CALL 001 Atch 1 STP-SIV Users Guide.docx

Space Test Program – Standard Interface Vehicle

(STP-SIV)

Payload User’s Guide

15 Jun 2008

BATC DWG 2258656 REV C

A6356_BR2281[QBS] Copies Are Uncontrolled BR 2281 9/10/04 DISTRIBUTION STATEMENT A: Approved for public release; distribution is unlimited.

CONTENTS

FIGURES

TABLES

1. PURPOSE

The purpose of this space test program-standard interface vehicle (STP-SIV) User’s Guide is to provide payload developers with the mission capabilities of the STP-SIV spacecraft and to document the standard interface provided. The information provided herein should provide the necessary detail for a payload provider to determine if STP-SIV is a candidate for their mission and to complete preliminary design work.

SCOPE

This document defines the support accommodations for a “generic payload” by the STP-SIV spacecraft bus. It defines the interface between the payload and the spacecraft (mechanical, electrical, data, thermal), the envelope of performance the spacecraft provides for the payload, the constraints that the payload must design within, an envelope of operational capability for the payload, documentation requirements and the integration and test philosophy of the STP-SIV program as it pertains to the payload. Experiments manifested on the STP-SIV bus are required to comply with the interface defined in this document unless provided with a waiver from the STP-SIV program office. For each STP-SIV spacecraft bus and payload suite, a payload unique interface control document is generated to document the specific interfaces required by each payload.

This document also provides additional information for designing a payload to operate from an orbiting platform, including information on space environment and operational considerations. This document does not specify payload performance or design criteria that affect only the individual payload reliability or performance. Designing to the interface defined herein ensures that the payload is compliant with the standard interface and that payload faults do not jeopardize the spacecraft or other payloads.

The STP-SIV spacecraft is designed to host up to four payloads on each mission. Given that many payloads may be competing for limited resources such as mass, volume and power, this document provides design guidelines that, when implemented, make a payload easier to manifest as part of a payload suite. Compatibility with a wide range of payloads and mission profiles increases the likelihood of being manifested on a STP-SIV spacecraft, especially for payloads that are flown as secondary missions. These guidelines are presented throughout this document, identified with a (GL), and summarized in Appendix C.

The use within this document of the terms "shall", "will", "must", "should" or "is" does not explicitly confer responsibility onto any particular entity or organization. A thorough reading of the document is necessary to comprehend and extract all applicable payload or spacecraft bus interface requirements. A compliance matrix is provided in Appendix B to identify paragraphs in this document that contain design information that requires payload verification versus those that may be taken as information only. A recommended verification methodology is also identified in the matrix.

0. Roles and Responsibilities The following roles and responsibilities have been established for the STP-SIV program.

0. Ball Aerospace & Technologies Corp (BATC) BATC has overall responsible for ensuring that the STP-SIV spacecraft bus is designed, manufactured and tested in compliance with the requirements found in the STP-SIV (ESPA Class) Technical Requirements Document (TRD). These requirements cover both the payload interface defined herein and space vehicle performance. BATC is responsible for generating and maintaining the payload specific interface control document (ICD) and integration of the individual experiments onto the STP-SIV spacecraft bus, conducting system-level testing, integrating the SV on the launch vehicle and initializing the SV for on-orbit operation.

BATC leads the Payload Integrated Product Team (IPT). BATC provides review of payload interface verification activities and communicates status and issues to STP.

Payload Supplier The payload suppliers are responsible for delivering a flight payload that is compliant to the STP-SIV standard interface and fulfilling the requirements in the Memorandum of Agreement between them and the DoD Space Test Program (STP). The process typically begins with the payload supplier generating an Experiment Requirements Document (ERD) and completing DD-1721 Flight Request form. This form along with the STP Experimenter’s Guide can be requested from the Space Development and Test Wing. Once a payload is manifest, the payload supplier is responsible for working with BATC to generate the interface control document and develop mission CONOPS. The Payload IPT works with payload suppliers for all the manifested payloads to address conflicts between payloads that may arise. The payload supplier provides support during SV testing at BATC for all payload related activities and post launch operations at the RSC. It is the payload supplier’s responsibility to ensure the payload satisfies all interfaces defined in the spacecraft to experiment ICD.

