Attachment 2 Systems Requirement Document (SRD).docx

Acquisition Delta, Innovation and Prototyping Space Test Program (STP) Space Test Experiment Platform 2.0 (STEP 2.0)

System Requirements Document (SRD)

Version 1.5 Space Systems Command Acquisition Delta, Innovation and Prototyping Space Test Program

(SSC/SZIS)

Kirtland Air Force Base Albuquerque, NM 87117

ATTACHMENT 2 – IDIQ SYSTEM REQUIREMENTS DOCUMENT (SRD)

ATTACHMENT 2 – IDIQ SYSTEM REQUIREMENTS DOCUMENT (SRD)

ii Acquisition Delta, Innovation and Prototyping Space Test Program (STP) Space Test Experiment Platform 2.0 (STEP 2.0)

System Requirements Document (SRD)

SUBMITTED BY:

STEVEN K. FUTCH, Maj, USSF Date Signed Program Manager, STEP 2.0

APPROVED BY:

JONATHAN P. SHEA, Lt Colonel, USSF Date Signed Director, Space Test Program

Change History

TABLE OF CONTENTS

1. GENERAL

1.1 DoD Space Test Program Background

The Department of Defense (DoD) Space Test Program (STP) is a multi-service program supporting Army, Navy, Air Force, and Space Force space test missions. DoD STP provides space access for Science and Technology (S&T) experiments throughout the DoD. S&T experiments are provided by the DoD Laboratories, (e.g., Air Force Research Lab (AFRL) and Naval Research Laboratory (NRL)) and any experimenter that has a DoD sponsor (e.g., universities, interagency cooperation, and international partners).

S&T is the backbone of emerging technologies therefore space experimentation is critical to continued Space Superiority support to the DoD Space Experiments Review Board (SERB) and International Space Access Review Board (ISARB). DoD STP currently has no in-house capability to procure and launch spacecraft with S&T payloads (PLs).

1.2 STEP 2.0 Scope

The Space Test Experiment Platform 2.0 (STEP 2.0) program is seeking to establish a multiple award Indefinite Delivery/Indefinite Quantity (IDIQ) contract with separate Delivery Orders (DOs) to procure proven spacecraft (SC) and space access for S&T experiments. The STEP 2.0 IDIQ contract vehicle provides STP the ability to execute DO(s) aligning PL(s), Space Vehicle (SV), and launch vehicle (LV) schedules.

The SC to be procured within each DO will range in size from 12U CubeSats to EELV (Evolved Expendable Launch Vehicle) Secondary Payload Adapter (ESPA) class, and any size in between. The SC shall be within the maximum mass requirements of a 12U to ESPA class vehicle. STP will define specific PL requirements in the Delivery Order (DO) Payload Requirements Document (PLRD). The SC will serve as the platform to integrate S&T PLs to operate in either Low Earth Orbit (LEO) or Geosynchronous Equatorial Orbit (GEO), as a class D system. The following provides a reference for bounding the range of mass and volumes for each SC form factor:

Minimum size and form factor: A 12U CubeSat with a size and form factor defined with dimensions not-to-exceed 226.3 x 226.3 x 366.0 mm (8.91 x 8.91 x 14.41 in) and a mass not-to-exceed 24 kg (53 lbs.) including PL(s). Acceptable 12U CubeSats shall be compatible with containerized dispensers for launch and separation (example either railed or tab-based dispenser). A typical 12U spacecraft (SC) should be able to host 8 kg (18 lbs., 1/3 of total mass) or more of PL mass.

Medium size and form factor: defined with dimensions in excess of 226.3 x 226.3 x 366.0 mm (8.91 x 8.91 x 14.41 in) and a mass in excess of 24 kg (53 lbs.) including PL; also defined with dimensions not to exceed the maximum size and form factor of the ESPA Class SC requirements defined in the immediate paragraph below.

Maximum size and form factor: A small SC (commonly considered an ESPA Class SC) having a mass up to 180 kg (400 lbs.) max including PL, with a size envelope of up to 61 cm (24 in) x 71.1 cm (28 in) x 92 cm (36.2 in). The SC for example may have a mechanical interface of 38.1 cm (15 in) bolt circle, using 24 – 0.635 cm (¼ in) fasteners, or may have alternate fastener system. A typical ESPA SC should be able to host 60 kg (132 lbs., 1/3 of total mass) or more of PL mass.

This contract includes the hardware, software and services needed to acquire a SC, integrate S&T PL, SV to LV integration, support launch operations, conduct on orbit experiment operations, and deliver data to STP’s PL customers. The classification will be identified within each DO DD254. The Space Vehicle Contractor (SVC) is to enter this effort with an existing ground segment to perform on orbit operations with spectrum approval, and infrastructure that meets DOD cyber security requirements. The SVC may be required to place the integrated SV into storage. The SVC shall deliver the integrated SV to the LV integration site. SVs procured and integrated under STEP 2.0 may fly on a dedicated small launch, on a small to heavy rocket as a rideshare on an integrated LV/PL stack, or via a launch broker rideshare service; the term LV represents all these options.

