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DRAFT

**Landsat Data Continuity Mission**

**Observatory Interface Requirements Document (Obs-IRD)**

**November 14, 2006**

**Revision - Draft**

![](media/image2.wmf){width="1.625in" height="1.375in"}

**CM FOREWORD**

This document is a Landsat Data Continuity Mission (LDCM) Project
Configuration Management (CM)-controlled document. Changes to this
document require prior approval of the applicable Configuration Control
Board (CCB) Chairperson or designee. Proposed changes shall be submitted
to the LDCM CM Office (CMO), along with supportive material justifying
the proposed change. Changes to this document will be made by complete
revision.

Questions or comments concerning this document should be addressed to:

LDCM Configuration Management Office

Mail Stop 427

Goddard Space Flight Center

Greenbelt, Maryland 20771

# Signature Page {#signature-page .Style-Heading-1}

+--------------------------+-------+---------------------------+-------+
| Prepared by:             | Date  |                           |       |
|                          |       |                           |       |
| Evan Webb                |       |                           |       |
|                          |       |                           |       |
| LDCM Mission Systems     |       |                           |       |
| Manager                  |       |                           |       |
|                          |       |                           |       |
| NASA-GSFC Code 599/427   |       |                           |       |
+--------------------------+-------+---------------------------+-------+
| Reviewed by:             | Date  | Reviewed by:              | Date  |
|                          |       |                           |       |
| Jeanine Morris-Murphy    |       |                           |       |
|                          |       |                           |       |
| LDCM OLI Manager         |       |                           |       |
|                          |       |                           |       |
| NASA GSFC Code 427       |       |                           |       |
+--------------------------+-------+---------------------------+-------+
| Reviewed by:             | Date  |                           |       |
|                          |       |                           |       |
| LDCM Deputy Project      |       |                           |       |
| Manager                  |       |                           |       |
|                          |       |                           |       |
| Del Jenstrom             |       |                           |       |
|                          |       |                           |       |
| NASA GSFC Code 427       |       |                           |       |
+--------------------------+-------+---------------------------+-------+
| Reviewed by:             | Date  |                           |       |
|                          |       |                           |       |
| LDCM Observatory Manager |       |                           |       |
|                          |       |                           |       |
| William Anselm           |       |                           |       |
|                          |       |                           |       |
| NASA GSFC Code 427       |       |                           |       |
+--------------------------+-------+---------------------------+-------+
| Approved by:             | Date  |                           |       |
|                          |       |                           |       |
| William Ochs             |       |                           |       |
|                          |       |                           |       |
| LDCM Project Manager     |       |                           |       |
|                          |       |                           |       |
| NASA GSFC Code 427       |       |                           |       |
+--------------------------+-------+---------------------------+-------+

#  Document Revision History {#document-revision-history .Style-Heading-1}

This document is controlled by the LDCM Project Management. Changes
require approval of the LDCM Project Manager, LDCM OLI Manager, and the
LDCM Mission Assurance Manager. Proposed changes shall be submitted to
LDCM Systems Engineering Manager.

  ------------- ---------- -------- ----------------------------------------
  **RELEASE**   **DATE**   **BY**   **DESCRIPTION**

  Draft         Nov.14,             Initial Version
                2006                
  ------------- ---------- -------- ----------------------------------------

# List of TBDs/TBCs/TBRs {#list-of-tbdstbcstbrs .Style-Heading-1}

This document contains information that is complete as possible. Items
that are not yet defined are annotated with TBD (To Be Determined).
Where final numerical values or data are not available, best estimates
are given and annotated TBC (To Be Confirmed). If there is an
inconsistency between two requirements then the best estimate is given
and annotated with a TBR (To Be Resolved). The following table
summarizes the TBD/TBC/TBR items in the document and supplements the
revision history.