Space Test Program (STP) STP Mission Space Test Program is a Department of Defense (DoD) activity under Air Force (AF) executive management. It is the primary spaceflight provider for the entire DoD space science & technology community. STP supports the pursuit of basic science of potential benefit to the DoD, and the reduction of mission risks inherent in technology transformation with respect to space control, situational awareness, communications, geolocation, surveillance and weather.

The Space Development and Test Wing (SDTW) manages STP. STP advances space system technology by providing spaceflight and related services for experiments generated within the entire DoD space science and technology community. Services provided by STP include flight requirements development, mission design, manifesting assistance, spacecraft design, fabrication, test, experiment integration, launch vehicle (LV) integration, launch services, ground and flight safety support, flight support, orbital operation, and transmittal of data to PIs for analysis/dissemination. STP provides these services to approved Space Experiments Review Board (SERB) payloads. Non-SERB payloads can also obtain the same services on a reimbursable basis.

The complexity of experiments flown by STP spans a wide range, from basic research through advanced technology payloads. The result of these experimental space demonstrations is substantial risk mitigation at relatively low cost for new space systems or sub-systems technologies, concepts, and designs before they are incorporated into operational systems.

Payload Manifest Process The payload suppliers begin the process of becoming manifested on STP-SIV either through the SERB, or for non-SERB payloads, by contacting STP mission design (MD). MD conducts initial payload survey to identify specific mission opportunities. If the payload is acceptable for an STP-SIV mission, STP requests BATC to further evaluate the compatibility of the payload with the spacecraft and any other payloads that may be manifested or planned for a particular vehicle. If the payload is compatible, STP manifests the payload. STP generates MOA with PI’s organization and the PI begins the integration process with STP and BATC. Additional detail on the manifesting process is documented in the Space Test Programs Experimenter’s Guide.

Interested payload suppliers should contact STP directly Commercial: (505) 853-5505

DSN: 263-5505

Email: dodstp.rideshare@kirtland.af.mil

Delivery Requirements

The data and hardware delivery requirements for payloads of the SIV space vehicles are summarized in Table 1.

Table 1 Data and Hardware Delivery Requirements

SV CDR – 2 mo P/L delivery

SV CDR – 2 mo P/L delivery

SV CDR – 2 mo P/L delivery

Payload inputs to ICD between SC and PL

SV CDR – 2 mo

Flight Payload and GSE required for system level tests

Inputs to system level test procedures

These are for initial planning purposes only. The final delivery dates are mutually agreed to by the STP Program Office, the payload provider, and BATC and documented in the program master schedule. Once the baseline schedule has been approved, change requests are submitted to the STP program office for approval.

Definitions For STP-SIV, the term Spacecraft (SC) refers to the space segment with no payloads. A payload is the space-borne hardware or software components of an experiment supplied to BATC as Government Furnished Equipment (GFE) through STP. The payload suite is the combination of all payloads on the mission. The Space Vehicle (SV) is the SC integrated with the payload suite.

Acronyms See Appendix F: Acronym List for a complete list of acronyms

APPLICABLE DOCUMENTS

The following documents form a part of this user’s guide to the extent specified herein. Unless a specific issue or revision is listed, the referenced document shall be of that issue or revision in effect on the date of the release of this document. In the event of a conflict between documents referenced in Section 4 and the contents of this document, the contents of this document shall apply.

0. Government Documentation Space Test Program (STP) Experimenter’s Guide – November 2004

0. Specifications and Standards

BATC Supplied Documentation No additional BATC Documents.

SYSTEM OVERVIEW

STP-SIV Characteristics STP-SIV is being developed to offer a small spacecraft optimized to host multiple SERB payloads. The spacecraft provides an agile, repeatable ESPA-Class spaceflight capability for DoD space technology demonstrations. The spacecraft has a design life of 13 months. The missions are primarily low Earth orbit at a variety of inclinations. The spacecraft is compatible with many launch vehicles to maximize launch opportunities as a secondary payload.

STP-SIV has a non-proprietary Standard Payload Interface (mechanical, thermal, data and power) for use by all experiments hosted on STP-SIV (ESPA Class) program spacecraft. The government has rights to this interface, allowing use on future DoD STP missions. The expectation is that non-recurring costs through the life of the STP-SIV program are reduced by using a repeatable spacecraft design and standard payload interface.