Flight Proven Spacecraft Each SC shall be flight proven. A flight proven SC shall have a flight history of 365 consecutive days of on orbit operations within the last 5 years. Flight proven SC may include incremental updates since last flight for which minor improvements are detailed and justified may meet the flight proven requirements (parts obsolescence, batteries, flight software updates, Internal Research and Development (IRAD) improvements, etc.).

CubeSats within the 12U form factor may be considered by demonstrating flight proven heritage via the SVC's 6U CubeSat with a flight history of 365 consecutive days of on orbit operations within the last 5 years. The SVC shall use flight proven subsystems from the SVC’s 6U CubeSat on the SVC’s 12U CubeSat form factor in order to utilize the 6U CubeSat flight heritage. The SVC’s 12U CubeSat shall be currently in production at least in integration and test phase.

Proven Ground Segment The SVC shall have a proven ground segment solution which consists of the ground station/system, operations facility, and communication network infrastructure providing command and control of a SCspacecraft and distributing payload data to the customer(s) for the proposed LEO or GEO SC orbit. The SVC or the SVC’s subcontractor shall have successfully operated an SV and distributed PL data for a minimum of 365 consecutive days with a 95 percent ground segment availability.

1.3 Document Overview

The Systems Requirements Document (SRD) defines minimum SV performance requirements. This document includes requirements for SC fabrication, PL integration, SV testing, LV integration, launch site testing, ascent, launch and early orbit check-out and one year of experiment mission operations support to include ground segment for a service.

STP, in collaboration with its respective mission partner(s), will develop individual mission requirements to be provided in a DO PLRD. Mission requirements include both mission-specific (standard) requirements, consistent with the Statement of Work (SOW), this document, and mission-unique requirements.

1.4 Key Definitions

The words “may”, “shall”, “will”, “spacecraft”, “space vehicle”, and “spacecraft contractor” have explicit meaning in this document. The definitions of these words as they apply to this document are listed here in Table 1 and in Appendix A Definitions:

Table 1 – Specific Definitions

1.5 Types of Requirements

This SRD discusses the division of responsibility between the PL and SC providers as well as common provisions for mechanical, thermal, power, and data interfaces that will be provided by the spacecraft. The combined SRD and PLRD will serve as the foundation for mission specific interface control documents (ICDs) to be delivered for each STEP 2.0 mission.

Standard Requirements All SVs shall meet the SRD Standard Requirements within this document.

Mission Requirements Mission requirements include both mission-specific and mission-unique requirements. The SVC shall meet DO PLRD requirements in addition to the SRD Standard Requirements.

Mission-Specific Requirements Mission-Specific Requirements are implementations of the requirements that are dependent on the specific mission being flown and will be specified in the DO PLRD. Mission-Specific Requirements are considered to be in scope of the SRD Standard Requirements and may tailor the SRD requirements as needed for the mission.

Mission-Unique Requirements Mission-Unique Requirements are not part of the SRD Standard but may be required and paid for by STP or its PL customer for a particular mission. Each mission will specify any mission-unique requirements in the DO PLRD and will be addressed within a DO SOW if required.

2. DOCUMENTS AND REQUIREMENTS

2.1 Applicable Documents

The documents listed in Appendix B, apply directly to the performance required, and contain provisions that constitute requirements of STEP 2.0 SOW and SRD to the degree specified in each section of this document. All mission specific applicable documents, including but not limited to those listed in Appendix B, Table 2, will be specified on each DO PLRD.

2.2 Compliance Reference and Guidance Documents

These documents listed in Appendix B, contain information relating to the work required by this SRD. The mission specific reference documents will be specified on each DO PLRD.

2.3 Requirements Precedence

In the event of a conflict in requirements between this SRD and any other referenced document, the following order of precedence applies: (1 DO SOW), (2 DO PLRD), (3 IDIQ SOW) and the (4 IDIQ SRD) this document. Applicable documents take precedence over Reference documents.

3. PROGRAM MANAGEMENT

This section is intentionally blank due to being covered in STEP 2.0 SOW Section 3.

4. GENERAL SYSTEMS ENGINEERING

4.1 Hardware and System Configuration Identification

Configuration Identification completely defines the as-built, as-designed, as-qualified deliverable hardware/system configuration and assures production of equivalent hardware for flight, this includes all development units upon which qualification related testing is performed. Procurement and/or product specifications shall be in general conformance with MIL-STD-961E. Drawings and associated lists shall conform to ASME Y14.100 except electrical schematic drawings which may conform to commercial drawing practice.

The SVC’s design activity number, drawing/part numbering system, and standard procedures for mono-multi detail drawings, separate/integral parts lists, and drawing format materials (i.e., pre-printed, stick-on, computer-graphic, etc.) shall be identified.

All other documentation (i.e., parts, materials and process standards, test procedures, test software, manufacturing procedures, etc.) shall conform to the SVC’s standard practices provided it is suitable for the purpose intended and complies with other SRD requirements (i.e., security, quality assurance, reliability). Non-deliverable drawings and data comprising the deliverable hardware/system configuration identification documentation, which exist at the SVC’s facility prior to contract award, may be used in existing form for manufacturing, procurement, inspection, and test purposes need not be upgraded as a condition of this paragraph.

