PACE-GNC-DRAFT-SPEC-0063A_02-01-2018.pdf

PACE Spacecraft GPS System Specification PACE-GNC-SPEC-0063 Revision -

Effective Date: February 1, 2018 i Check https://ipdtdms.gsfc nasa.gov to verify that this is the correct version prior to use

400-FORM-0002 (4/16/2014)

PACE Spacecraft GPS System Specification

Signature/Approval Page

Prepared By:

Agbontaen Imasuen

Reviewed By:

John Blackwood

Gary Davis

Bob Schweiss

Beth Weinstein

Jamie Eitnier

Bob Kraeuter

Craig Stevens

Dave Sohl

Jack Sanders

Mike Blau

Marlon Enciso

Tim Johnson

Dan McGuinness

Shavesha Rutledge

Melyane Ortiz-Acosta

Mike Xapsos

Adan Rodriguez-Arroyo

Dan Powers

Nikesha Davis

Approved By:

Andre Dress

Electronic Signatures available online at: https://ipdtdms.gsfc.nasa.gov/ ii

Preface

This document is under Plankton, Aerosol, Cloud, ocean Ecosystem (PACE) Mission configuration control. Changes to this document require prior approval of the PACE

Configuration Control Board (CCB) Chairperson or designee. Proposed changes shall be submitted to the PACE Configuration Management Office (CMO), along with supportive material justifying the proposed change. Changes to this document will be made by complete revision.

iii

Change History Log

Revision Effective Date Description of Changes

(Reference the CCR & Approval Date)

Revision - 12/06/2017 Baseline Release following the approval of PACE-CCR-0188

Revision A 02/01/2018 Updated following the approval of PACE-CCR-0226 iv

Table of TBDs/TBRs/TBSs

Action Item

No.