The space vehicle consists of two major elements: the payload suite and the spacecraft. Figure 1 illustrates the space vehicle with the standard payload interface. The space vehicle communicates with the mission operations control center via the AFSCN. Mission operations take place at the RSC at Kirtland, AFB in New Mexico.

Figure 1. STP-SIV provides a standard payload interface

Coordinate Systems There are three coordinate systems that the payload supplier should become familiar with.

2. SC Flight Coordinate System The SC flight coordinate system is defined in the nominal nadir pointing mode as Origin: located at the center of the LV interface ring and at the plane where the interface ring attaches to the spacecraft structure.

+ZFL: Nadir direction +YFL: Orbit normal, opposite the direction of angular momentum vector +XFL: In direction of velocity, completing the right-handed triad with +YFL and +ZFL

As described in section 5.2.2, the spacecraft periodically performs a ‘yaw-flip’ which places the +XFL in the opposite direction of the space vehicle velocity.

Figure 2. Spacecraft Flight Coordinate System

Payload Coordinate System The payload coordinate system is defined as Origin: located at the intersection of the spacecraft longitudinal axis with the mounting surface of the payload interface plate. The spacecraft longitudinal axis is centered on the LV interface ring.

+ZPL: coincident to the +ZFL axis of the SC flight coordinate system +XPL: parallel to the +XFL axis, offset in the +Z direction to lie in the plane of the payload mounting interface +YPL: parallel to the +YFL axis, offset in the +Z direction to lie in the plane of the payload mounting interface

Figure 3. Payload Coordinate System

SC Mechanical Coordinate System The SC mechanical system is primarily applicable to the spacecraft but is included here for reference. It is orientated to be consistent with the ESPA coordinate system. The payload coordinate system is the mechanical reference for the payloads. The SC mechanical coordinate system is defined as Origin: located at the center of the LV interface ring and in the plane separating the ESPA from the LV interface ring +XSC: In the direction of separation from the ESPA +YSC: Towards the nose of the EELV when the SC is mounted to the ESPA +ZSC: Completes the right-handed triad with +XSC and +YSC

Figure 4. Spacecraft Mechanical Coordinate System

Functional Block Diagram The functional block diagram of the interfaces between the payload and the spacecraft is shown in Figure 5.

Figure 5. Payload Interface Block Diagram

MISSION DESIGN

This section provides SC performance capability to include the range of operational orbits, SC modes and performance of the guidance, navigational & control (GN&C) subsystem.

0. Orbits The STP-SIV spacecraft is designed to operate in orbits between 400 km and 850 km altitude and 40° and 98.8° inclination.

GL – Payloads with a wide range of orbit possibilities are easier to manifest, especially as a secondary. Even primary payloads should have flexibility because ESPA launches are designed to put the primary space vehicle into the required orbit. Flexible orbit requirements from the payload leave open more launch possibilities.

3. Supported Orbit Types The STP-SIV spacecraft is designed for low earth orbits. Without a propulsion system, many of the specialized orbits such as sun-synchronous orbits, repeat-ground track orbits and frozen orbits cannot be maintained over the 1 year mission. Depending on the altitude, some of the benefits may still be realized, especially for shorter durations of the mission. BATC coordinates with the payload providers to determine what orbits best meet the objectives of the payload suite.

0. Elliptical Orbits The STP-SIV supports elliptical orbits so long as the orbit perigee remains above 400 km and the orbit apogee does not exceed 850 km. If it is desired to keep the apogee at the same latitude during the mission, an inclination of 63.4° needs to be selected to prevent apsidal rotation. If a different orbit inclination is selected, the orbit rotates causing the latitudes of apogee and perigee points to vary. The rate of the rotation depends on the orbit altitudes and inclination.

Sun-Synchronous Orbits A sun-synchronous orbit combines orbit altitude and inclination so that the space vehicle passes over any given point of the Earth's surface at the same local solar time. The orbit precesses at approximately 1 deg per day so the orbital plane maintains a constant angle with the earth-sun line. Figure 6 shows the inclination required for a particular altitude to achieve this precession rate. STP-SIV does not have a propulsion system so a true sun-synchronous orbit cannot be maintained. As atmospheric drag lowers the orbit altitude, the orbit precesses from the desired equatorial crossing. The angular displacement of the ascending node varies depending on solar activity and the initial altitude.