The configuration identification of special test equipment, Ground Support Equipment (GSE), or other non-flight articles used in support of deliverable hardware/system acceptance shall consist of sufficient engineering, manufacturing, and test documentation necessary to assure complete configuration definition and control for evaluation of its effectiveness in the intended use.

4.2 Configuration Accounting

The highest level of configuration accounting shall be the deliverable assembly that shall be assigned the Government issued serial numbers such as “STPSat-X” as described in the DO. Subassemblies/components comprising the deliverable hardware/system shall be identified per the SVC’s serial numbering system.

All Class I changes to approved baselines shall be documented through an engineering change proposal and be submitted to the Government (STP) for approval. Class I changes are defined as changes that affect form, fit, function, cost, or schedule of the final assembly. Class II changes are defined as changes to correct documentation or changes to hardware not defined as Class I changes. The threshold criteria will provide Class I and Class II guidance to the Configuration Control Board (CCB), this shall be done to prevent insignificant changes such as typographical changes, which do incur cost, from being considered Class I. Class II changes need only be listed by number and title in the monthly status reports for review. Departures from or non-compliance with approved baselines shall be requested using a deviation/waiver request. The SVC’s configuration accounting system shall provide for traceability of Class I change implementation to the lowest level of hardware and test procedure impacted.

4.3 System Safety

The SVC’s programmatic system safety program shall comply with AFI-91-202 and MIL-STD-882E as tailored. The system shall actively prevent harm to personnel or sensitive equipment that could be caused by a single component failure, or a failure propagated from one component to another. Hazardous ionizing/non-ionizing radiation energy shall not be generated without provisions to protect personnel or sensitive subsystems from damage or adverse effects. Packaging and handling procedures and characteristics shall provide controls to protect personnel or sensitive equipment. The SVC’s system safety program shall aid in the prevention of errors in assembly, installation, connections, or operations that could harm personnel or sensitive equipment. The SVC’s system safety program shall prevent the release of energy sufficient enough to harm personnel or sensitive equipment. The SVC’s system safety program shall control the use of hazardous materials.

4.4 SV Environmental Verification Testing

SV testing shall incorporate adequate safety margins consistent with maximum PL allotments and the SV testing approach. Mass and power growth contingencies for PL and SV components shall be consistent with program maturity as guided by AIAA/ANSI R-020A-1999 for mass and AIAA/ANSI G-020-1992 for power. SC/SV testing shall demonstrate the adequacy of design in a method consistent with SMC-S-016. The SC/SV shall follow a proto-qualification testing program at a minimum, consistent with guidance in SMC-S-016, unless a full qualification test program has been completed. The SV shall be built to operate in the natural space environment as defined in MIL-STD-1809.

4.5 Humidity Control

During all phases of integration, assembly, testing, shipping, handling, storage, and launch, the SV shall be maintained in an environment with a range of 45% to 65% relative humidity or as specified within the DO PLRD if more stringent. The SV shall control temperature and humidity to avoid condensation and electrostatic discharge on the space hardware for all phases of assembly, integration, transportation, and launch.

4.6 Transportation

SC/SV and required GSE transportation shall limit mechanical loads, vibration, and shock imparted on the SC/SV in accordance with values used for design and verification. SC/SV transportation shall maintain temperature, humidity, and contamination control as specified here or as in the DO PLRD.

4.7 Contamination and Environmental Controls

The SC, SV, and its PLs shall be controlled to a minimum cleanliness environment of ISO 8 (class 100,000 level) or as specified in the DO PLRD. DOs may have more stringent cleanliness requirements.

5. SPACECRAFT REQUIREMENTS

5.1 Spacecraft Reliability

Successful SC/SV operation is defined as meeting or exceeding all DO requirements for a period of a minimum of 365 consecutive days of on-orbit operations after Launch and Early Operations (LEOP) and Normal Operations Readiness Review (NORR) acceptance.

5.2 Structures and Mechanisms

The SC shall be capable of surviving and subsequently operating as specified after all mechanical environmental exposures. The SC structure and mechanisms shall be verified to withstand the static and dynamic stresses, strains, shocks, vibrations, temperature, pressure, and vibro-acoustic environments associated with assembly, integration, test, shipping, handling, storage, launch, and on orbit operations as associated with TOR-2004(8583) - 1. The SC shall undergo proto-qualification testing as defined within SMC-S-016. Launch loads for the purposes of proto-qualification verification testing for all SCs will be developed by the LV provider on a mission specific basis in the LV provider’s LV-to-SV ICD based on an analysis of the integrated PL stack and LV ascent.

