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431-SPEC-000008Revision CD

Effective Date: 7/20/2005Expiration Date: 7/20/2005

CHECK WITH LRO DATABASE AT:

https://lunarngin.gsfc.nasa.govTO VERIFY THAT THIS IS THE CORRECT VERSION PRIOR TO USE.

DRAFT

Robotic Lunar Exploration Program

Lunar Reconnaissance Orbiter Project

LRO Electrical Systems Specification

24 April01 May 2007

Goddard Space Flight CenterGreenbelt, Maryland

National Aeronautics andSpace Administration

Goddard Space Flight CenterGreenbelt, Maryland

National Aeronautics andSpace Administration

LRO Electrical Systems Specification 431-SPEC-000008Revision CD

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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 applicableConfiguration Control Board (CCB) Chairperson or designee. Proposed changes shall besubmitted to the LRO CM Office (CMO), along with supportive material justifying the proposedchange. Changes to this document will be made by complete revision.

Questions or comments concerning this document should be addressed to:

LRO Configuration Management OfficeMail Stop 431Goddard Space Flight CenterGreenbelt, Maryland 20771


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Signature Page

Prepared by:

Original signed by 7/20/2005Philip Luers DateRLEP Avionics SystemsGSFC/NASA, Code 561

Reviewed by:

Original signed by 7/20/2005Mike Pryzby DateLRO Spacecraft Systems EngineerSwales Aerospace/Code 870

Approved by:

Original signed by 7/20/2005Craig Tooley DateLRO Project ManagerGSFC/NASA, Code 431


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LUNAR RECONNAISSANCE ORBITER PROJECT

DOCUMENT CHANGE RECORD Sheet: 1 of 1

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DESCRIPTION OF CHANGEAPPROVED

BYDATE

APPROVED

Rev -

Rev A

Rev B

Rev C

Released per 431-CCR-000005




C. Tooley

C. Tooley

C. Tooley

C. Tooley

7/20/2005

6/22/2006

2/01/2007

4/10/2007


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List of TBDs/TBRs

ItemNo.

Location Summary Ind./Org. Due Date


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TABLE OF CONTENTS

Page1.0 Introduction..................................................................................................................1-1

1.1 Purpose ..............................................................................................................1-11.2 Definitions..........................................................................................................1-11.3 Electrical System Overview ..................................................................................1-1

1.3.1 Electrical System .....................................................................................1-21.3.2 Electrical System Drivers for LRO .............................................................1-2

2.0 Documentation ..............................................................................................................2-12.1 Applicable Documents .........................................................................................2-1

2.1.1 NASA Documents....................................................................................2-12.1.2 Non-NASA Documents ............................................................................2-1

2.2 Reference Documents ..........................................................................................2-12.2.1 NASA Documents....................................................................................2-12.2.2 Non-NASA Documents ............................................................................2-1

3.0 Electrical System Requirements .....................................................................................3-13.1 Power ................................................................................................................3-1

3.1.1 Power Distribution and Switching Scheme...................................................3-13.1.2 Power System Electronics Specifications.....................................................3-13.1.3 User (Subsystem) Specifications ................................................................3-3

3.2 System Grounding Requirements ...........................................................................3-83.2.1 Single-Point Primary Power Ground ...........................................................3-83.2.2 Distributed Signal Ground.........................................................................3-93.2.3 Common Mode Noise...............................................................................3-93.2.4 Bonding or Mating ...................................................................................3-93.2.5 Shield Grounds ......................................................................................3-103.2.6 Grounding of External Orbiter Surfaces.....................................................3-10

3.3 EMI/EMC Requirements ....................................................................................3-123.3.1 Conducted Emissions..............................................................................3-133.3.2 Radiated Emissions (RE02) .....................................................................3-153.3.3 Conducted Susceptibility.........................................................................3-163.3.4 Radiated Susceptibility (RS03).................................................................3-193.3.5 Orbiter RF Self-Compatibility..................................................................3-23

3.4 Data and Signal Interfaces...................................................................................3-233.4.1 Inter-Component Communications ...........................................................3-233.4.2 Pyrotechnic and Deployable Actuator Interfaces .........................................3-293.4.3 External Interfaces .................................................................................3-30

3.5 Mulipaction and Corona .....................................................................................3-313.6 Design For Radiation .........................................................................................3-313.7 Charging and Discharging Requirements...............................................................3-32

3.7.1 External Surface Charging.......................................................................3-32


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3.7.2 Surface Discharging Protection ................................................................3-333.7.3 Internal Charging ...................................................................................3-35

4.0 Harness Requirements ...................................................................................................4-14.1 General Harness Guidelines ..................................................................................4-1

4.1.1 Accessibility............................................................................................4-24.1.2 Harness Shields .......................................................................................4-24.1.3 Component Test Connector Panels .............................................................4-3

4.2 Electrical Materials ..............................................................................................4-44.2.1 Connectors..............................................................................................4-4

5.0 Qualification Assurance provisions .................................................................................5-15.1 General ..............................................................................................................5-1

5.1.1 Analysis .................................................................................................5-15.1.2 Demonstration .........................................................................................5-25.1.3 Inspection ...............................................................................................5-25.1.4 Test........................................................................................................5-2

5.2 Verification Matrix Table .....................................................................................5-36.0 Abbreviations and Acronyms ........................................................................................A-17.0 Appendix B. Traceability Matrix ................................................................................... B-1


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LIST OF FIGURES

Figure Page

Figure 3-1. SSPC In-rush and Trip Current Limits Curve, 1 Amp Service ..........................................3-6

Figure 3-2. SSPC In-rush and Trip Current Limits Curve, 2 Amp Service ..........................................3-6

Figure 3-3. SSPC In-rush and Trip Current Limits Curve, 5, 10, and 15 Amp Services.........................3-7

Figure 3-4. LRO Spacecraft Shielding Scheme.............................................................................3-11

Figure 3-5. LRO Spacecraft Power Isolation Scheme ....................................................................3-11

Figure 3-6. Narrowband Conducted Emissions CE01/CE03 Limits .................................................3-14

Figure 3-7. RE02 Limits for the Orbiter and Components that are ON from launch to vehicle separation(vertical axis is dBmicrovolts/meter) ...........................................................................................3-17

Figure 3-8. RE02 Limits for Components that are OFF from launch to vehicle separation (vertical axis isdBmicrovolts/meter) .................................................................................................................3-18

Figure 3-9. CS01/CS02 Limits ...................................................................................................3-18

Figure 3-10. CS06 Conducted Susceptibility Test Pulse.................................................................3-21

Figure 3-11. 1553 Data Bus Subsystem Diagram ..........................................................................3-24

Figure 3-12. 1553 Data Bus System Diagram...............................................................................3-25

Figure 3-13. Discrete Command Termination...............................................................................3-28

Figure 3-14. 1 pps Timing Characteristics....................................................................................3-29

Figure 3-15. External Harnesses and Charge Mitigation.................................................................3-35


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LIST OF TABLES

Table Page

Table 3-1. Switched Services Currents and Deratings ......................................................................3-2

Table 3-2. LRO Primary Power Voltage Range ..............................................................................3-4

Table 3-3. EMI/EMC Applicability and References ......................................................................3-12

Table 3-4. Atlas V Spacecraft RE Limits .....................................................................................3-15

Table 3-5. LRO Operational RS Test Limits ................................................................................3-20

Table 3-6. Launch Site/Vehicle RS Test Levels ............................................................................3-21

Table 3-7. LRO 1553 Remote Terminal Addresses .......................................................................3-25

Table 3-8. 1 pps Timing Characteristics ......................................................................................3-29

Table 4-1. Minimum and Maximum AWG ....................................................................................4-1


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1.0 INTRODUCTION

1.1 PURPOSE

This document provides the electrical and electronic requirements and some guidelines to thesubsystem designers for the Lunar Reconnaissance Orbiter (LRO) mission. The purpose of theserequirements and guidelines is to assure reliable and compatible operation of the elements thatmake up the electrical system, both during ground testing and on orbit, with margin for bothexpected and worst-case environmental conditions.Specific details of each subsystem interface will be documented in subsystem specifications,component specifications, and interface control documents (ICDs).

1.2 DEFINITIONS

Component: A functional subdivision of a subsystem and generally a self-containedcombination of items performing a function necessary for the subsystem's operation. Examplesare electronic box, transmitter, gyro package, actuator, motor, battery.Instrument: A spacecraft (SC) subsystem consisting of sensors and associated hardware formaking measurements or observations in space. For the purposes of this document, aninstrument is considered a subsystem (of the SC).Orbiter: An integrated assemblage of modules, subsystems, etc., designed to perform aspecified mission in space. For the purposes of this document, “Orbiter” and "SC" are usedinterchangeably. Other terms used to designate this level of assembly are Laboratory,Observatory, and satellite.Subsystem: A functional subdivision of a SC consisting of one or more components. Examplesare structural, thermal, attitude control, electrical power, command and data handling (C&DH),communication, science instruments or experiments.Segment: A functional subdivision of the overall system consisting of either the entire Orbiteror the entire Ground Segment.System: An integrated end-to-end assembly of all mission components, including both theGround Segment and the Orbiter Segment.

1.3 ELECTRICAL SYSTEM OVERVIEW

The electrical system includes electronics and electrical components, interconnect harnessing,structural chassis grounding system, grounding of external coatings, thermal blankets andelements that provide shielding.The environments that apply to the electrical system include self-generated, conducted, andradiated electromagnetic noise; ground-based electromagnetic emitters, the thermal andmechanical environments of integration and test (I&T), the launch environment, and the on-orbitenvironment.


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1.3.1 Electrical System

The LRO electrical system includes the electrical elements mounted on the Orbiter that areinterconnected to perform their defined functions to meet mission requirements. To the extentthat there are challenges in interconnecting the electrical system, whether built in-house orprocured from an external vendor, this specification is designed to define those design aspectsthat are critical to the integrated functioning of the system. This ICD defines the LRO generaland specific electrical requirements. The LRO subsystems should implement these requirementsduring their design process in order to assure proper system operation and will verify that therequirements are met.

1.3.2 Electrical System Drivers for LRO

The electrical system must be designed carefully to address electromagnetic interference (EMI),the orbital charging and radiation environments, as well as the functional requirements ofcollecting instrument data and transmitting it to Earth.

On-Orbit Charging Environment: While electrostatic discharge (ESD) threats may not betotally eliminated, they can be minimized and their effects mitigated through the use of sounddesign practices. An overview of the planned LRO integrated approach (Section 3.7,Charging and Discharging Requirements) includes limiting the number of discharge sources,limiting the size of discharges, implementing shielding between potential sources andpotential victims of discharges, and controlling victim susceptibility via filtering andbandwidth control.EMI: The key element of EMI control is the design and use of the SC structure as a Faradaycage. The Faraday cage concept provides shielding between the noisy outside environmentand the electronics and harnessing internal to the SC. Harnesses that transition through theFaraday cage will be grounded, shielded, and/or filtered to maintain the overall shieldintegrity. Noise sources external to the SC are expected to include unavoidable ESDs,ground-based Radio Frequency (RF) emitters, and self-generated RF from the LRO Ka- andS-band RF systems.Instrument Suite: The LRO instrument suite, as well as Star Trackers, will containinstruments with Charge-Coupled Device (CCD) detectors. These detectors are sensitive tocommon mode noise and can pick up ground noise.To minimize the total common mode noise environment, noise sources will be controlled atthe potential sources by limiting alternating current (AC) noise (Section 3.2, SystemGrounding Requirements). Coupling mechanisms between the potential sources and theCCD victims will be controlled by providing a low AC impedance to chassis ground.High Data Rate: The LRO downlink data rate of 125100 megabits per second (Mbps)requires relatively high-frequency clocks with corresponding fast rise and fall times that mayrepresent a significant noise source. High data quality and integrity requirements, along withshort bit times for the telemetry data, represent challenges to maintaining error-free data.These high-frequency clocks are basically RF signals and need to be treated as such when itcomes to both the grounding approach and interfacing via impedance matched transmissionlines.


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Total Dose Radiation and Single-Event Effects (SEEs): The LRO orbit represents amoderate radiation environment. The planned approach for total dose shielding will beshared between the structure and the component chassis to limit the radiation at the part level.SEEs will be controlled through the use of radiation-hard or radiation-tolerant parts andcircuit designs that can tolerate Single-Event Upsets (SEUs). Potentially destructive damagewill be controlled through the use of radiation-hardened parts, while upsets or soft failureswill be controlled through radiation-tolerant parts, circuit design, software design or othermitigation methods. Refer to the Radiation Requirements for the Lunar ReconnaissanceOrbiter (431-RQMT-000045) for additional details.Compatible and robust interfaces between electrical system components are key to meetingrequirements given the potential noise sources described in the above paragraphs a-d.Special attention should be given to the design and control interfaces. Most low data ratesignal interfaces are expected to use the 1553 data bus, which is inherently robust. Non-1553interfaces will be carefully controlled and reviewed for interface robustness andcompatibility, as well as evaluating the potential of being a noise source or noise susceptiblevictim. In order to minimize the potential for noise problems, each interface will be expectedto control its signal bandwidth only to that which is necessary to perform its function, subjectto review on a case-by-case basis by the LRO Project. This is expected to encompassrise/fall time control of edges on transmission signals, as well as filtering at the receivingend.RF Environment: The LRO RF environment will include self-generated S-band and Ka-bandradiation emitted by the High-Gain Antenna (HGA). At certain pointing angles, thispotential EMI source may illuminate the instruments and the solar array (SA). Other self-generated RF sources are expected to include the S-band downlink via the omni antennas.Ground-based RF sources are expected to include the launch pad transmitters and ascentsources, including the ground radars and the launch vehicle. The flight environment is alsoexpected to include ground radars and uplink sources to other SC in the vicinity of LRO.


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2.0 DOCUMENTATION

2.1 APPLICABLE DOCUMENTS

2.1.1 NASA Documents

GSFC-STD-7000 General Environmental Verification Standard for GSFC Flight Programsand Projects (GEVS)

2.1.2 Non-NASA Documents

MIL-STD-461C EMI/EMC Requirements Electromagnetic Emission and SusceptibilityRequirements for the Control of Electromagnetic InterferenceMIL-STD-462EMI/EMC Testing Methods, Notice 6ECSS-E-50-12 ESA SpaceWire SpecificationTIA/EIA-422 Electrical Characteristics of Balanced Voltage Digital Interface Circuits (formerlyknown as RS-422)TIA/EIA-644 Electrical Characteristics of Low Voltage Differential Signaling (LVDS) InterfaceCircuits

2.2 REFERENCE DOCUMENTS

2.2.1 NASA Documents

431-SPEC-000222 Lunar Reconnaissance Orbiter Project Power Distribution DiagramSpecification431-SPEC-000103 LRO SpaceWire Specification431-SPEC-000091 Thermal Systems SpecificationEEE-INST-002 Instructions for EEE Parts Selection, Screening, Qualification, andDerating431-SPEC-000013 Lunar Reconnaissance Orbiter Project Power Subsystem ElectronicsSpecification565-PG-8700.2.1 Design and Development Guidelines for Spaceflight Electrical HarnessesNASA-HDBK-4002 Avoiding Problems Caused by Spacecraft On-Orbit Internal ChargingEffectsNASA-HDBK-4001 Electrical Grounding Architecture for Unmanned Spacecraft431-REF-000273 Single Event Effect Criticality AnalysisTP2361 Design Guidelines for Assessing and Controlling Spacecraft Charging Effects431-SPEC-000TBD LRO Thermal Blanket Specification

2.2.2 Non-NASA Documents

AFSCM 91-710 Range Safety User Requirements ManualCLSB-0409-1109 Atlas Launch System Mission Planner's GuideMIL-HDBK-1553 Multiplexed Data Bus Handbook


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MIL-STD-1553B Multiplexed Data BusMIL-STD-1576 Electroexplosive Subsystem Safety Requirements and Test Methods forSpace Systems


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3.0 ELECTRICAL SYSTEM REQUIREMENTS

In this document, a requirement is identified by “shall,” a good practice by “should”, permissionby “may”, or “can”, expectation by “will”, and descriptive material by “is.”

