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
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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
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Electrical Systems ICD – DRAFT v.2
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Revision DRAFT
Signature Page
Prepared by:
_________
Date
Philip Luers
RLEP Avionics Systems
Code 561
Reviewed by:
Mike Pryzby
LRO Spacecraft Systems Engineer
Swales Aerospace
_________
Date
_________
Date
_________
Date
________
Approved by:
Craig Tooley
LRO Project Manager
Code 430
Concurred by:
Quang H. Nguyen
LRO C&DH Lead Engineer
Code 561
Jim Simpson
LRO ACS Lead Engineer
Code 561
_________
Date
_________
Date
Thomas Spitzer
LRO Power Systems Lead
Code 563
________
Date
________
And a bunch of others
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Electrical Systems ICD – DRAFT v.2
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Revision DRAFT
ROBOTIC LUNAR EXPLORATION PROGRAM
DOCUMENT CHANGE RECORD
REV
LEVEL
DESCRIPTION OF CHANGE
Sheet: 1 of 1
APPROVED
BY
-
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DATE
APPROVED
Electrical Systems ICD – DRAFT v.2
Item
No.
Location
431-ICD-000018
Revision DRAFT
Summary
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,
Table 3-2
Frequency and range of LRO S-band transmitter,
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,
Table 3-2
18 GHz is specified in GEVS for Ku-Band
transmitters. What should be the value for Ka (28?)
10
Section 3.3.4,
Table 3-2
RS level for Ka-band radiation expected on Orbiter
(if any)
11
Section 3.3.4,
Table 3-3
Range of RS03 testing (Delta II S and C band
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.
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Ind./Org.
Due Date
Electrical Systems ICD – DRAFT v.2
Item
No.
16
Location
Section 4.1.2
431-ICD-000018
Revision DRAFT
Summary
External harness shield – type of conductive tape
wrap
Table 1-1. List of TBDs/TBRs
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Revision DRAFT
TABLE OF CONTENTS
1.0
2.0
3.0
4.0
Page
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
Documentation .................................................................................................................12
2.1
Applicable Documents .......................................................................................... 12
2.2
Reference Documents ........................................................................................... 13
2.3
Other Documents .................................................................................................. 14
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
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
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Revision DRAFT
4.2
Electrical Materials ............................................................................................... 38
4.2.1 Connectors .................................................................................................39
4.2.2 Fuses ..........................................................................................................39
Appendix A. Abbreviations and Acronyms ................................................................................ i
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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
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464-ELEC-SPEC-0004
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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,
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464-ELEC-SPEC-0004
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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 highfrequency 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
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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 Workmanship Standard for Surface Mount
Technology
Soldered Electrical Connections
Crimping, Interconnecting Cables, Harness, and Wiring
Electrostatic Discharge Control (excluding electrically
NASA-STD-8739.2
NASA-STD-8739.3
NASA-STD-8739.4
NASA-STD-8739.7
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431-SPEC-000020
initiated explosive devices)
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 2-1. 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
<http://radhome.gsfc.nasa.gov/radhome/papers/seecai.htm>
Table 2-2. Reference Documents
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2.3
OTHER DOCUMENTS
MIL-HDBK-1553
NASA-HDBK-4002
NASA-HDBK-4001
NHB 5300
TP2361
Multiplexed Data Bus Handbook
Avoiding Problems Caused by Spacecraft On-Orbit Internal
Charging Effects
Electrical Grounding Architecture for Unmanned Spacecraft
NASA Workmanship Handbook
Design Guidelines for Assessing and Controlling Spacecraft
Charging
Table 2-3. 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.
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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
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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, nonshielded.
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 steadystate 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.
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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.
Ratio of In-rush/Steady State Current
5
4
1 A/uS rate of
change in the first
10 uSec
In-rush
current to
steady
state
Not to exceed 300 % of
nominal in the first 10 mS
after 10 uS surge
Steady
state
3
2
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.
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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 SinglePoint Ground (SPG) will be located within the PSE. All the primary power returns, solar array
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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.
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Example Component
Example Component
LVDS or
RS-422
LVDS or
RS-422
SpaceWire
Dout
SpaceWire
Din
SpaceWire
Sout
SpaceWire
Sin
SpaceWire
Din
SpaceWire
Dout
SpaceWire
Sin
SpaceWire
Sout
Coax or Triax
Coax or Triax
Example Component on exterior of
Orbiter
Outer Harness Shield 100%
coverage with copper
Spacecraft
Bulkhead
Outer Harness
Shield bonded to
connector backshells
Figure 3-2. LRO Spacecraft Grounding Scheme
Power System
Electronics
Example
Component
PSE Output Module
Input
Filter
Output
Filter
+28V
Output
Current
Shunt
Battery
Return
Solar Array
Return
Single
Point
Ground
Primary Power Twisted with
Primary Power Return
Secondary
Power and
Ground
>1 Mohm isolation from
Primary to Secondary
Chassis
Shunt
Figure 3-3. LRO Spacecraft Power Isolation Scheme
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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 lengthto-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 lowresistance 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.