5.3 Mass Properties

SC/SV mass properties (e.g., mass, volume, Center of Gravity (C/G), moments of inertia) shall comply with the design specifications defined in the “Evolved Expendable Launch Vehicle (EELV) Secondary Payload Adapter (ESPA) Rideshare Users Guide (RUG)” (May 2010) for SCs larger than 12U CubeSats. 12U CubeSat mass properties shall comply with design specifications as defined in the Cal Poly CubeSat Design Specification Rev 14 (CP-CDS-R14) or design specifications as defined in the proven CubeSat separation system’s specification documents.

5.4 Electrical Power Subsystem

The SC shall supply power to PLs for at least 365 consecutive days post LEOP and NORR acceptance. PL duty cycle scenarios will be defined in cooperation between the SVC, STP, and the PL providers on a preliminary basis during PL accommodation design. The SVC shall ensure SC batteries comply with SMC-S-017. The SC shall provide automatic dropout (shedding) of PL and bus nonessential loads to address low voltage and over-current conditions. The SC shall provide overcurrent protection for the SC components, subsystems and PLs. The SC shall monitor bus voltage and shall report this value as part of normal SC/SV telemetry. The SC shall provide a common, single-point ground for all SC and PL electrical elements. Electromagnetic Compatibility (EMC) grounding practices shall be in accordance with (IAW) MIL-STD-1542B Section 5.1.3.

5.5 Thermal Control Subsystem

The SC’s thermal control subsystem shall maintain and control SC temperatures within required operational and survival temperature limits during ground handling, shipping, testing, launch, and all on-orbit mission phases for all orbital conditions. The SC shall provide PL thermal interface provisions as specified in Section 6.2 of this document.

5.6 Allowable Emissions and Susceptibility

The SC shall maintain conducted and radiated emissions and susceptibility within limits set by MIL-STD-461G.

5.7 Bonding

SV bonding shall comply with MIL-STD-1542B Section 5.6.

5.8 Material Selection

All materials used in the construction of the SC shall be low outgassing with < 1% Total Mass Loss (TML) and < 0.1% Collected Volatile Condensable Material (CVCM) when tested under conditions of ASTM E595.

5.9 Corrosion Resistance

All materials used in SC construction shall be corrosion resistant IAW MIL-STD-1568D.

5.10 Mechanism Design

Debris resulting from any moving mechanical assemblies shall be contained.

5.11 Venting Provision

The SC shall provide for outgassing of internal SC and PL hardware and shall direct all venting areas away from the PL suite.

5.12 Attitude Determination Control

SC attitude knowledge shall be provided to +/- 0.03 deg (3 σ) accuracy in each axis, including all sources of alignment error, during nominal spacecraft operation through End-of-Life (EOL) and also sufficient to support the attitude requirements of the PL. The SC shall collect SC attitude data and include it in State of Health (SOH) and recorded PL telemetry sent to the ground. The SC shall be capable of sending attitude data to each experiment at least once per second. The SC shall send attitude data sent to each experiment; the data shall include a time stamp with an accuracy of 100 msec with respect to Universal Time Coordinated (UTC).

SC attitude control shall be provided to +/- 0.1deg (3 σ) in each axis during normal mission operations through EOL. High frequency motion (jitter) shall be held to within 0.01deg (3 σ) from peak to peak over frequencies between 1 and 25Hz. The SC shall be capable of autonomously acquiring a safe mission attitude upon separation from the LV. The SC shall be capable of maintaining a fixed axis in the direction of the SC velocity vector for normal operations. The SC shall provide the capability to slew to any commanded orientation, with any specific keep out zones defined in a DO. The SC shall be capable of slewing at an angular velocity of at least 2.0deg/sec in any axis. If the SC becomes disoriented or loses normal mission attitude at any time in the mission, the SC shall autonomously acquire and hold a safe attitude and power positive condition until it can be ground commanded to return to the normal mission attitude, or it shall autonomously reacquire the normal mission attitude without damage to the SC.

5.13 Command and Data Handling

The SC shall provide command and data handling (C&DH) capabilities utilizing a digital data bus for interface of the SC subsystem hardware and payload data.

Command Requirements The C&DH subsystem shall be responsible for command decryption, verification, execution, and storage. The SC shall be capable of immediately executing real-time commands by directing them to the appropriate interface. The SC shall be capable of storing commands and executing the command to within one second of the designated time. The SC shall report executed commands in the telemetry stream.

Data Storage Requirements The SC shall contain sufficient on-board memory to store PL data and sufficient on-board memory for SC attitude data and other SC data necessary for determining the health, status, and operation of the system at a minimum of a 48-hour period. Total SC on-board storage capabilities shall include a 100% margin on the 48 hours of SC data. The SC shall provide partitioning control for on-board SC data storage allotments for each PL contained within the PLRD. Additional information will be provided within each mission DO PLRD as required for each experiment set.

Data Latency The SC shall be capable of transmitting PL data within 24 hours of collection or as specified in the DO PLRD. Transmitted data shall include but not limited to SC health and status, timing, attitude, and ephemeris data.

