ROBOTIC LUNAR EXPLORATION PROGRAM CONFIGURATION CHANGE REQUEST

Transcription

1 Page 1 of 2_ ROBOTIC LUNAR EXPLORATION PROGRAM CONFIGURATION CHANGE REQUEST 431-CCR DATE INITIATED: 8/12/2005 CCR REV.: CCR REV DATE: CCB SECRETARY: Deb Yoder PROPRIETARY? Yes No ITAR SENSTIVE? Yes No CCR DUE DATE: 10/3/2005 CCR TITLE (Brief Description): Release the Baseline Version of the Lunar Reconnaissance Orbiter Mission Concept of Operations, ORIGINATOR NAME: Richard Saylor CODE/ORG.: 431/HONEYWELL Richard.Saylor@gsfc.nasa.gov PHONE: SPONSOR NAME: Craig Tooley CODE/ORG.: 431/GSFC ctooley@mscmail.gsfc.nasa.gov PHONE: EXTERNAL ORGANIZATION CCR#: EXTERNAL ORGANIZATION: DOCUMENT (INCLUDE THE DOC. # AND TITLE), CONTRACT, SOFTWARE AFFECTED: LUNAR RECONNAISSANCE ORBITER MISSION CONCEPT OF OPERATIONS EFFECTIVITY: ALL ACS ACS ANALYSIS C&DH COMMUNICATIONS ELECT. HARNESS FLT DYNAMICS GN&C I&T LAUNCH VEHICLE MECHANICAL MECH. ANALYSIS EGSE MGSE SPACECRAFT THERMAL MASS PAYLOAD/INSTR. POWER PROPULSION SOFTWARE FLIGHT OPERATIONS GROUND SYS. OTHER Change Class Class I Class II PROBLEM: Criticality Emergency Urgent Routine COST? NO YES If yes, select one basis for estimate: In-House Actuals ROM Historical Averages Other** ** If Other is chosen for the Basis of Estimate, please explain in Proposed Solution box below: The attached draft version of the LRO Mission Concept of Operations () requires baselining by Level 3A (LRO) CCB. PROPOSED SOLUTION: Release the draft version the LRO Mission Concept of Operations () by the Level 3A (LRO) CCB. Future changes will be initiated by submittal of CCRs. The LRO CMO/Code 431 shall maintain this document. TYPE OF CHANGE: Schedule Interface Document Weight Contract Other

2 Page 2 of 2_ ROBOTIC LUNAR EXPLORATION PROGRAM CONFIGURATION CHANGE REQUEST 431-CCR DATE INITIATED: 8/12/2005 CCR REV.: CCR REV DATE: CCB SECRETARY: Deb Yoder PROPRIETARY? Yes No ITAR SENSTIVE? Yes No CCR DUE DATE: 10/3/2005 BOARD ACTION: APPROVED APPROVED WITH CHANGE DISAPPROVED WITHDRAWN DEFERRED Comments CCB APPROVAL LEVEL REQUIRED [Check appropriate box(es)]: LEVEL 1 NASA HQ Signature: Date: LEVEL 2 RLEP Signature: Date: LEVEL 3A LRO Signature: Date: LEVEL 3B Mission 2 Signature: Date: April 4, 2005, Rev -

3 Effective Date: To be added upon Release Expiration Date: To be added upon Release Lunar Reconnaissance Orbiter Project Mission Concept of Operations August 5, 2005 Goddard Space Flight Center Greenbelt, Maryland National Aeronautics and Space Administration

4 LRO Mission Concept of Operations Draft CM FOREWORD This document is a Lunar Reconnaissance Orbiter (LRO) Project Configuration Management (CM)-controlled document. Changes to this document require prior approval of the applicable Configuration Control Board (CCB) Chairperson or designee. Proposed changes shall be submitted to the LRO CM Office (CMO), along with supportive material justifying the proposed change. Changes to this document will be made by complete revision. Questions or comments concerning this document should be addressed to: LRO Configuration Management Office Mail Stop 431 Goddard Space Flight Center Greenbelt, Maryland 20771

5 Signature Page Prepared by: Richard Saylor Date LRO Mission Operations System Engineer HTSI, Code 431 Reviewed by: William Marinelli Mini-RF Program Manager NAWC/WD Date Martin Houghton Mission Systems Engineer GSFC/NASA, Code 431 Date Approved by: Craig Tooley LRO Project Manager GSFC/NASA, Code 431 Date

6 LUNAR RECONNAISSANCE ORBITER PROJECT DOCUMENT CHANGE RECORD Sheet: 1 of 1 REV LEVEL DESCRIPTION OF CHANGE APPROVED BY DATE APPROVED Rev-

7 List of TBDs/TBRs Item No. Location Summary Ind./Org. Due Date 1 Sect. 2.5 Mini-RF Technology Demonstration Description R. Saylor/ GSFC 11/15/ Table 4-1 Update Launch Mode Configuration R. Saylor/ GSFC 11/15/ Sect Add details of the launch vehicle ascent activities 4 Table 4-3 Update orbiter configuration table for Separation and De-Spin R. Saylor/ GSFC R. Saylor/ GSFC 11/15/ /15/ Sect Add orbiter orientation description to phase R. Saylor/ GSFC 6 Table 4-4 Update orbiter configuration table R. Saylor/ GSFC 7 Sect Add orbiter orientation description to phase R. Saylor/ GSFC 8 Table 4-5 Verify SSR directory structure table R. Saylor/ GSFC 9 Table 4-6 Update orbiter configuration table R. Saylor/ GSFC 10 Sect Add orbiter orientation description to phase R. Saylor/ GSFC 11 Table 4-7 Update orbiter configuration table R. Saylor/ GSFC 12 Sect Add orbiter orientation description to phase R. Saylor/ GSFC 13 Table 5-2 Update orbiter configuration table R. Saylor/ GSFC 14 Sect Add orbiter orientation description to phase R. Saylor/ GSFC 15 Sect Add orbiter orientation description to phase R. Saylor/ GSFC 11/15/ /15/ /15/ /15/ /15/ /15/ /15/ /15/ /15/ /15/ /15/2005

8 Item No. Location Summary Ind./Org. Due Date 16 Sect Add LROC Commissioning Overview R. Saylor/ GSFC 17 Sect Add LOLA Commissioning Overview R. Saylor/ GSFC 18 Sect Add LEND Commissioning Overview R. Saylor/ GSFC 19 Sect Add CRaTER Commissioning Overview R. Saylor/ GSFC 20 Sect Add Diviner Commissioning Overview R. Saylor/ GSFC 21 Table 5-3 Update orbiter configuration table R. Saylor/ GSFC 22 Sect Verify time durations for LROC power cycle R. Saylor/ GSFC 23 Sect Add Mini-RF Tech. Demo ops concept R. Saylor/ GSFC 24 Sect Add orbiter orientation description to phase R. Saylor/ GSFC 25 Sect Update final Doppler tracking accuracy R. Saylor/ GSFC 26 Sect Add details on potential laser ranging network R. Saylor/ GSFC 27 Sect Verify the concept for EPV uplink R. Saylor/ GSFC 28 Sect Add orbiter orientation description to phase R. Saylor/ GSFC 11/15/ /15/ /15/ /15/ /15/ /15/ /15/ /15/ /15/ /15/ /15/ /15/ /15/ Sect. 6.4 Verify SA and HGA position for momentum management R Saylor 11/15/ Sect Add orbiter orientation description to phase R. Saylor/ GSFC 11/15/2005

9 Item No. Location Summary Ind./Org. Due Date 31 Sect Add orbiter orientation description to phase R. Saylor/ GSFC 32 Table 6-10 Update orbiter configuration table R. Saylor/ GSFC 33 Sect Add orbiter orientation description to phase R. Saylor/ GSFC 34 Sect. 6.7 Verify condition to initiate Yaw maneuver R. Saylor/ GSFC 35 Sect Add orbiter orientation description to phase R. Saylor/ GSFC 36 Sect Add orbiter orientation description to phase R. Saylor/ GSFC 37 Sect Add table showing ground network capabilities R. Saylor/ GSFC 38 Sect Add DSN scheduling Concept R. Saylor/ GSFC 39 Table 9-3 Verify or add bandwidth analysis for TBDs R. Saylor/ GSFC 11/15/ /15/ /15/ /15/ /15/ /15/ /15/ /15/ /15/ Figure 2-1 Updated based on latest mechanical configuration 41 Figure 2-2 Updated based on latest mechanical configuration 42 Figure 2-3 Updated based on latest mechanical configuration R. Saylor/ GSFC R. Saylor/ GSFC R. Saylor/ GSFC 11/15/ /15/ /15/ Table 6-9 Update table based on agreed cal/slew activities R. Saylor/ GSFC 11/15/2005

10 TABLE OF CONTENTS Page 1.0 Introduction Purpose And Scope Document Structure Applicable Documents LRO Mission Overview Mission Operations Phases Orbiter System Definitions Spacecraft Overview Instrument Module Propulsion Module Avionics Module Instrument Overview Lunar Reconnaissance Orbiter Camera (LROC) Lunar Orbiter Laser Altimeter (LOLA) Diviner Cosmic Ray Telescope for the Effects of Radiation (CRaTER) Lunar Exploration Neutron Detector (LEND) Lyman-Alpha Mapping Project (LAMP) Technology Demonstration Ground System Overview Pre-Launch Readiness/Operations Phase Space Segment Launch Readiness Ground Segment Readiness Pre-Launch Operations Implementation Procedure and Database Development Flat Sat Simulator Ground Network Site Testing and Verification Ground System Mission Readiness Testing Mission Simulators and Rehearsals Launch and Lunar Transfer Phase Launch and Ascent Sub-Phase Launch Countdown Sequence (L-24 hrs to Launch) Launch and Ascent Separation and De-Spin Orbiter Orientation Description Deployment and Sun Acquisition Sub-Phase Orbiter Orientation Description Lunar Cruise Sub-Phase Lunar Cruise Early Activities (Day 1) ii

11 4.4.2 Lunar Cruise Activities (Day 2) Lunar Cruise Activities (Day 3) Lunar Cruise Activities (Day 4) Orbiter Orientation Description Lunar Orbiter Insertion (LOI) Orbiter Orientation Description Orbiter Commissioning Spacecraft Commissioning Orbiter Orientation Description Integrated Instrument Commissioning Orbiter Orientation Description LROC Commissioning Overview LOLA Commissioning Overview LEND Commissioning Overview LAMP Commissioning Overview CRaTER Commissioning Overview Diviner Commissioning Overview Technology Demostration Commissioning Measurement Operations LRO s Universe at the Moon Routine Measurement Operations LRO Instruments Operation Mini-RF Technology Demo Operations Concept Orbiter Orientation Description (Routine Operations) Off-nadir Measurement Operations Daily Data Volume during Routine Operations Space to ground Link Communications Ground Network Laser Ranging Orbiter Commanding Operations Scenario Station-Keeping Orbiter Orientation Description Momentum Management Orbiter Orientation Description Instrument Calibrations Orbiter Orientation Description Lunar Eclipse Sub-Phase Orbiter Orientation Description Yaw Maneuver Sub-Phase Orbiter Orientation Description Safing / Safe-Mode Sub-Phase Power Induced Load Shedding Safe-Mode iii

12 6.8.3 Orbiter Orientation Description Extended Mission Operations Phase End-of-Mission Disposal Ground System Detailed Description Network, Facilities, and Personnel Security LRO Ground Network White Sands Complex Ground Station S-Band Ground Stations Back-Up/Emergency S-Band Support Communications Networks Data Networks Voice Networks Network Security Flight Dynamics Flight Dynamics Facility Flight Dynamics Analysis Branch (FDAB) Mission Operations Center (MOC) MOC Telemetry & Command (T&C) System Mission Planning System (MPS) Trending System Automated Operations Backup Mission Operations Center Science Operations Centers (SOC) Mission Operations Team Structure and Responsibilities Operations Team during Component Development Phase Operations Team during Orbiter Integration and Test Phase Operations Team during Launch and Commissioning Operations Team during Normal Operations Ground System Testing Ground System Readiness Testing Operations Testing Appendix A. Abbreviations and Acronyms... A-1 iv

13 LIST OF FIGURES Figure Page Figure 2-1. LRO Orbiter Definitions Figure 2-2. Orbiter Deployed Configuration Figure 2-3. Orbiter Deployed Configuration, View # Figure 2-4. Orbiter Stowed Configuration Figure 2-5. LRO System Block Diagram Figure 2-6. LRO Flight Software Modes Figure 2-7. LRO Ground System Overview Diagram Figure 3-1. Pre-Launch Readiness Phase Figure 3-2. Space Segment Readiness Flow Figure 4-1. Launch and Lunar Transfer Phase Sequence Figure 4-2. Typical Launch Day Information Flow Figure 4-3. Deployment and Sun Acquisition Events Figure 4-4. LRO Lunar Trajectory Figure 4-5. LRO Launch and Lunar Cruise Sequence of Events Figure 5-1. Orbiter Commissioning Phase Figure 6-1. LRO s Universe at the Moon Figure 6-2. LRO Integrated Instrument Operations Concept Figure 6-3. Candidate timeline for pulse and burst transmission at both 150-m (mapping) and 10-m (zoom) resolutions Figure 6-4. Candidate timeline for pulse transmission within a burst at both 150-m (mapping) and 10-m (zoom) resolutions Figure 6-5. Orbiter Measurement Operations Pointing Figure 6-6. Orbiter Data Path (Minimum Data Case) Figure 6-7. Orbiter Data Path (Maximum Data Volume Case) Figure 6-8. Ground Network Support Concept Figure 6-9. Ground Station Contact Concept Figure Typical Daily Operations Scenario Figure Weekly Operations Scenario v

14 LIST OF FIGURES (CONTINUED) Figure Page Figure 6-12: Typical Pass Scenario...Error! Bookmark not defined. Figure 9-1. LRO WSC Ground Station Block Diagram Figure 9-2. WSC Ground Antenna System Figure 9-3. CFDP Ka-Band Flow for LRO Figure 9-4. WSC CFDP Processor Control Figure 9-5. LRO MOC Architecture Figure 9-6. CFDP Messages Flow LIST OF TABLES Table Page Table 2-1. LRO Mission Phases Table 2-2. LRO Spacecraft Subsystem Summary Table 2-3. LRO Telemetry and Command Mode Definitions Table 3-1. LRO Activities by Mission Readiness Test Phase Table 4-1. LRO Launch Mode Configuration Table 4-2. LRO Launch Constraints Table 4-3. Orbiter Configuration after Separation and De-spin Table 4-4. Orbiter Configuration after Deployment and Sun Acquisition Table 4-5. LRO Spacecraft Recorder Directory Structure (TBR) Table 4-6. LRO Configuration during Cruise Sub-Phase Table 4-7. LRO Orbiter Configuration for Lunar Orbit Insertion Table 5-1. Spacecraft Commissioning Activities Table 5-2. LRO Orbiter Configuration for Spacecraft Commissioning Table 5-3. LRO Orbiter Configuration after Instruments Commissioning Table 6-1. Lunar Eclipse Predicts for during the LRO Mission Table 6-2. LRO Measurement Interruptions Table 6-3. LROC Flight Rules and Constraints (TBR) vi

15 LIST OF TABLES (CONTINUED) Table Page Table 6-4. LOLA Flight Rules and Constraints Table 6-5. LEND Flight Rules and Constraints (TBD) Table 6-6. LAMP Flight Rules and Constraints Table 6-7. CRaTER Flight Rules and Constraints (TBR) Table 6-8. Diviner Flight Rules and Constraints Table 6-9. LRO Minimum Daily Data Volume (LROC 9 NAC Pairs per Orbit) Table LRO Maximum Daily Data Volume (LROC 16 NAC Pairs per Orbit) Table LRO Space to Ground Telemetry Modes Table LRO CCSDS Virtual Channel Definitions Table Instrument Calibration Activities (TBR) Table Lunar Eclipse Predicts during the LRO Mission Table Orbiter Configuration during Lunar Eclipse Table 9-1. WSC Ka-Band Dump Concept Table 9-2. S-Band Support Sites Table 9-3. (Enter Table Title Here) Table 9-4. Ground System Interface Bandwidths Table 9-5. Ground System Operations Testing vii

16 1.0 INTRODUCTION The Lunar Reconnaissance Orbiter (LRO) is the first robotic mission of the Robotic Lunar Exploration Program (RLEP). The primary objective of the LRO mission is to conduct investigations that support future human exploration of the Moon. The launch readiness date for LRO is October PURPOSE AND SCOPE This document details the operations concept for the LRO mission. Its purpose is to describe how operations will be accomplished through all mission phases from pre-launch to end-ofmission disposal. This document is not a requirements document and is intended only to provide the concepts for how mission operations will be conducted. Requirements are defined in documents such as the Lunar Reconnaissance Orbiter Mission Requirements Document (431-RQMT ), the Detailed Mission Requirements for the Lunar Reconnaissance Orbiter Ground System (431- RQMT ) and in various interface control documents (ICDs). This document will serve as the basis from which more specific documents will be formed, such as the Lunar Reconnaissance Orbiter Flight Operations Plan (431-PLAN ) and the Lunar Reconnaissance Orbiter Launch and Commissioning Handbook (431-HDBK ). 1.2 DOCUMENT STRUCTURE The following describes the document breakdown structure: Section 1: Introduction This section describes the purpose and scope of the operations concept document. Section 2: Mission Overview This section provides a high-level overview of the LRO mission. It also provides a brief description of the spacecraft (SC) and instruments. Section 3: Pre-Launch Operations The pre-launch phase deals with integration and test (I&T) support as well as flight operations development and preparation. Section 4: Launch and Lunar Transfer The launch and lunar transfer phase starts with the launch vehicle (LV) countdown through Lunar Orbit Insertion (LOI). Section 5: Orbiter Commissioning The commissioning phase covers activation and calibration activities for both the SC and instruments. Section 6: Routine Operations This section describes the nominal operational activities. Section 7: Extended Mission Operations This section describes operations beyond the primary mission lifetime. 1-1

