56)

• 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. ⁵⁹⁾

• 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. ^{60) 61)}

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. 57). 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. ^{62) 63)}

Figure 50: 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. ⁶⁴⁾

Figure 51: Artist's view of LCROSS EDUS ready to separate from S-S/C (image credit: NASA)

Figure 52: LCROSS/EDUS heading-in with S-S/C in the foreground (image credit: NASA)

Figure 53: 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: ^{65) 66)}

• 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

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. ⁶⁷⁾

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

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