Next Generation Space Telescope

Electrical Systems ICD – DRAFT v.2 431-ICD-000018 Revision DRAFT

DRAFT v.2

Robotic Lunar Exploration Program (RLEP) Lunar Reconnaissance Orbiter (LRO)

Electrical Systems Interface Control Document (ICD)

3/8/2005

Goddard Space Flight Center Greenbelt, Maryland

National Aeronautics and Space Administration i CHECK WITH RLEP DATABASE AT: http://TBD1.gsfc.nasa.gov TO VERIFY THAT THIS IS THE CORRECT VERSION PRIOR TO USE. Electrical Systems ICD – DRAFT v.2 431-ICD-000018 Revision DRAFT

CM FOREWORD

This document is a Robotic Lunar Exploration Program (RLEP) Configuration Management (CM)-controlled document. Changes to this document require prior approval of the RLEP Program (or specific project if document only applies at project level) Manager. Proposed changes shall be submitted to the RLEP Configuration Management Office (CMO), along with supportive material justifying the proposed change. Changes to this document will be made by complete revision.

Questions or comments concerning this document should be addressed to:

RLEP Configuration Manager (TBD) RLEP Configuration Management Office Mail Stop 430 Goddard Space Flight Center Greenbelt, Maryland 20771

ii CHECK WITH RLEP DATABASE AT: http://TBD1.gsfc.nasa.gov TO VERIFY THAT THIS IS THE CORRECT VERSION PRIOR TO USE. Electrical Systems ICD – DRAFT v.2 431-ICD-000018 Revision DRAFT Signature Page

Prepared by:

______Philip Luers Date RLEP Avionics Systems Code 561

Reviewed by:

______Mike Pryzby Date Date LRO Spacecraft Systems Engineer Swales Aerospace

Approved by:

______Craig Tooley Date LRO Project Manager Code 430

Concurred by:

______Quang H. Nguyen Date Thomas Spitzer Date LRO C&DH Lead Engineer LRO Power Systems Lead Code 561 Code 563

______Jim Simpson Date And a bunch of others LRO ACS Lead Engineer Code 561

iii CHECK WITH RLEP DATABASE AT: http://TBD.gsfc.nasa.gov TO VERIFY THAT THIS IS THE CORRECT VERSION PRIOR TO USE. Electrical Systems ICD – DRAFT v.2 431-ICD-000018 Revision DRAFT ROBOTIC LUNAR EXPLORATION PROGRAM

DOCUMENT CHANGE RECORD Sheet: 1 of 1

REV APPROVED DATE DESCRIPTION OF CHANGE LEVEL BY APPROVED -

iv CHECK WITH RLEP DATABASE AT: http://TBD.gsfc.nasa.gov TO VERIFY THAT THIS IS THE CORRECT VERSION PRIOR TO USE. Electrical Systems ICD – DRAFT v.2 431-ICD-000018 Revision DRAFT

Item Location Summary Ind./Org. Due Date No.

1 Footer RLEP (or LRO) CM systems URL

2 Table 2-1 Document numbers for LRO Radiation Requirements, LRO Power Distribution Diagram, LRO SpaceWire ICD, LRO 1553 ICD, and LRO PSE Requirements documents

3 Section 3.1.2.6 Derating specification reference

4 Section 3.1.3.3 Verification method of in-rush current. Probably need to require use of SSPC in test circuit for proper line impedance.

5 Section 3.2.3 Measurement reference method for measuring common mode noise

6 Section 3.3.1.1 TBD curves for CS01 and CS02 limits

7 Section 3.3.2, Frequency and range of LRO S-band transmitter, Table 3-2 RE02 and RS03

8 Section 3.3.3.1 Verify CS01/CS02 curves, compare to GEVS-SE and 461C (need to understand slight differences)

9 Section 3.3.4, 18 GHz is specified in GEVS for Ku-Band Table 3-2 transmitters. What should be the value for Ka (28?)

10 Section 3.3.4, RS level for Ka-band radiation expected on Orbiter Table 3-2 (if any)

11 Section 3.3.4, Range of RS03 testing (Delta II S and C band Table 3-3 transmitters)

12 Section 3.3.6 These magnetic requirements were taken from SDO. Need to determine relevancy.

13 Section 3.4.1.4 Inter-Component Communications interfaces (besides 1553, SpaceWire, 422, LVDS, and relay commands) will go here if necessary.

14 Section 3.4.2 Pyrotechnics vs. QWKNUTS

15 Section 4.0 Minimum wire gauge, MDM connectors may want as small as 26 AWG.

CHECK WITH RLEP DATABASE AT: http://TBD.gsfc.nasa.gov TO VERIFY THAT THIS IS THE CORRECT VERSION PRIOR TO USE. Electrical Systems ICD – DRAFT v.2 431-ICD-000018 Revision DRAFT

Item Location Summary Ind./Org. Due Date No.

16 Section 4.1.2 External harness shield – type of conductive tape wrap

Table 1-1. List of TBDs/TBRs

CHECK WITH RLEP DATABASE AT: http://TBD.gsfc.nasa.gov TO VERIFY THAT THIS IS THE CORRECT VERSION PRIOR TO USE. Electrical Systems ICD – DRAFT v.2 431-ICD-000018 Revision DRAFT

TABLE OF CONTENTS

0 Introduction...... 10 1.1 Purpose...... 10 1.2 Electrical System Overview...... 10 1.2.1 Electrical System...... 10 1.2.2 Electrical System Drivers for LRO...... 10 1.3 Definitions and Terminology...... 12 2.0 Documentation...... 12 2.1 Applicable Documents...... 12 2.2 Reference Documents...... 13 2.3 Other Documents...... 14 3.0 LRO Electrical System Requirements...... 14 3.1 Power...... 14 3.1.1 Distribution and Switching Scheme...... 14 3.1.2 LRO Power System Electronics (PSE) Specifications...... 15 3.1.3 User (Subsystem) Specifications...... 16 3.2 System Grounding Requirements...... 18 3.2.1 Single-Point Primary Power Ground...... 18 3.2.2 Distributed Signal Ground...... 19 3.2.3 Common Mode Noise...... 21 3.2.4 Bonding or Mating...... 21 3.2.5 Grounding of External Orbiter Surfaces...... 21 3.3 EMI/EMC Requirements...... 22 3.3.1 Conducted Emissions (CE)...... 23 3.3.2 Radiated Emissions (RE02)...... 26 3.3.3 Conducted Susceptibility (CS)...... 28 3.3.4 Radiated Susceptibility (RS03)...... 31 3.3.5 Orbiter RF Self-Compatibility...... 31 3.3.6 Magnetic Requirements (TBD)...... 31 3.4 Data and Signal Interfaces...... 32 3.4.1 Inter-Component Communications...... 32 3.4.2 Pyrotechnic Interfaces (TBD)...... 34 3.4.3 External Interfaces...... 35 3.5 Mulipaction and Corona...... 36 3.6 Design For Radiation...... 36 4.0 Harness Requirements...... 36 4.1 General Requirements...... 36 4.1.1 Accessibility...... 37 4.1.2 Harness Shields...... 37 4.1.3 Test Connector Panels...... 38 vii

CHECK WITH RLEP DATABASE AT: http://TBD.gsfc.nasa.gov TO VERIFY THAT THIS IS THE CORRECT VERSION PRIOR TO USE. Electrical Systems ICD – DRAFT v.2 431-ICD-000018 Revision DRAFT 4.2 Electrical Materials...... 38 4.2.1 Connectors...... 39 4.2.2 Fuses...... 39 Appendix A. Abbreviations and Acronyms...... i

CHECK WITH RLEP DATABASE AT: http://TBD.gsfc.nasa.gov TO VERIFY THAT THIS IS THE CORRECT VERSION PRIOR TO USE. Electrical Systems ICD – DRAFT v.2 431-ICD-000018 Revision DRAFT LIST OF FIGURES