17 Section 8: End-of-Mission Disposal This section describes the activities associated with endof-mission. Section 9: Ground System (GS) Description This section describes in detail the various components of the GS as they are envisioned in the requirement definition and preliminary design phases of the mission development. 1.3 DOCUMENTS Applicable Documents 431-HDBK HDBK HDBK ICD ICD ICD ICD ICD ICD PLAN PLAN PLAN PLAN PLAN PLAN RQMT RQMT RQMT Lunar Reconnaissance Orbiter Launch and Commissioning Handbook Lunar Reconnaissance Orbiter Mission Command and Telemetry Database Style Guide Lunar Reconnaissance Orbiter Integrated Test and Operations System Procedure Style Guide Lunar Reconnaissance Orbiter Ground System Interface Control Document Cosmic Ray Telescope for Effects of Radiation Data Interface Control Document Lyman-Alpha Mapping Project Data Interface Control Document Lunar Exploration Neutron Detector Data Interface Control Document Lunar Orbiter Laser Altimeter Data Interface Control Document Lunar Reconnaissance Orbiter Camera Data Interface Control Document Lunar Reconnaissance Orbiter Ground System Product Development Plan Lunar Reconnaissance Orbiter Project Database Management Plan Lunar Reconnaissance Orbiter Flight Operations Plan Lunar Reconnaissance Orbiter Mission Readiness Testing Test Plan Lunar Reconnaissance Orbiter Integration and Test Plan Lunar Reconnaissance Orbiter Information Technology Security Plan Lunar Reconnaissance Orbiter Mission Requirements Document Detailed Mission Requirements for the Lunar Reconnaissance Orbiter Ground System Lunar Reconnaissance Orbiter Pointing and Alignment Requirements 1-2

18 431-SPEC SPEC Lunar Reconnaissance Orbiter Electrical System Specification Lunar Reconnaissance Orbiter CCSDS File Delivery Protocol Implementation Specification Reference Documents NASA FAR NPR NPR Security Requirements for Unclassified Automated Information Resources Security Information Technology NASA Procedural Requirements Planetary Protection Provisions for Robotic Extraterrestrial Missions NASA Procedural Requirements 1-3

19 2.0 LRO MISSION OVERVIEW LRO is the first mission of the RLEP. The goal for the RLEP is to prepare for future human exploration of the Moon. LRO specific objectives are: Characterize the lunar radiation environment, biological impacts, and potential mitigation Determine a high resolution global, geodetic grid of the Moon in 3 dimensions Assess in detail the resources and environments of the Moon s polar cap regions Perform high spatial resolution of the Moon s surface The LRO instrument complement includes six instruments. Together, all six instruments allow LRO to meet the mission objectives. Description of the six instruments is: Lunar Orbiter Laser Altimeter (LOLA): LOLA will determine the global topography of the lunar surface at high resolution, measuring landing site slopes and search for polar ice in shadow regions. Lunar Reconnaissance Orbiter Camera (LROC): LROC will acquire targeted images of the lunar surface capable of resolving small-scale features that could be landing site hazards. LROC will also produce wide-angle images at multiple wavelengths of the lunar poles to document the changing illumination conditions and potential resources. Lunar Exploration Neutron Detector (LEND): LEND will map the flux of neutrons from the lunar surface to search for evidence of water ice and provide measurements of space radiation environment which can be useful for future human exploration. Diviner Lunar Radiometer Experiment (DLRE): Diviner will map the temperature of the entire lunar surface at 300-meter horizontal scales to identify cold-traps and potential ice deposits. Lyman-Alpha Mapping Project (LAMP): LAMP will observe the entire lunar surface in the far ultraviolet (UV). LAMP will search for surface ice and frost in the Polar Regions and provide images of permanently shadowed regions illuminated only by starlight. Cosmic Ray Telescope for 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. 2-1

20 LRO will also fly a technology demonstration instrument called the Mini-Radio Frequency (RF). The purpose of the Mini-RF is to demonstrate new radar technology for future use in planetary resource mapping. The development of the LRO SC bus occurs at the Goddard Space Flight Center (GSFC). The orbiter team integrates the measurement instruments and performs system level testing at GSFC. LRO is scheduled to launch in October The orbiter is will be launched aboard an expendable launch vehicle (ELV) from the Eastern Range at the Kennedy Space Center (KSC). The LVwill inject LRO into a cis-lunar transfer orbit. After the lunar cruise, LRO will be required to perform a series of LOI maneuvers to capture into the commissioning orbit. After orbiter commissioning is complete, the final orbit maneuvers insert the orbiter into a nominal mission orbit of 50 kilometers (km). Once LRO is in the final mission orbit, the six instruments will start to collect measurement data for the mission. The SC collects and stores the data on the recorder. The SC dumps the data stored on the recorder using the Ka-Band system, and the dedicated dual feed antenna located at White Sands receives the data. Once the data is on the ground, the GS is responsible for distribution of the data for processing. The six instrument science operations centers (SOCs) will receive the data for higher-level measurement data processing. Once the data processing is complete, measurement data products are stored at the Planetary Data System (PDS) for longterm storage and use. The six instruments operate majority of the mission in the same operating mode. Periodic interruptions are expected to the nominal measurement operations and are identified later in this document. These would include orbiter station-keeping (SK) and momentum maneuvers, instrument calibration maneuvers, Lunar eclipses, and periodic yaw maneuvers. Along with the six instruments, the orbiter will fly a technology demonstration called the mini- RF payload. The mini-rf payload will operate on a non-interference basis throughout the mission. 2.1 MISSION OPERATIONS PHASES The LRO operations lifecycle spans the early planning stages of the mission through the 1-year routine measurement operations. An additional four years of extended mission operations are possible in a low fuel maintenance orbit. Table 2-1 summarizes the main LRO Mission Phases. 2-2

21 Table 2-1. LRO Mission Phases No Phase Sub-Phases Description 1 Pre-Launch/ Launch Readiness 2 Launch and Lunar Transfer 3 Orbiter Commissioning 4 Routine Operations 5 Extended Mission Operations 6 End-of-Mission Disposal Space Segment Readiness Ground Segment Readiness Launch and Ascent Separation and Despin Deployment and Sun Acq. Lunar Cruise Lunar Orbit Insertion Spacecraft Commissioning Integrated Instrument Commissioning Measurements (Routine Ops) Station-keeping Momentum Management Instrument Calibrations Lunar Eclipse Yaw Maneuver Safe Mode Includes instrument I&T, SC/orbiter I&T, space/ground segment testing as well as operations preparation and ground readiness testing leading up to launch. Includes all activities and operations from launch countdown sequence to LOI. LOI includes all maneuvers necessary to obtain the temporary parking orbit for Orbiter activation and commissioning. During the cruise phase, initial SC checkout will be performed to support activities for midcourse correction (MCC) and LOI. Configure and checkout the SC subsystems and ground systems prior to instrument turn-on. Instrument integrated activation will be developed to complete instruments turn-on and commissioning. Instrument commissioning includes any calibration activities needed in the temporary orbit. One year of nominal science collection in the 50 (+/- 15) km orbit. After 1-year of science observations, orbiter may be boosted into a higher orbit to reduce maintenance requirements. Main purpose for extended mission operations is to perform relay comm. operations. Additional measurement operations may be performed. Includes planning and execution of end-of-life (EOL) operations. Could include controlled/uncontrolled impact on the Lunar surface. 2-3

22 2.2 ORBITER SYSTEM DEFINITIONS The orbiter consists of several systems along with required ground support equipment (GSE) to perform testing. Figure 2-1 shows what components and systems make up the orbiter. The definitions defined below are used throughout this document. 2.3 SPACECRAFT OVERVIEW Figure 2-1. LRO Orbiter Definitions During the routine measurement data collection, the LRO SC has three primary responsibilities. The first is to keep the LRO instrument suite pointed nadir around the lunar orbit. The second is to collect the different measurement data and store it on the recorder for downlink. The third responsibility is to monitor the orbiter health and status. 2-4

23 The SC and subsystem description in this section provides background information for the operation concepts presented in this document. They reflect the SC and subsystem designs at the time this document was created. Table 2-2 summarizes the major SC subsystems functions. Table 2-2. LRO Spacecraft Subsystem Summary Command and Data Handling (C&DH) Provides SC processor for attitude control algorithms, command/telemetry processing. Communication cards provide the interfaces to the S-band/Ka-Band RF systems. Hardware command decoding for computer-free recovery Provides high speed and low speed data bus to the instruments and SC components Provides large volume recorder for measurement data and orbiter housekeeping (HK) Communication Ka-band transmitter for high rate measurement downlink using the HGA S-Band transponders connected to the omni antennae and HGA for receipt of ground commands and telemetry downlink Supports orbit determination via turnaround ranging Mechanical & Mechanism Design for ELV Modular design for flexibility at I&T Deployable SAs and HGA Guidance Navigation and Control (GN&C) Power Three axis control with reaction wheels Star Trackers (STs), Inertial Reference Unit (IRU), Coarse Sun Sensors (CSS) used for attitude control Momentum management is performed periodically with thrusters Controls pointing of the solar array (SA) and High Gain Antenna (HGA) gimbals SA located on gimbals for power generation One Lithium Ion battery for launch and 48 minute lunar occultations Power switching and distribution. Battery charging control Provide battery power for lunar eclipse Flight Software (FSW) Complex algorithms computed on central processor including Attitude Control System (ACS), stored commanding, telemetry and measurement data processing, and fault detection and correction. SC time distribution/maintenance 2-5

24 Thermal Combination of passive/active design Operational and survival heaters Thermostatic control of survival heaters Propulsion Monopropellant design (Hydrazine) for MCC, LOI, SK and momentum management. Two 20 pound force (lbf) axial thrusters for MCC and LOI Four dual thruster modules for SK and momentum management The orbiter system consists of two main modules that are the SC bus and instruments. The SC bus consists of the following main elements: SA System, HGA System, propulsion module, and avionics module. Figures 2-2 and 2-4 show the orbiter in the deployed and stowed configuration. CRaTER INSTRUMENT MODULE LROC LOLA AVIONICS MODULE LAMP Mini-RF PROPULSION MODULE SOLAR ARRAY Figure 2-2. Orbiter Deployed Configuration 2-6

25 INSTRUMENT MODULE AVIONICS MODULE LAMP HGA PROPULSION MODULE LEND Diviner SOLAR ARRAY Figure 2-3. Orbiter Deployed Configuration, View #2 2-7

26 2.3.1 Instrument Module Figure 2-4. Orbiter Stowed Configuration The instrument module contains the surface that the instruments are mounted. Along with the instruments, the SC gyro and ST units mount on the instrument bench to reduce distortion errors with the instruments. The instrument module contains the necessary thermal heaters and radiators to maintain the components within operating temperature range Propulsion Module The propulsion module contains all the components and plumbing required for the thrusters and contains the LV interface structure. The propulsion module includes one Hydrazine propellant tank, Helium pressurant tank, 4 sets of dual 5 lbf thrusters, two 20 lbf axial thrusters, valves, regulators and other plumbing fixtures. The axial thrusters will be used primarily for insertion burns to capture into lunar orbit. The dual thruster modules are used to perform routine SK and momentum management maneuvers. 2-8

27 2.3.3 Avionics Module The avionics module contains C&DH, GN&C, Power System Electronics (PSE), and Thermal Control Systems (TCS). Figure 2-5 shows the Avionics System Block Diagram and the primary interfaces between the components. New Instrument SA Gimbals LROC LAMP LOLA LEND Diviner CRaTER Battery Solar Array Sw. and Unsw. +28V Pwr Services Mini-RF SpaceWire Network LAMP Sci. & HK Unsw. + 28V H/W Decoded Command Discretes Discretes Thermistors Closed Loop Htrs Heat Pump Loop C&DH LVPC HK Unsw. + 28V SSR Sw. + 28V PMC SAM OM OM OM PSE Comm SSR SBC HK / IO Thermal LVPC C&DH MIL-STD-1553 Network Power Bus Low-Rate Cmds & Tlm Hi-Rate Tlm Gimbal Ctl Prop/Dep-A Prop/Dep-B Prop/Dep-C Prop/Dep-D Gimbal Ctl PDE * S-Xpndr Ka-Xmtr IMU * ST(2) * IRW(4) * SA & HG Deploy Actuation HGA Gimbals CSS(8) * Pressurant Tank Propulsion * NC P P P R R NC Propellant Tank Omnis HGA C&DH Comm CRaTER CSS EVD HGA H/W HK IO IMU IRW LAMP LEND LOLA LROC LVPC OM PDE PMC PSE SA SAM ST SBC SSR Xmtr Xpndr Mini-RF SpaceWire MIL-STD-1553 Command & Data Handling Communications Cosmic Ray Telescope for the Effects of Radiation Coarse Sun Sensor Engine Valve Driver High-Gain Antenna Hardware Housekeeping Input/Output Integrated Momentum Unit Integrated Reaction Wheel Lyman-Alpha Mapping Project Lunar Exploration Neutron Detector Lunar Orbiter Laser Altimeter Lunar Reconnaissance Orbiter Camera Low Voltage Power Converter Output Module Propulsion Deployable Electronics Power Management Controller Power System Electronics Solar Array Solar Array Module Star Tracker Single Board Computer Solid State Recorder Transmitter Transponder Figure 2-5. LRO System Block Diagram The C&DH system will collect and store the data on the SC recorder. Since the SC recorder is switchable, the C&DH system will also contain a small memory buffer located on the processor card that will stored limited amount of HK data. The C&DH system will also distribute commands to the other orbiter components. The SC will utilize the Consultative Committee for Space Data Systems (CCSDS) File Delivery Protocol (CFDP) to perform reliable transfers to the ground. 2-9

28 The single board computer (SBC) will execute the FSW that includes both C&DH and GN&C software (SW). The GN&C SW will control the SA and HGA gimbals. Figure 2-6 shows the current FSW modes. The modes in the figure are the high-level modes; each mode may have additional sub-modes. The key point regarding FSW transitions is that it requires a ground command to transition to higher-level modes while transitions to lower modes can occur through ground command or fault detection and correction (FDC) running onboard the SC. The SC will contain two banks of memory that contain the identical copy of the entire FSW. Startup Programmable Read-Only Memory (PROM) contains the initial boot code that tells the processor to load SW from one of the two memory banks. While the startup PROM cannot be updated, each FSW bank is programmable in flight. When the processor boots, the startup code looks for the selected bank and loads the FSW. Power On Sun Acquisition Sun Acquisition Mode ( Sun Pointing ( Null Body Rates ( IMU, Wheels, CSSs ( SA in Predefined position ( HGA in predefined position Maneuver Comp./ Gnd Cmd/ FDC Gnd Cmd Maneuver Comp./ Gnd Cmd/ FDC Delta-H Gnd Cmd Gnd Cmd/ FDC Delta-V Delta-H Mode ( 3-Axis Control ( Thrusters, IMU, Wheels ( SA in fixed position ( HGA in fixed position Gnd Cmd Maneuver Comp./ Gnd Cmd/ FDC Maneuver Comp./ Gnd Cmd/ FDC Gnd Cmd Delta-V Mode ( 3-Axis Control ( Thrusters, IMU, Wheels ( SA in fixed position ( HGA in fixed position Observing Observing Mode ( Inertial/Nadir pointing ( STs, IMU, Wheels ( SA Tracking Sun ( HGA Tracking Earth Figure 2-6. LRO Flight Software Modes The S-Band communication system is used primarily for real-time (R-T) HK and commands between the ground network and orbiter. The Ka-band system is used for dumping the measurement data from the recorder. Table 2-3 shows the current telemetry modes that are defined for LRO. 2-10

29 Table 2-3. LRO Telemetry and Command Mode Definitions Mode S-Band Command Mode 1 S-Band Command Mode 2 Data Rate (bps) (1) Coding Modulation Ranging Purpose 4x10 3 Uncoded PCM/PSK/ON Yes Standard Command Mode 4x10 3 Uncoded PCM/PSK/ON No Contingency Command Mode S-Band Telemetry Mode 1 S-Band Telemetry Mode 2 S-Band Telemetry Mode 3 S-Band Telemetry Mode 4 S-Band Telemetry Mode 5 S-Band Telemetry Mode 6 S-Band Telemetry Mode 7 S-Band Telemetry Mode 8 S-Band Telemetry Mode 9 S-Band Telemetry Mode 10 S-Band Telemetry Mode 11 S-Band Telemetry Mode x10 3 2x x x x x x10 3 Rate ½ and Reed- Solomon (RS) Rate ½ and RS Rate ½ and RS Rate ½ and RS Rate ½ and RS Rate ½ and RS Rate ½ and RS Rate ½ and RS 273.3x10 3 Rate ½ and RS 273.3x10 3 Rate ½ and RS 2.186x10 6 Rate ½ and RS PCM/PSK/PM No Contingency Telemetry Mode PCM/PSK/PM PCM/PSK/PM No Yes Contingency/Low Rate Telemetry Mode Contingency/Low Rate Telemetry Mode PCM/PSK/PM No Contingency Telemetry Mode PCM/PSK/PM PCM/PSK/PM PCM/PSK/PM PCM/PSK/PM PCM/PSK/PM BPSK OQPSK Yes Yes Yes Yes Yes No No 5x10 6 Uncoded OQPSK No Nominal R-T Telemetry Mode Nominal R-T Telemetry Mode High Rate R-T Telemetry Mode High Rate S-Band with S- Band Data Dump High Rate S-Band with S- Band Data Dump High Rate S-Band with S- Band Data Dump High Rate S-Band with S- Band Data Dump High Rate S-Band with S- Band Data Dump Ka-Band Telemetry Mode 1 Ka-Band Telemetry Mode 2 25x x10 6 Rate ½ and RS Rate ½ and RS OQPSK OQPSK No No Low Rate Ka-Band contingency rate Medium Rate Ka-Band contingency rate 2-11

30 Mode Ka-Band Telemetry Mode 3 Data Rate (bps) (1) Coding Modulation Ranging Purpose 100x10 6 Rate ½ and RS OQPSK No Nominal Ka-Band downlink rate Notes: (1) LRO Data Rate prior to any coding 2.4 INSTRUMENT OVERVIEW The LRO instrument complement is comprised of six instruments and one technology demo instrument. There will be a SOC for each instrument supporting the interface between the instrument Principal Investigator (PI) teams and the LRO GS Lunar Reconnaissance Orbiter Camera Mark Robinson leads LROC measurement investigation from Northwestern University. LROC comprises of two narrow angle cameras (NAC), wide-angle camera (WAC) and a Sequence and Compression System (SCS). LROC will utilize the high-speed data bus interface to transfer measurement data and HK to the SC recorder. LROC uses the high-speed data bus to retrieve control commands from a sequence file stored on the SC recorder. LROC measurement objectives include: Landing site identification and certification Unambiguous mapping of permanent shadows and sunlit regions Meter-scale mapping of polar regions with continuous illumination Overlapping observation to enable derivation of meter-scale topography Global multispectral imaging to map ilmenite and other minerals Global morphology base map Characterize regolith properties Determine current impact hazards by re-imaging 1-2m/pixel Apollo images The LROC SOC will be located at Northwestern University. It will interface to the Mission Operations Center (MOC) for mission products, measurement data files, and mission coordination. 2-12