Location Summary Individual/

Organization

Actionee v

Table of Contents

1.0 Introduction

1.1 General Information

1.2 Scope

2.0 Applicable Documents

3.0 Contract Description

3.1 GPS System Description

3.2 Ground Support Equipment Description

4.0 Functional/Performance Requirements

4.1 GPS System Flight Unit Functional/Performance Requirements

4.1.1 GPS Receiver Flight Unit Functional/Performance Requirements

4.1.1.1 Navigation solution

4.1.1.2 Time Tag Accuracy

4.1.1.3 Position Accuracy

4.1.1.4 Velocity Accuracy

4.1.1.5 1 PPS Accuracy

4.1.1.6 Time-at-Tone Message

4.1.1.7 Raw Data

4.1.1.8 Time to First Solution and Solution Availability

4.1.1.9 Initialization

4.1.1.10 Command and Data Handling Electrical Interface

4.1.1.11 1PPS Electrical Interface

4.1.1.12 Tracking Channels

4.1.1.13 Tracking

4.1.1.14 Navigation Solution Availability

4.1.1.15 Measurement Noise

4.1.2 GPS ANTENNA PERFORMANCE REQUIREMENTS

4.1.2.1 Operating Frequency

4.1.2.2 Polarization

4.1.2.3 Gain

4.1.2.4 Back Radiation

4.1.2.5 Axial Ratio

4.1.2.6 VSWR and Impedance

4.1.2.7 Resistive Losses

4.2 Resource Allocations

4.2.1 Total Mass Allocation

4.2.2 Nominal Power Allocation

4.2.3 Peak Power Allocation

4.2.4 Telemetry

4.3 Power

4.3.1 Primary (Unregulated) Power Input Requirements

4.3.1.1 Operating Voltage Range

4.3.1.2 Abnormal Voltages

4.3.1.3 Voltage Ripple

vi

4.3.1.4 Voltage Transients

4.3.1.5 Sudden Removal of Power

4.3.1.6 Polarity Reversal Protection

4.3.1.7 Over-Current Protection

4.3.1.8 Primary Power Return Ground

4.3.1.9 Power Wiring Redundancy

4.3.2 Load Induced Noise Requirements

4.3.2.1 Turn-on Current Transients

4.3.2.2 Operational Current Transients

4.3.2.3 Repetitive Transients

4.4 Electrical Grounding

4.4.1 Primary Power DC Isolation

4.4.2 Internally Generated Secondary Returns

4.4.3 Internally Generated Secondary to Primary DC Isolation

4.4.4 Mechanical Contact Resistance

4.4.5 Mating Method

4.4.6 Ground Strap, Non-Conductive Surface

4.4.6.1 Ground Lug Contact Area

4.4.7 NA

4.4.8 Connector and Back-Shell DC Resistance

4.5 Signal And Data Interfaces

4.5.1 Active Analog Telemetry

4.5.1.1 Active Analog Telemetry Short Circuit Protected

4.5.1.2 Active Analog Telemetry Source and Load Resistance

4.5.1.3 Separation of Active and Passive Analog Returns

4.5.2 Passive Analog Telemetry

4.5.2.1 Passive Analog Telemetry Interface

4.5.2.2 Passive Linearizing Resistor

4.5.2.3 Thermistor Resistor Conditioning

4.5.3 Digital Telemetry

4.5.3.1 NA

4.5.3.2 Digital Telemetry Voltage Range

4.5.3.3 Digital Telemetry Short Circuit Protected

4.5.3.4 Digital Telemetry Source and Load Resistance

4.5.4 Data and Clock Signals

4.5.4.1 Data and Clock Signal Interface

4.5.4.2 RS-422 Interface

4.5.4.3 NA

4.5.4.4 1553 Bus Interface

4.6 NA

4.7 Command and Data Services

4.7.1 Commands

4.7.2 Data

4.8 Flight Software

5.0 Physical Requirements

vii

5.1 Interface Documentation

5.1.1 Mechanical Interface

5.1.2 Electrical Interface

5.2 Mass Properties

5.2.1 Component Masses

5.2.2 Center of Mass Location

5.2.3 Center of Mass Accuracy

5.2.4 Determination of Moments and Products of Inertia

5.3 Physical Envelope

5.4 Mounting

5.5 Fields of View

5.6 Alignment

5.6.1 Antenna Offsets Knowledge

6.0 Environmental Requirements

6.1 Mechanical Factors of Safety

6.2 Quasi-Static Acceleration

6.3 Frequency Requirement

6.3.1 Fundamental Launch Frequencies

6.4 Vibration

6.4.1 Sinusoidal Vibration

6.4.2 Random Vibration

6.5 Shock

6.6 NA

6.7 Transportation

6.7.1 Transportation Cleanliness

6.8 Pressure

6.8.1 Operating Pressure Range

6.8.2 Maximum Depressurization Rate

6.8.3 Launch Vehicle Environmental Control System Impingement Velocity

6.9 On-Orbit Dynamic Environment

6.9.1.1 Dynamic Linear Acceleration

6.9.1.2 Dynamic Angular Acceleration

6.10 Ground Environments

6.11 Thermal Requirements

6.11.1 Flight Interface Design Temperature Limits

6.12 Charged Particle Radiation Requirements

6.12.1 Definitions

6.12.2 Total Ionizing Dose

6.12.2.1 Minimum TID Tolerance for EEE Parts and Materials

6.12.2.2 ELDR Evaluation

6.12.3 Displacement Damage Dose

6.12.4 Single Event Effects

6.12.4.1 Destructive Events (SELs)

6.12.4.2 Destructive Events (SEBs and SEGRs)

6.12.4.3 Non-Destructive Events (SEUs, SETs, SEFIs, SHEs, and MBUs) ... 33 viii

6.12.5 Charging Environment

6.13 Atomic Oxygen Fluence

6.13.1 Atomic Oxygen Analysis

6.13.2 Atomic Oxygen Testing

7.0 Cleanliness

7.1 Surface Contamination

7.1.1 Surface Contamination Levels at Delivery

7.1.1.1 Particulate Contamination

7.1.1.2 Molecular Contamination – Exposed Surfaces

7.1.1.3 Molecular Contamination – Covered Surfaces

7.1.2 Surface Contamination Generation

7.1.2.1 Particulate Generation

7.1.2.2 Molecular Generation

7.1.2.3 NA

7.1.2.4 Intentional and Unintentional Vents – Filters

7.1.2.5 Intentional Vents – Pressure Buildup Prevention

7.1.3 Cleanability

7.1.3.1 Cleanability – Sensitive Surfaces

7.1.3.2 Cleanability – Sensitive Surface Cleaning Methods

7.2 Electrostatic Cleanliness

7.2.1 Conductive Surface Ground Path

7.2.2 Conductive Surface Resistivity

7.2.3 Closeout of Gaps and Apertures

7.2.3.1 Conductive Tape Surface Resistivity

7.2.3.2 Conductive Tape, Grounding

8.0 Design & Construction Requirements

8.1 Parts, Materials & Processes (PMP)

8.1.1 EEE Parts

8.1.2 Materials

8.1.2.1 Material Conductivity

8.1.2.2 Material Limitations for Debris Casualty Area

8.1.3 Software Assurance

8.2 Electrical

8.2.1 Test Sensors

8.2.2 Interface Requirements

8.2.2.1 Connector Selection

8.2.2.2 Signal Segregation

8.2.2.3 Test and Flight Signal Isolation

8.2.2.4 Test Interfaces

8.2.3 Mitigation of Internal Charging

8.2.3.1 Mitigation Strategies for Internal Charging

8.2.3.2 Floating Conductors

8.2.3.3 Dielectric Structures

8.3 Safety

8.4 Electromagnetic Compatibility

ix

8.4.1 Conducted Emissions

8.4.1.1 Applicability of Conducted Emissions

8.4.1.2 CE101 – Differential Mode Current Emission Limits

8.4.1.3 CE102 – Differential Mode Current Emission Limits

8.4.1.4 Common Mode Current Emissions Limits

8.4.2 Conducted Susceptibility

8.4.2.1 Applicability of Conducted Susceptibility

8.4.2.2 CS101 – Power Leads, 30 Hz to 150 kHz Limit

8.4.2.3 CS114 – Power Leads, 150 kHz to 50 MHz Limit

8.4.2.4 CS106 - Transients

8.4.2.5 CS114 – Power and Signal Cables, Common Mode

8.4.3 Radiated Emissions

8.4.3.1 RE102 - Electric Field Emissions

8.4.4 Radiated Susceptibility

8.4.4.1 RS103 - Electric Field

8.4.5 RF Component Unique Tests

8.4.5.1 CE106 – Antenna Port Emissions

8.4.5.2 CS103 – Antenna Port Intermodulation

8.4.5.3 CS104 – Antenna Port Rejection of Undesired Signals

8.5 Identification and Marking

8.6 Workmanship

8.6.1 Workmanship Standards

8.6.2 Connector

8.6.2.1 GSE Cable Connectors

8.6.2.2 Prevention of Connector Mismating

8.6.2.3 Test Connectors

8.6.2.4 Connector Identification

8.6.2.5 Protection of Unused Test Connectors

8.6.2.6 Connector Savers

8.7 Reliability and Mission Lifetime

8.7.1 Mission Life

8.7.2 Operating Time

8.7.2.1 Total Time

8.7.2.2 Failure Free Time

8.8 Ground Handling

8.8.1 Ground Support Equipment (GSE)

8.8.2 NA

