B669436_ARACHNE _BUS_SPECIFICATION_FINAL.pdf

Blue Canyon Technologies X-SAT SATURN Bus Requirements Spec for the Pandora Program

Unclassified // Unlimited Distribution Lawrence Livermore National Laboratory 1

Document release number: LLNL-MI-2003212

L L N L X X X X X X X X X

ARACHNE Bus Specification

Unclassified // Distribution Unlimited

19 February 2025

Lawrence Livermore National Laboratory 2

Disclaimer This document was prepared as an account of work sponsored by an agency of the United States government. Neither the United States government nor Lawrence Livermore National Security, LLC, nor any of their employees makes any warranty, expressed or implied, or assumes any legal liability or responsibility for the accuracy, completeness, or usefulness of any information, apparatus, product, or process disclosed, or represents that its use would not infringe privately owned rights. Reference herein to any specific commercial product, process, or service by trade name, trademark, manufacturer, or otherwise does not necessarily constitute or imply its endorsement, recommendation, or favoring by the United States government or Lawrence Livermore National Security, LLC. The views and opinions of authors expressed herein do not necessarily state or reflect those of the United States government or Lawrence Livermore National Security, LLC, and shall not be used for advertising or product endorsement purposes.

Lawrence Livermore National Laboratory is operated by Lawrence Livermore National Security, LLC, for the U.S. Department of Energy, National Nuclear Security Administration under Contract DE-AC52-07NA27344.

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Contents

1 PURPOSE AND SCOPE

1.1 Requirement Descriptions

2 DEFINITIONS

2.1 Observatory

2.2 Bus and Payload

2.3 TBD/ TBR

3 REFERENCE DOCUMENTS

4 GENERAL REQUIREMENTS

4.1 Coordinate System

4.2 Bus Volume

4.3 Stowed Mass Properties

4.4 Orbit Definition

4.5 Lifetime

4.5.1 Minimum Design Lifetime

4.5.2 Minimum Storage Lifetime

4.6 Fault Management System

4.7 Payload

4.7.1 Payload Stowed CG

4.7.2 Payload Moments of Inertia

4.7.3 Payload Mounted Component Thermal Limits

4.8 Launch Powered Off

5 PAYLOAD INTERFACES

5.1 Payload Electrical Interfaces

5.2 Payload Mechanical Interfaces

6 ELECTRICAL POWER SYSTEM (EPS)

6.1 Bus Configuration

6.2 Payload Power

6.2.1 Payload Power Voltage

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6.2.2 Payload Power Consumption

