GEO_Attachment_3_417-R-GIRD-FOR-STUDY_Solar_Instrument_v3.pdf

Interface Requirements for Solar Instrument

Requirements Terminology The term “shall” designates a requirement that must be achieved and is synonymous with the term “threshold.”

The term “should” designates a desired level of performance the government would like the contractor to strive towards achieving and is synonymous with the term “goal.”

All other terms, including “will”, only designate statements of fact or intentions of the government and are not to be interpreted as contractor requirements.

An instrument may comprise more than one physical assembly, or unit. “Sensor unit” refers to the unit that contains the optics. “Instrument unit” means the sensor unit, electronics box (if applicable), or other units of the instrument.

Documents Applicable Documents ECSS-E-50-12A - Space Wire - Links, Nodes, Routers and Networks, 24 January 2003 European Cooperation for Space Standardization (ECSS)

ISO/DIS 14644-1 - Cleanrooms and Associated Controlled Environments, May 1, 1999

CCSDS 133.0-B-1 Space Packet Protocol, Blue Book, Issue 1, September 2003

IEEE/ASTM SI-10 - American National Standard for Use of the International System of Units (SI): The Modern Metric System, December 2002

417-R-RPT-0027 - The Radiation Environment for Electronic Devices on GOES-R Series Satellites, August 14, 2006

CCSDS 301.0-B-3 - Time Code Formats. Blue Book. Issue 3, January 2002

417-R-RPT-0050 - GOES-R Reliable Data Delivery Protocol Version 2.1 January 16, 2008

Requirements General Requirements On-Orbit Concept

Instrument-to-Spacecraft Science Data Rate The Solar Instrument science and engineering data rate, including all overhead associated with Consultative Committee for Space Data Systems (CCSDS) packetization by the instrument at the spacecraft interface, shall not exceed 3.5 million (106) bits per second.

Average Operational Power The Solar Instrument average operational mode power, including operational heater power, shall be no more than 180 Watts.

Mass Properties The mass of the Solar Instrument, including thermal blankets and cabling between units, shall not exceed 75 kg.

The length of the cable between the Solar Instrument sensor unit and the Solar Instrument electronics box shall not exceed 3 meters.

Volume

The stowed volume envelopes for the Instrument Sensor Unit, inclusive of all mounts, thermal blankets, and connectors, shall not exceed the dimensions provided below.

Height: 50 cm

Length: 100 cm

Width: 50 cm

Sensor Field of View The spacecraft shall provide the Solar Instrument sensor unit with a continual, unobstructed, and glint-free 10 degree full-width East-West and 10 degree North-South view of the Sun and centered on the Sun during normal on-orbit operations, as shown in Field-of-View Figure below. The FOV is measured from instrument sensor envelope's edges at the +X face. In this figure, rotations about the Z and Y are indicated by North-South and East-West, respectively.

x

5 deg y

5 deg x

5 deg

5 deg z

Sun-facing sideNorth-South FOV

East-West FOV

Field of View Figure

The satellite will operate in a geosynchronous orbit (Semi-major axis of approximately 42,164 Km) located at either 75° or 137° west longitude. Normal on-orbit operations entail periodic station keeping maneuvers that keep the satellite within a 0.2° inclination about the equator and within ±0.2° of the on-station longitude.

The instrument will be in survival mode during launch.

Coordinates The orbit reference frame (ORF) shall be defined as follows:

The ORF is orthogonal and right-handed.

The ORF origin is at the spacecraft center of mass.

The ORF +z axis points toward the center of the Earth.

The ORF +y axis points along the negative orbit normal.

The ORF +x axis completes the triad.

The body reference frame (BRF) shall be defined as follows:

The BRF is orthogonal and right-handed.

The BRF is fixed to the body of the spacecraft.

The location of the BRF origin will be specified by the spacecraft contractor.

The BRF axes are nominally parallel to the ORF axes when spacecraft attitude is in its nominal Earth-pointing, upright yaw attitude with zero attitude error.

The roll, pitch, and yaw axes are defined to be parallel to the BRF x, y, and z axes, respectively.

If there is a yaw-flip, the spacecraft will be flown upright (+Y BRF pointed in the +Y ORF direction) during northern hemisphere winter and inverted (+Y BRF pointed in the -Y ORF direction) during northern hemisphere summer. i.e., the +Y BRF axis is generally in the same hemisphere as the Sun.

If there is no yaw-flip, the spacecraft will be flown upright all year.

The reference coordinate system of each instrument unit shall be nominally parallel to the spacecraft BRF coordinate system, with the exception of solar-pointing instruments.

Yaw Flip Instruments shall meet all performance requirements whether or not the spacecraft performs a yaw flip.

Interface Requirements All instrument-to-spacecraft interfaces shall be single fault tolerant.

Mechanical The instrument contractor will provide all kinematic mounts, plus vibration isolation and thermal isolation mounting hardware.

