Attachment_A_-_ACS3_Payload_to_Spacecraft_Bus_ICD.pdf

Advanced Composite Solar Sail System (ACS3) Payload-to-Spacecraft Bus Interface Control Document (ICD)

ACS3 Project Payload-to-Spacecraft Bus ICD

ACS3-02-005

Table of Contents

1.0 Introduction

2.0 Units and Coordinates

2.1. Units

2.2. Payload Datum Planes

2.3. Mechanical Payload Coordinate System

3.0 Mechanical Interfaces

3.1. Primary Payload Volume

3.2. Primary Payload-to-Bus Interface

3.3. Payload-to-Solar Array Interface

3.4. Payload-to-Sun Sensors Interface

3.5. Payload-to-Patch Antennas Interface

3.6. Imaging Subsystem-to-Bus Interface

3.7. Payload Avionics-to-Bus Interface

3.8. UHF Transmitter-to-Bus Interface

3.9. Bonding

3.10. Electrical Connector Interface

3.11. Payload Cable Harness Routing

3.12. Mass

3.13. Center of Mass

3.14. Moments of Inertia

3.15. Debug Ports

3.16. Configuration

3.17. Mounting Hardware

3.18. Special Handling Requirements

3.19. Remove Before Flight Items

4.0 Electrical Interfaces

4.1. Payload Power

4.2. Payload Peak Power

4.3. Payload Commissioning Mode

4.4. Motor Pre-Heat Mode

ACS3-02-005

4.5. Deployment Start Mode

4.6. Deployment Mode

4.7. Deployment End Mode

4.8. Imagery Downlink Mode

4.9. Sail Snapshots Mode

4.10. Passivate and Power-Down

4.11. Payload Stand-By Mode

4.12. Bus Safe Mode and Commissioning

4.13. Payload Switching

4.14. Voltage to the Payload

4.15. Electrical Power Subsystem Telemetry

4.16. Power Quality

4.17. Over Current Circuit Limits

4.18. Maximum Current Draw Protection

4.19. Connector Interfaces

4.20. Primary Circuit Return Path

4.21. Primary Structure Grounding

4.22. UHF Transmitter Electrical Specifications

5.0 EMI/EMC

6.0 Communications Topology

7.0 Spacecraft Bus-to-Payload Communications Format

8.0 Telemetry Interfaces

8.1. Payload-to-Bus Telemetry Format

8.2. Time Synchronization Signal

9.0 Command Interfaces

9.1. Bus-to-Payload Command Format

9.2. Power-On Sequence

9.3. Power-Down Sequence

10.0 Thermal Interface

10.1. Thermal Models

10.2. Payload Thermal Control Operations

ACS3-02-005

10.3. Payload Thermal Boundary Conditions

11.0 Ground Support Equipment (GSE)

11.1. Mechanical

11.2. Electrical

12.0 Ground Handling

12.1. Cleanliness

12.2. ESD Protection

12.3. Temperature and Humidity

12.4. Monitoring

12.5. Hazardous Materials

12.6. Pressurized Systems

12.7. Ordinance

12.8. Radiation

13.0 Storage Conditions

13.1. Cleanliness

13.2. Temperature and Humidity

13.3. Monitoring

13.4. Access During Storage

ACS3-02-005

1.0 Introduction

For a thorough introduction to the project, please review the ACS3 Concept of Operations document.

Figure 1 is a notional concept of the 12U ACS3 spacecraft (payload integrated with the bus). There are independent spools for each sail and boom in the Sail Stowage Compartment and Boom Deployer, respectively. Four cameras are placed at the root of the spacecraft to provide imagery of the sail/boom system during and after deployment.

Figure 1. Notional Concept of Integrated Spacecraft with Booms Deployed (sail hidden for clarity).

Figure 2 shows the spacecraft with a fully deployed sail/boom system from the vantage point of an observer looking towards the sun (along the +x axis). The positive z-axis is defined as nadir-pointing.