6.2.3 Payload In-Rush Current

7 COMMAND & DATA HANDLING (C&DH)

7.1 Bus State of Health Telemetry

7.2 Telemetry Storage

7.3 Mission Data Storage

7.4 Data Downlink

7.5 Payload-Bus Data Formats

7.6 Watchdog Timer

7.7 Payload Data Interface

7.8 Payload PPS Interface

8 FLIGHT SOFTWARE

8.1 Boot Image Redundancy

8.2 Updating Software – On-Orbit

8.3 Packet Format

8.4 Payload TOD Message

8.5 Time Tagging

8.6 S-Band Radio Interface

8.7 X-Band Radio Interface

8.8 Absolute Time Stored Commanding

8.9 Relative Time Stored Commanding

8.10 EventChecks

8.11 Macros

8.12 Command History

8.13 Encryption

8.14 Error Correction/Scrubbing

8.15 Commandable Heater Set Points

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9 TELECOMMUNICATIONS

9.1 Ground System Capability – S-Band Radio

9.2 Ground System Capability – X-Band Transmitter

9.3 T&C Data Link

9.4 Mission Data Link

9.5 S-Band Rx Antennas

9.6 S-Band Tx Antennas

9.7 X-Band Medium Gain Antenna (MGA)

9.8 T&C Link Margin

9.9 Radiated Frequency

10 MECHANICAL

10.1 Launch Vehicle Interface

10.2 Bus Mass

10.3 Payload Mass

10.4 Debris Containment

10.5 Venting

10.6 Solar Array Drive Assembly

10.7 Bus Radiator

10.8 S-band Antennas (Rx and Tx)

10.8.1 Vendor Bus Mounted S-Band Antenna

10.8.2 Payload Mounted S-Band Antenna

10.9 GPS Antenna

10.10 First Mode

10.11 Solar Array Mode

10.12 NST Mounting

11 THERMAL

11.1 Thermal environments

11.2 Thermal Interfaces

11.2.1 Thermal Conductance

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11.2.2 Operational Limits

11.2.3 Non-operational (Survival) Limits

11.3 Temperature Margin

12 GUIDANCE, NAVIGATION, AND CONTROL (GN&C)

12.1 Slew Rate

12.2 Slew Acceleration

12.3 Fine Reference Orientation Knowledge

12.4 Fine Reference Pointing Control Accuracy

12.5 Fine Reference Pointing Control Stability

12.6 Position Knowledge

12.7 Safe Mode Pointing

12.8 Momentum Management

13 PROPULSION

14 TEST PROVISIONS

14.1 Hardline Command/Telemetry Connections

14.2 Remote Updating Software – Ground Testing

14.3 Remote Battery Charging

14.4 External power

15 ENVIRONMENTS AND TESTING

15.1 1g Test Environment

15.2 Vacuum

15.3 Self-Compatibility

15.4 Environmental Testing

15.4.1 Random Vibration Testing

15.4.2 Thermal Testing

15.4.3 Acoustic Environment

15.4.4 Self-Induced Shock Environment

16 MISSION ASSURANCE

16.1 Quality Assurance

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16.2 Material Usage

16.3 Workmanship

16.4 Cleanliness

17 RELIABILITY AND RADIATION

17.1 Radiation tolerance

17.2 LET Thresholds

17.3 Surface Charging

17.4 External Bus Structure Grounding

17.5 Internal Charging

17.6 Continuous Faraday Cage

17.7 Wiring ESD Protection

18 ACRONYMS

Bus Specification Lawrence Livermore National Laboratory 8

Section 1 1 Purpose and Scope

This specification defines and controls the technical requirements for design, manufacturing, test, and verification of the commercial vendor bus. The vendor bus will interface with a LLNS furnished payload to form an Observatory. The payload consists of a Cassegrain telescope assembly, a baffle, three detectors (UV, visible and infrared), associated electronics and thermal control.

1.1 Requirement Descriptions

The word “shall” is the only verb used for a binding requirement and requires verification of compliance by the vendor. The word “should/may” denotes a non-binding goal that is not formally verified by the vendor. The word “will” is used to state facts or declaration of purpose and is not a requirement to be verified by the vendor.

2 Definitions

2.1 Observatory

The term “Observatory” describes the vendor bus integrated with the "payload". [DEF-010]

2.2 Bus and Payload

The term “bus” refers to the vendor-provided bus. The term "payload" describes the payload(s) provided by LLNS [DEF-020].

2.3 TBD/ TBR

The term “To Be Determined (TBD)” is used in this specification where a performance value or a requirement is not known or cannot be stated in exact terms. LLNS will determine the unknown “TBD” requirement. The term “To Be Resolved (TBR)” means further review or revision of the established requirement may be necessary [DEF-030].

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3 Reference Documents

1. 1MICDxxxx-xxx Vendor Bus-Payload Mechanical Interface Control Document

2. xxxICD0002 (TBD) (Program) Vendor Bus-Payload Interface Control Document

3. xxxMICD0010 (TBD) (Program) MICD

4. xxxDOCxxxx (TBD) (Program) Vendor Bus Event Checks and Macros

5. 0PL0313 Vendor Contamination Control Plan

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4 General Requirements

4.1 Coordinate System

The Observatory Coordinate System shall be per the vendor Mechanical Interface Control Document (MICD) (1MICDxxxx-xxx) [GEN-010].

