GEO_Attachment_1_417-R-GIRD-FOR-STUDY_Imager_v3.pdf

Interface Requirements for Earth Imager 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 MIL-STD-461E Aug 99 - Requirements for the Control of Electromagnetic Interference Characteristics of Subsystems and Equipment

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

NASA/TM-2001-211221 - Guideline for the Selection of Near-Earth Thermal Environment Parameters for Spacecraft Design, October 2001

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

ISO/DIS 14644-3 Cleanrooms and associated controlled environments - Part 3: Metrology and test methods, 9/26/02

SN-C-0005 Revision D, Contamination Control Requirements, July 20,1998

Requirements General Requirements On-Orbit Concept

Instrument-to-Spacecraft Science Rate The 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 66.6 million (106) bits per second.

Average Power Operational The Instrument shall draw operational average power of no more than 400 watts.

Mass Properties The Instrument, including all units and cabling between units, shall have a mass less than 330 kilograms.

Volume The Instrument Sensor Unit, including mounts, thermal blankets and connectors for the stowed configuration, shall have dimensions that do not exceed the limits provided below.

Height: 1500 millimeters

Length: 1200 millimeters

Width: 1750 millimeters

Optical Port Field-of-View The spacecraft shall provide the Instrument Sensor Unit’s optical port a clear field-of-view within 65o of nadir as shown in the following figure.

Sun Shield

Sun Shield

65o 65o

Nadir

Optical Port

65o 65o

Sun Shield

Sun Shield

65o 65o

Nadir

Optical Port

65o 65o

North Field-of-View The spacecraft shall provide the Instrument Sensor Unit’s -Y face a 2π steradian field-of-view to space excepting a field-of-view blockage with less than 1.0W (TBR) back loading.

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.

Instruments shall survive an anomaly resulting in a static instrument line-of-sight (LOS) such that the sun passes through the LOS at orbit rate.

Instruments shall survive a spacecraft attitude anomaly resulting in the sun being at an arbitrary fixed location within the instrument field of regard (FOR).

The instrument will be in Survival Mode during launch.

Instruments shall survive a spacecraft attitude anomaly resulting in the sun sweeping through the field-of-view (FOV) of the instrument radiator from “horizon to horizon” at a rate of 6O/minute, passing through radiator normal.

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 For an instrument with passive cryogenic detector cooling, the GOES spacecraft will be rotated 180O about the Z-axis (yaw) twice per year to keep the -Y axis side of the spacecraft shaded, within ±4 days of the Sun crossing the orbit plane.

Instruments shall meet all performance requirements whether or not the spacecraft performs a yaw flip, except for instruments with cryogenic detectors cooled by a passive radiator.

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.

Nadir Pointed Instrument Alignment

Placement The placement of the instrument alignment reference frame with respect to the spacecraft IRU reference frame shall be to within 0.25 degrees per axis, including variation over all launch and on-orbit environments.

Initial Alignment Knowledge The prelaunch alignment knowledge of the instrument alignment reference frame with respect to the spacecraft IRU input axes shall be 300 microradians or better per axis.

Alignment Rate of Change The rate of change of the alignment of the instrument mounting frame with respect to the spacecraft IRU input axes shall not exceed 100 microradians per hour per axis. This requirement includes on-orbit environments and spacecraft structural and thermal stability.

Attitude and Disturbances for Nadir Pointed Instruments Attitude Error The attitude error of the instrument mounting frame relative to the desired ORF-referenced attitude shall not exceed 360 microradians, 3-sigma, per axis. Attitude error is defined as the difference between the desired attitude and the actual, or true, attitude of the instrument mounting frame.

The instrument mounting frame attitude shall be stable to within 500 microradians, peak-to-peak, per axis, over any 60 sec window.

Attitude Rate Error For a given total spacecraft gyro signal phase delay, the instrument mounting frame attitude rate error relative to the desired ORF-referenced attitude shall not exceed ± the corresponding limit from the Total Spacecraft Gyro Signal Phase Delay and Rate Error Trade Space Figure provided below, per axis, when the rate is filtered by a fourth order Butterworth low pass filter with a -3dB frequency of 15 Hz.

Spacecraft Translation Acceleration Limits The translational accelerations at the spacecraft side of each instrument sensor unit mount shall not exceed the limits specified in the Translational Acceleration Limits for Spacecraft to Instrument Table below. The limits apply to each orthogonal axis after the acceleration is bandpass-filtered using at least a fourth-order band-pass Butterworth filter with -3dB frequencies of f1 and f2. Note that the filter has fourth-order rolloff on both sides of the response. The accelerations can be present at any combination of the instrument sensor unit mounts and along any combination of the three orthogonal axes at each mount.