Figure 6. Altitude vs. inclination for sun-synchronous orbits

Figure 7. Annual precession of near-sun-synchronous orbit without propulsion

Flight Orientations and Pointing Modes Launch and Early Orbit Operations The SC is launched in an unpowered state. After the spacecraft separates from the launch vehicle, it goes through an initialization process to bring the avionics on line and deploy the solar arrays. It then goes into a safe attitude until after the initial ground contact. This ensures a stable thermal and power environment for the bus and payloads. The payload suite is provided a minimum of 20 watts of power for survival heaters during this phase of the mission. Initial payload checkout occurs during this initial phase. Launch and Early Orbit operations completes within 30 days of launch.

Nadir Pointing Mode (Operational) The Nadir Pointing Mode is the primary mission mode in which all flight systems are fully operational and the spacecraft is capable of supporting routine payload operations.

The SC points the payload mounting surface (+ZPL) towards nadir and the +XPL is maintained either in the velocity or anti-velocity direction. This ‘yaw-flip’ is performed periodically to avoid star-tracker outages due to interference from the sun and the resultant increase in attitude errors. The frequency of the ‘yaw flip’ is dependent on the mission orbit, but does not occur more frequently than once every 30 days.

Figure 8. Spacecraft Orientation in Nadir Pointing Mode

Sun Pointing Mode (Operational) The sun pointing mode is an operational mode in which all flight systems are fully operational and the spacecraft is capable of supporting routine payload operations. The SC is capable of pointing the payloads toward the sun for up to 5 minutes at least once per month.

Inertial Pointing Mode (Operational) The inertial pointing mode is an operational mode in which all flight systems are fully operational and the spacecraft is capable of supporting routine payload operations. The SC provides the capability to slew to any commanded orientation for up to 5 minutes and no more often than once per orbit. The orbit average power available to the payload suite is reduced by up to 30W during this pointing mode. This additional restriction must be taken into consideration during mission planning.

Safe Hold Mode Safe Hold mode is entered when spacecraft bus or computer faults prevent normal payload operations. While in the safe hold mode, the spacecraft turns off non-essential loads and points the solar array to achieve a positive power balance over each orbit. Payloads are turned off. Power to the payload suite cannot exceed 50W total for survival heaters. Once the safe orientation is achieved, the spacecraft nominally maintains an attitude that keeps the payloads +ZPL axis greater than 90 deg from the sunline. However, during the transition, payloads may be pointed towards the sun for short periods of time and should be designed to survive a sun pointing attitude.

Spacecraft Guidance, Navigation and Control This section of the user’s guide provides the payload supplier with the expected performance and accuracy of the spacecraft attitude and orbit ephemeris. The spacecraft provides a status message to each payload that includes spacecraft attitude, ephemeris and time. The format of the status message is described in 6.4.2.4.

3. Orbit Position Knowledge The spacecraft has a GPS receiver that computes both position and velocity. This ephemeris information is included in the spacecraft status message. The position error is less than 25 m (95th percentile) and the velocity error is less than 2 m/s (95th percentile). As a backup to GPS, the spacecraft is also equipped with a SGLS transponder that provides ranging measurements to support ground-based orbit determination and an on-board propagator to provide spacecraft ephemeris to the payloads. The accuracy of this ground based solution is dependent on many factors but 500 meters after three days is achievable.

Attitude Knowledge The spacecraft attitude is provided to the payloads as part of the spacecraft status message. The attitude accuracy is within +/- 0.03° (3σ) in all axes. The attitude knowledge is with respect to the payload mounting surface and the payload coordinate system.

Attitude Control The spacecraft is capable of controlling the attitude to within +/- 0.1° (3σ) about any axis.

Attitude Maneuverability The spacecraft is capable of slewing up to 0.15 deg/sec about any axis.

Stability SC high frequency motion (jitter), measured at the payload interface, is less than 0.01 deg (3) from peak to peak during any period of one second or less.

Launch Vehicles STP-SIV is compatible with the EELV Secondary Payload Adapter (ESPA) on both the Atlas V and Delta IV, Pegasus, Minotaur I and Minotaur IV. This list of launch vehicles provides STP significant flexibility to find timely and cost-effective means to get payloads into space.

