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Mission Requirements
Document (MRD) Hierarchy |
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1 |
Science
Requirements |
4 |
Spacecraft Requirements |
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1.1 |
Science Observations |
4.1 |
Structural /Thermal |
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1.2 |
Data Capture &
Completeness |
4.1.1 |
Launch
Vehicle Accommodation |
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1.3 |
Angular Resolution &
Coverage |
4.1.2 |
S/C &
Instr Accom. |
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1.5 |
Precision, Acc, & Dyn
Range |
4.1.3 |
Thermal
Monitoring and Control |
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1.6 |
Cadence |
4.1.4 |
Instrument
Optical Bench |
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4.2 |
Attitude Control
& Determination |
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2 |
Mission
Implementation |
4.2.1 |
Acquisition |
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2.1 |
Orbit |
4.2.2 |
Pointing
Knowledge |
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2.2 |
Mission Life |
4.2.3 |
Attitude
Control and Stability |
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2.3 |
Environment |
4.2.4 |
Propulsion
& Delta-V |
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2.4 |
Launch Vehicle (LV) |
4.2.5 |
Momentum
Management |
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2.5 |
Mission Implem/ Ops
Concept |
4.2.6 |
Safehold |
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2.6 |
Standard Spacecraft
Services |
4.3 |
Power |
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2.7 |
Development Approach |
4.3.1 |
Power
Generation |
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4.3.2 |
Power
Storage |
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3 |
Instrument Requirements |
4.3.3 |
Power
Distribution |
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3.1 |
EVE |
4.3.4 |
Load
Shedding |
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3.1.1 |
Alignment,
Jitter & Stability |
4.3.5 |
Constraints |
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3.1.2 |
Spectral
Resolution |
4.4 |
Comm & Data
System |
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3.1.3 |
Timing |
4.4.1 |
S-Band
Communications |
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3.1.4 |
Data
Completeness |
4.4.2 |
Ka-Band
Communications |
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3.1.5 |
Interface
Requirements |
4.4.3 |
Data System
Functions |
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3.2 |
HMI |
4.5 |
HGA Assembly |
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3.2.1 |
Alignment
& Jitter |
4.5.1 |
Operation,
Pointing and Stability |
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3.2.2 |
Angular
Resolution |
4.6 |
Deploy Actuation and
Verif |
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3.2.3 |
Timing |
4.6.1 |
Solar Array
Deploy and Verif |
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3.2.4 |
Data
Completeness |
4.6.2 |
HGA
Deployment and Verification |
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3.2.5 |
Interface
Requirements |
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3.3 |
SHARPP |
5 |
Ground Segment Requirements |
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3.3.1 |
Alignment
& Jitter |
5.1 |
Integration and Test
(I&T) |
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3.3.2 |
Angular
Resolution |
5.1.1 |
High Rate
Science GSE |
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3.3.3 |
Timing |
5.1.2 |
Low Rate GSE |
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3.3.4 |
Data
Completeness |
5.2 |
Ground Station
Implementation |
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3.3.5 |
Interface
Requirements |
5.2.1 |
Dedicated
Site Requirements |
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3.3.6 |
Dynamic
Range |
5.2.2 |
Ancillary
Site Requirements |
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5.2.3 |
Commanding |
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5.2.4 |
Housekeeping
Telemetry |
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5.2.5 |
Tracking |
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5.2.6 |
Science
Telemetry |
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5.3 |
Mission Operations Center
(MOC) |
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5.3.1 |
Cmd and Tlm
Functions |
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5.3.2 |
Data Products |
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5.4 |
Science Operations Center
(SOC) |
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5.4.1 |
Ops |
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5.4.2 |
Archive |
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5.4.3 |
Data Products |
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APS- Antenna Pointing
System |
INSTR- Instruments (EVE, HMI, SHARPP) |
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C&DH- Command &
Data Handling Subsystem |
LV- Launch Vehicle |
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CONTAM- Contamination
Subsystem |
MATL - Materials and Process Subsystem |
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DEPLOY- Deployment
Subsystem |
MECH- Mechanical Subsystem |
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ELEC- Electrical Systems |
PARTS- Parts Subsystem |
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FLT DYN- Flight Dynamics
Subsystem |
PROP- Propulsion Subsystem |
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FSW- Flight Software
Subsystem |
PWR- Power Subsystem |
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GN&C- Guidance,
Navigation and Control Subsystem |
RAD- Radiation Effects Subsystem |
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GND- Ground Segment |
RF- RF Communications Subsystem |
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GSE- Ground Support
Equipment Subsystem |
THERM- Thermal Subsystem |
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Observatory = Combination
of Spacecraft flight segment and Instrument flight segments |
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Instruments = EVE, HMI,
& SHARPP Flight segments |
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SOC = Science Operations
Center |
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Ground
System = Command and control facility and equipment located in the Mission
Operations Center (MOC) |
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Ground Station = Remotely
located antenna site and data distribution facility |
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Ground Segment = All
ground elements, including ground system, Instrument SOCs, SDO ground
station, and any ancillary ground stations |
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# |
Modified prior to CCB |
CCR#
& Date |
Title |
Functional Requirement |
Performance Requirement |
Comments |
Subsystem Allocation |
Trace From |
Verify Method |
Verf. Lead |
Verf. Status |
Verf. Data Ref. |
Sig Appr |
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1 |
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Science Requirements |
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1.1 |
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Science Observations |
The
science mission shall perform solar observations sufficient to characterize
solar activity as the sun exits from a period of solar minimum and progresses
to a period of solar maximum |
N/A |
N/A |
N/A |
N/A |
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1.1.1 |
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The
SDO Observatory will launch during or shortly after solar minimum (June 2007
to August 2008 (TBR)) in order to permit observations as the solar cycle
progresses towards solar maximum |
Defines
launch requirement within a range that allows initial observations at end of
solar minimum |
Project |
Lev.1 [Science Objectives, Mission Timeline Success Criteria
in 4.1.1] |
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1.1.2 |
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Science
observations shall cover 5 years of the progression from solar minimum to
solar maximum to meet full mission success science requirements |
Comments:
Minimum success will be achieved with a series of observations spanning both
minimum and maximum conditions, ideally met with a 3-year observation period
(c.f. : Level 1 Science Requirements Document for full details). |
ALL |
Lev.1 [Science Objectives, Mission Timeline Success Criteria
in 4.1.1] |
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1.2 |
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Data Capture & Completeness |
The end-to-end system of the
SDO Instrument, Spacecraft, and Ground System shall obtain and deliver solar
observations to the Investigator’s SOCs of sufficient quality to achieve the
mission science objectives. |
N/A |
N/A |
N/A |
N/A |
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1.2.1 |
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The
end-to-end HMI Data Capture budget requires 95% of all possible science data
over the SDO mission, including delivery of these data to the SOCs. The EVE and SHARPP data capture budgets
require 90% of all possible science data. |
For
minimum mission success, the Observatory shall return 80% of SHARPP and EVE
data, while a completeness of up to 95% (depending on the campaign) is
specified in the HMI Science Observation Requirements table in the Level 1
Science Requirements Document. Note that HMI is the driver of this
requirement as well as the completeness requirement. Requirement addressed through the use of a
configured data capture budget. Is
this requirement quantifiable? |
INSTR, C&DH, GN&C, APS, RF, GND |
Lev.1 [Science Meas. 1 in 4.1.1] |
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1.2.2 |
3/27/03 |
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The
combined HMI Instrument, Spacecraft, and Ground System shall provide an
end-to-end data completeness of 99.99% over periods of minutes to hours.The
requirement for EVE is 99.6% and SHARPP is 99.9% (TBR). |
The
HMI observable construction requires minimizing data loss in order to
calculate Dopplergrams and magnetograms from a series of filtergrams. The
completeness value for EVE and SHARPP will be reviewed to ensure that this
value does not unnecessarily drive BER for the instruments. |
INSTR, C&DH, RF, GND |
Lev.1 [Science Meas. 1 in 4.1.1] |
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1.2.3 |
3/27/03 |
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Individual
solar observations shall consist of a minimum of 10
minutes of continuous observation. |
Requirement
addressed minimum HMI Dopplergram observation duration. Planned gaps should
be kept short and as non-periodic as possible. |
INSTR, C&DH, APS, RF, GND |
Lev.1 [Science Meas. 1 in 4.1.1] |
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1.2.4 |
3/27/03 |
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A
data interruption of more than TBD (2.5 minutes??) is considered a data gap
and results in the beginning of a separate science observation |
|
INSTR, C&DH, APS, RF, GND |
Lev.1 [Science Meas. 1 in 4.1.1] |
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1.2.5 |
3/27/03 |
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Data
interuptions less than that specified in 1.2.4 are considered a data gap if they are not
preceded by TBD times (10x??) the duration of the interruption both prior to
and after the event, thereby resulting in the beginning of a separate science
observation |
|
INSTR, C&DH, APS, RF, GND |
Lev.1 [Science Meas. 1 in 4.1.1] |
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1.3 |
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Angular Resolution & Coverage |
The
Instruments shall provide adequate angular resolution and image field of view
to meet the Level 1 Science Requirements |
N/A |
N/A |
N/A |
N/A |
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1.3.1 |
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|
Dopplergrams
shall cover the full disk with a sampling of about 0.5 arcsec per pixel, and
image quality and stability that enable a resolution of <1.5 arcsec. |
Drives
high frequency jitter and instrument resolution. Full disk image defines FOV,
absolute pointing, and data rate. Using the Rayleigh criterion for
diffraction limited resolution, 0.5 arcsec pixels would give 1.2 arcsec
angular resolution that would then increase for non perfect optics, pointing,
photon statistics, charge spreading, etc. |
HMI |
Lev.1 [Science Meas. 1 in 4.1.1] |
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1.3.2 |
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|
Longitudinal
and vector magnetograms shall cover the full disk with a sampling of about
0.5 arcsec per pixel, and image quality and stability that enable a
resolution of <1.5 arcsec. |
Drives
high frequency jitter and instrument resolution. Full disk image defines FOV,
absolute pointing, and data rate. Using the Rayleigh criterion for
diffraction limited resolution, 0.5 arcsec pixels would give 1.2 arcsec
angular resolution that would then increase for non perfect optics, pointing,
photon statistics, charge spreading, etc. |
HMI |
Lev.1 [Science Meas. 2 in 4.1.1] |
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1.3.3 |
3/27/03 |
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Magritte atmospheric images shall cover the sun out to 1.4 solar radii
with a sampling of 0.66 arcseconds per pixel |
Drives
jitter but not pointing. 1.32 arcsecond resolution stems from the requirement
that a resolved feature should be less than 2 pixels wide (0.66
arcseconds/pixel). |
SHARPP |
Lev.1 [Science Meas. 4 in 4.1.1] |
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1.3.3.1 |
3/27/2003
- New |
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Spectre
atmospheric images shall cover the sun out to 1.2 solar radii with a sampling
of 0.6 arcseconds per pixel. |
|
SHARPP |
Lev.1 [Science Meas. 6 in 4.1.1] |
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1.3.4 |
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Coronographic
images shall cover the sun from 2 to 15 solar radii with a sampling of 15
arcseconds per pixel |
Drives
pointing. 30 arcsecond resolution stems from the requirement that a resolved
feature should be less than 2 pixels wide (15 arcsecond/pixel) |
SHARPP |
Lev.1 [Science Meas. 6 in 4.1.1] |
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1.3.5 |
3/27/03 |
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Spectral
Irradiance measurements must consist of integrated disk measurements over a
field of view extending to 1.8 solar radii from the center of the solar disk. |
Ensures overlap
of EVE and SHARPP measurements. |
EVE |
Lev.1 [Science Meas. 6 in 4.1.1] |
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1.4 |
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|
Spectral
Resolution & Wavelength Range |
The
Instruments shall provide adequate spectral resolution and range to meet
Level 1 spectral irradiance and atmospheric imaging measurements |
N/A |
N/A |
N/A |
N/A |
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1.4.1 |
3/27/03 |
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|
Solar
Spectral irradiance measurements shall be performed to
cover in the 0.1 to 105 nm range |
For
minimum mission success, 6 or more emissions to specify the chromosphere, TR,
and corona, plus the He II 30.4 nm emission. |
EVE |
Lev.1 [Science Meas. 5 in 4.1.1] |
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1.4.2 |
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|
Spectral
resolution of 0.1 nm for a minimum of 18 emission lines shall be achieved |
For
minimum mission success, .2 nm for 6 emissions. |
EVE |
Lev.1 [Science Meas. 5 in 4.1.1] |
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1.4.3 |
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Spectral
resolution of 5 nm for other emission lines shall be achieved |
No
corresponding measurements for minimum mission. |
EVE |
Lev.1 [Science Meas. 5 in 4.1.1] |
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1.4.4 |
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|
Atmospheric
images shall cover the temperature
range spanning 20,000 to 3,000,000 K with 7 wavebands |
AIA
requirement: 7 telescopes for 7 different wavebands |
SHARPP |
Lev.1 [Science Meas. 4 in 4.1.1] |
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1.5 |
|
|
Precision, Accuracy, & Dynamic Range |
The
Instruments shall provide adequate measurement accuracy over the require
measurement range to meet the Level 1 Requirements |
N/A |
N/A |
N/A |
N/A |
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1.5.1 |
3/27/03 |
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The
Doppler velocity noise shall be <25 m/sec over a period of 50 sec. The
Doppler velocity dynamic range shall be at least +/-6.5km/s in the presence
of a 3kG field. The orbital motion shall not cause power spectrum artifacts
with frequencies above 500 microHz. |
This
noise requirement flows down to a requirement that images be co-aligned on an
HMI sensor to 0.1 arcsec (3-sigma) over the 50-second interval needed for one
set of images to limit noise from intensity gradients and image shifts. Note
that 6.5km/sec Doppler shift only possible in orbits where Doppler shifts are
< 4 km/s. To avoid artifacts the orbital period must be of the order 5-10
times longer than the period corresponding to 500 microHz or about 5 hours.
The effictively requires orbits such as GEO or L1 orbits. |
HMI |
Lev.1 [Science Meas. 1 in 4.1.1] |
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1.5.2 |
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|
Longitudinal
magnetograms shall have a zero point error of <0.3 Gauss. |
The
zero point error refers to the disk-averaged line-of-sight field zero point
noise and drives the exposure duration accuracy requirement. |
HMI |
Lev.1 [Science Meas. 2 in 4.1.1] |
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1.5.3 |
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Longitudinal
magnetograms shall have a noise level of < 5 Gauss over a period of 10
minutes. |
This
flows down to a requirement that images be co-aligned on an HMI sensor to 0.1
arcsec (3-sigma) over the 50-second interval needed for one set of images to
limit noise from intensity gradients and image shifts. |
HMI |
Lev.1 [Science Meas. 2 in 4.1.1] |
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1.5.4 |
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|
Longitudinal
and vector magnetograms shall have a dynamic measurement range of +-3 kGauss |
|
HMI |
Lev.1 [Science Meas. 2, 3 in 4.1.1] |
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1.5.5 |
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|
The
polarimetric precision in Q, U, and V should be better than 0.3% in 10
minutes. |
Q,
U, and V are the basic polarimetry measurements from which the vector
magnetic field is derived. The noise level and systematic errors of vector
magnetic field measurements are complicated functions of solar conditions and
Q, U, and V. This also flows down to the image stability requirement of 0.1
arcsec (3 sigma). |
HMI |
Lev.1 [Science Meas. 2 in 4.1.1] |
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1.5.6 |
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|
|
Calibration
of the intensity of the atmospheric images shall be 10%. This absolute
calibration is defined to be the daily averaged integrated solar disk
intensity. |
Minimum
mission success shall be 20% (TBR). SHARPP calibration requirements should
not exceed those of EVE. |
SHARPP |
Lev.1 [Science Meas. 4 in 4.1.1] |
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1.5.7 |
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|
Signal
to noise of atmospheric images shall be 8 in quiet sun, 20 in active sun |
6
in quiet sun, 15 inactive sun for minimum mission success. |
SHARPP |
Lev.1 [Science Meas. 4 in 4.1.1] |
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|
|
1.5.8 |
|
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|
|
Coronagraphic
images shall have a calibrated intensity precision (absolute accuracy) of 10% |
20%
for minimum mission success. |
SHARPP |
Lev.1 [Science Meas. 6 in 4.1.1] |
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|
1.5.9 |
3/27/03 |
|
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|
Spectral
irradiance measurements shall have an absolute accuracy of 10% (1o) for the daily
average in 5 nm intervals and 10% (1o) for the brighter emission lines at the