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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.
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Components
(1)
Instruments
RF
Comp.
CE01
X
X
CE03
X
X
CE06
ICD
Section
GEVS
Section(2)
461C
Section(2)
X
3.3.1.1
2.5.2.1
2
X
3.3.1.1
2.5.2.1
3
X
3.3.1.2
2.5.2.1
4
3.3.2
2.5.2.2
17
Orbiter
RE02
X
X
X
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
3.3.3.6
2.5.3.1e
11
X
3.3.4
2.5.3.2
21
X
3.3.5
CS06
X
X
X
RS03
X
X
X
Self
Compatibility
X
Notes:
(1) X= Applicable
(2) GEVS and MIL-STD-461C sections provided for reference. See applicable ICD section for LRO specific
requirements.
Table 3-1. 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.
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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.
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CE01 (20 Hz - 14 KHz)
CE03 (14 KHz-50 MHz)
140
120
Instrument or Component Level
Test Limits (Differential Mode)
100
dBuA
80
Instrument or Component Level Test
Limits (Common Mode)
60
40
20
Bandwidth (Hz)
0
1.00E+01
1.00E+02
5 Hz
1.00E+03
500 Hz
1.00E+04
1.00E+05
5 KHz
1.00E+06
1.00E+07
FREQUENCY (Hz)
Figure 3-4. Narrowband Conducted Emissions CE01/CE03 Limits
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50 KHz
1.00E+08
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).
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80
30 GHz
70
60
50
14 kHz
Orbiter Levels
dB
uV/
m
40
408–430
MHz
(Delta II)
30
2038–2051
MHz
(LRO S-band)
20
Components and Instruments Levels
10
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
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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.
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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 MILSSTD-461C to the limits specified in GEVS-SE Rev A.
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CS06 Test
Spacecraft Power Bus With Positive Transient Superimposed
56
E
49
Vbus +
Vbus /2
42
35
Volts
28
21
10us
14
7
0
0
20
40
TIME
(microseconds)
60
80
Spacecraft Power Bus with Negative Transient Superimposed
56
49
42
10us
Volts
35
28
21
Vbus /2
14
7
0
0
20
40
TIME
(microseconds)
60
Figure 3-6. CS06 Conducted Susceptibility Test Pulse
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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
14 KHz – 2 GHz
2 GHz – 12 GHz
12 GHz – 18 GHz (TBD)
2.22 +/- TBD GHz
25.5 GHz – 27.5 GHz
Test Level
2 V/m
5 V/m
10 V/m
7 V/m
TBD V/m
Requirement Source
GEVS-SE Rev A
GEVS-SE Rev A
GEVS-SE Rev A
LRO S-Band Transmitter
LRO Ka-Band Indirect
Radiation
Table 3-2. LRO Operational RS Test Limits
Frequency Range
14 kHz - 40 GHz
Test Level
20 V/m
2241.5 +/- TBD MHz
40 V/m
5765 +/- TBD MHz
40 V/m
Requirement Source
Delta II Launch Pad
Environment
Delta II Second Stage Sband T/M
Delta II Second Stage Cband beacon (transmit)
Table 3-3. 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 inorbit 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
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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 transformercoupled 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.
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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.
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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 KaBand 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 S311-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 pyrotechnicactuated 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.
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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.
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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 EEEINST-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.
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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).
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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.
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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.
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Appendix A. Abbreviations and Acronyms
Abbreviation/
Acronym
AWG
C&DH
CCAS
CCD
CMO
DEFINITION
American Wire Gauge
Command and Data Handling
Cape Canaveral Air Station
Charge Coupled Device
Configuration Management Office
ESD
ELV
EMI
ESA
ETU
EWR
GEVS
GSE
GSFC
HGA
ICD
I&T
LRO
LVDS
MAR
MDM
MRD
NHB
PSE
RF
RLEP
rms
SEE
SEECA
SEU
SPG
SSPC
TBD
TBR
VDC
V p-p
Electro-Static Discharge
Expendable Launch Vehicle
Electro-Magnetic Interference
European Space Agency
Engineering Test Unit
Eastern and Western Range
General Environmental Verification Specification
Ground Support Equipment
Goddard Space Flight Center
High Gain Antenna
Interface Control Document
Integration and Test
Lunar Reconnaissance Orbiter
Low Voltage Differential Signaling
Mission Assurance Requirements
Micro-D Metal
Mission Requirements Document
NASA Handbook
Power Systems Electronics
Radio Frequency
Robotic Lunar Exploration Program
Root Mean Squared
Single Event Effects
Single Event Effect Criticality Analysis
Single Event Upset
Single Point Ground
Solid-State Power Control
To Be Determined
To Be Resolved
Volts Direct Current
Volts peak-to-peak
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