4.2 Bus Volume

The vendor stowed integrated observatory will fit into ESPA Grande or ESPA volume, as defined by NASA Science Mission Directorate (SMD) Launch Vehicle Secondary Payload Adapter Rideshare Users Guide with Do No Harm (SMD SPA RUG with DNH) (1MICDxxxx-

xxx) [GEN-020].

4.3 Stowed Mass Properties

The vendor bus shall have stowed mass properties as defined in the vendor MICD (1MICDxxxx-

xxx) [GEN-030].

4.4 Orbit Definition

The vendor bus shall be designed to operate in the following Sun-Synchronous orbit [GEN-040]:

• Altitude: 700 km

• Eccentricity: < 0.01

4.5 Lifetime

4.5.1 Minimum Design Lifetime

The vendor bus shall support a design lifetime of 26 months at LEO. [GEN-050].

4.5.2 Minimum Storage Lifetime

The vendor bus shall support a storage time of 1 year [GEN-060]. This is in addition to the design lifetime.

Note: The vendor bus will require maintenance, per the storage plan, during this period (e.g.

batteries, reaction wheels, etc.).

4.6 Fault Management System

The vendor bus shall be capable of autonomously responding to the detected fault conditions specified in xxxDOCxxxx, vendor bus EventChecks and Macros [GEN-070].

4.7 Payload

4.7.1 Payload Stowed CG

The payload will support the Observatory stowed CG that is < (TBD) cm from the Observatory Z-axis as centered on the Motorized Lightband (MLB). This shall be accomplished through a collaborative effort between the vendor and LLNS [GEN-080].

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4.7.2 Payload Moments of Inertia

The vendor bus shall support a maximum primary moment of inertia of 150 kg-m2 [GEN-090].

4.7.3 Payload Mounted Component Thermal Limits

LLNS will maintain the thermal limits, as defined in xxxICD0002, of payload mounted vendor components [GEN-100].

4.8 Launch Powered Off

The vendor bus shall support and maintain a powered-off mode for launch and not exit this mode until separation from the launch vehicle [GEN-110].

5 Payload Interfaces

5.1 Payload Electrical Interfaces

The vendor bus shall provide interfaces to the Payload as defined below, and as further specified herein [PLIF-010]. Quantity represents baseline configurations on the bus.

Table 1. Payload Electrical Interfaces

Description Quantity Note Payload Input Power

> 1

Payload preference is for input power to be delivered on a single line. With bus/payload accomodate, payload can support input power across multiple switched lines from the bus.

Switched Heater Power > 16 Heater power is booked under payload power consumption Thermistors > 16 Data Interface > 2 RS422 & ethernet, reef CDH 070 1 Pulse Per Second (1PPS) 1 Ref CDH 080

5.2 Payload Mechanical Interfaces

The vendor bus shall support the payload per the vendor bus standard mechanical interface that is defined in MICD, 1MICDxxxx-xxx [PLIF-020].

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6 Electrical Power System (EPS)

6.1 Bus Configuration

6.2 Payload Power

6.2.1 Payload Power Voltage

The vendor bus shall provide the payload with 27-32.8V unregulated, unfiltered voltage at the bus exterior payload power interface during nominal operations [EPS-030].

6.2.2 Payload Power Consumption

The vendor bus shall support a payload which consumes 200 W orbit average power (OAP), including heater power for the payload. The margin shall be dependent upon design maturity consistent with GSFC-STD-1000H [EPS-040].

Actual mission operations including payload peak power duration, duty cycle, vehicle pointing, transmitter on-time, etc. shall be adjusted as needed to ensure they remain within the fixed power resources provided by the bus (solar array and battery).

6.2.3 Payload In-Rush Current

The vendor bus shall support payload inrush current with a peak of 12.0A for less than 10 ms duration on the power load lines [EPS-050].

Rise Time

6.2.3.1 Payload Switch Rise Time – HC Power

The vendor bus rise time from 10% to 90% current on the power switches shall be 5 ms – 15 ms at with an input capacitance of less than 1 mF [EPS-060].