31 2.4.2 Lunar Orbiter Laser Altimeter David Smith leads LOLA measurement investigation from GSFC. LOLA uses a 1064 nanometer (nm) LASER that expands to provide a five spot pattern on the Moon s surface. A telescope receives the reflected light where the electronics processes the return. LOLA will use the low speed 1553 data bus for is data and command interface to the SC. LOLA has two prime objectives: Produce a high-resolution global topographic model and global geodetic framework that enables precise targeting, safe landing, and safe mobility on the Moon s surface. Characterize the polar illumination environment, and image permanently shadowed regions of the Moon to identify possible locations of surface ice crystals in shadowed polar craters. The LOLA SOC will be located at GSFC and will interface to the MOC for mission products, measurement data files, housekeeping data, and mission coordination Diviner David Paige leads diviner measurement investigation from University of California Los Angeles (UCLA). Diviner includes a 9-channel radiometer with a wavelength range from 0.3 to 200 microns. Diviner will make precise radiometric temperature measurements of the lunar surface with the following measurement goals: Map global day/night surface temperatures Characterize thermal environments for habitability Determine rock abundances at landing sites Identify potential polar ice reservoirs Search for near-surface and exposed ice deposits The Diviner SOC will be located at the Jet Propulsion Laboratory (JPL) and will interface to the MOC for mission products, measurement data files, R-T HK data, stored HK files, and mission coordination Cosmic Ray Telescope for the Effects of Radiation Harlan Spence leads CRaTER measurement investigation from Boston University (BU). The CRaTER instrument will measure cosmic ray sources from two different directions (looking nadir and zenith). The instrument telescope contains a series of five detectors spaced apart that will measure the different cosmic rays. CRaTER measurement goals are to: Measure and characterize the aspect of the deep space radiation environment and spectra of galactic and solar cosmic rays. 2-13

32 Characterize the biological impacts from the radiation environment The CRaTER SOC will be located at BU. The SOC will interface to the MOC for mission products, measurement data files, R-T HK data, stored HK files, and mission coordination Lunar Exploration Neutron Detector LEND measurement investigation is led by Igor Mitrofanov from the Russian Institute for Space Research. The LEND instrument includes collimated sensor and sensors to detect thermal, epithermal, and high-energy neutrons. LEND measurements include: Measurement of thermal neutrons with flux variation >1% and altitude-dependent spatial resolution about 50km. Measurement of epithermal neutrons >0.4 electron Volts (ev) with flux variation about 2% (pole) and 10% (equator). Measurement of high-energy neutrons Mega-electron Volts (MeV) with flux variations 4% (pole) and 10% (equator) The LEND SOC is located at the Russian Institute for Space Research in Moscow, Russia. The LEND SOC will interface to the MOC for mission products, measurement data files, stored HK files, and mission coordination Lyman-Alpha Mapping Project Alan Stern leads LAMP measurement investigation from Southwest Research Institute (SwRI). The LAMP instrument includes high and low power supplies and a double delay line detector. The LAMP instrument measurement goals are to: Provide landform mapping from (from Lyα albedos) at sub-km resolution in and around the permanently shadowed regions of the lunar surface. Identify and localize exposed water frost. Demonstrate the feasibility of using starlight and sky-glow for future surface mission applications. Detect the abundances of several atmospheric species. The LAMP SOC will be located at SwRI and interfaces with the MOC for mission products, measurement data files, R-T HK data, stored HK data files, and mission coordination. 2.5 TECHNOLOGY DEMONSTRATION The LRO carries a technology demonstration payload. The Mini-RF will take measurements to demonstrate the following technology: 2-14

33 Imaging from 50 km altitude surface areas that were imaged by Forerunner with same dual polarization, resolution, and S-Band frequency as was used by Forerunner. Imaging with both S- and X-Band surface areas that were imaged by Forerunner Imaging with both dual polarization and full, 4 channel polarization surface areas that were imaged by Forerunner. The Forerunner instrument is a current payload on the Chandrayaan-1 mission to the Moon. The LRO Mini-RF tech demo is synthetic aperture radar that consists of an electronics box, a fixed planar antenna mounted on an external SC surface, a cable harness between the electronics box and the antenna, and electrical interfaces to the SC, as described below. Two accommodations of the electronics box by the SC are considered at this time. The electronics box may be directly mounted to an external SC surface or it may be included in the SC avionics space. 2.6 GROUND SYSTEM OVERVIEW The LRO GS is comprised of five main elements as shown in Figure 2-7: The LRO Ground Network that consists of a dedicated S/Ka Band ground station at White Sands and various S-Band only ground stations located throughout the world. Mission Operations Center (MOC) Flight Dynamics (FD) Science Operations Center (SOC) for each instrument Communications network which provides voice and data connectivity between each of these elements Because of the high data volume that LRO will produce and the use of Ka-Band frequency, LRO elected to use a dedicated S/Ka ground station for primary measurement data downlink. The measurement data is received at the ground station and rate buffered to the MOC post-pass for data processing/accountability. The MOC at GSFC will distribute the measurement files to each of the instrument SOCs along with other mission products required for processing. The MOC is the focal point for all orbiter operations including health and safety monitoring. The MOC will generate all commands to the orbiter. The SOCs support instrument operations including instrument command sequence inputs, measurement data processing, transferring measurement products to the PDS, and instrument HK and performance trending. 2-15

34 Figure 2-7. LRO Ground System Overview Diagram The LRO GS elements are briefly described below: 1. The LRO ground network consists of a dedicated ground station at White Sands Complex (WSC) and S-Band ground stations. The WSC ground station is capable of receiving 100 megabits per second (Mbps) downlink of measurement data files produced by the instruments on Ka-Band frequency and supporting R-T commands and telemetry on S-Band frequency. Due to susceptibility to Ka from weather, White Sands provides the optimal location due to its minimal precipitation levels. Because LRO requires near continuous tracking data for orbit determination, additional S- Band sites are required. The S-Band only sites will provide R-T telemetry and commands (T&C) capabilities along with tracking data. The S-Band stations could be used to dump low rate measurement files in a contingency mode. LRO plans to use the Deep Space Network (DSN) for emergency/backup support. The emergency/backup support will utilize only the S-Band frequency. LRO has also accepted the requirement to provide high accuracy tracking data for measurement 2-16

35 data processing. At this time, it is not clear whether the high accuracy tracking data can provided using the S-Band system. Alternative concepts that LRO is looking at is using Laser ranging. If used, the ground based Laser tracking stations will be part of the ground network. 2. The MOC will be located at GSFC. It is the main T&C interface to the orbiter. The MOC will process HK data to monitor health and safety of the orbiter. The MOC will also distribute measurement data to the individual SOCs along with other required mission products. The MOC provides temporary measurement data storage until verification is received from the SOC on file delivery. The MOC will receive any required instrument command sequences from the SOCs and process them before uplink. The MOC will also distribute R-T telemetry to the SOCs. 3. The six SOCS provide the hardware and SW to support the following functions: Instrument health and safety monitoring Instrument command sequence generation/request Support orbiter calibration planning/coordination Measurement data processing (level 0 and higher) Measurement data product archiving and transfer to the PDS Maintain instrument FSW/tables 4. FD performs orbit determination, attitude processing, and maneuver/trajectory planning. For orbit determination, FD will receive the tracking data from the ground network and generate mission products. FD will also provide attitude verification and planning support for slews. Besides pre-mission trajectory and orbit planning, FD will also monitor and plan for trajectory maneuvers during the cruise, LOI burns, and SK maneuvers. 5. The GS communication network provides voice and data connectivity between each of the GS elements. It will provide the necessary data lines between the ground network, MOC, and SOCs. 2-17

36 3.0 PRE-LAUNCH READINESS/OPERATIONS PHASE This section and the following sections will provide details associated with each of the mission phases. The section will cover activities associated with preparing the SC, instruments, GS, and operations for launch and for mission operations. This phase starts with instrument, SC and ground I&T and ends at the start of the LV countdown sequence, refer to Figure 3-1. This phase last approximately 2.5 to 3 years prior to launch and is includes two sub-phases: Space Segment Readiness and Ground Segment Readiness. Figure 3-1. Pre-Launch Readiness Phase 3.1 SPACE SEGMENT LAUNCH READINESS Space Segment readiness includes the following activities: Instrument I&T SC subsystem assembly and test SC bus I&T Orbiter level I&T Launch site testing and processing Part of the space segment includes the necessary GSE that required for the various tests. GSE includes both mechanical, electrical, and simulators. Figure 3-2 shows a typical high-level flow for space segment readiness. As the I&T test plans develop, the exact sequence may change. The purpose of the figure is to provide the reader the basic concept for orbiter I&T activities. 3-1

37 Inst. Interface Check Out W/ Flat Sat IM Delivery Instrument Integration & Alignment Avionics Module integration Propulsion Module Integration Mate PM, AM IM, HGADAS, SADAS SC bus testing Alignment Baseline Bus Special Tests Instrument Special Tests CPT Sine Vibe W/ Aliveness Acoustics W/ Aliveness Separation/ Shock Test Alignment Check & Functional Test EMC Thermal Vac & Balance Alignment Check Post Environmental CPT Bus & Instrument Special Tests Pre-Ship Activities Ship To KSC Post-Ship Activities Post-Ship Functional Spin Balance Pad Readiness Activities Pad Installation Pad Aliveness Test Final Close Outs Fairing Installation LAUNCH Figure 3-2. Space Segment Readiness Flow SC simulators will be developed and provided to the instrument teams for early checkout of the interfaces. Instrument teams will also provide the SC team instrument simulators that will allow checkout of the data and electrical interfaces. Once the orbiter is integrated, an orbiter level environmental test program is performed prior to final checkout and ship to the launch site. At the launch site, final activities include: Post ship functional checkout Propulsion system fueling and spin balance test Pad installation activities Pad Aliveness test Fairing closeouts 3-2

38 Besides the instrument and SC hardware, a flat sat will be developed that includes SC subsystem engineering test units (ETUs) and instrument simulators. The flat sat will be used test interfaces, FSW, databases, and test procedures. The GS team will also use the flat sat for operations development and test. 3.2 GROUND SEGMENT READINESS The ground segment readiness includes the following activities: GS development, includes both hardware and SW I&T GS verification testing Mission operations development Operations testing The GS is broken into five basic elements as described in Section 2. The GS development schedule includes three releases. The first release will allow the ground team and operations team to support SC and orbiter level I&T. The second release will be a launch ready release of the GS. A third release is planned for final cleanup to close out any remaining issues with the ground system Pre-Launch Operations Implementation The LRO Mission Operations Team (MOT) performs preparation for flight operations. The team will be assembled early in the mission and assumes the dual role of both the I&T test engineering and on-orbit Flight Operations Team (FOT). Test engineers will support the subsystem component and instrument development. Test engineers are assigned to dedicated instruments to assist the instruments integration at GSFC. The test engineers will ensure that test procedures, databases, GSE interfaces, and documentation conform to the I&T standards. The test engineers will continue to follow I&T to the launch site processing activities. By utilizing test engineers from the operations team, it provides benefit to project in reducing effort to develop operations products since I&T products can be converted easily to flight procedures. The operations team benefit with increased knowledge into the instrument and SC systems. Some general guidelines for dovetailing orbiter I&T with flight operations are: Use the same ground system for flight operations, I&T and in the component development labs. Use the same personnel for both pre-launch I&T activities and on-orbit flight operations. Create operations oriented databases, procedures and other GS products right from the beginning. 3-3

39 Tests and test procedures should be developed based on real operational scenarios (test it like you fly it) Procedure and Database Development Several different T&C systems will be deployed on LRO. The SC and subsystems along with operations will be using the Integrated Test and Operations System (ITOS). Each instrument may be using its own T&C system, LOLA will utilize ITOS for its instrument I&T. If instrument development teams elect to use their own T&C system, it will be used during the instrument development phase. Instrument systems can be brought to orbiter level I&T but ITOS will be the primary T&C system used during all orbiter level testing. The primary ITOS can distribute data to each of the instruments GSE. Early in the development phase, databases are developed for components and instruments utilizing their T&C system. For SC components, ITOS is this system. Operationally oriented System Test and Operations Language (STOL) test procedures will begin to be authored in the component development labs in support of the test and integration of the component. At all support sites, the MOTs will provide expertise in the use of the ITOS T&C and transfer this knowledge to the component engineering staff. The LRO GS will develop several documents that provide guidelines regarding STOL and database development. The documents include: Lunar Reconnaissance Orbiter Integrated Test and Operations System Procedure Style Guide (431-HDBK ) Describe common STOL format, specify standard header information, and common STOL writing style. Lunar Reconnaissance Orbiter Telemetry and Command Database Style Guide (431- HDBK-431-HDBK ) Describes the nomenclature rules for specifying command and telemetry mnemonics. It will also provide details regarding limitations such as maximum length for each mnemonic type. Lunar Reconnaissance Orbiter (LRO) Mission Database Management Plan (431-PLAN ) Provides details how the official database will be developed and controlled throughout pre-mission and on-orbit. It will identify the various inputs and the process for updates. Once an instrument or component is delivered to orbiter I&T, all procedures and databases are controlled at the master copies located at GSFC. Inputs will not be accepted from the components development teams. If changes are performed during orbiter level I&T, changes can be sent to each of the subsystem components and instrument teams for tracking purposes Flat Sat Simulator Flat Sat is an electrical engineering model of the SC and assembled from ETUs. The purpose of the Flat Sat is to verify electrical interfaces prior to I&T, test and validate flight and GS SW, fine tune I&T procedures and scripts, train the I&T team and MOT, assist in the development of 3-4

40 mission operations procedures and scripts and to provide a realistic simulator of the orbiter and environment (particularly GN&C). Flat Sat is used during certain mission simulations, tests, and rehearsals. The simulator is maintained after launch for operations testing, training, and FSW update verification Ground Network Site Testing and Verification Once the White Sands dedicated ground station is constructed, the vendor will verify that the antenna meets certain specifications/requirements. Once the vendor has validated the site, acceptance and operational testing will be conducted. Verification of the S-Band system will be accomplished by tracking a satellite of opportunity. Flight Dynamics Facilities (FDF) certifies the tracking data. For the S-Band network sites, the ground network provider and the LRO GS will verify any upgrades. FD may still need to validate tracking data since LRO requires a higher accuracy. This requirement may drive station sites to upgrade their tracking systems Ground System Mission Readiness Testing LRO GS team will perform Mission Readiness Testing (MRT) to verify the readiness of the LRO GS to support the mission. Tests will be conducted to verify the performance of the system. All elements of the GS will be included in MRTs. MRT testing will take place after the first release of the GS architecture and conclude prior to mission Flight Operations Readiness Review (FORR). MRT includes several tests that are defined in the Mission Readiness Testing Test Plan (431-PLAN ). Goals for MRT include: Ensure the ground system supports all mission requirements as defined in the Detailed Mission Requirements for the Lunar Reconnaissance Orbiter Ground System (431- RQMT ) (Level 3) Verify the end-to-end GS functional capabilities Provide a documented test history and metrics data of system performance, discrepancies, and resolution Table 3-1 shows the different phases of mission readiness testing and test types for each phase. Table 3-1. LRO Activities by Mission Readiness Test Phase Phase #1 Element Testing Phase #2 Ground System Readiness Testing Phase #3 Operations Testing Test Ownership User Center Elements Ground System Readiness Project 3-5

41 Phase #1 Element Testing Phase #2 Ground System Readiness Testing Phase #3 Operations Testing Test Types Unit, System, Acceptance Integral Functional Operations Scenarios, Simulations, Rehearsals Purpose Internal Development Functional End-to-End verification Operational Readiness Problem Tracking Internal Unit Acceptance Discrepancy Tracking System Discrepancy Tracking System Mission Simulators and Rehearsals Test phase includes operations testing, mission simulators and rehearsals. This test phase will include SC, ground, and instrument teams. All operational products will be verified through the various tests that are defined in phase #3. Products will be validated using the Flat Sat and actual orbiter during test opportunities during orbiter I&T. Rehearsals will be planned for key mission events/phases. Currently there are five separate mission rehearsals: 1. Rehearsal #1 (Routine Operations) 2. Rehearsal #2 (Launch and Cruise Activities) 3. Rehearsal #3 (SC and Instrument Activation) 4. Rehearsal #4 (Instrument Activation) 5. Rehearsal #5 (Launch and Activation) During all mission rehearsals, the test team will inject both orbiter and GS anomalies which the mission team will need to detect and resolve. It is expected that the instrument teams will have representatives at the MOC for rehearsals that involve instrument activities. The instrument SOCs will be involved as well; data will be transferred to the SOC from the MOC according to operations plans. 3-6

42 4.0 LAUNCH AND LUNAR TRANSFER PHASE This phase covers activities from the LV countdown sequence through LOI into the commissioning orbit. There are five different sub-phases, which are: 1. Launch and Ascent 2. Separation (LRO separation from the LV third stage) 3. Deployment and Sun Acquisition 4. Lunar Cruise 5. LOI The phase is expected to last between 6 to 8 days, Figure 4-1 shows the sequence of the subphases. Start LV Countdown Sequence Orbiter in commissioning Orbit Launch and Lunar Transfer Launch & Ascent Start LV Countdown Sequence Separation Signals Detect Separation Separation Orbiter De-spun Deploy & Sun Acq Acquisition of Sun (Power Positive) Lunar Cruise Start of LOI Burns LOI Commissioning Orbit ~25-45 minutes ~10 minutes ~15 minutes ~4 days ~2 days Figure 4-1. Launch and Lunar Transfer Phase Sequence 4-1