8.8.3 NA

8.8.4 NA

8.8.5 NA

8.8.6 NA

8.8.7 GSE Cleanliness

8.8.8 GSE Bakeout

8.8.9 Test Harness

9.0 Mechanical Design Requirements

x

9.1 Structural Requirements

9.1.1 Component Fatigue

9.1.2 Fracture Control Requirements

9.2 Fastening Systems

9.2.1 Fastener Performance Analysis

9.2.1.1 Factors of Safety

9.2.1.2 Supplemental Factor

9.2.1.3 Ultimate Design Loads

9.2.1.4 Yield Design Loads

9.2.1.5 Design Separation Load

9.2.2 Fastener Locking and Retention

9.2.2.1 Locking Features

9.2.2.2 Verification

9.2.2.3 Locking Features

9.2.2.4 Installation Torque Specification and Control

9.2.3 Fastened Joints Criteria

9.2.3.1 Minimum and Maximum Preload

9.2.3.2 Analysis Addressing Potential Rupture

9.2.3.3 Ultimate Strength Analysis

9.2.3.4 Applied Shear Loading

9.2.3.5 Shear Loading

9.2.3.6 Simultaneous Applied Tensile and Shear Loads

9.2.3.7 Allowable Yield Tensile Load

9.2.3.8 Separation Analysis

9.2.3.9 Seal Analysis

10.0 Logistics

10.1 NA

10.2 Ground Support Equipment

10.3 Transportation Equipment

11.0 Verification Requirements

11.1 Verification Methods

11.1.1 Inspection

11.1.2 Analysis

11.1.3 Test

11.2 Inspection Requirements

11.2.1 Visual Inspection

11.2.2 Physical Measurement

11.2.3 Documentation Search

11.3 Analysis Requirements

11.4 Test Requirements

11.4.1 Definitions

11.4.2 Test Factors

11.4.3 Test Tolerances

11.4.4 Test Restrictions

11.4.4.1 Failure During Tests

xi

11.4.4.2 Modification of Hardware

11.4.4.3 External Adjustment

11.4.4.4 Re-Test Requirements

11.5 Required Tests

11.5.1 Performance Tests

11.5.1.1 Comprehensive Performance Test

11.5.1.2 Limited Performance Test

11.5.1.3 Abbreviated Functional Test

11.5.2 Mass Properties Measurement

11.5.3 Static Loads/Strength Test

11.5.3.1 Sine Burst

11.5.3.2 Static Pull

11.5.4 Sine Sweep Survey

11.5.5 Sine Vibration

11.5.6 Random Vibration

11.5.7 NA

11.5.8 NA

11.5.9 Shock

11.5.10 Thermal Vacuum Bake-out

11.5.11 Thermal Vacuum Test

11.5.11.1 Thermal Vacuum Test Parameters

11.5.11.2 Thermal Vacuum Test Profile

11.5.12 NA

11.5.13 NA

11.5.14 Harness Tests

11.5.14.1 Continuity/Hi-Pot Tests

11.5.14.2 Harness Bake-Out

11.5.15 EMI/EMC Tests

11.5.15.1 Conducted Emissions Tests

11.5.15.2 Conducted Susceptibility Tests

Appendix A Abbreviations and Acronyms xii

List of Figures

Figure Page

Figure 4-1 Inrush Current Limits Figure 6-1 Sine Vibration Environment Figure 6-2 GPS System Random Vibration Environment Figure 6-3 Shock Envelope Figure 6-4 Total Ionizing Dose-Depth Curve (includes x2 margin) Figure 6-5 Integral LET Spectra for Galactic Cosmic Ray Ions Figure 6-6 Solar Heavy Ion Linear Energy Transfer Spectrum Figure 8-1 CE101/CE102 Differential Mode Current Emission Limits Figure 8-2 Common Mode Current Emission Limits

Figure 8-3 Power Lead Conducted Susceptibility (CS101) Voltage Limit Figure 8-4 CS106 Transient Waveform Figure 8-5 Bulk Cable Injection (CS114) Calibrated Test Levels Figure 8-6 Component Level RE102 Radiated Electric Field Emission Limits Figure 11-1 Thermal Vacuum Profile

List of Tables

Table Page

Table 2-1 Applicable Documents

Table 4-1 Measurement Noise Limits

Table 6-1 Factors of Safety

Table 6-2 GPS System Design Limit Loads

Table 6-3 GPS System Sine Vibration Environment

Table 6-4 GPS System Random Vibration Environment, (22.7 kg, or less)

Table 6-5 NA

Table 6-6 GPS System Qualification Level Shock Response Spectrum

Table 6-7 Transportation Loads

Table 6-8 Temperature Limits at Mounting Interface

Table 6-9 Dose (including x2 margin) as a Function of Shielding

Table 6-10 Integral LET Spectra for Galactic Cosmic Ray Ions

Table 6-11 Solar Heavy Ion Linear Energy Transfer Spectrum

Table 8-1 Limited materials for debris casualty area

Table 8-2 Emission and Susceptibility Requirements

Table 8-3 Receiver Notch RE102 Test Limit Levels

Table 8-4 RS103 Radiated Susceptibility Levels

Table 11-1 Test Factors and Durations

Table 11-2 Test Tolerances

Table 11-3 Thermal Vacuum Test Parameters

1.0 INTRODUCTION

1.1 GENERAL INFORMATION

The Plankton, Aerosol, Cloud, ocean Ecosystem (PACE) mission is a strategic climate continuity mission that will extend the high quality ocean ecological, ocean biogeochemical, cloud, and aerosol particle data records begun by NASA in the 1990s. The mission will be capable of collecting radiometric and polarimetric measurements of the ocean and atmosphere, from which these biological, biogeochemical, and physical properties will be determined. PACE data products will not only add to existing critical climate and Earth system records, but also answer new and emerging advanced science questions related to Earth’s changing climate.

1.2 SCOPE

This specification describes the electrical, mechanical, environmental, and verification testing requirements for a space-qualified GPS System for the NASA Goddard Space Flight Center

(GSFC) PACE Mission.

2.0 APPLICABLE DOCUMENTS

The following documents and drawings in effect on the day this specification was signed shall apply to the fabrication and to the electrical, mechanical, and environmental requirements of the

GPS System to the extent specified herein. In the event of conflict between this specification and any referenced document, this specification will govern, with the exception of the PACE

GPS System Statement of Work (PACE-GNC-SOW-0029), in which case the Statement of Work takes precedence.

The following is a list of the applicable specifications and publications.

Table 2-1 Applicable Documents

Document Number Title

PACE-GNC-SOW-0029 PACE Spacecraft GPS System Statement of Work

ANSI/TIA/EIA-422-B Electrical Characteristics of Balanced Voltage

Digital Interface Circuits

NFPA 70 National Fire Protection Association National

Electric Code

NASA-STD-5001B Structural Design And Test Factors Of Safety For

Spaceflight Hardware

NASA-STD-8719.24 NASA Expendable Launch Vehicle Payload Safety

Requirements

NASA-STD-6016 Standard Materials and Processes Requirements for

Spacecraft

NASA-HDBK-7005 Dynamic Environment Criteria

NASA-STD-7001 Payload Vibroacoustic Test Criteria

IEST-STD-CC-1246E Product Cleanliness Levels And Contamination

Control Program

ASTM E-595-07 Standard Test Method for Total Mass Loss and

Collected Volatile Condensable Materials from

Outgassing in a Vacuum Environment

MIL-DTL-5541 Chemical Conversion Coatings on Aluminum and

Aluminum Alloys

MIL-A-8625F Anodic Coatings for Aluminum and Aluminum

Alloys

EEE-INST-002 Instructions for EEE Parts Selection, Screening, Qualification, and Derating

MIL-STD-461F Military Standard, Electromagnetic Emission And

Susceptibility Requirements For The Control Of

Electromagnetic Interference (EMI)

GSFC-STD-7000A General Environmental Verification Standard

(GEVS)

NASA-STD-5019A Fracture Control Requirements for Spaceflight

Hardware

NASA-STD-5020 Requirements for Threaded Fastening Systems in

Spaceflight Hardware

Document Number Title

MIL-STD-1553B Department of Defense Interface Standard for

Digital Time Division Command/Response

Multiplex Data Bus

FAA AC 20-71 Federal Aviation Administration Advisory Circular

(AC) 20-71, “Dual Locking Devices on Fasteners".