6.2.3.2 Payload Switch Rise Time – Deployment and Heater Power The vendor bus rise time from 10% to 90% current on deployment and heater switches shall be 5 ms – 15 ms at with an input capacitance of less than 10 μF [EPS-070].

7 Command & Data Handling (C&DH)

7.1 Bus State of Health Telemetry

The vendor bus shall be capable of recording bus State of Health (SOH) telemetry for subsequent downlink [CDH-010].

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7.2 Telemetry Storage

The vendor bus C&DH subsystem shall provide 6.5 GB (BOL) of non-volatile memory for bus telemetry and payload SOH data storage [CDH-020].

7.3 Mission Data Storage

The vendor bus C&DH subsystem shall provide 58 GB (BOL) of non-volatile memory for payload mission data storage [CDH-030].

7.4 Data Downlink

The vendor bus C&DH subsystem shall collect, package, and transmit telemetry, health and status, and mission data to the selected space-to-ground transmission link [CDH-040].

7.5 Payload-Bus Data Formats

The vendor bus shall send and receive data to and from the payload using the Consultative Committee for Space Data Systems (CCSDS) packet format as documented in xxxICD002

[CDH-050].

7.6 Watchdog Timer

The vendor bus C&DH subsystem shall provide a watch-dog circuit that operates independent of flight software and resets the Observatory to a known, commandable state when communication with flight software is lost [CDH-060].

7.7 Payload Data Interface

The vendor bus shall provide 3.3V RS422 protocol for simultaneous, bi-directional communication with the Payload [CDH-070].

In addition to the RS422 InterFace (IF), the vendor bus shall provide 1000 Base-T (Gigabit) Ethernet with sustained data rates of at least 80 (TBR) Mbps for simultaneous, bi-directional communication with the Payload [CDH-070].

7.8 Payload PPS Interface

The vendor bus shall provide a differential 1PPS signal to the payload per interface as documented in xxxICD002 [CDH-080].

8 Flight Software

8.1 Boot Image Redundancy

The vendor bus shall support multiple boot images of flight software and should include Cyclic Redundancy Check (CRC) or equivalent mechanism to check for boot image errors and ensure load of valid image [FSW-010].

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8.2 Updating Software – On-Orbit

The vendor bus shall support upload of new flight software and/or software patches while on-orbit [FSW-020].

8.3 Packet Format

The vendor bus software shall format and packetize bus and payload telemetry using CCSDS recommendations as shown in xxxICD0002 [FSW-030].

8.4 Payload TOD Message

The vendor bus shall provide a Time-of-Day (TOD) message to the payload in the format documented in 6ICD5178 [FSW-040].

Note: This message contains information such as "time at the tone," observatory attitude, orbit position and velocity, maneuver indication.

8.5 Time Tagging

Bus software shall support time tagging of Guidance, Navigation, & Control (GNC) telemetry (position & attitude) with a 1σ timing uncertainty in an external time frame (e.g. GPS) of less than 1.0 msec when GPS is valid [FSW-050].

8.6 S-Band Radio Interface

The vendor bus software shall provide command, control and data interfaces to a vendor S-band radio [FSW-060].

8.7 X-Band Radio Interface

The vendor bus software shall provide command, control and data interfaces to an X-band mission data vendor Software-Defined radio (transmit only) [FSW-070].

8.8 Absolute Time Stored Commanding

Bus software shall provide a stored command function which executes uploaded command loads time-tagged with absolute time to an accuracy less than 1 second from vehicle time for each command [FSW-080].

8.9 Relative Time Stored Commanding

Bus software shall provide a stored command function which executes uploaded command sequences time-tagged with relative time to an accuracy of 1 second from vehicle time for each command [FSW-090].

8.10 EventChecks

Bus software shall provide a telemetry monitoring function used to detect fault conditions against user-updateable table limits [FSW-100].

8.11 Macros

Bus software shall provide user-updateable fault safing macro command sequences that execute when triggered by telemetry monitoring function [FSW-110].