Sun-Pointing Accuracy The spacecraft shall point the Instrument Sun-Sensor line-of-sight to within 1.5 arcminutes, 3-sigma, of the center of the Sun in each of the E-W and N-S directions during any contiguous 60 minute period of the Operational Sun-pointing Time.

Sun-Pointing Stability The spacecraft shall point the Instrument Sun-Sensor LOS to the Sun such that the Instrument Sun-pointing jitter during any contiguous 60-second period of Operational Sun-pointing Time is less than 6.0 arcseconds, 3-sigma.

Thermal Thermal Control Concept The instrument units installed on the spacecraft bus fall under one of the following categories:

a) Thermally-independent units are conductively and radiatively decoupled from the spacecraft and reject their heat directly to space.

b) Thermally-coupled units dissipate their heat to the spacecraft.

In general:

• Instrument electronic units are thermally-coupled.

• Instrument sensor units are thermally-independent.

Independent Unit - Net Heat Transfer For Independent Units, the net heat transfer averaged over the instrument independent unit interface plane area shall be less than 15.5 watts/m².

Coupled Unit - Heat Transfer Flux Density For conductively-coupled units, peak local heat transfer fluxes conducted to the spacecraft shall be less than 0.25 watts per square centimeter.

Interface Temperatures For planning and preliminary design purposes, the interface temperature (spacecraft side) for Earth-viewing instruments shall be:

a) 0°C to 40°C during operation

b) -30°C to 50°C during non-operation

Instrument Electrical Power Electrical Power Interfaces The spacecraft shall supply functionally independent, redundant operational power buses to each instrument for normal instrument operation.

Power Definitions The following definitions shall be used when calculating average power, maximum power, and peak current values.

Average Power (Operational) - The total power into an instrument averaged over any 5 minute period.

Average Power (Survival) - The total power into an instrument averaged over any 72 minute period.

Maximum Power (Operational or Survival) - The total power into an instrument averaged over any 20ms period.

Peak Current (Operational) - The maximum (never to be exceeded) current drawn by the instrument.

Peak current is typically calculated as 1.5 times the maximum power at worst case nominal voltage (26 Vdc).

The instrument shall operate in accordance with the instrument performance specification within the range of 28 volts Vdc ± 2 volts at the instrument operational power input connector.

Instrument Operational Power Grounding The instrument electrical power grounding shall be in accordance with the Electrical Grounding Figure.

Electrical Grounding Figure

Data Transfer Between the Instrument and Spacecraft All data transferred between the instrument and the spacecraft shall use the European Cooperation For Space Standardization (ECSS) ECSS-E50-12A (Space Wire) standard through the packet layer as a minimum.

Guaranteed Delivery All data transferred between the instrument and the spacecraft shall provide guaranteed data delivery as defined in GOES R Reliable Data Delivery Protocol document.

Source Packet Format All data transferred over the SpaceWire shall use the CCSDS 133.0-B-1 Section 4.1 Protocol Data Unit

Pulse Per Second (PPS) The spacecraft shall provide the instrument a 1 PPS time code sequence accurate to ±10 microseconds relative to UTC.

Spacecraft Primary

Bus Unit(s)

Instrument Operational Power A Instrument Operational Power A Instrument Operational Power A Instrument Operational Power A

Instrument Operational Power B

Instrument Operational Power B Instrument Operational Power B

Instrument Operational Power B

Spacecraft Structure

Single Point

Ground Spacecraft Structure

Instrument Survival Heater Power A Instrument Survival Heater Power A Instrument Survival Heater Power B Instrument Survival Heater Power B Instrument Survival Heater Power A Instrument Survival Heater Power A Instrument Survival Heater Power B Instrument Survival Heater Power B

Low Impedance Bond Connections

Spacecraft Secondary

Bus Power Unit(s)

Primary Bus

Voltage

Instrument Converter

Control and Monitoring The spacecraft shall provide remote access to all critical telemetry and control.

Critical telemetry is defined as telemetry points that are required to monitor the instrument in powered off state.

Non-critical telemetry is defined as telemetry points that are required to monitor the instrument in powered on state.

Engineering telemetry are data required to process instrument sensor data to higher level products.

Housekeeping telemetry are data required to monitor instrument operation, health, and safety.

Environmental Conditions On-Orbit Radiation Environment The instruments and spacecraft shall comply with the on-orbit radiation requirements that are described in the GSFC document 417-R-RPT-0027 titled "The Radiation Environment for Electronic Devices on the GOES-R Series Satellites."

Launch Environment The instruments and spacecraft shall be designed to meet the launch environment described herein.

The baseline launch vehicle is planned to be an expendable SpaceX or Atlas V.

Eclipse Solar Eclipse shall be considered as part of the environmental variation. The Solar Eclipse season occurs twice yearly, with each eclipse season lasting approximately 45 days. The maximum eclipse duration is 72 minutes.

Attitude and Orbit Data All attitude, rate and orbit data shall be included in the spacecraft ancillary data packet.

Attitude Knowledge The spacecraft shall provide a periodic attitude estimate to the instrument.