Data Quality and Retransmission The SC shall allow for re-transmission of telemetry data, this mechanism shall permit the identification of missing or invalid telemetry blocks. The SC shall provide a mechanism for converting the identification of a missing or invalid block into a command requesting the re-transmission of that block. The SC shall retain all data for re-transmission, unless commanded otherwise from the ground, for a minimum of 48 hours. The SC shall discard data from oldest to newest. The minimum re-transmission size shall be a single telemetry block. The SC shall not need to examine the contents of any PL data block to perform re-transmission.

Data Formatting SC telemetry blocks shall be unambiguously formatted to permit identification of the data origin and the sequencing of the block within the (source) telemetry stream. It is permissible to have the telemetry reflect a global sequencing; however, the global and source sequencing (if both are present) shall be separate and distinct. Sequencing information shall be unchanged when the data is re-transmitted, allowing reconstruction of the original data stream.

5.14 Quasi-static Loads

The SC shall have the capability to withstand acceleration load factors (not including factors of safety) of 8.5 g’s LV axial direction and 8.5 g’s LV lateral direction applied simultaneously or as an alternate, 12 g’s applied individually in each axis. These loads must be applied at the SC’s C/G.

5.15 Random Vibration

The SC shall have the capability to withstand the random vibration environment following a proto-qualification or flight-proof testing strategy (3dB above acceptance for 2 minutes/axis (1 minute per axis to comply with flight-proof testing)) as defined within SMC-S-016.

5.16 Acoustic

The SC shall have the capability to withstand a launch acoustic environment, which may be demonstrated using vibration testing as defined within SMC-S-016.

5.17 Shock

The SC shall have the capability to withstand the maximum predicted mechanical shock environment for the mission.

5.18 Mechanical Fundamental Frequency

The SC and separation system shall meet a minimum first fundamental frequency of 35Hz in both the LV axial and lateral directions when supporting the required DO PL mass.

5.19 Avionics

Command Execution On-board SC software shall be capable of executing stored commands. SC software shall be capable of timing command execution to within one second.

Timing The SC shall maintain an on-board time reference time synchronization avionics capability to within 100 msec with respect to UTC and shall provide this data at least once per second to each experiment.

Processor Loading SC software peak processor loading (including process execution, interrupt, and interface handling) shall not exceed 70% of the processor capability.

Processer Requirement On-board SC command/telemetry software/firmware shall be capable of being updated and/or modified while on orbit through ground commands. This shall include the capability to update a specific memory address and verify to the ground that the memory reload is correct.

6. PAYLOAD INTERFACE REQUIREMENTS

6.1 Mechanical Interface

SC mechanical interface(s) to PLs shall provide mechanical support, and alignment with respect to the SV attitude reference for all experiments in the PL suite defined in the DO PLRD. Mechanical support shall be provided to attach and secure the PLs to the SC in compliance with SMC-S-016 and guidance TR-2004(8583)-1. The SV shall provide a contiguous payload suite volume allotment capable of providing as much as required in the DO PLRD. The SV shall be capable of supporting both exterior and interior mounted PL experiment components. The SV shall be capable of supporting, under all loading conditions, a PL suite center of mass with a distance meeting common LV requirements. The SV shall accommodate a PL suite where the moments of inertia about the SV longitudinal axis will fall within common LV requirements. The SV shall not interfere with a hemispherical field of regard from the PL external mounting locations relative to the nadir direction for the nominal flight attitude as defined by the SV provider. The SV shall not interfere with a hemispherical field of regard relative to the ram direction. The SV shall provide a means for aligning all instruments in the PL suite to within 0.01 deg (3 σ in all axes) with respect to the SV attitude reference.

6.2 Thermal Interface

The SV shall provide a means for transporting heat to and from all PLs within the bounds described in this document in order to maintain acceptable operating and survival temperature limits during all mission phases. The SV shall maintain the SC to PL interface minimum and maximum model temperature prediction (as defined in SMC-S-016) between –9 deg C to +39 deg C. The SV shall be capable of transferring heat away from the PL suite at a maximum heat transfer rate dependent upon SV Bus size and PL power. The SV shall monitor the PL bulk temperatures at the SV to PL interface and shall provide this data to each PL and in SV telemetry.

6.3 Power Interface

The SV shall be powered off during integration and launch. The SV shall power on and begin SV initialization following SV separation from the LV. Following SV initialization, a 12U CubeSat SV shall supply an orbit average power (OAP) of no less than 60W to the PL suite through EOL or an ESPA Class SV shall supply an OAP of no less than 100W to the PL suite through EOL. Other SV sizes shall supply an OAP of no less than 75W to the PL suite through EOL. Each DO PLRD will include the OAP requirement from initialization through EOL. The SV shall provide power lines and connectors for the PL(s) defined in the PLRD. The SV shall monitor the voltage and current provided to each PL with eight-bit resolution and shall report these values in SV telemetry. The SV shall be capable of automatically shutting off power to any PL(s) if the current exceeds a pre-specified limit as described in the DO PLRD. Duty cycle scenarios will be determined in cooperation between the SVC and STP on a preliminary basis during PL accommodation design and will be coordinated on a monthly frequency after on-orbit characterization of the SV power generation capacity. The SV shall provide a common, single-point ground for all SV electrical elements and for each of the PLs.