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8.12 Command History

The vendor bus telemetry shall provide command history to include 8 bytes of last accepted commands and 8 bytes of last rejected commands [FSW-120].

8.13 Encryption

The vendor bus shall use the Advanced Encryption Standard (AES) with 256-bit keys or equivalent [FSW-130].

8.14 Error Correction/Scrubbing

Static memory arrays used for critical functions (e.g. flight software, boot codes) shall have error correction codes and scrubbing to prevent the accumulation of errors [FSW-140].

8.15 Commandable Heater Set Points

The vendor bus shall provide the ability to change heater on/off set points by command, including commands issued by autonomous fault protection [FSW-150].

9 Telecommunications

9.1 Ground System Capability – S-Band Radio

The vendor bus shall support flight system transmit and receive communications with the ground station at a minimum ground elevation angle of 10° when the boresight of the Telemetry and Command (T&C) antenna is pointed directly at the ground station [TCOM-010].

9.2 Ground System Capability – X-Band Transmitter

The vendor bus shall support flight system transmit to the ground station at a minimum ground elevation angle of 10° when the boresight of the mission data link antenna is pointed directly at the ground station [TCOM-020].

9.3 T&C Data Link

The vendor bus shall provide a Radio Frequency (RF) uplink/downlink T&C system that supports the following characteristics [TCOM-030]:

Uplink Frequency Band 2025-2110 MHz Uplink Center Frequency (TBD) MHz Downlink Frequency Band 2200-2290 MHz Downlink Center Frequency (TBD) MHz Uplink: 100 kbps (nominal), 10 kbps (contingency) (TBR) Downlink: 100 kbps (nominal), 10 kbps (contingency) (TBR) Modulation: Binary Phase Shift Keying (BPSK) (uplink), Offset Quadrature Phase-shift Keying (OQPSK) (downlink) (TBD)

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Downlink FEC: Convolutional (K=7, R=1/2) (TBD)

9.4 Mission Data Link

The vendor bus shall provide an RF downlink system that supports the following characteristics

[TCOM-040]:

Downlink Frequency Band: 8025-8300 MHz Downlink: ≥ 20 Mbps (TBR) Modulation: OQPSK (downlink) (TBR) Downlink FEC: Convolutional (K=7, R=1/2) (TBD)

9.5 S-Band Rx Antennas

The vendor bus shall contain two primary S-band receive antennas, one located on the -X of the vendor bus and one located on the +X face of the Observatory, to provide near omni-coverage

[TCOM-050].

9.6 S-Band Tx Antennas

The vendor bus shall contain two primary S-band transmit antennas, one located on the -X of the vendor bus and one located on the +X face of the Observatory, to provide near omni-coverage

[TCOM-060].

9.7 X-Band Medium Gain Antenna (MGA)

The vendor bus shall furnish one X-band transmit antenna, nominally co-boresighted with the -X primary S-Band transmit antenna [TCOM-070].

9.8 T&C Link Margin

The vendor bus T&C RF link shall have greater than or equal to 3dB margin assuming a Kongsberg Satellite Service (KSAT) Lite 3.7m antenna [TCOM-080].

9.9 Radiated Frequency

The vendor bus shall radiate RF energy inside of the license-approved frequency range [TCOM- 090].

10 Mechanical

10.1 Launch Vehicle Interface

The vendor bus shall provide a mounting interface for a 24-inch Planetary System Corporation Motorized Lightband or equivalent [MECH-010].

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10.2 Bus Mass

The vendor bus and vendor items installed on the payload structure shall have a maximum dry mass of 175 kg [MECH-020].

10.3 Payload Mass

The vendor bus shall support a payload mass of up to 50 kg (TBR) [MECH-030].

10.4 Debris Containment

The vendor bus deployable devices shall deploy without generating visible debris [MECH-040].

10.5 Venting

The vendor bus shall be capable of surviving the ascent pressure decay environment either by showing that they are freely vented according to the following definition, or by analysis.

Freely vented components are defined as those with a ventable volume to venting area ratio (i.e.