Figure Page Figure 3-1. SSPC In-rush and Trip Current Limits Curve...... 17 Figure 3-2. LRO Spacecraft Grounding Scheme...... 20 Figure 3-3. LRO Spacecraft Power Isolation Scheme...... 20 Figure 3-4. Narrowband Conducted Emissions CE01/CE03 Limits...... 25 Figure 3-5. RE02 Radiated Narrowband Emission Limits for the Orbiter and Subsystems...... 27 Figure 3-6. CS06 Conducted Susceptibility Test Pulse...... 30 Figure 3-7. Transformer Coupled Stub Diagram (Current Mode)...... 33

LIST OF TABLES

Table Page

Table 1-1. List of TBDs/TBRs...... vi Table 2-1. Applicable Documents...... 13 Table 2-2. Reference Documents...... 13 Table 2-3. Other Documents...... 14 Table 3-1. EMI/EMC Applicability and References...... 23 Table 3-2. LRO Operational RS Test Limits...... 31 Table 3-3. Launch Site/Vehicle RS Test Levels...... 31

CHECK WITH RLEP DATABASE AT: http://TBD.gsfc.nasa.gov/index TO VERIFY THAT THIS IS THE CORRECT VERSION PRIOR TO USE. 464-ELEC-SPEC-0004 Revision A 1.0 INTRODUCTION 1.1 PURPOSE This document provides the electrical and electronic requirements and some guidelines to the subsystem and instrument designers for the Lunar Reconnaissance Orbiter (LRO) mission. The purpose of these requirements and guidelines is to assure reliable and compatible operation of the elements that make up the electrical system, both during ground testing and on orbit, with margin for both expected and worst-case environmental conditions. The LRO components, subsystems, instruments, and Ground Support Equipment (GSE) will adhere to the electrical and electronic requirements specified in the following paragraphs. Component procurement specifications may also refer to sections of this document. GSE requirements will be clarified more specifically in the GSE related documents. Specific details of each subsystem interface will be documented in subsystem specifications, component specifications, and Interface Control Documents (ICDs). 1.2 ELECTRICAL SYSTEM OVERVIEW The electrical system includes electronics and electrical components, interconnect harnessing, structural chassis grounding system, grounding of external coatings and thermal blankets, and elements that provide shielding. The environments that apply to the electrical system include self-generated, conducted, and radiated electromagnetic noise; ground-based electromagnetic emitters; and effects of the on-orbit charging environment. 1.2.1 Electrical System The LRO electrical system includes the electrical elements mounted on the Orbiter that are interconnected to perform their defined functions to meet mission requirements. To the extent that there are challenges in interconnecting the electrical system, whether built in-house or procured from an external vendor, this specification is designed to define those design aspects that are critical to the integrated functioning of the system. This electrical system specification defines the LRO general and specific electrical requirements. The LRO subsystems and instruments will implement these requirements during their design process in order to assure proper system operation and will verify that the requirements are met. 1.2.2 Electrical System Drivers for LRO LRO’s lunar orbit represents a moderate surface and internal charging environment. As a result, the electrical system must be designed carefully with this environment in mind and address integrated solutions for potential risks related to surface and internal charging. a. On-Orbit Charging Environment: While Electro-Static Discharge (ESD) threats may not be totally eliminated, they can be minimized and their effects mitigated through the use of sound design practices. An overview of the planned LRO integrated approach (Section 3.7, Charging and Discharging Requirements) includes limiting the number of discharge sources,

CHECK WITH RLEP DATABASE AT: http://TBD.gsfc.nasa.gov/index TO VERIFY THAT THIS IS THE CORRECT VERSION PRIOR TO USE. 464-ELEC-SPEC-0004 Revision A limiting the size of discharges, implementing shielding between potential sources and potential victims of discharges, and controlling victim susceptibility via filtering and bandwidth control. b. Electromagnetic Interference: The key element of electromagnetic interference (EMI) control is the design and use of the spacecraft structure as a Faraday cage. The Faraday cage concept provides shielding between the noisy outside environment and the electronics and harnessing internal to the spacecraft. Harnesses that transition through the Faraday cage will be grounded, shielded, and/or filtered to maintain the overall shield integrity. Noise sources external to the spacecraft are expected to include unavoidable ESDs, ground-based Radio Frequency (RF) emitters, and self-generated RF from the LRO Ka- and S-band RF systems. c. Instrument Suite: The LRO instrument suite, as well as Star Trackers, will contain instruments with Charge-Coupled Device (CCD) detectors. These detectors are sensitive to common mode noise and can pick up ground noise. To minimize the total common mode noise environment, noise sources will be controlled at the potential sources by limiting alternating current (AC) noise (Section 3.2, System Grounding Requirements). Coupling mechanisms between the potential sources and the CCD victims will be controlled by providing a low AC impedance to chassis ground. d. High Data Rate: The LRO downlink data rate of 125 Mbps requires relatively high- frequency clocks with corresponding fast rise and fall times that may represent a significant noise source. High data quality and integrity requirements, along with short 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 it comes to both the grounding approach and interfacing via impedance matched transmission lines. e. Total Dose Radiation and Single Event Effects (SEEs): The LRO orbit represents a moderate radiation environment. The planned approach for total dose shielding will be shared between the structure and the component chassis to limit the radiation at the part level to approximately 10 Krads behind 100 mils Aluminum (Al) total dose, for the one year mission lifetime, including a factor of 2 margin in order to utilize readily available parts. To limit the total mass required for shielding at the part level, the LRO structure may provide up to 40 mils of Al shielding, depending on component location, with the component chassis providing an additional 100 mils. The remaining difference necessary to meet an individual part’s total dose is expected to be achieved by other structural elements in the line of sight (verified by ray trace analysis) or localized spot shielding. SEEs will be controlled through the use of radiation-hard or radiation-tolerant parts and circuit designs that can tolerate Single Event Upsets (SEUs). Potentially destructive damage will be controlled through the use of radiation-hardened parts, while upsets or soft failures will be controlled through radiation-tolerant parts, circuit design, software design or other mitigation methods. f. Compatible and robust interfaces between electrical system components are key to meeting requirements given the potential noise sources described in the previous paragraphs. Special attention should be given to the design and control interfaces. Most low data rate signal interfaces are expected to use the 1553B bus, which is inherently robust. Non-1553B 11

CHECK WITH RLEP DATABASE AT: http://TBD.gsfc.nasa.gov/index TO VERIFY THAT THIS IS THE CORRECT VERSION PRIOR TO USE. 464-ELEC-SPEC-0004 Revision A interfaces will be carefully controlled and reviewed for interface robustness and compatibility, as well as evaluating the potential of being a noise source or noise susceptible victim. In order to minimize the potential for noise problems, each interface will be expected to control its signal bandwidth only to that which is necessary to perform its function, subject to review on a case-by-case basis by the LRO project. This is expected to encompass rise/fall time control of edges on transmission signals, as well as filtering at the receiving end. g. RF Environment. The LRO RF environment will include self-generated Ka-band radiation emitted by the High-Gain Antenna (HGA). At certain pointing angles, this potential electromagnetic interference (EMI) source may illuminate the instruments and solar arrays. 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 ascent sources, including the ground radars and the launch vehicle. The flight environment is also expected to include ground radars and uplink sources to other spacecraft in the vicinity of LRO. 1.3 DEFINITIONS AND TERMINOLOGY For the purposes of this document, the following definitions are valid: “Shall” is used to indicate a mandatory requirement. “Should” indicates a preferred alternative but is not mandatory. “May” indicates an option. “Will” indicates a statement of fact or intention. 2.0 DOCUMENTATION 2.1 APPLICABLE DOCUMENTS

