J1a_3_-_PTD_Spacecraft_to_Payload_ICD.pdf

Payload Technology Demonstrator (PTD)

Spacecraft to Payload

Interface Control Document

December 03, 2015

NASA Ames Research Center

Moffett Field, California

Pathfinder Technology Demonstrator (PTD) Spacecraft to Payload ICD ii

Spacecraft to Payload

Interface Control Document

Approved by:

Date Project Manager Pathfinder Technology Demonstrator Project

Spacecraft Lead

Project Systems Engineer

Propulsion System Lead iii

PTD PROJECT

DOCUMENT CHANGE LOG

Sheet: 1 of 1

REV

LEVEL

DESCRIPTION OF CHANGE

APPROVED

BY

DATE

APPROVED

Rev -

Baseline for RFP 02-09-2016 iv

TABLE OF CONTENTS

1.0 Scope

1.1 Purpose

1.2 Roles and Responsibilities

1.3 Document Organization

1.4 Acronyms and Definitions

1.5 Effectivity

1.6 Configuration Control

2.0 Documents

2.1 Applicable Documents

2.2 Reference Documents

2.3 Standards

General

3.0 Mechanical Interfaces

3.1 Mechanical Bus Overview

Mechanical Interface and Options

3.1.1.1 Payload MICD

3.1.1.2 Connector Locations and Cable Clamping

3.1.1.3 Payload Mechanical Safety Inhibit

3.1.1.4 Payload Subsystem-to-Spacecraft Electrical Isolation

Structural Design Geometric Relationships

3.1.3.1 Coordinate System

Interface Location and Dimensioning

3.1.4.1 Dimensions and Tolerances

3.2 Environments

4.0 Payload Thermal Control

4.1 Responsibilities

4.2 Design Considerations

Orbit Characteristics Orbit Thermal Modelling

4.3 Survival Heaters

4.4 Thermal Integration

4.5 Thermal Interfaces

MLI Interfaces Thermal Coatings Launch/Deployment Phase

4.6 Thermal Telemetry

5.0 Mass Properties

5.1 Payload and Thrust Axis

If design is a fixed thrust vector If design has a controllable thrust vector

6.0 Electrical Interfaces

6.1 Electrical Power Distribution

Connector Description Pin Definitions and Locations Battery Power Average Power Standby Power Survival Power Nominal Load Peak Load v

Commanding Voltage Over Current Circuit Limits Inrush Currents Grounding Wire Sizes Wiring EMI/EMC Requirements DC Power Input Ripple Current Overload Protection RS-422 Asynchronous Data Electrical Harnesses

6.1.20.1 Design and Fabrication Responsibility

7.0 Electromagnetic Compatibility (EMC) and Radiation

7.1 Conducted/Radiated EMISSION AND SUSCEPTIBILITY Interference

7.2 Radiation Tolerance

8.0 Contamination

8.1 Bakeout

9.0 DATA Interface

9.1 Overview

9.2 Spacecraft to Payload Command Data Interface

Command Bus Characteristics Command Format Command Function

9.3 Payload Subsystem to Spacecraft Telemetry Interface

Telemetry Bus Characteristics

9.3.1.1 Telemetry Format

Packet Priority Timestamps Fault Management

9.4 Integration Support

10.0 Appendix

10.1 Acronyms

vi

TABLE OF FIGURES

Figure 1-1: Pathfinder Technology Demonstrator Spacecraft Concept Figure 3-1: PTD 6U Volume Definition Figure 3-2: PTD Perspective View Figure 3-3: PTD MICD - Sheet 1 Figure 4-1: PTD Volume Divider Figure 6-2: Electrical Connector

TABLES

Table 4-1: Orbit Parameters Table 5-1: Payload Subsystem Mass Estimate (EXAMPLE)

1.0 SCOPE

The Pathfinder Technology Demonstrator Project (PTD) will be performing a series of flight demonstrations in LEO to expand Payload system capabilities for CubeSat-scale spacecraft with the capability to perform proximity operations, orbital maneuvers, deorbit, and to demonstrate electric Payload systems or other novel technologies. Demonstrations of 1 to 4 months will be performed in LEO to characterize and demonstrate Payload subsystems and raise their TRL.