Volume/Vent Area) of less than 2000 inches [MECH-050].

10.6 Solar Array Drive Assembly

The vendor bus shall provide a solar array drive motor which articulates the solar array wing about the Y-axis, as shown in 1MICDXXXX-XXX [MECH-060].

10.7 Bus Radiator

The vendor bus shall provide a nominal cold face of the Observatory for use as a radiator

[MECH-070].

10.8 S-band Antennas (Rx and Tx)

10.8.1 Vendor Bus Mounted S-Band Antenna

The vendor bus shall provide mounting for one primary S-band receive antenna and one primary S-band transmit antenna on the -X face of the Observatory [MECH-080].

10.8.2 Payload Mounted S-Band Antenna

The vendor shall furnish one primary S-band receive antenna and one primary S-band transmit antenna to be mounted to the +X face of the payload with a clear FOV [MECH-090].

10.9 GPS Antenna

The vendor bus shall provide a GPS antenna with a zenith field of view [MECH-100]. The Pandora antenna shall be payload mounted.

10.10 First Mode

The first mode (natural frequency) of the vendor bus (without the Payload) shall be > 45 Hz

(TBR) [MECH-110].

10.11 Solar Array Mode

The fundamental mode of the deployed solar array shall be greater than 1.2 Hz [MECH-120].

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10.12 NST Mounting

The vendor bus shall provide two-star trackers to be mounted on the payload with clear field of views located on the +Z side of the Observatory [MECH-130].

11 Thermal

11.1 Thermal environments

The vendor bus shall be designed to survive and operate in the expected thermal environment of the orbit defined in Section 4.4 [THRM-010].

11.2 Thermal Interfaces

11.2.1 Thermal Conductance

The vendor bus shall support a payload interface that has a thermal conductance of ≤ 0.01 W/°C

(TBR) [THRM-020].

Note: The intent of this requirement is to make the payload interface near-adiabatic / thermally isolated such that the payload and bus can be tested and verified independently before integrating the two.

11.2.2 Operational Limits

The vendor bus shall maintain the payload interface within a temperature range of -20°C to +60°C while in an operational mode on orbit [THRM-030].

11.2.3 Non-operational (Survival) Limits

The vendor bus shall maintain the payload interface within a temperature range of -30°C to +70°C while in a non-operational mode on orbit [THRM-040].

11.3 Temperature Margin

The vendor bus shall include thermal control as needed to ensure all bus components remain within their allowable temperature limits, with margins as defined in Figure 1 [THRM-050].

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Figure 1. Thermal Control Margins

12 Guidance, Navigation, and Control (GN&C)

12.1 Slew Rate

The vendor bus shall provide a slew rate of up to 1.25 deg/sec in any axis, assuming the Payload mass properties in Section 4.7.2 [GNC-010].

12.2 Slew Acceleration

The vendor bus shall provide a slew acceleration of up to 0.1 deg/sec2, assuming the Payload mass properties in Section 4.7.2 [GNC-020].

12.3 Fine Reference Orientation Knowledge

During nominal operations the Observatory orientation in the ECI frame (J2000) shall be known with 1σ accuracy 7 arcsec (LEO) in each of the three (3) Euler angles (e.g. yaw, pitch, roll). This assumes Observatory orientation is measured at the star trackers, both star trackers are producing valid data, and the vehicle is inertially staring (essentially no body rate) [GNC-030].

The bus shall be capable of using payload-generated control errors in the attitude control loop unless there is a written agreement from LLNS to remove this functionality.

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Note: The bus can use the attitude errors as computed by the payload, available at 5 Hz over the ethernet, to improve pointing knowledge, accuracy, and stability, if desired. Essentially, the payload can be used as an additional star tracker for the bus.

12.4 Fine Reference Pointing Control Accuracy

During nominal operations the Observatory 1σ (worst case RSS of 3 axes) pointing accuracy in the vendor star tracker coordinate frame shall be < 20 arc-sec. This assumes Observatory orientation is measured at the star trackers, both star trackers are producing valid data, and the vehicle is inertially staring (essentially no body rate) [GNC-040].