431-RQMT-000011 LRO Mission Requirements Document (MRD)

430-RQMT-000006 LRO Mission Assurance Requirements (MAR)

PPL-021 NASA GSFC Preferred Parts List (and derating)

GEVS-SE Rev. A GSFC General Environmental Verification Specification for STS & ELV payloads subsystems and components

NASA-STD-8739.1 Workmanship Standard for Staking and Conformal Coating of Printed Wiring Boards and Electronic Assemblies NASA-STD-8739.2 NASA Workmanship Standard for Surface Mount Technology NASA-STD-8739.3 Soldered Electrical Connections NASA-STD-8739.4 Crimping, Interconnecting Cables, Harness, and Wiring NASA-STD-8739.7 Electrostatic Discharge Control (excluding electrically 12

CHECK WITH RLEP DATABASE AT: http://TBD.gsfc.nasa.gov/index TO VERIFY THAT THIS IS THE CORRECT VERSION PRIOR TO USE. 464-ELEC-SPEC-0004 Revision A initiated explosive devices) 431-SPEC-000020 The Radiation Environment for the Lunar Reconnaissance Orbiter (LRO)

430- TBD-TBD-TBD LRO Radiation Requirements

ECSS-E-50-12 ESA SpaceWire Specification

TIA/EIA-644 Electrical Characteristics of Low Voltage Differential Signaling (LVDS) Interface Circuits

TIA/EIA-422 Electrical Characteristics of Balanced Voltage Digital Interface Circuits (formerly known as RS-422)

Table 2-2. Applicable Documents 2.2 REFERENCE DOCUMENTS

MIL-STD-461C EMI/EMC Requirements

MIL-STD-462 Note 6 EMI/EMC Testing Methods

MDC 00H0016 Delta II Payload Planners Guide

MIL-STD-975M NASA EEE Parts

MIL-STD-1553B Multiplexed Data Bus

431-TBD-TBD LRO Power Distribution Diagram

431-TBD-TBD LRO SpaceWire ICD

431-TBD-TBD LRO1553 ICD

431-TBD-TBD LRO Power System Electronics (PSE) Requirements

GSFC-733-HARN-01 Design and Manufacturing for Electrical Harnesses, Rev. A

SEECA Single Event Effect Criticality Analysis

Table 2-3. Reference Documents

CHECK WITH RLEP DATABASE AT: http://TBD.gsfc.nasa.gov/index TO VERIFY THAT THIS IS THE CORRECT VERSION PRIOR TO USE. 464-ELEC-SPEC-0004 Revision A 2.3 OTHER DOCUMENTS

MIL-HDBK-1553 Multiplexed Data Bus Handbook NASA-HDBK-4002 Avoiding Problems Caused by Spacecraft On-Orbit Internal Charging Effects NASA-HDBK-4001 Electrical Grounding Architecture for Unmanned Spacecraft NHB 5300 NASA Workmanship Handbook TP2361 Design Guidelines for Assessing and Controlling Spacecraft Charging Effects

Table 2-4. Other Documents

3.0 LRO ELECTRICAL SYSTEM REQUIREMENTS 3.1 POWER The power subsystem shall supply switched and unswitched unregulated power services per each side (A and B) with the specifications below. 3.1.1 Distribution and Switching Scheme Distributed power architecture shall be used to supply redundant over-current protected power to all the loads. The power subsystem will provide >5 amps high-power services to high power loads and low voltage power converters. +28 V unregulated power will be provided to each subsystem and instrument. No +28 V power returns shall be switched. The Command and Data Handling (C&DH), Power Systems Electronics (PSE), Servo Drive will provide power feeds to its low voltage components. The subsystem low voltage power converters will have the capability of switching power on and off to its loads where necessary by a ground command or onboard computer command. Unswitched power shall be supplied to only those critical functions necessary to receive commands and manage redundancy from the ground. Except for the previously mentioned critical functions, all other functions will be commandable to the off state in order to turn off a failed load or to conserve power in an emergency. The LRO Power Distribution Diagram shows a simplified LRO power distribution diagram with default power bus switch states. All power services shall be over-current protected (based on wire size and service capacity) using either solid-state power switch devices (electronic switches) or fuses. All solid-state power switch devices will act as circuit breakers, tripping off at preset current levels, and they are designed to be re-settable by command. Solid-state power switch devices shall be selected with sufficient margin (greater than 25%) to prevent false tripping during the normal power on/off operation and normal operational transients. All unswitched power services shall use fuses as over-current protection devices. Fuses shall be accessible for testing and verification without dismantling the flight harness or spacecraft structure.

CHECK WITH RLEP DATABASE AT: http://TBD.gsfc.nasa.gov/index TO VERIFY THAT THIS IS THE CORRECT VERSION PRIOR TO USE. 464-ELEC-SPEC-0004 Revision A Any circuit within the power subsystem and distribution hardware, which is connected directly to the unprotected bus, shall be designed and specifically inspected such that no single fault can cause complete loss of power to the spacecraft. 3.1.2 LRO Power System Electronics (PSE) Specifications For internal Power System Requirements, refer to the LRO PSE Requirements Document and the LRO Power Distribution Diagram. 3.1.2.1 Fault Tolerance Redundant power and return wires will be supplied to each user in each power feed. The power subsystem will provide A-side and B-side power feeds to all components with appropriate power isolation design incorporated in the event of power distribution faults. 3.1.2.2 Switching Profile of PSE Output Module When a service is switched on, the voltage shall rise from 0 to the steady-state voltage no faster than 50 µS to reduce the in-rush at the user circuitry. When a service is switched off or trips off due to a fault condition, the voltage shall fall to 0 V no faster than 25 µSec, prohibiting a sharp turn-off from producing an induced EMI emission. 3.1.2.3 Steady State Voltage The bus voltage at the PSE Output Module will have a nominal +28 VDC output voltage with a range from +23 VDC to +34 VDC (inclusive) at the component end of the electrical harness. 3.1.2.4 Spacecraft Power Bus Ripple The bus ripple contributed by PSE will be less than 0.3 Volts peak to peak (p-p). Nominal orbiter level power bus ripple resulting from contributions from all nominal sources will be less than 1.0 Volt p-p over the frequency range of 1.0 Hz to 10 MHz, and 0.5 V p-p over 10 MHz at the power system outputs, under any load condition. Spacecraft components shall meet operational performance requirements in the presence of a 2.8 Volt rms (at 50 watts) ripple superimposed on the steady-state voltage over the frequency range of 30 Hz to 50 kHz, and in the presence of a 1.0 volt rms (at 1 watt) from 50 kHz to 400 MHz superimposed on the steady-state at the power input. This requirement is tested as part of the EMI CS01 and CS02 tests. An instrument shall survive in the presence of a 2.8 Volt rms (at 50 watts) ripple superimposed on the steady-state voltage over the frequency range of 30 Hz to 50 kHz, and in the presence of a 1.0 volt rms (at 1 watt) from 50 kHz to 400 MHz superimposed on the steady-state at the power input. This requirement is tested as part of the EMI CS01 and CS02 tests. 3.1.2.5 Single Event Power Bus Transients Single event power bus transients superimposed on the power buses due to normal subsystem load switching will be limited to +/-3.0 V from the steady-state bus 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