NASM 33540 Safety Wiring, Safety Cabling, Cotter Pinning, General Practices for

SAE AS567 Safety Cable, Safety Wire, Key Washers, and Cotter

Pins for Propulsion Systems, General Practices for

Use of

541-WI-5330.1.41 Fastener Locking Using Arathane 5753

MSFC-STD-3029A Guidelines for the Selection of Metallic Materials for

Stress Corrosion Cracking Resistance in Sodium

Chloride Environments

MIL-STD-462, Notice 2 Electromagnetic Interference Characteristics, Measurement of, 1 May 1970

IRN-IS-200H-003 NAVSTAR GPS Space Segment/Navigation User

Segment Interfaces

IS-GPS-200H NAVSTAR GPS Space Segment/Navigation User

Interfaces

3.0 CONTRACT DESCRIPTION

3.1 GPS SYSTEM DESCRIPTION

The GPS System is capable of providing accuracy position, velocity, and time information to the

PACE Guidance, Navigation, and Control (GNC) System in a three-axis stabilized spacecraft operating in a sun-synchronous polar orbit at 676.5 km.

The GPS System consists of a GPS Receiver, GPS Antenna(s), low noise amplifier(s) and filter(s), if required.

The contractor may, for analysis purposes, assume the distance between the GPS receiver and each antenna to be 3 meters or less.

The unit shall compute three-dimensional position, velocity, and time in a user designated coordinate system.

3.2 GROUND SUPPORT EQUIPMENT DESCRIPTION

a. The vendor shall provide attachable hat-couplers for the GPS antennas to support ground testing that provide at least 30dB of attenuation to outside signals.

b. These components need not be space qualified but shall be thermal vacuum chamber compatible. The hat-coupler/antenna combination should be characterized for insertion loss over qualification temperature range at 10° granularity.

4.0 FUNCTIONAL/PERFORMANCE REQUIREMENTS

This section defines the functional and performance requirements for the GPS System as defined in Section 3.1.

4.1 GPS SYSTEM FLIGHT UNIT FUNCTIONAL/PERFORMANCE

REQUIREMENTS

a. The GPS System shall meet performance requirements in this section with GPS Satellites anywhere within the unobstructed FOV defined in Section 5.5 while the PACE spacecraft is nadir pointing in low earth orbit at 676.5 km altitude and 98 degrees inclination.

b. The GPS System shall have an operating frequency range (OFR) covering the GPS L1 band at a minimum.

c. The GPS System shall be immune to a field strength of 10 V/m (RMS) at the GPS antenna from an S-band signal operating between 2.2 and 2.3 GHz.

d. The GPS System shall be immune to a field strength of 1 V/m (RMS) at the GPS antenna from a Ka-band signal operating between 26.0 and 27.0 GHz.

e. The GPS System shall sustain no damage when subjected to an arbitrary input signal of average power +10 dBm, or less.

4.1.1 GPS Receiver Flight Unit Functional/Performance Requirements

The GPS Receiver (GPSR) is defined as the electronics assembly that receives GPS signals and produces the position, velocity, and time outputs to the PACE spacecraft. It does not include the antennas, low noise amplifiers, or any associated external cabling that are detailed in the subsequent sections.

4.1.1.1 Navigation solution

The GPSR shall generate a navigation solution including point position, velocity, and time estimates of the GPS antenna phase center in the ECEF frame at least once every second with an associated ‘solution valid flag’.

4.1.1.2 Time Tag Accuracy

The time provided by the GPSR at which the navigation solution is valid shall be accurate to within 1 microseconds (3-sigma), whenever the solution valid flag is set.

4.1.1.3 Position Accuracy

The GPSR position error shall be less than or equal to 50 meters (3-sigma) in each component of position, in-track, cross-track and radial, whenever the solution valid flag is set.

4.1.1.4 Velocity Accuracy

The GPSR velocity error shall be less than or equal to 0.4 meters/second (3-sigma) in each component of velocity, in-track, cross-track and radial, whenever the solution valid flag is set.

4.1.1.5 1 PPS Accuracy

The GPSR shall provide a 1 pulse per second (1PPS) signal aligned to the GPS whole second with an accuracy less than or equal to 300 nanoseconds (3-sigma), whenever the solution valid flag is set.

4.1.1.6 Time-at-Tone Message

The GPSR shall generate a “time-at-tone” telemetry message associated with the 1PPS signal indicating the accurate whole second of the 1PPS signal and a solution valid flag and be sent between 0.9 and 0.1 seconds before the 1PPS signal.

4.1.1.7 Raw Data

The GPSR shall provide raw pseudorange, pseudorange rate, and carrier phase telemetry as well as the antenna ID and an unbiased estimate of received C/N0 associated with the measurements.

4.1.1.8 Time to First Solution and Solution Availability

The GPSR shall produce valid navigation solutions compliant with the position, velocity, and time performance requirements of this document within 5 minutes of receipt of power, and any time thereafter while power is maintained, and when the receiver is tracking at least four

NAVSTAR GPS constellation space vehicles (SVs).

4.1.1.9 Initialization

The GPSR shall require no external orbit knowledge to acquire, track, and produce position, velocity, and time solutions.

4.1.1.10 Command and Data Handling Electrical Interface

The command and data interface, operating at a 1-Hz polling frequency, shall be either a MIL-

STD-1553 interface or an RS-422 UART interface.

4.1.1.11 1PPS Electrical Interface

The GPSR shall provide a differential 1 PPS (1 pulse per second) output to spacecraft using an

RS-422 interface.

4.1.1.12 Tracking Channels

The GPSR shall have at least 12 independent baseband digital tracking channels.

4.1.1.13 Tracking

The GPSR shall track a minimum of 80% of the instantaneous number of signals providing at least -158.5 dBW of signal power to a reference, ideal 3 dBi linearly polarized GPS antenna, after the receiver has been powered for at least 5 minutes.

4.1.1.14 Navigation Solution Availability

The navigation solution valid flag shall be set to ‘valid’ at least 99% of time after the receiver has been powered for at least 5 minutes and is tracking at least four NAVSTAR GPS SVs.

4.1.1.15 Measurement Noise

The GPSR RMS measurement noise level on the raw pseudorange observables as a function of receiver estimated C/No shall be below the limits listed in the Table 4-1.

Table 4-1 Measurement Noise Limits

C/No Level (dB-Hz)

RMS

Measurement

Noise (m)

40 4

45 3

50 2

4.1.2 GPS ANTENNA PERFORMANCE REQUIREMENTS

The GPS Antenna(s) is a component of the PACE GPS System.

4.1.2.1 Operating Frequency

The GPS antenna(s) shall have an operating frequency range (OFR) covering at least the GPS L1 band.

4.1.2.2 Polarization

The GPS antenna(s) shall have right hand circular polarization (RHCP).