measurement cadence of 20 sec or shorter. |
20%
for minimum mission success. |
EVE |
Lev.1 [Science Meas. 5 in 4.1.1] |
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1.5.10 |
3/27/03 |
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|
|
Spectral
irradiance measurements shall have a precision of 2% (1
sigma) per year |
5%
for minimum mission success. |
EVE |
Lev.1 [Science Meas. 5 in 4.1.1] |
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1.6 |
|
|
Cadence |
The
Instruments shall provide data sets with adequate cadence to study solar
phenomena on appropriate time scales |
N/A |
N/A |
N/A |
N/A |
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|
1.6.1 |
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|
|
The
dopplergrams shall have an observation cadence no slower than 50 seconds |
Drives
instr observation timing, instr processing & data rate |
HMI |
Lev.1 [Science Meas. 1 in 4.1.1] |
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1.6.2 |
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|
The
longitudinal magnetograms shall have an observation cadence no slower than
50 seconds |
Drives
instr observation timing, instr processing & data rate |
HMI |
Lev.1 [Science Meas. 2 in 4.1.1] |
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|
|
1.6.3 |
|
|
|
|
The
vector magnetograms shall have an observation cadence no slower than 5 minutes |
Drives
instr observation timing, instr processing & data rate |
HMI |
Lev.1 [Science Meas. 3 in 4.1.1] |
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|
1.6.4 |
|
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|
|
The
atmospheric images shall have an observation cadence of no higher than 10
seconds |
20
seconds for minimum success. Drives instr observation timing, instr
processing & data rate |
SHARPP |
Lev.1 [Science Meas. 4 in 4.1.1] |
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1.6.5 |
|
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|
|
The
coronagraphic images shall have an observation cadence of no higher than 60
seconds |
80
seconds for minimum success. Drives instr observation timing, instr
processing & data rate |
SHARPP |
Lev.1 [Science Meas. 6 in 4.1.1] |
|
|
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|
1.6.6 |
3/27/03 |
|
|
|
The
spectral irradiance measurements shall have an observation cadence of no
higher than 20 sec or better seconds |
60
seconds for minimum success. Drives instr observation timing, instr
processing & data rate |
EVE |
Lev.1 [Science Meas. 5 in 4.1.1] |
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2 |
|
|
Mission Implementation |
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2.1 |
|
|
Orbit |
SDO
orbit shall be selected to support the instrument science requirements and
overall ops concept |
N/A |
N/A |
N/A |
N/A |
|
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|
2.1.1 |
|
|
|
|
Geosynchronous
orbit with an initial RAAN of 200 +/- 7.5 degrees longitude |
Geo
orbit allows continuous ground contact for high rate downlink and introduces
an acceptable orbital Doppler shift and orbital period for HMI
measurements; RAAN selection minimizes
eclipse season impacts to science data collection; also has impacts on
antenna FOV and handovers in degraded mode. Tolerance on RAAN defines launch
window (each 1 hour widening of window results in 15 deg RAAN shift). Launch window opens at 192.5 degrees |
LV, GN&C, PROP, FLT DYN |
Lev.1
[Science Meas. 1], 1.2.1 (Data Capture & Comp.), 1.5.1 (Dopplergram
accuracy), 1.6 (Cadence) |
|
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|
2.1.2 |
|
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|
The
Observatory shall be designed to support operations in an orbit with two
yearly eclipse seasons with a maximum duration of 23 days, each with a
maximum daily shadow of 72 minutes |
Eclipse
seasons indicated in AO. Orbit eclipse characteristics derived from RAAN
selection |
GN&C, FLT, DYN, PROP, MECH, THERM, PWR |
Lev.1
[Science Meas.], 1.2.1 (Data Capture & Compl) |
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|
2.1.3 |
|
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|
|
The
orbit shall be selected to have an average longitudinal stationkeeping
position within a range of 100-110 deg W longitude |
Based
on gravity well at specified position which minimizes stationkeeping
maneuvers |
LV, GN&C, FLT, DYN, PROP |
Lev.1
[Science Meas.], 1.2.1 (Data Capture & Compl) |
|
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|
2.1.4 |
|
|
|
|
The
Observatory shall maintain a stationkeeping position within +/- 0.5 deg of
its allocated average longitudinal position |
Based
typical longitudinal orbit slot for Geo Spacecraft |
GN&C, FLT, DYN, PROP, GND |
Lev.1
[Science Meas.], 1.2.1 (Data Capture & Compl) |
|
|
|
|
|
2.1.5 |
|
|
|
|
The
mission shall be designed to remain within a maximum orbital eccentricity of
0.005 (TBR) over the course of a 5 year mission |
Defines
orbit characteristics- does not place prohibitive reqs on orbit; Affects
HGA/ground station pointing angle (minimizes antenna pointing angles) &
Delta-V stationkeeping budget; Eccentricity provides predictable range for
HMI doppler effects; Keeps observatory above outer Van Allen belt |
GN&C, FLT, DYN, PROP, GND |
Lev.1
[Science Meas.], 1.2.1 (Data Capture & Comp.), 1.5.1 (Dopplergram
accuracy) |
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|
2.1.6 |
|
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|
|
The
Observatory and mission shall be designed for a maximum orbital inclination
of < 30 (TBR) over the 5 year mission life |
Initial
inclination at launch planned to be 28.7 deg; Affects HGA/ground station
pointing angle & Delta-V stationkeeping budget |
GN&C, FLT, DYN, PROP, GND |
Lev.1
[Science Meas.], 1.2.1 (Data Capture & Comp.) |
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2.2 |
|
|
Mission Life |
Mission
life shall be sufficient to achieve science data collection to meet
fundamental science requirements |
N/A |
N/A |
N/A |
N/A |
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|
2.2.1 |
|
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|
|
5
Year Mission Life (after reaching on-station GEO orbit and post commissioning
activities) |
Derived
from Level 1 Reqs; Level 1 doc indicates three years as minimum mission life. |
ALL |
Lev.1
[Science Objectives, Mission Timeline Success Criteria in 4.1.1], 1.1.2
(Science Observ) |
|
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|
2.2.2 |
|
|
|
The
SDO Spacecraft shall be designed to be tolerant to a single fault tolerant
and still meet minimum mission success criteria or shall employ sufficient
testing or analysis to ensure system reliability where fault tolerance and
graceful degradation does not exist |
|
Project
guidelines direct the design of the most robust and fault tolerant system
within the constraints of allocated resources. Requirement reflects accepted design
practice commensurate with mission scope; spacecraft will employ the use of
redundancy, cross-strapping and spacecraft system design that promotes
graceful degradation of the SDO spacecraft functions in the event of an
anomaly or failure to allow spacecraft to meet 5 year goal; Anticipate that
the redundancy implementation required to meet project-level reliability
requirements for a 5 year mission |
ALL |
1.1.2
(Science Observ), 2.2.1 (Mission Life) |
|
|
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|
2.2.3 |
|
|
|
|
The
Observatory design shall support the inclusion of propellant and OFS gas for
a 10 year mission life |
Reflects
headquarters goal for operation if observatory is functional beyond mission
life requirement; subject to project resource constraints |
PROP, GN&C, FLT DYN, MECH |
Lev.1
[Science Objectives, Mission Timeline Success Criteria in 4.1.1], 1.1.2
(Science Observ) |
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2.3 |
|
|
Environment |
N/A |
N/A |
N/A |
N/A |
N/A |
|
|
|
|
|
2.3.1 |
|
|
Initial Orbital Insertion Environment |
The
Observatory shall be able to withstand and operate within the conditions of
the initial Geo Transfer Orbit through orbit change to on-station
geosynchronous position |
|
|
|
2.1
(Orbit) |
|
|
|
|
|
2.3.1.1 |
3/27/03 |
|
|
|
The
Observatory shall be able to survive TBD (10?) orbits at 185 km perigee
altitudes without any degradation to the Observatory or science mission |
Addresses
need to survive multiple low altitude perigee passes and the resultant
effects on the Observatory and spacecraft performance (Atomic oxygen
degradation, aeroheating, RW saturation, etc) |
CONTAM, GN&C, MATL, RAD, THERM, Electrical subsystems |
2.1
(Orbit) |
|
|
|
|
|
2.3.1.2 |
3/27/03 |
|
|
|
The
Observatory shall be able to survive TBD (50?) orbits through the outer Van
Allen Belts without any degradation to the Observatory or science mission |
Addresses
potential of extended radiation exposure through Van Allen belts during orbit
raising from GTO to Geo |
RAD, Electrical subsystems, MATL |
2.1
(Orbit) |
|
|
|
|
|
2.3.2 |
3/27/03 |
|
Radiation Effects & Design |
The
SDO Observatory shall survive the radiation environment of the mission |
N/A |
N/A |
RAD, INSTR, Electrical subsystems, PARTS, MATL |
1.5
(Precision), 2.1.2 (Mission life), 2.1.1 (Orbit) |
|
|
|
|
|
2.3.2.1 |
|
|
|
Parts
TID requirements shall be based on the effective shielding provided by the
SDO Observatory ray trace and dose/depth curve detailed in the SDO Radiation
Environment document (Doc # TBD) |
|
Radiation
calculations indicate 400 Krad TID behind 100mils Al for a 5 year mission
(includes a 2x margin for uncertainty of model and variability of
environment); reduces to 40 Krad over five years for an effective parts
shielding of 200 mils Al. SDO ray
trace provides specifics of effective shielding provided by Observatory |
RAD, INSTR, Electrical subsystems, PARTS, MATL |
2.1.1
(Orbit), 2.2.1 (Mission Life) |
|
|
|
|
|
2.3.2.2 |
|
|
|
|
The
parts selected shall be immune to destructive SEEs (part LET > 100MeV
{TBR} ) |
|
RAD, INSTR, Electrical subsystems, PARTS |
2.2.1
(Orbit), 2.1.2 (Mission Life) |
|
|
|
|
|
2.3.2.3 |
|
|
|
|
Parts
which demonstrate susceptibility to non-destructive SEE's at an LET lower
than 37 MeV (TBR) shall not degrade mission performance. Radiation effects analysis and a
criticality assessment should be conducted as part of the assessment of
mission performance effects. |
SEE
susceptible parts may be included if accompanied by design mitigation which
eliminates SEE consequences in S/C
performance (circuit design, non-critical usage, etc) |
RAD, INSTR, Electrical subsystems, PARTS |
2.1.1
(Orbit), 2.2.1 (Mission Life) |
|
|
|
|
|
2.3.3 |
|
|
Spacecraft Charging |
The
SDO Observatory design shall prevent surface and internal
charging/discharging effects that damage Observatory components or disrupt
Observatory operations |
N/A |
N/A |
N/A |
N/A |
|
|
|
|
|
2.3.3.1 |
3/27/03 |
|
Surface Charging |
The
Observatory shall be designed to dissipate collected charge on external
surfaces to prevent damaging ESD discharge |
|
Reference
NASA technical paper 2361 titled “Design Guidelines for Assessing and
Controlling Spacecraft Charging Effects” can be used as a guide for the
prevention of spacecraft charging. |
RAD, INSTR, ELEC, Electrical subsystems, MECH, DEPLOY,
THERM, MATL |
2.1.1
(Orbit), 2.1.2 (Mission Life) |
|
|
|
|
|
2.3.3.1.1 |
3/27/03 |
|
|
The
Observatory shall have sufficient conductive surface area on the sun-facing
side to allow dissipation of collected charge from the entire Observatory |
|
Reflects
the effect that surface charge collects on darkened conductive surfaces and
dissipates from sun-facing conductive surfaces; requires conductive path
between all Observatory surfaces to prevent isolated charge buildup |
RAD, INSTR, ELEC, Electrical subsystems, MECH, DEPLOY,
THERM, MATL |
2.1.1
(Orbit), 2.2.1 (Mission Life) |
|
|
|
|
|
2.3.3.1.2 |
3/27/03 |
|
|
|
TBD
% of all Observatory external surfaces shall be conductive (<= 1E9
Ohms/sq) with no single continuous non-conductive surface area greater than
TBD |
This
requirement addresses surface charge control and provides specfic
requirements to 2.3.3.1.2 above. Does
not apply to optical apertures and other similar surfaces. Refer to the SDO Electrical Systems Spec
(doc # TBD) for specific design implementation guidelines |
RAD, INSTR, ELEC, Electrical subsystems, MECH, DEPLOY,
THERM, MATL |
2.1.1
(Orbit), 2.2.1 (Mission Life) |
|
|
|
|
|
2.3.3.2 |
3/27/03 |
|
Internal Charging |
No
internal discharge shall cause permanent damage to the Observatory circuitry. |
|
NASA
HandBook-4002 on Spacecraft Charging can be used as a reference for internal
spacecraft charging. |
Rad, INSTR, ELEC, Electrical subsystems, MECH, MATL |
2.1.1
(Orbit), 2.2.1 (Mission Life) |
|
|
|
|
|
2.3.3.2.1 |
|
|
|
The
Observatory shall have sufficient shielding to prevent dielectric charge
buildup to prevent ESD discharge effects |
|
This
requirement addresses internal charge control. Refer to the SDO Electrical Systems Spec
(doc # TBD) for specific design implementation guidelines |
RAD, INSTR, ELEC, Electrical subsystems, MECH |
2.1.1
(Orbit), 2.2.1 (Mission Life) |
|
|
|
|
|
2.3.4 |
|
|
Contamination |
N/A |
N/A |
N/A |
N/A |
N/A |
|
|
|
|
|
2.3.4.1 |
|
|
Flight Configuration/Ops |
Contamination
of sensitive portions of the Observatory by condensables and particulates
shall not prevent the mission from meeting its requirements |
N/A |
N/A |
ALL |
Lev. 1
[Precision (Science Meas. 1-6)], 2.2.1
(Mission Life) |
|
|
|
|
|
2.3.4.1.1 |
|
|
|
Acceptable
levels of contamination shall be
maintained on the Observatory per the SDO Contamination Control Plan (Doc #
TBD) |
|
|
ALL |
Lev. 1
[Precision (Science Meas. 1-6)], 2.2.1
(Mission Life) |
|
|
|
|
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|
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|
|
2.3.4.2 |
|
|
I&T Activities |
Contamination
from I&T activities shall not prevent the mission from meeting its
requirements |
N/A |
N/A |
ALL |
Lev.
1 [Precision (Science Meas. 1-6)] |
|
|
|
|
|
2.3.4.2.1 |
|
|
|
|
Class
10,000 facilities are required for the instrument and propulsion integration
and test phases |
Requirement based on the proposal information
indicating that higher cleanliness standards are required during
instrument-level assembly and testing |
ALL |
Lev.
1 [Precision (Science Meas. 1-6)] |
|
|
|
|
|
2.3.4.2.2 |
|
|
|
|
Class
10,000 facilities are required for the Observatory integration and test
phases |
|
ALL |
Lev.
1 [Precision (Science Meas. 1-6)] |
|
|
|
|
|
2.3.4.2.3 |
|
|
|
The
Observatory shall be designed to incorporate and allow the use of Instrument
dry purge lines throughout the Observatory integration and test phases |
|
Requirement
based on instrument cleanliness requirements and specific requirements stated
in instrument proposals |
CONTAM, MECH, GSE, INSTR |
Lev.
1 [Precision (Science Meas. 1-6)] |
|
|
|
|
|
2.3.5 |
3/27/03 |
|
Magnetics |
Observatory
residual and induced magnetic fields and the natural orbital environmental
magnetic field shall not disrupt Observatory operations or corrupt mission
science |
N/A |
N/A |
CONTAM, INSTR, ELEC, Electrical subsystems, MATL |
Lev.
1 [Precision (Science Meas. 1-6)] |
|
|
|
|
|
2.3.5.1 |
3/27/03 |
|
|
|
The
EVE instrument shall be able to meet instrument operational and science
requirements in the presence of a magnetic field no greater than TBD (300
nTesla??) at the EVE instrument location |
Magnetic
environment consists of natural orbital magnetic field at GEO (~ 170 nTesla)
and EVE team needs to do a
reassessment of magnetic susceptibility and possible shielding. Variations in magnetic field through GEO
orbit are also a significant potential issue for EVE measurements. EVE is sensitive to magnetic fields -
affects EVE spectral resolution measurements at longest wavelength
photons. May need to do an on-orbit
magnetic shield assessment. |
CONTAM, INSTR, ELEC, Electrical subsystems, MATL |
Lev.
1 [Precision (Science Meas. 5)], 1.5.9, 1.5.10 (EVE Precision & Range) |
|
|
|
|
|
2.3.5.2 |
3/27/03 |
|
|
|
No
Observatory subsystem shall generate a magnetic field greater than TBD
(100nTesla??) at an equivalent distance to that on the Observatory between
the component and EVE |
Magnetic
field measured at component location.
Verification of this requirement may be by analysis or test on a
case-by-case basis |
CONTAM, INSTR, ELEC, Electrical subsystems, MATL |
Lev.
1 [Precision (Science Meas. 5)], 1.5.9, 1.5.10 (EVE Precision & Range) |
|
|
|
|
|
2.3.5.3 |
3/27/03 |
|
|
|
The
Observatory shall not generate a magnetic field greater than TBD (100
nTesla??) at the EVE instrument location |
|
CONTAM, INSTR, ELEC, Electrical subsystems, MATL |
Lev.
1 [Precision (Science Meas. 5)], 1.5.9, 1.5.10 (EVE Precision & Range) |
|
|
|
|
|
2.3.5.4 |
|
|
|
|
All
observatory subsystems (except as noted above) shall be able to meet full
mission requirements in the presence of a magnetic field of 40 µTesla (TBR) |
Estimated
magnetic field at expected minimum Geo perigee (185 Km) |
CONTAM, INSTR, ELEC, Electrical subsystems |
2.3.1.1
(Min perigee) |
|
|
|
|
|
|
|
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|
|
|
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|
|
|
|
|
|
|
2.4 |
|
|
Launch Vehicle (LV) |
The
launch vehicle (LV) shall provide sufficient performance capability and
reliability to place SDO Observatory in the desired initial transfer orbit |
N/A |
N/A |
LV, FLT DYN, PERF ASSUR. |
2.1.1
(Orbit) |
|
|
|
|
|
2.4.1 |
3/27/03 |
|
LV Selection |
|
The
Observatory design and ops concept shall be compatible with the use of the
DELTA 4040 or ATLASV 401 LV configuration |
Reflects
LV configuration options |
ALL |
2.1.1
(Orbit) |
|
|
|
|
|
2.4.2 |
|
|
LV Performance |
The
LV shall deliver the Observatory to a transfer trajectory from which the
Observatory-supplied propulsion system shall modify the orbit to its desired
final profile |
N/A |
N/A |
LV, FLT DYN |
2.1.1
(Orbit) |
|
|
|
|
|
2.4.2.1 |
3/27/03 |
|
|
|
The LV shall provide an 3200
kg (TBR) throw weight to a 28.7 deg (ETR)
inclination orbit |
Reflects
estimated Observatory mass allocation;
Preliminary allocation deliberately limited below LV capability of
3900 in order to leave open co-manifest options |
LV, FLT DYN |
2.1.1
(Orbit) |
|
|
|
|
|
2.4.2.2 |
|
|
|
|
The
LV orbit insertion errors shall be no greater than TBD (as defined in the "DELTA IV PAYLOAD PLANNERS GUIDE"). |
Expect
that LV ICD will supercede LV PPG as controlling configured document when
available. AO assumptions indicated
that Observatory will not budget fuel to correct for inclination insertion
errors- counting on LV to provide adequate insertion trajectory. PPG addresses deviation range in
inclination, apogee, perigee, and argument of perigee. |
LV, FLT DYN, Electrical subsystems |
2.1.1
(Orbit), 2.4.1 (LV) |
|
|
|
|
|
2.4.2.3 |
|
|
|
|
The
minimum perigee for initial orbital insertion trajectory shall nominally be
185 km. |
Low
perigee passes result in various effects on the Observatory and spacecraft
performance (Atomic oxygen degradation, aeroheating, RW saturation, magnetic
field effects, etc) |
LV, FLT DYN |
2.1.1
(Orbit), 2.4.1 (LV) |
|
|
|
|
|
2.4.2.4 |
3/27/03 |
|
|
|
The
LV-induced initial Observatory tipoff rates shall be no greater than 1, 2, 2 deg/sec (3 sigma) (x, y, z) at Observatory
separation |
Reflects
design requirement by which attitude control system must be able to unload
initial post-separation momentum |
LV, FLT DYN |
2.1.1
(Orbit), 2.4.1 (LV) |
|
|
|
|
|
2.4.3 |
|
|
LV Interfaces |
The
launch vehicle (LV) shall provide specified mechanical and electrical
interfaces to the Observatory and the ground system for testing,
verification, and flight ops |
|
|
LV, MECH, PWR, RF, C&DH, ELEC, GSE |
2.7.2
(Verify) |
|
|
|
|
|
2.4.3.1 |
|
|
Observatory umbilical connection |
The LV shall provide an umbilical interface for power and
communications from the Observatory to ground control station via the
blockhouse in order to allow remote test and monitoring of the Observatory
while it is integrated to the LV |
|
|
LV, MECH, PWR, RF, C&DH, ELEC, GSE |
2.7.2
(Verify) |
|
|
|
|
|
2.4.3.2 |
|
|
Environmental conditioning |
The LV shall provide access for the necessary environmental
conditioning of the Observatory while it is integrated to the LV |
|
|
LV, MECH, PWR, INSTR, CONTAM |
2.3
(Envir) |
|
|
|
|
|
2.4.3.2.1 |
|
|
|
|
THE LV fairing shall be Boeing VC6 (visibly clean, Level 6) or
equivalent |
Required to maintain contamination and environmental control
reflected in Contamination Control plan |
LV, MECH, CONTAM |
2.3
(Envir) |
|
|
|
|
|
2.4.3.2.2 |
|
|
|
|
The LV shall provide access for a continuous filtered purge of
grade TBD nitrogen up to T-0 (umbilical break-away line) |
