Instrument Host Overview    =========================    For all LRO experiments, data are collected by science instruments on    the spacecraft and then relayed via the spacecraft telemetry system    to stations of the Space Communications Network (SCN). The following    sections provide an overview of the spacecraft, the science    instruments, and SCN ground system.    Instrument Host Overview - Spacecraft    =======================================    The Lunar Reconnaissance Orbiter is designed for a one-year base mission    with a goal of an extended mission of up to four additional years.  LRO    will be launched on an Atlas 5 401 Launch Vehicle along with its    companion spacecraft, the Lunar CRater Observation and Sensing Satellite    (LCROSS), into a direct insertion trajectory to the Moon.  The on-board    mono-propellant hydrazine propulsion system will be used to capture into    a polar orbit at the Moon, timed to obtain an orbit plane that enables    optimum lighting conditions in polar regions during summer and winter    seasons.  Additional burns will circularize the orbit and maintain it    during the base mission at an altitude of 50 +/- 20 km.  A    lower-maintenance, elliptical orbit of 30x216 km will be used for    commissioning and may be used for the extended mission.  The spacecraft    carries enough fuel to provide over 1300 m/s of velocity change (delta    V) for orbit capture and maintenance.    LRO Description    ===================    The orbiter is a 3-axis stabilized, nadir-pointed spacecraft, designed    to operate continuously during the primary mission.  Four reaction    wheels provide attitude control to 60 arc sec and momentum storage of up    to 2 weeks, with thrusters providing momentum dumping once per month.    Two star trackers and an inertial reference unit provide attitude    knowledge of 30 arc sec.  Coarse sun sensors provide attitude    information in contingency modes, to enable and maintain proper attitude    with respect to the sun, keeping the spacecraft power positive and    thermally stable.    A 10.7 square meter solar array provides 1850 W end-of-life during the    sunlit portion of the orbit.  An 80 A-hr Lithium-ion battery maintains    the bus voltage and provides operational power during the orbit eclipses    and survival power during the rare, long eclipses of the sun by the    earth.  The power electronics distributes the raw 28+7 V to the    instruments and the spacecraft bus electronics, delivering over 800 W    average power each orbit.    The flight computer is a RAD-750 processor executing at 133 MHz.  Two    100 Gbyte recorders store science data for playback to the earth at 100    Mbps through a 40 W Ka-band transmitter and high-gain antenna.  An    S-band system provides command, engineering telemetry, and navigation    functions.  Laser ranging capability provides ~10 cm position precision    during four one-hour passes per day.  These data, when combined with    lunar measurements from LOLA, will improve the orbit determination    capability of LRO.    The thermal control system utilizes heat pipes to spread heat and move    it to the zenith-facing radiators.  A modular structure design enables    parallel assembly of the spacecraft.    The total mass of the observatory is less than 949 kg dry and 1846 kg    fully fueled.    Orbiter Features    ----------------    The LRO Orbiter has the following key features:    1. Single string with selected redundancy    2. C&DH hosts flight software, handles data interfaces    3. PSE controls battery charging and power distribution    4. PDE controls thrusters, deployment devices, and inhibits    5. S-band transponder and Ka-band transmitter support radio links    6. 2-axis gimbals on array and high-gain antenna track sun and earth    7. Omni antennas provide continuous receive capability    8. Standard interfaces       a. 1553       b. Spacewire    9. Modular Assembly       a. Propulsion module, instrument module, avionics module       b. Parallel development    Spacecraft Subsystems    ---------------------    The spacecraft subsystems are summarized briefly below.    Spacecraft Bus    ---------------    The LRO spacecraft bus consists of the following components:    1. Primary Structure (aluminum)       a. Heating/Cooling panel       b. Edge members, fittings, and spools       c. High Gain Antenna System (HGAS) support brackets       d. Solar Array Substrates (SAS) stanchions       e. Reaction wheel assemblies brackets    2. Instruments       a. CRaTER       b. DLRE       c. LAMP (on optical bench)       d. LEND       e. LOLA (on optical bench)       f. LROC (on optical bench)       g. MRF    3. Avionics Components       a. Reaction wheels       b. Battery       c. S-band OMNI antenna       d. Coarse Sun Sensors (CSSs)    High Gain Antenna System    -------------------------    The High Gain Antenna System (HGAS) includes a deployment system with    two latches requiring mechanical release, and three restraint areas.    