43 4.1 LAUNCH AND ASCENT SUB-PHASE Launch and Ascent Sub-Phase starts will the execution of the LV countdown sequence. Typically, the countdown sequence starts approximately L-24 hours prior to the targeted launch time. The sub-phase ends with just prior to the orbiter separation event from the third stage Launch Countdown Sequence (L-24 hrs to Launch) The LV team and KSC develop the coordinated launch countdown sequence. The launch countdown sequence controls the activities at KSC associated with LV and orbiter preparations. The LRO Project will provide management and engineering support at KSC. The LRO team will power, configures the orbiter into launch mode, and monitors the orbiter systems prior to launch. All launch critical elements such as the MOC, FD, and initial acquisition ground stations, relays status to the GS Lead. The GS Lead provides the overall GS readiness status to the LRO launch team at KSC. The LRO GS team will have a similar countdown sequence that control activities to get the ground elements ready for launch. The GS countdown sequence will identify when status checks are performed with all launch critical elements and when certain facilities are staffed and configured for launch. Figure 4-2 shows a typical launch countdown information flow. All readiness status/information is passed to the LRO Project representative at KSC who provides the LRO status to the KSC launch team. 4-2

44 LRO Mission Critical Ground Elements Status Orbiter Launch Config./Status Status LRO Project Personnel NASA/Launch Vehicle Launch Team Ground System Lead LRO Chief Engineer Go/No-Go LRO Mission Director (Payload POC) Go/No-Go Mission Integration Manager Flight Assurance Manager NASA Advisory Manager NASA Chief Engineer Safety Range Status Launch Vehicle Mission Director NASA Launch Manager Figure 4-2. Typical Launch Day Information Flow LRO Launch Site Team Activities The LRO Engineering Team at KSC will power and configure the orbiter into the launch configuration mode. The team will use the same GSE used during I&T to power and configure the orbiter. Once the orbiter is configured into the launch mode (refer to Table 4-1) the engineering team will continue to monitor the orbiter health and status throughout the launch countdown sequence providing status information to the LRO Project representative. Table 4-1. LRO Launch Mode Configuration LRO Launch Mode Configuration (TBR) Component/System Category Configuration Component/System Category Configuration C&DH Box C&DH On LROC Inst. Off 4-3

45 LRO Launch Mode Configuration (TBR) Component/System Category Configuration Component/System Category Configuration Solid State Recorder C&DH Off LOLA Inst. Off C&DH SW Mode C&DH Bank 1(TBR) LEND Inst. Off C&DH SW Config C&DH Launch data storage (DS) filter table R-T Telemetry C&DH 16 kbps (hardline through LV) CRaTER Inst. Off LAMP Inst. Off Cmd Rate C&DH 4kbps Diviner Inst. Off Comm. Cards C&DH On Battery Power Online S-Band Receivers Comm. On PSE Power On S/Ka Band Xmitter Comm. Off Servo Drive Bus Power Off (TBR) Transponder Config. Comm. RF Switches configured for Omni Deployment Bus Power Off Reaction Wheels GNC Off Isolation Valve Prop. Closed (TBR) Star Trackers GNC Off Prop. Heaters Prop. Enabled IMU/Gyro GNC On CATBED Heaters Prop. Off (TBR) ACS SW Mode GNC Sun-Acq (TBR) Prop. Thrusters Prop. Off Survival Heater Bus Thermal Enabled SC Ops Heaters Thermal Off Inst. Survival HTRs Thermal Enabled Inst. Ops Heaters Thermal Off Deployment HTRs Thermal Enabled Mini-RF Demo OFF After the orbiter is powered, R-T telemetry will be forwarded from the GSE located at KSC to the MOC at GSFC Ground System Activities The operations and orbiter engineering teams staffs the MOC. After launch, all operations and control transfers over to the team at GSFC. The GS team will execute the GS countdown sequence ensuring all required ground elements are configured and prepared for launch. The GS representative at the MOC will relay the status information to the LRO Project representative at KSC. The GS countdown sequence will include the following activities/events: Interface checks (Data and voice) with the ground networks 4-4

46 Voice network configuration to all required ground elements including launch team at KSC Polling and status checks with launch critical elements such as FD The Orbiter Engineering Team will receive the R-T HK data while the orbiter is powered prior to launch. The engineering team will be in constant contact with the KSC team to help identify and work any issues. The operations team will ensure that all MOC systems are configured and ready to support the mission including subsystem terminals used by the engineering team Launch and Ascent At lift-off, all R-T telemetry from the orbiter to the KSC GSE will be lost. The orbiter will continue to generate HK telemetry and the C&DH system will store the data in a short-term contingency buffer on the main processor card. The main processor memory buffer will have adequate size to store 16 kilobits per second (kbps) of HK data for 2 hours. Two hours will cover the time during ascent prior to initial acquisition with the ground network. LRO launch constraints are listed in Table 4-2. With the launch window constraints listed in the table, LRO will have two opportunities each month to launch. Each launch opportunity last approximately 3 days. Each day during the launch opportunity provides a single 10-minute window. The launch window assumes a minimum energy short coast launch inserting over the lunar southern pole. Table 4-2. LRO Launch Constraints No. Constraint Rationale 1. LRO Launch Window will be constrained such that the nominal spin direction at separation is within 15 degrees of either the sun or anti-sun. Protects against immediate sun exposure down the instrument boresights Launch Vehicle Ascent Activities TBD - Describe the sequence of events for LV during ascent The launch and ascent phase concludes with the start of the of LRO separation event Launch and Ascent Orbiter Orientation During ascent and prior to payload shroud deployment, the LV upper stage spins the orbiter up to 60 revolutions per minute (rpm) about the X-axis. Due to the launch constraint, the orbiter X- 4-5

47 axis is located within 15 of either the Sun or anti-sun direction to prevent damage to the instruments. 4.2 SEPARATION AND DE-SPIN Prior to the separation event, timers located on the LV will de-spin the third stage of the rocket to under 2 rpm. The de-spin event is accomplished using a third stage yo-yo. After the de-spin event occurs, additional timers will initiate the separation event. The orbiter will monitor three separation signals and when two out of three indicate separation, the orbiter will execute the separation script. The separation script controls all orbiter activities occurring from separation through initial acquisition with the ground. After separation is detected, the separation script will transition the ACS SW mode from Sun Acquisition to Delta-H. This provides the mission a redundant de-spin opportunity in case the third stage yo-yo fails to properly de-spin the orbiter to under 2 rpm. If the orbiter is below the 2-rpm spin rate, the ACS SW will automatically transition back to Sun Acquisition Mode. Table 4-3 shows the configuration of the orbiter after separation and de-spin has occurred. The separation script will continue to command and control activities for Deployment and Sun Acquisition. At this time, it is unknown whether the thruster CatBed heaters will be powered prior to separation. Normally, the thrusters are not used without the CatBed heaters first warming the thrusters on since cold firing the thrusters can reduce the lifetime of the thrusters. Thermal analysis will determine the temperature of the thrusters at separation. The LV provider will forward the actual separation vector to FD at GSFC. FD receives the separation vector and generates a separation performance report that shows errors between the target separation vector and actual. During this sub-phase, the operations team at GSFC will be preparing for initial acquisition. The MOC will prepare to send the R-T orbiter HK data to the GSE located at KSC for the KSC launch team. The MOC T&C system forwards orbiter HK telemetry data to the KSC GSE after initial acquisition occurs. Ground based RADAR may be used to assist in early orbit determination after separation and aid initial acquisition efforts. Table 4-3. Orbiter Configuration after Separation and De-spin LRO Separation and De-spin Configuration (TBR) Component/System Category Configuration Component/System Category Configuration C&DH Box C&DH On LROC Inst. Off Solid State Recorder C&DH Off LOLA Inst. Off C&DH SW Mode C&DH Bank 1 (TBR) LEND Inst. Off C&DH SW Config C&DH Launch DS filter table CRaTER Inst. Off R-T Telemetry C&DH N/A LAMP Inst. Off Cmd Rate C&DH 4kbps Diviner Inst. Off 4-6

48 LRO Separation and De-spin Configuration (TBR) Component/System Category Configuration Component/System Category Configuration Comm. Cards C&DH On Battery Power Online S-Band Receivers Comm. On PSE Power On S/Ka Band Xmitter Comm. Off Servo Drive Bus Power Off (TBR) Transponder Config. Comm. RF Switches Deployment Bus Power Off configured for Omni Reaction Wheels GNC Off Isolation Valve Prop. OPEN STs GNC Off Prop. Heaters Prop. Enabled IMU/Gyro GNC On CATBED Heaters Prop. OFF (TBR) ACS SW Mode GNC Sun-Acq (TBR) Prop. Thrusters Prop. OFF Survival Heater Bus Thermal Enabled SC Ops Heaters Thermal OFF Inst. Survival HTRs Thermal Enabled Inst. Ops Heaters Thermal OFF Deployment HTRs Thermal Enabled Mini-RF Demo OFF Orbiter Orientation Description After separation, the +X axis is within 15 of either the Sun or anti-sun. The orbiter will be spinning less than 2 per seconds. 4.3 DEPLOYMENT AND SUN ACQUISITION SUB-PHASE This sub-phase starts have the orbiter has separated and de-spun to under 2 rpm. The separation script continues to control the activities. The sub-phase activities include: SA Boom Deployment SA deployment Indexing the SA position Power reaction wheels Acquiring the Sun Perform initial acquisition with the ground Verify orbiter telemetry at the MOC Figure 4-3 shows an estimated sequence through deployment and initial acquisition. For each deployment, at least two attempts will be contained within the separation script. The SA boom will first deploy followed by the deployment of the SA. Once the deployment activities are complete, the SA gimbals are commanded to the default index. 4-7

49 Deployment and Sun Acquisition Sequence of Events Solar Array Deployment ~ 15 minutes Sun Acquisition ~ 20 minutes Initial Acquisition Communication Checks ~ 10 minutes SA Boom Deploy SA Index Sun Acq Acquired Sun Orbiter Configuration Checks Dump SBC HK Data SA Deploy Reaction Wheels Powered Initial Acq. Cmd Check Figure 4-3. Deployment and Sun Acquisition Events The separation script continues with powering the reaction wheels. The ACS SW will process the Sun position and start driving the wheels to acquire the Sun. Once the orbiter acquires the Sun, the battery will begin charging. HK data will be generated throughout the sequence and stored on the SBC memory buffer. Table 4-4 shows the final configuration of the orbiter after the Sun Acquisition event is completed. Table 4-4. Orbiter Configuration after Deployment and Sun Acquisition LRO Orbiter Deployment and Sun Acquisition Configuration (TBR) Component/System Category Configuration Component/System Category Configuration C&DH Box C&DH On LROC Inst. Off Solid State Recorder C&DH Off LOLA Inst. Off C&DH SW Mode C&DH Bank 1(TBR) LEND Inst. Off 4-8

50 LRO Orbiter Deployment and Sun Acquisition Configuration (TBR) Component/System Category Configuration Component/System Category Configuration C&DH SW Config C&DH Launch DS filter table CRaTER Inst. Off R-T Telemetry C&DH 16 kbps LAMP Inst. Off Cmd Rate Comm. Cards C&DH C&DH 4kbps On Diviner Battery Inst. Power Off Online (Charging) S-Band Receivers Comm. On PSE Power On S/Ka Band Xmitter Comm. Off Servo Drive Bus Power On Transponder Config. Comm. RF Switches Deployment Bus Power On configured for Omni Reaction Wheels GNC ON Isolation Valve Prop. Open STs GNC Off Prop. Heaters Prop. Enabled IMU/Gyro ACS SW Mode GNC GNC On Sun-Acq CATBED Heaters Prop. Thrusters Prop. Prop. Off Off Survival Heater Bus Thermal Enabled SC Ops Heaters Thermal Off Inst. Survival HTRs Thermal Enabled Inst. Ops Heaters Thermal Off Deployment HTRs Thermal Enabled Mini-RF Demo OFF At a designated point in the timeline, the ground network will initiate initial acquisition procedures. At this point, the ground network will raise the uplink carrier and begin searching for the orbiter. The orbiter will be configured through the separation script and searching for an uplink carrier based on receiver telemetry. If the receiver detects an uplink carrier, the SC will enable the telemetry downlink. The ground network will detect the downlink and lock on telemetry. The SC is configured to output 16 kbps HK telemetry. The HK telemetry is sent to the MOC in R-T. The MOC T&C system configures a connection to the KSC GSE. The R-T orbiter HK data flows over the connection that will allow the engineering team at KSC to view the R-T telemetry. MOC and KSC launch team engineers review and provide an early assessment on the orbiter health and status. After telemetry is established, the operations team verifies the command link by sending a non-operation command to the SC. The operations team begins activities as defined in the Lunar Reconnaissance Orbiter Launch and Commissioning Handbook (431-HDBK ). The early assessment and configuration checks performed by the operations team include: Orbiter configuration based on table above SA deployment status, contingencies will be prepared to manually deploy the SA boom and cells Sun Acquisition status, verify the SA is on the Sun and generating power. Check for any telemetry limits and configuration errors 4-9

51 The ground network will begin to generate Doppler and range tracking data for FD. The early tracking data is important in the assessment of the early trajectory and re-generation of mission planning products Orbiter Orientation Description During the SA deployment activity, the orbiter rotation may stop but the orbiter X-axis is +/-15 on the Sun line. During sun acquisition, the attitude control SW orients the orbiter to place the Sun on the SA. After Sun acquisition is complete, the orientation has the Sun line on the orbiter Y axis to within +/ LUNAR CRUISE SUB-PHASE This sub-phase starts after all activities are finished during the initial acquisition procedures. The main objectives during the cruise phase is to correct trajectory errors and perform a series of lunar orbit insertion burns to capture into the commissioning orbit around the Moon. The cruise phase is expected to last between 4-6 days. Besides the necessary maneuvers, other key activities include: Configuring the GNC and ACS SW Deploying the HGA Solid State Recorder (SSR) turn-on and commissioning Subsystem activation and configuration LEND and CRaTER Instrument early turn-on After separation, LRO s propulsion system will perform all necessary maneuvers to correct the lunar trajectory and capture into the commissioning orbit around the Moon. LRO has two different size thrusters. The axial thrusters (20 lbf) will be used for LOI while the smaller thruster modules (5 lbf) will be used for MCC and routine SK maneuvers. Figure 4-4 shows the expected trajectory during the lunar cruise phase. 4-10

52 Moon at encounter Cis-lunar transfer day transfer Launch C km 2 /s 2 1-day Lunar Orbit Sun direction Earth Nominal Cis-lunar Trajectory Solar Rotating Coordinates Figure 4-4. LRO Lunar Trajectory The activities performed during the phase begin the orbiter activation plan that the Lunar Reconnaissance Orbiter Launch and Commissioning Handbook (431-HDBK ) outlines. This document provides a detail sequence of activities and indicates any constraints. The operations and engineering teams uses this plan to perform daily planning of the mission. Figure 4-5 shows an early concept for the cruise events. The two main objectives during the phase are to prepare and configure the orbiter for the first MCC maneuver and LOI maneuvers. All other events are secondary objectives and are performed according to a priority order as outlined in the Lunar Reconnaissance Orbiter Launch and Commissioning Handbook (431- HDBK ). The operations and engineering team will try to perform the MCC maneuver as early as possible. The maneuver needs to occur less than 24 hours after launch, but if the maneuver is performed earlier, it may require less fuel. 4-11

53 Figure 4-5. LRO Launch and Lunar Cruise Sequence of Events Lunar Cruise Early Activities (Day 1) Day one priorities include the following: 1. Dumping and analyzing the HK data stored on the SBC memory from launch through initial acquisition 2. Performing GNC configuration to allow the FSW to transition to observing mode (observing mode is needed to perform the first MCC maneuver) 3. Perform propulsion system checks and calibration prior to MCC 4. Turn on the SC recorder 5. Monitoring systems and characterizing the power and thermal subsystems Throughout day 1, the ground network will be providing Doppler and range tracking data to FD during all ground contacts. 4-12

54 After the operations team established a T&C link with the orbiter in the previous sub-phase, the operations teams will start to dump the HK data stored in the SBC memory buffer. Data stored in the memory buffer will be in the form of files and the SBC will utilize the CFDP to transfer the files to the ground. CFDP is a reliable transfer protocol where the ground tracks and requests missing data units from the SC in order for the ground to verify all data was successfully transferred. The S-Band system will dump the HK data files. The operations team will increase the S-Band telemetry rate from 16 kbps to kbps. The increase in data rate will allow the ground team to continue to receive R-T HK data plus the stored files. By using the higher data rate, the contents of the SBC memory buffer is expected to be dumped within 8 minutes. The ground station relays the data in R-T to the MOC where the operations team performs trending and analysis on the data. The next priority of activities involves configuring the orbiter for the MCC maneuver planned for day 2. For the MCC maneuver, the orbiter will need to transition from Sun Acquisition to observing mode. The operations team will need to power on the STs to allow the orbiter to transition to observing mode. Once the STs are powered, the engineering team monitors and verifies the trackers have acquired the appropriate stars. FD will generate an early ephemeris that the operations team uplinks to the spacecraft. The ephemeris will allow the orbiter to transition to observing mode. The operation team proceeds to the configuration of the propulsion system for the MCC maneuver. If there is time, thruster calibration burns are performed. The configuration involves commanding the appropriate valves along with thermal control of the module. If there is time in day 1, the operations team will proceed with powering the SC recorder. The SC recorder will provide greater operational flexibility in mission planning and provide protection against ground network outages. After powering the recorder, the recorder directories are configured. Table 4-5 shows the current directory structure for the SC recorder. Table 4-5. LRO Spacecraft Recorder Directory Structure (TBR) Directory SSR/LAMP SSR/LROC SSR/LEND SSR/LOLA SSR/CRaTER SSR/Diviner Description Store all LAMP measurement data Store all LROC measurement data Store all LEND measurement data Store all LOLA measurement data Store all CRaTER measurement data Store all Diviner measurement data 4-13