CHECK WITH RLEP DATABASE AT: http://TBD.gsfc.nasa.gov/index TO VERIFY THAT THIS IS THE CORRECT VERSION PRIOR TO USE. 464-ELEC-SPEC-0004 Revision A 10% of its steady state value in 10 ms for a positive or negative load step of 10 amps with a maximum current rate change of 50 mA/us. 3.1.2.6 Over-Current Protection Deratings Solid-state power switching devices used for LRO will be selected based on the load current and the wiring for each load, and derated per the solid-state power switch device I2T trip curve characteristics, which can be found in the parts specification. Fuses shall be derated (per TBD) based on the user inrush and steady-state current for that subsystem. 3.1.2.7 Bus Impedance The impedance of the LRO power bus is a combination of the contributing impedance effects of the power source, distribution harness, switching devices, and connectors. For the purposes of modeling, the power subsystem impedance is approximated as an 80-milliohm resistor in series with 2 micro Henrys of inductance on each power and return line. The impedance of the power distribution harness must be added to this model to approximate the impedance of the power bus as seen at any given component power input. This impedance is component specific but may be approximated for test by 1 to 6 meters of wire, American Wire Gauge (AWG) 14, twisted, non- shielded. 3.1.3 User (Subsystem) Specifications The specifications in this section (3.1.3) apply to all subsystems and components, including instruments, whether built in-house or procured from an external source. 3.1.3.1 Steady-State Voltage The steady-state voltage supplied to components, subsystems and instruments by the LRO power system shall be within a range of +28+/-6 V DC, at their primary power inputs. The nominal power input at the components or instruments will be +28 V DC. 3.1.3.2 Ripple Each instrument shall meet all its performance requirements under the nominal spacecraft level bus ripple condition as specified in Section 3.1.2.4. All instruments and spacecraft subsystems shall survive in the presence of a 2.8 V rms (at 50 Watts) ripple superimposed on the steady-state voltage over the frequency range of 30 Hz to 50 KHz, and in the presence of a 1.0 V rms (at 1 Watt) from 50 KHz to 400 MHz superimposed on the steady- state voltage at the power input. 3.1.3.3 Turn-on Transients (In-Rush Current) The spacecraft PSE utilizes the Solid-State Power Control (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 rise time. The typical turn-on time of a solid-state power switch device is between 50 and 200 µs. The input voltage rises linearly with respect to the turn-on time.

CHECK WITH RLEP DATABASE AT: http://TBD.gsfc.nasa.gov/index TO VERIFY THAT THIS IS THE CORRECT VERSION PRIOR TO USE. 464-ELEC-SPEC-0004 Revision A This delay usually eliminates the need for a component, subsystem or instrument to employ an active means for reducing in-rush current at their power input. The LRO component transient in-rush current shall be within the following limits as listed below and as provided in Figure 3-1: (Verification method TBD) a) Not to exceed a rate of change of 1 A/µs in the first 10 µs. b) A maximum rate of change of 20 mA/µs after the initial 10 µs surge. c) A transient current shall never exceed 300% of the maximum steady-state current in the first 10 ms. d) In-rush current shall be reduced to nominal load at 100 ms after turn-on.

4 In-rush t 1 A/uS rate of Not to exceed 300 % of n current to Steady e

r change in the first nominal in the first 10 mS r steady state u 10 uSec after 10 uS surge

C state

t 3 a t S

y d a e t S / h

s 2 u r - n I

o i t a

R 1 After the first 10 uS, 20 mA/uS Max change rate

0 0.1 1 10 100 1000 10000 100000 1000000 TIME FROM TURN ON (Microseconds)

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

3.1.3.4 Susceptibility to Bus Transients All instruments and spacecraft subsystems shall be designed to not be damaged by any voltage in the range of -1 to +40 V DC for an indefinite time period applied to the power input during anomalistic operations. No flight component will be subjected to these tests. Verification will be by analysis or test on an engineering test unit (ETU) or at a board level only. 17

CHECK WITH RLEP DATABASE AT: http://TBD.gsfc.nasa.gov/index TO VERIFY THAT THIS IS THE CORRECT VERSION PRIOR TO USE. 464-ELEC-SPEC-0004 Revision A All instruments and spacecraft subsystems shall meet performance requirements during the single event transient specified in Section 3.1.2.5 3.1.3.5 Operational Bus Transients The rate of change of any operational current transients shall not exceed 20 mA/µs. The peak current shall not exceed 90% of the SSPC rated current. 3.1.3.6 Turn-Off Transients When the power service is switched off, the peak voltage transients induced on the power service shall not exceed +40 V, nor fall below -1 V. 3.1.3.7 Turn-Off Protection No subsystem or instrument shall be damaged by the unannounced removal of power. Any operations required on a routine basis prior to power turn off shall be listed in that individual subsystem or instrument ICD. 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 or instrument ICD. 3.1.3.8 Redundant Power Supplies Any subsystem or component that includes redundant power supply inputs shall not be damaged by the simultaneous application of power to both interfaces. Any fault within one of the redundant power converters shall not propagate to the other. 3.1.3.9 Polarity Reversal Protection All subsystems shall be designed to prevent damage due to polarity reversal of the input power. Either an in-line diode or a diode cross between power input and return may be used to meet this requirement. 3.1.3.10 Subsystem Over-Current Protection The use of non-resetting over-current protection (i.e. fuses) within the user subsystems shall be prohibited unless a waiver is requested and approved in accordance with the LRO Mission Assurance Requirements (MAR) 430-RQMT -000006. 3.2 SYSTEM GROUNDING REQUIREMENTS LRO will use proven grounding techniques that have been shown to reduce EMI and conducted noise from within the system. LRO will use a hybrid grounding approach with the primary structure serving as the low impedance zero voltage reference. Stray AC noise currents are encouraged to flow through structure in order to reduce common mode voltages. Primary power has a DC single point ground and AC multipoint grounds to chassis. Secondary power and grounds are multipoint grounds to chassis. 3.2.1 Single-Point Primary Power Ground LRO shall implement a single-point grounding scheme for the primary power bus. The Single- Point Ground (SPG) will be located within the PSE. All the primary power returns, solar array 18

CHECK WITH RLEP DATABASE AT: http://TBD.gsfc.nasa.gov/index TO VERIFY THAT THIS IS THE CORRECT VERSION PRIOR TO USE. 464-ELEC-SPEC-0004 Revision A returns, and battery grounds will be tied together at the SPG and connected to the spacecraft structure (or spacecraft chassis ground). At the component primary power interfaces, primary power (28 V DC) and primary power returns shall be isolated from the component chassis by greater than or equal to 1 Megohms DC. 3.2.2 Distributed Signal Ground LRO shall implement distributed signal grounding scheme for secondary power returns or signal grounds. Each subsystem signal ground shall be locally connected to the component chassis with low impedance paths (<= 5 milliohms DC) to minimize common mode noise. The component analog ground shall be isolated from the digital ground (or logic ground) at the DC/DC converter and backplane. If both analog and digital grounds must be referenced for the circuit operation, these grounds should be tied together at one point in the printed circuit board. The secondary return (power, signal, analog, or digital grounds) shall be locally connected to the component chassis with low impedance paths (<= 5 milliohms DC) to minimize stray current. Both secondary power (or signal) inputs and returns shall be isolated from primary power by equal to or greater than 1 Megohms DC. Secondary power returns connected to the component chassis should use the shortest path length possible to minimize the DC resistance and inductance. Wherever possible, this generally implies that the connections between signal grounds and component chassis ground cannot exceed 5 cm in length and should meet the length-to-width requirement of a 5-to-1 ratio. The printed circuit board signal grounds should be connected to the component chassis ground through low-inductance paths (traces), with no part of the signal ground more than 15 cm away from the component chassis ground. In order to keep the inductance at a minimum, the printed circuit board (PCB) ground traces should be kept as wide as possible. This allows for, and encourages, multiple connections between signal grounds and chassis on printed circuit cards within components.