The Bus shall accept data from the payload’s visible sensor as defined in the TBS Bus to PL ICD for use in attitude determination.

A 5 arcsec combined offset from the desired pointing of the payload boresight will greatly improve the performance of the IR channel and should be considered a goal of the attitude control system.

12.5 Fine Reference Pointing Control Stability

During nominal operations, the Observatory 1σ (worst-case RSS of 3 axes) pointing stability in the vendor star tracker coordinate frame shall be < 3 arc-sec over a 60-sec duration, with a goal of 2 arc-sec over a 60-sec duration. This assumes Observatory orientation is measured at the star trackers, both star trackers are producing valid data, and the vehicle is inertially staring (essentially no body rate) [GNC-050].

12.6 Position Knowledge

During nominal operations the Observatory position in the ECI frame (J2000) shall be known with 1σ accuracy ≤ 10 m (LEO) in each of the three (3) mutually perpendicular axes [GNC-060].

12.7 Safe Mode Pointing

During “safe mode,” the on-board ADCS shall autonomously orient the Observatory to within 15° of the desired attitude (1σ) [GNC-070].

12.8 Momentum Management

The vendor bus shall be capable of autonomously managing the total space vehicle angular momentum [GNC-080].

13 Propulsion

13.1 Orbit Phasing at BOL

The vendor bus shall provide for orbit phasing of 120 deg separation of the three Observatories within one month of separation from the launch vehicle. Fine maintenance of this position is not required over the mission life, as all vehicles are assumed to degrade approximately equally.

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13.2 Collision Avoidance Maneuvers

The vendor bus shall provide for a maximum of 5 collision avoidance maneuvers per vehicle, and assumed to have substantial lead time (1-3 months) to affect the orbit trajectory.

13.3 Deorbit Capability

The vendor bus shall provide for a deorbit capability at EOL.

14 Test Provisions

14.1 Hardline Command/Telemetry Connections

The vendor bus shall support the capability to send telemetry to and receive commands from Ground Support Equipment (GSE) via an externally accessible test port [TEST-010].

14.2 Remote Updating Software – Ground Testing

The vendor bus shall support the capability to update flight software with Ground Support Equipment (GSE) via an externally accessible test port [TEST-020].

14.3 Remote Battery Charging

The vendor bus shall support the capability to be powered by Ground Support Equipment (GSE) via an externally accessible test port [TEST-030].

14.4 External power

The vendor bus EGSE connectors shall be accessible with the solar arrays in both stowed and deployed configurations [TEST-040].

15 Environments and Testing

15.1 1g Test Environment The vendor bus shall provide the capability to be tested in a 1g environment [ENV-010].

Note: The full deployment of solar arrays may necessitate an offloading apparatus to deploy in a 1g environment.

15.2 Vacuum

The vendor bus shall be capable of operating in a vacuum as defined per GSFC-STD-7000A

(GEVS) [ENV-020].

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15.3 Self-Compatibility

The vendor bus shall perform an electromagnetic self-compatibility test. [ENV-030].

15.4 Environmental Testing

15.4.1 Random Vibration Testing

The vendor bus shall meet performance after exposure to the GEVS protoflight random vibration environment specified in GSFC-STD-7000 [ENV-040].

15.4.2 Thermal Testing

The vendor bus shall meet performance after exposure to the GEVS protoflight thermal environment specified in GSFC-STD-7000 [ENV-050].

15.4.3 Acoustic Environment

The vendor bus shall survive the acoustics environment per GSFC-STD-7000. This may be satisfied by vibe test, or acoustics test if structural components are susceptible to excitation by the acoustics environment (e.g., large panels) [ENV-060].

15.4.4 Self-Induced Shock Environment

The vendor bus shall meet performance requirements after exposure to self-induced shock events (solar array deployment and MLB separation) deployments [ENV-070].