ACS3-02-005

Figure 2. Sail Planform (+x is sunward, +z is nadir).

2.0 Units and Coordinates

2.1. Units

Units of millimeters, grams, milliwatts, milliamps, and Celsius shall be used, except where otherwise noted.

2.2. Payload Datum Planes

The payload datum planes are defined in Figure 3.

ACS3-02-005

Figure 3. Payload Datum Planes (+x is sunward, +z is nadir).

2.3. Mechanical Payload Coordinate System

The Payload Coordinate System (PCS) is shown in Figure 4. The origin is located at the midline of the spacecraft in the y- and z-directions, and in the plane of the elastic axes of the deployed booms in the x-direction. Specifically, it is 151.3 mm from datum I in the x-direction, 121.15 mm from datum J in the y-direction, and 119.15 mm from datum K in the z-direction.

ACS3-02-005

Figure 4. Payload Coordinate System (+x is sunward, +z is nadir).

3.0 Mechanical Interfaces

3.1. Primary Payload Volume

The spacecraft bus shall accommodate the payload volume as shown in Figures 5, 6, and 7, while also meeting spacecraft bus requirement 6.1.2, Volume/Dispenser. The ACS3 payload will likely be too large (in the y- and z-directions) to fit within the bus mechanical frame. Therefore, the structure typically reserved for payloads may need to be cut away, and the payload directly mounted to the resulting bus cross section. In addition, the bus vendor shall accommodate the payload cameras, avionics, and UHF transmitter within the bus volume (mechanical details can be found in Sections 3.6, 3.7, and 3.8). The details about specific placement will be discussed post contract award.

The dimensions in the y- and z-directions (242.3 mm and 238.3 mm, respectively) may preclude the ability for solar arrays to stow against the payload depending on the solar array design and/or selected launch dispenser. If this is the case, it may be possible to shorten the length of the solar arrays such that they only stow against the bus surfaces. However, it is understood this may not be possible depending on the power needs of both the payload and bus. Detailed discussions on this topic will take place post bus contract award.

ACS3-02-005

Figure 5. Payload Envelope (X-Z Plane)

Figure 6. Payload Envelope (X-Y Plane)

ACS3-02-005

Figure 7. Payload Envelope (Y-Z Plane)

3.2. Primary Payload-to-Bus Interface

The bus vendor shall propose a mechanical interface solution that is compatible with the ACS3 payload and minimizes customization of the vendor’s standard bus designs. The details of the proposed primary payload-to-bus mechanical interface will be discussed in detail with the bus vendor post contract award.

Figure 8. Payload Interface (isometric view). Electrical wiring layout is notional.

Figure 9. Payload Interface (top view). Electrical wiring layout is notional.

3.3. Payload-to-Solar Array Interface

The payload shall accommodate any components (such as release mechanisms) needed to retain the stowed solar array panels along datum D (assuming the panels are long enough to interface directly with the payload). This includes the possible routing of wires from the spacecraft bus towards datum D to trigger the release of the solar arrays.

The location planned to accommodate the bus provided hardware is shown in Figure 10. Details of this interface will be negotiated post bus contract award.

Figure 10. Location to Accommodate Bus-Provided Hardware to Secure Solar Panels.

3.4. Payload-to-Sun Sensors Interface

The payload shall accommodate the routing of wires from the spacecraft bus to any sun sensors placed on the payload side of the primary payload-to-bus interface.

3.5. Payload-to-Patch Antennas Interface

The payload shall accommodate the routing of wires from the spacecraft bus to any patch antennas placed on the payload side of the primary payload-to-bus interface.

3.6. Imaging Subsystem-to-Bus Interface

The spacecraft bus shall accommodate the placement of four (4) Government Furnished Equipment (GFE) cameras within the spacecraft bus to image the solar sail during and after deployment. Each camera can be estimated as a 34mm x 34mm x 34mm cube (to be refined as the design matures). To maximize the vantage point of the cameras with respect to the deployed boom/sail system, placement of the cameras as close to the base of the spacecraft bus (in the +X-direction) is desired (Figure 11).