These flights will be performed with a common 6U bus design that will accommodate a range of subsystems for demonstration of the selected Payload subsystems in LEO. The basic concept is in Figure 1-1

This concept is shown with solar arrays deployed and tracking in a sun-synchronous orientation. The Payload system shown is a notional Ion system just for perspective. This document defines the capabilities and PTD bus interfaces to allow incorporation of any Payload subsystem that can conform to these interface requirements.

Figure 1-1: Pathfinder Technology Demonstrator Spacecraft Concept

1.1 PURPOSE

The purpose of this document is to define and control the mechanical, electrical, thermal and data interfaces between the Pathfinder Technology Demonstrator bus (PTD) and the Payload subsystem (Payload).

1.2 ROLES AND RESPONSIBILITIES

NASA Ames Research Center (ARC) is responsible for the management of development and integration of the Cubesat bus, Ground Data System (GDS), launch support and Mission Operations. NASA Glenn Research Center (GRC) will procure and oversee vendor qualification of Payload subsystems and deliver flight-ready components for integration with the spacecraft bus. Payload system vendors will be managed by GRC to develop and deliver flight-ready Payload subsystems to the PTD project.

1.3 DOCUMENT ORGANIZATION

The PTD 6U Cubesat bus is effectively divided into two sections, an avionics volume and a Payload volume. This document defines the interfaces between these two volumes. The interfaces defined are:

Chapter 3.0 The Mechanical configuration the Payload system will have to fit and attach into

Chapter 4.0 Thermal design environment and options

Chapter 5.0 Mass Properties

Chapter 6.0 The Electrical and data interface characteristics and connector layout

Chapter 7.0 EMI/EMC environment and constraints

Chapter 8.0 Radiation

Chapter 9.0 Data Interface

1.4 ACRONYMS AND DEFINITIONS

Acronym definitions are defined in the Appendix.

1.5 EFFECTIVITY

This document is applicable to all PTD Payload subsystem configurations unless otherwise specified.

1.6 CONFIGURATION CONTROL

Configuration control of this document will be initiated upon signature approval and maintained in accordance with PTD Configuration Control Plan (TBD#).

2.0 DOCUMENTS

2.1 APPLICABLE DOCUMENTS

Documents applicable to this ICD are listed below. In the event of conflict between the documents referenced and the content of this ICD, the contents of this ICD shall be considered a superseding requirement.

1 Chemical Conversion Coatings on Aluminum and Aluminum Alloys, MIL-C-5541

2 National Space Transportation System Specification, Contamination Control Requirements, NASA SN-C-0005 (Current Issue)

3 EV2 Environments and Test Requirements Document for the PTD Program

4 NASA/GSFC std S-311-P-18K – Specification for Thermistor, (Thermally Sensitive Resistor), Insulated and Uninsulated, Negative Temperature Coefficient, Specifications for”

2.2 REFERENCE DOCUMENTS

Documents referenced in this ICD are listed below. In the event of conflict between the documents referenced and the content of this ICD, the contents of this ICD shall be considered a superseding requirement.

1 Electrical Characteristics of Unbalanced Voltage Digital Interface Circuits (ANSI/TIA/EIA- 423-B-96)

2 NASA MSFC-Spec-1238-A - Thermal Vacuum Bakeout Specification for Contamination Sensitive Hardware

3 NASA/GSFC std S-311-P-18K – Specification for Thermistor, (Thermally Sensitive Resistor), Insulated and Uninsulated, Negative Temperature Coefficient, Specifications for

4 NASA/ARC APR 8070.2 Class D Spacecraft Design and Environmental Test

2.3 STANDARDS

General

The Payload subsystem shall allow for a minimum 12-month shelf life prior to launch.

3.0 MECHANICAL INTERFACES

The mechanical interface for the PTD spacecraft is discussed in the following section.