The articulation system is a two axis gimbal system capable of slightly    greater than 180 degrees of rotation about each axis, and two rotating    cable wraps.  RF components include Ka-band antenna with an S-band    patch antenna, an S-band coax cable and a Ka band waveguide, and rotary    joints in each of two gimbal actuators. A small laser ranging    telescope is mounted on the wave guide with fiber optic cable providing    the connectivity to the LOLA instrument. Thermal control is provided    through blanketing, T-stats, heaters, and radiators dedicated to the    HGAS system.    The latching release assembly utilizes a 400 lb preload, to include a    1.3 gapping factor of safety. Release begins with actuation of the    non-explosive actuator (NEA) which initiates the deployment sequence.    The principle of operation of the NEA involves a tensile load reacting    against a spring-wound split spool. Upon fuse wire initiation, the    spring unwinds, allowing the halves of the spool to separate.  The slow    release of the tensile load minimizes the release of strain energy and    therefore minimizes shock.    Propulsion System    -----------------    The propulsion system has been designed to provide mid-course transit    corrections after separation from the launch vehicle, lunar orbit    capture, and station keeping for the remainder of the mission. The    propulsion system is a monopropellant hydrazine system. Fuel load is    894 kg of hydrazine (~ 1300 m/sec delta-V capability) in two identical    titanium diaphragm propellant tanks (40 in OD oblate spheroid TDRSS    type tanks in TDRSS configuration)    The system includes twelve dual coil catalytic hydrazine thrusters, four    of which are on-axis 80 Newton class insertion thrusters located around    the spacecraft center of gravity (in the x-axis). Eight canted 20    Newton class attitude control thrusters provide attitude control, lunar    orbit maintenance maneuvers, and momentum dumping.    Isolation valves with redundant coils are used to isolate thruster banks    in the event of a thruster failure. Flow control orifices prevent    water hammer surges.    The Helium Pressure Regulated System is a 4200 psi COPV Helium    pressurant tank (17 in outside diameter by 29.6 in Length ). A two    stage regulator (single fault tolerant) is set at 270 psi nominal.    Redundant pyro valves provide for high pressure isolation during ground    and launch operations with a high pressure latch valve to isolate high    pressure source during mission  operations.    Electrical Power System    -----------------------    The LRO Electrical Power System (EPS) provides the following functions:    1. Load power requirements for all nominal mission modes    2. Battery charging and control    3. Interface to the C&DH subsystem for power system performance,       monitoring, configuration and control    4. Power interface to the electrical subsystem    5. Capability to respond to Special Commands    6. Power distribution and load switching capability    7. External power interfaces for I&T, ground, and pre-launch operations    8. Bus protection by automatically shedding loads if a defined fault       condition occurs    9. Restoration of power to shed loads shall be remotely controllable via       ground command.    The power system contains three sub-elements: the solar array, Power    Sub-system Electronics (PSE), and the lithium-ion battery.    Solar Array System (SAS)    ------------------------    Three low-CTE graphite composite solar array panels are contained within    one articulated wing. The system is mounted on flexures to aluminum    spacecraft structure to allow for differential thermal growth. Four    NEAs are used to release the array.    Articulation is provided through a two-axis gimbal with four panel    hinges, two at each hinge line connecting center panel to two outer    panels. When stowed, cells of the central panel face outward. For    deployment, panel springs (less than 10 lb of stored force) are driven    with viscous dampers to dissipate energy.    The solar array mounts to a single structure. Benefits of such a design    include straightforward integration, alignment, and testing. Further,    the  design provides for reduced and less-sensitive loads on gimbal    bearings, while structural loads are carried at cup/cone fittings    rather than through more-flexible gimbal paths.    