  ---------- --------------- ------------------------------------------------
  **ITEM**   **REFERENCE**   **DESCRIPTION**

                             Data not supplied

                             

                             

                             
  ---------- --------------- ------------------------------------------------

Contents

[Signature Page i](#signature-page)

[Document Revision History ii](#document-revision-history)

[List of TBDs/TBCs/TBRs ii](#list-of-tbdstbcstbrs)

[1 SCOPE 8](#scope)

[1.1 IDENTIFICATION 8](#identification)

[1.2 Applicability 8](#applicability)

[1.3 TERMINOLOGY 8](#terminology)

[1.4 DOCUMENT OVERVIEW 9](#document-overview)

[1.5 CONFLICTS 11](#conflicts)

[1.6 REQUIREMENT WEIGHTING FACTORS 11](#requirement-weighting-factors)

[2 APPLICABLE DOCUMENTS 12](#applicable-documents)

[2.1 GOVERNMENT REQUIREMENT DOCUMENTS
12](#government-requirement-documents)

[2.2 NON-GOVERNMENT REQUIREMENT DOCUMENTS
12](#non-government-requirement-documents)

[2.3 REFERENCE DOCUMENTS 12](#reference-documents)

[3 Observatory Interface REQUIREMENTS
13](#observatory-interface-requirements)

[3.1 General 13](#general)

[3.2 Spacecraft to Ground Interface Requirements
13](#spacecraft-to-ground-interface-requirements)

[3.2.1 S-band to Ground Interface 13](#s-band-to-ground-interface)

[3.2.1.1 LANDSAT Ground Network (LGN) 13](#landsat-ground-network-lgn)

[3.2.1.1.1 Transmit (downlink): 13](#transmit-downlink)

[3.2.1.1.2 Receive (uplink) 13](#receive-uplink)

[3.2.1.2 Space Network (SN) 13](#space-network-sn)

[3.2.1.2.1 Transmit (downlink): 13](#transmit-downlink-1)

[3.2.1.2.2 Receive (uplink) 14](#receive-uplink-1)

[3.2.1.3 NASA Ground Network (GN) 14](#nasa-ground-network-gn)

[3.2.1.3.1 Transmit (downlink): 14](#transmit-downlink-2)

[3.2.1.3.2 Receive (uplink) 14](#receive-uplink-2)

[3.2.2 X-band to Ground Interface 15](#x-band-to-ground-interface)

[3.2.2.1 Landsat Ground Network (LGN) 15](#landsat-ground-network-lgn-1)

[3.2.2.2 International Cooperators (ICs)
15](#international-cooperators-ics)

[3.3 Spacecraft to Instrument Interface Requirements
16](#spacecraft-to-instrument-interface-requirements)

[3.3.1 Mechanical Interface Requirements
16](#mechanical-interface-requirements)

[3.3.1.1 Dimensions 17](#dimensions)

[3.3.1.2 Reference Frames 17](#reference-frames)

[3.3.1.2.1 Observatory Reference Frame 17](#observatory-reference-frame)

[3.3.1.2.2 Spacecraft Body Frame 17](#spacecraft-body-frame)

[3.3.1.3 Instrument Fields of View 18](#instrument-fields-of-view)

[3.3.1.3.1 OLI Fields of View 18](#oli-fields-of-view)

[3.3.1.3.1.1 Unobstructed Field of View 18](#unobstructed-field-of-view)

[3.3.1.3.1.2 Glint-Free Field of View 18](#glint-free-field-of-view)

[3.3.1.3.1.3 Unobstructed Solar Calibration Field of View
18](#unobstructed-solar-calibration-field-of-view)

[3.3.1.3.1.4 Glint-Free Solar Calibration Field of View
18](#glint-free-solar-calibration-field-of-view)

[3.3.1.3.2 TIRS Fields of View 19](#tirs-fields-of-view)

[3.3.1.4 Mass Properties 19](#mass-properties)

[3.3.1.4.1 Mass Allocation 19](#mass-allocation)

[3.3.1.4.1.1 OLI Mass Allocation 19](#oli-mass-allocation)

[3.3.1.4.1.2 TIRS Mass Allocation 19](#tirs-mass-allocation)

[3.3.1.4.1.3 I-SSR Mass Allocation 19](#i-ssr-mass-allocation)

[3.3.1.4.2 Center of Mass 19](#center-of-mass)

[3.3.1.4.2.1 Center of Mass Location 19](#center-of-mass-location)

[3.3.1.4.3 Moments and Products of Inertia
20](#moments-and-products-of-inertia)

[3.3.1.4.3.1 Moments and Products of Inertia Accuracy
20](#moments-and-products-of-inertia-accuracy)