3.1 POWER

ESS-1 The power subsystem shall supply switched and unswitched unregulated powerservices with the specifications below.

3.1.1 Power Distribution and Switching Scheme

ESS-2 Distributed power architecture shall be used to supply over-current protectedpower to all the loads. The power subsystem will provide +31 volts (V)unregulated power to each subsystem.

ESS-3 In the spacecraft nominal configuration, no +31V power returns shall be switched.Power returns may be switched to provide safety inhibit controls prior toseparation from the launch vehicle.

Unswitched power will be supplied to only those critical functions necessary to receivecommands and manage redundancy from the ground. Except for the previously mentionedcritical functions, all other functions will be commandable to the off state in order to turn off afailed load or to conserve power in an emergency. The Lunar Robotic Orbiter PowerDistribution Diagram Specification (431-SPEC-000222) shows a simplified LRO powerdistribution diagram with default power bus switch states.All power services will be over-current protected (based on service capacity) using solid-statepower controllers (SSPCs). All SSPCs will act as circuit breakers, tripping off at preset currentlevels, and they are designed to be re-settable by command. See Section 3.1.3.5 b) for derating.

ESS-4 Critical unswitched power services shall use fuses as over-current protectiondevices during integration and test only. Non-critical or redundant unswitchedservices may be fused in flight.

The current sensing shunts for telemetry are located in the return portion of the switched andunswitched services. Current must be returned on the originating service.

3.1.2 Power System Electronics Specifications

For internal Power System Requirements, refer to the Lunar Reconnaissance Orbiter ProjectPower Subsystem Electronics Specification (431-SPEC-000013) and the Lunar ReconnaissanceOrbiter Project Power Distribution Diagram Specification (431-SPEC-000222).


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3.1.2.1 LRO Power Redundancy

ESS-5 The power system shall provide the capability for redundant wires (more thanrequired for current carrying for each service).

3.1.2.2 Power System Electronics Output Switching Profile

ESS-6 When a service is switched on, the voltage shall rise from 0 to the steady-statevoltage no faster than 50 microseconds (µs) to reduce the in-rush at the usercircuitry.

ESS-7 When a service is switched on, the voltage shall rise from 0 to the steady-statevoltage no slower than 3 milliseconds (ms) to allow for proper operation ofpower-on reset circuitry.

ESS-8 When a service is switched off or trips off due to a fault condition, the voltageshall fall to 0 V no faster than 20 µs, prohibiting a sharp turn-off from producingan induced EMI emission.

3.1.2.3 Power Steady State Power

ESS-9 The bus voltage at the Power System Electronics (PSE) Output Module shall havea nominal +31 volts direct current (VDC) output voltage with a range from +21VDC to +35 VDC (inclusive) at the component end of the electrical harness.

ESS-10 The peak current shall not exceed 80% of the SSPC maximum sustainablecurrent, shown below in Table 3-1.

Table 3-1. Switched Services Currents and Deratings

Switched Service Type (A) SSPC Maximum SustainableCurrent (A)

80% Derating (A)

1 1.2 0.962 2.4 1.95 6 4.8

10 12 9.615 18 14.4

3.1.2.4 Power Bus Ripple

ESS-11 The bus ripple contributed by PSE shall be less than 0.3V peak-to-peak (p-p).ESS-12 Nominal orbiter level power bus ripple resulting from contributions from all

nominal sources shall be less than 1.0V p-p over the frequency range of 1.0 hertz(Hz) to 10 megahertz (MHz), and 0.5V p-p over 10 MHz at the power systemoutputs, under any load condition.


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3.1.2.5 Single-Event Power Bus Transients

ESS-13 Single-event power bus transients superimposed on the power buses due tonormal subsystem load switching shall be limited to +/-3.0V from the steady-statebus value. An example of this would be a subsystem that controls its own loads,turning them on and off.

The bus will recover to within 10% of its steady state value in 10 ms for a positive or negativeload step of 10 amps with a maximum current rate change of 50 milliamps per microsecond(mA/µs). The bus will recover to within 10% of its steady state value in 50 ms for a positive ornegative load step of 15 amps with a maximum current rate change of 300 mA/µs.

3.1.2.6 Over-Current Protection Deratings

ESS-14 SSPC devices used for LRO shall be selected based on the load current for eachload, and derated per the solid-state power switch device I_T trip curvecharacteristics, which can be found in the parts specification.

3.1.2.7 Bus Impedance

The impedance of the LRO power bus is a combination of the contributing impedance effects ofthe power source, distribution harness, switching devices, and connectors. For the purposes ofmodeling, the power subsystem impedance is approximated as an 80-milliohm resistor in serieswith 2 micro Henrys of inductance on each power and return line. The impedance of the powerdistribution harness must be added to this model to approximate the impedance of the power busas seen at any given component power input. This impedance is component specific but may beapproximated for test by 3 meters of wire, American Wire Gauge (AWG) 20, twisted, non-shielded.

3.1.2.8 Insulation

ESS-15 Exposed battery terminals shall be conformally coated to reduce likelihood ofaccidental short circuits to the SC structure.

3.1.3 User (Subsystem) Specifications

The specifications in this section apply to all subsystems whether built in-house or procured froman external source.

3.1.3.1 Steady-State Voltage

ESS-16 SC subsystems shall operate in the presence of a primary power input with thevoltage range as given in Table 3-2 as measured at their primary power inputs.The nominal power input at the subsystems will be +31 VDC.


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Table 3-2. LRO Primary Power Voltage Range

Subsystem Voltage Range (VDC) CommentCRaTER 27 - 35DLRE 27 - 35LAMP 27 - 35LEND 27 - 35LOLA 27 - 35LROC 27 - 35Mini-RF 27 - 35

C&DH 21 - 35TT&C 21 - 35PDE 21 - 35ST-1 & ST-2 27 - 35IMU 21 - 35Propulsion Components 21 - 35DIRWA1-4 21 - 35SAS GCE & HGAS GCE 21 - 35Ka-Band Modulator 27 - 35Ka-Band TWTA 27 - 35

Survival Heaters 27 - 35 <70% duty cycleInstrument OperationalHeaters

31 - 35 <70% duty cycle

Instrument OperationalHeaters

27 - 35 <100% duty cycle

ESS-243 Deleted. Superseded by ESS-16.

3.1.3.2 Ripple

ESS-17 SC subsystems shall meet operational performance requirements in the presenceof ripple as described in Conducted Susceptibility (CS) 01/CS02 testing inSection 3.3.3.1.

3.1.3.3 Turn-on Transients (In-Rush Current)

The SC PSE utilizes SSPC devices to control the power. Unlike the electromechanical switches,the solid-state power switch devices control the turn-on time by limiting the input voltage risetime. The typical turn-on time of a solid-state power switch device is between 800 microseconds(µs) and 3 milliseconds. The input voltage rises linearly with respect to the turn-on time. Thisdelay may eliminate the need for a subsystem to employ an active means for reducing in-rushcurrent at their power input. The subsystem should plan, analyze and test to verify that a currentlimiter is not required.


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ESS-18 The LRO component transient in-rush current shall be within the following limitsas listed below and as provided in Figure 3-1 for 1 Amp Services, Figure 3-2 for 2Amp Services or Figure 3-3 for 5, 10 and 15 Amp Services.

Note that Figures 3-1 (1 Amp Services) and 3-2 (2 Amp Services) are taken directly from thedata sheet for the Micropac Industries model number 53278 28 Volt High Side Latching Solid-State Power Controller. For more information regarding the operation of the 53278, refer to thedatasheet. Figure 3-3 (5, 10, 15 Amp Services) is taken directly from the data sheet for theMicropac Industries series 53503 Latching Isolated Output Solid-State Power Controllers. Formore information regarding the operation of the 53503, refer to the datasheet.

ESS-19 If the inrush current is measured with an electromechanical device as the inputpower switching device, the inrush current of a subsystem shall not to exceed arate of change of 1 amp per microsecond (A/µs) in the first 10 µs.

ESS-20 If the inrush current is measured with an electromechanical device as the inputpower switching device, the inrush current of a subsystem shall not have amaximum rate of change of greater than or equal to 20 mA/µs after the initial 10µs surge.

ESS-21 If the inrush current is measured with an electromechanical device as the inputpower switching device, the subsystem transient current shall never exceed 300%of the maximum steady-state current in the first 10 ms.

ESS-235 If the inrush current is measured with an SSPC as the input power switchingdevice, the inrush current of a subsystem shall not exceed a rate of change of 200milliamp per microsecond (mA/µs) until the voltage reaches the nominal level.

ESS-236 If the inrush current is measured with an SSPC as the input power switchingdevice, the inrush current of a subsystem shall not exceed a rate of change of 20mA/µs after the voltage reaches the nominal level.

ESS-237 If the inrush current is measured with an SSPC as the input power switchingdevice, the transient current shall never exceed 300% of the rated output currentof the SSPC in the first 70 ms.

ESS-22 In-rush current shall be reduced to nominal load at 100 ms after turn-on.


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2

3

4

5

6

7

8

1

0.01 0.1 1 10 TIME (Seconds)

9

10 MUST TRIP LIMIT

GUARANTEED MUST TRIP LIMIT

AT TA > -18O C

NO TRIP LIMIT

2.4 A

1.2 A

CURRENT (Amps)

1 Amp Service

Figure 3-1. SSPC In-rush and Trip Current Limits Curve, 1 Amp Service

4

6

8

10

12

14

16

2

0.01 0.1 1 10 TIME (Seconds)

18

20 MUST TRIP LIMIT


AT TA > -18O C

NO TRIP LIMIT

4.8 A

2.4 A

CURRENT (Amps)

2 Amp Service

Figure 3-2. SSPC In-rush and Trip Current Limits Curve, 2 Amp Service


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500%

100%

0.01

0.1 1 10 TIME (Seconds)

900% MUST TRIP LIMIT


AT TA > -18O C

NO TRIP LIMIT

2 x I

1.2 x I

CURRENT (Amps)

5, 10, 15 Amp Services

Figure 3-3. SSPC In-rush and Trip Current Limits Curve, 5, 10, and 15 Amp Services

3.1.3.4 Survival of Anomalous Voltage

ESS-23 All subsystems shall be designed to not be damaged by any voltage in the rangeof 0 to +40 VDC for an indefinite time period applied to the power input duringanomalistic operations.

Verification of ESS-23 will be by analysis or test on an engineering test unit (ETU) or at a boardlevel only.

ESS-24 Deleted.ESS-25 Deleted.

No flight component will be subjected to these tests.

ESS-26 All subsystems shall meet performance requirements during the single-eventtransient (SET) specified in Section 3.1.2.5.

3.1.3.5 Operational Bus Transients

ESS-27 The rate of change of any operational current transients shall not exceed 20mA/µs.


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3.1.3.6 Turn-Off Transients

ESS-28 When the power service is switched off, the peak voltage transients induced onthe power service shall not exceed +40V, nor fall below -1V.

3.1.3.7 Turn-Off Protection

ESS-29 No subsystem shall be damaged by the unannounced removal of power.ESS-30 Any operations required on a routine basis prior to power turn off shall be listed

in that individual subsystem ICD.ESS-31 Any minimum time following power turn-off that the component must remain off

prior to power turn-on shall be listed in that individual subsystem ICD.

3.1.3.8 Redundant Power Supplies

ESS-32 Any subsystem or component that includes redundant power supply inputs shallnot be damaged by the simultaneous application of power to both interfaces.

3.1.3.9 Polarity Reversal Protection

ESS-33 All subsystems shall not be damaged by polarity reversal of the input power.

3.1.3.10 Subsystem Over-Current Protection

ESS-34 The use of non-resetting over-current protection (i.e., fuses) within the usersubsystems shall be prohibited unless a waiver is requested and approved inaccordance with the Robotic Lunar Exploration Program Mission AssuranceRequirements (430-RQMT-000006).

3.2 SYSTEM GROUNDING REQUIREMENTS

LRO will use proven grounding techniques that have been shown to reduce EMI and conductednoise from within the system. LRO will use a hybrid grounding approach with the primarystructure serving as the low impedance zero voltage reference. Stray AC noise currents areencouraged to flow through structure in order to reduce common mode voltages. Primary powerhas a DC single point ground and AC multipoint grounds to chassis. Secondary power andgrounds are multipoint grounds to chassis.

3.2.1 Single-Point Primary Power Ground

ESS-35 LRO shall implement a single-point grounding scheme for the primary power bus.ESS-36 The Single-Point Ground (SPG) shall be located within the PSE.ESS-37 All the primary power returns, SA returns, and battery grounds shall be tied

together at the SPG and connected to the SC structure (or SC chassis ground).ESS-38 At the component primary power interfaces, primary power (28 VDC) and

primary power returns shall be isolated from the component chassis by greaterthan or equal to 1 Megohms (Mohms) direct current (DC).


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3.2.2 Distributed Signal Ground

ESS-39 LRO shall implement distributed signal grounding scheme for secondary powerreturns and/or signal grounds.

ESS-40 The secondary return (power, signal, analog, or digital grounds) shall be locallyconnected to the component chassis with low impedance paths (<= 2.5 milliohmsDC per joint) to minimize stray current.

ESS-41 Both secondary power (or signal) inputs and returns shall be isolated fromprimary power by equal to or greater than 1 Mohms DC.

3.2.3 Common Mode Noise

ESS-42 Common mode noise for the primary and secondary power, as well as digital,analog and signal grounds, shall be less than 100 millivolts (mV) p-p.

ESS-43 The LRO structure shall not be used intentionally to carry DC current or commonmode current.

3.2.4 Bonding or Mating

ESS-44 The LRO structure shall be used as a common reference point for all electronics.ESS-45 All conductors shall be grounded to the SC structure, no floating conductors.ESS-46 The electrical DC resistance of a mechanical contact between two conductive

mating surfaces shall not exceed 5 milliohms DC.ESS-47 The primary mating method for a component shall be the metal-to-metal contact

between component mounting feet (or baseplate) and the LRO structure. Mating(electrically bonding) surfaces should be free from nonconductive finishes andshould establish sufficient conductive contact surface area to satisfy electricalgrounding requirement (ESS-51).

ESS-48 If a component is to be mounted on a composite or other low conductive material,a grounding strap shall be attached from the component chassis to an Orbiterconductive structure.

ESS-49 The electrical DC resistance of the grounding strap as measured at the componentchassis tie-point and the Orbiter structure tie-point shall not exceed 5 milliohmsDC.

ESS-50 All ground straps for components or subsystems shall have at most a length-to-width ratio of 5 to 1 (width is at least 1/5 of length), be made of copper at least 1mil thick, and still be flexible enough to allow for bending to occur.

ESS-51 The grounding lug location on the component chassis or the tie points in contactwith the ground strap shall have a minimum contact area of 80 millimeterssquared (mm2). This is satisfied by a 1/4" diameter fastener and NAS1149C04-series washer.

ESS-52 The electrical DC resistance between any two mated electronic components onthe LRO structure, measured at the foot of each component, shall not exceed 10milliohms DC.


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ESS-53 Component connectors and backshells shall be electrically mated to the chassisthrough an electrical resistance not exceeding 5 milliohms DC.

3.2.5 Shield Grounds

ESS-54 Shield grounds associated with cables (twisted shielded pairs, coaxial cables, etc.)shall be terminated to chassis ground per Figures 3-4 and 3-5.

3.2.6 Grounding of External Orbiter Surfaces

ESS-55 All external surfaces shall be grounded. Where this is not possible, it will beidentified in the individual component Electrical ICD.

3.2.6.1 Thermal Blankets

ESS-56 All layers of thermal blankets shall be redundantly grounded to the SC structureusing a low-resistance method as follows:

• have at least two grounding tabs for blankets less than 1 meter squared(m**2).• have an additional grounding tab for every square meter greater than 1 m**2ESS-57 Any point on the blanket shall be less than 1 meter from the nearest grounding

point.ESS-58 Insulating films such as Kapton and other dielectric materials shall be less than 5

mil thick, assembled to minimize surface charge build-up, and grounded to bleedsurface charge.