6.4 Data Interface

The SV shall be capable of supporting a minimum of two PLs. The SV shall support analog, discrete, and / or digital interfaces to the PLs (example RS-422 / RS-485). The SV shall allocate supplied data interfaces between PLs. The DO PLRD defines all PL data interface requirements. Interfaces shall be capable of simultaneously operating at maximum capability without adversely affecting the SV or any of the PLs under normal operating conditions.

The SV shall support a minimum of 12 analog outputs from the PL suite to the SV. These outputs shall be capable of being sampled at least once per second. Each analog output shall be capable of being resolved to a resolution of at least 12 bits. The SV shall be capable of providing an associated time tag for each analog output, accurate to at least 100 msec.

The SV shall provide a minimum of 10 input bi-level discrete signals to the PL suite. The SV shall provide a minimum of 9 discrete output bi-level discrete signals from the PL suite to the SV. All discrete signals from the SV to each PL shall be capable of transmission within 100msec accuracy of the intended transmission time. The SV shall support time tagging for both input and output discrete signals to a minimum accuracy of 100 msec. The SV shall be capable of separately time tagging each discrete signal.

The SV shall provide a minimum of one digital interface for each individual PL. The SV shall provide for data transfer capability at a sustained rate of 200 kbps for each individual PL or as defined in the DO PLRD. The SV shall be capable of sustaining consistent data rates at all times during nominal operations. Experiments will be responsible for buffering data to maintain data rates at or below the SV supported rate. The SV shall be the controller. The SV shall provide appropriate software synchronization in accordance with the needs of the individual PLs or as defined in the DO PLRD. Desired synchronization shall include acknowledgement of the transfer and receipt of data across the SV to PL data interface. The SV shall detect and appropriately respond to communication faults with the PLs. At a minimum, these faults shall include interface failure and invalid message. The SV shall communicate with the PL at a frequency of at least once per second, when the PL is in normal operating mode. Each PL will indicate to the SV when it is in normal operating mode or standby mode.

The SV shall be capable of transferring data from the ground and the SV to the appropriate PL via the digital interface(s), this will be specified further in the DO PLRD. The SV shall convert the transferred data from uplink format to the format specified in the mission specific SC-to-PL ICD. There is no requirement for the SV to interpret PL commands. The SV shall support the capability to transfer SV data (including but not limited to time and/or position) to each PL.

The SV shall be capable of transferring data from each PL to the SV and to the ground. The SV shall convert the data from the PL to the appropriate format for on-board storage or downlink as necessary. There is no requirement for the SV to interpret PL data. There is no requirement for the SV to inject SV data into the PL data stream. Formatting of the SV data to the individual experiments shall be governed by the mission specific SC-to-PL ICD. Processing of the analog and or digital data streams shall not adversely affect the SV or any of the experiments under any nominal operating condition. Payload side connectors will be determined as documented in the individual experiment to SC-to-PL ICDs.

6.5 Functional Test

The functional testing includes validation of PL connectivity to the SV. Testing shall include mechanical (fit-check), power and data flow between the SC and PL.

6.6 Payload Mounting Provisions

SV mounting interfaces shall be provided as described in the SC-to-PL ICD. SV mounting surfaces shall provide unobstructed field of view for the PL.

7. STANDARD REQUIREMENTS – SPACE VEHICLE

7.1 Environmental Testing

The environmental testing shall include vibration, separation/pyro shock, thermal vacuum, thermal balance, alignment and mass properties in accordance with SMC-S-016 and TOR-2011(8591)-2 vol 1 and 2. The SV shall be subject to Electromagnetic Interference/ Electromagnetic Compatibility (EMI/EMC) testing of the SV in a fully functional configuration. The SVC shall verify use the appropriate inspection, demonstration, test, and analysis techniques to verify all requirements. Tailoring shall be in accordance with the DO PLRD.

SV testing shall be IAW flight-proof methodology as defined in SMC-S-016. This testing shall verify, at a minimum, the capability of the SV to survive vibration, separation/pyro shock, thermal vacuum, thermal balance, alignment, and mass properties experienced throughout all mission phases. Pressure and leakage tests are only required if the vehicle contains a propulsion system or other pressurized subsystems or components. EMI/EMC testing shall be IAW MIL-STD-461G.

Functional Testing The SV shall be subjected to and successfully pass thorough functional testing prior to integration of the experiments. Tests of mechanisms shall be in accordance with TOR-2004(8583)-1. Tests of mechanisms shall verify non-interference, proper lubrication, adequate torque margins, and strength under worst case loading as defined within SMC-S-016. Environmental testing of the SV including EMI/EMC, vibration, separation/pyro shock, thermal vacuum, thermal balance, alignment and mass properties is included in this effort. Functional testing shall be completed following transportation to the LV integration site. SV and PL engineering units and models of all deployable structures shall be used as applicable. SV structural and thermal testing shall validate models used for design and analysis to justify the flight qualification status of the vehicle as defined within SMC-S-016.