Required to maintain contamination and environmental control
reflected in Contamination Control plan |
LV, MECH, CONTAM |
2.3
(Envir), 2.2.5 (Safety) |
|
|
|
|
|
2.4.3.2.3 |
|
|
|
|
The LV shall provide continuous Class 1,000 HEPA filtered air
supply into fairing up to launch |
Required to maintain contamination and environmental control
reflected in Contamination Control plan |
LV, MECH, CONTAM |
2.3
(Envir), 2.2.5 (Safety) |
|
|
|
|
|
2.4.3.2.4 |
|
|
|
The LV interface shall provide access and an implementation
approach for battery A/C to maintain
Observatory batteries within required operational and safety margins |
|
Reflects Battery Handling Plan (doc # TBD) and SDO Launch site
operations plan |
LV, MECH, PWR, CONTAM |
2.3
(Envir), 2.2.5 (Safety) |
|
|
|
|
|
2.4.3.3 |
|
|
Separation signals |
|
|
|
|
|
|
|
|
|
|
2.4.3.3.1 |
|
|
|
|
The
Observatory shall support the use of 3
separation signals from the LV third stage to the Observatory |
|
LV, MECH, ELEC |
4.2.1
(Acquisition)) |
|
|
|
|
|
2.4.3.4 |
|
|
LV Access |
The LV shall provide vehicle access to allow servicing of
Observatory components (battery, prop system, Instruments, etc) as required
on the launch pad with the fairing in place |
|
|
LV, MECH |
2.7.2
(Verif), 2.7.5 (Safety) |
|
|
|
|
|
|
|
|
|
|
|
|
|
|
|
|
|
|
|
2.5 |
|
|
Mission Implementation/ Ops Concept |
The SDO Ops Concept shall combine the Science collection and
distribution requirements as well as the Observatory design implementation
approach and constraints into an implementation plan that allows successful
mission planning and operation |
|
|
ALL |
Lev.1
[Sci Req 1-6], 1.1 (Sci Obs), 1.2 (Data Capture & Compl) |
|
|
|
|
|
2.5.1 |
|
|
Continuous Contact |
The Observatory shall maintain (near) continuous science data
downlink contact with the ground station in order to capture the science data
within the capture budget |
|
Traced from both the high DL data rates required as well as the
prohibitive impact of implementing a SSR approach to both store then DL the
data at a rate that would clear the SSR in a reasonable time frame. Any gaps in DL coverage must be encompassed
within the data capture budget margins. |
C&DH, GND, APS, RF, GN&C |
Lev.1
[Sci Req 1-6], 1.2 (Data Capture & Compl), 1.6 (Cadence) |
|
|
|
|
|
2.5.2 |
|
|
Dedicated Ground Station Sites |
The Mission implementation shall employ the use of a dedicated
ground station to meet science data downlink completeness requirements |
|
Reflects the high data rate and data capture budget
implementation approach |
GND |
Lev.1
[Sci Req 1-6], 1.2 (Data Capture & Compl), 1.6 (Cadence) |
|
|
|
|
|
2.5.3 |
|
|
Data Completeness |
The Mission implementation shall meet the data capture and
completeness requirements needed to meet the science requirements |
|
Addressed through the use of a configured Data Capture Budget
(Doc # TBD) which allocates data loss throughout the data capture and
distribution components |
INSTR, C&DH, GN&C, APS, RF, GND |
Lev.1
[Sci Req 1-6], 1.2 (Data Capture & Compl) |
|
|
|
|
|
2.5.4 |
|
|
Data Delivery |
The Ground Stations shall route the science data directly to the
Instrument Science Operations Centers (SOCs) after receiving the data (per
data latency requirements in 5.2.6.6) |
|
Addresses the need for continuous transport of science data
directly to the SOCs. Observatory
housekeeping data will be routed to the SOCs via the MOC. |
GND, INSTR |
Lev.1
[Science Objectives D & E: Support Forecasting] |
|
|
|
|
|
2.5.5 |
|
|
Observatory Pointing and Jitter Control |
The combination of the Observatory attitude control, the
Observatory and Instrument mechanical and thermal design, and Instrument
internal pointing compensation shall provide the necessary pointing control
and jitter performance required to meet the instrument science requirements |
|
|
INSTR, MECH, GN&C, APS |
Lev.1
[Sci Req 1-6], 1.3 (Ang Res. & Coverage) |
|
|
|
|
|
2.5.5.1 |
3/27/03 |
|
|
|
The Observatory shall designate KCOR as the science reference
boresight and shall point this reference boresight to the target (sun center)
to an absolute accuracy of 30 arcsec (3 sigma) using the AIA GT signal |
Since KCOR has no internal pointing capability and has the most
severe science degradation impacts due to pointing errors, the observatory shall use the GT signal,
with a suitable bias applied, to point KCOR at sun center |
SHARPP, MECH, GN&C, APS |
Lev.1
[Sci Req 1-6], 1.3 (Ang Res. & Coverage) |
|
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|
2.5.5.2 |
3/27/03 |
|
|
|
The AIA Magritte and SPECTRE boresights will be maintained
within 70 arcsec (3 sigma)
in the Y and Z axes over a period of not less than one week |
|
SHARPP, MECH, GN&C, APS |
Lev.1
[Sci Req 1-6], 1.3 (Ang Res. & Coverage) |
|
|
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|
|
2.5.5.3 |
|
|
|
|
The HMI boresight will be maintained within 200 arcsec of sun
center through a combination of on-ground alignment and optical bench
stability |
|
HMI, MECH, GN&C, APS |
Lev.1
[Sci Req 1-6], 1.3 (Ang Res. & Coverage) |
|
|
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|
|
2.5.5.4 |
|
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|
|
The
HMI boresight will be adjusted to within 14 arcsec (3 sigma) through the
adjustment of the HMI on-orbit alignment system (legs), internal guiding
correction provided by the HMI Image Stabilization System and the Spacecraft
pointing control |
14
arcsec is the limit of the HMI image motion compensation system |
HMI, MECH, GN&C, APS |
Lev.1
[Sci Req 1-6], 1.3 (Ang Res. & Coverage) |
|
|
|
|
|
2.5.5.5 |
3/27/03 |
|
|
|
The Observatory shall maintain the EVE boresight to an absolute
accuracy of 450 arcsec (3
sigma) to the target (sun center) through a combination of on-ground
alignment accuracy and optical bench stability |
The Spacecraft will maintain the EVE mounting surface (reference
optical surface) to within an accuracy of 150 arcsec (3 sigma) to the target
(sun center); need to determine EVE absolute pointing req (see 3.1.1.1) |
EVE, MECH, GN&C, APS |
Lev.1
[Sci Req 1-6], 1.3 (Ang Res. & Coverage) |
|
|
|
|
|
2.5.5.6 |
|
|
|
|
The Observatory jitter at the HMI and SHARPP mounting interface
to the Observatory optical bench shall be 5 arcsec (3 sigma) over frequencies
of 0.02 Hz to 50 Hz in the X, Y & Z axes |
Addresses requirement for control of torque disturbances and
torque noise |
GN&C, MECH, APS |
1.3
(Ang Res & Coverage), 2.5.5 (Pointing & Jitter Control), 2.6.4
(Attitude Control) |
|
|
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|
|
2.5.5.7 |
|
|
|
|
EVE requires a stability of 140 arcsec (TBR) over a period of
minutes to days (TBR). |
EVE defines stability as longer period drift in the control
point (in the control bandwidth of the S/C) and alignment, both preflight and
inflight offset errors of a sensor bore to the 'reference' control point (See
section 4). Can we come up with an acceptable definition? |
GN&C, MECH, APS |
1.5
(Precision), 2.5.5 (Pointing & Jitter Control), 2.6.4 (Attitude Control) |
|
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|
2.5.6 |
|
|
Spacecraft Autonomy |
The Spacecraft shall possess sufficient onboard autonomy to
allow basic fault detection and correction. |
|
Addresses the need for onboard autonomy for basic operational
reliability and fault tolerance |
C&DH, GN&C, FSW, PWR |
2.2.1
(Mission Life), 2.2.2 (Fault Tol) |
|
|
|
|
|
2.5.7 |
|
|
Critical Telemetry Monitoring |
All mission and time critical activities (separation, solar
array & HGA deployment, acquisition, critical Delta-V maneuvers, etc)
shall be performed within ground contact to allow telemetry monitoring |
|
|
C&DH, FSW, RF,
GND |
2.2.1
(Mission Life), 2.2.2 (Fault Tol) |
|
|
|
|
|
2.5.8 |
|
|
Mission Phases |
The SDO Observatory and mission shall be designed to support the
various phases of SDO flight operations |
|
|
|
|
|
|
|
|
|
2.5.8.1 |
|
|
|
The SDO Observatory and mission shall be designed to support
Launch and Acquisition |
|
Phase covering pre-launch configuration until Observatory is
power-positive and pointing at the Sun |
|
|
|
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|
|
2.5.8.2 |
|
|
|
The SDO Observatory and mission shall be designed to support
In-Orbit Checkout |
|
Phases covering the first weeks to check out and calibrate the
Observatory. Portions of this will occur concurrently with the Orbit
Circularization phase |
|
|
|
|
|
|
|
2.5.8.3 |
|
|
|
The SDO Observatory and mission shall be designed to support
Orbit Circularization |
|
Occurs during In-Orbit Checkout and covers the mission elements
required to conduct several apogee thruster burns to raise orbit to Geo |
|
|
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|
|
|
|
2.5.8.4 |
|
|
|
The SDO Observatory and mission shall be designed to support
Nominal Mission Mode |
|
Primary on-station operational phase, where science data is
collected and transferred to the ground station, where it is distributed to
the SOCs. Observatory housekeeping
data is also collected and transferred to the MOC for monitoring/trending/etc
and distribution to the SOCs |
|
|
|
|
|
|
|
2.5.8.5 |
|
|
|
The SDO Observatory and mission shall be designed to support
Periodic Calibrations and Housekeeping |
|
Interruptions in minimal science phase needed to maintain
science quality through calibration observations and alignment adjustments |
|
|
|
|
|
|
|
2.5.8.6 |
|
|
|
The SDO Observatory and mission shall be designed to support
Eclipse periods |
|
Consequence of Geo orbit- principal Observatory requirement is
to survive and minimize impact on science operations |
|
|
|
|
|
|
|
2.5.8.7 |
|
|
|
The SDO Observatory and mission shall be designed to support
Stationkeeping and Momentum Management |
|
Required operations to keep observatory within designated
orbital position and angular momentum capabilities |
|
|
|
|
|
|
|
2.5.8.8 |
|
|
|
The SDO Observatory and mission shall be designed to support
Safehold and Emergency modes |
|
Required to place the Observatory in a safe mode in the event of
an anomaly and allow anomaly investigation and correction from the ground |
|
|
|
|
|
|
|
2.5.8.9 |
|
|
|
The SDO Observatory and mission shall be designed to support
Observatory Disposal |
|
Required to dispose of Observatory per NASA guidelines |
|
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|
2.6 |
|
|
Standard Spacecraft Services |
The SDO Spacecraft shall provide standard spacecraft
functions and services required for orbital science operations |
|
|
|
|
|
|
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|
|
2.6.1 |
|
|
Structure |
The Spacecraft mechanical structure shall accommodate the
requirements of the Launch Vehicle, Spacecraft and Instrument |
|
|
MECH |
2.1
(Orbit), 2.4 (LV) |
|
|
|
|
|
2.6.1.1 |
|
|
|
All
Observatory components shall be allocated and remain within a specified mass
budget |
|
Requirement will be tracked by use of configured mass resources
budget spreadsheet (Doc # TBD) ; changes to mass budget allocation only by
project approved configuration change request |
All Observatory Components |
2.1
(Orbit), 2.4 (LV), 2.4.2.1 (Observ. Mass Allocation) |
|
|
|
|
|
2.6.2 |
|
|
Thermal Control |
The Spacecraft shall provide a thermal environment which meets
limits required by each element of the Observatory for all mission phases |
|
|
THERM, MECH, All Observatory Components |
2.1
(Orbit) |
|
|
|
|
|
2.6.3 |
|
|
Power Distribution |
The Spacecraft shall provide and distribute sufficient power to
support all phases of the SDO mission |
|
|
PWR |
2.1
(Orbit), 2.2.1 (Mission Life) |
|
|
|
|
|
2.6.3.1 |
|
|
|
All Observatory components shall be allocated and remain within
a specified power budget (which adresses end-of-life power requirements) |
|
Requirement will be tracked by use of configured power budget
spreadsheet; changes to power budget allocation only by project approved
configuration change request |
All Observatory Components |
2.1
(Orbit) |
|
|
|
|
|
2.6.4 |
|
|
Guidance, Navigation, & Control |
The Observatory and
Ground System shall provide the knowledge & pointing capability to
determine and control Observatory orientation, position, and velocity through
all phases of the mission within the constraints of mission science goals |
|
|
GN&C, FLT DYN, GND |
2.1
(Orbit) |
|
|
|
|
|
2.6.5 |
|
|
Communications |
Observatory shall have the capability to receive and execute
commands and transfer data between the spacecraft and ground system in order
to carry out mission operations |
|
|
C&DH, FSW, RF, GND |
2.1
(Orbit) |
|
|
|
|
|
2.6.6 |
|
|
Data Processing, Storage, & Timekeeping |
The Spacecraft shall possess sufficient data processing and
storage capability and Spacecraft timekeeping functions to meet mission
operations requirements |
|
|
C&DH, FSW |
2.1
(Orbit) |
|
|
|
|
|
2.6.7 |
|
|
Housekeeping Telemetry |
The Spacecraft shall downlink sufficient housekeeping
engineering data to the ground to allow nominal spacecraft operation and
performance evaluation, as well as anomaly investigation and resolution |
|
Addresses the need for Observatory telemetry for ground
monitoring and troubleshooting during testing and flight |
ALL |
2.1
(Orbit), 2.2.2 (Fault Tol) |
|
|
|
|
|
2.6.8 |
|
|
Spacecraft Architecture |
The Observatory shall utilize the 1553 bus as the primary method
for distributing observatory commands and collecting Observatory telemetry
for downlink |
|
|
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|
2.7 |
|
|
Development Approach |
N/A |
N/A |
N/A |
|
|
|
|
|
|
|
2.7.1 |
|
|
Performance Assurance & Safety |
The development of flight hardware, software, and GSE for the
SDO mission shall adhere to the SDO Mission Assurance Requirements (MAR)
document |
N/A |
N/A |
ALL |
2.1.2
(Mission Life), 2.6 (Standard S/C Services) |
|
|
|
|
|
2.7.2 |
|
|
Verification |
The Observatory shall undergo sufficient testing or analysis to
verify that it meets all Mission requirements |
N/A |
N/A |
ALL |
2.1.2
(Mission Life), 2.6 (Standard S/C Services) |
|
|
|
|
|
2.7.2.1 |
|
|
|
The Observatory shall be subjected to environments prior to
launch as part of a comprehensive test program to verify that it meets launch
and mission requirements based on the SDO environmental verification test
matrix |
|
|
ALL |
2.1
(Orbit), 2.6 (Standard S/C Services) |
|
|
|
|
|
2.7.2.2 |
|
|
|
The design shall provide for adequate visibility to accommodate
effective subsystem and system
functional and performance verification at all stages of development |
|
|
ALL |
2.6 (Standard S/C Services), 2.7.2
(Verification) |
|
|
|
|
|
2.7.2.2.1 |
|
|
|
Sufficient telemetry shall be available through the 1553 bus to
fully verify electrical subsystem development and system performance. Requirement does not cover the monitoring
and verification of high rate instrument science data over the 1553 bus |
|
Derived from development philosophy of reducing effort and risk
by designing system architecture and development approach for the same
testability process that can be used throughout the various Observatory
development phases |
Electrical subsystems |
2.6 (Standard S/C Services), 2.7.2
(Verification) |
|
|
|
|
|
2.7.2.2.2 |
|
|
|
Critical internal test points shall be identified by all
subsystems and instruments and brought out to external test points and/or
skin connectors to allow monitoring during test |
|
Addresses the need for intermediate test points for verification
as well as tools for ground test/debugging process both at subsystem and
Observatory level |
Electrical subsystems |
2.6 (Standard S/C Services), 2.7.2
(Verification) |
|
|
|
|
|
2.7.3 |
|
|
Configuration Management |
SDO hardware, software and operations concept development shall
employ procedures which enable the establishment and tracking of product
implementation traceability to the configured designs and approaches |
N/A |
N/A |
ALL |
2.6 (Standard S/C Services), 2.7.2
(Verification) |
|
|
|
|
|
2.7.4 |
|
|
Electrical Specification |
The SDO subsystems, instruments, components and GSE shall adhere
to the electrical and electronic requirements specified in the SDO Electrical
Systems Specification (doc # TBD) |
|
N/A |
ALL |
2.7.3
(CM), 2.7.2 (Verification) |
|
|
|
|
|
2.7.5 |
|
|
Mechanical Specification |
The SDO subsystems, instruments, components and GSE shall adhere
to the mechanical requirements specified in the SDO Mechanical Subsystem
Specification (doc # TBD) |
|
|
All Observatory Components |
2.7.3
(CM), 2.7.2 (Verification) |
|
|
|
|
|
2.7.6 |
|
|
Thermal Specification |
The SDO subsystems, instruments, components and GSE shall adhere
to the thermal requirements specified in the SDO Thermal Subsystem
Specification (doc # TBD) |
|
|
All Observatory Components |
2.7.3
(CM), 2.7.2 (Verification) |
|
|
|
|
|
2.7.7 |
|
|
End-of Life Disposal |
The Observatory design shall provide the capability for
controlled end of life disposal in accordance with NASA guidelines (NSS
1740.14, NPD 8710.3, revisions as of Jan 2003) |
|
|
GN&C, FLT DYN, PROP, MECH, GND |
2.2.5
(Safety), NASA EOL Guidelines |
|
|
|
|
|
3 |
|
|
Instrument Requirements |
|
|
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|
|
3.1 |
|
|
EVE |
|
|
|
|
|
|
|
|
|
|
3.1.1 |
|
|
Alignment & Stability |
|
|
|
|
|
|
|
|
|
|
3.1.1.1 |
3/27/03 |
|
|
|
The EVE components shall meet the performance requirements
specified in the SDO Observatory Alignment Budget (Doc. # TBD). |
Reference Optical Surface is the mounting feet. |
EVE |
1.5
(Precision) |
|
|
|
|
|
3.1.1.2 |
|
|
|
|
During nominal science operations, the EVE instrument shall
limit any disturbance contribution to the Observatory jitter to within TBD (1
arcsec??) |
LASP
- can not give an accurate assessment at this time. Obviously this leads to a
disturbance torque spec which would be something we could flow down without
s/c interia knowledge. Perhaps a better organization would be if this is
kicked up to Level 2 as in "All dynamic components shall limit their
disturbance contribution to less than TBD (?1 arcsec/sec > 2 Hz or
whatever the control bandwidth for the S/C is). If defined in this manner, it
would then be possible to flow down the jitter spec to the various components
as NTE disturbance torques in Level 3 |
EVE |
2.5.5
(Pointing & Jitter Control) |
|
|
|
|
|
3.1.2 |
|
|
Spectral Resolution |
|
MEGS A shall provide solar irradiances 0.5 - 36 nm with 0.1 nm
resolution, MEGS B shall provide 35 - 105 nm solar irradiances with 0.1 nm
resolution, OFS shall provide relative solar irradiance measurements (TBD) nm
with > 1 nm resolution, ESP shall provide solar irradiance measurements
with 0.2nm resolution. |
EVE may prefer to break out the precision, accuracy and time
cadence for each sensor as well. They also may prefer to state this with less
detail. |
EVE |
1.4
(Spectral Resolution) |
|
|
|
|
|
3.1.3 |
|
|
Timing |
|
|
|
|
|
|
|
|
|
|
3.1.3.1 |
|
|
|
The EVE Instrument shall utilize the relative Spacecraft timing
signals and accuracy provided over the 1553 bus |
|
EVE has a need to receive absolute time broadcasts over the 1553
bus in addition to relative time pulses. |
EVE |
1.5
(Precision) |
|
|
|
|
|
3.1.4 |
|
|
Data Completeness |
|
|
|
|
|
|
|
|
|
|
3.1.4.1 |
3/27/03 |
|
|
|
The EVE science data bit error rate shall be less than 2.5x10-7 (TBR). |
Reflects the instrument component of the EVE
99.6% data science completeness budget. |
EVE |
2.5.3
(Data Capture & Compl) |
|
|
|
|
|
3.1.5 |
|
|
Interface Requirements |
|
|
|
|
|
|
|
|
|
|
3.1.5.1 |
|
|
|
|
The EVE science data shall not exceed of maximum data rate
allocation of 2 Mbps over the IEEE 1355 high rate science data bus |
Defines the EVE allocation of the 130 Mbps (150 Mbps
post-encoding & margin) science data downlink |
EVE, C&DH |
1.6.6
(Cadence) |
|
|
|
|
|
3.1.5.2 |
|
|
|
The EVE instruments shall adhere to the high speed bus data rate
and interface requirements detailed in the SDO/EVE High Rate Science Bus
Interface Specification (Doc # TBD) |
|
|
EVE, C&DH |
1.2.2
(Data Capture & Compl) |
|
|
|
|
|
3.1.5.3 |
|
|
|
The EVE Instrument shall receive all Commands and distribute all
housekeeping telemetry over the Observatory 1553 interface |
|
|
EVE, C&DH, FSW |
2.6.8
(S/C Arch) |
|
|
|
|
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|
|
3.2 |
|
|
HMI |
|
|
|
|
|
|
|
|
|
|
3.2.1 |
|
|
Alignment & Jitter |
|
|
|
|
|
|
|
|
|
|
3.2.1.1 |
3/27/03 |
|
|
|
The HMI components shall meet the performance requirements
specified in the SDO Observatory Alignment Budget (Doc # TBD). |
The two HMI imaging detectors should be co-aligned to have the
entire solar disk on each detector with a minimum of 10 arcseconds separation
to any detector edge, and to have the detector rows aligned to within TBD
arcseconds. Knowledge of roll is an
important parameter for HMI. Do we
need to define a reference optical surface for use relative to the s/c? |
HMI |
1.3.1,