[3.3.1.4.3.2 Moments and Products of Inertia Variation Documentation
20](#moments-and-products-of-inertia-variation-documentation)

[3.3.1.5 Stowed and Critical Clearances
20](#stowed-and-critical-clearances)

[3.3.1.6 Documentation in MID 21](#documentation-in-mid)

[3.3.1.7 Mounting Provisions 21](#mounting-provisions)

[3.3.1.7.1 Instrument Optical Bench 21](#instrument-optical-bench)

[3.3.1.7.2 Instrument Structural Support
22](#instrument-structural-support)

[3.3.1.7.3 I-SSR Accommodation 22](#i-ssr-accommodation)

[3.3.1.7.4 Mounting Method 22](#mounting-method)

[3.3.1.7.5 Mounting Interface 22](#mounting-interface)

[3.3.1.7.5.1 Mounting Interface Documentation
22](#mounting-interface-documentation)

[3.3.1.7.6 Drill Templates 23](#drill-templates)

[3.3.1.7.6.1 Drill Template Usage 23](#drill-template-usage)

[3.3.1.7.6.2 Drill Template Fabrication Requirements
23](#drill-template-fabrication-requirements)

[3.3.1.7.7 Mounting Hardware 23](#mounting-hardware)

[3.3.1.7.7.1 Mounting Hardware Provider 23](#mounting-hardware-provider)

[3.3.1.7.7.2 Mounting Hardware Documentation
24](#mounting-hardware-documentation)

[3.3.1.7.7.3 Mounting Hardware Delivery 24](#mounting-hardware-delivery)

[3.3.1.7.7.4 Mounting Surface Requirements
24](#mounting-surface-requirements)

[3.3.1.7.8 Mounting Location and Documentation
24](#mounting-location-and-documentation)

[3.3.1.7.9 Special Mounts 24](#special-mounts)

[3.3.1.8 Orbit Control, Pointing and Alignment
24](#orbit-control-pointing-and-alignment)

[3.3.1.8.1 Orbit Control 30](#orbit-control)

[3.3.1.8.1.1 Spacecraft State Knowledge 30](#spacecraft-state-knowledge)

[3.3.1.8.2 Pointing 30](#pointing)

[3.3.1.8.2.1 Spacecraft Attitude Knowledge
30](#spacecraft-attitude-knowledge)

[3.3.1.8.2.2 Spacecraft Attitude Control
31](#spacecraft-attitude-control)

[3.3.1.8.2.3 Jitter Sources 31](#jitter-sources)

[3.3.1.8.3 Alignment 31](#alignment)

[3.3.1.8.3.1 Installation Alignment Responsibilities
31](#installation-alignment-responsibilities)

[3.3.1.8.3.2 Installation Alignment References
32](#installation-alignment-references)

[3.3.1.8.3.3 Alignment Knowledge 32](#alignment-knowledge)

[3.3.1.8.3.4 Alignment Control 32](#alignment-control)

[3.3.1.8.3.5 Alignment Stability 33](#alignment-stability)

[3.3.2 Thermal Interface Requirements
33](#thermal-interface-requirements)

[3.3.2.1 General 33](#general-1)

[3.3.2.2 Thermal Recovery 34](#thermal-recovery)

[3.3.2.3 Instrument Thermal Fields of View
34](#instrument-thermal-fields-of-view)

[3.3.2.3.1 OLI Thermal Fields of View 34](#oli-thermal-fields-of-view)

[3.3.2.3.1.1 OLI Thermal FOV 34](#oli-thermal-fov)

[3.3.2.3.2 TIRS Thermal Fields of View 34](#tirs-thermal-fields-of-view)

[3.3.2.4 Heat Transfer 34](#heat-transfer)

[3.3.2.5 Temperature 35](#temperature)

[3.3.2.5.1 Spacecraft Mounting Interface Temperature Requirements
35](#spacecraft-mounting-interface-temperature-requirements)

[3.3.2.5.2 Temperature Monitoring 35](#temperature-monitoring)

[3.3.2.5.2.1 Mechanical Mounting Interface Temperature Monitoring
35](#mechanical-mounting-interface-temperature-monitoring)

[3.3.2.5.2.2 Instrument Temperature Monitoring
36](#instrument-temperature-monitoring)