Each camera will have a minimum field of view of 90 degrees in both the horizontal (rotation about the x-axis) and vertical (rotation about the y- or z-axis, depending on location) directions.

ACS3-02-005

Figure 11. Notional Placement of the Payload Cameras within the Spacecraft Bus.

3.7. Payload Avionics-to-Bus Interface

The spacecraft bus shall accommodate the placement of payload avionics boards within the spacecraft bus. The maximum expected envelope of these boards are 150-mm x 100-mm x 25-mm and 55mm x 40mm x 20mm. The larger board can be broken up into multiple smaller boards to help with packaging. The boards will include a System on Module (SOM), memory, multiplexing, de-serializer for the four cameras, motor controller, power regulation, and drive electronics for the solar sail motor. The notional layout of the larger board is shown in Figure 12.

Figure 12. Notional Layout of the 150mm x 100mm x 25mm Payload Avionics Board.

3.8. UHF Transmitter-to-Bus Interface

ACS3-02-005

The spacecraft bus shall accommodate the placement of the Stensat Radio Beacon (Figure 13) within the spacecraft bus. Mechanical specifications are provided in Table 1.

Table 1. Stensat Mechanical Specifications.

Dimensions 44.45 mm x 78.74 mm Mass See Table 2 Mounting Holes 0.125 inches, 4-40 mounting hardware

Figure 13. Stensat Radio Beacon.

3.9. Bonding

Use of bonded interfaces shall be coordinated and approved by NASA.

3.10. Electrical Connector Interface

Electrical interfacing between the bus and payload will occur at the primary payload avionics board (Figure 12) within the bus volume. The connection type, location, and orientation is dependent on the final placement of the payload avionics boards, and will be discussed in further detail post contract award.

3.11. Payload Cable Harness Routing

The spacecraft bus shall implement keep-out volumes to allow for cable routing between the following components (as necessary):

- Each of the four GFE cameras (Section 3.6) to the payload avionics boards (Section 3.7)

- Payload avionics boards to the primary interface with the payload (Section 3.2)

- UHF transmitter (Section 3.8) to the payload avionics boards

ACS3-02-005

Keep-out volume details will be finalized after contract award, as they rely on knowledge of both payload component placement within the bus and the layout of the bus itself.

3.12. Mass

The spacecraft bus shall not exceed a mass of 10 kg.

The payload shall not exceed a mass of 15 kg.

The 12U launch dispenser should accommodate up to 25 kg of mass. The spacecraft bus mass shall include payload avionics cards, cameras, and UHF transmitter (as specified in Table 2).

Table 2. Bus-Housed Payload Component Masses.

Component(s) Mass (grams) Cameras 20 (5 each) Payload Avionics Card(s) 120 UHF Transmitter 50

3.13. Center of Mass

The spacecraft bus shall accommodate payload centers of mass (in both the stowed and deployed state) as defined in Tables 3 and 4.

Table 3. Payload Center of Mass (Stowed)

X-Direction (PCS, mm) -12.0 ± 4.0 Y-Direction (PCS, mm) 0.1 ± 0.1 Z-Direction (PCS, mm) 0.0 ± 0.1

Table 4. Payload Center of Mass (Deployed)

X-Direction (PCS, mm) -9.7 ± 3.0 Y-Direction (PCS, mm) 0.1 ± 0.1 Z-Direction (PCS, mm) 0.0 ± 0.1

3.14. Moments of Inertia

The payload inertias are listed in Tables 5 and 6. They are reported about the center of mass for their respective configurations (see Section 3.12), with the x-, y-, and z-directions consistent with the PCS.