55 Directory SSR/HK Description Store all orbiter HK data Lunar Cruise Activities (Day 2) The main priority for day 2 is performing the MCC maneuver. After the MCC maneuver, secondary priorities include deploying the HGA and powering CRaTER and LEND. Additional day 2 activities include performing incomplete day 1 activities. The ground network will provide Doppler and Range tracking data to FD during all ground contacts. FD will deliver the MCC maneuver plan to the operations team prior to day 2. The maneuver plan will include the attitude and burn duration information. The MCC maneuver is needed to correct any launch vehicle insertion errors. The maneuver is expected to be approximately 70 meters per second (m/s). The operations team will generate the maneuver load command sequence and uplink it. The constraints for the MCC maneuver are: Orbiter is in observing mode Propulsion system is configured Thruster maneuver is performed when ground coverage is available The sequence of activities for MCC includes: 1. Turn on the propulsion CATBED heaters at least 45 minutes prior to the maneuver 2. Lock the SA and HGA gimbals to desired position 3. The orbiter will slew to the designated attitude based on the command load, orbiter will perform maneuver in observing mode using the reaction wheels 4. After attitude is reached for the maneuver, the FSW will transition to Delta-V mode with the appropriate burn information based on the maneuver plan 5. After burn is complete, the FSW will transition back to observing mode automatically 6. The command load will return the orbiter to the prior attitude 7. Propulsion system will be configured including turning off the CATBED heaters. After the MCC maneuver, LEND and CRaTER can be turned on for early calibration and measurement collection during the lunar cruise phase. Based on the turn-on sequence provided by the LEND and CRaTER SOC, the operations team will follow the turn-on procedures for each instrument. Once powered, the operations activities will focus on configuring and activating the spacecraft. LEND and CRaTER does not require any uplink of command sequences to control 4-14

56 the instruments. Once powered, the instruments operate autonomously. Constraint for turning on either CRaTER or LEND is: SC recorder must be on or the SC is operating at the kbps data rate. LEND measurement collection during cruise are necessary to characterize the neutron background. Measurements collected during this phase will be useful since the ground team will be able to identify and characterize the SC component of the background. CRaTER measurement collection during cruise phase provides the CRaTER ground team calibration data on background environment. CRaTER data rate will be variable during the phase but will not exceed 100 kbps. If the recorder is powered prior to instrument turn-on, all measurement data is stored on the recorder and is dumped when time permits. The data collected may not be dumped until after LRO reached the commissioning orbit. If the recorder is not powered, the SC needs to be operating at the kbps S-Band rate. At this rate, the bandwidth allows R-T HK and LEND/CRaTER measurement data to be downlink simultaneously. The SBC will store CRaTER and LEND measurement files using the SBC memory buffer. With the higher S-Band rate, the file is dumped as soon as it is closed on the SBC memory. This operation will not increase the storage volume of the SBC memory. The next activity for Day 2 is deployment of the HGA. The deployment of the HGA is needed for thermal reasons and it will provide flexibility to operations since all ground network assets can track LRO using the HGA. The ground will command the deployment of the HGA boom. The operations team deploys the HGA during a ground station view period. After deployment, the operations team proceeds with checking the HGA gimbal operations Lunar Cruise Activities (Day 3) Day 3 priorities include additional SC configurations and propulsion system characterizing in preparation for LOI maneuver. 1. Propulsion System Characterization 2. C&DH Configuration 3. GN&C Characterization 4. TCS Configuration 5. Ka-Band Checkout Throughout day 3, the ground network will provide FD Doppler and Range tracking data during all S-Band contacts. Once the SC recorder are powered, the CRaTER and LEND data will be stored on the recorder. The R-T S-Band rate is decreased to using either the 16 or 32 kbps rate. 4-15

57 Since the LOI burns are potentially long duration maneuvers, the propulsion system should be characterized to increase the efficiency, which reduces fuel usage. Since the majority of the fuel is used for the LOI burn, any fuel that can be saved during the burns can be used during the extended mission phase. Characterization of the propulsion system involves performing small burns with various thrusters to calculate thruster efficiency and alignment. After proper characterization, updated thruster efficiency and alignment tables can be loaded to the SC. C&DH configurations include the following: Configuring and verifying the on-board safing SW Verifying the status of the low/high speed data bus Verifying the operation of the SC recorder Time correlation and correction Removing launch tables GN&C characterization includes performing slews for calibration of the STs and gyro performance. Through the activities, updated alignment and bias tables can be loaded to the SC to improve the GN&C performance. TCS configuration is performed throughout the cruise phase. As each new component or load is powered, the TCS needs to accommodate the additional load to maintain the proper thermal environment. Thermal configurations include configuring the heater buses (instrument and SC operational heaters), and setting different temperature set points. Day 3 may also give the opportunity to turn on some of the instrument contamination heaters. The lowest priority for day 3 is performing a checkout of the Ka-band system. Checkout will be dependent on ground coverage and pointing accuracy of the orbiter during the cruise phase. If checkout can occur, the Ka-Band transmitter will be turned on over the Ka-Band ground station and data will be dumped. The space to ground transmission will use CFDP to dump data files from the recorder to the ground station Lunar Cruise Activities (Day 4) The primary focus of day 4 is preparing for the LOI burns. LOI includes a series of 5 maneuvers which will be spaced over two days. This allows all the burns to occur during the prime shift. Secondary activities will include continuing activities from day 3. Table 4-6 shows the orbiter configuration at the end of Lunar Cruise. 4-16

58 Table 4-6. LRO Configuration during Cruise Sub-Phase LRO Orbiter Lunar Cruise Configuration (TBR) Component/System Category Configuration Component/System Category Configuration C&DH Box C&DH ON LROC Inst. OFF (Contam HTRs On) Solid State Recorder C&DH ON LOLA Inst. OFF C&DH SW Mode C&DH Bank 1 (TBR) LEND Inst. ON C&DH SW Config C&DH Nominal Config CRaTER Inst. ON R-T Telemetry C&DH 16 or 32 kbps LAMP Inst. OFF (Contam HTRs On) Cmd Rate C&DH 4kbps Diviner Inst. OFF Comm. Cards C&DH ON Battery Power Online (Charging) S-Band Receivers Comm. ON PSE Power ON S/Ka Band Xmitter Comm. ON Transponder Config. Comm. RF Switches Servo Drive Bus Deployment Bus Power Power ON OFF configured for HGA Reaction Wheels GNC ON Isolation Valve Prop. Open Star Trackers GNC ON Prop. Heaters Prop. Enabled IMU/Gyro GNC ON CATBED Heaters Prop. Cycle ACS SW Mode GNC Observing Prop. Thrusters Prop. OFF Survival Heater Bus Thermal Enabled SC Ops Heaters Thermal ON Inst. Survival HTRs Thermal Enabled Inst. Ops Heaters Thermal ON Deployment HTRs Thermal Enabled Mini-RF Demo OFF Orbiter Orientation Description Throughout the Lunar Cruise phase, the orbiter control SW keeps the Y axis on the Sun line. Some commissioning activities such as the MCC and ACS sensor calibrations may require orbiter slews from this attitude. The sensor calibration slews may require a series of 90 slews off nominal attitude. During the phase, any slews will not place the Sun on the instrument boresights. 4.5 LUNAR ORBITER INSERTION LOI maneuver plan calls for up to five separate maneuvers to capture into the commissioning orbit. The commissioning orbit is a low maintenance orbit that will allow the mission to save fuel while the remaining SC and instruments are configured and checked out. LRO will attempt to find the frozen orbit around the Moon, which is believed to be at approximately 30x216 km orbit. By finding the frozen orbit, it will benefit the extended mission since LRO may return to the frozen orbit for the extended mission phase. The series of LOI maneuvers are: 4-17

59 1. Maneuver 1 Captures LRO into a 12-hour orbit. Approximately 333 m/s delta-v maneuver. 2. Maneuver 2 Captures LRO into a 6-hour orbit. Approximately 112 m/s delta-v maneuver. 3. Maneuver 3 Captures LRO into the 30x216 frozen orbit. Approximately 385 m/s delta-v maneuver. 4. Maneuver 4 Trim maneuver 5. Maneuver 5 Trim maneuver The initial plan is to perform the maneuvers 1 through 3 on one day and the two trim maneuvers on the second day. The actual sequence may change to minimize fuel usage and risk to the mission during the LOI maneuvers. For each maneuver, FD will deliver a plan to the operations team for generation and uplink to the SC. The maneuvers will be performed within view of the ground station. The HGA will be locked into a position that still allows S-Band coverage. If the HGA can not provide the coverage, the omni antenna will be used. The sequence of events for each maneuver includes: 1. Turn on the propulsion CATBED heaters at least 45 minutes prior to the maneuver 2. Lock the SA and HGA gimbals to desired position 3. The orbiter will slew to the designated attitude based on the command load, orbiter will perform maneuver in observing mode using the reaction wheels 4. After attitude is reached for the maneuver, the FSW will transition to Delta-V mode with the appropriate burn information based on the maneuver plan 5. After burn is complete, the FSW will transition back to observing mode automatically 6. The command load will return the orbiter to the prior attitude 7. Propulsion system will be configured including turning off the CATBED heaters. Between each LOI maneuver, the ground network will provide FD with Doppler and Range tracking data. LRO will be tracking continuously on every ground contact. Table 4-7 shows the configuration of the orbiter after the LOI maneuvers are complete. Table 4-7. LRO Orbiter Configuration for Lunar Orbit Insertion LRO Orbiter Configuration for Lunar Orbit Insertion (TBR) Component/System Category Configuration Component/System Category Configuration C&DH Box C&DH ON LROC Inst. OFF (Contam HTRs On) Solid State Recorder C&DH ON LOLA Inst. OFF 4-18

60 LRO Orbiter Configuration for Lunar Orbit Insertion (TBR) Component/System Category Configuration Component/System Category Configuration C&DH SW Mode C&DH Bank 1 LEND Inst. ON C&DH SW Config C&DH Nominal Config CRaTER Inst. ON R-T Telemetry C&DH 16 or 32 kbps LAMP Inst. OFF (Contam HTRs On) Cmd Rate C&DH 4kbps Diviner Inst. OFF Comm. Cards C&DH ON Battery Power Online (Charging) S-Band Receivers Comm. ON PSE Power ON S/Ka Band Xmitter Comm. ON Servo Drive Bus Power ON Transponder Config. Comm. RF Switches Deployment Bus Power OFF configured for HGA Reaction Wheels GNC ON Isolation Valve Prop. Open Star Trackers IMU/Gyro GNC GNC ON ON Prop. Heaters CATBED Heaters Prop. Prop. Enabled Cycle ACS SW Mode GNC Observing Prop. Thrusters Prop. OFF Survival Heater Bus Thermal Enabled SC Ops Heaters Thermal ON Inst. Survival HTRs Thermal Enabled Inst. Ops Heaters Thermal ON Deployment HTRs Thermal Enabled Mini-RF Demo OFF With the completion of the LOI maneuvers, LRO transitions to the Orbiter Activation and Commissioning Phase Orbiter Orientation Description TBD Diagrams and description of the orbiter orientation during the phase. 4-19

61 5.0 ORBITER COMMISSIONING Orbiter activation starts after LRO captures into the commissioning orbit of 30x216 km. If LRO finds the frozen orbit, it may not need to perform any maintenance maneuvers throughout the activation and commissioning phase. The phase is broken into two sub-phases; SC commissioning and integrated instrument commissioning, see Figure 5-1. The final activity during this phase is to transition to the nominal mission orbit of 50 km. Figure 5-1. Orbiter Commissioning Phase For the purposes of this document, the term commissioning involves two activities: 1. Activation Activities Includes powering components and configuration. It also includes functionally checking out the components. 2. Calibration/Validation (Cal/Val) Activities The main purpose for Cal/Val is to verify system performance. 5.1 SPACECRAFT COMMISSIONING SC commissioning actually starts during the lunar cruise phase. Activities during this phase are just the continuation of SC activities as defined in the Lunar Reconnaissance Orbiter Launch and Commissioning (431-HDBK ). Table 5-1 summarizes spacecraft commissioning activities. The main goal for SC commissioning is to provide a calibrated SC bus to support instruments activities. In order to minimize the amount of time for orbiter commissioning, some activities involving instrument activation may be performed during this sub-phase. Table 5-2 shows the orbiter configuration after SC commissioning is complete. Table 5-1. Spacecraft Commissioning Activities C&DH Recorder function checks Safing implementation configuration Power SA pointing calibration Power analysis Configuring power control parameters 5-1

62 Thermal Configure thermal heater buses and set points Monitor and trend orbital temperatures Communications Verify Ka comm. links Verify all S-Band rates Propulsion GNC Calculate fuel usage from LOI and determine fuel remaining Update thruster calibrations based on LOI data Monitor and calibration pointing algorithms for HGA and SA Continue component calibrations for STs and gyro Configure ACS Fault Detection Handling (FDH) Table 5-2. LRO Orbiter Configuration for Spacecraft Commissioning LRO Orbiter Configuration for Spacecraft Commissioning (TBR) Component/System Category Configuration Component/System Category Configuration C&DH Box C&DH ON LROC Inst. OFF (Contam HTRs On) Solid State Recorder C&DH ON LOLA Inst. OFF C&DH SW Mode C&DH Bank 1 (TBR) LEND Inst. ON C&DH SW Config C&DH Nominal Config CRaTER Inst. ON R-T Telemetry C&DH 16 or 32 kbps LAMP Inst. OFF (Contam HTRs On) Cmd Rate C&DH 4kbps Diviner Inst. OFF Comm. Cards S-Band Receivers C&DH Comm. ON ON Battery PSE Power Power Online (Charging) ON S/Ka Band Xmitter Comm. ON Servo Drive Bus Power ON Transponder Config. Comm. RF Switches Deployment Bus Power OFF configured for HGA Reaction Wheels GNC ON Isolation Valve Prop. Open Star Trackers GNC ON Prop. Heaters Prop. Enabled IMU/Gyro GNC ON CATBED Heaters Prop. Cycle ACS SW Mode GNC Observing Prop. Thrusters Prop. OFF Survival Heater Bus Thermal Enabled SC Ops Heaters Thermal ON Inst. Survival HTRs Thermal Enabled Inst. Ops Heaters Thermal ON Deployment HTRs Thermal Enabled Mini-RF Demo OFF 5-2

63 5.1.1 Orbiter Orientation Description During SC commissioning, the orbiter pointing is nominally nadir towards the Moon surface. Some commissioning activities such as the ACS sensor calibrations may require orbiter slews from this attitude. The sensor calibration slews may require a series of 90 slews off nominal attitude. During the phase, any slews will not place the Sun on the instrument bore-sights. Slews during the phase should not last more than 15-minutes. 5.2 INTEGRATED INSTRUMENT COMMISSIONING Instrument commissioning includes both activation and calibration/validation activities. During instrument commissioning, the instrument teams will bring GSE and engineering personnel to the MOC to work closely with the operations team. The MOC will provide adequate space to accommodate instrument personnel and equipment. During this phase, the MOC will also forward R-T and stored instrument data to the instrument SOCs. The instruments will be commissioning in a very controlled manner, under the coordination of the mission directory, the operations team, and instrument teams. Procedures will have been prepared in advance and coordinated with the operations team. The procedures will also have been exercised during the various simulations and rehearsals. Each instrument engineering team is expected to develop the initial instrument activation sequence. The sequence may include some of the items: Turn-on sequence Deployment of doors and covers Configuring electrical systems Configuring instrument SW Functionally checking out instrument components The operations team will integrate each of the individual instrument activation plans and form an integrated instrument activation plan. The integrated activation plan will be captured in the Lunar Reconnaissance Orbiter Launch and Commissioning (431-HDBK ). As the second part of instrument commissioning, it is expected that the science working group (SWG) will develop the mission calibration and validation plan for the instruments. The purpose of the Ca1/Val activities is to verify the performance of each instrument. In order to complete Cal/Val, the instruments will most likely need to perform a series of activities that involve slewing to different targets. As details become clear, the SWG and operations team will look at the different activities and determine the best plan if any activities can be done in parallel. The activities along with the description will be in the Lunar Reconnaissance Orbiter Launch and Commissioning (431-HDBK ). 5-3

64 5.2.1 Orbiter Orientation Description During instrument commissioning, the orbiter pointing is nominally nadir towards the Moon surface. Some instrument commissioning requires the orbiter to slew to different targets such as stars. The slews generally will last less than 15-minutes and will not place the Sun on the instrument bore-sights Lunar Reconnaissance Orbiter Camera Commissioning Overview TBD Lunar Orbiter Laser Altimeter Commissioning Overview TBD Lunar Exploration Neutron Detector Commissioning Overview TBD Lyman-alpha Mapping Project Commissioning Overview LAMP activation will start once the orbiter has reached the commissioning orbit. The first steps include power and HK tests. These activities include the following: Low-Voltage power on Low and high-speed data link checkout Test detector acquisition modes (stim pixels/test image pattern) HK checks SC interface checks LAMP Terminator Sensor (LTS) checks Open entrance aperture door (fire TiNi launch latch actuator and command door to open/close states) LAMP commissioning activities will include the following activities: Perform optics decontamination Open detector door using the wax pellet actuator Slow ramp-up of the detector high voltage to full operating level Test images of ISM H Lyα background, dark rates, and Ultraviolet (UV) star scans Verify optical boresight pointing using UV star scans 5-4

65 Verify/calibrate in-band sensitivity using UV stars with known in-band flux calibrations Cosmic Ray Telescope for effects of Radiation Commissioning Overview TBD Diviner Commissioning Overview TBD Table 5-3 shows the orbiter configuration at the end of instrument commissioning. Table 5-3. LRO Orbiter Configuration after Instruments Commissioning LRO Orbiter Configuration after Orbiter Commissioning (TBR) Component/System Category Configuration Component/System Category Configuration C&DH Box C&DH ON LROC Inst. ON Solid State Recorder C&DH ON LOLA Inst. ON C&DH SW Mode C&DH Bank 1 LEND Inst. ON C&DH SW Config C&DH Nominal Config CRaTER Inst. ON R-T Telemetry C&DH 16 or 32 kbps LAMP Inst. ON Cmd Rate C&DH 4kbps Diviner Inst. ON Comm. Cards C&DH ON Battery Power Online (Charge/Discharge) S-Band Receivers Comm. ON PSE Power ON S/Ka Band Xmitter Comm. ON Servo Drive Bus Power ON Transponder Config. Comm. RF Switches Deployment Bus Power OFF configured for HGA Reaction Wheels GNC ON Isolation Valve Prop. Open Star Trackers GNC ON Prop. Heaters Prop. Enabled IMU/Gyro GNC ON CATBED Heaters Prop. OFF ACS SW Mode GNC Observing Prop. Thrusters Prop. OFF Survival Heater Bus Thermal Enabled SC Ops Heaters Thermal ON Inst. Survival HTRs Thermal Enabled Inst. Ops Heaters Thermal ON Deployment HTRs Thermal Enabled Mini-RF Demo OFF The final mission activity for commissioning phase is to insert LRO into the mission orbit. It is expected that an additional two burns will be needed to insert into the final mission orbit of 50km. The maneuver sequence is similar to the plan during LOI and is listed below: 1. Turn on the propulsion CATBED heaters at least 45 minutes prior to the maneuver 2. Lock the SA and HGA gimbals to desired position 3. The orbiter will slew to the designated attitude based on the command load, orbiter will perform maneuver in observing mode using the reaction wheels 5-5