3.1 MECHANICAL BUS OVERVIEW

The PTD bus structure is a 2U x 3U form factor divided into two separate volumes. The first volume is the Avionics 3.6U volume for spacecraft avionics, sensors, and attitude control system. The second 2.4U volume is for the Payload subsystem, and the focus of this document. A 2D view of the Payload and Avionics volumes can be seen in Figure 3-1. The arrow on the left of the graphic is the thrust direction for propulsion payloads. The Spacecraft vendor may increase the Payload volume size if space is available, but shall not go below this minimum size.

Figure 3-1: PTD 6U Volume Definition

Mechanical Interface and Options

The primary structure is an aluminum frame with an aluminum divider plate between the Avionics and Payload volumes. Aluminum covers wrap the frame to create enclosed volumes and act as structural stiffeners and also as radiators for thermal control. Based on the wide variety of Payload options and thermal loads, the cover plates on the Payload volume side faces (+Y, -Y) can be solid, or partially open (as shown in (Figure 3-2). Based on thermal properties of the Payload system, it is up to the Payload developer to specify an open or closed design for implementation.

The Payload volume end plate (-X) is tailored and provided by the Payload vendor. This is presumed to be the primary mount point for the Payload subsystem with Payload-unique cutouts provided as needed for the Payload subsystem. Other supports and brackets can be attached to the existing frame or cover plates as needed for structural support of the Payload subsystem. Screws or bolts used must be flush-mounted with the exterior surface. For bonded mounts, NASA approved low-outgassing bonding agents must be utilized with application and curing procedures defined.

Payload Volume Avionics Volume

Figure 3-2: PTD Perspective View

3.1.1.1 Payload MICD

Notional mechanical drawings are shown on the following pages showing dimensions, for a notional primary structure configuration. These are intended for perspective purposes only and not to imply a particular design solution, the Spacecraft vendor may utilize any design solution they choose that will meet contract requirements. Center of geometry is indicated and three axis indicators shown.

Propulsive force will push the bus in the +X direction

Figure 3-3: PTD MICD - Sheet 1

3.1.1.2 Connector Locations and Cable Clamping

The Payload Subsystem will install cable support interfaces such that the cables can be clamped within 5 inches of the connectors and a minimum of every 6 inches along the harness length (locations subject to review) adhering to standard bend radius specifications for selected wiring. The cable support provisions will be either a No. 10 (0.1900)-32 UNF-threaded receptacle or a tie strap retention device subject to NASA approval.

3.1.1.3 Payload Mechanical Safety Inhibit

The Payload vendor shall include a physical inhibit plug/connector that disables all possibility of the Payload system from activating. This plug will be red-flagged with a remove before flight ribbon. When the plug is removed, the Payload subsystem shall be in a flight-ready state for nominal activation.

3.1.1.4 Payload Subsystem-to-Spacecraft Electrical Isolation

Each Payload subassembly, connector, and harness shall be electrically isolated from the spacecraft bus, with electrical grounding only provided through the bus electrical interface in accordance with NASA specifications.

Structural Design

The Payload subsystem shall be designed to survive shock and vibration loads specified in GEVS (GSFC-STD-7000A) levels, with minimum safety factors for Payload system elements.

Geometric Relationships

3.1.3.1 Coordinate System

The Spacecraft coordinate system shall be in conformance with Figure 3-2 as shown by the X/Y/Z orientation in the upper right of the graphic.

Origin: The Origin for all measurements is the Geometric center of the primary structure as shown in the MICD in Figure 3-3 with the circular black/white indicator in each view.

Orientation: XS/C is along the long axis of the 6U volume, YS/C the short axis, and ZS/C the vertical axis.

A right-handed Cartesian coordinate system is employed.

Interface Location and Dimensioning

3.1.4.1 Dimensions and Tolerances

Unless otherwise specified, all linear dimensions are in millimeters, all angular dimensions are in degrees, and the tolerances for these are as follows:

Decimals: X.X = ± 0.1 mm X.XX = ± 0.01 mm X.XXX = ± 0.002 mm

Angles: ± 0°30' degrees

3.2 ENVIRONMENTS

The Payload Subsystem will be designed and tested to GEVS (GSFC-STD-7000A) space environments and in accordance with NASA APR 8070.2. The payload subsystem shall be tested through at least 8 (TBR) hot and cold thermal vacuum cycles to protoflight levels as part of vendor acceptance testing prior to delivery to PTD for integration.