[3.3.2.6 Thermal Control Design 36](#thermal-control-design)

[3.3.2.6.1 Thermal Control Hardware 36](#thermal-control-hardware)

[3.3.2.6.2 Survival Heaters 36](#survival-heaters)

[3.3.2.6.3 Operational Heaters 36](#operational-heaters)

[3.3.2.6.4 Multilayer Insulation 37](#multilayer-insulation)

[3.3.2.6.5 Surface Cleanliness 37](#surface-cleanliness)

[3.3.3 Electrical Interface Requirements
37](#electrical-interface-requirements)

[3.3.3.1 Power 37](#power)

[3.3.3.1.1 Power Bus Requirements 37](#power-bus-requirements)

[3.3.3.1.1.1 Electrical Interface Definitions
37](#electrical-interface-definitions)

[3.3.3.1.1.2 Power Allocation 37](#power-allocation)

[3.3.3.1.1.2.1 OLI Power Allocations 38](#oli-power-allocations)

[3.3.3.1.1.2.1.1 Operational Average Power Allocation
38](#operational-average-power-allocation)

[3.3.3.1.1.2.1.2 Operational Peak Power Allocation
38](#operational-peak-power-allocation)

[3.3.3.1.1.2.1.3 Survival Power Allocation
38](#survival-power-allocation)

[3.3.3.1.1.2.1.4 Launch-phase Power Allocation
38](#launch-phase-power-allocation)

[3.3.3.1.1.2.1.5 Power Interface Allocation
38](#power-interface-allocation)

[3.3.3.1.1.2.2 TIRS Power Allocations 38](#tirs-power-allocations)

[3.3.3.1.1.2.2.1 Operational Average Power Allocation
38](#operational-average-power-allocation-1)

[3.3.3.1.1.2.2.2 Operational Peak Power Allocation
38](#operational-peak-power-allocation-1)

[3.3.3.1.1.2.2.3 Survival Power Allocation
39](#survival-power-allocation-1)

[3.3.3.1.1.2.2.4 Launch-phase Power Allocation
39](#launch-phase-power-allocation-1)

[3.3.3.1.1.2.2.5 Power Interface Allocation
39](#power-interface-allocation-1)

[3.3.3.1.1.2.3 I-SSR Power Allocations 39](#i-ssr-power-allocations)

[3.3.3.1.1.2.3.1 Operational Average Power Allocation
39](#operational-average-power-allocation-2)

[3.3.3.1.1.2.3.2 Operational Peak Power Allocation
39](#operational-peak-power-allocation-2)

[3.3.3.1.1.2.3.3 Survival Power Allocation
39](#survival-power-allocation-2)

[3.3.3.1.1.2.3.4 Launch-phase Power Allocation
39](#launch-phase-power-allocation-2)

[3.3.3.1.1.2.3.5 Power Interface Allocation
39](#power-interface-allocation-2)

[3.3.3.1.1.3 Power Control 40](#power-control)

[3.3.3.1.1.3.1 Power Connections 40](#power-connections)

[3.3.3.1.1.3.2 Power Application 40](#power-application)

[3.3.3.1.1.3.3 Power Fault Tolerance 40](#power-fault-tolerance)

[3.3.3.1.1.3.4 Instrument Heater Power Separation
41](#instrument-heater-power-separation)

[3.3.3.1.1.3.5 Instrument Internal Power 41](#instrument-internal-power)

[3.3.3.1.1.3.6 Instrument External (Spacecraft) Power
41](#instrument-external-spacecraft-power)

[3.3.3.1.1.3.7 Unannounced Removal of Power
41](#unannounced-removal-of-power)

[3.3.3.1.1.4 Electrical Power Interface Requirements
41](#electrical-power-interface-requirements)

[3.3.3.1.1.4.1 Impedance 42](#impedance)

[3.3.3.1.1.4.2 Survival Heater Bus 42](#survival-heater-bus)

[3.3.3.1.1.4.2.1 Safety Bus Operation 43](#safety-bus-operation)

[3.3.3.1.2 Grounds, Returns, and References
43](#grounds-returns-and-references)

[3.3.3.1.2.1 Grounding Responsibility 43](#grounding-responsibility)

[3.3.3.1.2.2 External Ground Tie Point 43](#external-ground-tie-point)