SV End of Life After the SIV mission is complete, the SV is configured to a safe condition and eventually re-enters the earth’s atmosphere due to natural orbital decay. To assess the re-entry footprint, the payload suppliers shall provide a complete list of materials that make up the payload components no later than two months prior to space vehicle CDR.

PAYLOAD INTERFACE REQUIREMENTS

This section of the user’s guide documents the standard payload interface. Designing the payload to the interface defined herein ensures low risk integration and prevents a payload failure from propagating to the spacecraft and other payloads.

0. Mechanical Interface Requirements This section of the user’s guide defines the mechanical interface between the spacecraft and the payloads.

4. Mechanical Interfaces Payload/Platform The STP-SIV spacecraft provides a standard payload mechanical interface (Figure 9). The payloads mount to the aluminum plate using #10 fasteners to holes on 2.54 cm (1 inch) centers. See Appendix D for a complete drawing of the payload interface plate (PIP). The payload supplier shall make all mounting points compatible with the standard mounting grid. Payload mounting locations on the PIP are determined by BATC to ensure requirements of each individual payload are satisfied.

Figure 9. Payload Interface Plate with Payload Coordinates

Payload Volume The allowable payload envelope in the launch configuration is shown in Figure 10. Payloads are mounted ‘on-top’ of the spacecraft and must fit within this volume for launch. Once the spacecraft is on orbit and deploys the solar arrays, payloads may deploy elements as necessary to perform their missions. All deployed elements shall remain above (+ZPL) the payload interface plane.

Figure 10. Payload envelope for launch configuration

Mass Properties Mass The spacecraft is capable of supporting up to 60 kg of payload mass. This mass is allocated between multiple payloads for most missions. Final payload mass shall be measured to +/- 1%.

GL – It is important to keep mass as low as reasonably possible. STP-SIV aims to fly multiple payloads on each mission so a lighter payload provides more flexibility in selecting complementary payloads, thus increasing flight opportunities for an individual payload.

Centering The payload suite must have a center of mass less than 4 cm radial distance from the longitudinal axis (Zpl axis) of the spacecraft and less than 22 cm above (+Zpl) the payload interface plane. These limits apply to both the launch and on-orbit configuration. The balancing of payload mass is performed by BATC at the payload suite and space vehicle level. When multiple payloads are flown on one SV, payload mounting locations are selected to keep the center of mass within limits while satisfying individual requirements of each payload. If balance mass is needed to bring the payload suite into compliance with this requirement, it is counted against the payload suite mass allocation. Although this is a system issue, there are several things an individual payload can do to maximize system level flexibility.

GL – Keep the footprint to a minimum. This provides more flexibility for arranging all payloads on the payload interface plate to achieve the required center of mass relative to the spacecraft longitudinal axis.

GL – Keep payload center of mass less than 22 cm above the payload baseplate and strive to keep it as low as possible.

The center of mass requirements has some flexibility for a payload suite with a total mass of less than the 60 kg maximum. Specific capability is assessed on a mission-by-mission basis.

The payload supplier shall provide the location of the center of mass of the payload flight unit, to an accuracy better than 0.5 cm in all three axes for both the launch and deployed configuration. The center of mass shall be determined both by analysis and by test.

Inertia The SC accommodates a payload suite where the maximum self moment of inertia about an axis perpendicular to the SC longitudinal axis is less than 3.37 kg*m2 and the maximum self moment of inertia about the SC longitudinal axis is less than 5.11 kg*m2.

The payload supplier shall provide the payload principle moments of inertia to accuracy better than 10% of the total inertia in both the launch and deployed configuration.

GL – Designing a compact box provides BATC additional flexibility for arranging payloads to satisfy the payload suite inertia limits.

Structural Loading Structural loading is evaluated for each SV, ensuring that the payload suite does not exceed qualification limits of the SV structure.

Stiffness Stiffness in Launch Configuration In the launch configuration, the payload primary lateral and axial resonances shall be greater than 75Hz. The primary torsional frequency in the launch configuration shall be greater than 75 Hz. This frequency shall be measured with the payload constrained in the same manner as mounted onto the spacecraft.