Allowable spacecraft phase delay and rate error trade space

Rate Error Limit 100 µrad/sec

Phase Delay Limit 35 msec

Allowable spacecraft phase delay and rate error trade space

Rate Error Limit 100 µrad/sec

Phase Delay Limit 35 msec

Total Spacecraft Gyro Signal Phase Delay and Rate Error Trade Space Figure

Translational Acceleration Limits for Spacecraft to Instrument Table f1

(Hz) f2

(Hz)

Peak Limit (mg) f1

(Hz) f2

(Hz)

Peak Limit (mg) f1

(Hz) f2

(Hz)

Peak Limit (mg)

0.0 0.9 6.3

20.2 64.0 203.2 9.0 9.5

10.1 10.7 11.3 12.0 12.7 13.5 14.3 15.1 16.0 17.0 18.0 19.0 20.2 21.4 22.6 24.0 25.4

512.0 10.1 32.0 101.6 322.5 512.0 10.1 10.7 11.3 12.0 12.7 13.5 14.3 15.1 16.0 17.0 18.0 19.0 20.2 21.4 22.6 24.0 25.4 26.9 28.5

18.44 1.50 1.00 4.29 10.55 15.16 0.40 0.40 0.40 0.40 0.40 0.40 0.40 0.40 0.40 0.40 0.40 0.40 0.40 0.40 0.40 0.40 0.40 0.40 0.40

26.9 28.5 30.2 32.0 33.9 35.9 38.1 40.3 42.7 45.3 47.9 50.8 53.8 57.0 60.4 64.0 67.8 71.8 76.1 80.6 85.4 90.5 95.9 101.6 107.6

30.2 32.0 33.9 35.9 38.1 40.3 42.7 45.3 47.9 50.8 53.8 57.0 60.4 64.0 67.8 71.8 76.1 80.6 85.4 90.5 95.9 101.6 107.6 114.0 120.8

0.40 0.40 1.40 1.40 1.40 1.40 1.40 1.40 1.40 1.40 1.40 3.18 3.18 3.18 3.18 3.18 3.18 1.40 1.40 1.40 1.40 1.40 1.40 1.40 2.40

114.0 120.8 128.0 135.6 143.7 152.2 161.3 170.9 181.0 191.8 203.2 215.3 228.1 241.6 256.0 271.2 287.4 304.4 322.5 341.7 362.0 383.6 406.4 430.5 456.1

128.0 135.6 143.7 152.2 161.3 170.9 181.0 191.8 203.2 215.3 228.1 241.6 256.0 271.2 287.4 304.4 322.5 341.7 362.0 383.6 406.4 430.5 456.1 483.3 512.0

2.40 2.40 2.40 1.40 1.40 1.40 3.02 3.02 3.02 3.02 2.56 2.56 2.56 2.56 2.56 2.56 2.56 2.56 2.56 2.56 2.56 2.56 2.56 2.56 2.56

(CCR 00163)

The translational accelerations at the spacecraft side of each instrument sensor unit mount shall produce an absolute peak acceleration Shock Response Spectra (SRS) less than the limits set in the On Orbit Operational SRS Acceleration Limits Table below. The limits apply to each orthogonal axis for SRS natural frequencies when using a quality factor, Q, of 50. The SRS is computed after the acceleration is high pass filtered with a fourth order Butterworth filter with a -3dB cut off at 1.0 Hz. Use a logarithmic interpolation for both frequency and acceleration terms in the On Orbit Operational SRS Acceleration Limits Table.

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

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

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.

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.

This requirement together with the next two define separate static, diurnal and dynamic attitude knowledge components. These requirements bound the knowledge error, which is the difference between the estimated attitude and the true attitude.

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.

Accuracy The integrated rate error of the spacecraft angular rate estimate shall not exceed 1 microradian, 3-sigma, in the x and y axes, and 1.5 microradians, 3-sigma, in the z-axis over any 1 second window.

The integrated rate error of the spacecraft angular rate estimate shall not exceed 2 microradians, 3-sigma, in the x and y axes, and 2.5 microradians, 3 sigma in the z-axis, over any 30 second window.

The integrated rate error of the spacecraft angular rate estimate shall not exceed 7 microradians, 3-sigma, per axis, over any 300 second window.

The integrated rate error of the spacecraft angular rate estimate shall not exceed 20 microradians, 3-sigma, per axis, over any 900 second window.

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

Update Rate The spacecraft shall update the angular rate estimate at a rate no less than 100 Hz.