1.3.2 (Ang Res & Coverage) |
|
|
|
|
|
3.2.1.2 |
3/27/2003
- Delete |
|
|
|
The internal HMI alignment shift due to launch shift, thermal
effects, and uncorrected 1 g sag shall be less than TBD (30 arcsec??) in the Yand Z axes |
Instrument allocation of 200 arcsec adjustment range |
HMI |
1.3.1,
1.3.2 (Ang Res & Coverage) |
|
|
|
|
|
3.2.1.3 |
|
|
|
|
The
HMI instrument shall provide 40 dB disturbance rejection at the focal plane
in pitch and yaw with a servo bandwidth of >30 Hz (TBR). |
In order to meet 1 arcsecond image resolution, the image jitter
during exposure must be less than 0.1 pixel (0.05 arcsecond). Based on attenuation required from 5 arcsec
spacecraft jitter req to HMI 1/10 pixel (3 sigma) jitter req |
HMI |
1.3.1,
1.3.2 (Ang Res & Coverage) |
|
|
|
|
|
3.2.1.4 |
|
|
|
|
The range of the ISS (image stabilization system) for both drift
and jitter combined shall be at least +/-14 by +/- 17 arcsec . |
The HMI internal pointing must be capable of TBD arcsecond
adjustment over periods of 1 week to compensate for medium-term drift
relative to the spacecraft boresight. |
HMI |
1.3.1,
1.3.2 (Ang Res & Coverage) |
|
|
|
|
|
3.2.1.5 |
|
|
|
|
The HMI instrument shall utilize adjustable rear mounting legs
to provide pitch and yaw adjustment of the HMI alignment with respect to the
Observatory over an adjust range of +/- 200 arcseconds with 2 arcsecond
increments. |
The HMI instrument pointing must be capable of TBD arcsecond
adjustment to provide initial co-alignment with the spacecraft boresight and
to compensate for relative drift over the course of the mission. |
HMI |
1.3.1,
1.3.2 (Ang Res & Coverage) |
|
|
|
|
|
3.2.1.6 |
|
|
|
|
During nominal science operations, the HMI instrument shall
limit any disturbance contribution to the observatory jitter to within 1
arcsec (TBR). |
Opening the door and moving the legs (both done very
infrequently) may make larger disturbances. |
HMI |
2.5.5
(Pointing & Jitter Control) |
|
|
|
|
|
3.2.2 |
|
|
Angular Resolution |
|
|
|
|
|
|
|
|
|
|
3.2.2.1 |
|
|
|
|
The HMI Instrument shall provide a camera with at least
4096x4096 pixels to achieve the required dopplergram, longitudinal and vector
magnetogram resolutions. |
The HMI instrument shall provide a camera with square pixels
with at least 3920 x 3920 pixels
to achieve the required resolution. 1960 arc-sec diameter Sun) |
HMI |
1.3.1,
1.3.2 (Ang Res & Coverage) |
|
|
|
|
|
3.2.3 |
|
|
Timing |
|
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|
|
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|
|
3.2.3.1 |
3/27/03 |
|
|
|
The
basic reference clock controlling the HMI observing cycle shall be to
maintain a stability of 10^-6 over periods of days to
months, with adjustments to be made smoothly in
order to insure that there are no discontinuities or step changes in the
clock. Time must be known to
within 100 (50) ms at all times. |
An HMI reference clock
may be provided internally to the instrument. HMI must receive the spacecraft
absolute time broadcasts in order to initially set the internal clock, and to
monitor drifts of the HMI internal clock relative to the spacecraft clock. |
HMI |
1.5.1 (Precision) |
|
|
|
|
|
3.2.4 |
|
|
Data Completeness |
|
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|
|
3.2.4.1 |
3/27/03 |
|
|
|
The HMI science data bit error rate shall be less than 1x10-10 (TBR) |
HMI's allocation of the 99.99% data completeness. |
HMI |
2.5.3
(Data Capture & Compl) |
|
|
|
|
|
3.2.5 |
|
|
Interface Requirements |
|
|
|
|
|
|
|
|
|
|
3.2.5.1 |
|
|
|
|
Any compression of the HMI science data shall be in a manner
such that HMI does not exceed a maximum data rate allocation of 55 Mbps over
the IEEE 1355 high rate science data bus |
Defines the HMI allocation of the 130 Mbps (150 Mbps
post-encoding & margin) science data downlink |
HMI, C&DH |
1.6.1
, 1.6.2, 1.6.3 (Cadence) |
|
|
|
|
|
3.2.5.2 |
|
|
|
Any
compression of the HMI data shall be in a manner that preserves the data
quality commensurate with the science analysis and completeness requirements |
|
This
addresses need for HMI science data compression to meet the science data rate
req while still meeting data completeness req |
HMI |
1.2.2
(Data Capture & Compl) |
|
|
|
|
|
3.2.5.3 |
|
|
|
The HMI instruments shall adhere to the high speed bus data rate
and interface requirements detailed in the SDO/HMI High Rate Science Bus
Interface Specification (Doc # TBD) |
|
|
HMI, C&DH |
1.2.2
(Data Capture & Compl) |
|
|
|
|
|
3.2.5.4 |
|
|
|
The HMI Instrument shall receive all Commands and distribute all
housekeeping telemetry over the Observatory 1553 interface |
|
|
HMI, C&DH, FSW |
2.6.8
(S/C Arch) |
|
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3.3 |
|
|
SHARPP |
|
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|
|
3.3.1 |
|
|
Alignment & Jitter |
|
|
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|
|
3.3.1.1 |
3/27/03 |
|
|
|
The SHARPP components shall meet the performance requirement
specified in the SDO Observatory Alignment Budget (Doc. # TBD). |
|
SHARPP |
1.3.1,
1.3.2 (Ang Res & Coverage) |
|
|
|
|
|
3.3.1.2 |
3/27/2003
- Delete |
|
|
|
The internal AIA alignment shift due to launch shift, thermal
effects, and uncorrected 1 g sag shall be less than TBD (30??) arcsec in the
Y and Z axes |
Instrument allocation of 60 arcsec absolute pointing error |
SHARPP |
1.3.1,
1.3.2 (Ang Res & Coverage) |
|
|
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|
|
3.3.1.3 |
|
|
|
|
The SHARPP AIA Instruments shall provide 12 dB disturbance
rejection from mounting feet to focal plane |
Based on attenuation required from 5 arcsec spacecraft jitter
req to AIA 1.32 arcsec (3 sigma) jitter req |
SHARPP |
1.3.3
(Ang Res & Coverage) |
|
|
|
|
|
3.3.1.4 |
|
|
|
|
SHARPP shall provide two guide telescopes (primary and redundant
units), each with a noise equivalent angle of 1 arcsec and an update
frequency of at least 10 Hz. |
|
SHARPP |
1.3.4
(Ang Res & Coverage) |
|
|
|
|
|
3.3.1.5 |
|
|
|
|
During nominal science operations, the SHARPP instrument shall
limit any disturbance contribution to the Observatory jitter to within TBD (1
arcsec??) |
|
SHARPP |
2.5.5
(Pointing & Jitter Control) |
|
|
|
|
|
3.3.1.6 |
03/27/03
- New |
|
|
|
The Guide Telescope shall be designed such that the science
sun acquisition can be performed given an initial pointing error of 60 arcsec
(3 sigma) in the y and z axes. |
|
SHARPP |
2.5.5
(Pointing and Jitter Control) |
|
|
|
|
|
3.3.2 |
|
|
Angular Resolution |
|
|
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|
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|
3.3.2.1 |
|
|
|
|
SHARPP AIA Instruments shall provide cameras with a width of at
least 1.4*960*2/.66=4073 pixels to achieve the required atmospheric image
resolution |
|
SHARPP |
1.3.3
(Ang Res & Coverage) |
|
|
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|
|
3.3.2.2 |
|
|
|
|
SHARPP KCOR Instrument shall provide a camera with a width of at
least 15*960*2/14 =2057 pixels to achieve required coronagraphic image
resolution |
Initial
information indicates that SHARPP will only be 2048 by 2048- need to address
this |
SHARPP |
1.3.4
(Ang Res & Coverage) |
|
|
|
|
|
3.3.3 |
|
|
Timing |
|
|
|
|
|
|
|
|
|
|
3.3.3.1 |
|
|
|
The SHARPP Instrument shall utilize the relative Spacecraft
timing signals and accuracy provided over the 1553 bus |
Time
Accuracy .01 seconds for relative. Absolute? |
Does the SHARPP instr have a need to receive absolute time
broadcasts over the 1553 bus in addition to relative time pulses? |
SHARPP |
1.5
(Precision) |
|
|
|
|
|
3.3.3.2 |
03/27/03
- New |
|
|
|
Absolute time accuracy 0.1 seconds. |
|
SHARPP |
1.5
(Precision) |
|
|
|
|
|
3.3.4 |
|
|
Data Completeness |
|
|
|
|
|
|
|
|
|
|
3.3.4.1 |
3/27/03 |
|
|
|
The SHARPP science data bit error rate shall be less than 5x10-8 (TBR) |
Reflects the instrument component of the 99.9% data science
completeness budget (1x10-7 total budget) |
SHARPP |
2.5.3
(Data Capture & Compl) |
|
|
|
|
|
3.3.5 |
|
|
Interface Requirements |
|
|
|
|
|
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|
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|
|
3.3.5.1 |
|
|
|
|
The SHARPP science data shall not exceed of maximum data rate
allocation of 72 Mbps over the IEEE 1355 high rate science data bus |
Defines the SHARPP allocation of the 130 Mbps (150 Mbps
post-encoding & margin) science data downlink |
SHARPP, C&DH |
1.6.4
, 1.6.5 (Cadence) |
|
|
|
|
|
3.3.5.2 |
|
|
|
Any
compression of the SHARPP data shall be performed in such a manner that
SHARPP preserves the data quality commensurate with the science analysis and
completeness requirements. |
|
This
addresses need for SHARPP science data compression to meet the science data
rate req |
SHARPP |
1.2.2
(Data Capture & Compl) |
|
|
|
|
|
3.3.5.3 |
|
|
|
The SHARPP instruments shall adhere to the high speed bus data
rate and interface requirements detailed in the SDO/SHARPP High Rate Science
Bus Interface Specification (Doc # TBD) |
|
|
SHARPP, C&DH |
1.2.2
(Data Capture & Compl) |
|
|
|
|
|
3.3.5.4 |
|
|
|
The SHARPP Instrument shall receive all Commands and distribute
all housekeeping telemetry over the Observatory 1553 interface |
|
|
SHARPP, C&DH, FSW |
2.6.8
(S/C Arch) |
|
|
|
|
|
3.3.6 |
3/27/03 |
|
Dynamic Range |
SHARPP AIA instruments shall provide cameras with a dynamic
range of at least 13 bits. SHARPP KCOR shall provide a camera with a dynamic
range of at least 14 bits. |
SHARPP AIA instruments shall provide cameras with a dynamic
range of at least 13 bits. SHARPP KCOR shall provide a camera with a dynamic
range of at least 14 bits. |
This addresses need for SHARPP science data to meet the required
contrast levels based upon photon statistics, noise and expected count rates. |
SHARPP |
1.5
(Precision) |
|
|
|
|
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|
|
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|
|
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|
|
4 |
|
|
Spacecraft Requirements |
|
|
|
|
|
|
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|
|
4.1 |
|
|
Structural /Thermal |
|
|
|
|
|
|
|
|
|
|
4.1.1 |
|
|
Launch Vehicle Accommodation |
The mechanical structure of the Observatory shall accommodate
the constraints of the launch vehicle |
N/A |
N/A |
MECH |
2.4
(LV) |
|
|
|
|
|
4.1.1.1 |
3/27/03 |
|
|
|
The SDO Observatory shall fit within the static and dynamic
envelope of the Delta 4040 or the ATLASV 401 fairing |
Reference the "DELTA IV PAYLOAD PLANNERS GUIDE" for
this and the subsequent req |
MECH |
2.4
(LV) |
|
|
|
|
|
4.1.1.2 |
3/27/03 |
|
|
|
The SDO Observatory shall
meet the interface requirements of the Delta 4040 or the ATLASV
401 PAF |
|
MECH |
2.4
(LV) |
|
|
|
|
|
4.1.1.3 |
|
|
|
|
The stowed Observatory shall have a C.M. < 4 m above the
separation plane |
|
MECH |
2.4
(LV) |
|
|
|
|
|
4.1.1.4 |
3/27/03 |
|
|
|
The SDO Observatory shall be designed to meet the mechanical,
thermal, and EMI requirements of the Delta 4040 or the ATLASV
401 LV as defined in the "DELTA IV PAYLOAD
PLANNERS GUIDE" |
|
MECH |
2.4
(LV) |
|
|
|
|
|
4.1.1.5 |
|
|
|
|
The
stowed Observatory shall be static and dynamically balanced to meet the
requirements defined in the "DELTA IV PAYLOAD PLANNERS GUIDE" |
|
MECH |
2.4
(LV) |
|
|
|
|
|
|
|
|
|
|
|
|
|
|
|
|
|
|
|
4.1.2 |
|
|
Spacecraft and Instrument Accommodation |
|
N/A |
N/A |
|
2.6.1
(Stand S/C Services- Mech) |
|
|
|
|
|
4.1.2.1 |
|
|
|
The structure shall provide sufficient area to mount all
electronics boxes and components and provide necessary provision for required
thermal radiators |
|
|
MECH, THERM |
2.4
(LV) |
|
|
|
|
|
4.1.2.2 |
|
|
|
The Observatory shall be designed such that the CM migration
over the life of the mission does not
eliminate attitude control authority from any required thruster |
|
All SDO mass changes shall be regularly monitored and assessed
against the CM requirement and a CM reassessment conducted when warranted to
verify that CM still remains within specified parameters |
MECH, GN&C, PROP |
2.4
(LV), 4.2.4 (Delta-V) |
|
|
|
|
|
4.1.2.3 |
|
|
|
|
The structure shall be electrically conductive to a measured
resistance of less than 2.5 mOhms across electronic component mounting
surfaces |
Addresses Observatory electronics grounding requirements |
MECH, THERM, ELEC |
2.7.4
(Elec) |
|
|
|
|
|
4.1.2.4 |
|
|
|
The structure shall provide a clear field of view for all
instrument and relevant components (Instrument CCD thermal radiators, ACS
sensors, thrusters, and omni and high gain antennas) within the specified
parameters required by that component |
|
|
MECH, THERM |
2.6.1
(Stand S/C Services- Mech), 3.1.1, 3.2.1, 3.3.1 (Instr FOV) |
|
|
|
|
|
4.1.2.5 |
|
|
|
|
Thruster
plume impingement avoidance angle from all Observatory components shall be
> 30 deg (TBR) half angle measured from the center line of the thruster
nozzle |
Addresses
plume impingement resulting in disturbance effects. Contamination is affected
by this requirement and needs to be evaluated |
CONTAM, PROP, MECH, GN&C |
Lev. 1
[Precision (Science Meas. 1-6)], 2.2.1
(Mission Life) |
|
|
|
|
|
4.1.2.6 |
3/27/03 |
|
|
|
Thruster
plume impingement heating effect avoidance angle from all Observatory
components shall be > 45 deg (TBR) half angle measured from the center
line of the thruster nozzle |
Addresses
thruster heating effects; may be waived if appropriate thermal protection
materials are used. Contamination is
affected by this requirement and needs to be evaluated |
CONTAM, PROP, MECH, GN&C, THERM, MATL |
Lev.
1 [Precision (Science Meas. 1-6)],
2.2.1 (Mission Life) |
|
|
|
|
|
4.1.2.7 |
|
|
|
The Observatory structure shall be designed for appropriate
shipment accommodation |
|
Vehicle shipment, lifting & handling, etc |
MECH |
2.6.1
(Stand S/C Services- Mech) |
|
|
|
|
|
4.1.3 |
|
|
Thermal Monitoring and Control |
|
N/A |
N/A |
|
2.6.2
(Stand S/C Services- Therm) |
|
|
|
|
|
4.1.3.1 |
|
|
|
The Observatory thermal control system shall be designed to
provide Observatory components with thermal radiators (where required) with
adequate radiator area to keep component thermal rejection requirements |
|
|
THERM, MECH |
2.6.2
(Stand S/C Services- Therm) |
|
|
|
|
|
4.1.3.2 |
|
|
|
The Observatory shall provide operational temperature control
capability in order to allow the Observatory to maintain the Observatory
components and structure within operational ranges |
|
|
THERM |
2.6.2
(Stand S/C Services- Therm) |
|
|
|
|
|
4.1.3.3 |
|
|
|
The Observatory shall provide autonomous temperature control
within survival temperature limits at all times even in the event of
operational heater monitoring and control failure |
|
|
THERM |
2.6.2
(Stand S/C Services- Therm) |
|
|
|
|
|
4.1.3.4 |
|
|
|
|
Survival heaters shall be sized
for a minimum bus voltage of 24V (TBR) and above this voltage they
shall keep components within their survival limits indefinitely |
Based on minimum bus voltage due to battery cell failures |
THERM, PWR |
2.6.2
(Stand S/C Services- Therm), 2.1.2.2 (Fault Tol) |
|
|
|
|
|
4.1.3.5 |
|
|
|
|
Unless otherwise noted, the operational baseplate temperature
for spacecraft electronics components shall be 0 deg to +40 deg C (consistent
with appropriate thermal-induced reliability concerns) |
|
THERM, Electrical Subsystems |
2.6.2
(Stand S/C Services- Therm) |
|
|
|
|
|
4.1.3.6 |
|
|
|
|
The propulsion system components containing propellant or
pressurant shall be maintained within the range of 10-40 deg C at all times |
Prevents damage of prop system due to effects of temperature
extremes on prop tank and lines. |
THERM, PROP |
2.6.2
(Stand S/C Services- Therm) |
|
|
|
|
|
4.1.3.7 |
|
|
|
|
The reaction wheels shall be maintained within a range of TBD (10-30 deg??) C at all times |
Recommended temperature range for optimal performance and
extended reliability and operational life |
THERM, GN&C |
2.6.2
(Stand S/C Services- Therm) |
|
|
|
|
|
4.1.3.8 |
3/27/03 |
|
|
|
The battery shall be maintained within a range of 10-30o C (TBR) at all times |
Recommended temperature range for optimal performance and
extended reliability and operational life |
THERM, PWR |
2.6.2
(Stand S/C Services- Therm) |
|
|
|
|
|
4.1.3.9 |
|
|
|
|
The EVE telescope assemblies shall be maintained within a
temperature range of TBD and will not exceed a rate of change of TBD |
Maintains EVE Instrument optics within nominal operational
temperature range and change rates to prevent
thermal deformation |
THERM, EVE |
2.6.2
(Stand S/C Services- Therm) |
|
|
|
|
|
4.1.3.10 |
|
|
|
|
The HMI telescope assemblies shall be maintained within a
temperature range of TBD and will not exceed a rate of change of TBD |
Maintains HMI Instrument optics within nominal operational
temperature range and change rates to prevent
thermal deformation |
THERM, HMI |
2.6.2
(Stand S/C Services- Therm) |
|
|
|
|
|
4.1.3.11 |
|
|
|
|
The SHARPP telescope assemblies shall be maintained within a
temperature range of TBD and will not exceed a rate of change of TBD |
Maintains SHARPP Instrument optics within nominal operational
temperature range and change rates to prevent
thermal deformation |
THERM, SHARPP |
2.6.2
(Stand S/C Services- Therm) |
|
|
|
|
|
4.1.3.12 |
|
|
|
|
The Observatory shall provide Instrument CCDs radiators with an
equivalent sink temperature of -150 deg (TBR) in order to meet Instrument CCD
temperature requirements |
May need to break this req into separate instrument reqs- will
be done after individual instrument CCD temp requirements become known |
THERM, INSTR |
2.6.2
(Stand S/C Services- Therm) |
|
|
|
|
|
4.1.3.13 |
|
|
|
Nominal
on-station Observatory maneuvers shall not result in direct sunlight on
Instrument thermal radiators for a time period such that critical internal
instrument temperatures rise beyond acceptable levels |
|
Addresses instrument keep-out zones during pre-planned
on-station maneuvers |
GN&C, THERM |
1.2.1
(Data Capture & Compl), 4.1.3.9, 4.1.3.10 (Instr Therm limits) |
|
|
|
|
|
4.1.3.13.1 |
3/27/03
- New |
|
|
|
Stationkeeping and momentum unwading maneuvers at GEO shall
be designed to limit the maximum angle between the Sun and the x-axis to less
than 45 degrees (TBR), for a duration of no longer than 30 minutes
(TBR). This constraint may be violated
in the event of a thruster failure or spacecraft emergency. |
|
GN&C, THERM |
1.2.1
*Data Capture and Completeness, 4.1.3.9, 4.1.3.10 (Instrument Thermal Limits) |
|
|
|
|
|
4.1.3.14 |
|
|
|
|
The Solar Arrays shall be designed to operate within a
temperature range of TBD |
|
THERM, PWR |
2.6.2
(Stand S/C Services- Therm) |
|
|
|
|
|
4.1.3.15 |
|
|
|
|
The coarse sun sensors (CSSs) shall be designed to operate
within a temperature range of TBD |
|
THERM, GN&C |
2.6.2
(Stand S/C Services- Therm) |
|
|
|
|
|
4.1.3.16 |
|
|
|
|
The HGAs, RF waveguide system, and gimbal assemblies shall be
designed to operate within a temperature range of TBD |
|
THERM, APS, DEPLOY, RF |
2.6.2
(Stand S/C Services- Therm) |
|
|
|
|
|
4.1.3.17 |
|
|
|
|
All Solar Array deployment components (hinges, deployment
devices, etc) shall be designed to operate within a temperature range of TBD |
|
THERM, PWR, DEPLOY |
2.6.2
(Stand S/C Services- Therm) |
|
|
|
|
|
4.1.3.18 |
|
|
|
|
All HGA deployment components (hinges, deployment devices, etc)
shall be designed to operate within a temperature range of TBD |
|
THERM, DEPLOY, APS, RF |
2.6.2
(Stand S/C Services- Therm) |
|
|
|
|
|
4.1.3.19 |
|
|
|
|
All Instrument aperture cover mechanisms (hinges, deployment
devices, reclosable mechanism control, etc) shall be designed to operate
within a temperature range of TBD |
|
THERM, INSTR |
2.6.2
(Stand S/C Services- Therm) |
|
|
|
|
|
4.1.3.21 |
|
|
|
|
The Spacecraft shall provide temperature monitoring thermistors
for all Observatory components and critical structural surfaces in order to
maintain knowledge of Observatory temperatures |
|
THERM, Electrical Subsystems |
2.6.2
(Stand S/C Services- Therm), 22.6.7 (H/K Telemetry) |
|
|
|
|
|
4.1.3.22 |
|
|
|
|
The Instrument optical bench and instrument telescope assemblies
shall be designed to survive any angle of offpointing after launch and
through orbit circularization for period TBD (1.5 hrs) and return to proper
alignment with no degradation of science performance |
Addresses worst case offpointing for GEO-insertion maneuvers and
the resultant thermally induced alignment errors. Need to clarify whether Instruments are
powered or unpowered during this time |
MECH, THERM, INSTR, GN&C |
1.3
(Ang Res & Coverage), 4.1.2.4 (Struct Instr Accomm) |
|
|
|
|
|
|
|
|
|
|
|
|
|
|
|
|
|
|
|
4.1.4 |
|
|
Instrument Optical Bench |
The Observatory structure shall provide an optical bench to
mount all instrument telescopes & star trackers in such a manner that
meets mounting, FOV, and thermal requirements |
N/A |
N/A |
MECH, THERM |
2.6.1.