For more information regarding thermal blanket grounding methods refer to the LRO BlanketSpecification (451-SPEC-000TBD)


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Figure 3-4. LRO Spacecraft Shielding Scheme

Power System Electronics

Example Component

PSE Output Module

Output Current Shunt

Single Point

Ground

Chassis Shunt

Primary Power Twisted with Primary Power Return

Input Filter

Output Filter

Secondary Power and

Ground

>1 Mohm isolation from Primary to Secondary

Solar Array Return

Battery Return

+28V

Figure 3-5. LRO Spacecraft Power Isolation Scheme

3.2.6.2 Hinges

ESS-59 A ground strap per Section 3.2.4 shall carry the ground across any hinged joints.


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3.2.6.3 Solar Array (SA) Panels

ESS-60 Solar array (SA) panels and substrates shall be electrically grounded to the SCstructure.

ESS-61 Ground straps shall be implemented per Section 3.2.4.

3.2.6.4 Antennas and Antenna Booms

ESS-62 The HGA assembly shall employ a grounding scheme to assure HGA metalsurfaces and waveguides are grounded directly or indirectly to the SC structurethrough less than 100 milliohms DC resistance.

The HGA boom gimbal rotating joints and deployment hinges should not be considered adequatein providing a good ground path. Therefore, separate ground connections will be provided inslip-rings on a rotating joint or a dedicated ground strap.

ESS-63 The omnidirectional antenna metal surface and cable shields shall be groundeddirectly or indirectly to the SC structure through less than 5 milliohms DCresistance.

3.3 EMI/EMC REQUIREMENTS

ESS-64 Emissions and susceptibility testing shall be performed per this document, whichhas tailored the General Environmental Standards for GSFC Flight Programs andProjects (GEVS) (GSFC-STD-7000) test levels for the LRO mission.

The EMI/Electromagnetic Compatibility (EMC) tests required below are meant to cover theLRO mission environments including Orbiter RF self compatibility, launch site, launch pad,launch/ascent, lunar transfer, and lunar orbit.

ESS-65 Table 3-3 indicates the tests that shall be performed on each component and at theOrbiter level.

Table 3-3. EMI/EMC Applicability and References

Components(1) (2)

RFComp

(3)

Orbiter ESSSection

GSFC-STD-7000Section(4)

461CSection

(4)CE01 X X 3.3.1.1 2.5.2.1 2CE03 X X 3.3.1.1 2.5.2.1 3CE06 X 3.3.1.2 2.5.2.1 4RE02 X X X 3.3.2 2.5.2.2 17CS01 X X 3.3.3.1 2.5.3.1a 6CS02 X X 3.3.3.2 2.5.3.1a 7


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CS03 X 3.3.3.3 2.5.3.1b 8CS04 X 3.3.3.4 2.5.3.1c 9CS05 X 3.3.3.5 2.5.3.1d 10CS06 X X 3.3.3.6 2.5.3.1e 11RS03 X X X 3.3.4 2.5.3.2 21Self

CompatibilityX 3.3.5

Notes: (1) X= Applicable(2) Subsystems may be tested as individual components or assemblies where applicable.Subsystems include instruments (see section 1.2).(3) RF Components includes all RF receiving components and instruments(4) GSFC-STD-7000 and MIL-STD-461C sections provided for reference. Seeapplicable ICD section for LRO specific requirements.Test levels of emissions and susceptibility defined in applicable figures may differ fromGSFC-STD-7000 and MIL-STD-461C, this document takes precedence.

The EMI/EMC tests shall be performed with primary power at the LRO nominal +31VDC,unless otherwise stated.ESS-66 The EMI/EMC test methods shall be per the requirements of MIL-STD-462

(Notice 6) unless noted in this document. EMI/EMC requirements will beimposed on individual components.

ESS-67 Deleted.

When fully integrated into the SC, this will ensure that these components will not interfere witheach other. In addition, the radiated emission and susceptibility requirements are imposed on afully integrated SC. This will ensure that the SC will not adversely affect the launch vehicle,will not be affected by the external emissions (particularly at the launch site), and will notinterfere with any sensitive instruments making science measurements.

ESS-68 All tests shall be performed in ambient with either the component or system in itsmost sensitive mode for susceptibility testing and in its most noisy mode asappropriate for the EMI emission test.

3.3.1 Conducted Emissions

3.3.1.1 CE01/CE03

ESS-69 Conducted emissions (CE) from components shall not exceed the values shown inFigure 3-6 when subjected to CE01 (30 Hz - 20 kHz) and CE03 (20 kHz - 50MHz) narrowband testing.

ESS-70 CE01/CE03 shall be performed on all primary power and return lines to eachcomponent.

ESS-71 CE01/CE03 shall be performed in differential and common mode.


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Conducted emissions testing will be performed only at the subsystem or component levels.

ESS-72 Each component shall meet the transient current pulse limits, both single event(excluding turn-on) and recurring, as specified in Section 3.1.3. Applicable testparameters and limits are as follows for narrowband conducted emissions:

a.Interface lines to be measured are differential mode current lines: primary powerinputs, primary power returns.b.Interface lines to be measured are common mode current lines: primary powerinputs with return including heater circuits.c.Differential mode narrowband test limits are 120 decibel microamps (dBuA)(1.0 A rms) from 30 Hz to 450Hz, then decreasing to 50 dBuA (0.316 mA rms) at20 KHz, then decreasing to 20 dBuA (10uA rms) at 2 MHz, and then continuingat that level to 50 MHz, as shown in Figure 3-6.d.Common mode narrowband test limits are 50 dBuA (0.316 mA rms) from 30Hz to 20 KHz, then decreasing to 20 dBuA (10uA rms) at 2 MHz, and thencontinuing at that level to 50 MHz, as shown in Figure 3-6.

0

20

40

60

80

100

120

140

1.00E+01 1.00E+02 1.00E+03 1.00E+04 1.00E+05 1.00E+06 1.00E+07 1.00E+08

FREQUENCY (Hz)

dBuA

Instrument or Component Level Test Limits (Differential Mode)

Instrument or Component Level Test Limits (Common Mode)

CE01 (30 Hz - 14 KHz) CE03 (14 KHz-50 MHz)

Bandwidth (Hz) 5 Hz 500 Hz 5 KHz 50 KHz

Figure 3-6. Narrowband Conducted Emissions CE01/CE03 Limits


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3.3.1.2 CE06

ESS-73 All RF receivers and transmitters shall perform the additional CE06 EMI test tothe limits contained in MIL-STD-461C.

3.3.2 Radiated Emissions (RE02)

ESS-74 Radiated emissions (RE) from subsystem or components shall not exceed thevalues shown in Figure 3-7 or Figure 3-8 when subjected to RE02 narrowbandtesting.

ESS-75 Radiated electric field emissions from any components that are ON from launchto vehicle separation shall not exceed the limits shown in Figure 3-7 (lower line).

ESS-76 Radiated electric field emissions from any components that are OFF from launchto vehicle separation shall not exceed the limits shown in Figure 3-8.

ESS-77 The aggregate RE from the Orbiter shall not exceed the limits shown in Figure 3-7 (upper line).

The Orbiter receiver has a center frequency at 2091.3967 MHz and the notches in Figure 3-7 andFigure 3-8 will protect the receiver with at least 6 MHz on both sides of the centerfrequency.

The Mini-RF Technology Demonstration receiver has a center frequency at 2380 MHz (S-Band)and 7140 MHz (X-Band), and the notches in Figure 3-7 and 3-8 will protect thereceiver with at least 15 MHz on both sides of the center frequency.

The Atlas V maximum allowable payload RE levels are provided for reference only in Table 3-4.Since the Atlas allowable RE levels are higher than GSFC GEVS levels, the LROOrbiter and subsystems shall be required to meet the GEVS levels per Figure 3-7and 3-8.

Table 3-4. Atlas V Spacecraft RE Limits

Frequency Range dB uV/m14 kHz - 410 MHz 114410 - 430 MHz 37430 MHz - 1 GHz 1141 - 5.660 GHz 1545.660 - 5.720 GHz 735.720 - 18 GHz 154


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3.3.3 Conducted Susceptibility

ESS-78 Undesirable response, malfunction, or degradation of performance shall not beproduced in any components during CS testing with the tests specified below.Performance deviation of Instruments is acceptable as long as the componentunder test survives the component CS test. CS testing will not be performed atthe Orbiter level.

3.3.3.1 CS01/CS02

ESS-79 The CS01 and CS02 (injection of energy into power lines) shall be performed onall components that contain the DC/DC converters or power regulation devices.

ESS-80 The CS01 test limits for the components level tests shall be 2.8 V rms at thefrequency range of 30 Hz to 1.5 kHz, and ramping in a straight line down to 1.0volt at 50 kHz.

ESS-81 The CS02 limit for the component level test shall be 1.0 V rms at the frequencyrange of 50 KHz to 400 MHz.

These limits, which are defined by MIL-STD-461C, are shown in Figure 3-9 below.ESS-82: The CS01 and CS02 (injection of energy into power lines) performance shall

be verified at the nominal +31V only.

3.3.3.2 CS03

ESS-83 The CS03 (Two Signal Intermodulation) test shall be performed on all RFreceiving components.

ESS-84 The CS03 (Two Signal Intermodulation) test performed on all RF receivingequipment shall not cause the RF equipment to exhibit any intermodulationproducts from two input signals, beyond those permitted in the RF componentspecification.

ESS-85 The CS03 test for RF receiving components shall be conducted per MIL-STD-462to the limits specified in GSFC-STD-7000.

3.3.3.3 CS04

ESS-86 The CS04 (Rejection of Undesired Signals) test shall be performed on all RFreceiving components.

The CS04 (Rejection of Undesired signals) test for RF receiving components consists of a 0.0decibel milliwatt (dBm) (decibels referenced to 1 milliwatt) signal applieddirectly to the receiver input terminals and notched around the receiver inputbandwidth at 80.0 decibels (dB) above its threshold.

ESS-87 The input notch center shall be at the receiver-tuned frequency and in the centerof the notch.


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ESS-88 The CS04 test for RF receiving components shall be conducted per MIL-STD-461C to the limits specified in GSFC-STD-7000.

20

30

40

50

60

70

80

10

0.01 0.1 1 10 100 1,000 10,000 100,000 Frequency (MHz)

dB uV/m

14 kHz

Components and Instruments Levels

Orbiter Levels

30 GHz

20 dB uV/m 2091.3967 +/- 6 MHz

(LRO S-band RX)

37 dB uV/m 410–430 MHz

Atlas V

8 dB uV/m 2380 +/- 15 MHz

(LRO Mini-RF S-band RX)

8 dB uV/m 7140 +/- 15 MHz (LRO Mini-RF X-

band RX)

Figure 3-7. RE02 Limits for the Orbiter and Components that are ON from launch tovehicle separation (vertical axis is dBmicrovolts/meter)


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20

30

40

50

60

70

80

10

0.01 0.1 1 10 100 1,000 10,000 100,000 Frequency (MHz)

dB uV/m

14 kHz

Components and Instrument Levels (for components and instruments that are OFF from launch to vehicle separation)

30 GHz

20 dB uV/m 2091.3967 +/- 6 MHz

(LRO S-band RX)

8 dB uV/m 2380 +/- 15 MHz

(LRO Mini-RF S-band RX)

8 dB uV/m 7140 +/- 15 MHz (LRO Mini-RF X-

band RX)

Figure 3-8. RE02 Limits for Components that are OFF from launch to vehicle separation(vertical axis is dBmicrovolts/meter)

2

3

4

5

6

1

0.0001(100 Hz)

0.001(1 kHz)

0.01(10 kHz)

0.1(100 kHz)

1 10 100 1,000

Frequency (MHz)

V r

ms

30 Hz

50 kHz

0.00001(10 Hz)

400 MHz

1.5 kHz2.8 Vrms

1.0 Vrms

2

3

4

5

6

1

0.0001(100 Hz)

0.001(1 kHz)

0.01(10 kHz)

0.1(100 kHz)

1 10 100 1,000

Frequency (MHz)

V r

ms

30 Hz

50 kHz

0.00001(10 Hz)

400 MHz

1.5 kHz2.8 Vrms

1.0 Vrms

Figure 3-9. CS01/CS02 Limits


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3.3.3.4 CS05

ESS-89 The CS05 (Cross Modulation) test shall be performed on all RF receivingcomponents.

ESS-90 The CS05 (cross-modulation) test performed on all RF receiving equipment shallnot cause the RF equipment to exhibit any cross-modulation from two inputsignals.

ESS-91 The CS05 test for RF receiving components shall be conducted per MIL-STD-461C to the limits specified in GSFC-STD-7000.

3.3.3.5 CS06

ESS-92 The CS06 (Powerline Transient) shall be performed on all components thatcontain the DC/DC converters or power regulation devices.

The CS06 (Powerline Transient) test consists of both a positive transient test and a negativetransient test, having amplitude as shown in Figure 3-10 superimposed on the +31V power busas shown in Figure 3-10.

ESS-93 The CS06 positive transient pulse shall be limited to +56V peak absolute valueand a 10 µs width at the +31V steady-state bus crossing point as shown in Figure3-10.

The CS06 negative transient pulse shall be limited to 0V value and a 10 µs width at the the +31Vsteady-state bus crossing point as shown in Figure 3-10.ESS-94 The CS06 test shall be conducted per MIL-STD-462 to the limits specified in

GSFC-STD-7000.

3.3.4 Radiated Susceptibility (RS03)

ESS-95 Undesirable response, malfunction, or degradation of performance shall not beproduced during component, or Orbiter Radiated Susceptibility (RS) testing withthe E-field levels shown in Table 3-5.

The LRO Expendable Launch Vehicle (ELV) will be serviced at the Astrotech facility andlaunched from the Cape Canaveral Air Station (CCAS) and can be exposed to the maximumtransmitter limits shown in Table 3-6. These limits include those expected at payload processingat the Astrotech facility, transport to the Vertical Integration Facility and the pad environment.

ESS-96 The Orbiter components and instrument shall survive the RS test levels of the"Payload Processing and Launch Pad Environments” as given in Table 3-6.

ESS-239 All subsystems and instruments shall be tested with power ON to the levels andfor the frequencies given in Table 3-6 as “Payload Processing and Launch PadEnvironments”

ESS-240 Instruments shall operate in the most typical mission operational state during the"Payload Processing and Launch Pad Environments" test, unless the Instrument


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System Engineer determines that the test will damage the instrument, in that case,the instrument shall then be placed in a mode where no damage will occur.

ESS-241 Instruments and subsystems that are powered ON from launch to launch vehicleseparation shall be powered ON during the RS test at the ELV S-band and C-bandtransmitter frequencies shown in Table 3-6.

ESS-242 Instrument and subsystems that are powered OFF from launch to launch vehicleseparation shall be tested to survive, but may be powered OFF during the RS testat the ELV S-band and C-band transmitter frequencies shown in Table 3-6.

Table 3-5. LRO Operational RS Test Limits

Frequency Range Test Level Requirement Source14 KHz - 2 GHz 2 V/m GSFC-STD-70002 GHz - 12 GHz 5 V/m GSFC-STD-7000

12 GHz - 28 GHz 10 V/m GSFC-STD-70002.271 GHz +/- 5 MHz 12 V/m LRO S-Band Transmitter25.5 GHz - 28.0 GHz 28 V/m LRO Ka-Band Indirect

Radiation2.380 GHz +/- 15 MHz 5 V/m LRO Mini-RF S-band Indirect

Radiation7.14 GHz +/- 15 MHz 5 V/m LRO Mini-RF X-band Indirect

RadiationFrequency Range Test Level (V/m) Test Level (dBuV/m) Requirement Source14 KHz - 2 GHz 2 126 GSFC-STD-70002 GHz - 12 GHz 5 134 GSFC-STD-700012 GHz - 28 GHz 10 140 GSFC-STD-7000

2.271 GHz +/- 5 MHz 12 142 LRO S-BandTransmitter

25.5 GHz - 28.0 GHz 28 149 LRO Ka-BandIndirect Radiation

2.380 GHz +/- 15MHz

5 134 LRO Mini-RF S-bandIndirect Radiation

7.14 GHz +/- 15 MHz 5 134 LRO Mini-RF X-bandIndirect Radiation


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Time useconds

40 30 10 50 60 70 80 0 20 0

8

16 Volts

31

40

48

56

64

24

NOTE: The SC Power Bus Pulse shall not exceed 0Vdc at anytime.