4.0 PAYLOAD THERMAL CONTROL

4.1 RESPONSIBILITIES

The Payload thermal control environment is to be designed and specified by the Payload subsystem vendor. NASA will coordinate and support overall thermal design with the Payload vendor, but thermal behavior and aspects of the Payload system itself are the responsibility of the Payload vendor.

4.2 DESIGN CONSIDERATIONS

Orbit Characteristics

Table 4-1: Orbit Parameters

Season Altitude

(km) Inclination

(deg)

Beta Angle (deg)

Solar Flux (W/m2)

Albedo Earth IR (W/m2)

Cold Case Summer 800 98 90 1320 0.2 218

Hot Case Winter 350 98 0 1420 0.4 240

Orbit Thermal Modelling

A 98-deg Polar Sun-Synchronous orbit should be modelled. For the cold case, an altitude of 800 km, a sun angle parallel to the orbit plane (12pm orbit) with sun on the narrow “edges” of the spacecraft should be assumed. For the hot case, an altitude of 300 km, a sun angle perpendicular to the orbit plane (6am orbit) with sun on the wide “faces” of the spacecraft should be assumed. Solar flux, Albedo and Earth IR radiation shall be per APR 8070.2.

Side covers on the Payload volume are optionally open. The divider plate between the Payload and Avionics volumes will be insulated to limit any Payload waste heat from affecting the Avionics volume and sensor elements. See Figure 4-1 showing the divider plate highlighted.

4.3 SURVIVAL HEATERS

Survival heaters and control are the responsibility of the Payload vendor whether the Payload system is powered on or off. Bus design accommodation provides a dedicated unswitched power supply line to support this capability, even when primary power to the Payload subsystem is switched off. See Electrical pinouts in section 6.1. Survival heater(s) shall be thermostatically controlled and not require the Payload primary power to be active.

4.4 THERMAL INTEGRATION

The Spacecraft vendor shall be responsible for integrating the full spacecraft thermal model and will incorporate the vendor supplied Payload thermal design solution into the integrated system model.

Based on the integrated spacecraft thermal design, specific thermal accommodations may be required and coordinated between NASA, the Spacecraft Vendor, and the Payload vendor. Modification or attachments to the divider plate specified by the Payload vendor will be the responsibility of the Spacecraft vendor and require NASA approval. All coatings used by the Payload vendor must be approved by NASA.

Figure 4-1: PTD Volume Divider

4.5 THERMAL INTERFACES

This section defines the general Payload Subsystem to PTD bus TCS interfaces. The information within will ensure conformity with the NASA thermal analyses, and ensure all element interfaces and implementation responsibilities are addressed.

The primary thermal interface between the Payload and Avionics volume is via conduction of heat through the primary frame structure and surface cover plates. Ideally, the Payload thermal environment is independent and isolated from the Avionics volume, so thermal design of the Payload volume should minimize any energy flow into the Avionics volume.

No more than 4W (TBR) should transfer through conductive heat into or out of the Avionics volume.

The avionics volume bulk temperature may be assumed to be 0C for cold case and 40C for hot case analysis purposes.

MLI Interfaces

MLI may be used for the divider plate between the Payload and Avionics Volumes to isolate heat flow from the Payload subsystem in to the avionics volume. Specific MLI accommodations for the Payload subsystem and volume (not including the divider plate) are the responsibility of the Payload vendor to design and model. NASA as the bus integrator will install MLI in accordance with Payload vendor’s design and installation procedures.

X -X

-Y

Thermal Coatings

Any necessary thermal coatings or optical properties on surfaces that are required for the Payload subsystem are the responsibility of the Payload system vendor. Coatings intended for use must receive approval from NASA.