Deployment is initiated through a software command delivered to the    propulsion and deployment electronics (PDE), which then fire    non-explosive actuators (NEA's) with a design similar to that shown for    the high gain antenna. Separation nuts release rods that restrain the    array in the stowed configuration. Restraining bolts are pulled by    strings to the outer panel and captured panel-to-panel latches prevent    panels from immediately unfurling. Firing the prime and redundant    actuators on one unit at the  same time, then moving sequentially on to    the next unit, ensures that the actuators will release the solar array    in a deterministic order. Individual panel deployment is controlled by    cam latches. The three panels unfold perpendicular to the spacecraft    x-axis until all three panels are coplanar with panel surfaces,    perpendicular to the x-axis.    Once fully deployed, the wing rotates about two axes, with power going    through Gimbal Actuators at each axis. Rotation required is 180 degrees    for azimuth actuator and 90 degrees for elevation actuator. Position    knowledge is obtained via incremental encoders at actuators and    potentiometers at panel hinges.    Power Sub-system Electronics    ----------------------------    The basic design and operation of the LRO PSE architecture is modeled    after  the Microwave Anisotropy Probe (MAP) and utilizes the solar    arrays to convert sunlight energy into electrical energy while in the    sunlit, or solstice, period. The electrical power is then transferred    to the PSE where it is conditioned and directed to all of the    electrical loads connected to the spacecraft bus.  During the eclipse    seasons, the PSE will also direct a  portion of the sunlight generated    electrical power to the battery for energy storage recharging.  During    the eclipse portion of the orbit the battery will provide all of the    energy to the spacecraft. In order to achieve minimal electrical    losses due to power converters as well as a maintaining a stable    voltage range, the battery will be connected directly to the electrical    bus.    Any excess power from the solar arrays not needed for battery charging    or spacecraft loads will be shunted back to the solar array. The PSE    will perform all of the functions related to power distribution as well    as battery charging and will be designed for single fault tolerance    while still being capable of meeting all mission requirements.    Lithium-Ion Battery    -------------------    A single Lithium Ion battery will provide power to the spacecraft. The    battery energy storage capacity shall be greater than 80 Ampere-hours at    Beginning of Life (BOL) within a 24 - 34V voltage range. The battery    dimensions are (L x H x W) 700mm x 300mm x 180mm (max value). The    battery is attached to the spacecraft structure at 15 locations around    the perimeter of the battery, roughly evenly spaced. Thermal control    is provided at the spacecraft level though a heat pipe assembly.    The battery is constructed from cells arranged in blocks. The basic    block  size is 8s12p (i.e. 12 strings of 8 cells each) and seven blocks    are used, to give a total of 8s84p. Within each block the cells are    connected in  strings with eight cells in each string in series. Four    blocks are arranged to create the top deck, with three blocks in the    lower deck.  Four positive connections for each block in a deck are led    out to a 25-way positive power connector. Four negative connections    for each block and half-string taps in  a deck are led to 35-way    negative/telemetry connectors, In addition, PRT's  and thermistors    installed in the battery are run out to the internal harness    connectors.    Internal 25- and 35 -way harnesses are connected to the power    distribution system in the lower deck. Voltage monitoring lines are    protected with resistors. The positive side of the battery is    connected via two relays arranged in parallel to give redundancy    against open-circuit failure of relays.  The relay system also includes    status-monitoring lines and may be commanded from several sources.    Command lines are isolated from each other with diodes.    The SONY US18650HC Li-Ion Cell is 18 mm diameter by 65 mm high with a    mass of 40.3 grams. The cell is a steel can (anode) containing a roll    of interleaved electrodes soaked in electrolyte, with the necessary    connections attached and cell safety features in the top cap (cathode).    The insulation desks are used to isolate the electrode roll from the    can top with electrical connections provided by tabs and the central    anode pin. The cathode cap is crimped to the can. The cells are each    identical SONY 18650HC from the same manufacturing lot, with a high    level of uniformity in performance characteristics.    Command and Data Handling (C&DH)    --------------------------------    The command and data handling sub-system includes a single board    computer (SBC), a mass storage system, a space wire (SpW) interface    which serves as a high speed interface for the instruments, and a 1553B    Mil-standard low speed bus. Each has some level of heritage with    previous NASA and/or DoD missions.    