2.6.2 (Stand S/C Services- Mech & Therm) |
|
|
|
|
|
4.1.4.1 |
|
|
|
The instrument optical bench shall provide all instrument
telescopes with an optical FOV free of obstruction and glint from other
portions of the Observatory |
|
Instrument FOV requirements detailed in Instrument/Observatory
mechanical ICDs (doc #'s TBD). See Instrument FOV
requirements in Section 3 |
MECH |
4.1.2.4
(Struct Instr Accomm) |
|
|
|
|
|
4.1.4.2 |
|
|
|
The instrument optical bench shall provide all instrument
thermal radiators with suitable thermal FOV required to maintain instrument
CCDs and telescope electronics to a temperature within their specified
operating temperature ranges. |
|
See 4.1.3.10 for instrument CCD thermal reqs |
MECH, THERM |
4.1.2.4
(Struct Instr Accomm) |
|
|
|
|
|
4.1.4.3 |
|
|
|
The instrument optical bench shall provide a stable mounting
alignment environment between all instruments and the guide telescope that
meets the mission pointing requirements during nominal operations |
|
Reflects the instrument mechanical alignment component of the
overall instrument pointing budget |
MECH |
1.3
(Ang Res & Coverage), 4.1.2.4 (Struct Instr Accomm) |
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4.1.4.3.1 |
3/27/03 |
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Alignment Errors |
All alignment errors between each of the instrument mounting
interfaces and the GT interface shall meet the performance requirement
specified in the SDO Observatory Alignment Budget (Doc. Number TBD). |
Static error sources may include 1G effects, on-orbit thermal
settling, launch shifts, measurements errors, etc |
MECH |
1.3
(Ang Res & Coverage), 4.1.2.4 (Struct Instr Accomm) |
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4.1.4.3.2 |
3/27/2003
- Delete |
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Dynamic Alignment Errors |
The dynamic alignment error between each of the instrument
mounting interfaces and the GT interface shall be limited to TBD (5 arcsec??)
over a period of one week |
Dynamic error based on daily and seasonal effects |
MECH |
1.3
(Ang Res & Coverage), 4.1.2.4 (Struct Instr Accomm) |
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4.1.4.3.3 |
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High Frequency Alignment Errors |
The out of phase relative motion between the SHARPP instrument
boresights and the GT boresights shall not exceed an amplitude of TBD (0.05
arcsec??) for frequencies between 0.1 Hz and 10 Hz |
Addresses the need to avoid significant modes that would
interfere with the SHARPP instrument motion compensation system (HMI has own
internal closed-loop IMC system, while EVE has no jitter requirements) |
MECH |
1.3
(Ang Res & Coverage), 4.1.2.4 (Struct Instr Accomm) |
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4.1.4.3.4 |
3/27/03 |
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The instrument optical bench design shall allow all instruments
to return to within their specified alignment within 60 minutes after a
planned maneuver or eclipse |
Based on preliminary Instrument inputs. Reflects the
thermal-induced alignment error (instrument optics, mechanical alignment)
during eclipses, offpointing, maneuvers, etc
This also places a requirement on sun impingement and shadowing due to
offpointing. |
MECH, THERM, GN&C |
1.3
(Ang Res & Coverage), 4.1.2.4 (Struct Instr Accomm), 1.2 (Data Capture
& Compl) |
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4.1.4.4 |
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Instrument
Ground Test Access |
The
instrument optical bench shall provide access for all required GSE and test
& support equipment |
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MECH, GSE |
4.1.2.4
(Struct Instr Accomm) |
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4.1.4.4.1 |
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The Instrument telescopes shall have access provided for all GSE
and purge lines required during I&T |
Req may need to be modified to indicate that all purge lines
connect to single accessible connector plate |
MECH, CONTAM |
2.3
(Envir), 4.1.2.4 (Struct Instr Accomm) |
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4.2 |
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Attitude Control & Determination |
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4.2.1 |
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Acquisition |
The Observatory shall have the capability to acquire the sun and
maneuver to a sun pointing orientation |
N/A |
N/A |
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2.2.2
(Fault Tol), 2.6.3 (Power Dist), 2.6.4 (Attitude Control) |
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4.2.1.1 |
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The Observatory shall be pointed to the sun within the accuracy
defined in 4.2.6.1 from any initial orientation within 30 minutes after a
spacecraft pointing anomaly or a specific ground command to enter the sun
acquisition state |
|
GN&C |
2.2.2
(Fault Tol), 2.6.3 (Power Dist), 2.6.4 (Attitude Control) |
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4.2.1.2 |
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The
Observatory shall be pointed to the sun within the accuracy defined in
4.2.6.1 from any initial orientation within 45 minutes after Observatory
separation from the LV. |
This
requirement infers 15 minutes unload momentum with use of thrusters (as
referenced in 4.2.5.1) and 30 minutes to acquire sun. Req driven by battery charge depletion
concerns. Safehold pointing accuracy defined in 4.2.6.1. |
GN&C, LV |
2.2.2
(Fault Tol), 2.6.3 (Power Dist), 2.6.4 (Attitude Control), 2.4.2.4 (LV Perf) |
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4.2.1.3 |
3/27/03 |
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Sun acquisition shall be accomplished without the use of
thrusters for initial post LV separation rates of up to 1,
2, 2 deg/sec
(3 sigma) (x, y, z) |
Drives sizing of RWs.
Initial momentums determined by the LV |
GN&C, LV |
2.1.2
(Mission Life), 2.1.4.2 (LV Perf) |
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4.2.2 |
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Pointing Knowledge |
The Observatory shall provide sufficient pointing knowledge both
for enabling science observations and for orbit maintenance |
N/A |
N/A |
|
1.3
(Ang Res & Coverage), 2.6.4 (Attitude Control) |
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4.2.2.1 |
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The Observatory shall provide pointing knowledge of 25 arcsec (3
sigma) with regard to the geocentric inertial frame in the X, Y and Z axes
during science observation modes. |
Need to understand the need for this requirement from the
instrument team- currently no known rationale |
GN&C, MECH, INSTR |
?? |
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4.2.3 |
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Attitude Control and Stability |
The
Spacecraft shall provide sufficient attitude control and stability to enable
instrument science observations |
N/A |
N/A |
|
1.3
(Ang Res & Coverage), 2.5.5 (Pointing & Jitter Control) |
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4.2.3.1 |
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The attitude control system shall provide a ground commanded
capability to point to any point in inertial space. |
|
Used not only to acquire sun pointing, but also for offpointing
and calibration maneuvers |
GN&C |
1.5
(Precision), 2.6.4 (Attitude Control) |
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4.2.3.2 |
3/27/03 |
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The Spacecraft shall point the instrument science reference
boresight (KCOR boresight, as designated in 2.5.5.1) to a target (nominally
the center of the sun) to an accuracy 60 arcsec (3 sigma) with regard to the geocentric inertial frame in
the X, Y, & Z axes over the period of the mission. Pointing requirement shall be met without
the use of the GT. |
Addresses
the need for the Observatory to first allow the GT to acquire the sun to
allow its utilization in Observatory pointing. 50 arcsec req derived from the GT linear
range. |
GN&C, SHARPP |
1.3
(Ang Res & Coverage), 2.5.5 (Pointing & Jitter Control), 2.6.4
(Attitude Control) |
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4.2.3.3 |
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The attitude control system shall provide 2 arcsec (3 sigma)
steady-state pointing accuracy (exclusive of jitter effects) relative to a GT
error signal |
Required to meet HMI 1 week pointing |
GN&C |
1.3
(Ang Res & Coverage), 2.5.5 (Pointing & Jitter Control), 2.6.4
(Attitude Control) |
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4.2.3.4 |
3/27/03 |
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The Spacecraft ACS shall be able to execute a 180 degree slew
and settle to Observatory pointing accuracy (as indicated in 4.2.3.2) within
a period of 20 minutes |
Used to bound the accuracy and time period required for
Observatory calibration maneuvers and setup for orbit
adjust maneuvers. 20 minute req driven by worst case Geo burn
offpointing and 90 min battery charge allocation (20 min initial slew, 40 min
burn, 20 return slew). Excludes
momentum stored as result of initial tipoff rates |
GN&C |
2.6.4
(Attitude Control) |
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4.2.3.5 |
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During science operations, the rotation of the Spacecraft YZ
plane shall be maintained at a constant angle (within the accuracy
requirements of 4.2.3.2) relative to the solar north pole |
Addresses roll control requirement in relation to maintaining
the Observatory X axis (instrument boresight) in the proper sun pointing
orientation |
GN&C |
1.3
(Ang Res & Coverage), 2.6.4 (Attitude Control) |
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4.2.4 |
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Propulsion & Delta-V |
The Spacecraft shall provide the capability for orbit insertion,
orbit maintenance, and Observatory disposal |
N/A |
N/A |
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2.1
(Orbit), 2.4 (LV) |
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4.2.4.1 |
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Propulsion System |
The Spacecraft shall implement an onboard propulsion system with
sufficient propellant budget for all phases of the mission |
|
Mission Phases referenced in 2.5.8 |
GN&C, PROP |
2.1
(Orbit), 2.4 (LV) |
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4.2.4.1.1 |
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The thruster locations and orientations shall be selected to
provide Delta-V functionality over the mass properties configuration
throughout the mission life |
|
Addresses the effect of changing CM over the mission life (due
to fuel depletion) and the effect on thruster placement and effectiveness |
GN&C, PROP, FLT DYN, MECH |
4.1.2.2
(S/C CM vs Thruster placement) |
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4.2.4.1.2 |
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The detailed propellant budget shall be documented in the SDO
Propulsion System Delta-V Budget Spreadsheet (doc # TBD) |
|
Configured Delta-V budget spreadsheet covers all phases of
mission operation and Observatory disposal in accordance with a five year
mission life. Covers thruster calibration, tip-off correction, orbit
circularization, periodic RW momentum unloading and on-station maneuvers, and
Observatory disposal |
GN&C, PROP, FLT DYN |
4.2.4.1
(Delta-V) |
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4.2.4.1.3 |
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The Propulsion system shall include thruster & isolation
valve inhibits for ground test and pre-launch operations |
|
Prevents inadvertent firing of thrusters. Inhibits include arm/fire command sequence,
ground safing procedures, as well as arm plugs which must be in place to fire
thrusters ( test plugs provide GSE simulation of thruster firing in ground
testing in lieu of actual thruster valve actuation, allowing testing of
thruster hardware and S/W). See 4.6.1.3 for related guidance |
GN&C, PROP, GSE |
2.2.2
(Fault Tol), 4.2.4.1 (Delta-V), 2.7.7 (Safety) |
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4.2.4.1.4 |
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The thruster catalyst bed heaters shall have proper thermal
control protection in place to prevent inadvertent overheating during ground
testing. |
|
Prevents catalyst bed damage due to overheating outside vacuum
environment. Protection includes
ground safing procedures, catbed temperature monitoring and alerts. See 4.6.1.3 for related guidance |
GN&C, PROP, THERM |
2.2.2
(Fault Tol), 4.2.4.1 (Delta-V), 2.7.7 (Safety) |
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4.2.4.2 |
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Delta-V Maneuver Predictability |
The execution of delta-V maneuvers shall be sufficiently
predictable to achieve the required mission trajectory within the propellant
allotment |
|
Requirement reflects the fact that the predictability of the
maneuver outcome affects the size of the delta-V maneuvers and the propellant
required |
GN&C, PROP, FLT DYN, GND |
2.4
(LV), 2.4.2 (Delta-V) |
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4.2.4.3 |
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Notification |
The Observatory shall provide notification to the instruments
over the 1553 bus prior to any thruster operations |
|
Any planned thruster operations will be included to the SOCs as
part of mission planning notification.
In preplanned thruster modes, the req implementation will provide some
advance lead time prior to the action.
This req also addresses unplanned emergency thruster firings and
therefore in those cases will result in extremely short notification lead
time. Allows instruments the opportunity
to reconfigure in response to pending action |
GND, FSW, C&DH |
2.7.7 (Safety) |
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4.2.5 |
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Momentum Management |
The Spacecraft shall be designed to provide the necessary
momentum management for all phases of the Observatory lifetime |
N/A |
N/A |
|
2.5.5
(Pointing & Jitter Control) |
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4.2.5.1 |
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The Spacecraft shall have the capability to use thrusters to
remove residual angular momentum within 15 minutes after LV separation to
within the rates specified by 4.2.1.3 |
Addresses
the need to use ground commanded thrusters to reduce Observatory rates to a
level within initial acquisition capability in case of higher initial LV
induced rates. |
GN&C, PROP |
2.2.2
(Fault Tol), 2.6.3 (Power Dist), 2.6.4 (Attitude Control), 2.4.2.4 (LV Perf) |
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4.2.5.2 |
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The combined momentum unloading and stationkeeping maneuvers
shall have a frequency of no more than once every 4 weeks |
Defines analysis, accuracy, and predictability for Delta H and
Delta V effects |
GN&C, PROP, FLT DYN, GND |
2.5.3 (Data Comp), 4.2.4 (Delta-V),
4.2.4.1.2 (Delta-V Budget) |
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4.2.5.3 |
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Pointing control for all on-station momentum maneuvers shall be
within +/- 5 deg (TBR) (3-sigma) per axis |
Bounds the perturbation of the nominal Observatory pointing
during momentum unloading |
GN&C, PROP |
2.5.3 (Data Comp), 2.5.5 (Pointing &
Jitter Control) |
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4.2.6 |
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Safehold |
The Spacecraft design shall provide an autonomous mode used to
maintain a power-safe, thermal-safe, and instrument-safe environment in the
event of Observatory anomalies |
|
Instrument safing provided prior to offpointing, load-shedding,
etc when possible. |
GN&C |
2.2.2
(Fault Tol), 2.5.6 (Autonomy), 2.6.3 (Power Dist), 2.6.4 (Attitude Control) |
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4.2.6.1 |
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The Spacecraft shall be designed to achieve an orientation
placing the solar arrays within +/- 15 degrees normal to the sun from any
initial orientation |
Addresses
need to remain power positive during safehold. Infers no requirement for rotation about
the sun line while in S/H. |
GN&C |
2.2.2
(Fault Tol), 2.6.3 (Power Dist), 2.6.4 (Attitude Control) |
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4.2.6.2 |
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The pointing attitude within the accuracy defined in 4.2.6.1
shall be attained within 30 minutes of entering this operational mode or
exiting an eclipse period |
Based on preliminary GN&C team estimate. Addresses limited battery life, drives RW
sizing. Note that eclipse is estimated
to be up to 70 minutes in duration (plus 30 minute sun acq safehold for worst
case condition), affecting battery life reqs. |
GN&C |
2.2.2
(Fault Tol), 2.6.3 (Power Dist), 2.6.4 (Attitude Control) |
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4.2.6.3 |
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The Spacecraft shall be designed with the capability to remain
in safehold indefinitely with ground intervention |
Addresses the autonomous nature of maintaining safehold
operational mode. Implies RW momentum
unloading (which requires ground commanding) in order to maintain required
pointing after momentum buildup |
GN&C, GND |
2.2.2
(Fault Tol), 2.6.3 (Power Dist), 2.6.4 (Attitude Control) |
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4.2.6.4 |
3/27/03 |
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The Spacecraft shall be designed to maintain safehold pointing
without any ground intervention for a period of at least 7 days during GEO operations. |
Addresses momentum management margin required as part of regular
unloading period- need to plan unloading with sufficient margin to still
maintain pointing if safehold entered immediately prior to regularly
scheduled momentum unloading maneuver.