16 Mission Assurance

16.1 Quality Assurance

The vendor's approved and controlled quality assurance requirements shall be used [MA-010].

16.2 Material Usage

The vendor bus shall comply with vendor best practices and standards for material usage [MA- 020].

16.3 Workmanship

The unit shall be free of defects that would interfere with operational use such as excessive scratches, nicks, burrs, loose material, contamination, and corrosion [MA-030].

16.4 Cleanliness

The vendor bus shall, with the exception of environmental testing, be maintained to at least a Visibly Clean (VC) Level in accordance with 0PL0313 [MA-040].

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17 Reliability and Radiation

17.1 Radiation tolerance

All electronics shall be tolerant to >20 krad(Si) if at least 0.2" of equivalent aluminum shielding is available; otherwise additional shielding or mitigation shall be used [REL-010].

17.2 LET Thresholds

Destructive Single Event Effects (SEE) modes shall have a Linear Energy Transfer (LET) threshold of the following [REL-020]:

• Single Event Latchup (SEL) (e.g. bulk complementary metal-oxide-semiconductor (CMOS)) ≥ 37 MeV∙cm2/mg

• Single Event Gate Rupture (SEGR), Single Event Burnout (SEB), Single Event Dielectric Rupture (SEDR) (i.e. normally incident phenomena) ≥ 37 MeV∙cm2/mg

17.3 Surface Charging

All external surfaces directly exposed to the external space environment shall be conductive (or dissipative), conductively tied together (R<1E9 ohm, <1E6 ohm for conductors), conductively tied to the Observatory electrical ground, and conductively tied to the adjacent external materials

[RAD-030].

17.4 External Bus Structure Grounding

For vendor bus structural elements not inside a chassis, all conductive elements shall be connected to the Observatory ground. Note: Using a charge bleed off resistance of 1E12 Ohm is sufficient, but 1E8 Ohm is easier to verify [REL-040].

17.5 Internal Charging

All dielectrics and electronics shall have 0.200” of equivalent aluminum shielding in all directions, or shall be shown to have a mission worst case 24 hour charging flux below 0.1 pA/cm^2 [REL-050].

17.6 Continuous Faraday Cage

The vendor bus component module shall create a continuous Faraday cage for bus electronics

[REL-060].

17.7 Wiring ESD Protection

All wiring coming into the chassis/Faraday cage shall have ESD protection [REL-070].

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18 Acronyms

Acronym Full Form ADCS Attitude Determination and Control System AES Advanced Encryption Standard BOL Beginning of Life CMOS Complementary Metal-Oxide-Semiconductor CRC Cyclic Redundancy Check CCSDS Consultative Committee for Space Data Systems C&DH Command and Data Handling ECI Earth-Centered Inertial EGSE Electrical Ground Support Equipment EOL End of Life EPS Electrical Power System ESD Electrostatic Discharge FEC Forward Error Correction FOV Field of View FS Flight Software GEVS General Environmental Verification Standard GN&C Guidance, Navigation, and Control GPS Global Positioning System GSE Ground Support Equipment KSAT Kongsberg Satellite Services LEO Low Earth Orbit LET Linear Energy Transfer LLNL Lawrence Livermore National Laboratory LLNS Lawrence Livermore National Security LOB Line of Business MGA Medium Gain Antenna MICD Mechanical Interface Control Document MLB Motorized Lightband OAP Orbit Average Power OQPSK Offset Quadrature Phase-Shift Keying PPS Pulse Per Second RSS Root Sum Square RS422 Recommended Standard 422 (Serial Communication Standard) SEB Single Event Burnout

Lawrence Livermore National Laboratory 25

Acronym Full Form SEGR Single Event Gate Rupture SEL Single Event Latchup SEDR Single Event Dielectric Rupture SEE Single Event Effects SOH State of Health TBD To Be Determined TBR To Be Resolved TBS To Be Supplied T&C Telemetry and Command TOD Time of Day UV Ultraviolet VC Visibly Clean

4.8 Launch Powered Off

11.3 Temperature Margin

18 Acronyms