Stiffness in Flight Configuration If payloads have deployable assemblies, the minimum resonance frequency of these assemblies, in the deployed configuration, shall be greater than 7 Hz in any axis. Resonance frequencies lower than 7 Hz can be evaluated for compatibility with SIV on a case-by-case basis.

Field of View Payloads are provided an unobstructed field of view of 2 steradian, oriented towards the +ZPL (nadir) axis and originating at the payload interface plane. Additionally, the payloads are provided a 2 steradian unobstructed field of view towards the +XPL (velocity direction until seasonal yaw flip), originating 55 cm from the leading edge of the PIP in the +XPL direction. This provides nearly 3 steradians of unobstructed views for payloads as shown in Figure 11. Field of view obstructions from other payloads must be negotiated on a case-by-case basis.

Figure 11. Payload Field of View

Alignment STP-SIV provides the ability to align payloads to less than 0.1° (3σ) in each axis. To achieve this accuracy about the xPL and yPL axes, the payload mounting interface must be flat to less than 0.005 inches. To achieve the alignment about the zPL axis, the mounting holes must have positional tolerances of less than 0.010 inches. For payloads less than 5 inches on a side, additional considerations may be required. BATC works with the payload supplier to identify any necessary design requirements for payloads to meet less stringent alignment.

The spacecraft also provides the capability to measure the alignment between the spacecraft control frame and the payload boresight so the uncertainty in payload pointing, under ambient conditions, is known to less than 0.01° (3σ) in each axis. To achieve the 0.01° alignment knowledge, the payload must include an optical cube, optical flats, tooling balls, or similar alignment device that can be measured optically by a theodolite. The payload provider must also provide the alignment offset between the alignment device and the payload boresight. The spacecraft bus has alignment cubes on both the star tracker and the payload interface plate for mechanical measurement of the payload boresight to the spacecraft mechanical reference frame and the spacecraft flight reference frame. The alignment device must either be vacuum compatible or must have the capability to be removed and replaced for thermal vacuum testing and removed for flight. The payload alignment device must be located such that there is clear line of sight. BATC coordinates with all of the payload suppliers manifested for a given mission to ensure that the alignment devices are not visibly obstructed by other mission payloads. The agreed locations are documented in the spacecraft to experiment ICDs.

BATC performs the alignment measurement between the payload reference cube and the spacecraft bus alignment reference cubes during the integration and test phase of the program. Measurements are taken both before and after thermal testing.

Environments for Mechanical Design This section defines the mechanical environments the payload is exposed to during environmental testing of the space vehicle, during launch and on orbit. The space vehicle undergoes a protoqualification test program. Since the payloads are integrated to the spacecraft prior to testing, they should be designed with sufficient margin to ensure they survive and operate normally after being exposed to the environments described in this section.

This document does not specify payload specific design and test requirements except where such requirements are necessary to ensure the spacecraft or other payloads are not put at risk. It is expected that the payload organization determines payload requirements with knowledge of these environments in accordance with experiment objectives and risk tolerance. Section 8 of this document provides design and test recommendations that correspond with the requirements levied on spacecraft components.

0. Vibration Environment

Table 2 provides the maximum predicted environment (MPE) and protoqualification levels the payloads may be exposed to during space vehicle testing and launch. The table is applicable to components weighing less than 22.7 kg. The power spectral density may be reduced for payloads weighing more using the following calculations.

Where W is component weight in kilograms. The slope identified in Table 2 remains + and – 6 dB/oct.

Table 2. Payload random vibration levels

Shock Environment

Table 3 provides the maximum predicted environment (MPE) for the shock environment at the payload mounting interface that payloads may be exposed to during launch and space vehicle environmental testing. During test, the payloads are exposed to the protoqualification levels two times.

Table 3. Payload Shock Environment

Mass Acceleration

The mass acceleration curve (Figure 12 and Table 4) provides the load factors to ensure the payload can withstand the expected acceleration environment. For SIV, this load factor applies in each of the three orthogonal axes. The space vehicle does not undergo static testing.

Figure 12. Mass acceleration curve

Table 4. Mass Acceleration Load Factors

Acoustics

The expected acoustic environment is defined in Table 5. Acoustic testing is not performed during space vehicle environmental testing because the random vibration spectrum encompasses the expected launch acoustic environment.