Infers that Safehold pointing requirement is in addition to 4 week
momentum unloading req in 4.2.5.2 (resulting in 5 weeks of momentum buildup) |
GN&C |
2.2.2
(Fault Tol), 2.6.3 (Power Dist), 2.6.4 (Attitude Control) |
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4.2.6.4.1 |
03/27/03
- New |
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Maintain safehold pointing for >/- 2 orbits with no ground
intervention during transfer orbit. |
|
GN&C |
2.2.2
(Fault Tol), 2.6.3 (Power Dist), 2.6.4 (Attitude Control) |
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4.2.6.5 |
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Autonomous safehold operations should not be impeded by the loss
of the central spacecraft flight processor or the 1553 spacecraft data bus
and shall be altogether independent of these elements |
|
Addresses the autonomous and independent performance aspects
defined by safehold. Indicates SDO
implementation decision that safehold shall be controlled by an independent
processor and a path of safehold component control independent of the 1553
bus or similar shared data bus. |
GN&C |
2.2.2
(Fault Tol), 2.6.3 (Power Dist), 2.6.4 (Attitude Control) |
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4.2.6.6 |
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The Observatory shall provide notification to the instruments
over the 1553 bus when Safehold has been entered |
|
Allows instruments the
opportunity to reconfigure in response to action. Notification will probably be implemented
as the non-transmission of S/C message (since 1553 bus may be disabled prior
to entering S/H |
GND, FSW, C&DH, GN&C |
2.2.2
(Fault Tol), 2.6.3 (Power Dist), 2.6.4 (Attitude Control) |
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4.3 |
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Power |
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N/A |
N/A |
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4.3.1 |
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|
Power Generation |
Provide sufficient power generation to support the Observatory
through all the phases of the mission |
N/A |
N/A |
PWR |
2.1
(Orbit), 2.2 (Mission Life), 2.6.3 (Standard S/C Services- Power) |
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4.3.1.1 |
3/27/03 |
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The power system shall have the capability to support 1325 W (TBR) orbital average
power load at the end of the 5 year mission life for the geosynchronous orbit
selected |
Assumes 20% margin for growth |
PWR |
2.1
(Orbit), 2.2 (Mission Life), 2.6.3 (Standard S/C Services- Power) |
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4.3.2 |
|
|
Power Storage |
Provide sufficient energy storage to support the Observatory
through all the phases of the mission |
N/A |
N/A |
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4.3.2.1 |
3/27/03 |
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The battery shall not exceed a total of 550 cycles of 60% depth of discharge as part of qualification and acceptance
testing, pre-launch testing and mission ops. |
Provides protection against Li-Ion battery degradation through
large number of deep discharge cycles |
PWR |
2.1
(Orbit), 2.2 (Mission Life), 2.6.3 (Standard S/C Services- Power) |
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4.3.2.2 |
3/27/03 |
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100 battery cycles shall be allocated
to pre-launch and acceptance testing, with the remaining cycles reserved for
flight accommodation |
Addresses the number of battery cycles in 4.3.2.1 allocated for
ground operations |
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4.3.2.3 |
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The battery shall be able to provide at least 600 watts (TBR)
for 90 minutes (TBR) as part of the launch phase in order to allow the
Observatory to acquire power positive orientation after launch |
Defined by launch worst case assumptions- include 15 minute internal power hold prior
to launch, 30 minutes post launch prior to separation, 15 minutes to unload
momentum with use of thrusters, and 30 minutes to achieve safehold pointing
(90 minutes) |
PWR |
2.1
(Orbit), 2.2 (Mission Life), 2.6.3 (Standard S/C Services- Power) |
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4.3.2.4 |
3/27/03 |
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The battery shall be able to provide at least TBD (1325 W) for 72 minutes at EOL
to support worst case eclipse condition without load shedding
instruments |
Addresses worst case eclipse conditions |
PWR |
2.2
(Mission Life), 2.5.3 (Data Comp),
2.6.3 (Standard S/C Services- Power) |
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4.3.2.5 |
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The battery shall be able to provide power for Spacecraft
operations for 102 minutes at EOL to support worst case eclipse
conditions. In order to meet this
requirement, load shedding is permitted |
Defined by the worst-case eclipse duration (72 minutes), worst
case safehold acquisition (30 minutes) plus margin. Note that instruments will be powered at
this time and may need to be load-shed if necessary. |
PWR |
2.2
(Mission Life), 2.6.3 (Standard S/C Services- Power) |
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4.3.3 |
|
|
Power Distribution |
The Spacecraft design shall provide the capability for power
distribution within the Observatory |
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4.3.3.1 |
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The
output of the power subsystem shall provide for the distribution 29V +/- 6V
power to all Observatory electronic subsystems and components. |
Defined
by power regulation specification of power subsystem. |
PWR |
2.6.3
(Standard S/C Services- Power) |
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4.3.3.2 |
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All electronic subsystems shall be qualified to a voltage bus
range of 22-35 V |
Difference between 4.3.3.1 assumes a 1 volt drop between PSE
output and component input |
Electrical Subsystems |
2.2.2
(Fault Tol), 2.6.3 (Standard S/C Services- Power) |
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4.3.3.3 |
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The spacecraft power system shall utilize a distributed power
architecture with primary high-current power distributed to the Observatory
subsystems for further distribution as needed |
|
Higher power components may receive power directly from PSE due
to current draw considerations |
PWR |
2.2.2
(Fault Tol), 2.6.3 (Standard S/C Services- Power), SDO Arch. Design |
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4.3.3.4 |
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|
|
The power subsystem shall provide dual power feeds to all
components as part of the operational and survival distribution, with
appropriate power isolation design incorporated in the event of power
distribution faults |
|
Reflects redundancy and reliability requirements need to meet
mission life |
PWR |
2.2.2
(Fault Tol), 2.6.3 (Standard S/C Services- Power) |
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4.3.3.5 |
|
|
|
Unswitched power services shall be appropriately fused to
prevent damage to or loss of the Spacecraft power system due to excess
current draw. Switched services shall
have remotely resetable "circuit breaker" for power system
protection. |
|
Safety and reliability design implementation in support of
redundant spacecraft design.
"Circuit breaker" function resettable by ground command |
PWR |
2.2.2
(Fault Tol), 2.6.3 (Standard S/C Services- Power), 2.7.7 (Safety) |
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4.3.3.6 |
3/27/03 |
|
|
|
The power subsystem shall provide the capability for a minimum of 24 (TBR) Observatory-controlled
switched and 6 (TBR) unswitched high power
services on both the primary and redundant power
output modules to meet Observatory requirements. |
Based on initial power services assessment. Implementation
detailed in power distribution diagram |
PWR |
2.2.2
(Fault Tol), 2.6.3 (Standard S/C Services- Power), SDO Arch. Design |
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4.3.3.7 |
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All
Electronic components shall be designed with redundant power inputs with
diode protection on each line |
|
Electrical Subsystems |
2.2.2
(Fault Tol), 2.6.3 (Standard S/C Services- Power), SDO Arch. Design |
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4.3.3.8 |
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|
|
The GSFC in-house subsystems/components shall use a common power
design implementation consisting of a processor-controlled secondary power
distribution system for lower power switched services, based on a common
dc-dc converter/power switching design implementation |
|
Requirement reflects project design decision to utilize common
distributed processor-controller power system design; offers advantages in
design commonality along with associated benefits in testability,
reliability, cost and manpower
reduction |
PWR |
2.2.2
(Fault Tol), 2.6.3 (Standard S/C Services- Power), SDO Arch. Design |
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4.3.3.9 |
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|
In addition to the primary and redundant power services, the
Spacecraft power system shall provide TBD (3??) switched CCD decontamination
heater service(s) to the SDO instruments |
Provides Observatory-controlled switched service to prevent
contamination deposits on instrument detectors during instrument powered-off
conditions in post-launch and survival modes, as well as opportunity to
perform on-flight CCD decontamination as necessary. Still need to work out the details of
Observatory vs Instrument in-flight heater control and total number of
services provided |
PWR, INSTR |
1.5
(Precision), 2.1.2 (Mission Life), 2.3.3 (Contam) |
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4.3.4 |
|
|
Load Shedding |
The Observatory shall be designed with the capability to execute
a pre-determined hierarchy of load shedding operations based on power system
status telemetry. This capability
shall be independent of the spacecraft central processor and the 1553 bus |
|
|
PWR |
2.2.2
(Fault Tol), 2.6.3 (Standard S/C Services- Power), SDO Arch. Design |
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4.3.5 |
|
|
Constraints |
Appropriate limits will be provided to avoid exceeding peak
discharge rates on the battery |
N/A |
N/A |
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4.3.5.1 |
3/27/03 |
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|
The torque commanded to each reaction wheel assembly shall be
software-limited such that the maximum power draw per wheel shall not exceed 300 watts/sec (TBR). |
Prevents RW current draw from exceeding Power subsystem output
module specifications. Is there an
additional requirement for power-saving power limitations during initial
post-launch sun acquisition to limit power drain on battery prior to
power-positive operations? |
GN&C, FSW, PWR |
2.6.3
(Standard S/C Services- Power) |
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4.3.5.2 |
3/27/03 |
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|
Simultaneous thruster firings shall be limited to the
combinations of either one 110lb engine and fpur 5lb
thrusters or eight 5lb thrusters in order to
limit the maximum instantaneous power draw |
Assumes 5 lb thrusters with ~ 10-15 W draw each and 25 lbs
thrusters with ~ 25 W draw each |
PWR, GN&C, FSW, PROP |
2.6.3
(Standard S/C Services- Power) |
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4.3.5.3 |
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|
All components shall design to appropriate in-rush current
requirements as defined in the SDO Electrical System Specification (doc #
TBD) |
|
Prevents power-on transients that could damage or loss to
Observatory power system or subsystem components |
PWR, ELEC, Electrical Subsystems |
2.7.4
(Elec) |
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4.4 |
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Comm & Data System |
N/A |
N/A |
N/A |
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2.6.5
(Standard S/C Services- Comm) |
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4.4.1 |
|
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S-Band Communications |
N/A |
N/A |
N/A |
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4.4.1.1 |
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|
Uplink |
A forward command link shall be provided to support Observatory
commanding |
N/A |
N/A |
GND, RF, C&DH |
2.6.5
(Standard S/C Services- Comm) |
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4.4.1.1.1 |
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|
The command uplink shall be compatible with the SDO ground
system, TDRSS, DSN and USN networks.
This requirement can be met using the GSTDN format. |
|
Addresses the potential need for multiple ground station
networks either for nominal ops or potential contingency operations |
GND, RF, C&DH |
2.6.5
(Standard S/C Services- Comm), 2.5.1 (Continuous Contact), 2.2.2 (Fault Tol) |
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4.4.1.1.2 |
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The command link shall utilize the CCSDS command path service
protocol; both COP-1 and bypass modes shall be supported |
|
|
GND, RF, C&DH, FSW |
2.6.5
(Standard S/C Services- Comm), 2.2.2 (Fault Tol) |
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4.4.1.1.3 |
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|
Full spherical command link coverage shall be supported through
all mission phases |
|
Defines
the use of dual omni antennas for spherical coverage. |
GND, RF, C&DH |
2.6.5
(Standard S/C Services- Comm) |
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4.4.1.1.4 |
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The mission shall support an uplink command rate of up to 2 Kbps |
|
GND, RF, C&DH, FSW |
2.6.5
(Standard S/C Services- Comm) |
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4.4.1.1.5 |
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|
The uplink bit error rate and command checking shall be such to
preclude the execution of an invalid command |
|
Implies bit error rate, link margin, command verification, and
command retransmission to maintain a robust command link |
GND, RF, C&DH |
2.6.5
(Standard S/C Services- Comm), 2.2.2 (Fault Tol) |
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4.4.1.1.6 |
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The S-Band command uplink shall have a link margin of ≥ TBD dB through all mission phases |
Margin helps ensure a stable and reliable RF uplink
communications path |
GND, RF |
2.6.5
(Standard S/C Services- Comm) |
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4.4.1.1.7 |
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The uplink command path shall include appropriate security
measures to prevent the execution of any unauthorized command sequences |
|
Reflects the need for Observatory security measures (need to
reference specific NASA requirement here) |
GND, C&DH, FSW |
2.6.5
(Standard S/C Services- Comm), NASA Security Guidance |
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4.4.1.1.8 |
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|
Each uplink receiver shall receive unswitched power from the
power system and shall not have the capability to be powered off |
Reflects robust spacecraft design requirement to maintain
command receive capability at the Observatory at all times |
C&DH, PWR |
2.6.5
(Standard S/C Services- Comm), 2.2.2 (Fault Tol) |
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4.4.1.2 |
|
|
Downlink |
A return link shall be provided to support Observatory
housekeeping telemetry transfer to the ground station |
N/A |
N/A |
GND, RF, C&DH |
2.6.5,
2.6.7 (Standard S/C Services- Comm & H/K Telem) |
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4.4.1.2.1 |
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|
The downlink shall be compatible with the SDO ground system,
TDRSS, DSN and USN networks. This
requirement can be met using the GSTDN format. |
|
Addresses the potential need for multiple ground station
networks either for nominal ops or potential contingency operations |
GND, RF, C&DH, FSW |
2.6.5,
2.6.7 (Standard S/C Services- Comm & H/K Telem) |
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4.4.1.2.2 |
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|
The housekeeping downlink telemetry shall support the CCSDS AOS
protocol |
The SDO baseline is CCSDS protocol |
GND, RF, C&DH |
2.6.5,
2.6.7 (Standard S/C Services- Comm & H/K Telem) |
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|
4.4.1.2.3 |
|
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|
|
Housekeeping telemetry shall comply with the CCSDS AOS Grade 2
service |
|
GND, RF, C&DH |
2.6.5,
2.6.7 (Standard S/C Services- Comm & H/K Telem) |
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|
4.4.1.2.4 |
|
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|
The S-Band downlink shall have a link margin of ≥ 3 dB
through all mission phases |
Margin helps ensure a stable and reliable RF downlink
communications path |
GND, RF, C&DH |
2.6.5,
2.6.7 (Standard S/C Services- Comm & H/K Telem) |
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4.4.1.2.5 |
|
|
|
The downlink rate shall support
the simultaneous downlink of both real time and stored housekeeping
data through all mission phases |
|
|
GND, RF, C&DH |
2.6.5,
2.6.7 (Standard S/C Services- Comm & H/K Telem) |
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|
4.4.1.2.6 |
|
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|
|
The downlink shall have the capability to allow recovery of 24
hours of stored housekeeping data while simultaneously downlinking real-time
data to the ground station |
Implies the need for acceptable bit error rates, SSR EDAC, data
retransmission, and recorder management such that Observatory data is not
released until valid data receipt on the ground is confirmed |
GND, RF, C&DH |
2.6.5,
2.6.7 (Standard S/C Services- Comm & H/K Telem), 2.2.2 (Fault Tol) |
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|
4.4.1.2.7 |
|
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|
The downlink rate and mission design shall support the
simultaneous downlink of real-time housekeeping and 24 hours of stored
housekeeping data at Geo within 60 minutes (TBR) |
Propose time duration to be 30-60 minutes; determines the
maximum rate required for what is expected to be worst case housekeeping DL
req; implies the use of serial bus to
dump SSR due to 1553 bus bandwidth limitations. Will likely require scheduled
ground pass with larger antenna in order to meet DL rate requirement |
GND, RF, C&DH |
2.6.5,
2.6.7 (Standard S/C Services- Comm & H/K Telem) |
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|
4.4.1.2.8 |
|
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|
|
The Downlink shall support a minimum downlink data rate of 1
Kbps (TBR) and a maximum downlink data rate of 500 Kbps (TBR) |
Bounds DL rate requirements.
Minimum rate driven by TDRSS emergency mode and network compatibility
testing. Maximum rate derived from worst case of 4.4.1.2.6 |
GND, RF, C&DH |
2.6.5,
2.6.7 (Standard S/C Services- Comm & H/K Telem) |
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4.4.1.2.9 |
|
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|
|
The downlink hardware shall provide the capability to latch and
record the Vc0 frame sync time with TBD accuracy when it is transmitted in
the downlink data stream and transmit this time to the ground |
Reflects the need for Vc0 frame sync time for Vc0 Range Data
Delay (RDD) spacecraft/ground time correlation method |
C&DH |
4.4.3.8
(Time), 5.2.5 (Tracking) |
|
|
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|
4.4.1.2.9 |
|
|
|
The Observatory transponder shall support turnaround ranging |
|
Required for orbit determination and prediction |
RF |
5.2.5
(Tracking) |
|
|
|
|
|
4.4.1.2.10 |
|
|
|
|
The Observatory transponder shall support simultaneous telemetry
and turnaround ranging on the RF downlink by utilizing a carrier and
subcarrier to provide separate channels for each. The transponder shall also have the
capability to disable the subcarrier and directly modulate the downlink
carrier with telemetry |
This requirement is driven by the need to allow long tracking
data arcs to be captured following GTO manuevers while still maintaining S/C
telemetry. Note that the req to
disable the subcarrier is designed to maximize data rates for recorder dumps
when ranging is not required. See
5.2.1.5 and 5.2.2.5 for ground system reqs |
C&DH, RF |
2.6.5,
2.6.7 (Standard S/C Services- Comm & H/K Telem), 5.2.5 (Tracking) |
|
|
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|
|
4.4.1.2.11 |
|
|
|
The Downlink shall have the capability to operate both in
coherent and non-coherent modes |
|
Non-coherent- DL freq based on XPNDR internal oscillator freq
(fixed); Coherent- DL freq based on uplink freq modulation (variable)- needed
for ranging |
RF |
5.2.5
(Tracking) |
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|
4.4.2 |
|
|
Ka-Band Communications |
A return link shall be provided to support Science telemetry
transfer to the ground station |
|
|
|
1.6
(Cadence), 2.5.3 (Data Capture & Compl), 2.6.5 (Standard S/C Services-
Comm) |
|
|
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|
|
4.4.2.1 |
|
|
|
|
The science telemetry shall support the CCSDS protocol
recommendations |
|
C&DH, GND |
2.6.5
(Standard S/C Services- Comm) |
|
|
|
|
|
4.4.2.2 |
|
|
|
|
The downlink rate shall support the continuous 150 Mbps downlink
of science data (300 Mbps symbol rate w/ Conv encoding) while on station in a
geosynchronous orbit |
Reflects science data rate of 130 Mbps plus overhead and
downlink encoding (130 + headers + RS + Conv = ~ 300 Mbps) |
C&DH, RF, GND |
1.6
(Cadence), 2.5.3 (Data Capture & Compl), 2.6.5 (Standard S/C Services-
Comm) |
|
|
|
|
|
4.4.2.3 |
|
|
|
|
The downlink bit error rate shall be less than 1x10-8 (including
downlink encoding benefits) |
Reflects the preliminary calculation of the DL portion of the
99.99% data science data completeness budget requirement; Reflects error rate
of data recovered after transmission and recovery at ground station |
C&DH, RF, GND |
1.6
(Cadence), 2.5.3 (Data Capture & Compl), 2.6.5 (Standard S/C Services-
Comm) |
|
|
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|
|
4.4.2.4 |
|
|
|
|
Ka-Band link calculations shall include the assumption of 99.25%
availability due to atmospheric and rain attenuation |
Reflects data capture budget assumption of only 66 hours lost
due to rain attenuation [1 - 66 hours/(24 hrs/day* 365 days/year)= 0.9925] |
C&DH, RF, GND |
1.6
(Cadence), 2.5.3 (Data Capture & Compl), 2.6.5 (Standard S/C Services-
Comm) |
|
|
|
|
|
4.4.2.5 |
|
|
|
|
The Ka-Band Science downlink shall have a link margin of ≥
3 dB |
Margin helps ensure a stable and reliable RF downlink
communications path |
RF, GND |
1.6
(Cadence), 2.5.3 (Data Capture & Compl), 2.6.5 (Standard S/C Services-
Comm) |
|
|
|
|
|
4.4.2.6 |
3/27/2003
- Delete |
|
|
|
The Observatory shall support Ka-Band antenna handovers with
a maximum downlink data gap of 2.5 minutes |
Driven by HMI data gap requirement in 1.2.4. Data capture budget provides for up to 2.5
minute handovers |
C&DH, RF, GND, APS |
1.6
(Cadence), 1.2.3 (Science Data Gaps), 2.5.3 (Data Capture & Compl), 2.6.5
(Standard S/C Services- Comm) |
|
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|
|
4.4.3 |
|
|
Data System Functions |
|
|
|
|
|
|
|
|
|
|
4.4.3.1 |
|
|
Architecture |
The Spacecraft shall employ the use of a distributed processing
architecture with a central spacecraft processor and subsystem common nodal
processing elements for commanding and telemetry collection using the same
core design and layout |
|
Requirement reflects project design decision to utilize common
distributed processor-controller power system design; offers advantages in
design commonality along with associated benefits in testability,
reliability, cost and manpower
reduction |
C&DH, GN&C, PWR, APS, FSW |
2.1.2.2
(Fault Tol), 2.6.6 (Standard S/C Services- Data processing), SDO Arch. Design |
|
|
|
|
|
4.4.3.2 |
|
|
Attitude Control Processing |
The Spacecraft shall support the attitude control system
processing requirements |
|
ACS Single Board Computer requirements encompassed in S/W SBC
reqs provided to C&DH |
C&DH, GN&C, FSW |
2.6.6
(Standard S/C Services- Data processing) |
|
|
|
|
|
4.4.3.2.1 |
|
|
|
|
The data system and 1553 bus shall provide sufficient capability
to support a 5 Hz closed loop control capability for Observatory attitude
control |
Requirement based on instrument pointing requirements |
C&DH, GN&C, FSW |
2.6.6
(Standard S/C Services- Data processing), 4.2.3 (Control & Stability) |
|
|
|
|
|
4.4.3.2.2 |
3/27/03 |
|
|
|
The maximum allowable delay in the closed
loop attitude control is one ACS closed loop control cycle as defined in
4.4.3.2.1. |
Derived from 5 Hz loop requirement and pointing
requirement. This value is typically
what is assumed for S/C controller stability & robustness |
C&DH, GN&C, FSW |
2.6.6
(Standard S/C Services- Data processing), 4.2.3 (Control & Stability) |
|
|
|
|
|
4.4.3.3 |
|
|
Memory Management |
The Observatory shall provide memory management functions |
|
|
|
|
|
|
|
|
|
4.4.3.3.1 |
|
|
|
All on-board processors shall provide the capability to load
code and data from the ground to on-board processors |
|
Requirement allows flexibility to adjust for on-orbit conditions |
C&DH, GN&C, PWR, INSTR, FSW |
2.6.6
(Standard S/C Services- Data processing) |
|
|
|
|
|
4.4.3.3.2 |
|
|
|
The Spacecraft central processor shall provide the capability to
load code and data from the ground to
into non-volatile memory |
|
Reflects the need to retain updated data and code despite resets
and power cycling |
C&DH, FSW |
2.6.6
(Standard S/C Services- Data processing) |
|
|
|
|
|
4.4.3.3.3 |
|
|
|
The Observatory shall provide the capability to dump onboard
processor memory to the ground station |
|
|
C&DH, FSW |
2.6.6
(Standard S/C Services- Data processing), 2.6.5, 2.6.7 (Standard S/C
Services- Comm & H/K Telem) |
|
|
|
|
|
4.4.3.3.4 |
|
|
|
All on-board processors shall have adequate memory protection
against memory errors |
|
Memory
errors may be induced by radiation effects or hardware failures (stuck
bits). S/C processor may elect to
utilize checksumming of static memory to detect, correct and report bit
flips, while other processors may utilize rad hard memory and memory
management |
C&DH, GN&C, PWR, INSTR, FSW |
2.1.2.2
(Fault Tol), 2.6.6 (Standard S/C Services- Data processing) |
|
|
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|
|
4.4.3.4 |
|
|
Commands |
|
|
|
|
|
|
|
|
|
|
4.4.3.4.1 |
|
|
Real Time Commands |
The Spacecraft shall provide a capability to receive, decode,
validate and distribute real time commands |
|
|
C&DH, RF, FSW |
2.6.5
(Standard S/C Services- Comm) |
|
|
|
|
|
4.4.3.4.1.1 |
|
|
|
|
The Observatory shall use the 1553 bus to distribute commands to
the relevant Observatory subsystems |
|
C&DH, RF, FSW |
2.6.5
(Standard S/C Services- Comm) |
|
|
|
|
|
4.4.3.4.1.2 |
|
|
|
|
The uplink command decode function shall receive unswitched
power from the power subsystem and shall not have the capability to be
powered off |
|
C&DH, RF |
2.6.5
(Standard S/C Services- Comm), 2.2.2 (Fault Tol) |
|
|
|
|
|
4.4.3.4.2 |
|
|
Hardware Decoded Commands |
The Spacecraft shall provide a capability to support hardware
decoded commands, which are received, decoded, and distributed by the
Observatory hardware without the intervention of flight software |
|
|
C&DH, RF |
2.6.5
(Standard S/C Services- Comm), 2.2.2 (Fault Tol) |
|
|
|
|
|
4.4.3.4.2.1 |
|
|
|
|
The uplink hardware shall have the capability to decode and
distribute a minimum of 16 (TBR) hardware decoded commands to Observatory
subsystems |
Required to reset, recover and reconfigure the Observatory in
the absence of processor or flight software operation. Specific # of cmds will be determined by
reconfiguration needs of the observatory.