These numbers represent the upper bound on deployed inertia (7.0 meter-long booms with a sail area of 82 m2). The sail normal (when deployed) is in the [1 0 0] direction.

ACS3-02-005

Table 5. Payload Inertias (Stowed Solar Sail)

Ixx (kg-m2) 0.1 ± 0.1 Iyy (kg-m2) 0.09 ± 0.1 Izz (kg-m2) 0.09 ± 0.1

Table 6. Payload Inertias (Deployed Solar Sail)

Ixx (kg-m2) 17.25 ± 2.0 Iyy (kg-m2) 8.75 ± 1.0 Izz (kg-m2) 8.75 ± 1.0

3.15. Debug Ports

The spacecraft bus shall accommodate access to a payload debug port connector for reprogramming during flight system testing. The type and location of the port is to be negotiated post contract award.

3.16. Configuration

Payload CAD files shall be provided to the bus provider, and updated versions shall be sent as necessary and available.

3.17. Mounting Hardware

The mounting fasteners between the payload and bus will be negotiated post bus contract award.

The material of the payload mechanical interface is anodized 6061 aluminum (TBR).

3.18. Special Handling Requirements

Special handling requirements of NASA-provided GFE (e.g., payload avionics board, cameras, UHF transmitter, etc.) will be discussed in further detail post contract award.

3.19. Remove Before Flight Items

The spacecraft bus vendor shall inform NASA of “Remove Before Flight” items needed for the bus.

4.0 Electrical Interfaces

4.1. Payload Power

The spacecraft bus shall accommodate a payload power draw as specified in Table 7, which includes margin.

ACS3-02-005

This is a one-time event driven by the time it takes a pair of frangibolts to reengage the launch lock mechanism. It is estimated to take 30 seconds, but may take up to 60 seconds if colder than expected.

4.8. Imagery Downlink Mode

This is a long-term, continuous event that drives the mission duration (defined in requirement 6.1.1).

4.9. Sail Snapshots Mode

This is a repetitive event. It is driven by the time the cameras need to take photos of the deployed sail. Images of the deployed sail are expected to be taken once every 8.5 minutes for the duration of a full orbit, and then repeated every week (i.e., every week images are taken every 8.5 minutes for one orbit).

4.10. Passivate and Power-Down

Once all mission success criteria are met, and the final target orbit is achieved, the payload (and bus) shall passivate and power down (requirement 6.3.8).

4.11. Payload Stand-By Mode

The payload’s nominal mode is stand-by. During this time minimal payload avionics are powered on.

4.12. Bus Safe Mode and Commissioning

During spacecraft bus safe mode and commissioning, the payload will be powered off.

4.13. Payload Switching

The payload shall be operated with a single 12V nominal load switch from the bus.

Note: Load switching on 12V, 5V, and 3.3V lines is preferred if those voltages can easily be implemented at the payload interface (see Section 4.14).

4.14. Voltage to the Payload

The spacecraft bus shall provide regulated 12V +/- 1.0V power at the payload interface under nominal conditions.

Note: The payload is capable of stepping-down the regulated 12V power signal to 5.0V and 3.3V.

However, the ability for the bus to provide regulated 12V, 5.0V, and 3.3V power to the payload directly may be a viable alternative. Further discussions about this will take place post bus contract award.

4.15. Electrical Power Subsystem Telemetry

ACS3-02-005

The spacecraft bus shall collect, store, and telemeter to the ground payload current consumption on each power line (as applicable).

The payload shall provide the spacecraft bus with current consumption data on each power line (if the switching occurs on the payload side of the interface).

Note: The side of the interface with the power switches is most likely where this information will originate (Section 4.13).

4.16. Power Quality

The peak-to-peak ripple in the power supply shall be smaller than 250 mV (TBR).

4.17. Over Current Circuit Limits

(TBD) payload current limiting specification shall be provided by the spacecraft bus provider and included in the ICD post spacecraft bus contract award.