CHECK WITH RLEP DATABASE AT: http://TBD.gsfc.nasa.gov/index TO VERIFY THAT THIS IS THE CORRECT VERSION PRIOR TO USE. 464-ELEC-SPEC-0004 Revision A Example Component Example Component

LVDS or LVDS or RS-422 RS-422

SpaceWire SpaceWire Dout Din

SpaceWire SpaceWire Sout Sin

SpaceWire SpaceWire Din Dout

SpaceWire SpaceWire Sin Sout

Coax or Triax Coax or Triax

Example Component on exterior of Orbiter Outer Harness Shield 100% coverage with copper

Outer Harness Shield bonded to Spacecraft connector backshells Bulkhead

Figure 3-2. LRO Spacecraft Grounding Scheme

Power System Example Electronics Component PSE Output Module Input Output Filter Filter +28V Secondary Power and Ground Output Current Shunt Primary Power Twisted with Single >1 Mohm isolation from Battery Primary Power Return Point Primary to Secondary Return Ground Solar Array Return Chassis Shunt

Figure 3-3. LRO Spacecraft Power Isolation Scheme

CHECK WITH RLEP DATABASE AT: http://TBD.gsfc.nasa.gov/index TO VERIFY THAT THIS IS THE CORRECT VERSION PRIOR TO USE. 464-ELEC-SPEC-0004 Revision A 3.2.3 Common Mode Noise Common mode noise for the primary and secondary power, as well as digital, analog and signal grounds, shall be less than 100 mV p-p (TBD: what to measure, what to reference, and how to test). The LRO structure shall not be used intentionally to carry DC current or common mode current. 3.2.4 Bonding or Mating The LRO structure shall be used as a common reference point for all electronics. All conductors shall be grounded to the spacecraft structure, no floating conductors. 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 and should establish sufficient conductive contact surface area to satisfy electrical grounding and contact requirement. The electrical DC resistance of a mechanical contact between two conductive mating surfaces shall not exceed 5 milliohms DC. 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 Orbiter conductive structure. The electrical DC resistance of the grounding strap as measured at the component chassis tie-point and the Orbiter structure tie-point shall not exceed 5 milliohms DC. All ground straps shall have at least a length- to-width ratio of 5 to 1, be made of copper at least 1 mil thick, and still be flexible enough to allow for bending to occur. The grounding lug location on the component chassis or the tie points in contact with the ground strap shall have a minimum contact area of 80 mm2. This is satisfied by a 1/4" diameter fastener and NAS1149C04-series washer. The electrical DC resistance between any two mated electronic components on the LRO structure, measured at the foot of each component, shall not exceed 10 milliohms DC. Component connectors and backshells shall be electrically mated to the chassis through an electrical resistance not exceeding 5 milliohms DC. 3.2.5 Grounding of External Orbiter Surfaces All external surfaces shall be grounded. 3.2.5.1 Thermal Blankets All layers of thermal blankets shall be redundantly grounded to the spacecraft structure using a low- resistance method as follows: at least two grounding tabs for blankets less than 1m2, and at least two grounding tabs with an additional grounding for every square meter for blankets greater than 1m2. Any point on the blanket shall not be more than 1 meter from the nearest grounding point.

CHECK WITH RLEP DATABASE AT: http://TBD.gsfc.nasa.gov/index TO VERIFY THAT THIS IS THE CORRECT VERSION PRIOR TO USE. 464-ELEC-SPEC-0004 Revision A 3.2.5.2 Solar Array Panels Solar array panels and substrates shall be electrically grounded to the spacecraft structure. Ground straps shall be implemented per Section 3.2.4. 3.2.5.3 Hinges Hinges have poor incidental contact and should not be considered adequate in providing a good ground path. Therefore, a ground strap per Section 3.2.4 shall carry the ground across the joint. 3.2.5.4 Antennas and Antenna Booms The high-gain antenna (HGA) assembly shall employ a grounding scheme to assure HGA metal surfaces and waveguides are grounded directly or indirectly to the spacecraft structure through less than 100 milliohms DC resistance. The HGA boom gimbal rotating joints and deployment hinges should not be considered adequate in providing a good ground path. Therefore, separate ground connections shall be provided in slip-rings on a rotating joint or a dedicated ground strap. The omnidirectional antenna metal surface and cable shields shall be grounded directly or indirectly to the spacecraft structure through less than 5 milliohms DC resistance. 3.3 EMI/EMC REQUIREMENTS Emissions and susceptibility testing shall be performed per this document, which has tailored the GEVS test levels for the LRO mission. The electromagnetic interference/electromagnetic conduction (EMI/EMC) tests required below are meant to cover the LRO mission environments including Orbiter RF self compatibility, launch site, launch pad, launch/ascent, lunar transfer, and lunar orbit. The EMI/EMC test methods shall be per the requirements of MIL-STD-462 Rev.C (Notice 6) unless noted in this specification. Table 3-1 indicates the tests that shall be performed on each component and instrument, and at the Orbiter level.

CHECK WITH RLEP DATABASE AT: http://TBD.gsfc.nasa.gov/index TO VERIFY THAT THIS IS THE CORRECT VERSION PRIOR TO USE. 464-ELEC-SPEC-0004 Revision A

Components Instruments RF Orbiter ICD GEVS 461C (1) Comp. Section Section(2) Section(2)

CE01 X X X 3.3.1.1 2.5.2.1 2

CE03 X X X 3.3.1.1 2.5.2.1 3

CE06 X 3.3.1.2 2.5.2.1 4

RE02 X X X X 3.3.2 2.5.2.2 17

CS01 X X X 3.3.3.1 2.5.3.1a 6

CS02 X X X 3.3.3.2 2.5.3.1a 7

CS03 X 3.3.3.3 2.5.3.1b 8

CS04 X 3.3.3.4 2.5.3.1c 9

CS05 X 3.3.3.5 2.5.3.1d 10

CS06 X X X 3.3.3.6 2.5.3.1e 11

RS03 X X X X 3.3.4 2.5.3.2 21

Self X 3.3.5 Compatibility

Notes:

(1) X= Applicable

(2) GEVS and MIL-STD-461C sections provided for reference. See applicable ICD section for LRO specific requirements. Table 3-5. EMI/EMC Applicability and References Test levels of emissions and susceptibility defined in applicable figures may differ from GEVS-SE, and MIL-STD-461C and MIL-STD-462 and take precedence. All tests shall be performed in ambient with either the component or system in its most sensitive mode for susceptibility testing and in its most noisy mode as appropriate for the EMI emission test. 3.3.1 Conducted Emissions (CE) 3.3.1.1 CE01/CE03 Conducted emissions from subsystem components and instruments shall not exceed the values shown in Figure 3-5 when subjected to CE01 (20 Hz – 14 kHz) and CE03 (14 kHz – 40 MHz) narrowband testing.

CHECK WITH RLEP DATABASE AT: http://TBD.gsfc.nasa.gov/index TO VERIFY THAT THIS IS THE CORRECT VERSION PRIOR TO USE. 464-ELEC-SPEC-0004 Revision A a) CE01/CE03 shall be performed on all +28 V primary power and return lines to each component. b) CE01/CE03 shall be performed in differential and common mode. c) Conducted emissions testing will be performed only at the instrument and component levels. Each subsystem component shall meet the transient current pulse limits, both single event (excluding turn-on) and recurring, as specified in Section 3.1.3. Applicable test parameters and limits are as follows for narrowband conducted emissions: a) Interface lines to be measured are differential mode current lines: +28 V inputs, +28 V input returns. b) Interface lines to be measured are common mode current lines: +28 V power inputs with return including heater circuits. c) Differential mode narrowband test limits are 120 dBuA (1.0 A rms) from 30 Hz to 450 Hz, then decreasing to 50 dBuA (10mA rms) at 20 KHz, then decreasing to 20 dBuA (10uA rms) at 2 MHz, and then continuing at that level to 50 MHz (see Figure 3-5). d) Common mode narrowband test limits are 50 dBuA (0.316 mA rms) from 30 Hz to 20 KHz, then decreasing to 20 dBuA (10uA rms) at 2 MHz, and then continuing at that level to 50 MHz (see Figure 3-5). 3.3.1.2 CE06 All RF receivers and transmitters shall perform the additional CE06 EMI test to the limits contained in MIL-STD-461C.