The functions of the C&DH system include:    1. Hosting the attitude control system (ACS) and flight software (FSW)    2. Command and data handling functions (command acceptance and       distribution, telemetry collection, science data storage) for the       instruments    3. Provide the interface for low rate telemetry, and command and control       of spacecraft subsystems    4. Provide the interface to the spacecraft communications transmitter       and transponder for high speed telemetry (selected instruments and       subsystems  requiring high data transfer rates)    5. Science data formatting    The C&DH subsystem has the following features:    1. Ten Printed Wiring Assemblies (PWAs) in a single housing.    2. The extensive use of SpaceWire European Cooperation for Space       Standardization (ECSS-E-50-12A) and MIL-STD-1553 allows expandability       and scalability.    3. S-Band Communication (SComm)/ Ka-Band Communication (KaComm) cards are       two independent functions, allowing the addition or deletion of either       without redesign. They are referred to as a single assembly,       Communications  (Comm) card, since they share a common backplane       connector.    4. The Single Board Computer (SBC) is a British Aerospace Engineering       (BAE) off the shelf product, with the addition of MIL-STD-1553 Bus       Controller (BC)/Remote Terminal (RT) and 4-port SpaceWire interface.    5. Four Data Storage Boards (DSBs) operate as a mass storage system for       the spacecraft. The DSB boards are designed to interface with the SBC       via a Compact Peripheral Component Interconnect (cPCI) backplane       interface.    6. The Housekeeping / Input Output (HK/IO) card provides an interface to       the LAMP instrument as well as providing synchronization to all       instruments on the spacecraft.    7. The Multi-function Analog Card (MAC) has been custom designed to       manage the wide range of thermal environments in the lunar orbit.    8. The Low Voltage Power Converter (LVPC) provides power to all the       subassemblies except the Primary Ultra Stable Oscillator (USO).    9. The backplane provides interconnectivity to all the cards within the       C&DH enclosure.    10. The primary and redundant USOs are components of the C&DH subsystem,       but are externally mounted separate assemblies.    Guidance, Navigation, and Control (GN&C) Attitude Control System (ACS)    -------------------------------------------    The GN&C Attitude Control System (ACS) controls the pointing of the LRO    spacecraft. Through the use of Star Trackers and Coarse Sun Sensors,    the  ACS will determine where the spacecraft is currently pointing in    fine and coarse accuracies, respectively. A three axis, ring laser    gyro (called an Inertial Measurement Unit) measures the rate at which    the spacecraft is rotating. The use of Reaction Wheels allows the    spacecraft to smoothly point into any desired direction as well as    compensate for disturbance torques (such as High Gain Antenna movement    and Solar Array movement). Eight small thrusters (20 N) are available    to provide steering (attitude control) during large thruster firings    (80 N) for Lunar Orbit Insertion (LOI). Additionally, the small    thrusters are available to use every two weeks to zero out the momentum    buildup in the Reaction Wheels and to perform station-keeping maneuvers    while in the mission orbit.    The Attitude Control System (ACS) will determine spacecraft attitude,    guidance to reach the desired pointing vector, and use actuators to    achieve the desired pointing vector. Additionally, the ACS will    provide pointing support for High Gain Antenna and Solar Array.    The ACS consists of:    1. Sensors       a. Star Trackers (2)       b. Inertial Reference Unit (1)       c. Coarse Sun Sensors (10)    2. Actuators       a. Reaction Wheels (4)       b. 20-Newton ACS thrusters (8)    The Propulsion Deployment Electronics (PDE) component will provide    thruster  control as well as inhibit during launch. GN&C Flight    software (FSW) will  be executed within the C&DH Single Board Computer    (SBC).    Four reaction wheels have the primary purpose of managing nadir pointing    requirement for the spacecraft, and will be used initially as the    primary  means of orienting the spacecraft after separation from the    launch vehicle. The mass of each wheel is 13.2 kg, 18 inches in    diameter, and 6 inches in  height. The fourth reaction wheel allows for    redundancy throughout the mission, such that any one reaction wheel    failure will allow the mission to continue. Momentum dumping is    forecast at approximately two-week intervals throughout the mission,    performed by the attitude control system (small)  thrusters.    The wheels are mounted on the spacecraft structure. Operationally they    will be 'off' at launch and will not be turned 'on' until after launch    vehicle separation. At that time, the wheels will be powered and will    begin to adjust the angular momentum vector that results from the    separation sequence.    