4.1.2.3 Gain

The GPS antenna(s) shall have sufficient gain to close the RF link to unobstructed GPS Satellites anywhere within the unobstructed FOV defined in Section 5.5

4.1.2.4 Back Radiation

The GPS antenna(s) shall have a minimum boresight to back ratio ≥ 11 dB, at all angles greater than 120° from boresight in the OFR.

4.1.2.5 Axial Ratio

The GPS antenna shall have an axial ratio of less than 7dB within 90° of boresight in the OFR, and less than 2dB at boresight.

4.1.2.6 VSWR and Impedance

The VSWR (Voltage Standing Wave Ratio), as measured at the GPS antenna RF interface shall be a maximum of 1.5:1 in the OFR, with respect to the nominal 50 Ohm impedance.

4.1.2.7 Resistive Losses

The GPS antenna resistive losses shall be ≤ 0.1 dB over the OFR, to minimize system noise.

4.2 RESOURCE ALLOCATIONS

4.2.1 Total Mass Allocation

The GPS System shall have a mass of less than or equal to 4 kg.

4.2.2 Nominal Power Allocation

The GPS Receiver shall have a nominal (continuous) power consumption of less than or equal to

9 W for a duration of 100 minutes.

4.2.3 Peak Power Allocation

The GPS Receiver shall have a peak (instantaneous) power consumption of less than or equal to

12 W for a duration of 60 seconds.

4.2.4 Telemetry

The GPS Receiver shall generate housekeeping telemetry of less than or equal to TBR bps.

4.3 POWER

4.3.1 Primary (Unregulated) Power Input Requirements

4.3.1.1 Operating Voltage Range

The GPS Receiver shall be designed to operate over the bus voltage range of +27 to +35 VDC at their primary power inputs during all normal mission phases and for all expected load conditions

(except when turned off).

4.3.1.2 Abnormal Voltages

The GPS Receiver shall survive without performance degradation after indefinite exposure to an anomalous voltage range of 0 to +40 VDC. This requirement is to be verified by analysis or test of a non-flight unit.

4.3.1.3 Voltage Ripple

The GPS Receiver shall be designed to meet its performance requirements in the presence of ripple voltages specified in the CS101 and CS114 requirements in paragraphs 8.4.2.2 and 8.4.2.3, respectively.

4.3.1.4 Voltage Transients

The GPS Receiver shall be designed to meet its performance requirements in the presence of transient voltages specified in the CS106 requirements in paragraph 8.4.2.4.

4.3.1.5 Sudden Removal of Power

The GPS Receiver shall meet its performance requirements without degradation after exposure to an abrupt, unannounced removal of power.

4.3.1.6 Polarity Reversal Protection

The GPS Receiver should have built-in protection to prevent damage due to polarity reversal at the power inputs (up to -35V), where feasible.

4.3.1.7 Over-Current Protection

The GPS Receiver should not use non-resetting over-current protection (i.e., fuses) internal to the unit, but if they are required by the contractor to isolate an internal fault, prior approval shall be obtained from the PACE Contracting Officer’s Representative (COR).

4.3.1.8 Primary Power Return Ground

The GPS Receiver shall provide a dedicated Primary Power return in the same connector as the primary power.

4.3.1.9 Power Wiring Redundancy

The GPS Receiver should utilize redundant power wiring, if possible.

4.3.2 Load Induced Noise Requirements

4.3.2.1 Turn-on Current Transients

With an input voltage (0 –35V) rise time no faster than 500us, the inrush current shall not exceed the no-trip limit in Figure 4-1 below for durations greater than 100 microseconds.

4.3.2.1.1 Turn-on Current Transients (less than 100 microseconds)

For durations less than 100 microseconds, inrush current shall not exceed 600% of switch current rating of 2 amps.

Figure 4-1 Inrush Current Limits

The compliance of this requirement can be demonstrated using a solid-state power switch with the similar turn on characteristics as the flight power control switch.

4.3.2.2 Operational Current Transients

a. The GPS Receiver operational current transients shall be less than or equal to 5A above the steady state current for a period of less than 1 millisecond.

b. The rate of change of the current shall be less than or equal to 20 milliamperes/microsecond.

4.3.2.3 Repetitive Transients

Any repetitive current transients, in the frequency range between 1 Hz to 200 Hz, shall not exceed 0.2 A peak-to-peak.

This measurement will be performed with an oscilloscope current probe in time domain.

4.4 ELECTRICAL GROUNDING

4.4.1 Primary Power DC Isolation

At the GPS Receiver primary power interfaces, primary power and primary power returns shall be isolated from signal grounds and from the component chassis by a DC resistance of greater than or equal to 1 MΩ.

4.4.2 Internally Generated Secondary Returns

Secondary returns (power and signal grounds) shall be referenced to the component chassis ground by connecting them at one or more places.

4.4.3 Internally Generated Secondary to Primary DC Isolation

Secondary power (or signal) inputs shall be isolated from primary power by a DC resistance of greater than 1 MΩ.

4.4.4 Mechanical Contact Resistance

The DC resistance of the mechanical contact between two conductive mating surfaces (internal to the component, and at the spacecraft interface) shall be less than or equal to 10 mΩ DC resistance.

4.4.5 Mating Method

The primary mating method for a PACE component is metal-to-metal contact between component mounting feet (or base plate) and the spacecraft structure. When the use of this method is not possible as determined by the Mechanical Interface Control Document (MICD), the use of a ground strap is necessary.

4.4.6 Ground Strap, Non-Conductive Surface

The GPS Receiver shall contain provisions for a grounding strap to be attached from the component chassis for connection to the spacecraft conductive structure

4.4.6.1 Ground Lug Contact Area

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. The ground lug contact area must remain free of any material finish that may affect the reliability of the ground connection and be shown in the MICD.

4.4.7 NA

4.4.8 Connector and Back-Shell DC Resistance

Component connectors and back-shells of components that contain electrical circuits with secondary ground tied to chassis ground, shall be electrically connected to chassis with a DC resistance less than or equal to 10 mΩ.

4.5 SIGNAL AND DATA INTERFACES

4.5.1 Active Analog Telemetry

These signals do not require conditioning and will meet the following requirements:

4.5.1.1 Active Analog Telemetry Short Circuit Protected

The sources of active analog signals shall be short circuit protected.

4.5.1.2 Active Analog Telemetry Source and Load Resistance

The source and load resistances (at the source and destination, respectively) shall be documented in the component Electrical Interface Control Document (EICD) such that the overall system compatibility of the telemetry signal is reviewed and controlled.