66 4. After attitude is reached for the maneuver, the FSW will transition to Delta-V/H mode with the appropriate burn information based on the maneuver plan 5. After burn is complete, the FSW will transition back to observing mode automatically 6. The command load will return the orbiter to the prior attitude 7. Propulsion system will be configured including turning off the CATBED heaters 5.3 TECHNOLOGY DEMOSTRATION COMMISSIONING During the commissioning period, the operations team will allocate time to power and functionally check out the Mini-RF payload prior to orbiter maneuvers to insert into the final mission orbit. A summary of activities for Mini-RF functional checkout are below: TBD 5-6

67 6.0 MEASUREMENT OPERATIONS A measurement operation is the routine mission phase for LRO. The phase starts when commissioning is complete and the orbiter is inserted into the nominal mission orbit of 50km. Measurement operations are broken into five different sub-phases, which are: Routine Operations Nominal day to day orbiter operations SK Monthly orbit maintenance maneuver along with required instrument calibrations Momentum Management Twice a month, the SC is commanded to Delta-H to adjust system momentum. Instrument Calibrations Monthly activities to perform instrument calibrations. Calibrations will be coordinated before SK maneuvers. Lunar Eclipse Long duration eclipse when the Earth passes between the Moon and the Sun. Yaw Maneuver Twice a year LRO will perform an 180 yaw maneuver Safe Mode Low power mode to protect the orbiter systems while the ground investigates the problem 6.1 LUNAR RECONNAISSANCE ORBITER S UNIVERSE AT THE MOON Figure 6-1 shows a graphical representation of LRO s universe at the Moon. The figure shows the key events that occur during monthly and yearly cycles. 6-1

68 Figure 6-1. LRO s Universe at the Moon Some of the key events and descriptions are below: Twice a year, LRO will be in full Sun for roughly one month for each event. The full Sun condition occurs when the orbit Beta angle reaches ~76. During the one month period, the orbiter should not experience any lunar occultation. During the eclipse season (the shaded portion of the Sun circle), LRO is expected to have a maximum lunar occultation of 48 minutes. The maximum duration lunar occultation will occur when the orbit beta angle reaches

69 Twice a year, LRO will be required to perform a 180 yaw maneuver. The purpose is to keep the Sun on the correct side of the SC for power and thermal reasons. The yaw maneuvers will be performed as the orbit approaches the beta 0 condition. Twice a month, LRO s orbit will be in full view of the Earth for a period of ~2 days. During the 2 days, ground stations on the Earth will have continuous view of LRO. However, due to the HGA range of motion limitation, the HGA ground contacts will still be limited to approximately 56 minutes. The omni antennas can provide continuous coverage. Twice a month, LRO will need to perform momentum management. The SC will be commanded to Delta-H mode. The maneuver will occur when the ground has complete coverage of the orbit (as defined above). Once a month, LRO will be required to perform SK maneuvers to maintain the mission orbit. SK maneuvers are a series of small maneuvers to adjust the orbit. The SK maneuver will be performed when the ground has complete coverage of the orbit. Twice a year (on average), the Earth will pass between the Moon and the Sun (Lunar Eclipse). The actual eclipse duration will vary, Table 6-1 shows all of the predicted lunar eclipses that might impact LRO. Table 6-1. Lunar Eclipse Predicts for during the LRO Mission Date Eclipse Type Eclipse Duration February 9, 2009 Penumbral - July 7, 2009 Penumbral - August 6, 2009 Penumbral - December 31, 2009 Partial 62 minutes June 26, 2010 Partial 164 minutes December 21, 2010 Total 73 minutes (Total Eclipse) 209 minutes (Total + Partial) June 15, 2011 Total 101 minutes 220 minutes (Total + Partial) December 10, 2011 Total 52 minutes 213 minutes (Total + Partial) June 4, 2012 Partial 128 minutes November 28, 2012 Penumbral - April 25, 2013 Partial 32 minutes May 25, 2013 Penumbral - October 18, 2013 Penumbral - 6-3

70 Date Eclipse Type Eclipse Duration Penumbral Eclipse The Moon passes through the Earth s penumbral shadow. These events are subtle and quire difficult to observe. Partial Eclipse Total Eclipse A portion of the Moon passes through the Earth s umbral shadow. The entire Moon passes through the Earth s umbral shadow. Two times are provide, first is the time for the total eclipse period and the second is the start to finish time that includes partial and total eclipse times. 6.2 ROUTINE MEASUREMENT OPERATIONS LRO will operate in this sub-phase for the majority of the mission. LRO mission orbit will have the following characteristic: Orbit: 50km (+/- 20km) Orbit Inclination: ~90, inclination drifts approximately 0.5 per year Orbit period: ~113 minutes The nominal pointing for the orbiter is nadir looking at the Moon s surface. It is expected that some instruments may require off-slewing. These activities will be coordinated across all instrument teams. Although the main goal is to collect measurement data for one year of measurement operations, throughout the year there are known interruptions to the measurement data capture. Table 6-2 tries to capture the total orbits LRO will fly during the one-year prime measurement period and identify known interruptions during the one year. The table does not take into account unplanned interruptions such as orbiter safing and other contingencies. The table also assumes varies durations for some activities that have not been clearly defined yet. As the mission definition progress, the table will be updated. 6-4

71 Table 6-2. LRO Measurement Interruptions LRO Measurement Data Interruptions Length of Measurement Phase: Number of Orbits during Phase: 1 year orbits Interruptions Type Frequency Duration Total orbits per Year (Orbit) Station-Keeping 1 month 1 12 orbits Yaw-Maneuver 2 year orbits Monthly Cals 1 month 3 36 orbits Momentum Management 2 month orbits Total Interruption to Measurement Collection 61 orbits Percentage of Time to 1 year Mission: 1.31% LRO Instruments Operation Instrument operations will vary from instrument to instrument on LRO. Some instruments operate only over a portion of the orbit while others operate continuously. Figure 6-2 shows a graphical representation of each of the instruments. The figure shows when instruments are collecting data and any key events. Additional details on instrument operations will be contained in the following sections. 6-5

72 LRO Baseline Instruments Operating Modes ~105 W ~96 W ~103 W ~105 W Instrument Avg. Orbital Power ~90 W ~90 W ~88 W ~88 W ~6.35 Mbps Instrument Avg. Data Rate ~0.14 Mbps ~0.14 Mbps LAMP Operating Mode Constant Data Rate ~20.2 kbps HV Disable HV Enable Constant Data Rate ~20.2 kbps CRaTER Operating Mode Diviner Operating Mode LEND Operating Mode LOLA Operating Mode Constant data rate ~100 kbps (During Flares), Non flares ~0.2 kbps Constant data rate ~10.6 kbps Constant data rate ~0.035 kbps Constant data rate ~10 kbps LROC Operating Mode WAC (1 Mbit Image/Sec) NAC 2 (Each Image ~256 MB) NAC 1 (Each Image ~256 MB) WAC LROC Power Cycle Non Sun Lit, Total time ~56.5 minutes Moon s Pole Sun Lit ~56.5 minutes Non Sun Lit, Total time Moon s ~56.5 minutes Pole Figure 6-2. LRO Integrated Instrument Operations Concept LROC Instrument Operations Concept LROC consists of three cameras; two NAC and one WAC. The two NACs will take a maximum of 16 image pairs (total of 32 images) each orbit. Each NAC image has a maximum size of 256 Megabytes (MB). It requires the NAC approximately 15 seconds to fill the NAC image and another 206 seconds to transfer the image to the SC recorder. Before transferring the image to the recorder, the LROC SCS has the option to compress the image. The LROC operations concept is based on images being compressed. If images are not compressed, the number of images LROC can take per orbit is reduced so that the total data volume is not exceeded. The WAC operates continuously over the Sun-lit portion of the orbit. While the NACs and WAC will operate majority of the time over the Sun-lit portion of the orbit, occasionally images are taken on the night portion of the orbit for special measurements and calibrations. 6-6

73 The images the NAC collect will be stored on the SC recorder. Each file will contain one NAC image providing a maximum of 32 files per orbit from the NACs. The WAC operates continuously; the current plan calls for the WAC file to be closed every 10 minutes. LROC will produce approximately 55 files per orbit. LROC requires a daily command sequence to control when the WAC and NACs image. The LROC SOC generates the command sequence and delivers the file to the MOC located at GSFC. Details for the command sequence are TBD and will be defined in the Lunar Reconnaissance Orbiter Ground System Interface Control Document (431-ICD ). The MOC will process the LROC command sequence and uplink it to the orbiter. To protect the instrument from single-event upsets (SEUs), LROC will be powered cycle every orbit. The power cycle will occur over the night portion of the orbit and the exact time for each orbit will be provided by the LROC SOC. The LROC SOC will send the MOC times each orbit to power cycle. The MOC will incorporate the power cycle sequence in the daily command load. The following outlines the steps required to power cycle the LROC instrument: 1. Remove main 28 Volts (V) power supply to LROC, such that the voltage drops below 1V for a minimum of 10 (TBR) seconds. 2. Cease all SpaceWire transmissions to LROC. 3. Close any open files in the SSR/LROC/ directory of the SSR. 4. Re-apply main 28V power to LROC and wait for 10 (TBR) seconds. 5. Resume timecode and bufferstatus command transmission. LROC initial flight rules and constraints are provided in Table 6-3. Table 6-3. LROC Flight Rules and Constraints (TBR) No. Type Rule/Constraint LOLA Instrument Operations Concept LOLA operates continuously over the entire orbit. The instrument pulses the 1064 nanometer (nm) LASER and measures the reflected return. Once the instrument is commissioning, it requires minimum operations control. The instrument will operate autonomously collect measurement data at approximately 10 kbps rate. 6-7

74 LOLA generates the measurement data along with HK data and sends both types to the SC processor. The SC processor receives the data and stores the data on the SC recorder. Each data file on the recorder will contain approximately 10 minutes worth of measurement data. Since LOLA operates autonomously, the instrument does not require a daily command sequence. The LOLA SOC will send any command request to the MOC as defined in the Lunar Reconnaissance Orbiter Ground System Interface Control Document (431-ICD ). LOLA initial flight rules and constraints are provided in Table 6-4. Table 6-4. LOLA Flight Rules and Constraints No. Type Rule/Constraint 1 Nadir Pointing Maintain SC pointing to within +/-1 of nadir for greater than 97% of the measurement phase LEND Instrument Operations Concept LEND operates continuously over the entire orbit. Once commissioned, the instrument requires minimum operations control. The standard operations for LEND is to turn on all anticoincidence levels on the high-energy neutron sensor. The standard operations mode will provide the main data products for LEND. In reduced operations mode, the instrument switches between different levels of anticoincidence for the high-energy neutron sensor. While the LEND instrument is on, the instrument will generate approximately 0.6 kbps of measurement data. The SC processor receives the data and stores the data on the SC recorder. The processor will open a new data file on the recorder every orbit. This concept generates a LEND file size of approximately 4.7 Megabits (Mbits). Currently there is no command sequence required for LEND. Once the instrument is commissioning, the instrument will operate continuously without further control. A process will be defined in the Lunar Reconnaissance Orbiter Ground System Interface Control Document (431-ICD ) that will allow the LEND SOC to make command requests. LEND initial flight rules and constraints are provided in Table 6-5. Table 6-5. LEND Flight Rules and Constraints (TBD) No. Type Rule/Constraint 6-8

75 LAMP Instrument Operations Concept LAMP collects measurement data only over the night portion of the orbit. LAMP will be in pixel list acquisition mode all the time. When the orbiter crosses the lunar terminator (day to night), the high voltage (HV) is ramped up to full operating level. The ramp-up is triggered by either a LTS or time-tagged command sent by the SC. Prior to crossing the terminator again (night to day), the high voltage is ramped down to safe level. Again, the LTS or a SC command triggers the change in high voltage level. While LAMP is operating over the night portion of the orbit, the instrument will generate measurement data at a rate of 20.2 kbps. The SC will collect and store the data on the SC recorder. Currently, the SC will store 10 minutes worth of measurement data in each data file. Each LAMP measurement data file will be approximately 12 Mbits. The LAMP SOC will generate any required command sequence updates. At this time, LAMP does not require any daily command sequence. The LAMP Terminator Sensor (LTS) or the daily command load that the mission operations team generates controls high voltage control. The daily command load will trigger commands based on the terminator predict times generated by FD. Table 6-6 provides a list of flight rules and safety constraints for the LAMP instrument. Table 6-6. LAMP Flight Rules and Constraints No. Type Rule/Constraint 1 Sun Avoidance LAMP shall not point to within ± 20 of the Sun when the aperture door and/or the failsafe door is/are open and the detector high voltage power supply (HVPS) are operating. 2 Sun Avoidance LAMP shall not point to within ±15 of the Sun when the aperture of failsafe doors are open and the detector HVPS are not operating. 3 Hot UV Star Avoidance No portion of the LAMP slit shall never be pointed to within ±0.5 of a set of identified hot UV stars when the aperture door is open and the detector HVPS are operating. Estimated number of hot UV stars: ~ Dayside Avoidance LAMP detector HV shall not be operated to full operating level during the dayside portion of the orbit. 5 Dayside Avoidance LTS will be used to determine the terminator crossings and potentially damaging peaks of eternal sunshine at the terminator. 6 HV Operation Ambient pressure must be less than 5x10-6 Torr for safe detector HV operation. 7 Temperature If instrument temperatures (6) exceed the set threshold value (one threshold per temperature sensor) during operation, the instrument will go to safe mode and then power off. 6-9

76 CRaTER Instrument Operations Concept CRaTER operates continuously over the entire mission orbit. CRaTER generates a minimum data rate of 0.2 kbps but the instrument can generate up to 100 kbps during solar flare events. If flares occur, the event could last for several days. The SC processor receives the instrument measurement data and stores the data in files on the SC recorder. CRaTER data files will be limited to 8 Mbits during flare conditions (data rate of 100 kbps) or one file per orbit during low rate measurement data collection. The CRaTER instrument operates autonomously once powered and configured. A daily command sequence is not required for the instrument. However, the CRaTER instrument team wants to reset the instrument once per day. A stored command load executing on the SC processor sends the reset command. The CRaTER SOC will send either a file with daily times or setup a predetermine time each time for the reset. The operations team at GSFC will trigger the reset command in the stored command sequence based on the inputs provide by the CRaTER SOC. The CRaTER SOC will send any command request to the MOC as defined in the Lunar Reconnaissance Orbiter Ground System Interface Control Document (431-ICD ). CRaTER initial flight rules and constraints are provided in Table 6-7. Table 6-7. CRaTER Flight Rules and Constraints (TBR) No. Type Rule/Constraint Diviner Instrument Operations Concept Diviner operates continuously over the entire mission orbit collecting data at a rate of 10.6 kbps. Diviner performs automated calibration checks several times per orbit. The sequence is trigger from a command notification coming from the stored command load on the SC indicating an equator crossing. Diviner performs between 10 to 12 in-flight calibrations per orbit. When the in-flight calibration sequence is initiated, the instrument rotates the telescope off-nadir. The total sequence takes approximately 10 seconds (2 seconds to move telescope off nadir, 4-6 seconds to stare, 2 seconds to rotate the telescope back to nadir). A table within the instrument controls the sequence. The Diviner teams expect to update this table periodically throughout the mission. The SC processor will collect the instrument measurement data and store the data in files on the SC recorder. Currently, the SC is planning to store 10 minutes of measurement data per file. This gives an average data file size for Diviner of 6.5 Mbits. 6-10

77 The Diviner SOC will forward any updated control table to the MOC for uplink to the Diviner instrument. The operations team will generate a set of equator crossing times based on FD predicts for the daily command sequence. Any other command request from the Diviner SOC is performed as defined in the Lunar Reconnaissance Orbiter Ground System Interface Control Document (431-ICD ). Diviner initial flight rules and constraints are provided in Table 6-8. Table 6-8. Diviner Flight Rules and Constraints No. Type Rule/Constraint 1 Sun Avoidance The instrument shall never look at the Sun, even for a few milliseconds (ms) during nominal or emergency operations. Damage to focal plane detectors will be destroyed Mini-RF Technology Demo Operations Concept Mini-RF operates on a non-interference basis with on-going instrument operations. Mini-RF cannot operate while LROC is collecting and transferring data to the SC due to data bus bandwidth constraints. At this time, the basic operational concept is to provide the Mini-RF with a minimum of 12 data collection opportunities during the 1-year mission. Each data collection opportunity has a data collection duration of 4-minutes. To accomplish the minimum goal, a data collection period is planned in conjunction with the monthly SK maneuvers. The SK maneuvers will interrupt nominal instrument data collection. This concept minimizes the impact to instrument operations while providing the mini-rf the minimal data sets. Additional data collection opportunities will be coordinated with the operations team on a non-interfering basis with LROC operations and the SC power and data storage constraints. The Mini-RF will operate in one of eight data collection modes. These are defined by the radar frequency (2.385 gigahertz [GHz], S-band or 7.14 GHz, X-band), the image resolution (150 meters [m] or 10m), and the polarization matrix (dual polarization or quad polarization). The pulse bandwidths are 25 MHz at 10-m resolution and 1.7 MHz at 150-m resolution. Either phase or frequency modulation may be employed. The particular radar frequencies may be adjusted as the design and mission constraints evolve, however the modulation bandwidth will remain fixed. The Mini-RF instrument radiates pulses in at a fixed rate within a burst. While the detailed design is still underway, a candidate timeline is shown in Figure 6-3. In all imaging modes, the Mini-RF instrument will radiate 42.9 microseconds (μs) pulses with a peak power of 300 watts (w) at a pulse rate of 4200 or 9300 Hz during a burst. Bursts consist of between 90 and 2967 pulses, last between 21 and 319ms, and occur at a rate between 3.8 and 0.9 Hz, depending on the image resolution. The burst duty cycle varies between 8% and 30%. The burst schedule may be 6-11