Thermal Control System    ----------------------    The Thermal Control system basic design contains four major elements:    the  spacecraft bus (avionics, reaction wheels, and battery), instrument    module, propulsion module, and deployables (high gain antenna and solar    array). The overall orbiter is designed with mainly zenith directed    radiators providing minimal exposure to the lunar environment.    Most avionics are thermally coupled into an embedded constant    conductance heat pipe (CCHP) aluminum honeycomb panel.  Dual bore    header heat pipes couple the isothermal panel to a CCHP radiator that    is separately mounted on the zenith surface of the spacecraft.    For the reaction wheel assemblies, two baseplate heat pipes are    connected to two header pipes, that are then routed to the back of the    Iso-Thermal Panel through a hole in the structural panel.    The Lithium Ion battery is maintained on a separate CCHP network with    two heat pipes mounted on its sidewalls, and an additional pipe coupled    to the sidewalls to provide additional reliability (one fault    tolerance).    A de-coupled instrument optical bench (low thermal distortion) uses low    coefficient of thermal expansion (CTE) M55J material, is fully    blanketed, and is heated with low density heaters to maintain cold    limit temperatures. The CRaTER and LEND instruments are provided with    individual radiators that are thermostatically controlled. Mini-RF    electronics are attached to the avionics header. The Diviner    electronics box is mounted on the iso-thermal panel, while the    remaining instruments are thermally isolated.    Within the propulsion module, most components are attached to a    temperature controlled cylinder that is part of the spacecraft bus    structure, to include the upper propellant tank, pyro valves, and high    and low pressure panels. Thruster valves are coupled to the spacecraft    and tailored to survive soak-back after burn sequences. The lower    propellant tank is thermally coupled to the aft deck for structural    reasons. Remaining lines not on the structural cylinder are    independently heated.    The High Gain Antenna uses multi-layer insulation (MLI), thermal    isolation, and heaters to maintain temperatures within allowable    limits. The gimbal assembly has four heater circuits and two radiators    to serve the actuators and rotary joints. The laser ranging telescope    (mounted on the HGA) has a dedicated heater circuit with MLI thermal    isolation. Optical fiber from the laser ranging telescope has a copper    strap coupled to the high gain antenna structure. The solar array    assembly uses MLI and heater control to maintain actuators and dampers.    Communications System    ---------------------    The communications system consists of S-band forward and return links to    support command, communications, tracking and telemetry, and a Ka-band    return link for telemetry and science data transfer.    S-Band forward link data rate is fixed at 4 kbps. The S-Band return    link data rate is selectable on orbit and varies from 125 bps to 256    kbps. The Ka-band return link is also selectable on orbit and varies    from 25 Mbps to 100 Mbps.    The Ka-band sub-system includes a Ka-Band modulator, a Traveling Wave    Tube Amplifier (TWTA) consisting of a traveling wave tube (TWT), an    Electronic Power Conditioner (EPC), and a High Power Isolator.  The    S-Band subsystem consists of one Spacecraft Tracking and Data Network    (STDN) compatible transponder, an S-band RF Switch, and the RF paths to    and from the TT&C omni-directional antennas and the S-band feed on the    High Gain Antenna (HGA).    The Ka-band uses only the High Gain Antenna. The S-band can utilize    either the Omni-directional or the High Gain Antennae as controlled    through the RF transfer switch.    Diplexers allow the transponder's receiver and transmitter to connect to    a common antenna port that in turn connects to the appropriate RF    network for each antenna system. The diplexers also include band pass    and band reject filters in the transmit channel to suppress any receive    signal component in the output of the transmitting power amplifier.    The traveling wave tube amplifier (TWTA) operates in a bandwidth of 300    MHZ  (+/- 150 MHz) with a voltage standing wave ratio (VWSR) of 2:1.    Output power is a minimum of 40 Watts.  On/off commands are generated    through a 28V signal into the electronic power conditioner (EPC).    Data Storage Boards (DSB)    -------------------------    The four DSB cards are designed as a file system which handles the    storage and retrieval of files.  The DSBs provide 384 Giga bits (Gbits)    at Beginning of Life (BOL) of memory capacity for incoming data files    for a 17.5 hour minimum science data and HK data collection.    