CHECK WITH RLEP DATABASE AT: http://TBD.gsfc.nasa.gov/index TO VERIFY THAT THIS IS THE CORRECT VERSION PRIOR TO USE. 464-ELEC-SEC-0004 Revision A

CE01 (20 Hz - 14 KHz) CE03 (14 KHz-50 MHz) 140

120 Instrument or Component Level Test Limits (Differential Mode) 100

A Instrument or Component Level Test u Limits (Common Mode) B d 60

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

0 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)

Figure 3-4. Narrowband Conducted Emissions CE01/CE03 Limits

CHECK WITH RLEP DATABASE AT: http://TBD.gsfc.nasa.gov/index TO VERIFY THAT THIS IS THE CORRECT VERSION PRIOR TO USE. 464-ELEC-SEC-0004 Revision A 3.3.2 Radiated Emissions (RE02) Radiated emissions from subsystem components and instruments shall not exceed the values shown in Figure 3-6 when subjected to RE02 narrowband testing. a) Radiated electric field emissions from any Orbiter components or instruments shall not exceed the limits shown see Figure 3-6 (lower line). b) The aggregate radiated emissions from the Orbiter shall not exceed the limits shown in Figure 3-5 (upper line). c) The Orbiter receiver has a center frequency at 2044.250 MHz (TBD), and the notch in Figure 3-5 will protect the receiver with at least 6 MHz on both sides of the center frequency. d) The Delta II maximum allowable payload radiated emissions levels are: 36dB uV/m in the 408 - 430 MHz range, (see Figure 3-6) and 92.4dB uV/m at 5.687 - 5.693 GHz range, which is not shown (amplitude off the scale).

CHECK WITH RLEP DATABASE AT: http://TBD.gsfc.nasa.gov TO VERIFY THAT THIS IS THE CORRECT VERSION PRIOR TO USE. 464-ELEC-SEC-0004 Revision A

30 GHz 70

50 14 kHz

Orbiter Levels 40 m / V u

B 30 d 408–430 2038–2051 MHz MHz (Delta II) (LRO S-band)

20 Components and Instruments Levels

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

Figure 3-5. RE02 Radiated Narrowband Emission Limits for the Orbiter and Subsystems 27

CHECK WITH RLEP DATABASE AT: http://TBD.gsfc.nasa.gov TO VERIFY THAT THIS IS THE CORRECT VERSION PRIOR TO USE. 464-ELEC-SEC-0004 Revision A 3.3.3 Conducted Susceptibility (CS) Undesirable response, malfunction, or degradation of performance shall not be produced in any subsystems or components during CS testing with the tests specified below. For the instruments, performance deviation is acceptable as long as the unit under test shall survive the component CS test. CS testing will not be performed at the Orbiter level. 3.3.3.1 CS01/CS02 The CS01 and CS02 (injection of energy into power lines) shall be performed on all components and instruments that contain the DC/DC converters or power regulation devices. The CS01 and CS02 test limits for the components and instrument level tests shall be 2.8 and 1.0 V rms at the frequency range of 30 Hz to 50 KHz (CS01) and 50 KHz to 400 MHz (CS02), respectively. [TBD – this was taken from SDO, compare this to GEVS-SE Section 2.5.3.1a, p. 2.5-7 and MIL-STD-461C, Section 6, p. 2-5, fig 2-4] 3.3.3.2 CS03 The CS03 (Two Signal Intermodulation) test shall be performed on all RF receiving components and instruments. The CS03 (Two Signal Intermodulation) test performed on all RF receiving equipment shall not cause the RF equipment to exhibit any intermodulation products from two input signals, beyond those permitted in the RF component specification. The CS03 test for RF receiving components shall be conducted per MILS-STD-461C to the limits specified in GEVS-SE Rev A. 3.3.3.3 CS04 The CS04 (Rejection of Undesired Signals) test shall be performed on all RF receiving components and instruments. The CS04 (Rejection of Undesired signals) test for RF receiving components consists of a 0.0 dBm (1 milliwatt) signal applied directly to the receiver input terminals and notched around the receiver input bandwidth at 80.0 dB above its threshold. The input notch center shall be at the receiver-tuned frequency and in the center of the notch. The CS04 test for RF receiving components shall be conducted per MIL-STD-461C to the limits specified in GEVS-SE Rev A 3.3.3.4 CS05 The CS05 (Cross Modulation) test shall be performed on all RF receiving components and instruments. The CS05 (cross-modulation) test performed on all RF receiving equipment shall not cause the RF equipment to exhibit any cross-modulation from two input signals. The CS05 test for RF receiving components shall be conducted per MIL-STD-461C to the limits specified in GEVS-SE Rev A. 3.3.3.5 CS06 The CS06 (Powerline Transient) shall be performed on all components and instruments that contain the DC/DC converters or power regulation devices. 28

The CS06 (Powerline Transient) test consists of both a positive transient test and a negative transient test, having amplitude of +28 V superimposed on the +28 V power bus as in Figure 3-3. This pulse shall be limited to +56 V peak absolute value and 10 µs from 0.5E (42 Volts) to the +28 V steady-state value crossing point (see Figure 3-3). The CS06 test shall be conducted per MILS- STD-461C to the limits specified in GEVS-SE Rev A.

CHECK WITH RLEP DATABASE AT: http://TBD.gsfc.nasa.gov TO VERIFY THAT THIS IS THE CORRECT VERSION PRIOR TO USE. 464-ELEC-SEC-0004 Revision A CS06 Test Spacecraft Power Bus With Positive Transient Superimposed

Vbus + Vbus /2 42

35 Volts 28

21 10us 14

0 0 20 40 60 80 TIME (microseconds) Spacecraft Power Bus with Negative Transient Superimposed

42 10us 35 Volts

Vbus /2 14

0 0 20 40 60 80 TIME (microseconds) Figure 3-6. CS06 Conducted Susceptibility Test Pulse

3.3.4 Radiated Susceptibility (RS03) Undesirable response, malfunction, or degradation of performance shall not be produced during component, instrument, or Orbiter RS testing with the E-field levels shown in Table 3-1. The LRO Expendable Launch Vehicle (ELV) will be serviced and launched from the Cape Canaveral Air Station (CCAS) and can be exposed to the maximum transmitter limits shown in Table 3-3. The Orbiter components and instrument shall survive the radiated susceptibility (RS) test levels of the launch site transmitters. Instrument electronics may be powered off during the RS test at the ELV S-band and C-band transmitter frequencies.