10us

-31Vdc

CS06 Negative 31Vdc 10uS Pulse

Time useconds

40 30 10 50 60 70 80 0 20 0

8

16 Volts

31

40

48

56

64

24

NOTE: The SC Power Bus Pulse shall not exceed 56Vdc at anytime.

10us

25Vdc

CS06 Positive 25Vdc 10uS Pulse

Figure 3-10. CS06 Conducted Susceptibility Test Pulse

Table 3-6. Launch Site/Vehicle RS Test Levels


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Payload Processing andLaunch Pad Environment

Frequency Range V/m Source14 kHz - 2.876 GHz 202.876 - 2.892 GHz 25 GPN-20 & WSR-88D2.892 - 5.547 GHz 205.547 - 5.553 GHz 20 Television Channel 95.553 - 5.684 GHz 205.684 - 5.696 GHz 20 Radar 19.145.696 - 5.794 GHz 205.794 - 5.806 GHz 20 Radar 19.39

12 - 40 GHz 20Atlas V Radiated Emissions

Frequency Range V/m Source14 kHz - 2.206 GHz 202.206 - 2.216 GHz 20 Atlas V Second Stage S-band

T/M2.216 - 5.759 GHz 205.759 - 5.771 GHz 58 Atlas V Second Stage C-band

beacon (peak transmit)5.771 - 40 GHz 20

Payload Processingand Launch Pad

EnvironmentFrequency Range V/m dBuV/m Source

14 kHz - 2.700 GHz 20 1462.700 - 2.900 GHz 25 148 GPN-20 & WSR-88D2.900 - 5.400 GHz 20 1465.400 - 5.900 GHz 40 152 C-Band Tracking

Radars5.900 - 30 GHz 20 146

Atlas V RadiatedEmissions

Frequency Range V/m dBuV/m Source14 kHz - 2.206 GHz 20 146

2.206 - 2.216 GHz 20 146 Atlas V Second StageS-band T/M

2.216 - 5.759 GHz 20 146


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5.759 - 5.771 GHz 58 155 Atlas V Second StageC-band beacon (peak

transmit)5.771 - 30 GHz 20 146

3.3.5 Orbiter RF Self-Compatibility

ESS-97 The Orbiter RF self-compatibility test shall be included in the Orbiter-levelEMI/EMC test.

ESS-98 During Orbiter Self-Compatibility, the Orbiter shall be configured to a nominalscience mode to simulate the in-orbit operation. Ka-band and S-band transmitterswill free-radiate from their antennae during this test.

3.4 DATA AND SIGNAL INTERFACES

ESS-99 The presence or absence of any combination of the input signals applied in anysequence shall not cause damage to a component, reduce its life expectancy, orcause any malfunction, whether the component is powered or not.

3.4.1 Inter-Component Communications

ESS-100 All signals between boxes shall be controlled to limit signal bandwidth so that nosignal should be given more bandwidth than needed to communicate thenecessary functions under all expected on-orbit environmental conditions.

ESS-101 Subsystems connected to the C&DH subsystems via the 1553 data bus shallcommunicate commands and housekeeping telemetry per Section 3.4.1.1.

ESS-102 Subsystems connected to the C&DH subsystem via the SpaceWire network shalluse four twisted pair wires (100+2% ohm cable) with a separate shield aroundeach twisted pair and an overall shield per ESA Space Wire Specification (ECSS-E-50-12).

ESS-103 Subsystems connected using RS-422 differential signals shall adhere to theelectrical terminations as given in the Electrical Characteristics of BalancedVoltage Digital Interface Circuits (TIA/EIA-422).

ESS-104 Subsystems connected using LVDS signals shall adhere to the electricalterminations as given in the Electrical Characteristics of Low Voltage DigitalSignaling (LVDS) Circuits (TIA/EIA-644).

3.4.1.1 LRO 1553 Data Bus

ESS-105 All 1553 Bus Controller (BC), Remote Terminal (RT), and coupling transformerdevices shall comply with MIL-STD-1553B requirements.

3.4.1.1.1 LRO 1553 Data Bus Topology


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ESS-106 The transformer-coupled (long stub) interface shall be implemented for the LRO1553B data bus, as specified in MIL-STD-1553B.

A typical subsystem RT will interface to the data bus with an isolation transformer and acoupling transformer, as shown in Figure 3-11. The complete 1553 data bus willbe assembled as shown in Figure 3-12. The LRO-unique 1553B busimplementation requirements and clarifications are described below.

Isolation Transformer

Transceiver

Remote Terminal

Stub < 10 feet

Coupling Transformer

Isolation

Resistors 59 Ù

(typical)

Data Bus Wire Pair

Bus Shield

Bus Coupler

Bleed Resistor 1 M Ù

(typical)

Bleed Resistor 1 M Ù

(typical)

Figure 3-11. 1553 Data Bus Subsystem Diagram


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Figure 3-12. 1553 Data Bus System Diagram

3.4.1.1.2 LRO 1553 Remote Terminal Address Assignments

ESS-107 The Remote Terminal Addresses shall be assigned per Table 3-7.

Table 3-7. LRO 1553 Remote Terminal Addresses

Address Component or Subsystem0 Reserved (Mode Code)1 Reserved (Single Board Computer (SBC) #1 RT operations if necessary)2 Reserved (SBC #2 if necessary)3 C&DH Multifunction Analog Card4 PSE5 Spare6 Spare7 Spare8 Integrated Reaction Wheel (IRW) 19 IRW 2

10 IRW 311 IRW 4


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12 Star Tracker (ST) 113 ST 214 Inertial Measurement Unit (IMU)15 Reserved (Backup IMU if necessary)16 CRaTER17 Diviner Lunar Radiometer Experiment (DLRE)18 LEND19 Spare20 Spare21 LOLA22 Spare23 Spare24 Propulsion / Deployment Electronics (PDE) A25 PDE B26 PDE C27 PDE D28 SA Gimbal Control29 HGA Gimbal Control30 Spare31 Reserved (Broadcast)

3.4.1.1.3 LRO 1553 Redundancy

ESS-108 A dual, standby-redundant bus as defined in MIL-STD-1553B shall be used.

3.4.1.1.4 LRO 1553 Data Bus Connectors

ESS-109 All 1553 Bus Controller(s) and Remote Terminals (RTs) shall utilize a connectorcompatible with the LRO 1553 harness which uses a Trompeter PL3155ACconnector.

3.4.1.1.5 LRO 1553 Data Bus Cable

ESS-110 The 1553 data bus shall use M17/176-00002 cable.

3.4.1.1.6 LRO 1553 Stub Coupling

ESS-111 The 1553 bus shall use transformer coupling as defined in Section 4.5.1.5.1 ofMIL-STD-1553B.

ESS-112 The 1553 bus coupling transformer shall have a turn ratio of 1:1.41+/-3% (stub tobus).

ESS-113 The 1553 bus coupling transformer shall include fault isolation resistors inaccordance with Section 4.5.1.5.1.2 of MIL-STD-1553B with R = 59 ohms (R =0.75*Zo, with Zo = 78 ohms +/- 10%)


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ESS-114 The 1553 bus shall have a characteristic impedance of 78 ohms +/- 10% and beterminated at both ends of the bus with 78 ohms +/- 10% termination.

ESS-115 The 1553B bus cabling and couplers shall have 100% shielding coverage.

The 1553 stub should be as short as possible (less than 10 feet is preferred).

ESS-116 All RT isolation transformers shall be designed to provide an output signal levelof 18 to 27V p-p at the component output interface.

3.4.1.2 SpaceWire Network Requirements

LRO-unique interface requirements for the high-speed instrument data are to be specified in theLRO Project SpaceWire Specification (431-SPEC-000103).

3.4.1.3 Ka-Band Downlink I&Q Data Interfaces

ESS-117 The Ka-band I-channel and Q-channel (I&Q) downlink data interface from theC&DH to the Ka-Band RF transmitter shall be fabricated per the ElectricalCharacteristics of Low Voltage Differential Signaling (LVDS) Interface Circuits(TIA/EIA-644).

3.4.1.4 Analog Telemetry Interfaces

ESS-118 The SC C&DH shall be capable of monitoring the passive analog, active analogand discrete telemetry signals.

ESS-119 The specific thermistor part number utilized to generate the passive analogtelemetry shall be identified in the Thermal Systems Specification (431-SPEC-000091)

ESS-120 4.02 kilohms bias resistor in the C&DH shall be used in parallel with anythermistor read by the C&DH. The range of the temperature that can be sensed bythe thermistor with the bias resistor is -40 to +75 degree Centigrade (ºC), and witha Platinum Resistance Thermometer (PRT) the range is -170 ºC to + 145 ºC, asrepresented by 0 - 4V telemetry output.

ESS-121 To prevent the current switching noise from the passive analog returns, the activeanalog signal ground or reference shall not be shared with the passive analogreturns.

3.4.1.5 Discrete Command Interfaces

ESS-122 A discrete pulse command shall be used to actuate digital actuators.ESS-123 A discrete pulse shall be transmitted using RS-422 differential signalsESS-124 A discrete pulse shall adhere to the electrical terminations as given in the

Electrical Characteristics of Balanced Voltage Digital Interface Circuits(TIA/EIA-422).

ESS-125 Discrete command receivers shall use a 100 ohm termination resistor to match theimpedance of the RS-422 cable.


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ESS-126 The required RS-422 termination configuration is shown in Figure 3-13. Sending Component Receiving Component

RS-422 Transmitter

RS-422 Receiver

100 ohm

Figure 3-13. Discrete Command Termination

3.4.1.6 +28 V Discrete Command Interfaces

The +28V discrete pulse command will be used to actuate relays or other non-digital actuators.ESS-127 The +28V discrete command pulse circuit shall be designed to minimize coupling

of the actuator switching noise into the digital or logic portion of the electronics.ESS-128 Unregulated +28 VDC from the PSE may be used to generate the +28 V pulse

command; however, the isolation requirements between primary power andsecondary power (logic power) of 1 Mohm shall always be maintained.

3.4.1.7 Pulse per Second (1 pps)

ESS-129 The LRO SC shall provide a 1 Pulse per Second (pps) to subsystems that requireSC timekeeping information.

ESS-130 The 1 pps shall be transmitted using RS-422 differential signals and shall adhereto the electrical terminations as given in Section 3.4.1.5 above.

ESS-131 Users of the 1 pps shall use receivers and termination compatible with the C&DHwhich is using an Intersil HS-26CLV31RH differential driver.

ESS-132 The 1 pps shall have timing characteristics as shown in Figure 3-14 and Table 3-8(below). The fiducial occurs on the rising edge of 1pps_int.


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1pps_int (internal to C&DH)

tLH tHL

t1s

tPLH

10%

90%

1pps+

1pps-

tPLH

tH

Figure 3-14. 1 pps Timing Characteristics

Table 3-8. 1 pps Timing Characteristics

Parameter Value DescriptiontPLH, tPHL 2-30 nanoseconds Delay through drivertTLH, tTHL 1-13 nanoseconds Rise & Fall times

tH 34-36 microseconds Pulse widtht1S 1s +/- 10 nanoseconds Pulse-to-pulse accuracy

3.4.1.8 Other Inter-Component Communications

ESS-133 Other component interfaces not specified in the above Sections 3.4.1.1 through3.4.1.5 shall be identified in the specific applicable subsystem to subsystemElectrical ICDs.

3.4.2 Pyrotechnic and Deployable Actuator Interfaces

ESS-134 The pyrotechnic and deployable control circuits utilized for the LRO pyrotechnic-actuated valves and deployable devices shall meet the Range Safety UserRequirements (AFSCM 91-710).

The following LRO-unique requirements are derived from these mission success requirements:


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ESS-135 The actuator initiation circuit shall have three independent inhibits, two separationbreakwires, and at least one C&DH command to protect against any catastrophichazards.

ESS-136 The actuator initiation circuit shall have two independent inhibits, one separationbreakwire, and at least one C&DH command to protect against any criticalhazards.

ESS-137 The pyrotechnic actuation circuits shall limit stray currents at the pyrotechnicdevice to less than 20 dB below the no-fire current level.

ESS-138 The actuation device and its control circuits shall be isolated from any externalenergy source by greater than 20 dB.

ESS-139 Electroexplosive devices shall be protected from electrostatic hazards by theplacement of resistors from line-to-line and from line-to-ground (Structure). Theparallel combination of all resistors must be 10k ohms or more (MIL-STD-1576).

ESS-238 The open/close position of the pyrotechnic valves shall be remotely monitoredduring pre-launch operations, or the functional power shall be deenergized (anadditional fourth inhibit shall be in place between the power source and the threerequired inputs) and the control circuits for the three required inhibits shall bedisabled (no single failure in the control circuitry will result in the removal of aninhibit) until the hazard potential of inadvertent activation no longer exists.

3.4.3 External Interfaces

ESS-140 LRO shall provide test points and other external interfaces to support thecomponent integration, anomaly investigation, environmental testing, and launchsite operations.

External electrical interfaces and access panels will be defined in the individual subsystemsElectrical ICDs.

The LRO SC is to be launched on an expendable launch vehicle. Specific interface requirementsbetween Orbiter GSE and the launch vehicle, and between the SC and the launch vehicle, will bedefined in the applicable launch vehicle ICD(s).

3.4.3.1 Umbilical Interface Connector

ESS-141 The electrical interconnection between Orbiter and the launch vehicle shall bemade through the umbilical connectors.

ESS-142 There shall be two umbilical interface connectors that are capable of carrying upto 61 each of 22 AWG or larger contacts to support the launch pad operationswith the vehicle.

ESS-143 The umbilical connector(s) shall, at a minimum, include the following functions:hard-line control of the battery relay, battery charging/SA Simulator input and


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monitoring, hard-line communications, health-and-safety critical hard-line (H/L)telemetry data, and launch vehicle interface signals.

ESS-144 Umbilical signals shall be protected from ESD threats. Covering the umbilicalconnector(s) following the separation from the launch vehicle is difficult toachieve, therefore filtering is a preferred method of the ESD protection.

3.4.3.2 Launch Vehicle Separation Signals

ESS-145 Two independent separation break wires shall be required to sense the validseparation from the launch vehicle.

These signals will be utilized to inhibit the SA deployment, high-gain antenna deployment, RFtransmitters, thruster valves and the Reaction Wheel power on following the successfulseparation.

3.4.3.3 Component Test Interfaces

ESS-146 Component test points that require access during Orbiter-level testing shall bebrought to the Spacecraft Test Panel.

ESS-147 All test points shall be protected or isolated from the facility-induced noise, ESD,and GSE malfunction.

ESS-148 GSE cable connectors that mate with flight test connectors shall be flight-approved connectors.

ESS-149 Test connectors shall be capped with flight-approved RF and static control coverswhen not in use, including in orbit.

ESS-150 Wherever possible, component power shall not be applied or accessed at orthrough a test connector.

Test signals and flight signals should not be located together in the same connector.

3.5 MULIPACTION AND CORONA

ESS-151 Damage or measurable degradation due to RF breakdown (corona and arcing)shall be prevented by design.

ESS-152 The RF design shall preclude measurable degradation due to multipaction andcorona in RF systems that must operate during the launch and ascent stages (e.g.,filters, switches, and antenna elements), at critical pressures, or in a vacuumenvironment.

ESS-153 Components with high-voltage circuits shall be immune to corona and arcingwhile in a nominal orbital vacuum environment.

3.6 DESIGN FOR RADIATION

ESS-154 LRO components shall meet their performance requirements with acceptabledegradation due to radiation induced effects. The LRO radiation requirements are


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contained within the Radiation Requirements for the Lunar ReconnaissanceOrbiter (431-RQMT-000045).