Accuracy The attitude knowledge static error shall not exceed 1200 microradians, 3-sigma, per axis, pre-calibration.

The attitude knowledge diurnal error shall not exceed 45 microradians, 3-sigma, per axis. Diurnal error refers to that portion of the error that repeats from day-to-day.

The attitude knowledge dynamic error shall not exceed 30 microradians, 3-sigma, per axis. Dynamic error includes all non-static and non-diurnal errors.

Update Rate The spacecraft shall update the attitude estimate at a rate no less than 1 Hz.

Latency The attitude estimate latency shall not exceed 100 milliseconds.

Spacecraft Angular Rate The spacecraft shall provide a periodic angular rate estimate to the instrument.

Bandwidth The spacecraft angular rate estimate shall have a minus 3dB bandwidth of greater than 25 Hz.

Spacecraft Orbit The spacecraft shall provide a periodic spacecraft orbit estimate to the instrument via the ancillary data packet.

The spacecraft position estimate shall be accurate to 100 meters 3-sigma, in-track (ORF x-axis), cross-track (ORF y-axis) and radial (ORF z-axis) directions.

The spacecraft velocity estimate shall be accurate to within 6 cm/sec, 3-sigma, per axis.

Update Rate The spacecraft shall update the orbit estimate at a rate no less than 1 Hz.

Instrument-to-Spacecraft Disturbances Instrument Disturbance Torque Limits At any time during the Operational Sun-pointing Time, the sum of the magnitude of the Instrument uncompensated torques and the magnitude of its uncompensated linear forces multiplied by a lever arm of 2 meters shall not exceed 0.2 N-m.

Instrument Allowable Angular Momentum The magnitude of Instrument's uncompensated angular momentum shall not exceed 0.1 N-m-sec.

Instrument Sun-Pointing Data to Spacecraft The Instrument shall send two axis Sun-pointing attitude data to the spacecraft.

The Sun-pointing data shall be provided with a resolution of .25 arcsec, per axis or better.

The Sun-pointing data shall be accurate to 2 arcsec, 3-sigma, per axis, or better.

The Sun-pointing sensor shall have a bandwidth greater than 50 Hz.

Instrument shall provide sun-pointing data to the spacecraft with data rate and the noise equivalent angle (NEA) relationship shown in the Figure provided below.

Data Rate - Noise Equivalent Angle Relationship Figure

0.05

0.1

0.15

0.2

0.25

0.3

0.35

0.4

0.45

0 5 10 15 20 25 Data Rate (Hz)

NE

A

(a rc se c)

, 1 -R

M S

The Sun-pointing sensor shall have a latency less than 0.04 sec.

The sampling period for Sun-pointing data shall be uniform to 0.005 sec or better.

The Sun-pointing sensor shall have an acquisition field of view of +/- 24 arcmin, per axis.

The Sun-pointing sensor shall have a central linear field of view of +/- 95 arcsec, per axis.

Acronyms %AC percent area coverage

A Ampere(s)

ASD Acceleration Spectral Density

BOL beginning of life

BRF body reference frame bps bits per seconds

C Celsius (Degrees) cm centimeter db decibel dc Direct Current

ECSS European Cooperation for Space Standardization

EED Electro-Explosive Device

EEE Electrical, Electronic and Electromechanical

ESD electro static discharge

FOR field of regard

FOV field-of-view g Earth’s gravitational acceleration

GIRD General Interface Requirements Document

GOES Geosynchronous Operational Environmental Satellite

GSE Ground Support Equipment

GSFC Goddard Space Flight Center

Hz hertz

ICD Interface Control Document

IDD Instrument Description Document

IRU inertial reference unit

ISO International Office for Standardization

K kelvin kg kilogram

Km kilometer kPa Kilo Pascals

LOS line-of-sight

LSB Least Significant Bit m meter ma milli-ampere(s)

Mbps Mega bits per seconds

MHz Mega hertz mm millimeter

MAT Mission Allowable Temperatures

MEFL Maximum Expected Flight Level m-g milli-g (Earth’s gravitational acceleration)

MLI Multi-layer insulation

Ms milli-seconds

NASA National Aeronautics and Space Administration

N-m Newton-meter

N-m-sec Newton-meter-second nPa Nano Pascals

NVR Non-Volatile Residue

ORF Orbit reference frame

OSR Optical solar reflectors

PPB Parts Per Billion

PPM Parts Per Million

PPS Pulse Per Second

PSD power spectral density

QCM Quartz Crystal Microbalance

RIU Remote Interface Unit

SCF SPP Coordinate Frame

SI International System of Units

SINDA System Integrated Numerical Differential Analyzer

SIS Solar Imaging Suite

SPP Sun-Pointing Platform

SRS Shock Response Spectra

TBD to be determined

TBR to be refined/reviewed

THC Total Hydrocarbons

UIID Unique Instrument Interface Document

UTC Universal Time Code µsec mico-seconds

V Volts

Vdc Volts-Direct Current

W Watts

Interface Temperatures