Frequency Range Test Level Requirement Source 14 KHz – 2 GHz 2 V/m GEVS-SE Rev A 2 GHz – 12 GHz 5 V/m GEVS-SE Rev A 12 GHz – 18 GHz (TBD) 10 V/m GEVS-SE Rev A 2.22 +/- TBD GHz 7 V/m LRO S-Band Transmitter 25.5 GHz – 27.5 GHz TBD V/m LRO Ka-Band Indirect Radiation Table 3-6. LRO Operational RS Test Limits

Frequency Range Test Level Requirement Source 14 kHz - 40 GHz 20 V/m Delta II Launch Pad Environment 2241.5 +/- TBD MHz 40 V/m Delta II Second Stage S- band T/M 5765 +/- TBD MHz 40 V/m Delta II Second Stage C- band beacon (transmit)

Table 3-7. Launch Site/Vehicle RS Test Levels

3.3.5 Orbiter RF Self-Compatibility The Orbiter RF self-compatibility test shall be included in the Orbiter-level EMI/EMC test. During this test, the Orbiter shall be configured to a nominal science mode to simulate the in- orbit operation. Ka-band and S-band transmitters will free-radiate from their antennas during this test. 3.3.6 Magnetic Requirements (TBD) Magnetic dipole moment and field intensity due to stray magnetic fields shall be minimized by proven design and manufacturing techniques that are consistent with the GSFC in-house EMI/EMC control process. The maximum DC dipole moment produced by the Orbiter shall not 31 CHECK WITH RLEP DATABASE AT: http://TBD.gsfc.nasa.gov TO VERIFY THAT THIS IS THE CORRECT VERSION PRIOR TO USE. 464-ELEC-SEC-0004 Revision A exceed 15 AM2 or 15,000 pole.cm. Any component or piece part exceeding 3 AM2 or 3,000 pole.cm dipole moment shall be identified and approved for its use. 3.4 DATA AND SIGNAL INTERFACES The presence or absence of any combination of the input signals applied in any sequence shall not cause damage to a component, reduce its life expectancy, or cause any malfunction, whether the component is powered or not. 3.4.1 Inter-Component Communications All signals between boxes shall be controlled to limit signal bandwidth so that no signal should be given more bandwidth than needed to communicate the necessary functions under all expected on-orbit environmental conditions. Subsystems connected to the C&DH subsystems via the 1553 bus shall communicate commands and housekeeping telemetry per the LRO 1553 data bus ICD. Subsystems connected to the C&DH subsystem via the SpaceWire network shall use four twisted pair wires (100+2% ohm cable) with a separate shield around each twisted pair and an overall shield per ESA Space Wire Specification (ECSS-E-50-12). Subsystems connected using RS-422 differential signals shall adhere to the electrical terminations as given in TIA/EIA-422. Subsystems connected using LVDS signals shall adhere to the electrical terminations as given in TIA/EIA-422. 3.4.1.1 LRO 1553 Data Bus The transformer-coupled (long stub) interface shall be implemented for the LRO 1553B data bus, as specified in MIL-STD-1553B. Figure 3-6 shows a typical interface between 1553B bus terminal and the data bus with a coupling transformer (or stub). The LRO unique 1553B bus implementation requirements and clarifications are described below. The remote terminal isolation transformer shall be designed to provide an output signal level of 18 to 27 V p-p at the component output interface. The bus shall have a characteristic impedance of 78 ohms and be terminated at both ends of the bus with 78 ohms termination. The bus coupler transformer shall have a turn ratio of 1:1.41+/-3% (stub to bus). Even though the MIL-STD-1553B standard requirement is for less than 20 feet (6 meters) for a transformer- coupled stub, LRO recommended that the stub design be as short as possible (less than 10 feet is preferred). The 1553B bus cabling and couplers shall have 100% shielding coverage. Concentric twinax (tri-axial) threaded-type female connectors are preferred. Assembled data buses shall be tested for impedance, polarity phasing, signal levels, rise/fall times, and shield termination per MIL-STD-1553B.

32 CHECK WITH RLEP DATABASE AT: http://TBD.gsfc.nasa.gov TO VERIFY THAT THIS IS THE CORRECT VERSION PRIOR TO USE. 464-ELEC-SEC-0004 Revision A

Figure 3-7. Transformer Coupled Stub Diagram (Current Mode)

3.4.1.2 SpaceWire Network Requirements LRO-unique interface requirements for the high-speed instrument data will be specified in the LRO SpaceWire ICD.

3.4.1.3 Ka-Band Downlink I&Q Data Interfaces The Ka-band I-channel and Q-channel (I&Q) downlink data interface from the C&DH to the Ka- Band RF transmitter shall be fabricated per TIA/EIA-644. 3.4.1.4 Other Inter-Component Communications (TBD) 3.4.1.5 Analog Telemetry Interfaces The spacecraft C&DH shall be capable of monitoring the passive analog, active analog and discrete telemetry signals. The standard thermistor part number utilized to generate the passive analog telemetry shall be S- 311-P-18-05A. 5 K Ohms bias resistor shall be used in parallel with the thermistor. The range of the temperature can be sensed by the thermistor with the bias resistor is -30 to +70 degree C as represented by 0 – 5 V telemetry output. To prevent the current switching noise from the passive analog returns, the active analog signal ground or reference shall not be shared with the passive analog returns. 3.4.1.6 Discrete Command Interfaces The +28 V discrete pulse command will be used to actuate the relay or other non-digital actuators. The discrete command pulse circuit shall be designed to minimize coupling of the actuator switching noise into the digital or logic portion of the electronics. Unregulated +28 V from the PSE may be used to generate the pulse command; however, the isolation requirements between primary power and secondary power (logic power) of 1Mohm shall always be maintained. 3.4.2 Pyrotechnic Interfaces (TBD) The pyrotechnic control circuits utilized for the LRO deployable devices and pyrotechnic- actuated valves shall meet the Eastern and Western Range requirements (EWR 127-1). The following LRO-unique requirements are derived from the Eastern Range Safety and mission success requirements:

a) The actuator initiation circuit shall have three independent inhibits, group enable, individual arm, and individual actuation (or fire) functions.

b) The group enable command, which are four separate enables, shall be utilized for enabling one of four actuator circuits : HGA, S/A, pyrotechnic valve NO, and pyrotechnic valve NC.

c) The pyrotechnic actuation circuits shall limit stray currents at the pyrotechnic device to less than 20 dB below the no-fire current level.

d) The actuation device and its control circuits shall be isolated from any external energy source by greater than 20 dB.

e) Each actuator shall have a slow blow device (i.e., fusistor) to protect the spacecraft power bus from pyrotechnic shorts.

3.4.3 External Interfaces LRO shall provide test points and other external interfaces to support the component integration, anomaly investigation, environmental testing, and launch site operations. The LRO spacecraft is to be launched on an expendable launch vehicle. Specific interface requirements between system ground support equipment (GSE) and the launch vehicle, and between each spacecraft and the launch vehicle, will be defined in launch vehicle ICDs. 3.4.3.1 Umbilical Interface Connector The electrical interconnection between Orbiter and the launch vehicle shall be made through the umbilical connectors. There shall be 2 umbilical interface connectors that are capable of carrying up to up to 98 of 22 AWG or larger contacts to support the launch pad operations with the vehicle. The umbilical connector(s) shall, at a minimum, include the following functions: hard-line control of PSE relays, battery charging/ Solar Array Simulator input and monitoring, hard-line communications, health-and-safety critical hard-line (H/L) telemetry data, and launch vehicle interface signals. Umbilical signals shall be protected from ESD threats. Covering the umbilical connector(s) following the separation from the launch vehicle is difficult to achieve, therefore filtering is a preferred method of the ESD protection. 3.4.3.2 Launch Vehicle Separation Signals Three independent separation signals (continuity loops in the launch vehicle interface connectors) shall be required to sense the valid separation from the launch vehicle. These signals shall be utilized to initiate the autonomous solar array deployment and the Reaction Wheel power on following the successful separation. The separation signals will be split between the 2 different umbilical connectors, so that a failure in one of two connections will not result in premature or anomalous separation. Sensing the loss of current flow and a 10 Megohms or greater impedance on all three separation loops shall signify successful separation of the launch vehicle from the Orbiter. 3.4.3.3 Component Test Interfaces Component test points that require access during Orbiter-level testing shall be brought to the Orbiter skin connector: a) All test points shall be protected or isolated from the facility-induced noise, ESD, and GSE malfunction.