The DSB receive dedicated +3.3VDC MEM power for the main power source    and +15 SWITCH power to power the switches that allow the SBC to turn    off individual cards. Power is received via the backplane cPCI    connector.    Data transfer to and from the SBC is via a cPCI interface on the    backplane.    Instrument Module    -----------------    The instrument module includes a graphite composite optical bench with    honeycomb panels, inserts, and spools.  Flexures for mounting the    instruments are titanium. Three instruments are mounted on the optical    bench (LAMP, LOLA, LROC) and the remainder are on the spacecraft bus    (Diviner, Crater, LEND, and the mini-RF antenna). Star Trackers    complete the instrument module.    The Instrument module (IM) is of a 'Wine Box' design, which reduces mass    (fewer fasteners) and increases stiffness with continuous bond lines    (not just connected at fasteners). The entire structure has a low    coefficient of thermal expansion (CTE) made of graphite material    (M55/CyanateEster Q.I. sheets). Insert material is Titanium with BR127    electrical conductive primer.    LRO SCIENCE INSTRUMENTS    ---------------------------    LRO investigations provide the following high-priority measurement sets:    1. Characterization of the deep space radiation environment in lunar       orbit, including neutron albedo (especially at energies in excess of 10       MeV, as well as:       a. Characterization of biological effects caused by exposure to the          lunar orbital radiation environment       b. Characterization of changes in the properties of multifunctional          radiation shielding materials caused by extended exposure to the          lunar orbital environment    2. Geodetic lunar global topography (at landing-site relevant scales)    3. High spatial resolution hydrogen mapping of the Moon's surface    4. Temperature mapping in the Moon's polar shadowed regions    5. Landform-scale imaging of lunar surfaces in permanently shadowed       regions    6. Identification of putative deposits of appreciable near-surface water       ice in the Moon's polar cold traps    7. Assessment of meter and smaller-scale features to facilitate safety       analysis for potential lunar landing sites    8. Characterization of the illumination environment in the Moon's       polar regions at relevant temporal scales (i.e., in terms of hours)    Six instruments were selected for LRO to provide these measurement sets    as well as ancillary datasets relevant to numerous outstanding lunar    science questions. These instruments are:    1. Cosmic Ray Telescope for the Effects of Radiation (CRaTER):    CRaTER will investigate the effect of galactic cosmic rays on tissue-    equivalent plastics as a constraint on models of biological response to    background space radiation. Specific science and measurement objectives    are:    a. Measure and characterize the Linear Energy Transfer (LET) spectra of       galactic and solar cosmic rays (particularly above 10 MeV) in the deep       space radiation environment most critically important to the       engineering and modeling communities to assure safe, long-term human       presence in space.    b. Develop a simple, compact, and comparatively low-cost instrument,       based on previously flown instruments, with a sufficiently large       geometric factor to measure LET spectra and its time variation       globally in the lunar orbit.    c. Investigate the effects of shielding by measuring LET spectra behind       different amounts and types of areal density materials, including       tissue- equivalent plastic.    d. Test models of radiation effects and shielding by verifying/validating       model predictions of LET spectra with LRO measurements, using       high-quality galactic cosmic rays (GCR) and solar energetic protons       (SEP) spectra available contemporaneously with ongoing/planned NASA       (ACE, STEREO, SAMPEX) and other agency spacecraft (NOAA-GOES).    2. Diviner Lunar Radiometer Experiment (DLRE):    DLRE will chart the temperature of the entire lunar surface at    approximately 500 meter horizontal scales to identify cold-traps and    potential ice deposits. Specific science and measurement objectives are:    a. Map Global Day/Night Surface Temperature    b. Characterize Thermal Environments for Habitability    c. Determine Rock Abundances Globally and at Landing Sites    d. Identify Potential Polar Ice Reservoirs    e. Map Variations in Silicate Mineralogy    3. Lyman-Alpha Mapping Project (LAMP):    The Lyman-Alpha Mapping Project (LAMP) will observe virtually the entire    lunar surface in the far ultraviolet. LAMP will search for surface ices    and frosts in the polar regions and provide frost abundance, landform    and surface UV spectral maps of permanently shadowed regions illuminated    only by starlight and interplanetary Lyman alpha. Specific science and    measurement objectives are:    a. Identify and pinpoint surface exposed frost in Permanently       Shadowed Regions (PSRs).    