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

Accuracy 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 At any time during the operational mode of the spacecraft, the sum of the magnitude of the instrument sensor unit’s uncompensated torques and the magnitude of its uncompensated linear forces multiplied by a lever arm of 2 meters shall not exceed 1.0 N-m.

The magnitude of the instrument unit’s uncompensated angular momentum shall not exceed 1.0 N-m-sec.

Integrated Angular Momentum The absolute value of the high pass filtered angular momentum integrated with respect to time shall not exceed the values listed in Filtered Integrated Angular Momentum Limits table. The filter is a second order Butterworth filter with a -3 dB cutoff frequency at 0.01 Hz. The spacecraft center of mass will be within 2 meters of the centroid of the sensor unit mounts.

Angular Momentum Limits The absolute value of the high pass filtered angular momentum produced by the sensor unit referenced at the spacecraft center of mass shall not exceed the values listed in the Angular Momentum Limits table. The high pass filter is a second order Butterworth filter with a -3 dB cutoff frequency at 0.01 Hz.

The spacecraft center of mass will be within 2 meters of the centroid of the sensor unit mounts.

External Sourced Translation Acceleration Limits The translational accelerations responses from sources external to the instrument sensor unit at the spacecraft side of each instrument sensor unit mount shall not exceed the limits specified in the

External Source Translational Acceleration Limits for Spacecraft to Instrument Table below. The limits apply to each orthogonal axis after the acceleration is filtered using at least a fourth-order band pass Butterworth filter with -3dB frequencies at f1 and f2 and with a fourth order roll off. Note that the filter has a fourth-order roll-off on both sides of the response.

External Source Translational Acceleration Limits for Spacecraft to Instrument Table

(Hz) f2

Peak Limit (mg)

Peak Limit (mg)

Peak Limit (mg)

0.0 512.0 6.71 26.9 30.2 0.18 114.0 128.0 0.63

0.9 10.1 0.67 28.5 32.0 0.18 120.8 135.6 0.63

6.3 32.0 0.45 30.2 33.9 0.63 128.0 143.7 0.63

20.2 101.6 1.34 32.0 35.9 0.63 135.6 152.2 0.63

64.0 322.5 3.13 33.9 38.1 0.63 143.7 161.3 0.63

203.2 512.0 6.26 35.9 40.3 0.63 152.2 170.9 0.63

9.0 10.1 0.18 38.1 42.7 0.63 161.3 181.0 0.63

9.5 10.7 0.18 40.3 45.3 0.63 170.9 191.8 0.63

10.1 11.3 0.18 42.7 47.9 0.63 181.0 203.2 0.63

10.7 12.0 0.18 45.3 50.8 0.63 191.8 215.3 0.63

11.3 12.7 0.18 47.9 53.8 0.63 203.2 228.1 0.63

12.0 13.5 0.18 50.8 57.0 0.63 215.3 241.6 0.63

12.7 14.3 0.18 53.8 60.4 0.63 228.1 256.0 1.00

13.5 15.1 0.18 57.0 64.0 0.63 241.6 271.2 1.00

14.3 16.0 0.18 60.4 67.8 0.63 256.0 287.4 1.00

15.1 17.0 0.18 64.0 71.8 0.63 271.2 304.4 1.00

16.0 18.0 0.18 67.8 76.1 0.63 287.4 322.5 1.00

17.0 19.0 0.18 71.8 80.6 0.63 304.4 341.7 1.00

18.0 20.2 0.18 76.1 85.4 0.63 322.5 362.0 1.00

19.0 21.4 0.18 80.6 90.5 0.63 341.7 383.6 1.00

20.2 22.6 0.18 85.4 95.9 0.63 362.0 406.4 1.00

21.4 24.0 0.18 90.5 101.6 0.47 383.6 430.5 0.63

22.6 25.4 0.18 95.9 107.6 0.47 406.4 456.1 0.63

24.0 26.9 0.18 101.6 114.0 0.47 430.5 483.3 0.63

25.4 28.5 0.18 107.6 120.8 0.63 456.1 512.0 0.63

(CCR 00162) (CCR 02039D)

The translational acceleration responses from sources external to the instrument sensor unit at the spacecraft side of each instrument sensor unit mount shall produce an absolute peak acceleration Shock Response Spectra (SRS) less than the limits set in the External Source On Orbit Operational SRS Acceleration Limits Table below. The limits apply to each orthogonal axis for SRS natural frequencies when using a quality factor, Q, of 50. The SRS is computed after the acceleration is high pass filtered with a fourth order Butterworth filter with a -3dB cut off at 1.0 Hz. Use a logarithmic interpolation for both frequency and acceleration terms in the External Source On Orbit Operational SRS Acceleration Limits Table.

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

Acronyms