3.7 CHARGING AND DISCHARGING REQUIREMENTS

ESS-155 The Orbiter shall meet all its performance requirements in the presence of variouscharging and discharging environment effects induced by the LRO trans-lunar andlunar orbit environments.

3.7.1 External Surface Charging

The Orbiter will be in a direct insertion lunar orbit, including one pass through the Van Allenbelts, a high-charging plasma environment. It has been shown that in these areas, charge buildsup and discharges, causing either a direct hit (lightning strike) or a possible large ground currentthat can disrupt the SC electronics. These can discharge between Orbiter elements (differentialcharging) or can discharge back to space (bulk charging).

3.7.1.1 Differential Charging

ESS-156 The Orbiter shall minimize differential charging of the external surfaces.ESS-157 The Orbiter, when exposed the photoelectron environment shall exhibit maximum

differential charging less than the insulation breakdown voltage, with a factor of 2margin.

3.7.1.2 Charging Mitigation

ESS-158 The Orbiter design shall prevent surface and internal charging/ discharging effectsthat can damage the SC components or disrupt SC operations.

ESS-159 External surfaces >6 cm**2 shall be conductive with a resistivity of less than10**9 ohms per square (ohms/sq.) and grounded to the Observatory structure perSection 3.2.5, so that charge can bleed from that surface faster than the charge canbuild up on that surface.

ESS-160 There shall be no more than 10 surfaces of >6 cm**2 in size with surfaceresistivity greater than 10**9 ohms/sq. on the outer surface per 1 square meter ofexternal surface.

ESS-161 Insulating films such as Kapton and other dielectric materials on the externalsurface shall be less than 5 mil thick and assembled to minimize surface chargebuild-up and grounded to bleed surface charge.

Non-conductive surface areas on the SAs and instrument optics will be minimized to the extentpossible.ESS-162 System impacts of the discharges from any unavoidable non-conductive surfaces

shall be assessed and approved. If necessary, a waiver request will be submittedfor each specific non-compliant application.


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ESS-163 A comprehensive list of non-conductive surfaces on the Orbiter shall bemaintained and reviewed periodically throughout the LRO development toidentify any new ESD threats to the Orbiter.

ESS-164 The list shall include assessment of each dielectric region of the Orbiter surfacefor its breakdown voltage, its ability to store energy, and its effects onneighboring electronics (disruption or damage) and surfaces (erosion orcontamination).

ESS-165 The Orbiter design shall minimize bulk charging of the Orbiter relative to thespace plasma.

ESS-166 Surfaces exposed to the Sun shall contain >1m**2 conductive surface that isgrounded to Orbiter structure in order to bleed off accumulated charges back tospace.

ESS-167 Proper handling, assembly, inspection, and test procedures shall be developed toinsure the electrical continuity of the Orbiter surface grounding.

ESS-168 Surface conductivity shall be verified by measuring resistance from any one pointof material surface to the Orbiter structure.

ESS-169 All grounding methods shall be demonstrated to be acceptable over the servicelife of the Orbiter.

3.7.2 Surface Discharging Protection

ESS-170 The Orbiter shall be designed to prevent discharges on the external surfaces frompermanently damaging components or upsetting science data collection.

ESS-171 The electrical system shall be designed to carry discharge currents and to shieldfrom the electric field from the discharge without any permanent damage to theOrbiter.

ESS-172 SA panels shall use materials and fabrication techniques to minimize ESD effects.ESS-173 Power system electrical design shall incorporate features to protect against

transients due to electrical discharge from the SA.ESS-174 SC transmitters and receivers (command line and data line) shall be immune to

transients produced by ESD.ESS-175 The transmitter, receiver, and antenna system shall be tested for immunity to ESD

near the antenna feed.ESS-176 The repetition rate shall be selected to be consistent with estimated arc rates of

nearby materials.ESS-177 The Orbiter structure shall be designed to serve as a Faraday cage providing 40

dB overall shielding of the internal electronics from the external environment:• The Faraday cage consists of the K1100 structure panels or an RF shieldproviding 40 dB attenuation for any aperture not covered with panels.• An RF shield can be the equivalent of at least four layers of VDA2 material,with each layer grounded to chassis along its entire perimeter.• Joints, seams, and seals between panels can be constructed from copper tapewith conductive adhesive or EMI gaskets. Gaps should be less than 2.5 cm inlength.


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ESS-178 The Faraday cage shall not have any gaps or holes larger or longer than 2.5 cmrequired to maintain 40 dB shielding to 500 MHz from the external noisyenvironment.

ESS-179 All internal harnesses and electronics shall be kept 20 cm away from uncoveredor shielded structure openings or apertures that will be present in flight.

ESS-180 There shall be no openings that provide a direct path for charged plasma to passinto the Observatory structure (Faraday cage).

ESS-181 The Orbiter electrical elements and harnesses outside of the Orbiter mechanicalbus structure (Faraday cage) shall be protected from the discharges generated byESD sources.

ESS-182 Sufficient shielding shall be used around cables to protect against circuit damageor operational upsets from discharges and their RF emissions.

A minimum 1 mil aluminum foil or aluminum tape wrapped with a 50% overlap around externalharnesses up to the entrance of the Faraday cage can be used to seal off the externalenvironments. It should be noted that additional shielding will be necessary to prevent theinternal charging due to the Teflon insulation of the harness (see Section 3.7.3).ESS-183 External circuits shall be filtered at entrance to the Faraday cage. If filtering is

not practical at the entrance to the Faraday cage, the harness shield shall becontinued internal to the Faraday cage after being sealed and grounded at theentrance, and the harness wires shall be shielded and separately routed to theirdestination component and then filtered at the component.

ESS-184 External sensors and harness (i.e., thermistors) shall have an outer shield that isgrounded 360 degrees circumferentially at the entrance to the Faraday cage.

ESS-185 Shields terminated at the entrance to the Faraday cage shall not protrude morethan 2.5 cm into the Faraday cage.

ESS-186 Inner shields connected to the outer shield shall not be brought through aconnector pin into the Faraday cage.

ESS-187 Exposed voltages above 50V shall be protected from shorts to ground caused by aplasma cloud generated by a discharge.

External harness charge mitigation is represented pictorially in Figures 3-4 (above) and 3-15(below).


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( </=5 miliohmresistance between

box chassis tocable sheilds )

Thermistoror PRT

Filter Connectorsfor thermal sensors

ExternalElectronics

External Load

Outer HarnessShield 100%

coverage

Outer Harness Shields arebonded 360 degrees to theComponents and to the SC

Multi WireHarness

INSIDEOBSERVATORYFADADAY CAGE

ExternalEnvironment

All wires/cables shall have an overall Shield with100% coverage for harnesses exterior to the SCStructure.

Shields shall have their ends tied to chassis 360degrees around the connector backshells at bothends of the harness.

External sensors (i.e., Thermistors, CSS, etc.)shall require filter connectors at the input to theObservatory "Faraday Cage".

Figure 3-15. External Harnesses and Charge Mitigation

3.7.3 Internal Charging

ESS-188 The Orbiter design shall prevent internal charging/discharging effects that maydamage the Orbiter components or disrupt Orbiter operations. Internal dielectricsand metal charging/discharges occur from those energetic electrons that penetratethe Orbiter surfaces.

ESS-189 Internal charging effects shall be controlled by shielding all electronics elementswith sufficient aluminum equivalent thickness (100 mil Al for bulk dielectrics orto 60 mil Al equivalent for Teflon harness insulation) so that the internal chargingrate is benign.

ESS-190 Internal dielectrics materials with bulk resistivity of >10**12 ohm-cm that do notmeet the shielding requirement shall be controlled via one of the methodsdescribed in the following paragraphs below. Materials typically include


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unavoidable dielectrics such as the SA insulators, antennas, connectors, thermalisolators, thermistor mounting, thermal blankets, thermal blanket tape, and othermaterials such as Kapton and Teflon insulators, which have a bulk resistivity ofhigher than 10**18 ohm-cm. The following techniques are to be used:

• Limit the electron flux to insulators by shielding to 10**10 electrons/cm**2 in10 hours. (This can be met with plate shielding with 110 mil Al for bulkdielectrics or to 60 mil Al equivalent for Teflon harness insulation.)• Filter nearby circuitry to withstand a 900-V, 3-pico Farad (pF), 1500-ohmdischarge.• Detailed analysis of discharge could result in smaller or larger discharge sourcethan above. Assess discharge threat to circuits that cannot be totally shielded suchas antennas, umbilical connector, thermistors, platinum resistance temperature(PRT) sensors, heaters, SA, Coarse Sun Sensor (CSS), etc.• Coat the exterior surface of the dielectric with a grounded layer with aresistivity of <10**9 ohm/sq.• Prevent the discharge from reaching a victim circuit by EMI shielding and orgrounded conductive barrier that will safely absorb and dissipate the discharge.

ESS-191 If none of above control techniques can be applied, the impacts of the dischargefrom the dielectrics material shall be assessed for an approval.

ESS-192 Ungrounded (floating) conductors shall not be allowed in the Orbiter. Thisincludes unused wires in harnesses; ground test sensors; ground use signals inumbilical cables; unused or unpopulated circuit board traces; ungroundedintegrated circuit (IC), relay, transistor, or capacitor cases; spare pins inconnectors; thermal blankets; aluminum or copper tape; ungrounded bracketry forharness or connectors; TC105 harness tie-down clips; harness P-clamps;conductive epoxy; thermostat cases; screws; or nut plates. Exceptions are allowedby waiver if analysis shows that no direct or radiated path to victim circuitryexists or that the victim can survive discharge.

ESS-193 Leakage impedance of conductive internal parts shall be less than 10,000 ohms.This requirement applies to conductive fittings on dielectric structural parts.Further investigation into these effects and mitigations of internal charging can befound in the Avoiding Problems Caused by Spacecraft On-Orbit InternalCharging Effects Handbook (NASA-HDBK-4002).


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4.0 HARNESS REQUIREMENTS

ESS-194 Orbiter harnesses shall satisfy the requirements of this section and the electricalsystems requirements described in Section 3 of this specification.

ESS-195 Harnesses shall be developed Design and Development Guidelines for SpaceflightElectrical Harnesses (565-PG-8700.2.1).

ESS-196 Qualified wire, cable, and connectors specified in the Instructions for EEE PartsSelection, Screening, Qualification, and Derating (EEE-INST-002) shall be usedfor the SC flight harness or any of the component harnesses.

ESS-197 Wires, connectors, connector contacts, and other harness piece parts shall bederated per the Instructions for EEE Parts Selection, Screening, Qualification, andDerating (EEE-INST-002).

ESS-198 Minimum wire sizes shall be per Table 4-1, subject to the limitations inherent inthe Power System Electronics (PSE) connectors.

Table 4-1. Minimum and Maximum AWG

Type Maximum AWG(minimum size)

Minimum AWG(maximum size)

2 Amp Unswitched Services 22 225 Amp Unswitched Services 22 201 Amp Switched Services 22 222 Amp Switched Services 22 20 (limit 3)5 Amp Switched Services 22 2010 Amp Switched Services 20 2015 Amp Switched Services 20 20Heater Circuits 22 22SpaceWire 28 26Signals (other thanSpaceWire)

24 24

4.1 GENERAL HARNESS GUIDELINES

Harnesses should be designed to meet the functional requirements of the subsystems, and therequirements given in Section 3.0, while adhering to the following guidelines:ESS-199 Wherever possible, the harnesses shall be grouped by common electrical

characteristics of the signals carried in the harness.ESS-200 All power lines and power return lines to a particular component shall be twisted

together in a single bundle. Power harnesses should be routed away from signalharnessing, and power and signal wires should not be run in the same bundle.

ESS-244 Length of wire twist shall be between 8 and 16 times the outer diameter of theHarness, in accordance with the Section 7.3 (20) of Crimping, InterconnectingCables, Harness, and Wiring (NASA-STD-8739.4)


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ESS-201 Wherever two or more power conductors are used to increase the current carryingcapability, or provide redundant wires, the component interfaces to the twoconductors shall be of the same design

ESS-202 Wherever two or more power conductors are used to increase the current carryingcapability, or provide redundant wires, the wires shall be routed in the sameharness to keep path length the same for both

ESS-203 Wherever two or more power conductors are used to increase the current carryingcapability, or provide redundant wires, there shall be the same number of powerand return lines, and each return wires shall be twisted with a power wire

ESS-204 Wherever possible, the power and signal shall not share the same componentinterface and harness connectors.

ESS-205 The connector half that sources power to another component shall be female(socketed) to protect against inadvertent grounding prior to mating.

ESS-206 The pyrotechnic harness shall be a twisted, shielded cable without anyinterruptions in the outer shield.

4.1.1 Accessibility

Harnesses should be designed with accessibility and manufacturability in mind:ESS-207 The component connectors, on each box, shall be spaced far enough apart to

access the harness connector with EMI backshells by a hand or with an extractiontool.

Any harness cable or connector should not touch any other adjacent connectors or harnesses.Electrical boxes should be spaced 6 inches (15 cm) from a panel or structure mounting strutwhen the box has one or more connectors that face that structure.Electrical boxes should be spaced 6 inches (15 cm) apart when one of these boxes has connectorsthat face the other box.Electrical boxes should be spaced 9 inches (23 cm) apart when both of these boxes haveconnectors that face each other.

ESS-208 Mechanical support for harnesses shall be designed in accordance with theCrimping, Interconnecting Cables, Harness, and Wiring (NASA-STD-8739.4).

ESS-209 Harness splices (except for heaters) shall not be allowed without LRO projectapproval.

ESS-210 Wires that have tin coating shall not be used for flight due to the possibility of tinwhisker growth that could cause a short. Silver is the preferred coating.

4.1.2 Harness Shields

ESS-211 Harness shields shall be terminated at the connector backshell at both ends ofeach harness:

ESS-212 All shields shall be grounded to the SC structure.ESS-213 Wire harness shields shall not carry current by design.


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ESS-214 Outer shields shall not be tied to component connector pins or wires in the cablebundle.

Inner-shield pig-tail lengths to ground should be minimized, and should not exceed 5 cm.

ESS-215 Shield termination pig-tails shall be bonded to the connector backshell.

Pig-tails should only pass through a connector pin when used as a special noise reductiontechnique (as in the case of LVDS [or SpaceWire Cable] inner shields).

ESS-216 Shields shall not be daisy-chained from one shield to another. Each internalharness shielded wire should connect its shield directly to the ground point.

ESS-217 EMI backshells shall be required for all harness connectors on the externalsurfaces of the Orbiter.

ESS-218 All external harnesses shall be shielded with an overall outer bundle shield andterminated with 360° outer shield bond to the backshell.

ESS-219 Aluminum tape (LG-1055 or equivalent) at least 1 mil thick, wrapped with 50%overlap on each 360-degree wrap over the previous wrap shall be used as harnessbundle shield. For the external harnesses, the additional shielding may berequired to protect from the deep dielectric charging.

ESS-220 All internal interconnections and cables between components shall be bundleshielded and terminated.

4.1.3 Component Test Connector Panels

ESS-221 Component test connector panels shall be located on an external Orbiter surface,and these panels should be accessible during various I&T and launch padoperations.

ESS-222 All electrical connectors (with the exception of the spacecraft umbilicalconnectors) shall be covered such that they are not exposed to the trans-lunar andlunar orbit environments.

4.1.3.1 Safing Plugs or Arming Plugs

ESS-223 Safing and arming plugs shall be incorporated in the cable or harness that controlordinance, deployable actuators, propulsion valves, RF transmitters, SA power,Battery power, electromechanical actuator devices and lasers.

4.1.3.2 Fairing Access to Connector Plugs

ESS-224 Access through the launch vehicle fairing shall be provided for contingencyoperations such as fuel off-loading, safety related inhibit/critical circuit arm plugremoval, contamination event inspection, long term storage/maintenance andlimited troubleshooting.


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The emphasis is on completing all required SC/payload access requirements prior to fairinginstallation to simplify pad operations, access through fairing doors should be not be planned forroutine pad operations.