Explore the possible use of limited H/W cmds and more SDN S/W cmds |
C&DH, RF |
2.6.5
(Standard S/C Services- Comm), 2.2.2 (Fault Tol) |
|
|
|
|
|
4.4.3.4.3 |
|
|
Stored Commands |
The Spacecraft shall provide a stored command capability to
receive, store, and later execute sequences of commands |
|
|
C&DH, RF, FSW |
2.6.5
(Standard S/C Services- Comm), 2.2.2 (Fault Tol) |
|
|
|
|
|
4.4.3.4.3.1 |
|
|
|
|
The Spacecraft shall provide at least two absolute time stored
command sequence buffers with a relative accuracy of 1 second command
execution |
Use of two or more buffers allows capability to load one buffer
whith utilizing the other |
C&DH, FSW |
2.6.5
(Standard S/C Services- Comm), 2.2.2 (Fault Tol) |
|
|
|
|
|
4.4.3.4.3.2 |
|
|
|
|
The Spacecraft shall provide relative time stored command
sequences with a relative accuracy of 1 second command execution |
|
C&DH, FSW |
2.6.5
(Standard S/C Services- Comm), 2.2.2 (Fault Tol) |
|
|
|
|
|
4.4.3.4.3.3 |
|
|
|
|
The
Spacecraft shall provide a capability to initiate relative time sequence
commands triggered on telemetry events |
Requirement supports using stored telemetry commands for health
and safety functions |
C&DH, FSW |
2.6.5
(Standard S/C Services- Comm), 2.2.2 (Fault Tol) |
|
|
|
|
|
|
|
|
|
|
|
|
|
|
|
|
|
|
|
4.4.3.5 |
|
|
Health and Safety |
|
The Spacecraft shall provide for Telemetry Statistic Monitor
(TSM) on-board telemetry monitoring points |
Allows the use of TSM/RTS capability for the autonomous
initiation of relative time command sequences upon detection of a
pre-specified set of conditions |
C&DH, FSW |
2.6.5
(Standard S/C Services- Comm), 2.2.2 (Fault Tol) |
|
|
|
|
|
|
|
|
|
|
|
|
|
|
|
|
|
|
|
4.4.3.6 |
|
|
Housekeeping Telemetry |
The Observatory shall provide the capability to collect and
downlink housekeeping telemetry through all phases of the mission |
|
|
C&DH, RF, FSW |
2.6.5,
2.6.7 (Standard S/C Services- Comm & H/K Telem) |
|
|
|
|
|
4.4.3.6.1 |
|
|
|
The data system shall provide error encoding to meet downlink
data error rate requirements |
|
|
C&DH, RF, FSW |
2.6.5,
2.6.7 (Standard S/C Services- Comm & H/K Telem), 2.2.2 (Fault Tol) |
|
|
|
|
|
4.4.3.6.2 |
|
|
Real time Telemetry |
The Spacecraft shall collect and transmit Observatory
housekeeping telemetry to the ground station in (near) real time. |
|
Corresponding Ground System requirement is 5.2.4.1 |
C&DH, RF, FSW |
2.1.5.5
(H/K Telem) |
|
|
|
|
|
4.4.3.6.3 |
|
|
Stored Telemetry |
The Spacecraft shall collect and store Observatory housekeeping
telemetry for later transmission to the ground station |
|
|
C&DH, FSW |
2.6.5,
2.6.7 (Standard S/C Services- Comm & H/K Telem) |
|
|
|
|
|
4.4.3.6.3.1 |
|
|
|
|
The Spacecraft shall have the capability to store and dump up to
24 hours of housekeeping telemetry in nominal mission mode |
Required for dump and evaluation of on-board anomalies where
loss of telemetry occurs; Recorder size restrictions may drive telemetry
sampling rate |
C&DH, FSW |
2.6.5,
2.6.7 (Standard S/C Services- Comm & H/K Telem) |
|
|
|
|
|
4.4.3.6.3.2 |
|
|
|
|
The spacecraft shall provide 90 Mbytes (TBR) spacecraft recorder
for storage of housekeeping telemetry through all phases of mission
operations |
Recorder size defined by ops concept and bounded by largest
storage requirements in either launch and IOC, GTO circularization, or 24
storage during Observatory on-station anomalies. Initial assumption of 8 Kb/sec rate for 24
hours, resulting in 86.4 Kbyte recorder |
C&DH, FSW |
2.6.5,
2.6.7 (Standard S/C Services- Comm & H/K Telem) |
|
|
|
|
|
4.4.3.6.3.3 |
|
|
|
|
The Spacecraft shall have the capability to record telemetry at
a sampling rate of up to 32kb/sec for shorter intervals where higher sampling
rates are required |
24 hour storage requirement does not apply to higher telemetry
sampling rate |
C&DH, FSW |
2.6.5,
2.6.7 (Standard S/C Services- Comm & H/K Telem) |
|
|
|
|
|
4.4.3.6.3.4 |
|
|
|
|
The data system recorder shall have a bit error rate of less
than TBD (1x10-9??) |
Drives need for recorder error scrubbing. Initial recorder error rate based on
roughly equivalent error rate (per bit) at science data |
C&DH, FSW |
2.6.5,
2.6.7 (Standard S/C Services- Comm & H/K Telem) |
|
|
|
|
|
4.4.3.6.3.5 |
|
|
|
The Spacecraft shall have the capability to playback stored
telemetry to the ground station upon command |
|
|
C&DH, FSW |
2.6.5,
2.6.7 (Standard S/C Services- Comm & H/K Telem) |
|
|
|
|
|
4.4.3.6.3.6 |
|
|
|
The Spacecraft shall support
the simultaneous downlink of both real time and stored housekeeping
data through all mission phases |
|
|
C&DH, RF, FSW |
2.6.5,
2.6.7 (Standard S/C Services- Comm & H/K Telem) |
|
|
|
|
|
4.4.3.6.3.7 |
|
|
|
|
The Spacecraft shall support the capability to collect and store
housekeeping telemetry while playing back stored telemetry to the ground
station |
Allows a running buffer to be kept of housekeeping telemetry in
the event of an anomaly during playback |
C&DH, FSW |
2.6.5,
2.6.7 (Standard S/C Services- Comm & H/K Telem) |
|
|
|
|
|
4.4.3.6.3.8 |
|
|
|
|
The Housekeeping telemetry collection/storage rate and content
shall be reconfigurable via ground command |
Implies the use of Filter Tables to set
collection/record/playback parameters through the various phases of mission
operations |
C&DH, FSW |
2.6.5,
2.6.7 (Standard S/C Services- Comm & H/K Telem) |
|
|
|
|
|
|
|
|
|
|
|
|
|
|
|
|
|
|
|
4.4.3.7 |
|
|
Science Data |
The Observatory shall route science data for downlink
transmission to the ground station |
|
The corresponding Ground System requirement is 5.2.5.1 |
INSTR, C&DH |
2.6.5
(Standard S/C Services- Comm) |
|
|
|
|
|
4.4.3.7.1 |
|
|
|
The
Observatory shall collect continuous fixed-length instrument science data
frames via the IEEE 1355 interface protocol and provide encoding for
continuous transmission to the science ground station |
|
Reflects science data DL requirements and lack of instrument
science data storage; Ongoing discussion on whether Instruments provide
complete VCDUs or simply fixed length frames |
INSTR, C&DH |
1.6
(Cadence), 2.6.5 (Standard S/C Services- Comm) |
|
|
|
|
|
4.4.3.7.2 |
3/27/03 |
|
|
|
The spacecraft onboard science data bit error rate shall be less than 1x10-10 and applies to the data path
from the instrument CCD output to the conversion to RF signal |
Reflects the spacecraft hardware portion of the 99.99% science data completeness budget
requirement. Applies primarily to HMI data completeness reqs- EVE and
SHARRP reqs Based on total data losses from Instr CCD through DL interface. Worst case assumption is
that single bit flip results in packet frame loss- can support up to TBD
packet frame losses per day and meet data completeness req. Implies use of significant rad hard
components to meet this req. |
INSTR, C&DH |
1.2.2
(Data Compl) |
|
|
|
|
|
4.4.3.7.3 |
|
|
|
The science data downlink shall provide forward error correction
encoding to meet downlink data error rate requirements |
|
|
C&DH |
2.6.5
(Standard S/C Services- Comm), 1.2.2 (Data Compl) |
|
|
|
|
|
4.4.3.7.4 |
|
|
|
|
The maximum science observation data transferred to the S/C for
encoding and downlink to the ground station shall not exceed an aggregate
total of 130 Mbps (150 Mbps downlink once error encoding and margin added) |
Places upper limit on Science data rate based on Instrument
collection and data rate allocations.
130 Mbps includes science data packet headers |
C&DH |
1.6
(Cadence), 2.6.5 (Standard S/C Services- Comm) |
|
|
|
|
|
4.4.3.7.5 |
|
|
|
The Science data downlink function shall have the flow-control
capability of limiting the data from each instrument input to its allocated
data rate |
|
Capability prevents data from one instrument from inadvertently
flooding the downlink data stream and interfering with the data downlink from
other instruments |
C&DH |
2.6.5
(Standard S/C Services- Comm), 1.2.2 (Data Compl) |
|
|
|
|
|
4.4.3.7.6 |
|
|
|
The science data flow control capability shall be reconfigurable
in flight to allow reallocation of
instrument science data telemetry bandwidth |
|
Allows
reconfiguration for optimal use of B/W |
C&DH |
2.6.5
(Standard S/C Services- Comm), 2.2.2 (Fault Tol) |
|
|
|
|
|
4.4.3.7.7 |
|
|
|
The science data downlink function shall assure that the data
from each instrument source is uniquely identified |
|
Implies that the DL function inserts virtual channel frame
headers or verifies that the virtual channel ID in each fixed frame is
correct. This allows correct science
data routing on the ground via science frame headers |
C&DH |
2.5.4
(Data Delivery) |
|
|
|
|
|
|
|
|
|
|
|
|
|
|
|
|
|
|
|
4.4.3.8 |
|
|
Time Implementation |
The Observatory shall maintain time to the accuracy required by
all aspects of the mission |
|
|
|
2.6.6
(Timing) |
|
|
|
|
|
4.4.3.8.1 |
|
|
Relative Time Accuracy |
|
The Spacecraft shall maintain time and make it available to
components over the 1553 bus with a relative accuracy of 10 ms (TBR) over a
period of 1 sec |
Reflects preliminary implementation decision to use 1553 bus as
time distribution approach- need to investigate relative time accuracy that
can be achieved using this approach and verify relative accuracy |
C&DH, FSW |
2.6.6
(Timing), 4.2 (Attitude Cntrl & Det), 4.4.3.4.3 (Stored Cmds) |
|
|
|
|
|
4.4.3.8.2 |
|
|
Absolute Time Accuracy |
The mission shall support the ability to adjust and maintain the
Observatory time clock to compensate for onboard time drift and stability
effects |
|
|
C&DH, FSW, GND |
2.6.6
(Timing), 4.2 (Attitude Cntrl & Det), 4.4.3.4.3 (Stored Cmds) |
|
|
|
|
|
4.4.3.8.2.1 |
|
|
|
|
The mission shall support the ability to maintain Observatory
time to ground time to within 100 msec (TBR) |
Based on HMI requirement of 100 msec (with goal of 10 msec)
which appears to be more stringent than (and therefore encapsulate) the other
instrument reqs. Planning on using Vc0 time correlation approach. In order to meet this req, need to measure
to ~ 10 ms and use to predict and adjust time drift. |
C&DH, FSW, GND |
3.2.4
(HMI timing), 2.6.6 (Timing) |
|
|
|
|
|
4.4.3.8.2.2 |
|
|
|
The Spacecraft shall provide a "smooth" time adjust
capability to allow instrument time to be updated without time jumps or
discontinuities |
|
|
C&DH, FSW, INSTR, GND |
3.2.4
(HMI timing), 2.6.6 (Timing) |
|
|
|
|
|
4.4.3.8.2.3 |
|
|
|
|
During science operations, the maximum time adjust shall be no
greater than 100 µsec (TBR) over a period of 1 sec (TBR) |
|
C&DH, FSW, INSTR, GND |
3.2.4
(HMI timing), 2.6.6 (Timing) |
|
|
|
|
|
|
|
|
|
|
|
|
|
|
|
|
|
|
|
4.5 |
|
|
High Gain Antenna (HGA) Assembly |
The Spacecraft shall provide a HGA function that meets the
science downlink rate and reliability requirements of the mission |
|
|
|
1.6 (Cadence), 2.5.1 (Cont. Contact), 2.5.3 (Data Capture &
Compl) |
|
|
|
|
|
4.5.1 |
|
|
Operation, Pointing and Stability |
|
|
|
|
|
|
|
|
|
|
4.5.1.1 |
3/27/03 |
|
|
The HGAs implementation shall provide sufficient FOV coverage with 2 high gain antennas in the
nominal mode of operation such that Observatory attitude adjustments are not
needed for full antenna ground station coverage |
|
|
APS, RF, DEPLOY |
1.2
(Data Capture & Compl), 2.1.5.3 (Data Comp Budget), 2.2.2 (Fault Tol) |
|
|
|
|
|
4.5.1.2 |
3/27/03 |
|
|
The HGAs implementation shall provide sufficient degraded
mode coverage such that minimum mission science data requirements are met
with the use of a single HGA |
|
Need further information on actual science minimum mission
requirements and operation constraints; anticipate that single HGA will not
meet 95% completeness req; Single HGA
option may include the use of S/C roll manuevers to optimize coverage |
APS, RF, DEPLOY, GN&C |
Lev.1
[Science Obj 2.1], 1.2 (Data Capture & Compl), 2.1.5.3 (Data Comp Budget) |
|
|
|
|
|
4.5.1.3 |
|
|
|
|
The HGA implementation shall provide a pointing accuracy of +/-
0.25 deg to the ground station |
This requirement may actually need to be broken out into three
separate reqs: ACS knowledge, deployment repeatability/knowledge, and HGA
gimbal pointing accuracy. The 0.25 deg
offset angle corresponds to a 0.25 dB gain reduction for 1/2 meter HGA
dish. Budget does not include DC error
components that can be statically removed |
APS, RF, DEPLOY, GN&C |
1.2
(Data Capture & Compl), 2.1.5.3 (Data Comp Budget) |
|
|
|
|
|
4.5.1.4 |
|
|
|
|
The contribution of HGA operation to nominal Observatory
attitude control disturbance effects shall not exceed 0.5 arcsec (TBR) |
Initial HGA ROM indicated < 0.1 arcsec contribution per
gimbal- further investigation needed |
APS |
4.2.2
(Pointing Acc), 4.2.3 (Attitude Cntrl & Stability) |
|
|
|
|
|
4.5.1.5 |
|
|
|
|
The HGA implementation shall be able to calibrate out antenna
deployment errors of up to +/- 1 deg (TBR) |
|
GND, DEPLOY, APS, RF |
4.2.2
(Pointing Acc), 4.2.3 (Attitude Cntrl & Stability) |
|
|
|
|
|
|
|
|
|
|
|
|
|
|
|
|
|
|
|
4.6 |
|
|
Deployment Actuation and Verification |
The Spacecraft shall support the capability to initiate and
monitor solar array and High Gain Assembly (HGA) deployment functions |
N/A |
N/A |
|
|
|
|
|
|
|
4.6.1 |
|
|
Solar Array Deployment and Verification |
|
The Spacecraft shall provide a nominal deployment capability
based of receipt of 2/3 of the
LV/Observatory separation signals |
Baseline deployment option utilizes processor/software control
to detect separation and initiate deployment and confirmation |
PWR, DEPLOY, FSW |
2.6.3 (Standard S/C Services- Power),
2.4.3.3 (Sep Sigs) |
|
|
|
|
|
4.6.1.1 |
|
|
|
|
The Spacecraft shall provide an independent hardware capability
to autonomously deploy the solar arrays upon positive receipt of 3/3
separation signals between the LV/Observatory |
Reflects the need for independent deployment of solar arrays in
the event of processor or Software anomaly |
PWR, DEPLOY |
2.6.3 (Standard S/C Services- Power),
2.4.3.3 (Sep Sigs), 2.2.2 (Fault Tol) |
|
|
|
|
|
4.6.1.2 |
|
|
|
The
Spacecraft shall provide a ground-based deployment command initiation
capability |
|
Allows independent deploy command capability independent of
separation signal control |
PWR |
2.6.3 (Standard S/C Services- Power),
2.4.3.3 (Sep Sigs), 2.2.2 (Fault Tol) |
|
|
|
|
|
4.6.1.3 |
|
|
|
The Spacecraft shall provide sufficient deployment
blocks/interrupts to prevent inadvertent deployment initiation during ground
testing |
|
Reflects ground test safety needs to prevent inadvertent
deployment |
PWR, GSE |
2.6.3 (Standard S/C Services- Power),
2.4.3.3 (Sep Sigs), 2.2.2 (Fault Tol), 2.7.7 (Safety) |
|
|
|
|
|
4.6.1.4 |
|
|
|
The Spacecraft shall implement a redundant deployment function
such that no reconfiguration of the Observatory is required to initiate
deployment in the event of a failed deployment attempt |
|
Reflects the design need to place back-up deployment functions
within subsystem components powered and command-accessible as part of the
nominal launch configuration; eliminates the need to first reconfigure
Observatory in the event of an unsuccessful deployment attempt, further
risking Observatory survival. Also
implies need for redundant deployment mechanism |
PWR, DEPLOY |
2.6.3 (Standard S/C Services- Power),
2.4.3.3 (Sep Sigs), 2.2.2 (Fault Tol) |
|
|
|
|
|
4.6.1.5 |
|
|
|
The Spacecraft shall provide the capability to detect and
monitor solar array deployment as part of the deployment verification process |
|
|
PWR |
2.5.7
(Critical Telem monitoring)) |
|
|
|
|
|
4.6.1.6 |
|
|
|
The Spacecraft shall block the power-on and initiation of
reaction wheel operation until on-board telemetry confirms the positive
deployment of Observatory solar arrays; this capability will have an override
feature to allow ground override of this blocking approach |
|
Reflects dual implementation needs; the need to prevent initial
safehold acquisition until positive separation to prevent RW damage due to
attached third stage; the need for an emergency RW safing capability for
ground testing |
PWR, GN&C, FSW |
2.6.3
(Standard S/C Services- Power), 2.7.7 (Safety) |
|
|
|
|
|
|
|
|
|
|
|
|
|
|
|
|
|
|
|
4.6.2 |
|
|
High Gain Antenna (HGA) Deployment and Verification |
N/A |
N/A |
N/A |
|
|
|
|
|
|
|
4.6.2.1 |
|
|
|
The Spacecraft shall provide the capability for commanded
deployment of the HGA |
|
HGA deployment is not required immediately after separation and
therefore is not an autonomous function |
APS, DEPLOY, FSW |
1.6
(Cadence), 2.5.1 (Continuous Contact) |
|
|
|
|
|
4.6.2.2 |
|
|
|
The Spacecraft shall implement a redundant commanded HGA
deployment function such that no reconfiguration of the Spacecraft is
required to initiate deployment in the event of a failed deployment attempt |
|
Reflects the design need to place back-up deployment functions
within subsystem components powered and command-accessible as part of the
nominal deployment configuration; eliminates the need to first reconfigure
Observatory in the event of an unsuccessful deployment attempt |
APS, DEPLOY, FSW |
1.6
(Cadence), 2.5.1 (Continuous Contact), 2.1.2.2 (Fault Tol) |
|
|
|
|
|
4.6.2.3 |
|
|
|
The Spacecraft shall provide the capability to detect and
monitor HGA deployment as part of the deployment verification process |
|
|
APS, DEPLOY, FSW |
2.5.7
(Critical Telem monitoring)) |
|
|
|
|
|
4.6.2.4 |
|
|
|
|
The
Spacecraft shall provide the capability to deploy the HGA and verify HGA
deployment angle to an accuracy of TBD (1 deg??) |
Assume that HGA deploy angles can be calibrated out in orbit
(see 4.5.1.5) |
APS, DEPLOY |
1.6
(Cadence), 2.5.1 (Continuous Contact),2.5.7 (Critical Telem monitoring)) |
|
|
|
|
|
|
|
|
|
|
|
|
|
|
|
|
|
|
|
5 |
|
|
Ground Segment Requirements |
|
|
|
|
|
|
|
|
|
|
5.1 |
|
|
Integration and Test (I&T) |
N/A |
N/A |
N/A |
|
|
|
|
|
|
|
5.1.1 |
|
|
High Rate Science GSE |
The high rate science GSE shall have the capability to support
testing and evaluation of the Observatory operation and instrument telemetry
as part of ground testing |
|
|
GND, GSE, C&DH |
2.7.2
(Verif) |
|
|
|
|
|
5.1.1.1 |
|
|
|
|
The ground system high rate GSE shall capture and decode science
data telemetry in real time |
|
GND, GSE, C&DH |
2.7.2
(Verif) |
|
|
|
|
|
5.1.1.2 |
|
|
|
|
The ground system high rate GSE shall deliver decoded science
data to instrument test equipment in real time |
|
GND, GSE, C&DH |
2.7.2
(Verif) |
|
|
|
|
|
5.1.1.3 |
|
|
|
|
The ground system high rate GSE shall have the capability to
record TBD hours of encoded and/or unencoded high rate data |