Table 5. Expected Acoustic Levels 1/3 Octave Center Frequency (Hz) Sound Pressure Level (dB) re: 2x10-5 N/m2

Pressurized Structures Pressurized structures shall be designed to comply with the requirements of AFSPCMAN91-710.

Magnetic Design The payload may be exposed to a magnetic field of up to 65 A-m2. The payload may also be exposed to a 25 gauss demagnetization field for up to 5 minutes. For payloads that produce a magnetic dipole of greater than 0.1 A-m2, BATC may position compensating magnets, totaling less than 1 A-m2, to reduce the magnetic field to less than 0.1 A-m2.

Vacuum and Pressure Decay During launch vehicle ascent, the payloads may be exposed to a depressurization rate up to 5500 Pa/sec. Typically, analysis or testing is not necessary if the component has a vent area of at least 60 cm2 per cubic meter of volume.

Thermal Interface This section of the user’s guide defines the thermal interface between the spacecraft and the payloads.

4. Thermal Control of the Payload Interface SIV is designed to dissipate waste heat from the payloads through thermal radiators located on the spacecraft bus. The spacecraft maintains the temperature at the payload interface between -20°C and +50°C through all mission phases, regardless of the operational state of the payload suite, provided the total payload waste heat dumped into the payload mounting interface is less than 100W.

The heater design is composed Minco Kapton Strip Heaters, and High Reliability Honeywell Thermostats. There are separate switches for controlling nominal heaters, survival heaters. Figure 13 shows the heaters and thermostats in their approximate positions. The thermostats are represented by the small circles, the heaters by the long rectangles. The orange and blue circuit represents the nominal heater circuit; the red and black represent the survival heater circuit

Figure 13. Location of thermostats and heaters on PIP

Payload Thermal Monitoring The spacecraft monitors the temperature of the PIP using four temperature sensors mounted to the underside (-ZPL side) of the payload interface plate. Each payload receives the temperature read by these sensors once per second as part of the spacecraft status message. The spacecraft status message is described in section 6.4.2.4. The temperature measurements are also provided to the ground as part of spacecraft state of health (SOH).

The accuracy of the temperature measurements provided to the payload is +/- 8°C. Depending on the location of the payload relative to the temperature sensor, the accuracy of the reported temperature may be better. The location of the payloads relative to the temperature sensors is defined in the spacecraft to payload ICDs developed for each mission.

Payload Module Thermal Blankets BATC designs and manufactures thermal blankets for the payload module based on the thermal SV thermal analysis. The SV thermal model incorporates the provided payload thermal models. This allows the payload module blanketing to be designed from the system perspective. The blankets are grounded to chassis using fasteners attached to the PIP. The payload specific ICDs define applicable surface finishes on the payloads and materials used for the thermal interface to the SC.

Payload Power This section of the user’s guide defines the power available to the payloads during the different phases of the mission and during different spacecraft modes of operation. This section also describes the characteristics of the power feeds provided to each payload.

4. Mean Orbital Power During normal mission operations, the spacecraft provides a minimum 100W orbit average power (OAP) to the payload suite. This power is allocated between the different payloads on each mission. If the total power draw for all payloads combined exceeds 100W, it may be necessary to duty cycle payload operations in order to reduce the payload suite power consumption to within the capability of the spacecraft.

GL – Provide a minimum power standby mode.

GL – Design the experiment so continuous payload operation is not required.

Power Peaks Limitations The spacecraft supports a peak power to the payload suite of 100W. Any payload design exceeding this limit is evaluated on a mission specific basis.

Mission Power Profile Launch Phase The spacecraft is unpowered from launch until separation from the launch vehicle. After separation from the launch vehicle, the spacecraft begins an automated deployment sequence which includes turning on the spacecraft avionics, deploying the solar arrays and acquiring the safe mode of operation. Once this initial deployment sequence has completed, up to 20W of orbit average power may be provide to the payload suite through the completion of the launch and early orbit activation sequence at which time the full capabilities described herein become available.

Inertial Pointing Modes The spacecraft has the capability to point in any inertial attitude for up to 5 minutes per orbit. During these off-nadir activities, the payload orbit average power is reduced by up to 30W, depending on the required orientation of the spacecraft to support the payload operations. The peak power capability remains 100W, but duty cycling and thermal impacts are evaluated for specific payload designs.