4.1.3.3 Fuse Pigtails

ESS-225 Power harnesses shall use break-out-boxes (BOB) for fuses to permit theinterruption of the unswitched power bus to a faulty component or harness duringI&T.

4.2 ELECTRICAL MATERIALS

ESS-226 All subsystem materials list shall be reviewed for their electrical properties andassessed for its compatibility with electrical systems requirement defined in thisspecification.

ESS-227 Parts with unstable materials (in terms of the LRO orbital environment) thatcannot be stabilized through additional processing for the proposed applicationshall not be used.

4.2.1 Connectors

ESS-228 All connectors on a component shall be different sizes, pin counts and/or gendersto prevent any mismating of harness to component connectors.

ESS-229 Wherever possible, keying shall be used.ESS-230 Connector savers shall be used during integration and test to minimize wear on

connector contacts.ESS-231 Connector mate/demate logs shall be used to record mates and demates.ESS-232 Instrument developers are expected to provide the mating connectors to their

flight components for use on the SC harness.ESS-233 All subsystems shall provide a list of connectors to the SC electrical systems

engineer prior to connector part procurement.ESS-234 All component input/output (I/O) connector interfaces shall be designed to

accommodate an EMI backshell.


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5.0 QUALIFICATION ASSURANCE PROVISIONS

5.1 GENERAL

All requirements in this document shall be verified by one of the four methods defined below.

5.1.1 Analysis

The analysis method is used when:• A rigorous, representative, and conclusive analysis is possible• Test is not cost effective, and• Inspection and demonstrations are not adequateAnalyses may include, but are not limited to, engineering analysis (which includes models andsimulations), review of record, and similarity analysis.

5.1.1.1 Engineering Analysis

Engineering analysis may be quantitative, qualitative, or a combination of the two. Quantitativeanalysis involves the study and modeling of the physical entity whose performance is to beverified. Examples of quantitative analyses include end-to-end link analysis, structural (staticand dynamic) analysis, thermal models, pointing knowledge and stability. Qualitative analysesare non-numerical and related to qualitative measure of performance, such as failure modes andeffects analyses (FMEA), maintainability, and redundancy.

5.1.1.2 Validation of Records and Other Documentation Analysis

This kind of analysis uses design and manufacturing documentation to show compliance ofdesign features and manufacturing processes. Validation of design documentation, e.g.,engineering drawings, verifies that the “as-designed” hardware complies with contractual designand construction requirements. Validation of manufacturing records at end-item acceptanceverifies that the “as-built” hardware has been fabricated per the approved design and associateddocumentation. Review and analysis of other documentation such as acceptance data packagesand other compliance documentation of lower levels of assembly are valid analysis techniques.

5.1.1.3 Similarity Analysis

Similarity is included as a valid verification/qualification method. Qualification by similarity isused in lieu of test when it can be shown that an item is similar to, or identical in design toanother item that has been previously qualified to equivalent, or more stringent requirements.Formal qualification documentation of the previously qualified item must be available forassessment when planning to qualify by similarity. Furthermore, an item whose design has beenqualified by similarity must undergo acceptance verification to assess workmanship.


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5.1.2 Demonstration

Demonstration is a verification method that provides a qualitative determination, rather thandirect quantitative measurement, of the properties or functional characteristics of an end-item.The qualitative determination is made through observation with, or without test equipment orinstrumentation.

5.1.3 Inspection

Inspection is the verification method used to verify construction features, workmanship,dimension, physical characteristics, and spacecraft conditions such as configuration, cleanliness,and locking hardware. Inspection also includes simple measurements such as length, and it isperformed without the use of special laboratory or precision equipment. In general, requirementsspecifying function or performance are not verified by inspection.

5.1.4 Test

Verification by test consists of direct measurement of performance parameters relative tofunctional, electrical, mechanical, and environmental requirements. These measurements areobtained, during or after controlled application of functional and environmental stimuli to the testarticle, e.g., payload or satellite, and using instrumentation or special test equipment that is notan integral part of the test article being verified. The test activities include reduction andanalysis of the test data, as appropriate. The following paragraphs define different categories oftests including performance, functional, environmental, interface, and structural tests.

5.1.4.1 Performance Test

A performance test consists of an individual test or series of electrical and/or mechanical testsconducted on flight, or flight-configured hardware and software at conditions equal to, or lessthan design specifications. Its purpose is to verify compliance of the test article with the statedapplicable specification requirements that are verifiable by test. Typically, a full performancetest is conducted at ambient conditions at the beginning and the end of a test sequence duringwhich the test article is subjected to applicable environmental conditions, e.g., vacuum, high/lowtemperature extremes, or acoustics/random mechanical excitation.

5.1.4.2 Functional Tests

A functional test is a suitably chosen subset of a performance test. Typically, functional tests areconducted at ambient conditions between environmental exposures during the qualification oracceptance test sequence. The objective is to verify that prior to application of the nextenvironment, exposure to the environment has not adversely affected the test article. Whenappropriate, functional tests, or a portion thereof, are conducted while the test article is exposedto a particular thermal or vacuum environment. Functional test, or a portion thereof, may also beconducted to assess the state of health of the hardware after major operations, such astransportation of flight hardware from one location to another.


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5.1.4.3 Environmental Tests

Environmental testing is an individual test or series of tests conducted on flight, or flight-configured hardware to assure that flight hardware will perform satisfactorily after it is subjectedto the induced launch environments, as well as its flight environment. Examples are: vibration,acoustic, temperature cycling, thermal vacuum and vacuum outgassing certification, andElectromagnetic Interference/Compatibility. Depending on the severity of the chosenenvironmental conditions, the purpose of the environmental exposure is to sufficiently stress thehardware so as to verify the adequacy of the design (protoflight levels and durations) orworkmanship during fabrication (acceptance levels and durations).

5.1.4.4 Special Tests

Special tests are individual tests, or a series of tests conducted on flight, or flight-configuredhardware to assure satisfactory performance of a particular critical element of the system, e.g.,optical alignment. The special test verification category includes structural, mechanism andcommunication tests. Special tests may, or may not be performed in conjunction withenvironmental exposure.

5.1.4.5 Interface Tests

Interface tests verify the mechanical, electrical, and/or hardware-software interface between unitsand elements integrated into a higher level of assembly such as a module, subsystem, element, ora system.

5.1.4.6 Structural Tests

These tests are performed on structural elements, components, or assembled subsystems beforedelivery of the assembled structure to the integration and test organization. Structural testsdesigned to verify requirements of this specification may include: (1) static structural proof tests(to verify the strength/stiffness adequacy of the primary load path), and (2) dynamic tests, suchas a modal survey or acoustic response test.

5.2 VERIFICATION MATRIX TABLE

Requirement Number Heading I A D T ResponsibleOrg

ESS-1 3.1.0-1 Power 4ESS-2 3.1.1.0-1 Power Distribution and Switching Scheme 3ESS-3 3.1.1.0-2 Power Distribution and Switching Scheme 3ESS-4 3.1.1.0-3 Power Distribution and Switching Scheme 2ESS-5 3.1.2.1.0-1 LRO Power Redundancy 4ESS-6 3.1.2.2.0-1 Power System Electronics Output Switching

Profile4

ESS-7 3.1.2.2.0-2 Power System Electronics Output SwitchingProfile

4


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ESS-8 3.1.2.2.0-3 Power System Electronics Output SwitchingProfile

4

ESS-9 3.1.2.3.0-1 Power Steady State Power 4ESS-10 3.1.2.3.0-2 Power Steady State Power 3ESS-11 3.1.2.4.0-1 Power Bus Ripple 4ESS-12 3.1.2.4.0-2 Power Bus Ripple 2ESS-13 3.1.2.5.0-1 Single-Event Power Bus Transients 3ESS-14 3.1.2.6.0-1 Over-Current Protection Deratings 4ESS-15 3.1.2.8.0-1 Insulation 4ESS-16 3.1.3.1.0-1 Steady-State Voltage 3ESS-243 3.1.3.1.0-4 Steady-State Voltage 3ESS-17 3.1.3.2.0-1 Ripple 3ESS-18 3.1.3.3.0-2 Turn-on Transients (In-Rush Current) 4ESS-19 3.1.3.3.0-3 Turn-on Transients (In-Rush Current) 4ESS-20 3.1.3.3.0-4 Turn-on Transients (In-Rush Current) 4ESS-21 3.1.3.3.0-5 Turn-on Transients (In-Rush Current) 4ESS-235 3.1.3.3.0-6 Turn-on Transients (In-Rush Current) 4ESS-236 3.1.3.3.0-7 Turn-on Transients (In-Rush Current) 4ESS-237 3.1.3.3.0-8 Turn-on Transients (In-Rush Current) 4ESS-22 3.1.3.3.0-9 Turn-on Transients (In-Rush Current) 4ESS-23 3.1.3.4.0-1 Survival of Anomalous Voltage 3ESS-24 3.1.3.4.0-2 Survival of Anomalous VoltageESS-25 3.1.3.4.0-3 Survival of Anomalous VoltageESS-26 3.1.3.4.0-4 Survival of Anomalous Voltage 3ESS-27 3.1.3.5.0-1 Operational Bus Transients 4ESS-28 3.1.3.6.0-1 Turn-Off Transients 4ESS-29 3.1.3.7.0-1 Turn-Off Protection 3ESS-30 3.1.3.7.0-2 Turn-Off Protection 3ESS-31 3.1.3.7.0-3 Turn-Off Protection 3ESS-32 3.1.3.8.0-1 Redundant Power Supplies 3ESS-33 3.1.3.9.0-1 Polarity Reversal Protection 3ESS-34 3.1.3.10.0-1 Subsystem Over-Current Protection 3ESS-35 3.2.1.0-1 Single-Point Primary Power Ground 2ESS-36 3.2.1.0-2 Single-Point Primary Power Ground 2ESS-37 3.2.1.0-3 Single-Point Primary Power Ground 2ESS-38 3.2.1.0-4 Single-Point Primary Power Ground 4ESS-39 3.2.2.0-1 Distributed Signal Ground 3ESS-40 3.2.2.0-2 Distributed Signal Ground 4ESS-41 3.2.2.0-3 Distributed Signal Ground 4ESS-42 3.2.3.0-1 Common Mode Noise 3


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ESS-43 3.2.3.0-2 Common Mode Noise 2ESS-44 3.2.4.0-1 Bonding or Mating 2ESS-45 3.2.4.0-2 Bonding or Mating 4ESS-46 3.2.4.0-3 Bonding or Mating 2ESS-47 3.2.4.0-4 Bonding or Mating 3ESS-48 3.2.4.0-5 Bonding or Mating 2ESS-49 3.2.4.0-6 Bonding or Mating 2ESS-50 3.2.4.0-7 Bonding or Mating 4ESS-51 3.2.4.0-8 Bonding or Mating 4ESS-52 3.2.4.0-9 Bonding or Mating 2ESS-53 3.2.4.0-10 Bonding or Mating 3ESS-54 3.2.5.0-1 Shield Grounds 3ESS-55 3.2.6.0-1 Grounding of External Orbiter Surfaces 4ESS-56 3.2.6.1.0-1 Thermal Blankets 4ESS-57 3.2.6.1.0-2 Thermal Blankets 4ESS-58 3.2.6.1.0-3 Thermal Blankets 4ESS-59 3.2.6.2.0-1 Hinges 4ESS-60 3.2.6.3.0-1 Solar Array (SA) Panels 4ESS-61 3.2.6.3.0-2 Solar Array (SA) Panels 3ESS-62 3.2.6.4.0-1 Antennas and Antenna Booms 4ESS-63 3.2.6.4.0-2 Antennas and Antenna Booms 4ESS-64 3.3.0-1 EMI/EMC Requirements 4ESS-65 3.3.0-2 EMI/EMC Requirements 4ESS-66 3.3.0-6 EMI/EMC Requirements 3ESS-67 3.3.0-7 EMI/EMC RequirementsESS-68 3.3.0-8 EMI/EMC Requirements 3ESS-69 3.3.1.1.0-1 CE01/CE03 3ESS-70 3.3.1.1.0-2 CE01/CE03 3ESS-71 3.3.1.1.0-3 CE01/CE03 3ESS-72 3.3.1.1.0-4 CE01/CE03 3ESS-73 3.3.1.2.0-1 CE06 3ESS-74 3.3.2.0-1 Radiated Emissions (RE02) 3ESS-75 3.3.2.0-2 Radiated Emissions (RE02) 3ESS-76 3.3.2.0-3 Radiated Emissions (RE02) 3ESS-77 3.3.2.0-4 Radiated Emissions (RE02) 2ESS-78 3.3.3.0-1 Conducted Susceptibility 3ESS-79 3.3.3.1.0-1 CS01/CS02 3ESS-80 3.3.3.1.0-2 CS01/CS02 3ESS-81 3.3.3.1.0-3 CS01/CS02 3ESS-82: 3.3.3.1.0-4 CS01/CS02 3


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ESS-83 3.3.3.2.0-1 CS03 3ESS-84 3.3.3.2.0-2 CS03 3ESS-85 3.3.3.2.0-3 CS03 3ESS-86 3.3.3.3.0-1 CS04 3ESS-87 3.3.3.3.0-2 CS04 3ESS-88 3.3.3.3.0-3 CS04 3ESS-89 3.3.3.4.0-1 CS05 3ESS-90 3.3.3.4.0-2 CS05 3ESS-91 3.3.3.4.0-3 CS05 3ESS-92 3.3.3.5.0-1 CS06 3ESS-93 3.3.3.5.0-2 CS06 3ESS-94 3.3.3.5.0-4 CS06 3ESS-95 3.3.4.0-1 Radiated Susceptibility (RS03) 3ESS-96 3.3.4.0-2 Radiated Susceptibility (RS03) 3ESS-239 3.3.4.0-3 Radiated Susceptibility (RS03) 3ESS-240 3.3.4.0-4 Radiated Susceptibility (RS03) 3ESS-241 3.3.4.0-5 Radiated Susceptibility (RS03) 3ESS-242 3.3.4.0-6 Radiated Susceptibility (RS03) 3ESS-97 3.3.5.0-1 Orbiter RF Self-Compatibility 2ESS-98 3.3.5.0-2 Orbiter RF Self-Compatibility 2ESS-99 3.4.0-1 Data and Signal Interfaces 4ESS-100 3.4.1.0-1 Inter-Component Communications 2ESS-101 3.4.1.0-2 Inter-Component Communications 3ESS-102 3.4.1.0-3 Inter-Component Communications 2ESS-103 3.4.1.0-4 Inter-Component Communications 4ESS-104 3.4.1.0-5 Inter-Component Communications 4ESS-105 3.4.1.1.0-1 LRO 1553 Data Bus 4ESS-106 3.4.1.1.1.0-1 LRO 1553 Data Bus Topology 2ESS-107 3.4.1.1.2.0-1 LRO 1553 Remote Terminal Address

Assignments3

ESS-108 3.4.1.1.3.0-1 LRO 1553 Redundancy 2ESS-109 3.4.1.1.4.0-1 LRO 1553 Data Bus Connectors 4ESS-110 3.4.1.1.5.0-1 LRO 1553 Data Bus Cable 2ESS-111 3.4.1.1.6.0-1 LRO 1553 Stub Coupling 2ESS-112 3.4.1.1.6.0-2 LRO 1553 Stub Coupling 4ESS-113 3.4.1.1.6.0-3 LRO 1553 Stub Coupling 4ESS-114 3.4.1.1.6.0-4 LRO 1553 Stub Coupling 2ESS-115 3.4.1.1.6.0-5 LRO 1553 Stub Coupling 4ESS-116 3.4.1.1.6.0-6 LRO 1553 Stub Coupling 4