Component Environmental Testing Identical SV units and components previously qualified to flight levels and successfully flown; may be accepted for space flight upon approval by the Government (STP). Documentation of the proposed SV and component qualification status and flight worthiness shall be provided to the Government for approval. Units without previous higher qualification shall be subject to proto-qualification as defined within SMC-S-016.

Quasi-static Loads Acceleration load factors (not including factors of safety) for ESPA Class weighing 140 kg to 180 kg (308 lbs. to 397 lbs.) are 8.5 g’s LV axial direction and 8.5 g’s LV lateral direction applied simultaneously or as an alternate, 12 g’s applied individually in each axis may be used. These loads should be applied at the C/G of the SV using the maximum PL mass as specified. Individual SV components shall be able to withstand the acceleration loads specified in a mass acceleration curve to be provided by LV contractor. SC weighing <140 kg (308 lbs.) shall be tested to loads defined in the mass acceleration curve for SV mass.

Random Vibration The SV shall be tested to withstand the random vibration environment following a flight proof testing strategy (3 dB above acceptance for 1 minute/axis) as defined within SMC-S-016. SMC-S-016 defines acceptance levels to be an envelope of the maximum predicted environment (MPE). If the MPE of the LV is unknown, the random vibration acceptance levels shall be derived based on the SV mass following the National Aeronautics and Space Administration (NASA) General Environmental Verification Standard (GEVS) (GSFC-STD-7000A). Strategies for force limiting specific resonant frequencies of primary structures can be permitted with government review and approval.

Acoustic SV acoustic testing shall be IAW SMC-S-016 and TOR-2011(8591)-2 vol 1 and 2, for systems with a vibration response due to direct acoustic excitation.

Shock The maximum predicted mechanical shock environment will be developed in cooperation between the SVC, STP, and STP’s PL customers.

Thermal Vacuum and Bake-Out SV thermal vacuum testing shall include a minimum of four cycles with full functional tests conducted at the high and low extremes of the first and last cycles in accordance with SMC-S-016. The SV shall undergo a thermal bake-out in a vacuum of 10-4 torr at an agreed upon temperature and duration as specified in the DO SC-to-PL ICD to minimize the outgassing rate of the SV. Options include vacuum for four hours at 70 deg C or eight hours at 50 deg C.

7.2 Electronic Compatibility Test

Emissions and susceptibility testing shall be IAW MIL-STD-461G procedures. All tailoring of this standard will require approval by STP.

7.3 Safe Hold

The SV shall be capable of going into a safe hold mode under anomalous conditions and survive without intervention after the beginning of normal flight operations.

7.4 Disposal

The SVC shall comply with United States Government Orbital Debris Mitigation Standard Practices 2019 (ODMSP) and the DO.

7.5 Storage

The SC/SV shall be capable of being stored up to six months while awaiting launch.

8. STANDARD REQUIREMENTS – LAUNCH INTEGRATION

8.1 Electrical Interface

Electrical interconnections and cables from the LV to the SV electrical interfaces and separation systems will be provided to the SVC. Electrical cable harness connector part numbers will be decided by agreement between the LV provider, SC provider, and STP on a mission specific basis. Electrical signals between the SV and LV shall include primary and redundant separation signals.

8.2 Launch Vehicle Separation System

The SV’s separation system shall be compatible with the LV-to-SV ICD. The separation system shall be capable of providing a safe separation (no re-contact during separation) from the LV, including means to prevent premature separation. The separation system shall accommodate any stay-out zones as defined by the LV provider. Debris resulting from all separations shall be contained.

8.3 Space Vehicle to Launch Vehicle Integration

SV to LV integration shall be conducted in accordance with this document and an ICD agreed to and signed by STP, the SVC, and the LV provider. If applicable, discrepancies between this document and the LV-to-SV ICD will be resolved by STP.

8.4 Launch Pad Testing and Battery Charging

The SV shall be in a safe condition at all times prior to launch and during all LV operations. All launch site operations shall be compliant with the required range safety documents. All ground operations shall meet the requirements of AFSPCMAN-91-710 Volume 3 as applicable to the launch site to include power on and deployment inhibits.

8.5 Launch and Early Orbit Operations

The SV shall be capable of being launched in an unpowered state. The SV shall be capable of maintaining autonomous operation in a safe hold for up to 48 hours after orbit insertion. All SV deployments and RF transmissions shall not occur within 10 minutes from separation of the LV. Additional requirements shall be called out in the LV-to-SV ICD. The SV shall be subjected to and successfully pass functional check-out testing prior to the end of LEOP.

8.6 Ground Support Equipment

GSE, including any PL unique GSE, shall perform the functions required to: inspect, test, operate, evaluate, calibrate, measure, assemble, disassemble, handle, transport, safeguard, store, service, repair, and maintain the SV during all phases of ground operations at the SVC’s facilities, test sites, LV integration site and launch site.

9. STANDARD REQUIREMENTS – GROUND/MISSION OPERATIONS

9.1 Ground Segment Availability

The ground segment shall have a system availability of 95% for real-time satellite contact functions to include data storage, in any sliding 30-day window. The ground segment shall be available a minimum of 22.6 hours out of 24 hours per day (averaged over one month) with no loss of capability.