35 CHECK WITH RLEP DATABASE AT: http://TBD.gsfc.nasa.gov TO VERIFY THAT THIS IS THE CORRECT VERSION PRIOR TO USE. 464-ELEC-SEC-0004 Revision A b) GSE cable connectors that mate with flight test connectors shall be flight-approved connectors. c) Test connectors shall be capped with flight-approved RF and static control covers when not in use, including in orbit. d) Component power shall not be applied or accessed at or through a test connector. e) Test signals and flight signals should not be located together in the same connector. 3.5 MULIPACTION AND CORONA Damage or measurable degradation due to RF breakdown (corona and arcing) shall be prevented by design. The RF design shall preclude measurable degradation due to multipaction and corona 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 vacuum environment. Components with high-voltage circuits shall be immune to corona and arcing while in a nominal orbital vacuum environment. 3.6 DESIGN FOR RADIATION LRO components shall meet their performance requirements with acceptable degradation due to radiation induced effects. The LRO radiation requirements are contained within the LRO Radiation Requirements Document. 4.0 HARNESS REQUIREMENTS Orbiter harnesses shall satisfy the requirements of this section and the electrical systems requirements described in Section 3 of this specification. Qualified wire, cable, and connector specified in EEE-INST-002 should be used for the spacecraft flight harness or any of the instrument and subsystem harnesses: a) Minimum wire size (max. gauge) for power and heaters should be 22 AWG. b) Minimum wire size (max. gauge) for signals should be 24 AWG (TBD). Wires, connectors, connector contacts, and other harness piece parts shall be derated per EEE- INST-002. 4.1 GENERAL REQUIREMENTS a) The harnesses shall be grouped by common electrical characteristics of the signals carried in the harness. b) 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 signal harnessing, and power and signal wires should not be run in the same bundle.

c) The power and signal shall not share the same component interface and harness connectors. d) The connector half that sources power to another component shall be female (socketed) to protect against inadvertent grounding prior to mating. e) The pyrotechnic harness shall be a twisted, shielded cable without any interruptions in the outer shield. 4.1.1 Accessibility Harnesses should be designed with accessibility and manufacturability in mind: a) 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 extraction tool. b) Any harness cable or connector should not touch any other adjacent connectors or harnesses. c) Electrical boxes should be spaced 6 inches (15 cm) from a panel or structure mounting strut when the box has one or more connectors that face that structure. d) Electrical boxes should be spaced 6 inches (15 cm) apart when one of these boxes has connectors that face the other box. e) Electrical boxes should be spaced 9 inches (23 cm) apart when both of these boxes have connectors that face each other. f) Mechanical support for harnesses shall be designed in accordance with NASA-STD-8739.4. g) Harness splices shall not be allowed without LRO project approval. h) Wires that have tin coating shall not be used for flight due to the possibility of tin whisker growth that could cause a short. Silver is the preferred coating. 4.1.2 Harness Shields Harness shields shall be terminated at the connector backshell at both ends of each harness: a) All shields shall be grounded to the spacecraft structure. b) Wire harness shields shall not carry current by design. c) Outer shields shall not be tied to component connector pins or wires in the cable bundle. d) Inner-shield pig-tail lengths to ground should be minimized, and should not exceed 5 cm. Long pig-tails are the leading cause of EMI noise due to the high resistances in pig-tail wires. e) Shield termination pig-tails shall be bonded to the connector backshell. f) Pig-tails should only pass through a connector pin when used as a special noise reduction technique (as in the case of LVDS [or SpaceWire Cable] inner shields). 37 CHECK WITH RLEP DATABASE AT: http://TBD.gsfc.nasa.gov TO VERIFY THAT THIS IS THE CORRECT VERSION PRIOR TO USE. 464-ELEC-SEC-0004 Revision A g) Shields shall not be daisy-chained from one shield to another. Each internal harness shielded wire should connect its shield directly to the ground point. h) EMI backshells shall be required for all harness connectors. i) All external harnesses shall be shielded with an overall outer bundle shield and terminated with 360° outer shield bond to the backshell. Aluminum tape (TBD) at least 1 mil thick, wrapped with 50% overlap on each 360-degree wrap over the previous wrap shall be used as harness bundle shield. For the external harnesses, the additional shielding may be required to protect from the deep dielectric charging. j) All internal interconnections and cables between components shall be bundle shielded and terminated. 4.1.3 Test Connector Panels Test connector panels shall be located on an external Orbiter surface, and these panels should be accessible during various integration and test (I&T) and launch pad operations. All electrical connectors shall be covered such that they are not exposed to the GEO environments. 4.1.3.1 Safing Plugs or Arming Plugs Safing and arming plugs shall be incorporated in the cable or harness that control ordinance, propulsion valves, electromechanical actuator devices and lasers. 4.1.3.2 Fairing Access to Connector Plugs Access through the launch vehicle fairing shall be provided for arming of electrical or deployment release mechanisms enable plugs. Fairing access shall also allow access to Orbiter for communication testing while attached to the launch vehicle. 4.1.3.3 Fuse Plugs Power harnesses shall use connector plugs that contain fuses (see Section 4.2.2, Fuses) to permit the interruption of the unswitched power bus to a faulty component or harness. The fuse plug will be located at the PSE unswitched output connector. 4.2 ELECTRICAL MATERIALS All subsystem/instrument materials list shall be reviewed for their electrical properties and assessed for its compatibility with electrical systems requirement defined in this specification. Parts with unstable materials (in terms of the LRO orbital environment) that cannot be stabilized through additional processing for the proposed application shall not be used.

4.2.1 Connectors a) All connectors on a component shall be different sizes, pincounts and/or genders to prevent any mismating of harness to component connectors. Wherever possible, keying shall be used. b) Connector savers shall be used during integration and test to minimize wear on connector contacts. c) Connector mate/demate logs shall be used to record mates and demates. d) Instrument developers are expected to provide the mating connectors to their flight components for use on the spacecraft harness. e) All subsystems shall provide a list of connectors to the spacecraft electrical systems engineer prior to connector part procurement. f) All component I/O connector interfaces shall be designed to accommodate the EMI backshell for the D and Micro-D Metal (MDM) connectors. 4.2.2 Fuses The Orbiter shall use the MIL-PRF-23419/12 (FM12A) fuse type, which is approved for space flight use by GSFC. Fuses shall be chosen in order to protect the harness wiring from exceeding its worst-case derated current-carrying capabilities. When redundant fuses are applied in the circuit, a current steering device such as in-line diode or resistor shall be used with the redundant fuse in order to control current flow though the unintended path.

Appendix A. Abbreviations and Acronyms

Abbreviation/ Acronym DEFINITION AWG American Wire Gauge C&DH Command and Data Handling CCAS Cape Canaveral Air Station CCD Charge Coupled Device CMO Configuration Management Office

ESD Electro-Static Discharge ELV Expendable Launch Vehicle EMI Electro-Magnetic Interference ESA European Space Agency ETU Engineering Test Unit EWR Eastern and Western Range GEVS General Environmental Verification Specification GSE Ground Support Equipment GSFC Goddard Space Flight Center HGA High Gain Antenna ICD Interface Control Document I&T Integration and Test LRO Lunar Reconnaissance Orbiter LVDS Low Voltage Differential Signaling MAR Mission Assurance Requirements MDM Micro-D Metal MRD Mission Requirements Document NHB NASA Handbook PSE Power Systems Electronics RF Radio Frequency RLEP Robotic Lunar Exploration Program rms Root Mean Squared SEE Single Event Effects SEECA Single Event Effect Criticality Analysis SEU Single Event Upset SPG Single Point Ground SSPC Solid-State Power Control TBD To Be Determined TBR To Be Resolved VDC Volts Direct Current V p-p Volts peak-to-peak

i CHECK WITH RLEP DATABASE AT: http://TBD.gsfc.nasa.gov/index.jsp TO VERIFY THAT THIS IS THE CORRECT VERSION PRIOR TO USE.