b. Map all permanently shadowed regions with resolutions down to 100m.    c. Demonstrate the feasibility of natural starlight and Lyman-Alpha       sky-glow illumination for future lunar surface mission applications.    d. Assay the lunar atmosphere and its variability.    4. Lunar Exploration Neutron Detector (LEND):    LEND will map the flux of neutrons from the lunar surface to search for    evidence of water ice, and will provide space radiation environment    measurements that may be useful for future human exploration. Specific    science and measurement objectives are:    a. Determine hydrogen content of the subsurface at the polar regions with       spatial resolution of 10km and with sensitivity to concentration       variations of 100 parts per million (ppm) at the poles.    b. Characterization of surface distribution and column density of       possible near-surface water ice deposits in the Moon's polar cold       traps.    c. Global mapping of Lunar neutron emissions at an altitude of 30-50 km       above Moon's surface, with a spatial resolution of 5 km (pixel radius)       at the spectral range of thermal energies up to 15 MeV.    5. Lunar Orbiter Laser Altimeter (LOLA):    LOLA will determine the global topography of the lunar surface at high    resolution, measure landing site slopes, surface roughness, and search    for possible polar surface ice in shadowed regions. Specific science and    measurement objectives are:    a. Global Geodetic Lunar Topography.    b. Characterize Polar Region Illumination.    c. Image Permanently Shadowed Regions.    d. Contribute to the assessment of meter-scale features to facilitate       landing-site selection.    e. Identify surface polar ice, if present.    6. Lunar Reconnaissance Orbiter Camera (LROC):    LROC will acquire targeted narrow angle images of the lunar surface    capable of resolving meter-scale features to support landing site    selection, as well as wide-angle images to characterize polar    illumination conditions and to identify potential resources. Specific    science and measurement objectives are:    a. Landing site identification and certification, with unambiguous       identification of meter-scale hazards.    b. Mapping of permanent shadows and sunlit regions.    c. Meter-scale mapping of polar regions.    d. Repeat observations to enable derivation of meter-scale topography.    e. Global multispectral imaging to map ilmenite and other minerals.    f. Global black and white morphology base map.    g. Characterize regolith properties.    h. Determine recent small impactor rates by re-imaging regions       photographed with the Apollo Panoramic Camera (1-2 meter m/pixel).    Instrument Host Overview - SCN    ================================    The primary function of the SCN is to provide two-way communications    between the Earth and the LRO spacecraft. S-band is used to support    commanding (4 kbps), telemetry, and tracking. Ka-band telemetry (100    Mbps) is used to downlink science instrument data. During operations    after commissioning, the SCN consists of the White Sands 1 (WS1) ground    station, S-Band Sites, and Laser Ranging (LR) site. WS1 transmits s-band    commands to the spacecraft and receives both s-band telemetry and    Ka-band data from the spacecraft. The S-Band Sites consist of 4    Universal Space Network (USN) sites and Deep Space Network (DSN) as    back-up, for transmitting s-band commands to the spacecraft and    receiving s-band telemetry from the spacecraft. The Laser Ranging    facility transmits laser pulses to the spacecraft as part of the    capability of computing the range to the spacecraft.    The WS1 facility includes an 18-meter S/Ka-band antenna at White Sands    Complex (WSC), New Mexico.  The USN facilities include: a USN 13-meter    s-band antenna at South Point, Hawaii; a German Aerospace Center (DLR)    15-meter s-band antenna at Weilheim, Germany; a Swedish Space    Corporation (SSC) 13-meter s-band antenna at Kiruna, Sweden; and, a USN    13-meter s-band antenna at Dongara, Australia.  The DSN facilities    include: a 34-meter s-band antenna at Goldstone, California; a 34-meter    s-band antenna at Madrid, Spain; and, a 34-meter s-band antenna at    Canberra, Australia.    The LR site in Greenbelt, Maryland transmits 532 nm laser pulses to the    LRO spacecraft. The receiver telescope on the spacecraft High Gain    Antenna System (HGAS) provides the Lunar Orbiter Laser Altimeter (LOLA)    instrument with the LR signal via LOLA channel 1. LR range data are sent    from the LOLA instrument to the spacecraft and are then included in LOLA    telemetry sent to the LRO Mission Operations Center (MOC), which in turn    provides these data to the LOLA Science Operations Center (SOC). The SOC    computes the range to the spacecraft.