78 interrupted to produce frames in place of continuous strip-map images as according to mission requirements and constraints. The particular values cited here (peak power, pulse length, etc.) are subject to change as the system design evolves, but any changes will not increase the direct current (DC) power or thermal loads on the SC. 266 ms one burst 21 ms 90 pulses mapping resolution dual polarization 198 pulses mapping resolution quad polarization 1063 ms one burst 319 ms 1340 pulses zoom resolution dual polarization 2967 pulses time zoom resolution quad polarization Figure 6-3. Candidate timeline for pulse and burst transmission at both 150-m (mapping) and 10-m (zoom) resolutions. The pulse schedule within a burst is shown in Figure 6-4. The peak-radiated power is 300w during an individual pulse. The pulse duty cycle is 15% at both mapping and zoom resolutions. The averaged radiated power during a burst is therefore 45w. The average radiated power averaged over the burst period is 3.6 w at the mapping resolution where the burst duty cycle is 8%, and 13.5 w at the zoom resolution where the burst duty cycle is 30% 6-12

79 238 µs 300 w peak radiated power 15% pulse duty cycle during a burst 45 w average radiated power during a burst 35.7 µs 4200 Hz dual polarization 107 µs 16.1 µs 9300 Hz quad polarization time Figure 6-4. Candidate timeline for pulse transmission within a burst at both 150-m (mapping) and 10-m (zoom) resolutions The main modes of operation are off (no power consumption), standby (only the STALO is powered, and possibly the processor in a sleep mode), and data collection (all elements powered). Calibrate modes will be implemented, but these will be entered for short periods of time (less than 10 seconds) and are essentially the same as data collection mode with respect to power and thermal considerations Mini-RF Operations Details The length of the radar image is determined by the amount of time the radar is operating. In the 50km mission orbit, the Mini-RF will image a swath width between 3 and 6km on the lunar surface, depending on the imaging mode. The goal is be operate the radar at latitudes larger than 80 of at least one lunar pole. The two primary reasons for this are (1) to assure overlap with the Chandrayaan-1 Forerunner S-Band data, and (2) to assure overlap between sequential passes at the same or contracting modes of the Mini-RF data. All data is collected at a 45 incidence angle, nominal beam center for the S-band and X-band. The current plan is to mount the radar antenna on the SC body such that the radar is at the 45 incidence angle when the SC is nadir pointing. This mounting of the antenna will not drive a requirement to roll the SC. It is required to collect sufficient data near the poles to support analyses of data from all modes. To help assure sufficiency, it is desired to operate the Mini-RF in one or more modes during passes as the orbit approaches maximum latitude, through polar closest approach, and down to nominally 83 o (or 80 o ). The length of radar on time may depend on mode, and resources available from the LRO host. In the event that radar operational time may be limited during any one pass, the neighborhoods at highest latitudes have highest priority. 6-13

80 Data Generation Rates The Mini-RF instrument will generate data at peak rates of 6.8 Mbps (S-band, 150-m resolution), 5.9 Mbps (X-band, 150 m resolution), or 88 Mbps (10-m resolution), averaged over the 286 µs pulse repetition interval. The corresponding rates averaged over the burst period are 0.54, 0.47, and 26.4 Mbps. Data will be buffered within the Mini-RF instrument so as to keep the data transfer rates within limits imposed by the mission. These limits must not result in less than (TBD) bits of data being transferred to the SSR during a single SC orbit cycle Data Transfer Protocol The mini-rf sends data to the SC as CCSDS fixed length packets via file transfer Orbiter Orientation Description (Routine Operations) During routine operations, the orbiter points nadir looking (+Y-axis) at the Moon s surface throughout the entire orbit (see Figure 6-5). In order to accomplish this, the orbit performs a one revolution per orbit about the Y-axis. Depending on the year, the orbiter velocity vector will be either +X or X axis. Figure 6-5. Orbiter Measurement Operations Pointing 6-14

81 LRO will fly in a mean 50km orbit that will have a +/-15km variation from the mean altitude. The Sun will always be located in the Y hemisphere and the +Y side of the orbiter is the coldest since it will always point towards deep space Off-nadir Measurement Operations During routine operations, some instruments desires to slew off nadir to collect additional measurement data. At this time, all potential off-nadir slews have not been defined. In general, any off-nadir slews will be limited to 15-minutes in duration and less than 20 from nadir pointing. Any slews will not violate any flight constraints such as Sun avoidance, thermal and power Daily Data Volume during Routine Operations The daily data volume generated from LRO should be fairly consistence with the exception of the number of LROC images and CRaTER s data rate based on flare conditions. LROC can generate a maximum of 16 NAC pairs per orbit, but a minimum of 9 NAC pairs per orbit. Table 6-9 and Table 6-10 show the LRO daily data volume calculations based on the LROC minimum and maximum data generation. Table 6-9. LRO Minimum Daily Data Volume (LROC 9 NAC Pairs per Orbit) 6-15

82 Table LRO Maximum Daily Data Volume (LROC 16 NAC Pairs per Orbit) The data calculations above also assume that SC HK data will be stored on the recorder at a rate of 32 kbps. Figure 6-6 and Figure 6-7 show a high-level data path based on the numbers from the tables above. The figures show that some that the SC recorder does not store all of the data overhead. The Reed-Solomon (RS) encoding and convolutional encoding gets applied to all data at the communication cards. The ground station will receive the telemetry stream and perform convolutional and RS decoding on the data stream. 6-16

83 Figure 6-6. Orbiter Data Path (Minimum Data Case) 6-17

84 Figure 6-7. Orbiter Data Path (Maximum Data Volume Case) Space to ground Link Communications LRO requires a Ka-band station for high rate measurement data dumps and S-Band for R-T telemetry, tracking and control (TT&C). To support the daily data volume, LRO plans to have a dedicated Ka-Band station located at White Sands. The dedicated station will also provide S- Band communications with the SC. The WSC ground station will provide between 4 to 6 contacts a day. The number of contacts are determined by the view period the ground station has with the Moon. The plan is to maximize the use of the WSC ground station and use existing S- Band sites to provide the required 24-hr S-Band coverage. Figure 6-8 shows the current concept for the ground network. 6-18

85 Figure 6-8. Ground Network Support Concept One note of the figure, pass durations are the minimum pass duration requirements. Figure 6-9 shows the actual pass contact concept that will be used to ensure that the required pass durations are obtained. For nominal operations, the SC recorder will be dumped when the orbiter is in contact with the Ka-band station. The data files stored on the recorder will be dumped using CFDP. CFDP is a reliable transportation method that will ensure all data packets are received on the ground before the SC deletes the file. For nominal operations, the S-Band sites forwards the entire S-Band telemetry stream to the MOC at GSFC. Nominally, the highest rate for the S-Band stations will be 128 kbps. This rate allows R-T HK data along with any high priority stored data files. In order to meet the orbit determination requirements, flight dynamics need approximately 30 minutes worth of tracking data (Doppler and Ranging) every orbit. The plans call for LRO to have an S-Band contact for at least 30 minutes every orbit. When using the WSC antenna, LRO 6-19

86 will have concurrent S and Ka coverage. All S-Band sites will collect tracking data and forward the data to flight dynamics at GSFC for processing. The required accuracy for tracking data is: Doppler Tracking: 8 mm/sec (TBR) Range Tracking: 15m During the S-Band contacts, the operations team will perform daily commanding and monitor the health and status of the orbiter. LRO will use a 4 kbps uplink rate to perform all ground to orbiter commanding. All commands are generated and sent from the MOC located at GSFC. In order to ensure that the pass duration requirements are met, the actual pass concept is shown in Figure 6-9. Due to limited range of motion of the HGA gimbals, the HGA will only have view of the Earth ground stations for approximately half the orbit. The S-Band transmitter will be configured to turn-on at the start of the predicted acquisition of signal (AOS) based on FD predicts. Once the ground station locks onto the telemetry stream, the MOC will verify the uplink by sending a test command. After the test command verifies, the Ka-Band transmitter will be turned on. The Ka-Band transmitter needs to warm up before data is dumped. The warm-up period will last approximately 4 minutes. After the warm-up period, the SC recorder will be commanded to dump data files. The operations team will plan at least a 45 minutes data dump and pause file transactions within a few minutes of predicted Loss of Signal (LOS). At LOS, the Ka and S Band transmitter will be powered off through the stored command load. With this concept, the transmitters S-Band transmitter are on for the entire view period and the Ka-band transmitter are on for at least 50 minutes. Throughout the mission LRO will be required to use different telemetry modes. The orbiter will provide different telemetry rates for both Ka and S band frequency. The different rates provide the operations team flexibility during certain activities. Table 6-11 shows the current telemetry mode definition. The LRO mission uses CCSDS to package the various data types collected on-board the SC. Table 6-12 shows the current virtual channel assignments based on the data type and where the data is originating. 6-20

87 Ka-Band Transmitter Off Ka-Band Transmitter On Ka-Band Transmitter Off AOS LOS Turn-off Ka Warm-Up Period Ka-Data Dump (~45 mins) Pause File Dumps S-Band Transmitter Off S-Band Transmitter On S-Band Transmitter Off AOS S-Band 2-Way Link Established LOS Turn-off Non-View Time ~58 minutes HGA View of Earth Ground Site Lunar Orbit Non-View Time ~113 minutes HGA has view of an Earth Ground Station, Based on HGA Gimbal range of motion HGA does not see Earth Ground Station, due to orbit or HGA Gimbal range of motion Figure 6-9. Ground Station Contact Concept 6-21

88 Table LRO Space to Ground Telemetry Modes Mode S-Band Command Mode 1 S-Band Command Mode 2 Data Rate Coding Modulation Ranging Purpose (bps) (1) 4x10 3 Uncoded PCM/PSK/ON Yes Standard Command Mode 4x10 3 Uncoded PCM/PSK/ON No Contingency Command Mode S-Band Telemetry Mode 1 S-Band Telemetry Mode 2 S-Band Telemetry Mode 3 S-Band Telemetry Mode 4 S-Band Telemetry Mode 5 S-Band Telemetry Mode 6 S-Band Telemetry Mode 7 S-Band Telemetry Mode 8 S-Band Telemetry Mode 9 S-Band Telemetry Mode 10 S-Band Telemetry Mode 11 S-Band Telemetry Mode x10 3 2x x x x x x10 3 Rate ½ and RS Rate ½ and RS Rate ½ and RS Rate ½ and RS Rate ½ and RS Rate ½ and RS Rate ½ and RS Rate ½ and RS 273.3x10 3 Rate ½ and RS 273.3x10 3 Rate ½ and RS 2.186x10 6 Rate ½ and RS PCM/PSK/PM No Contingency Telemetry Mode PCM/PSK/PM PCM/PSK/PM No Yes Contingency/Low Rate Telemetry Mode Contingency/Low Rate Telemetry Mode PCM/PSK/PM No Contingency Telemetry Mode PCM/PSK/PM PCM/PSK/PM PCM/PSK/PM PCM/PSK/PM PCM/PSK/PM BPSK OQPSK Yes Yes Yes Yes Yes No No 5x10 6 Uncoded OQPSK No Nominal R-T Telemetry Mode Nominal R-T Telemetry Mode High Rate R-T Telemetry Mode High Rate S-Band with S- Band Data Dump High Rate S-Band with S- Band Data Dump High Rate S-Band with S- Band Data Dump High Rate S-Band with S- Band Data Dump High Rate S-Band with S- Band Data Dump Ka-Band Telemetry Mode 1 25x10 6 Rate ½ and RS OQPSK 6-22 No Low Rate Ka-Band contingency rate Ka-Band Telemetry 50x10 6 Rate ½ and OQPSK No Medium Rate Ka-Band

89 Mode Data Rate Coding Modulation Ranging Purpose (bps) (1) Mode 2 RS contingency rate Ka-Band Telemetry Mode 3 100x10 6 Rate ½ and RS OQPSK No Nominal Ka-Band downlink rate Notes: (1) LRO Data Rate prior to any coding Table LRO CCSDS Virtual Channel Definitions LRO Virtual Channel Definition VC ID Location Type Data Type 0 SBC Real-time - Real-Time HK generated by SBC (SC and Instruments), controlled by telemetry output (TO) filter tables - Can include CFDP package data units (PDUs) for table/memory dump operations from SBC - Real-time data is normal CCSDS packets/frames 1 SBC Stored 2 SSR Stored 3 SSR Stored - Files stored on SBC "Contingency Memory Buffer" ' Files can be either HK files, instrument science files (CRaTER/LEND), and tables/memory dump file - Files will use CFDP for downlink - Files stored in engineering/hk directory on the SSR - Contains mostly stored engineering data but can include table/memory dump files - Files stored in measurement directory on the SSR - Includes all measurement data generated by the instruments Space to Ground S-Band Contingency Communications Besides the nominal Ka and S band ground stations that will support the LRO mission, LRO will require backup and contingency support from the DSN. LRO will use the 34-meter DSN antennas for S-Band contacts only. LRO does not plan to use DSN for any Ka-band operations. The DSN ground stations can provide enough link to use some of the higher rate S-Band rates as well as the margin to receive telemetry and send commands when the spacecraft is in a tumble. It will be possible due to ground network outage or SC Ka-band failure to dump measurement data from the record using the S-Band system. This mode will not allow LRO to dump the nominal data volume, but it will allow a portion of the data to be dumped. In this mode, a modified instrument operations scenario is needed along with a priority scheme to management which data is collected and dumped. 6-23

90 Space to Ground Ka-Band Contingency Communications With the one dedicated Ka-Band stations, the station may have down periods due to routine maintenance or equipment failure. The ground station will be implemented with redundant front end processing equipment, but the antenna system will not have any redundancy. One benefit of building the dedicated antenna at White Sands is that the ground station design is similar to the antenna design that Solar Dynamics Observatory (SDO) mission is using. SDO will have two Ka-band antennas at White Sands. The LRO mission plans to have an operations agreement with the SDO mission over the possible use of one of the two SDO antennas for backup operations. In this mode, the SDO antenna would route the data stream to the LRO front-end equipment. This will reduce the complexity of implementing the backup scenario Ground Network Laser Ranging TBD Orbiter Commanding The MOC at GSFC supports all orbiter commanding activities. There are two basic types of commanding, R-T and load files. The operations team at the MOC will use the T&C system to generate and send all R-T commands to the orbiter. The mission planning system (MPS) will be used to generate load files that will be uplinked to the orbiter using the T&C system. The load files may contain the daily stored command load, FSW tables, or FSW patches. All R-T command will be perform using pre-built and tested command procedures. Below list some of the routine activities that involve orbiter commanding: Stored Command Load Uplink Extended Precision Vector (EPV) uplink Orbiter Clock Correlation and Adjustment Recorder Management and Data Transfers Stored Command Load Uplink The operations team will generate the daily load each day, Monday through Friday. The daily load will contain stored commands that control the orbiter for the following day. The command load is generated based on inputs from FD, operations team, and instrument SOCs. The load built on Friday will cover the weekend operations. The Friday load will contain 3 days worth of operations. Additional details on the command load uplink scenario are provided in Section Extended Precision Vector Uplink To reduce the orbit determination knowledge error, the FDF will process tracking data on a daily basis, 7 days a week (TBR). By processing tracking data on a daily basis, the knowledge error 6-24

91 can be reduced to approximately 3 to 4 seconds. FD will generate the new set of SC vector and transfer the file to the MOC. The MOC will detect the incoming file and select the appropriate vector for uplink. During the week, this operation will occur during nominal hours when the operations staffed the MOC. On weekends, this operation will be performed by the MOC automation SW Orbiter Clock Correlation and Maintain The SC clock correlation is performed by taking the difference between the time stamps of when a command left the ground station and when the command is received by the SC. The formula the operations team will use is listed below: ΔTime = GrndCmdTime sccmdtime xfertime errorfactor The GrndCmdTime will be supplied by the station. Currently, the White Sands and Australia ground stations will contain the necessary equipment to provide the command time stamping information. The sccmdtime is provided by the SC and is applied when the communication board receives the command. This information will be sent back to the ground through HK telemetry. The xfertime is the time it takes for the transmission of the command to the orbiter. The MOC will apply a correction factor based on range estimates provided by FD. The errorfactor will extract out any other known sources of error such as delays in SC time tagging or ground time tagging. The current plan is to perform the clock correlation once a time. This activity is automated for weekend operations. The MOC GS performs the clock correlation activity throughout the day. If the clock difference is off by more than 100 ms, a command is sent to adjust the clock to the correct Universal Time Code (UTC) time Operations Scenario The current concept is to staff the MOC Monday through Friday, with one 8-hour shift. During the off hours, the orbiter and ground system operations are monitored through the automation systems developed during pre-mission. The stored command load will control most of the orbiter activities throughout the day. The operations team receives inputs from various sources to generate the command file each day. Friday s command file will cover three days. Figure 6-10 shows a high-level concept for a typical day and Figure 6-11 shows a typical week for LRO. 6-25

92 Figure Typical Daily Operations Scenario Figure Weekly Operations Scenario For each S-Band contact, a particular sequence of activities is used (refer to Figure 6-12). It starts with the pre-pass activities and ends with the post-pass activities. During the 30 to 60 minute pass, the operations team will go through the same sequence. 6-26

93 LRO Typical Pass Scenario Pre-Pass Activities Pass Sequence Pass Activities Post-Pass Activities Pre-Pass Activities Acquisition Of Signal (AOS) Health & Safety Command & Control Operations Activities Loss of Signal (LOS) Post-Pass Activities Configure Ground System Prepare pass plans and procedures Perform initial connectively check with ground station Stored command sequence turns on S-Band transmitter and configures spacecraft for pass Ground station locks on signal and forwards real-time tlm to MOC MOC performs initial test command AOS Orbiter configuration checks are performed by the MOC Orbiter events are dumped and checked Real-time telemetry is checked vs. limits in database Load daily command load (if needed) Turn on Ka transmitter Initiate dump of measurement data Perform clock maintenance Perform file maintenance Perform table loads Real-time instrument commanding Special activities Monitor operations Stored command turns off S and Ka transmitter after predicted LOS Ground station sends data quality stats to the MOC LOS Ground station starts transfers of stored files to MOC MOC processes stored data for trending and health and safety checks Perform trending Distribute data products to SOCs Perform Level 0 Processing Data analysis Health & Safety Checks Command & Control Operational Activities 6.3 STATION-KEEPING Figure Typical Pass Scenario In order to maintain the mission orbit, LRO is required to perform monthly SK maneuvers. The maneuvers will cause an interruption to measurement collection for approximately one orbit. Since the maneuvers can be predicted in advance, instrument calibrations can be performed either before or after the SK maneuver. By coordinating the calibrations to SK maneuvers, it makes coordination between the measurement teams and operations team easier. The SK maneuvers are currently planned for every 349 orbits or approximately every 27.4 days. The maneuver requires a sequence of two burns. The first burn circularizes the orbit at either apoapsis or periapsis. The second burn corrects the orbit eccentricity. The two burns are performed in the velocity vector direction and will require 180 yaw maneuver to perform the burns. Below list the complete sequence: 1. Turn on the propulsion CATBED heaters at least 45 minutes prior to the maneuver 6-27