Power during Safe-Mode If the spacecraft enters safe-mode, all non-essential spacecraft loads are turned off. Primary power to the payloads is also turned off. For payloads requiring their own heater circuits, up to 50W of power is available to the payload suite, using one of three available power lines that is dedicated to the payload specific heaters.

Power Feeds Each payload is provided three switched power feeds. Payloads which have multiple power modes of operation, may wish to divide up the three spacecraft power services among the differing modes. For example, if the payload has a nominal or idle state of operation and occasionally powers up focal plane arrays, the payload may dedicate a single spacecraft power service to that nominal mode and a separate power service to the focal plane electronics. For payloads with their own heater circuits, one of these three feeds must be dedicated to heaters only for them to be active during the reduced power modes associated with safe mode and launch and early orbit ops.

2. Voltage The spacecraft provides each payload 28 +6/-6 Vdc unregulated power from the spacecraft bus on each switched service.

Ripple The ripple voltage is less than 0.5 V peak-to-peak as measured over the bandwidth from DC to 10 MHz when the power system is delivering the maximum rated current into resistive loads for 28V payload inputs.

Voltage Transients The payload shall not generate voltage transients on the primary DC power lines in excess of the limits shown in Figure 14 when power is applied. This limit shall be applicable during turn on, turn off, and during all mode changes.

Figure 14. Voltage transient limits

Inrush & Capacitive Loading When spacecraft power is applied to the payload, the payload’s inrush current amplitude, as a function of time, shall not exceed the following limits:

a) Current rate of change less than 1 amp/microsecond

b) Maximum current amplitude less than 10 amps

c) The current returns to nominal operating current by 40 milliseconds

Fault Protection The spacecraft provides over-current protection on each of the power lines provided to the payload. Each switch providing power to the payload is equipped with a circuit to actively disable the switch if the current exceeds 5.94 amps. After the initial 40 milliseconds from switch turn-on, if payloads exceed the 5.94 amp limit for greater than 20 microseconds, the fault protection turns the power switch off, removing power from the payload. The current on these switches is also monitored by the spacecraft software. The software can be configured to remove power if the current exceeds a pre-determined current less than 5.94 amps.

Undervoltage The payload should design means to prevent permanent damage or degradation by application of any input voltage in the range of 0 to +34 volts for any time duration.

Raw Input Application and Removal The payload should design means to prevent being damaged during removal of power due to the opening of switches on the switched service supply.

Payload Data Interface The spacecraft Integrated Avionics Unit (IAU) functions as the main data and command interface between the payloads. All payload command, data collection and data storage is via the Payload Interface Board (PIB) which resides within the IAU. The PIB provides each payload one “Data Port” which consists of a 62 pin HDD connector on the payload module. All payload commands and collection of High Rate data, Real-Time data, analogs and discretes is through the single 62 pin Data Port. All payload data (high rate and real-time) is polled and ingested in a “Round Robin” scheme ensuring that no payload can monopolize the bus. Both the payload high rate and real-time data is time-stamped at the time of receipt.

The PIB ingests payload high rate mission data, encapsulates this data in a CCSDS compliant CADU format and stores the formatted CADU for subsequent transmission to the ground. All High Rate data is transferred via either an asynchronous UART EIA-422 link or a synchronous EIA compliant RS-422 link. The choice of synchronous or asynchronous data transfer method is selectable for each payload and is fixed prior to launch. Details of High Rate data transfers are provided in subsequent sections.

The PIB provides for collection of payload real-time data. One EIA-422 UART is provided on each payload data port. Payload real-time data is collected and interleaved into the real-time spacecraft downlink and is also stored on the PIB for retransmit. Each payload’s real-time data is limited to 240 bytes/sec.

The PIB provides to each payload data port; 8 analog inputs, 8 discrete inputs, and 8 discrete outputs. The term “input” or “output” used throughout this section is referenced from the spacecraft point of view (i.e. input refers to data from a payload into the spacecraft).The analog and discrete telemetry channels are sampled by the PIB once a second. Analog data is converted to a digital format using a 12 bit A-to-D converter. The PIB interleaves the discrete and analog data into the spacecraft real-time telemetry, in CP_PDU format.