4.5.1.3 Separation of Active and Passive Analog Returns

External to the box, active analog return currents shall not share the same wire with passive analog return currents

4.5.2 Passive Analog Telemetry

These include temperature sensors such as thermistors and platinum resistance thermometers

(PRTs). Requirements are as follows:

4.5.2.1 Passive Analog Telemetry Interface

If the GPS System is supplied with passive internal temperature sensors, they shall be #311P18-

02A101 thermistor parts, with a resistance of 2252 ohms at 25 degrees C.

4.5.2.2 Passive Linearizing Resistor

The linearizing resistor for the thermistors, if required, will be located with the temperature sensor conditioning circuit in the Spacecraft avionics.

4.5.2.3 Thermistor Resistor Conditioning

Temperature sensors will be conditioned to provide a voltage range which represents a temperature range that includes a margin of greater than or equal to 5 degrees C beyond the unit qualification or non-operational survival limit, whichever is worse.

4.5.3 Digital Telemetry

4.5.3.1 NA

4.5.3.2 Digital Telemetry Voltage Range

The digital telemetry voltage range shall be within 0 to +5 VDC.

4.5.3.3 Digital Telemetry Short Circuit Protected

The sources of digital telemetry signals shall be short circuit protected.

4.5.3.4 Digital Telemetry Source and Load Resistance

The source and load resistances (at the source and destination, respectively) for digital telemetry signals shall be such that the intended logic levels from the source are not violated.

4.5.4 Data and Clock Signals

4.5.4.1 Data and Clock Signal Interface

a. The GPSR shall provide the following digital telemetry signals to the spacecraft at a 1 Hz frequency.

1. The position information shall be provided as a vector in Earth Centered Earth Fixed

(ECEF) coordinates.

2. The velocity information shall be provided as a vector in ECEF coordinates.

3. The GPS time corresponding to the supplied position and velocity information.

b. The command and telemetry interface to the GPSR shall be either digital differential RS-

422 (ANSI/TIA/EIA-422-B) or a MIL-STD-1553B interface that accommodates the following signals:

GPSR 1 PPS (RS-422)

GPSR Transmit/Receive (RS-422 UART or MIL-STD-1553B)

4.5.4.2 RS-422 Interface

The RS-422 interface circuit and line termination shall be documented in the GPS Receiver

Electrical ICD, using ANSI/TIA/EIA-422-B, Electrical Characteristics of Balanced Voltage

Digital Interface Circuits as a guide.

Signals that use RS-422 interface should be designed to minimize the chassis ground current loop. The chassis current due to the line imbalance of the 3.3V CMOS interface should be less than 30 micro-amps.

4.5.4.3 NA

4.5.4.4 1553 Bus Interface

Any 1553 Bus interface shall be designed per MIL-STD-1553B, Digital Time Division

Command/Response Multiplex Data Bus.

4.6 NA

4.7 COMMAND AND DATA SERVICES

4.7.1 Commands

The GPS system shall accommodate the following set of commands or their equivalents, as a minimum.

a. Reset: This command shall cause the software to reset from the beginning.

b. Ground Test Mode: This command shall cause the GPS Receiver to acquire and track

GPS satellites and to report a position fix while on the ground.

c. Self-Test: This command shall cause the GPS Receiver to perform tests of internal components.

d. Start Acquisition: This command shall cause the GPS Receiver to start GPS satellite acquisition.

e. Set Mask Angle: This command shall cause the GPS Receiver to exclude the GPS satellites within the mask angle from the acquisition algorithm.

f. Parameter Update: This command shall cause the GPS Receiver to update default parameter values that modify GPS Receiver operation.

4.7.2 Data

The GPS system shall provide all the data listed below, once per second, as a minimum.

a. Position: The position shall be provided as a vector in Earth Centered Earth Fixed

(ECEF) Coordinates

b. Velocity: The velocity shall be provided as a vector in ECEF Coordinates

c. Time: The GPS Time (elapsed time from GPS epoch) at which the measured position and velocity apply shall be provided.

d. Data Validity Flag: A flag indicating the availability of a sufficient number of GPS satellites to produce a 3D fix shall be provided.

e. Dilution of Precision (DOP): A figure for a selected DOP figure of merit shall be provided. The selection set shall include Position Dilution of Precision and Time

Dilution of Precision at a minimum.

f. Number and Identification of GPS tracked: The number of GPS satellite tracked and their identification shall be provided.

g. Pseudo-Range and Carrier Phase: The pseudo-range and carrier phase from each track

GPS satellite shall be accessible.

h. Sensor Status: The GPS system shall provide health and status information, including the Flight Software revision number.

4.8 FLIGHT SOFTWARE

The contractor will provide the flight software within the GPS Receiver System.

a. Housekeeping Status: The GPS System flight software shall provide operational and health and safety monitoring. This includes memory checksum and watchdog timer servicing.

b. Software Version Control: The GPS Receiver System flight software at launch shall be uniquely identifiable (i.e., be able to verify SW version).

c. Software Updates: The GPS Receiver System shall have the capability to update flight software during integration and test phases as well as during flight operations.

d. Software Table Updates: The GPS Receiver System shall have the capability to update software tables during integration and test phases as well as during flight operations.

e. Memory contents shall be downloadable to the ground to verify configurable parameters, tables, etc. There must be methods to verify settings (e.g., via a checksum of that region or via dump of the configurable items).

f. Non-Volatile Memory Storage of Flight Code: Flight software executable code shall reside in non-volatile storage devices that can be reprogrammed during flight operations.

g. Boot Code: Boot (Start-up or Kernel) code shall reside in non-volatile storage devices that are not alterable in flight.

5.0 PHYSICAL REQUIREMENTS

5.1 INTERFACE DOCUMENTATION

The contractor shall use metric units when interfacing with NASA GSFC including any drawings, documents, and models, except for the following cases:

Heritage Hardware: Hardware that has been previously qualified, or of similar design heritage, may be specified in English units where use of metric equivalents would lead to additional cost to the program.

Fasteners: Although bolt patterns will be defined using metric dimensioning, use of

English fasteners (with hole dimensioning and tolerancing) is permitted.

Angular Measurement: Angular measurement may be expressed in degree of arc or in an appropriate subdivision of degree of arc such as second of arc (arc-sec) when advantageous to application.

5.1.1 Mechanical Interface

The mounting interface shall be defined in the Mechanical Interface Control Drawing (MICD), which will be developed between the contractor and NASA GSFC.

5.1.2 Electrical Interface

The electrical interface shall be defined in the Electrical Interface Control Document (EICD), which will be developed between the contractor and NASA GSFC.