Launch/Deployment Phase

During launch and container deployment into LEO, the PTD bus will be completely unpowered for up to 45 minutes (TBR) and tumbling in orbit. During that period, no survival heater power will be available.

4.6 THERMAL TELEMETRY

Critical Payload subsystem areas and components should have their temperatures reported in Payload subsystem telemetry, with High and Low Yellow and Red limits specified so they can be monitored by Mission Operations. Unless otherwise directed, achievement of a Red limit temperature (which indicates imminent thermal failure) will result in removal of power and forced shutdown of the Payload system.

Pins are provided for an analogue temperature sensor to be wired so the survival temperature can be monitored by the PTD spacecraft. Placement and wiring of the survival temperature sensor is the responsibility of the Payload Vendor. The temperature sensor should be compliant to “NASA/GSFC std S- 311-P-18K – Specification for Thermistor, (Thermally Sensitive Resistor), Insulated and Uninsulated, Negative Temperature Coefficient, Specifications for”

5.0 MASS PROPERTIES

The mass budget and center of gravity for the Payload subsystem must be documented using NASA design contingencies. For new designs, typical contingency should be 30%, with higher maturities like build-to-print elements as low as 5%. An example table is shown that should be populated for your Payload subsystem.

Item Unit Mass

(kg) Qty

Current Mass Est.

Maturity-based Contingency

Projected Growth (kg)

Total Projected Mass (kg)

Totals

Table 5-1: Payload Subsystem Mass Estimate (EXAMPLE)

For PTD, the Payload system is allocated up to 3kg of mass for the completely integrated Payload subsystem

5.1 PAYLOAD AND THRUST AXIS

Thrust dynamics are dependent on the Payload design and quantity of thrusters, whether fixed or gimbaled, etc. In general, the thruster force applied must move the spacecraft in the X-axis direction with minimal body torque due to CG offset.

If design is a fixed thrust vector

The net Payload thrust vector shall not exceed an offset of 1.5mm (3 sigma) from the spacecraft X-axis at the point(s) of application of the thrust.

The direction of the net Payload thrust vector shall be aligned to within 2deg (3 sigma) of the spacecraft X axis.

If multiple thrusters are used, effective body torques that would be caused by a thruster failure must be documented. If thrusters can be fired individually, then effective torques for each thruster firing independently must be documented.

If design has a controllable thrust vector

The Payload thrust vector shall have the capability to provide a torque that will compensate for any body torques resulting from the angular misalignment or offset of the Payloads thrust vector relative to the spacecraft center-of-geometry.

6.0 ELECTRICAL INTERFACES

6.1 ELECTRICAL POWER DISTRIBUTION

The PTD Bus provides DC electrical power for the Payload Subsystem. The electrical and data interface for the Payload subsystem is supplied through connector J1 which will be a pass-thru connector through the payload and avionics volume divider. A notional connector solution with a pinout and supply constraint concept is described in the following subsections. Specific connector type and design will be negotiated by the Spacecraft and Payload vendor when the ICD is baselined.

Connector Description

Connector P/N: PTB-22-21PS

Manufacturer: Amphenol Industrial

Figure 6-1: Electrical Connector

Pin Definitions and Locations

Pin# Signal Description

A RS422TX+ Data from Payload Subsystem to S/C B RS422TX- Data from Payload Subsystem to S/C C RS422RX+ Data from S/C to Payload Subsystem D RS422RX- Data from S/C to Payload Subsystem E Digital GND Digital GND

F PWR1A Unregulated power from bus 1

G Return_PWR1A Return for unregulated power from bus 1 (GND)

H PWR2A Unregulated power from bus 2

J Return_PWR2A Return for unregulated power from bus 2 (GND)

K PWR1B Unregulated power from bus 1

L Return_PWR1B Return for unregulated power from bus (GND)

M PWR2B Unregulated power from bus 2

N Return_PWR2B Return for unregulated power from bus 2 (GND)

P SH1 Survival Heater Power

R Return_SH1 Return for Survival Heater power

S TSense Survival Heater Temperature Sensor

T Return_TSense Survival Heater Temperature Sensor Return

U Spare5 Spare pin

V Spare6 Spare pin

W Spare7 Spare pin

X Spare8 Spare pin

Battery Power

Unregulated battery power will be provided to the Payload Subsystem through redundant power connections. The Payload Subsystem will down regulate to any required voltages as needed. The Bus provides power on one of two switched services, A or B. Each service has a redundant path, 1 or 2. Thus when the ‘A’ service is powered, both PWR1A and PWR2A will be on, and conversely ‘B’ service will be off. The B side is a separate circuit and switch path that will only be used in the event of a failure on Service A. The Payload Vendor shall accept power from either A or B services.