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ESS-117 3.4.1.3.0-1 Ka-Band Downlink I&Q Data Interfaces 4ESS-118 3.4.1.4.0-1 Analog Telemetry Interfaces 3ESS-119 3.4.1.4.0-2 Analog Telemetry Interfaces 4ESS-120 3.4.1.4.0-3 Analog Telemetry Interfaces 4ESS-121 3.4.1.4.0-4 Analog Telemetry Interfaces 4ESS-122 3.4.1.5.0-1 Discrete Command Interfaces 2ESS-123 3.4.1.5.0-2 Discrete Command Interfaces 4ESS-124 3.4.1.5.0-3 Discrete Command Interfaces 4ESS-125 3.4.1.5.0-4 Discrete Command Interfaces 4ESS-126 3.4.1.5.0-5 Discrete Command Interfaces 4ESS-127 3.4.1.6.0-2 +28 V Discrete Command Interfaces 4ESS-128 3.4.1.6.0-3 +28 V Discrete Command Interfaces 4ESS-129 3.4.1.7.0-1 Pulse per Second (1 pps) 2ESS-130 3.4.1.7.0-2 Pulse per Second (1 pps) 4ESS-131 3.4.1.7.0-3 Pulse per Second (1 pps) 4ESS-132 3.4.1.7.0-4 Pulse per Second (1 pps) 2ESS-133 3.4.1.8.0-1 Other Inter-Component Communications 3ESS-134 3.4.2.0-1 Pyrotechnic and Deployable Actuator

Interfaces4

ESS-135 3.4.2.0-3 Pyrotechnic and Deployable ActuatorInterfaces

2


2


4


4


4


3

ESS-140 3.4.3.0-1 External Interfaces 2ESS-141 3.4.3.1.0-1 Umbilical Interface Connector 2ESS-142 3.4.3.1.0-2 Umbilical Interface Connector 2ESS-143 3.4.3.1.0-3 Umbilical Interface Connector 2ESS-144 3.4.3.1.0-4 Umbilical Interface Connector 2ESS-145 3.4.3.2.0-1 Launch Vehicle Separation Signals 2ESS-146 3.4.3.3.0-1 Component Test Interfaces 4ESS-147 3.4.3.3.0-2 Component Test Interfaces 4ESS-148 3.4.3.3.0-3 Component Test Interfaces 4


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ESS-149 3.4.3.3.0-4 Component Test Interfaces 4ESS-150 3.4.3.3.0-5 Component Test Interfaces 4ESS-151 3.5.0-1 Mulipaction and Corona 4ESS-152 3.5.0-2 Mulipaction and Corona 4ESS-153 3.5.0-3 Mulipaction and Corona 4ESS-154 3.6.0-1 Design For Radiation 4ESS-155 3.7.0-1 Charging and Discharging Requirements 2ESS-156 3.7.1.1.0-1 Differential Charging 2ESS-157 3.7.1.1.0-2 Differential Charging 2ESS-158 3.7.1.2.0-1 Charging Mitigation 2ESS-159 3.7.1.2.0-2 Charging Mitigation 2ESS-160 3.7.1.2.0-3 Charging Mitigation 2ESS-161 3.7.1.2.0-4 Charging Mitigation 2ESS-162 3.7.1.2.0-6 Charging Mitigation 2ESS-163 3.7.1.2.0-7 Charging Mitigation 2ESS-164 3.7.1.2.0-8 Charging Mitigation 2ESS-165 3.7.1.2.0-9 Charging Mitigation 2ESS-166 3.7.1.2.0-10 Charging Mitigation 2ESS-167 3.7.1.2.0-11 Charging Mitigation 2ESS-168 3.7.1.2.0-12 Charging Mitigation 2ESS-169 3.7.1.2.0-13 Charging Mitigation 2ESS-170 3.7.2.0-1 Surface Discharging Protection 2ESS-171 3.7.2.0-2 Surface Discharging Protection 2ESS-172 3.7.2.0-3 Surface Discharging Protection 3ESS-173 3.7.2.0-4 Surface Discharging Protection 3ESS-174 3.7.2.0-5 Surface Discharging Protection 3ESS-175 3.7.2.0-6 Surface Discharging Protection 3ESS-176 3.7.2.0-7 Surface Discharging Protection 2ESS-177 3.7.2.0-8 Surface Discharging Protection 2ESS-178 3.7.2.0-12 Surface Discharging Protection 2ESS-179 3.7.2.0-13 Surface Discharging Protection 2ESS-180 3.7.2.0-14 Surface Discharging Protection 2ESS-181 3.7.2.0-15 Surface Discharging Protection 2ESS-182 3.7.2.0-16 Surface Discharging Protection 3ESS-183 3.7.2.0-18 Surface Discharging Protection 3ESS-184 3.7.2.0-19 Surface Discharging Protection 2ESS-185 3.7.2.0-20 Surface Discharging Protection 2ESS-186 3.7.2.0-21 Surface Discharging Protection 2ESS-187 3.7.2.0-22 Surface Discharging Protection 3ESS-188 3.7.3.0-1 Internal Charging 2


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ESS-189 3.7.3.0-2 Internal Charging 2ESS-190 3.7.3.0-3 Internal Charging 2ESS-191 3.7.3.0-4 Internal Charging 2ESS-192 3.7.3.0-5 Internal Charging 4ESS-193 3.7.3.0-6 Internal Charging 4ESS-194 4.0-1 Harness Requirements 2ESS-195 4.0-2 Harness Requirements 3ESS-196 4.0-3 Harness Requirements 2ESS-197 4.0-4 Harness Requirements 2ESS-198 4.0-5 Harness Requirements 2ESS-199 4.1.0-2 General Harness Guidelines 2ESS-200 4.1.0-3 General Harness Guidelines 2ESS-244 4.1.0-4 General Harness Guidelines 3ESS-201 4.1.0-5 General Harness Guidelines 2ESS-202 4.1.0-6 General Harness Guidelines 2ESS-203 4.1.0-7 General Harness Guidelines 2ESS-204 4.1.0-8 General Harness Guidelines 4ESS-205 4.1.0-9 General Harness Guidelines 2ESS-206 4.1.0-10 General Harness Guidelines 2ESS-207 4.1.1.0-2 Accessibility 4ESS-208 4.1.1.0-7 Accessibility 2ESS-209 4.1.1.0-8 Accessibility 4ESS-210 4.1.1.0-9 Accessibility 4ESS-211 4.1.2.0-1 Harness Shields 4ESS-212 4.1.2.0-2 Harness Shields 4ESS-213 4.1.2.0-3 Harness Shields 4ESS-214 4.1.2.0-4 Harness Shields 4ESS-215 4.1.2.0-6 Harness Shields 4ESS-216 4.1.2.0-8 Harness Shields 4ESS-217 4.1.2.0-9 Harness Shields 4ESS-218 4.1.2.0-10 Harness Shields 4ESS-219 4.1.2.0-11 Harness Shields 4ESS-220 4.1.2.0-12 Harness Shields 4ESS-221 4.1.3.0-1 Component Test Connector Panels 2ESS-222 4.1.3.0-2 Component Test Connector Panels 2ESS-223 4.1.3.1.0-1 Safing Plugs or Arming Plugs 2ESS-224 4.1.3.2.0-1 Fairing Access to Connector Plugs 2ESS-225 4.1.3.3.0-1 Fuse Pigtails 2ESS-226 4.2.0-1 Electrical Materials 3ESS-227 4.2.0-2 Electrical Materials 4


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ESS-228 4.2.1.0-1 Connectors 4ESS-229 4.2.1.0-2 Connectors 4ESS-230 4.2.1.0-3 Connectors 4ESS-231 4.2.1.0-4 Connectors 4ESS-232 4.2.1.0-5 Connectors 4ESS-233 4.2.1.0-6 Connectors 4ESS-234 4.2.1.0-7 Connectors 4


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6.0 ABBREVIATIONS AND ACRONYMS

Abbreviation/Acronym

DEFINITION

A/µs Amps per microsecond(s)AC Alternating CurrentAFSPCMAN Air Force Space Command ManualAl AluminumAm2 Amp meters squaredANSI American National Standards InstituteAWG American Wire GaugeBC Bus ControllerBOB Breakout BoxC CentigradeC&DH Command and Data HandlingCE Conducted EmissionsCCAS Cape Canaveral Air StationCCB Configuration Control BoardCCD Charge-Coupled DeviceCCR Configuration Change RequestCM Configuration Managementcm**2 Centimeters squaredCMO Configuration Management OfficeCRaTER Cosmic Ray Telescope for the Effects of RadiationCS Conducted SusceptibilityCSS Coarse Sun SensordB DecibeldBm Decibel meterdBµA Decibel microampsDC Direct CurrentDLRE Diviner Lunar Radiometer ExperimentECSS European Cooperation for Space StandardizationEEE Electrical, Electronic, ElectromechanicalEIA Electronic Industries AllianceELV Expendable Launch VehicleEMC Electromagnetic CompatibilityEMI Electromagnetic InterferenceESA European Space AgencyESD Electrostatic DischargeESS Electrical Systems SpecificationETU Engineering Test Unit


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GEVS General Environmental Verification Standard for GSFC FlightPrograms and Projects

GHz GigahertzGSE Ground Support EquipmentGSFC Goddard Space Flight CenterH/L Hard-lineHDBK HandbookHGA High Gain AntennaHz HertzI&Q I-channel and Q-channelI&T Integration and TestI/O Input/OutputIC Integrated CircuitICD Interface Control DocumentIMU Inertial Measurement UnitINST InstructionIRW Integrated Reaction WheelLEND Lunar Exploration Neutron DetectorLOLA Lunar Orbiting Laser AltimeterLRO Lunar Reconnaissance OrbiterLVDS Low Voltage Differential SignalingkHz kilohertzkohmKohm kilohmskradKrad kiloradsµV/m Microvolts per meterµs Microsecond(s)mA/µs milliamps per microsecondsMbps Megabits per secondMDM Micro-D MetalMHz megahertzMIL Militarym**2 Meter squaredmm**2 Millimeters squaredMohm MegohmsmsMs millisecondmV millivoltsNASA National Aeronautics and Space AdministrationNHB NASA HandbookNC Normally ClosedNO Normally Openohms/sq. Ohms per squarep-p peak-to-peak


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PCB Printed Circuit BoardPDE Propulsion / Deployment ElectronicspF pico FaradPRT Platinum resistance temperaturePSE Power Subsystems ElectronicsRF Radio FrequencyRLEP Robotic Lunar Exploration ProgramrmsRms Root Mean SquaredRQMT RequirementRS Radiated SusceptibilityRT Remote TerminalSA Solar ArraySBC Single Board ComputerSC SpacecraftSEE Single-Event EffectsSEECA Single-Event Effect Criticality AnalysisSEU Single-Event UpsetSPG Single-Point GroundSPEC SpecificationSSPC Solid-State Power ControlST Star TrackerSTD StandardSTS Space Transportation SystemT/M TelemetryTBD To Be DeterminedTBR To Be ResolvedTIA Telecommunications Industry AssociationV Volt(s)V/m Volts per meterVDC Volts Direct CurrentVpp Volts peak-to-peak


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7.0 APPENDIX B. TRACEABILITY MATRIX

Parent Requirement Requirement Child Requirement 3.1 Power 3.1.1 Power Distribution andSwitching Scheme 3.1.1 Power Distribution andSwitching Scheme 3.1.1 Power Distribution andSwitching Scheme

EPS-55 3.3.5.1.12 Un-SwitchedServices

3.1.2.1 LRO Power Redundancy ESR-2: 3.1.2 Power HarnessRequirements

3.4.1.14 Component PowerRedundancy



3.1.2.2 Power System ElectronicsOutput Switching Profile

3.4.1.4 Power SystemElectronics Output SwitchingProfile



3.1.2.2 Power System ElectronicsOutput Switching Profile





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Parent Requirement Requirement Child Requirement 3.1.2.2 Power System ElectronicsOutput Switching Profile




3.1.2.3 Power Steady State Power EPS-49 3.3.5.1.7 VoltageTransients

3.1.2.3 Power Steady State Power 3.4.1.3 Maximum SustainedInput Current

3.4.1.3 Maximum SustainedInput Current

3.4.1.3 Maximum SustainedInput Current

3.1.2.4 Power Bus Ripple EPS-50 3.3.5.1.8 Output Ripple 3.1.2.4 Power Bus Ripple 3.1.2.5 Single-Event Power BusTransients

3.4.1.9 Single-Event Power BusTransients

EPS-49 3.3.5.1.7 VoltageTransients



3.1.2.6 Over-Current ProtectionDeratings

EPS-54 3.3.5.1.11 HardwareFault Detection

3.1.2.8 Insulation 3.1.3.1 Steady-State Voltage 3.4.1.1 Component Power Input

Voltage

3.4.1.1 Component Power InputVoltage


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Parent Requirement Requirement Child Requirement

3.4.1.1 Component Power InputVoltage

3.1.3.1 Steady-State Voltage 3.1.3.2 Ripple 3.4.1.7 Ripple

3.4.1.7 Ripple

3.4.1.7 Ripple 3.1.3.3 Turn-on Transients (In-Rush Current)



















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Parent Requirement Requirement Child RequirementRush Current) Rush Current)















3.1.3.4 Survival of AnomalousVoltage


EPS-21 3.1.2.2.8 SubsystemVoltage Tolerance Capability




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Parent Requirement Requirement Child Requirement 3.1.3.4 Survival of AnomalousVoltage












3.1.3.5 Operational Bus Transients 3.4.1.8 Operational BusTransients

3.4.1.8 Operational BusTransients

3.4.1.8 Operational BusTransients

3.1.3.6 Turn-Off Transients 3.4.1.10 Turn-Off Transients

3.4.1.10 Turn-Off Transients

3.4.1.10 Turn-Off Transients 3.1.3.7 Turn-Off Protection 3.4.1.11 Turn-Off Protection




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Parent Requirement Requirement Child Requirement 3.1.3.7 Turn-Off Protection 3.4.1.11 Turn-Off Protection


3.4.1.11 Turn-Off Protection 3.1.3.7 Turn-Off Protection 3.4.1.11 Turn-Off Protection


3.4.1.11 Turn-Off Protection 3.1.3.8 Redundant Power Supplies 3.4.1.15 Redundant Power

Supplies

3.4.1.15 Redundant PowerSupplies

3.4.1.15 Redundant PowerSupplies

3.1.3.9 Polarity ReversalProtection




3.1.3.10 Subsystem Over-CurrentProtection




3.2.1 Single-Point Primary PowerGround

ESR-3: 3.1.3 Power IsolationRequirements








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3.4.2.1 Primary Input Isolation

3.4.2.1 Primary Input Isolation

3.4.2.1 Primary Input Isolation 3.2.2 Distributed Signal Ground 3.2.2 Distributed Signal Ground 3.4.2.2 Secondary Ground

3.4.2.2 Secondary Ground

3.4.2.2 Secondary Ground 3.2.2 Distributed Signal Ground 3.4.2.2 Secondary Ground

3.4.2.2 Secondary Ground

3.4.2.2 Secondary Ground 3.2.3 Common Mode Noise 3.4.2.3 Common Mode Noise

3.4.2.3 Common Mode Noise

3.4.2.3 Common Mode Noise 3.2.3 Common Mode Noise ESR-3: 3.1.3 Power Isolation

Requirements 3.2.4 Bonding or Mating 3.2.4 Bonding or Mating ESR-4: 3.2.1 Conductor

Grounding Requirements

ESR-5: 3.2.1 ConductorGrounding Requirements

3.4.2.5 Grounding of Conductors

3.4.2.5 Grounding of Conductors

3.4.2.5 Grounding of Conductors 3.2.4 Bonding or Mating 3.4.2.4 Bonding or Mating

3.4.2.4 Bonding or Mating

3.4.2.4 Bonding or Mating 3.2.4 Bonding or Mating ESR-4: 3.2.1 Conductor



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Parent Requirement Requirement Child RequirementGrounding Requirements