5.2 MASS PROPERTIES

5.2.1 Component Masses

The GPS Receiver System component masses are as follows:

a. The mass of 1 GPS Receiver shall not exceed 3 kilograms (kg)

b. The total mass of GPS Antenna(s) shall not exceed 0.75 kg

5.2.2 Center of Mass Location

The contractor shall define the center of mass in the Mechanical ICD.

5.2.3 Center of Mass Accuracy

The center of mass of the GPSR shall be determined to within ±3 mm relative to an external reference.

5.2.4 Determination of Moments and Products of Inertia

The final moments and products of inertia of GPS Receiver shall be calculated to within 10%.

5.3 PHYSICAL ENVELOPE

a. The GPS Receiver outside dimensions, including mounting flanges and connectors shall not exceed the volume envelope of less than: 0.31 m length x 0.31 m width x 0.14 m height as defined in the GPS System Mechanical ICD(s).

b. Each GPS Antenna and mount assembly shall not exceed the volume envelope of less than 0.125 m in diameter and 0.110 m high as defined in the GPS System Mechanical

ICD(s).

5.4 MOUNTING

a. The package will be hard-mounted on a mechanical surface of the spacecraft structure.

Proper fit/alignment of the GPS Receiver to the structure shall be inherent in its design, fabrication, and assembly to the structure, through the use of close dimensional control in the location of mounting holes and the use of correct mounting hardware.

b. Mounting interface flatness, and co-planarity requirements for the component side of the interface (including brackets, if any, and shims) shall be as defined in the MICD.

c. The GPS Receiver shall have a flatness of 0.25 mm from any 2 points on the mounting interface for a component <0.3 m2 and 0.50 mm from any 2 points on the mounting interface for a component >0.3 m2 and <0.6 m2.

5.5 FIELDS OF VIEW

a. The GPS Antenna(s) shall meet the performance requirements in Section 4.1 while configured in an unobstructed view of the GPS satellites within 90 degrees of the spacecraft zenith axis, while nadir pointing in low earth orbit at 676.5 km altitude and 98 degrees inclination.

b. All GPS Antennas fields of view shall be as shown in the MICD.

5.6 ALIGNMENT

5.6.1 Antenna Offsets Knowledge

Preflight calibrations shall determine the locations of the phase centers of all GPS antennas with respect to an appropriately defined reference frame, to within 0.03 m, with 99% probability.

6.0 ENVIRONMENTAL REQUIREMENTS

Environmental design requirements for the spacecraft components are specified in this section.

The GPS System shall meet its performance requirements in section 4.0 during and after exposure to the environments specified in this section.

6.1 MECHANICAL FACTORS OF SAFETY

a. The GPS Receiver System as well as Mechanical Ground Support Equipment (MGSE) shall demonstrate positive Margins of Safety under limit loads for all yield and ultimate failures using the Factors of Safety (FS) defined in Table 6-1. Margin of Safety (MS) is defined as follows:

MS = (Allowable Stress (or Load) / (Applied Limit Stress (or Load) x FS)) -1

Table 6-1 Factors of Safety

b. Primary and secondary structure comprised of composite materials, Beryllium, bonded joints and/or bonded inserts shall be proof tested to 1.25 x Limit Load; qualification by analysis only is not acceptable. Actual flight component testing is preferred, but testing of representative sets of hardware with a similarity qualification argument can be used if approved by the NASA/GSFC COR.

6.2 QUASI-STATIC ACCELERATION

Quasi-static acceleration represents the combination of steady-state accelerations and the low frequency mechanically transmitted dynamic accelerations that occur during launch.

The GPS System shall be designed to withstand the quasi-static design limit loads defined in the mass-acceleration curve (MAC) shown in Table 6-2 without damage or degradation of performance. The design loads shown below will be updated based on the results of coupled loads analysis.

Linear interpolation should be used between breakpoints to determine the appropriate limit load as a function of GPS System weight. Note that these design limit loads are intended to cover only the low frequency launch environment and must be used in conjunction with the random vibration environments to assess structural margins.

Table 6-2 GPS System Design Limit Loads

6.3 FREQUENCY REQUIREMENT

6.3.1 Fundamental Launch Frequencies

The GPS System shall have a fundamental frequency greater than 100 Hz when hard mounted at its spacecraft interface. Any component, which fails to meet the specified fundamental frequency, must supply a finite element model, correlated to modal survey test results up to 50

Hz, to be used in coupled loads analyses. Requirements for the submitted finite element model are in the SOW and discussed in Section 11.5.4.

6.4 VIBRATION

6.4.1 Sinusoidal Vibration

The GPS System shall undergo qualification, protoflight or acceptance (level depends on qualification status of unit) sine vibration testing on all three axes at the levels shown in Table

6-3. Stiff components showing a first mode greater than 150Hz can be exempted from sine vibration testing upon approval by the NASA/GSFC COR. A generic sine vibration specification is provided for both Protoflight (PFT), Qualification (QT) and Acceptance (AT) levels applied at the PACE to GPS System interface. See Section 11.4.1 for definitions of

Protoflight, Qualification, and Acceptance.

Table 6-3 GPS System Sine Vibration Environment

Frequency Protoflight, Qual Level Flight, Acceptance Level

5-20 Hz 0.63 in (double amplitude) 0.50 in (double amplitude)

20-100 Hz 12.5 G 10.0 G

Figure 6-1 Sine Vibration Environment

This environment will be updated with random vibration analysis. Note for lightweight GPS

System, the highest design loads may be from this random vibration environment.

b. The contractor shall provide random vibration analysis along with static loads analysis.

Please see NASA-HDBK-7005 and NASA-STD-7001 for more information.

During the test, the test input level will be reduced (notched) at critical frequencies, if required, to limit the random vibration loads and/or acceleration responses to 3 dB above design limit levels.

c. Notching shall be limited to -12 dB of the original input and to a bandwidth of less than

100 Hz to limit the random vibration responses to 3dB above design limit levels.

Notching beyond these limits will require NASA/GSFC COR approval.

6.5 SHOCK

a. The GPS System shall be designed to meet its performance requirements after being subjected to the shock environment in Table 6-6, applied at the GPS System interface to the PACE spacecraft structure.

Table 6-6 GPS System Qualification Level Shock Response Spectrum

Freq (Hz) SRS (G)

100 100

800 1700

8000 1700

10000 2200

b. A shock susceptibility and attenuation assessment shall be performed on all Spacecraft components. If the flight shock environment as shown on a Shock Response Spectra

(SRS) plot (Q=10) is enveloped by the curve shown below (Figure 6-3), then the shock environment can be considered benign and there is low risk in deferring the shock test to the Observatory level.