Average Power

The Payload subsystem on-orbit-average continuous electrical power required shall be less than or equal to 20W (TBD). If operational use is more than this limit, then the Payload operation shall be capable of being duty-cycled to ensure this is not exceeded.

Standby Power

When the Payload subsystem is powered and not active, the electrical power required shall be documented in Watts. This number includes any expected power used for survival or operational heater loads.

Survival Power

The Bus shall provide to the Payload subsystem a minimum of 5W (TBR) orbit average (10 W peak (TBR)) unswitched power at any given time for used in supplying survival heater(s).

Nominal Load

The estimated Payload subsystem power consumption for each mode of operation needs to be defined. Produce a table defining modes of operation and power utilization for each. Examples might be initial power on, warmup, ready to fire, firing, stand-by, survival heating, etc. If there are different modes of Payload operation, e.g., different power levels to for different modes, this would need to be captured as well.

Insert table with the operation modes and power consumption

Peak Load

The Payload subsystem peak electrical power required shall not be greater than 180W (TBD). Note that this would require significant duty-cycling to ensure the on-orbit-average power does not exceed 20W (TBD).

Commanding

The Payload subsystem must accept commands from the PTD spacecraft to control Payload system operation.

Voltage

The voltage level provided from the Bus to the Payload subsystem will be unregulated

11.5V +/-2V DC.

Over Current Circuit Limits

The Payload subsystem will be switched off if it draws more than 5.5A for more than 250 milliseconds (TBR).

Inrush Currents

The Payload subsystem inrush current will be less than 5.5A for less than 250ms

(TBR).

Grounding

The Payload system is provided two grounds from the PTD bus; one dedicated to digital signals, and the other for analog. The Payload subsystem may not connect ground to the bus structure and must use provided grounds through the electrical connector.

Wire Sizes

Wiring shall be in accordance with manufacturer recommendations for voltage and current requirements and de-rated appropriately.

Wiring

Wiring design, selection, and testing should be done in conformance to the “SMC-S-020 standard “TECHNICAL REQUIREMENTS FOR WIRING HARNESS, SPACE VEHICLE”, latest issue.

EMI/EMC Requirements

The Payload subsystem shall be compatible with the PTD bus. Radiation and conduction levels need to be documented for the Payload subsystem so Spacecraft vendor can perform compatibility analysis to determine if filtering or shielding is required.

DC Power Input Ripple

The PTD unregulated power maximum input ripple is 200 mV p-p (TBR).

Current Overload Protection

The PTD power subsystem (EPS) provides over-current protection between the PTD bus and the Payload subsystem that will isolate a power system malfunction between the two subsystems. On overcurrent limit being reached, the power supply will be switched off to switched sources.

RS-422 Asynchronous Data

The S/C C&DH communicates asynchronously with the Payload subsystem via RS-422 asynchronous data transfer. The command and telemetry data transfer to and from Payload subsystem will be at 19200bps (TBD). See section 9.2.1 and 9.3.1

Electrical Harnesses

The Payload subsystem is responsible for providing the electrical harness and connectors required for operation of the Payload subsystem. Harness connection to the spacecraft will be integrated by NASA.

Pinouts, and safe-to-mate and installation procedures for the Payload Subsystem must be provided for the Payload subsystem by the Payload vendor. Harnesses are particular off-gassing contamination concerns; refer to Section 8.1 for bakeout requirements.