4.18. Maximum Current Draw Protection

The spacecraft bus shall limit current draw on each power line to a design configurable threshold.

Note 1: The side of the interface regulating the power will typically do this (Section 4.13).

Note 2: The value should be set high enough not to inadvertently trigger during nominal operations.

Note 3: Ideally this would be on-orbit resettable.

4.19. Connector Interfaces

Data

Connectors and pin-outs will be negotiated post contract award.

Note: Likely to be UART, differential RS-422, or TTL

Power

Connectors and pin-outs will be negotiated post contract award.

4.20. Primary Circuit Return Path

The spacecraft bus vendor shall ensure wires are used for the primary circuit return path. Structure or shields shall not be used.

ACS3-02-005

The payload shall ensure wires are used for the primary circuit return path. Structure or shields shall not be used.

Note 1: From ARC-STD-8070.1 – 3.3.8.2

4.21. Primary Structure Grounding

The spacecraft bus vendor shall ensure each subsystem ground tree has a local single point DC ground to spacecraft chassis via the shortest practical wire length.

Note: ARC-STD-8070.1 – 3.3.8.1

4.22. UHF Transmitter Electrical Specifications

The spacecraft bus shall accommodate the operation of the Stensat radio beacon (Figure 13) within the spacecraft bus. Electrical and Radio Frequency (RF) specifications are provided in Table 8.

Table 8. Stensat Electrical / RF Specifications.

Band 70 cm RF Output Power 0 to 3 Watt programmable Operating Voltage 5.0 to 12.0 Volts Operating Current 650 to 2000 mA (transmitting)

40 mA (idle) Serial Interface Rate 38.4 Kbaud UART 8 bit, no parity, one stop bit Digital Input Signal Specifications High signal > 0.7*Vdd

Low signal < 0.3*Vdd, 10 uA Digital Output Signal Specifications High signal > 2.0 volts

Low signal < 0.4 volts, 3 ma sink Frequency Range 420-450 MHz

5.0 EMI/EMC

The spacecraft bus shall be electromagnetically self-compatible (per spacecraft bus requirement 6.1.8). This includes when assembled with the Ames-provided payload avionics board, UHF transmitter, and cameras.

The payload shall be electromagnetically self-compatible.

6.0 Communications Topology

Figure 14 shows the communication topology for the ACS3 mission.

ACS3-02-005

10.1. Thermal Models

The payload thermal models are in a format usable in Thermal Desktop and shall be delivered to the bus provider for integrated verification.

10.2. Payload Thermal Control Operations

The payload’s thermal control system (TCS) is expected to be largely passive. A heater may be required to bring the motor up to operational temperature prior to solar sail deployment. During this time, a command shall be sent to the payload to enter the mode, and temperature data shall be sent from the payload to the bus to downlink to the ground.

10.3. Payload Thermal Boundary Conditions

The payload shall be thermally coupled to the bus at the primary interface plane.

11.0 Ground Support Equipment (GSE)

11.1. Mechanical

The spacecraft bus vendor shall provide mechanical GSE needed to handle the bus during payload integration.

11.2. Electrical

The spacecraft bus vendor shall provide electrical GSE needed to integrate the bus with the payload.

12.0 Ground Handling

12.1. Cleanliness

The spacecraft bus vendor shall provide guidance on bus ground handling procedures post contract

12.2. ESD Protection

12.3. Temperature and Humidity

ACS3-02-005

12.4. Monitoring

12.5. Hazardous Materials

N/A

12.6. Pressurized Systems

12.7. Ordinance

12.8. Radiation

13.0 Storage Conditions

The spacecraft bus vendor shall provide guidance on bus storage conditions post contract award.

13.1. Cleanliness

13.2. Temperature and Humidity

13.3. Monitoring

13.4. Access During Storage

Access to the payload/bus during storage shall be restricted to personnel designated for payload/bus handling, integration, and test.