94 2. Lock the SA and HGA gimbals to desired position 3. The orbiter will slew to the designated attitude based on the command load, orbiter will perform maneuver in observing mode using the reaction wheels. Depended on the orbiter orientation, an initial 180 slew may not be needed. 4. After attitude is reached for the maneuver, the FSW will transition to Delta-V mode with the appropriate burn information based on the maneuver plan (1 st burn to circularize the orbit) 5. After burn is complete, the FSW will transition back to observing mode automatically 6. The command load will command the orbiter to perform a 180 yaw maneuver to reverse direction. 7. The command load will command the orbiter into Delta-V mode for the 2 nd burn (corrects the eccentricity). 8. After burn is complete, the FSW transitions back to observing mode automatically. 9. The command sequence will return the orbiter to the correct attitude using the wheels. 10. Propulsion system will be configured such as to turn-off the CATBED heaters Prior to the maneuver day, FD will develop and transfer the maneuver plan to the operations team. The operations team will use the mission planning system to create the maneuver command sequence. Both the operations team and FD will verify the maneuver command sequence. The operations team will ensure that the maneuver is coordinated with ground network S-Band coverage. Due to limitations on the ground network, the operations team may need to schedule DSN to view the maneuver if the HGA cannot be used during the maneuver. Prior to the SK maneuver, the instruments are commanded to the appropriate modes for thruster burns. Some instruments will need to be placed in a safe configuration to prevent contamination. The details of the instrument configuration will be defined later and coordinated between the operations team and each of the instrument SOCs. After the maneuver is complete, FD will process the tracking data and release a maneuver burn performance report. The operations team will calculate the fuel used and fuel remaining based on thruster firing counts and pressure/temperature information Orbiter Orientation Description Station-keeping maneuvers require two burns that require the orbiter to slew 180 between maneuvers. The slews takes less than 10 minutes to perform. 6-28

95 6.4 MOMENTUM MANAGEMENT Twice a month, the operations team will perform momentum management. The activity will be scheduled when the periods when the Earth ground stations have full coverage of the orbit. FD will generate the delta-h plan and transfer to the MOC. The operations team will ingest and generate the command load sequence and ensure the maneuver has appropriate ground tracking for the maneuver. The sequence of activities for the maneuver is provided below: 1. Turn on the propulsion CATBED heaters at least 45 minutes prior to the maneuver 2. Lock the SA and HGA gimbals to desired position (TBR) 3. The FSW will transition to Delta-H mode with the appropriate burn information based on the maneuver plan (1 st burn to circularize the orbit) 4. After burn is complete, the FSW will transition back to observing mode automatically Orbiter Orientation Description Momentum management should not require a change in SC attitude. The orbiter should remain pointing nadir throughout the activity. 6.5 INSTRUMENT CALIBRATIONS Instrument calibrations are performed in conjunction of the monthly SK maneuvers. Since the SK maneuvers events can be predicted in advance, this allows easier coordination with the instrument teams. A integrated and coordinated calibration plan is generated between each instrument, operations and FD. Any instrument calibration will still need to conform to any flight rules and constraints. Currently calibrations have been allocated three orbits, but the duration is up to the measurement team. Additional calibration time can be performed but it will increase interruption to nominal measurement data collection. Most likely the instrument calibration activities will take place before the SK maneuver since some instruments will need to placed in safe modes. The calibration activities will occur during the operational hours. It is expected that the SWG will develop a coordinated calibration plan and work with the operations team to define the sequence. Since calibrations will most likely require various slews, coordination between the instrument teams, operations team, and FD is needed. Table 6-13 shows the current list of potential calibration activities for each of the instruments. 6-29

96 Table Instrument Calibration Activities (TBR) Instrument Cal. Description Status LROC Image stars and deep space. Useful for WAC absolute calibration Orbiter Orientation Description TBD Diagrams and description of the orbiter orientation during the phase. 6.6 LUNAR ECLIPSE SUB-PHASE Twice a year, on average, the Earth will pass between the Moon and the Sun causing partial/full lunar eclipse. Based on predicts (refer to Table 6-14) the first significant lunar eclipse occurs on December 31, This event may or may not occur during the prime measurement phase. For any significant lunar eclipse, the SC will need to be placed in a low power mode. Table Lunar Eclipse Predicts during the LRO Mission Date Eclipse Type Eclipse Duration February 9, 2009 Penumbral - July 7, 2009 Penumbral - August 6, 2009 Penumbral - December 31, 2009 Partial 62 minutes June 26, 2010 Partial 164 minutes December 21, minutes (Total Total Eclipse) 209 minutes (Total + Partial) June 15, 2011 December 10, 2011 Total Total 101 minutes 220 minutes (Total + Partial) 52 minutes 213 minutes (Total + Partial) June 4, 2012 Partial 128 minutes November 28, 2012 Penumbral

97 Date Eclipse Type Eclipse Duration April 25, 2013 Partial 32 minutes May 25, 2013 Penumbral - October 18, 2013 Penumbral - Penumbral Eclipse Partial Eclipse Total Eclipse The Moon passes through the Earth s penumbral shadow. These events are subtle and quire difficult to observe. A portion of the Moon passes through the Earth s umbral shadow. The entire Moon passes through the Earth s umbral shadow. Two times are provide, first is the time for the total eclipse period and the second is the start to finish time that includes partial and total eclipse times. The orbiter low power mode will be determined by the length of eclipse, but it most likely will include turning instrument off and turning certain SC components off to conserve energy. Table 6-15 shows the worst-case low power mode. One note, the configuration in the following table is just one example, each lunar eclipse will be analyzed and the appropriate SC configuration will be determined. Table Orbiter Configuration during Lunar Eclipse LRO Orbiter Configuration for Lunar Eclipse (TBR) Component/System Category Configuration Component/System Category Configuration C&DH Box C&DH ON LROC Inst. OFF Solid State Recorder C&DH OFF LOLA Inst. OFF C&DH SW Mode C&DH Bank 1 LEND Inst. OFF C&DH SW Config C&DH Nominal Config CRaTER Inst. OFF R-T Telemetry C&DH 16 or 32 kbps LAMP Inst. OFF Cmd Rate C&DH N/A Diviner Inst. OFF Comm. Cards C&DH ON Battery Power Online (Discharge) S-Band Receivers Comm. ON PSE Power ON S/Ka Band Xmitter Comm. OFF Servo Drive Bus Power OFF 6-31

98 LRO Orbiter Configuration for Lunar Eclipse (TBR) Component/System Category Configuration Component/System Category Configuration Transponder Config. Comm. RF Switches configured for Omni Deployment Bus Power Reaction Wheels GNC ON Isolation Valve Prop. Open Star Trackers GNC OFF Prop. Heaters Prop. Enabled IMU/Gyro GNC ON CATBED Heaters Prop. OFF ACS SW Mode GNC Observing (TBR) Prop. Thrusters Survival Heater Bus Thermal Enabled SC Ops Heaters Thermal OFF Inst. Survival HTRs Thermal Enabled Inst. Ops Heaters Thermal OFF Deployment HTRs Thermal Enabled Mini-RF Demo OFF Prop. OFF OFF Orbiter Orientation Description TBD Diagrams and description of the orbiter orientation during the phase. 6.7 YAW MANEUVER SUB-PHASE Twice a year, the orbiter needs to perform a 180 yaw maneuver to maintain the Sun on the correct side of the SC. The reason for the maneuver is driven by the gimbal range of motion and thermal design of the orbit. When the yaw maneuver is performed, the orbiter will reverse the orientation in the velocity direction. Based on power and thermal constraints, the maneuver will be performed within ± 5 (TBR) of the orbit beta angle of 0. During the maneuver, the orbiter will not sweep the instruments field of views (FOVs) through the Sun. The maneuver should be completed within 1 orbit. Since the SC is maneuvering, interruption to some of the instrument measurement collection will occur. The maneuver will be done using the reaction wheels. The operations team will schedule the maneuver during ground coverage. Since the maneuver may take the HGA off the Earth ground station, the operation may need to use the omni antennas Orbiter Orientation Description The yaw maneuver will reverse the X-axis orientation. At the completion of the maneuver, the orbiter will return to nadir pointing. During the maneuver, the Sun will be kept off the anti-sun. 6-32

99 6.8 SAFING / SAFE-MODE SUB-PHASE The LRO fault protection system includes several types and levels of safing including loadshedding and safe-mode which are dependent on entrance conditions. If LRO enters the safemode, it could cause days worth of interruption to measurement collection since the instruments will be powered off. The actual duration of interruption will be dependent on the type of anomaly Power Induced Load Shedding This mode is initiated by several possible power conditions, which may include: Low bus voltage Low battery state of charge High current draw The end result of the load shedding is the SC will be put into a low power mode where nonessential loads will be powered off. Typically, this is a tiered approach where the first action results in minimum reconfiguration of the orbiter and the final action results in the orbiter reconfiguration into safe low power mode. The orbiter will be held on the Sun-line to keep the arrays on the Sun and provide a power-positive orientation. If required, a notice can be sent to the instruments prior to the removal of power Safe-Mode Entry into this mode may be caused by the processor detecting a problem (such as an attitude excursion). The orbiter may go into a Sun-pointing mode. Additionally, if communications with the ground is lost for a specified time, the SC will enter safe-mode. Safe-mode is a low power mode that will ensure the Sun is illuminating the arrays so the SC is in a power-positive orientation Orbiter Orientation Description TBD Diagrams and description of the orbiter orientation during the phase. 6-33

100 7.0 EXTENDED MISSION OPERATIONS PHASE The prime mission phase of LRO is one year of nominal measurement collection. After the one year, the mission enters the extended mission operations phase. The goals for extended phase have not been defined yet, but some initial goals may be: Support future RLEP missions Perform extended measurement collections One option to expand the extended mission operations phase to the maximum duration is to change the orbit to a frozen elliptical orbit. The frozen orbit is believed to be at 3-x216 km. The main driver for the frozen orbit is to reduce fuel usage to extend the time LRO can remain in the lunar orbit. Another option is to continue to collect measurement data from all or limited number of instruments based on future RLEP objectives. Regardless of the option, LRO does have limited resources to maintain the lunar orbit and eventually planning will occur to dispose the mission. 7-1

101 8.0 END-OF-MISSION DISPOSAL The current plan is to fly LRO until the fuel is exhausted. Over time, the orbit will decay and LRO will crash into the surface of the Moon. At this time, there are no plans to have a controlled crash on the surface. Guidelines outlined in the Planetary Protection Provisions for Robotic Extraterrestrial Missions NASA Procedural Requirements (NPR ) provide the NASA guidelines for mission disposal. 8-1

102 9.0 GROUND SYSTEM DETAILED DESCRIPTION As described in the previous sections, the LRO GS is comprised of five main elements (see Figure 2-7 above): The LRO Ground Network which consists of a dedicated S/Ka Band ground station at White Sands and various S-Band only ground stations located throughout the world. Mission Operations Center (MOC) Flight Dynamics (FD) Science Operations Center (SOC) for each instrument Communications network which provides voice and data connectivity between each of these elements 9.1 NETWORK, FACILITIES, AND PERSONNEL SECURITY Network, facilities and personnel security requirements are important considerations in the operations environment. Further details on network, facility and personnel security will be included in future documentation such as the Lunar Reconnaissance Orbiter Information Technology Security Plan (431-PLAN ) to address the Security of Information Technology NASA Procedural Requirements (NPR ) and the Security Requirements for Unclassified Automated Information Resources (NASA FAR ). 9.2 LRO GROUND NETWORK The LRO ground network includes a dedicated ground station at White Sands Complex, existing S-Band ground stations, and the DSN. The three different components of the ground network will support all phases and modes of the LRO mission White Sands Complex Ground Station The LRO WSC Ground Station includes an 18-meter dual Ka/S Band antenna system and frontend equipment for data processing. The ground station can handle a Ka-Band rate of 100 Mbps and S-Band telemetry rate of 300 kbps. The ground station will also support the 4 kbps command uplink rate. The LRO ground station will be built on the grounds of the NASA WSC located in White Sands, New Mexico (NM). This location was chosen for several reasons. Since Ka-Band RF is highly susceptible to rain interference, a site with minimal annual precipitation was required. Also, the WSC has much of the required infrastructure already in place to support the new LRO ground station. Figure 9-1 shows a block diagram for the White Sands Ground Station. 9-1

103 Figure 9-1. LRO WSC Ground Station Block Diagram The antenna is an 18-meter parabolic dish that can simultaneously support Ka-Band telemetry and S-Band T&C uplink (refer to Figure 9-2). The antenna system will be similar to the ground stations being built for SDO. By having a similar antenna systems, costs to operate the antennas are lower since personnel and spare parts can be shared across three antenna systems. 9-2

104 Figure 9-2. WSC Ground Antenna System White Sands Complex Ka-Band Component The Ka-Band components of the ground stations receive SC Ka-Band RF downlink at a frequency of GHz. The RF signal is down-converted to an Intermediate Frequency (IF) that can be input to a demodulator and Bit-Sync receiver within the front-end processor. The Ka-band antenna system will nominally auto-track the SC RF signal to keep it within the narrow Ka-Band beamwidth of the 18-meter dish. However, for initial acquisition of signal, the auto-track system first finds and tracks the SC based on the signal strength of the wider S-Band beamwidth. Once locked, it will transition over to the closed loop auto-tracking of the Ka-band signal. The antenna will also have the capability to program-track the SC signal by following pre-described Sc position data calculated from the acquisition data provided by FD Data Distribution System The two main components of the Data Distribution System (DDS) are comprised of the high rate front-end processor (FEP) and the CFDP processor. The FEP receives the down-converted IF signal from the Ka-Band system and performs viterbi and RS corrections and sorts the data by VC. There is redundant FEP for the antenna. The CFDP processor receives the Virtual Channel Data Unit (VCDU) from the FEP. The FEP will be configured to pass only certain VCs to the CFDP processor. The CFDP processor performs the CFDP processing on data files and generates status packets to be sent back to the SC indicating whether the file contains missing pieces or all file pieces have been received on the 9-3

105 ground. The CFDP processor sends the status packets to the MOC. The MOC T&C system receives the packets and forwards them on the command line for uplink to the SC. The CFDP processor also contains the storage device that will store the data files at the ground station for the required amount of time. Figure 9-3 shows the end-to-end flow for the Ka-Band CFDP process. Figure 9-3. CFDP Ka-Band Flow for LRO During early mission operations and contingencies, the Ka-Band FEP will be configured to forward certain VC data back to the MOC directly. In this mode, the CFDP processor back at the MOC will perform the file processing. Only a limited portion of the data stream can be sent back to the MOC directly since the ground network will not be able to support the 100 Mbps. The current plan is to deploy two CFDP processor at the Ka-Band ground site for redundancy. Since the CFDP processor is mission unique, the MOC will control and configure the CFDP 9-4

106 processors over the network. Figure 9-4 shows how the MOC systems will control and receive status information from the CFDP processors located at WSC. Data Storage Encrypted CFDP Naks, Ack EOF, Finish Cmds (TCP/IP) File Data CFDP Status, EOF, Ack Finish (TCP/IP) CFDP Processor (ITOS) CFDP Processor Status CFDP control & reponses VCDUs w/cfdp PDUs, EOF, Ack-Finish Optional CFDP PDUs, Engineering Data EOF, Ack-Finish Ka-Band FEP S-Band FEP VCDUs w/cfdp PDUs, EOF, Ack-Finish Data Storage?? Encrypted Commands Ground Network Command Encrypter (routed back through T&C) CFDP Naks, Ack EOF, Finish Cmds T&C System CFDP Status, EOF, Ack Finish CFDP Processor Status CFDP Processor ITOS CFDP control & reponses File Data Files w/ CCSDS Pkts/Raw Data Delivered post support File Data Mission Operations Center Data Storage Figure 9-4. WSC CFDP Processor Control One additional note, LRO will also be capable of supporting the unreliable transfer method of CFDP. In certain contingencies, the operations team can configure the SC and ground for this mode. In this mode, the SC will dump the files by the ground will not command missing packets Ka-Band Dump Concept The White Sands ground station provides a minimum of four contacts each day to dump the recorder. The two most critical contacts are the first and last support since these two contacts define the total time that the SC recorder needs to store data. Referring to Table 9-1, all the expected data on the first contact will not be dumped during the first 45 minutes. It actually 9-5

107 takes two supports to dump all the data stored before the last and first contact, ~1044 minutes. The remaining contacts provide adequate time for CFDP retransmits. Table 9-1. WSC Ka-Band Dump Concept White Sands Complex S-Band Component The purpose of the S-Band equipment is threefold: 1. Initially acquire the SC 2. Provide the means to transmit commands and range codes to the SC 3. Provide the means to downlink the R-T HK data and range tones. The S-Band HK is then forwarded to the MOC for processing White Sands Complex S-Band Transmit and Receive The S-Band receiver (RX) and transmitter (TX) are housed in the same facility as the Ka-Band system and share the same 18-meter dish. The nominal R-T S-Band downlink rate from the orbiter is either 16 or 32 kbps. While other modes are possible, they will be used during certain operations or contingency activities. The S-Band receiver operates on a frequency of MHz. It demodulates and separates the signals from the downlink, which carry health and safety data and ranging data. The S-Band transmitter operates on a frequency of MHz. It provides the means to uplink data to the SC. Commands received from the MOC are impressed upon a 16 KHz frequency sub-carrier and then modulated on the S-Band for uplink to the orbiter. The command uplink rate for LRO is 4 kbps White Sands Complex S-Band Housekeeping Data Distribution S-Band telemetry contains the orbiter housekeeping or engineering data and is needed for health and safety monitoring. The S-Band HK telemetry is sent in R-T from the LRO Ground Station at WSC to the MOC at GSFC. 9-6