6.1.20.1 Design and Fabrication Responsibility

The design and fabrication of the electrical harness for the Payload subsystem is the responsibility of the Payload Vendor. This is to include any connectors, mount tie-downs, strain relief, etc.

7.0 ELECTROMAGNETIC COMPATIBILITY (EMC) AND RADIATION

7.1 CONDUCTED/RADIATED EMISSION AND SUSCEPTIBILITY INTERFERENCE

The Payload Subsystem shall be capable of nominal operations with expected externally or internally induced electromagnetic fields. It shall be self-compatible and operate in a space environment.

The Payload subsystem will measure and document generated fields and strengths and note any susceptibility concerns at specific frequencies. This will allow NASA to perform analysis to determine if there may be a need to implement any specific radiation or conductance mitigation to ensure self-compatibility of the entire integrated spacecraft.

7.2 RADIATION TOLERANCE

The Payload vendor will need to provide an analysis (or better actual testing) to show expected Total Ionizing Dose (TID) radiation susceptibility for critical and susceptible parts or assemblies. Since the PTD demonstration is only 1-4 months in a LEO orbit, TID will be relatively low.

The Payload subsystem TID susceptibility must be over 3kRad for critical components.

Single Event Latchup resistance should be at least 20MeV for critical components or assemblies.

8.0 CONTAMINATION

The Payload subsystem components shall be delivered VC+UV (Visibly clean, plus ultraviolet) at a cleanliness Level of 500 B.

Visibly clean: Cleanliness inspection which specifies an incident light of 50 foot-candles. The surface to be inspected shall be observed by the unaided eye (except for corrected vision) at a distance of 2 to 5 feet.

Plus Ultraviolet: Visibly clean (as defined above) and inspected with the aid of an ultraviolet light (black light) of 3200 to 3800 Angstrom wavelength (320 to 380 nanometers).

8.1 BAKEOUT

Payload components with high off-gassing concerns, like wiring insulation, will be baked out prior to delivery to PTD for integration.

The payload subsystem shall be tested through at least 4 hot and cold thermal vacuum cycles in reference to GEVS as part of vendor acceptance testing prior to delivery to PTD for integration.

All flight hardware shall meet outgassing certification rates prior to integration to the spacecraft.

Outgassing rates shall be measured with a TQCM, when available, set at the temperature 10 degrees C below the minimum on-orbit temperature of the “critical contamination sensitive element”. Measurements shall be made at a pressure less than 1 x 10-5 Torr during certification testing. The TQCM must be mounted within the chamber such that the TQCM has a representative view of the flight hardware or is monitoring the hardware vent. Bakeout criteria requires that when the rate of increase of the TQCM frequency data level out at a rate of 1-Hz/hour or less, that the rate shall be maintained (on the average) for a period of 36-hours. Temperature of the bakeout shall be the highest value permitted without endangering the hardware (i.e. below survival temperature), but should be at least 10 degrees C above the maximum orbital operating temperature. The target offgassing rate should be a rate of 5.7E-8 g/cm2/s or less.

The following test data shall be collected and delivered to the project: Chamber configuration (i.e., chamber size, use of shrouds, TQCM location, cold finger/scavenger plate locations (if used), and general test setup), TQCM readings (taken as a minimum every 0.5 hours), hardware temperature, chamber/shroud temperature, TQCM temperature, and pressure.

Following bakeout, the hardware shall be protected from surface recontamination resulting from subsequent handling and environments to which it will be exposed.

9.0 DATA INTERFACE

9.1 OVERVIEW

The PTD spacecraft FSW will manage Payload activities and operation of the Payload subsystem as directed/scripted from Operations Controllers on the ground. When commanded, the FSW will execute predetermined scripts and sequences to operate the Payload subsystem to exercise the goals of the Payload demonstration.

The capabilities of the Payload subsystem and modes of operation and how to perform them are the Vendor’s responsibility to document. Document any potential limits or dependencies., operational considerations for temperatures or voltages to monitor, etc. NASA will use these characteristics and work with the vendor to develop the on-orbit demonstration plan for the Payload subsystem to demonstrate its operation effectively.