Figure 6-3 Shock Envelope

c. Analysis supporting this conclusion (e.g., deferral of shock testing) shall be provided to the NASA/GSFC COR for review.

d. Any component determined to be susceptible to the shock environment (e.g., shock levels are above the curve) shall have shock testing performed at the component level

(preferably on a qualification unit).

6.6 NA

6.7 TRANSPORTATION

In addition to the launch loads shown above, the GPS System shall also be designed to withstand the maximum transportation loads shown in Table 6-7 without damage or degradation of performance.

Table 6-7 Transportation Loads

6.7.1 Transportation Cleanliness

Materials and enclosure used for transportation and storage shall not generate molecular or particle contaminants or degrade the surface cleanliness of the item or adjacent items

6.8 PRESSURE

6.8.1 Operating Pressure Range

The GPS System shall be designed to meet all performance requirements while operating over a pressure range of 1.08 x 105 N/m2 (813 Torr) to 1.3 x 10-12 N/m2 (1 x 10-14 Torr).

6.8.2 Maximum Depressurization Rate

The GPS System shall be designed to meet all performance requirements after exposure to a maximum depressurization rate of -50 mbar/sec (-0.72 psi/sec) experienced during launch and ascent.

6.8.3 Launch Vehicle Environmental Control System Impingement Velocity

The GPS System exterior surfaces shall not suffer damage or degradation when exposed to the

LV ECS airflow velocity of 10 m/sec.

6.9 ON-ORBIT DYNAMIC ENVIRONMENT

The GPS System shall be designed to handle all linear and angular acceleration requirements at the same time, i.e. the linear and angular acceleration are not mutually exclusive.

6.9.1.1 Dynamic Linear Acceleration

The GPS System shall survive on-orbit when subjected to the maximum linear acceleration due to nominal thruster firing of 0.060 m/sec2.

6.9.1.2 Dynamic Angular Acceleration

The GPS System shall survive the maximum angular acceleration of 2.5E-5 rad/sec2.

6.10 GROUND ENVIRONMENTS

a. The GPS System shall meet all of their performance requirements during exposure to air temperature between +5 and +30 degrees C and relative humidity between 30% and 70%.

b. After being powered “OFF” and” and exposed to air temperatures of +5 to +30 degrees C and relative humidity of 0 to 70%, the GPS System shall meet all of its performance requirements.

6.11 THERMAL REQUIREMENTS

6.11.1 Flight Interface Design Temperature Limits

a. When powered “OFF”, the GPS Receiver shall be capable of surviving indefinitely when its temperatures are within the survival limits shown in Table 6-8, without damage or permanent performance degradation.

b. The GPS Receiver shall meet all performance requirements when powered “ON” anywhere within the Operational and Protoflight/Qualification limits shown in Table 6-8.

c. The GPS Receiver shall demonstrate turn on at the Minimum and Maximum Protoflight

/Qualification (in spec) limits shown in Table 6-8.

d. When the GPS System is powered “OFF”, the GPS Antenna shall be capable of surviving indefinitely when its temperatures are within the survival limits shown in Table

6-8, without damage or permanent performance degradation.

e. When the GPS System is powered “ON”, the GPS Antenna shall meet all performance requirements anywhere within the Operational and Protoflight/Qualification limits shown in Table 6-8.

f. The GPS Antenna shall demonstrate turn on at the Minimum and Maximum Protoflight

/Qualification (in spec) limits shown in Table 6-8.

Table 6-8 Temperature Limits at Mounting Interface

Minimum Temperature (ºC) Maximum Temperature (ºC)

GPS Receiver Operational (In Spec) -10 +50 Protoflight/Qualification

(In Spec) -20 +60

Survival (Unpowered) -30 +60

GPS Antenna

Operational (In Spec) -20 +50

Protoflight/Qualification

(In Spec) -30 +60

Survival (Unpowered) -40 +70

6.12 CHARGED PARTICLE RADIATION REQUIREMENTS

Components containing electronic parts will be exposed to a natural space radiation environment that consists of: (1) trapped particles which include electrons, protons, and heavier ions; (2) particles from solar events (coronal mass ejections and flares); and (3) galactic cosmic ray particles.

6.12.1 Definitions

Total Ionizing Dose (TID) - the mean energy deposited by ionizing radiation in a device region divided by the mass of the region. This is often given in units of rad (Si), where 1 rad (Si) = 100 erg deposited per gram of silicon.

Radiation Design Margin (RDM) - the ratio of the derated hardness capability of the part to the estimated dose at the part location.

Enhanced Low Dose Rate Sensitivity (ELDRS) - used to refer to a part that shows enhanced radiation-induced damage at dose rates below 50 rad (Si)/s. The enhancement is the result of true dose rate effects.

Displacement Damage Dose (DDD) - the mean energy deposited by ionizing radiation in a device region that goes into atomic displacements divided by the mass of the region. There is no official unit for DDD. One such unit is MeV/g

Non-Ionizing Energy Loss (NIEL) - a measure of the rate of energy loss due to atomic displacements as a particle traverses a material.

Single Event Effect (SEE) - any measurable effect to a circuit due to an ion strike. This includes, but is not limited to, single event upsets (SEUs), single event transients (SETs), single hard errors (SHEs), single event latchups (SELs), single event functional interrupts (SEFIs), single event burnouts (SEBs), single event gate ruptures (SEGRs), and single event dielectric ruptures

(SEDRs).

Single Event Upset (SEU) - a change of state or transient induced by an energetic particle such as a cosmic ray or proton in a device. This may occur in digital or analog, circuits and may have effects in surrounding interface circuitry (a subset known as SETs). These are “soft” errors in that a reset or rewriting of the device will usually return the device to normal behavior thereafter.

The general goal for non-destructive events such as SEUs or SETs is not to avoid them completely, but to manage their impact through robust circuit design, automatic correction, and/or operational activities based on knowledge from ground radiation tests and circuit/system analysis.

Single Hard Error (SHE) - a SEU that causes a permanent change to the operation of a device.

An example is a stuck bit in a memory device.

Multiple Bit Upset (MBU) - an event induced by a single energetic particle such as a cosmic ray or proton that causes multiple upsets or transients during its path through a device or system in a single logical structure (ex., 2 bits affected in a single 16-bit word).

Single Event Functional Interrupt (SEFI) - a condition that causes loss of device functionality due to a single event in a control portion of a device. It generally requires a device reset or a re-initialization to resume normal device operations, but, for some devices, a power cycle is necessary to resume normal device operations. The general goal for non-destructive events such as SEFI is to avoid them, however, managing their impact through robust circuit design, automatic correction, and/or operational activities may be considered.

Single Event Latchup (SEL) - a condition that may cause device failure due to a single event induced high current state.