Provides capability to
record and playback science data from local ground station in addition
to instrument SOCs |
GND, GSE, C&DH |
2.7.2
(Verif) |
|
|
|
|
|
|
|
|
|
|
|
|
|
|
|
|
|
|
|
5.1.2 |
|
|
Low Rate GSE |
The Observatory GSE shall have the capability to decode,
deliver, and display Observatory housekeeping telemetry in real time through
all phases of ground testing and mission operation |
|
|
GND, GSE, C&DH |
2.6.7
(Standard S/C Services), 2.7.2 (Verif) |
|
|
|
|
|
5.1.2.1 |
|
|
|
The observatory GSE shall have the capability to format and send
commands to the Observatory through all phases of ground testing and mission
operation |
|
|
GND, GSE, C&DH |
2.6.7
(Standard S/C Services), 2.7.2 (Verif) |
|
|
|
|
|
|
|
|
|
|
|
|
|
|
|
|
|
|
|
5.2 |
|
|
Ground Station Implementation |
The Ground Station implementation and operation shall support
SDO mission implementation and operation requirements through all phases of
mission life |
N/A |
N/A |
GND |
1.2
(Data Capture & Compl), 1.6 (Cadence), 2.5 (Ops Concept) |
|
|
|
|
|
5.2.1 |
|
|
Dedicated Site Requirements |
The ground station sites shall meet the housekeeping and science
data capture and loss budget requirements |
|
Reflects need to locate site in area that meets atmospheric
attenuation and FOV coverage requirements defined in the capture budget to
meet the 95% science data requirement |
GND |
1.2
(Data Capture & Compl), 1.6 (Cadence), 2.5 (Ops Concept) |
|
|
|
|
|
5.2.1.1 |
|
|
|
The ground segment shall provide sufficient redundancy to meet
the ground station data capture budget requirements |
|
Addressed by implementation decision to use dual antenna and
ground sites to meet data capture budget requirements |
GND |
1.2
(Data Capture & Compl), 1.6 (Cadence), 2.5 (Ops Concept), 2.2.2 (Fault
Tol) |
|
|
|
|
|
5.2.1.2 |
|
|
|
Each ground station shall provide S-Band frequency command,
telemetry and tracking functions in
support of SDO mission operations |
|
|
GND |
2.5
(Ops Concept) |
|
|
|
|
|
5.2.1.3 |
|
|
|
Each ground station shall provide local Ka-Band frequency
telemetry storage in support of SDO science operations |
|
Required to handle temporary line outages between antennas and
DDC |
GND |
1.2
(Data Capture & Compl), 1.6 (Cadence), 2.5 (Ops Concept) |
|
|
|
|
|
5.2.1.4 |
|
|
|
Ground station functions shall be capable of remote reconfiguration from the
MOC |
|
Allows MOC to modify Ground station as needed to support mission
operations |
GND |
2.5
(Ops Concept) |
|
|
|
|
|
5.2.1.5 |
|
|
|
|
The Ground station shall be able to generate range tones on the
uplink and receive simultaneous S-Band telemetry and ranging data modulated
on an RF downlink subcarrier and carrier respectively. The station must be able to receive
telemetry on either the carrier or subcarrier |
Required for tracking capability. See 4.4.1.2.10 for S/C req |
GND |
2.5
(Ops Concept) |
|
|
|
|
|
5.2.1.6 |
|
|
|
End-to-end compatibility testing shall be conducted to verify
the compatibility of all ground stations and any RF networks utilized with
the Observatory and launch vehicle |
|
|
GND |
2.5 (Ops Concept), 2.5.7 (Critical
Telemetry), 2.7.2 (Verification) |
|
|
|
|
|
5.2.2 |
|
|
Ancillary Site Requirements |
The mission shall demonstrate verified compatibility with other
NASA and/or commercial ground stations required during launch, early mission
phases, and contingency support |
N/A |
N/A |
GND |
2.5 (Ops Concept), 2.5.7 (Critical
Telemetry) |
|
|
|
|
|
5.2.2.1 |
|
|
|
|
All ancillary ground sites shall only be required to support the
Observatory at the S-Band frequency and functions (command, telemetry,
tracking for orbit determination) |
|
GND |
2.5 (Ops Concept), 2.5.7 (Critical
Telemetry) |
|
|
|
|
|
5.2.2.2 |
|
|
|
|
All commands routed through the ancillary S-band ground sites
shall originate from the SDO MOC for uplink to the Observatory |
Requirement does not preclude SOC commanding, but requires
commands to pass through MOC for checking and forwarding to the Observatory |
GND |
2.5 (Ops Concept), 2.5.7 (Critical
Telemetry) |
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5.2.2.3 |
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All Observatory housekeeping telemetry collected by the
ancillary S-band ground sites shall be forwarded in near real-time to the SDO
MOC for archive and distribution to the respective SOCs |
|
GND |
2.5 (Ops Concept), 2.5.7 (Critical
Telemetry) |
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5.2.2.4 |
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End-to-end compatibility testing shall be conducted to verify
the compatibility of all ground stations and any RF networks utilized with
the Observatory and launch vehicle |
|
GND |
2.5 (Ops Concept), 2.5.7 (Critical
Telemetry), 2.7.2 (Verification) |
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5.2.2.5 |
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The Ground station shall be able to generate range tones on the
uplink and receive simultaneous S-Band telemetry and ranging data modulated
on an RF downlink subcarrier and carrier respectively. The station must be able to receive
telemetry on either the carrier or subcarrier |
Required for tracking capability. See 4.4.1.2.10 for S/C req |
GND |
2.5
(Ops Concept) |
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5.2.3 |
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Commanding |
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5.2.3.1 |
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The ground station shall support a maximum S-Band command uplink
data rate of 2 Kbps |
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GND |
2.5 (Ops Concept), 2.6.5 (Standard S/C
Services- Comm) |
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5.2.4 |
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Housekeeping Telemetry |
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N/A |
N/A |
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5.2.4.1 |
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The ground station shall collect observatory housekeeping data
in the S-Band frequency and distribute in real-time to the MOC |
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GND |
2.5 (Ops Concept), 2.6.5 (Standard S/C
Services- Comm) |
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5.2.4.2 |
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The ground station shall support a maximum housekeeping downlink
data rate of TBD Kbps |
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GND |
2.5 (Ops Concept), 2.6.5 (Standard S/C
Services- Comm) |
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5.2.4.3 |
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The ground station shall have the capability to latch and record
the Vc0 frame sync time when it is received at the ground station with TBD (10 msec??) accuracy |
Used for Vc0 time correlation method. 10 msec accuracy required to maintain S/C
to ground time to 100 msec (4.4.3.8.2.1) |
GND |
2.5 (Ops Concept), 2.6.5 (Standard S/C
Services- Comm) |
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5.2.5 |
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Tracking |
Sufficient orbit tracking capabilities shall be provided for the
mission |
N/A |
N/A |
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5.2.5.1 |
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Definitive orbit position determination to 120 meters (TBR)
accuracy shall be achieved throughout the geosynchronous orbital phase |
Required for stationkeeping maneuvers and for HMI orbital
Doppler prediction data products |
FLT DYN, GND |
2.1
(Orbit), 1.5.1 (HMI Dopplergram precision) |
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5.2.5.2 |
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Definitive orbit velocity determination to < 1cm/sec accuracy
shall be achieved throughout the geosynchronous orbital phase |
Required for stationkeeping maneuvers and for HMI orbital
Doppler prediction data products |
FLT DYN, GND |
2.1
(Orbit), 1.5.1 (HMI Dopplergram precision) |
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5.2.5.3 |
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Orbit position prediction to 120 m (TBR) accuracy for a period
of a week shall be achieved throughout the geosynchronous orbital phase |
Required for stationkeeping maneuvers and for HMI orbital
Doppler prediction data products |
FLT DYN, GND |
2.1
(Orbit), 1.5.1 (HMI Dopplergram precision) |
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5.2.5.4 |
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Orbit velocity prediction to < 1 cm/sec (TBR) accuracy for a
period of a week shall be achieved throughout the geosynchronous orbital
phase |
Required for stationkeeping maneuvers and for HMI orbital
Doppler prediction data products |
FLT DYN, GND |
2.1
(Orbit), 1.5.1 (HMI Dopplergram precision) |
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5.2.5.5 |
03/27/03
- New |
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Provide acquisition data to all ground networks supporting
post-separation and transfer orbit ops. |
|
FLT DYN, GND |
2.5
(Ops Concept) |
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5.2.6 |
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Science Telemetry |
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5.2.6.1 |
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The ground station shall collect Instrument science data in the
Ka-Band frequency and directly distribute to the respective SOCs |
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5.2.6.2 |
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The ground station shall support the (near) continuous
collection of an aggregate science data downlink of 150 Mbps + margin |
Based on 130 Mbps science data prior to encoding and margin |
GND, INSTR |
2.5.4
(Data Delivery) |
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5.2.6.3 |
3/27/03 |
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The ground station shall be capable of autonomously relocking to
the downlink in 30 (TBR)
seconds after a dropout |
Allows rapid reacquisition and accommodation of an onboard
Observatory antenna handover |
GND, INSTR |
2.5.4
(Data Delivery) |
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5.2.6.4 |
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The ground station shall employ and demonstrate a data
distribution implementation with sufficient margin to support science data
retransmissions |
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GND, INSTR |
2.5.4
(Data Delivery) |
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5.2.6.5 |
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The ground station shall employ and demonstrate a data
distribution implementation with sufficient reliability to achieve error-free
data distribution including science data retransmissions |
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GND, INSTR |
2.5.4
(Data Delivery) |
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5.2.6.6 |
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The ground station shall directly distribute instrument science
data to the respective SOCs with a delivery latency of 3 min (TBR) |
Bases on initial Ground system assessment of 1 minute to capture
telemetry after transmission from Observatory and 2 minutes to process and
send to SOCs |
GND, INSTR |
2.5.4
(Data Delivery) |
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5.2.6.7 |
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The ground station shall provide 30 days of temporary data
storage to allow science data retransmission if required |
Reflects AO requirement |
GND, INSTR |
2.5.4
(Data Delivery) |
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5.2.7.8 |
3/27/03 |
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The ground station shall provide science data retransmission
from the temporary science data archive with a latency of TBD after requested by the SOCs |
Implies redundancy in data preservation at ground station or DDC
to prevent data loss |
GND, INSTR |
2.5.4
(Data Delivery) |
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5.2.7.9 |
03/27/03
- New |
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To ensure retransmission of desired data prior to deletion
from ground station data storage, the SOC shall request data retransmission
within TBD of initial ground station receipt of science data. |
|
GND, INSTR |
2.5.4
(Data Delivery) |
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5.3 |
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Mission Operations Center (MOC) |
The
MOC shall provide all ground data system interfaces required to support
Observatory I&T, launch, IOC, and in-orbit operations |
N/A |
N/A |
|
2.5
(Ops Concept) |
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5.3.1 |
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Command and Telemetry Functions |
The
MOC shall send all commands to the Observatory via the ground stations |
|
Requirement does not preclude SOC commanding, but requires
commands to pass through MOC for checking and forwarding to the Observatory |
GND |
2.5
(Ops Concept) |
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5.3.1.1 |
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The
MOC shall support the (near) continuous receipt of all Observatory
housekeeping data via S-Band telemetry downlink |
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GND |
2.5
(Ops Concept) |
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5.3.1.2 |
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The
MOC shall provide both real-time and long-term Observatory health and safety
ground monitoring functions |
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GND |
2.5
(Ops Concept) |
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5.3.1.3 |
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The
MOC shall archive and maintain Observatory housekeeping data over the life of
the mission |
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GND |
2.5
(Ops Concept) |
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5.3.1.4 |
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The
MOC shall forward the instrument housekeeping telemetry to the respective
SOCs with a maximum nominal latency of 1 min (TBR) |
|
GND |
2.5
(Ops Concept) |
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5.3.1.5 |
|
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The
MOC shall monitor dedicated ground station operations including housekeeping
and science data capture, housekeeping and science data distribution |
|
GND |
2.5
(Ops Concept) |
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5.3.1.6 |
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The
MOC shall provide real-time autonomous health and safety telemetry monitoring
and notification |
Allows
fault detection and unattended operations capability |
GND |
2.5
(Ops Concept), 2.2.2 (Fault Tol) |
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5.3.2 |
|
|
Data Products |
The
MOC shall provide the necessary data products required for the operation and
maintenance of the Observatory |
N/A |
N/A |
GND |
2.5
(Ops Concept) |
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5.3.2.1 |
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The
MOC shall provide housekeeping trending capabilities for ongoing monitoring
and evaluation of Observatory housekeeping data |
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|
GND |
2.5
(Ops Concept) |
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5.3.2.2 |
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|
The
MOC shall receive science operations plans from the SOCs and generate and
uplink science operations plans to the Observatory |
|
|
GND |
2.5
(Ops Concept) |
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5.3.2.3 |
|
|
|
The
MOC shall provide flight dynamics capabilities and products supporting
attitude determination, orbit prediction and determination, acquisition data
generation and delivery, and maneuver planning and execution |
|
|
GND |
2.5
(Ops Concept) |
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5.3.2.4 |
|
|
|
The
MOC shall provide ongoing flight software maintenance support to the SDO
spacecraft bus |
|
|
GND |
2.5
(Ops Concept) |
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|
5.3.2.5 |
|
|
|
The
MOC shall have the capability to maintain time correlation between the
Observatory and the ground segment |
|
|
GND |
2.5
(Ops Concept) |
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|
5.3.2.5.1 |
|
|
|
|
Time
correlation between the Observatory and the ground system shall be maintained
to 100 msec accuracy |
|
GND |
2.5
(Ops Concept) |
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5.4 |
|
|
Science Operations Center (SOC) |
Each
SOC shall provide and coordinate all instrument science data collection,
archiving, analysis, and science product generation/distribution capability
for the Instrument science team |
N/A |
N/A |
INSTR |
Lev.1
[Science Meas. 1-6, 2.1 (Science Obj)] |
|
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|
5.4.1 |
|
|
Ops |
The
SOC shall provide the capability for science mission planning and command
generation |
|
See
5.3.1 |
INSTR |
Lev.1
[Science Meas. 1-6, 2.1 (Science Obj)] |
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5.4.2 |
|
|
Archive |
The
SOC shall provide a long term archive of all instrument science and science
data products |
N/A |
Need
clarification from science teams on definition of "long term" - can not be longer than duration of mission.
Should be transferred to NASA specified archive at end of contract. |
INSTR |
Lev.1
[Science Meas. 1-6, 2.1 (Science Obj)] |
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|
|
Science
data products shall be archived consistently with LWS data storage, access,
and search requirements |
|
|
INSTR |
Lev.1
[Science Meas. 1-6, 2.1 (Science Obj)] |
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5.4.3 |
|
|
Data Products |
The
SOC shall formulate all primary science data products and provide a means for
their distribution to the science community |
N/A |
Comments:
"Trace to the SDO Program Data Management Plan (PDMP)." |
INSTR |
Lev.1
[Science Meas. 1-6, 2.1 (Science Obj)] |
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