9.2 SPACECRAFT TO PAYLOAD COMMAND DATA INTERFACE

Commands will be issued to the Payload subsystem via the RS-422 interface from the PTD Bus.

All commands received and processed by the Payload subsystem shall, at a minimum, increment a telemetered command counter to show execution of a received command.

Command Bus Characteristics

The command bus is implemented via an asynchronous, bidirectional RS-422 connection. The Universal Asynchronous Receiver / Transmitter (UART) connection settings are as follows:

Data Rate 19200 bits / second

Start bits 1

Data bits 8

Parity bits 0

Stop bits 1

Data bits are transmitted least-significant bit first.

Command Format

Commands are binary packets containing a command identifier, length, arguments, and checksum at a minimum. Use of a fixed or variable length solution is acceptable. If a larger number of states need to be simultaneously updated, or if a state is in the control loop that needs to be continuously updated then a fixed size array might be more appropriate. The states are encoded in bitfields or bit arrays and can be updated on each control tick so there is minimal parsing of commands to set up sequential individual commands. Specific implementation on the interface will be negotiated with NASA and the Payload and Spacecraft vendors after contract award.

Command Function

The Payload Vendor shall document available commands, command ordering or precedence dependencies, critical commands that have constraints to be observed before issuing, delays between commands, etc. This document is referred to as the Command Data Dictionary, and lists command formats, any arguments, any responses, and constraints for issue of the command.

The vendor is also responsible for documenting any procedures for subsystem startup, warmup, arming, firing, shutdown, etc. If the Payload subsystem can store command sequences or scripts for later execution, any built-in scripts and procedures for modifying scripts must be documented.

9.3 PAYLOAD SUBSYSTEM TO SPACECRAFT TELEMETRY INTERFACE

Data sent from the Payload subsystem to the spacecraft as Telemetry will be used to monitor Payload subsystem state, status, health and operation. Which data is made available is the responsibility of the Payload vendor, but should be enough to understand the operation of the subsystem, its health and any critical safety parameters.

Telemetry Bus Characteristics

The Telemetry bus is implemented via an asynchronous, bidirectional RS-422 connection. The Universal Asynchronous Receiver / Transmitter (UART) connection settings are as follows:

Data Rate 19200 bits / second

Start bits 1

Data bits 8

Parity bits 0

Stop bits 1

Data bits are transmitted Least-Significant Bit first.

9.3.1.1 Telemetry Format

Telemetry format will be sent as packets with a packet identifier, packet type, sequential counter, telemetry data (fixed length), and checksum. Endianess must be defined by the Payload vendor. Protocol for fixed and multi packet (if necessary) will be defined by the Protocol vendor and negotiated with NASA at program award.

Packet Priority

The PTD Operations team and SEs will determine the priority and frequency to downlink individual telemetry items to the ground for Payload subsystem data. If the Payload system vendor would like to recommend a prioritization and data decimation hierarchy for Payload subsystem data, NASA will work with the vendor to accommodate that recommendation.

Timestamps

All telemetered data will be time stamped by the PTD FSW at time of receipt from the data interface. The Payload Subsystem does not need to timestamp data. If the Payload subsystem vendor is interested in time-stamping data based on time criticality and resolution issues for effective analysis, this can be negotiated with NASA. This negotiation would include how time is set and maintained on the Payload subsystem and formats for representation.

Fault Management

Built-in Payload system fault management needs to be documented by the Payload vendor. Critical parameter monitoring that should be done by the Spacecraft to see if critical thresholds have been exceeded (Fault Detection) and the corresponding spacecraft reaction (Fault Response) to correct the condition or safe the Payload subsystem from potential damage must also be documented. Time criticality of the response is also necessary, since low criticality issues may just be monitored on the ground and not require a spacecraft direct response.

9.4 INTEGRATION SUPPORT

The Payload Vendor shall provide a test unit that supports the Command and Data Interface to allow early test and debugging of the communication interface.

10.0 APPENDIX

10.1 ACRONYMS

FSW Flight Software PTD Pathfinder Technology Demonstrator TID Total Ionizing Dose (Radiation)