9.2 Ground Segment Architecture

The SVC’s ground segment architecture shall consist of uplink, downlink, and data transfer in addition to the following:

· Be compliant with DoDIs 5000.02T, 8310.01, 8320.07, 8330.01, 8581.01, 4630.09, 4650.01, and 3222.03 in addition to CNSSP-12 and NIST SP 800-53 Rev 5.

· Demonstrate compliance with applicable IT standards, protocols, and interfaces for the sharing of DoD data, information, and IT services.

· Demonstrate compliance with performance specifications.

· Ensure interoperability and electromagnetic compatibility.

10. SECURITY AND COMMUNICATIONS SECURITY (COMSEC) REQUIRMENTS

10.1 Regulation

SVC shall comply with processing DoD CUI (NIST SP 800-171 revision 2 and, NIST SP 800-172) and classified Secret DoD information (DoD NISPOM). The SVC shall comply with DCSA AO accreditation.

10.2 Encryption Requirements

Command uplink and data telemetry downlink shall be capable of implementing encryption per National Security Agency (NSA) guidelines. SVC shall acquire NSA (Type 1 or Type 2) approved equipment or request GFE encryption/decryption equipment that will be made available IAW the DO PLRD. SVC shall be responsible for secure storage and installation of keying material required to operate the space vehicle, and the SVC's command and control ground segment equipment and flight hardware. The SVC shall coordinate with the appropriate government COMSEC custodian to obtain key at the appropriate classification level IAW the DO PLRD.

APPENDIX A – DEFINITIONS

Capability: The ability to complete a task or execute a course of action under specified conditions and level of performance.

Command: Deliver instructions to the space vehicle systems in order to configure, reconfigure the space vehicle and or the space vehicle’s subsystems to include payloads.

Electrical Interface Panel: The electrical interface between a space vehicle to a payload.

FlatSat: The FlatSat is a high fidelity electrical and functional representation of the spacecraft bus. It is a high-fidelity test bed for Integration and Test, flight software, and flight operations.

Flight Proven: Spacecraft has successfully completed test, demonstration, and 365 days of mission operations including payload operations.

Functional Test: A functional test is one that varies the set of direct mechanical or electrical input stimuli and observes the resulting output end states or signals. Output end states or signals are compared to defined requirements for verification that the system passes a functional test. Functional tests are performed before and after the system undergoes a stress event (e.g., random vibration testing, shock testing, transportation) to ensure that the system has not been damaged causing anomalous behavior. In some instances, functional tests are performed during a stress event (e.g., thermal vacuum testing).

Ground Segment: All the ground-based elements of a Space vehicle system used by operators and support personnel, to include facilities, equipment, software, utilities, and infrastructure necessary to support management of the spacecraft, telemetry, tracking, control, and distribution of payload data. Primary elements consist of ground stations, mission control centers, ground networks, remote terminals and spacecraft integration and test facilities.

Integrated Payload Stack: The Integrated PL Stack consists of all the SVs delivered to space by the launch provider and integrated together with their associated adapters and interfaces.

Integration: Is the activity required to bring the PL and the SC together to operate as a unified flight vehicle. Integration may include early integration studies. In addition, integration is the activity to bring the SV and LV together to enable space lift.

Interface: A boundary or point common to two or more similar or dissimilar command and control systems, sub-systems, or other entities against which or at which necessary information flow takes place.

Issue: A problem that has been realized and must be addressed in order to achieve program goals.

Item Unique Identification: Item Unique Identification (IUID) is a DoD program to identify and track government-furnished property (GFP) through the use of Unique Item Identification (UII) in transaction-derived data from electronic business transactions. IUID applies to all items for which the government’s unit acquisition cost is five thousand dollars or more; items for which the government’s unit acquisition cost is less than five thousand dollars, when identified by the requiring activity as DoD serially managed, mission essential or controlled inventory; when the government’s unit acquisition cost is less than five thousand dollars and the requiring activity determines that permanent identification is required; regardless of value for (a) any DoD serially managed subassembly, component, or part embedded within an item and, (b) the parent item that contains the embedded subassembly, component or part.

Launch Vehicle: A integrated hardware system that includes structural elements, propulsion, guidance and control, electrical power, tracking, telemetry and control, ordnance, flight termination, software, payload fairing, and payload separation initiation, designed and produced to responsively deliver payloads to orbit. Description includes: dedicated small launch, on a small to heavy rocket as a rideshare PL on an integrated PL stack, or via a rideshare service via a launch broker

Launch and Early Operation Phase: The period encompassing LV ascent, SV separation, SV turn on, deployment of satellite systems (e.g., solar arrays, antennas, etc.), maneuvers to the mission orbit, and space vehicle on-orbit checkouts. On-orbit checkouts are used to assess the space vehicle and payload functionality and states-of-health and prepare for the nominal experiment operations phase.

May: Denotes a non-mandatory…