3.4.2.4 Bonding or Mating 3.2.4 Bonding or Mating ESR-4: 3.2.1 Conductor



3.2.4 Bonding or Mating 3.2.4 Bonding or Mating 3.2.4 Bonding or Mating ESR-4: 3.2.1 Conductor





3.4.2.4 Bonding or Mating 3.2.4 Bonding or Mating 3.2.4 Bonding or Mating 3.4.2.7 Connector Grounding

3.4.2.7 Connector Grounding

3.4.2.7 Connector Grounding 3.2.5 Shield Grounds ESR-6: 3.2.2 Shield Grounding

Requirements 3.2.6 Grounding of ExternalOrbiter Surfaces

ESR-7: 3.2.3 External SurfaceRequirements

3.4.2.6 Grounding of ExternalSurfaces


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Parent Requirement Requirement Child Requirement 3.4.2.6 Grounding of ExternalSurfaces

3.4.2.6 Grounding of ExternalSurfaces

3.2.6.1 Thermal Blankets ESR-7: 3.2.3 External SurfaceRequirements

3.2.6.1 Thermal Blankets ESR-7: 3.2.3 External SurfaceRequirements

3.2.6.1 Thermal Blankets 3.2.6.2 Hinges ESR-7: 3.2.3 External Surface

Requirements 3.2.6.3 Solar Array (SA) Panels ESR-7: 3.2.3 External Surface

Requirements 3.2.6.3 Solar Array (SA) Panels ESR-7: 3.2.3 External Surface

Requirements 3.2.6.4 Antennas and AntennaBooms 3.2.6.4 Antennas and AntennaBooms

ESR-6: 3.2.2 Shield GroundingRequirements

3.3 EMI/EMC Requirements 3.3 EMI/EMC Requirements ESR-9: 3.2.5 Electromagnetic

Compatibility Requirements 3.3 EMI/EMC Requirements 3.4.3.1 Test Methods

3.4.3.1 Test Methods

3.4.3.1 Test Methods 3.3 EMI/EMC Requirements 3.3 EMI/EMC Requirements 3.4.3.2.1 CE01/CE03

3.4.3.2.1 CE01/CE03

3.4.3.2.1 CE01/CE03 3.3.1.1 CE01/CE03 3.4.3.2.1 CE01/CE03

3.4.3.2.1 CE01/CE03

3.4.3.2.1 CE01/CE03 3.3.1.1 CE01/CE03 3.4.3.2.1 CE01/CE03

3.4.3.2.1 CE01/CE03


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3.4.3.2.1 CE01/CE03 3.3.1.1 CE01/CE03 3.4.3.2.1 CE01/CE03

3.4.3.2.1 CE01/CE03

3.4.3.2.1 CE01/CE03 3.3.1.1 CE01/CE03 3.3.1.2 CE06 3.4.3.2.2 CE06

3.4.3.2.2 CE06

3.4.3.2.2 CE06 3.3.2 Radiated Emissions (RE02) 3.4.3.3 Radiated Emissions

(RE02)

3.4.3.3 Radiated Emissions(RE02)

3.4.3.3 Radiated Emissions(RE02)

3.3.2 Radiated Emissions (RE02) 3.3.2 Radiated Emissions (RE02) 3.3.2 Radiated Emissions (RE02) 3.3.3 Conducted Susceptibility 3.4.3.4 Conducted Susceptibility

3.4.3.4 Conducted Susceptibility

3.4.3.4 Conducted Susceptibility 3.3.3.1 CS01/CS02 3.4.3.4.1 CS01/CS02

3.4.3.4.1 CS01/CS02

3.4.3.4.1 CS01/CS02 3.3.3.1 CS01/CS02 3.4.3.4.1 CS01/CS02

3.4.3.4.1 CS01/CS02

3.4.3.4.1 CS01/CS02 3.3.3.1 CS01/CS02 3.4.3.4.1 CS01/CS02

3.4.3.4.1 CS01/CS02


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3.4.3.4.1 CS01/CS02 3.3.3.1 CS01/CS02 3.4.3.4.1 CS01/CS02

3.4.3.4.1 CS01/CS02

3.4.3.4.1 CS01/CS02 3.3.3.2 CS03 3.4.3.4.2 CS03

3.4.3.4.2 CS03

3.4.3.4.2 CS03 3.3.3.2 CS03 3.4.3.4.2 CS03

3.4.3.4.2 CS03

3.4.3.4.2 CS03 3.3.3.2 CS03 3.4.3.4.2 CS03

3.4.3.4.2 CS03

3.4.3.4.2 CS03 3.3.3.3 CS04 3.4.3.4.3 CS04

3.4.3.4.3 CS04

3.4.3.4.3 CS04 3.3.3.3 CS04 3.4.3.4.3 CS04

3.4.3.4.3 CS04

3.4.3.4.3 CS04 3.3.3.3 CS04 3.4.3.4.3 CS04

3.4.3.4.3 CS04

3.4.3.4.3 CS04 3.3.3.4 CS05 3.4.3.4.4 CS05

3.4.3.4.4 CS05

3.4.3.4.4 CS05


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Parent Requirement Requirement Child Requirement 3.3.3.4 CS05 3.4.3.4.4 CS05

3.4.3.4.4 CS05

3.4.3.4.4 CS05 3.3.3.4 CS05 3.4.3.4.4 CS05

3.4.3.4.4 CS05

3.4.3.4.4 CS05 3.3.3.5 CS06 3.4.3.4.5 CS06

3.4.3.4.5 CS06

3.4.3.4.5 CS06 3.3.3.5 CS06 3.4.3.4.5 CS06

3.4.3.4.5 CS06

3.4.3.4.5 CS06 3.3.3.5 CS06 3.3.4 Radiated Susceptibility(RS03)

3.4.3.5 Radiated Susceptibility(RS03)



3.3.4 Radiated Susceptibility(RS03)




3.3.4 Radiated Susceptibility(RS03) 3.3.4 Radiated Susceptibility(RS03) 3.3.4 Radiated Susceptibility(RS03)


11-24-05

B-13



Parent Requirement Requirement Child Requirement(RS03) 3.3.4 Radiated Susceptibility(RS03)




3.3.5 Orbiter RF Self-Compatibility 3.3.5 Orbiter RF Self-Compatibility 3.4 Data and Signal Interfaces 3.4.4 Data and Signal Interfaces

3.4.4 Data and Signal Interfaces

3.4.4 Data and Signal Interfaces 3.4.1 Inter-ComponentCommunications 3.4.1 Inter-ComponentCommunications 3.4.1 Inter-ComponentCommunications

ESR-10: 3.3.1 SpaceWireRequirements

3.4.1 Inter-ComponentCommunications

3.4.4.1 Inter-ComponentCommunications



3.4.1 Inter-ComponentCommunications




3.4.1.1 LRO 1553 Data Bus 3.4.4.1.1 1553 Data Bus


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Parent Requirement Requirement Child Requirement 3.4.4.1.1 1553 Data Bus

3.4.4.1.1 1553 Data Bus 3.4.1.1.1 LRO 1553 Data BusTopology

ESR-11: 3.3.2 Requirements

3.4.4.1.1.1 LRO 1553 Data BusTopology



3.4.1.1.2 LRO 1553 RemoteTerminal Address Assignments

3.4.4.1.1.2 LRO 1553 RemoteTerminal Address Assignments



3.4.1.1.3 LRO 1553 Redundancy ESR-11: 3.3.2 Requirements

3.4.4.1.1.3 1553 Redundancy

3.4.4.1.1.3 1553 Redundancy

3.4.4.1.1.3 1553 Redundancy 3.4.1.1.4 LRO 1553 Data BusConnectors


3.4.4.1.1.4 1553 Data BusConnectors



3.4.1.1.5 LRO 1553 Data BusCable


3.4.1.1.6 LRO 1553 Stub Coupling ESR-11: 3.3.2 Requirements


11-24-05

B-15



Parent Requirement Requirement Child Requirement 3.4.4.1.1.5 LRO 1553 StubCoupling

3.4.4.1.1.5 LRO 1553 StubCoupling










3.4.1.1.6 LRO 1553 Stub Coupling ESR-11: 3.3.2 Requirements 3.4.1.1.6 LRO 1553 Stub Coupling ESR-11: 3.3.2 Requirements 3.4.1.1.6 LRO 1553 Stub Coupling 3.4.4.1.1.5 LRO 1553 Stub

Coupling



3.4.1.3 Ka-Band Downlink I&QData Interfaces 3.4.1.4 Analog TelemetryInterfaces


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Parent Requirement Requirement Child Requirement 3.4.1.4 Analog TelemetryInterfaces 3.4.1.4 Analog TelemetryInterfaces 3.4.1.4 Analog TelemetryInterfaces

3.4.4.1.2 Analog TelemetryInterfaces



3.4.1.5 Discrete CommandInterfaces

ESR-12: 3.3.3 +5V DiscreteCommand Requirements

3.4.4.1.3 Discrete CommandInterfaces













11-24-05

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3.4.1.6 +28 V Discrete CommandInterfaces


3.4.4.1.4 +28 V DiscreteCommand Interfaces



3.4.1.6 +28 V Discrete CommandInterfaces




11-24-05

B-18



Parent Requirement Requirement Child Requirement 3.4.4.1.4 +28 V DiscreteCommand Interfaces

3.4.1.7 Pulse per Second (1 pps) 3.4.4.1.5 Pulse per Second (1pps)

3.4.4.1.5 Pulse per Second (1pps)


3.4.1.7 Pulse per Second (1 pps) ESR-14: 3.3.5 Pulse-per-second(pps) Requirements










3.4.1.8 Other Inter-ComponentCommunications

3.4.4.1.6 Other Inter-ComponentCommunications



11-24-05

B-19





3.4.2 Pyrotechnic and DeployableActuator Interfaces 3.4.2 Pyrotechnic and DeployableActuator Interfaces 3.4.2 Pyrotechnic and DeployableActuator Interfaces 3.4.2 Pyrotechnic and DeployableActuator Interfaces 3.4.2 Pyrotechnic and DeployableActuator Interfaces

ESR-18: 3.5 Pyrotechnic SystemRequirements

3.4.2 Pyrotechnic and DeployableActuator Interfaces

ESR-19: 3.5 Pyrotechnic SystemRequirements

3.4.2 Pyrotechnic and DeployableActuator Interfaces 3.4.3 External Interfaces ESR-20: 3.6 External interface

Requirements

EPS-85 3.5.1.1 Launch VehicleUmbilical Interface Requirements

3.4.3.1 Umbilical InterfaceConnector

ESR-20: 3.6 External interfaceRequirements












11-24-05

B-20



Parent Requirement Requirement Child Requirement 3.4.3.2 Launch Vehicle SeparationSignals


3.4.3.3 Component Test Interfaces ESR-20: 3.6 External interfaceRequirements


3.4.4.2 Component TestInterfaces














11-24-05

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5.04 5 3.5 Mulipaction and Corona 3.4.5 Multipaction and Corona

3.4.5 Multipaction and Corona

3.4.5 Multipaction and Corona5.04 5 3.5 Mulipaction and Corona 3.4.5 Multipaction and Corona


3.4.5 Multipaction and Corona 3.5 Mulipaction and Corona 3.4.5 Multipaction and Corona


3.4.5 Multipaction and Corona 3.6 Design For Radiation 3.7 Charging and DischargingRequirements 3.7.1.1 Differential Charging ESR-7: 3.2.3 External Surface

Requirements 3.7.1.1 Differential Charging ESR-7: 3.2.3 External Surface

Requirements 3.7.1.2 Charging Mitigation ESR-7: 3.2.3 External Surface


Requirements

3.4.6 Charging Mitigation


3.4.6 Charging Mitigation 3.7.1.2 Charging Mitigation 3.7.1.2 Charging Mitigation 3.4.6 Charging Mitigation



11-24-05

B-22




3.4.6 Charging Mitigation 3.7.1.2 Charging Mitigation 3.4.6 Charging Mitigation


3.4.6 Charging Mitigation 3.7.1.2 Charging Mitigation 3.4.6 Charging Mitigation


3.4.6 Charging Mitigation 3.7.1.2 Charging Mitigation 3.7.1.2 Charging Mitigation 3.7.1.2 Charging Mitigation 3.7.1.2 Charging Mitigation ESR-7: 3.2.3 External Surface



Requirements 3.7.2 Surface DischargingProtection


3.7.2 Surface DischargingProtection






3.7.2 Surface DischargingProtection 3.7.2 Surface DischargingProtection



3.4.6 Charging Mitigation 3.7.2 Surface DischargingProtection 3.7.2 Surface DischargingProtection





11-24-05

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Parent Requirement Requirement Child Requirement 3.7.2 Surface DischargingProtection
















3.7.2 Surface DischargingProtection 3.7.3 Internal Charging ESR-8: 3.2.4 Internal Charging

Requirements 3.7.3 Internal Charging ESR-8: 3.2.4 Internal Charging




Requirements



3.4.6 Charging Mitigation 3.7.3 Internal Charging ESR-8: 3.2.4 Internal Charging

Requirements





11-24-05

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Parent Requirement Requirement Child Requirement 4 Harness Requirements ESR-22: 3.7 General Harness

Requirements 4 Harness Requirements ESR-22: 3.7 General Harness

Requirements 4 Harness Requirements ESR-22: 3.7 General Harness

Requirements

3.4.7 Harness Requirements


3.4.7 Harness Requirements 4 Harness Requirements ESR-22: 3.7 General Harness

Requirements



3.4.7 Harness Requirements 4 Harness Requirements ESR-22: 3.7 General Harness

Requirements 4.1 General Harness Guidelines ESR-22: 3.7 General Harness


Requirements 4.1 General Harness Guidelines 4.1 General Harness Guidelines ESR-22: 3.7 General Harness




Requirements

3.4.7.1.8 Separation of Signals

3.4.7.1.8 Separation of Signals

3.4.7.1.8 Separation of Signals 4.1 General Harness Guidelines ESR-22: 3.7 General Harness

Requirements


11-24-05

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Parent Requirement Requirement Child RequirementRequirements

3.4.7.1.9 Connector Gender

3.4.7.1.9 Connector Gender

3.4.7.1.9 Connector Gender 4.1 General Harness Guidelines ESR-22: 3.7 General Harness

Requirements 4.1.1 Accessibility ESR-22: 3.7 General Harness

Requirements

3.4.7.1.1 Accessibility

3.4.7.1.1 Accessibility

3.4.7.1.1 Accessibility 4.1.1 Accessibility ESR-22: 3.7 General Harness



Requirements 4.1.2 Harness Shields ESR-22: 3.7 General Harness










Requirements


11-24-05

B-26



Parent Requirement Requirement Child RequirementRequirements

4.1.3 Component Test ConnectorPanels

ESR-22: 3.7 General HarnessRequirements

4.1.3 Component Test ConnectorPanels


4.1.3.1 Safing Plugs or ArmingPlugs


4.1.3.2 Fairing Access toConnector Plugs


4.1.3.3 Fuse Pigtails ESR-22: 3.7 General HarnessRequirements

4.2 Electrical Materials ESR-22: 3.7 General HarnessRequirements

4.2 Electrical Materials ESR-22: 3.7 General HarnessRequirements

2.13 2 4.2.1 Connectors ESR-22: 3.7 General HarnessRequirements

3.4.7.1.2 Sizing

3.4.7.1.2 Sizing

3.4.7.1.2 Sizing 4.2.1 Connectors ESR-22: 3.7 General Harness

Requirements

3.4.7.1.3 Keying

3.4.7.1.3 Keying

3.4.7.1.3 Keying 4.2.1 Connectors ESR-22: 3.7 General Harness

Requirements

3.4.7.1.4 Connector Savers

3.4.7.1.4 Connector Savers

3.4.7.1.4 Connector Savers 4.2.1 Connectors ESR-22: 3.7 General Harness

Requirements


11-24-05

B-27



Parent Requirement Requirement Child Requirement 3.4.7.1.5 Mate/Demate Logs

3.4.7.1.5 Mate/Demate Logs

3.4.7.1.5 Mate/Demate Logs 4.2.1 Connectors ESR-22: 3.7 General Harness

Requirements 4.2.1 Connectors ESR-22: 3.7 General Harness

Requirements

3.4.7.1.6 Connector List

3.4.7.1.6 Connector List

3.4.7.1.6 Connector List 4.2.1 Connectors ESR-22: 3.7 General Harness

Requirements

3.4.7.1.7 EMI Backshell