[3.3.3.1.2.3 Thermal Blanket Grounding 43](#thermal-blanket-grounding)

[3.3.3.1.3 Electrical Harnesses and Connectors
43](#electrical-harnesses-and-connectors)

[3.3.3.1.3.1 Electrical Harnesses 43](#electrical-harnesses)

[3.3.3.1.3.1.1 Harness Wiring Requirements
43](#harness-wiring-requirements)

[3.3.3.1.3.1.2 Harnesses Provider 44](#harnesses-provider)

[3.3.3.1.3.1.3 Harness Documentation 44](#harness-documentation)

[3.3.3.1.3.1.4 Harness Tie Points 44](#harness-tie-points)

[3.3.3.1.3.2 Electrical Connectors 44](#electrical-connectors)

[3.3.3.1.3.2.1 General Considerations 44](#general-considerations)

[3.3.3.1.3.2.2 Connector Location and Types
45](#connector-location-and-types)

[3.3.3.1.3.2.3 Keying 45](#keying)

[3.3.3.1.3.2.4 Flight Plugs 45](#flight-plugs)

[3.3.3.1.3.2.5 Buffer Connectors and Connector Savers
45](#buffer-connectors-and-connector-savers)

[3.3.3.1.3.2.6 Test Connectors: 45](#test-connectors)

[3.3.3.1.3.3 Breakout Boxes 46](#breakout-boxes)

[3.3.3.2 Command and Data Handling 46](#command-and-data-handling)

[3.3.3.2.1 Electrical Interfaces 46](#electrical-interfaces)

[3.3.3.2.1.1 Interface Fault Tolerance 48](#interface-fault-tolerance)

[3.3.3.2.2 Data Bus Requirements 48](#data-bus-requirements)

[3.3.3.2.2.1 Bus Functions 48](#bus-functions)

[3.3.3.2.3 Interface Characteristics 48](#interface-characteristics)

[3.3.3.2.3.1 High Speed Science Data Bus (HSSDB)
48](#high-speed-science-data-bus-hssdb)

[3.3.3.2.3.1.1 HSSDB Data Rate 49](#hssdb-data-rate)

[3.3.3.2.3.1.2 HSSDB Physical Layer 49](#hssdb-physical-layer)

[3.3.3.2.3.1.3 HSSDB Data Format 49](#hssdb-data-format)

[3.3.3.2.3.2 High Speed I-SSR Data Bus (HSIDB)
49](#high-speed-i-ssr-data-bus-hsidb)

[3.3.3.2.3.2.1 HSIDB Data Rate 49](#hsidb-data-rate)

[3.3.3.2.3.2.2 HSIDB Physical layer 50](#hsidb-physical-layer)

[3.3.3.2.3.2.3 HSIDB Data Format 50](#hsidb-data-format)

[3.3.3.2.3.3 MIL-STD-1553 Characteristics
50](#mil-std-1553-characteristics)

[3.3.3.2.3.3.1 Electrical Interface 50](#electrical-interface)

[3.3.3.2.3.4 Pulse Commands 50](#pulse-commands)

[3.3.3.2.3.5 Time of Day Pulse 50](#time-of-day-pulse)

[3.3.3.2.3.6 Synchronization 51](#synchronization)

[3.3.3.2.4 Instrument Commands and Data Load
51](#instrument-commands-and-data-load)

[3.3.3.2.4.1 Commands 51](#commands)

[3.3.3.2.5 Instrument Health and Status Telemetry
52](#instrument-health-and-status-telemetry)

[3.3.3.2.5.1 Point-to-Point Telemetry 52](#point-to-point-telemetry)

[3.3.3.2.5.2 Command Verification 52](#command-verification)

[3.3.3.2.5.3 Instrument Memory Dump 52](#instrument-memory-dump)

[3.3.3.2.5.4 Telemetry Monitor 52](#telemetry-monitor)

[3.3.3.2.6 Command & Data Interface Test Packets
53](#command-data-interface-test-packets)

[3.4 Environmental Conditions 53](#environmental-conditions)

[3.4.1 Radiation Environment 54](#radiation-environment)

[3.4.2 Meteoroid and Debris Environments
54](#meteoroid-and-debris-environments)

[3.4.3 Spacecraft Magnetic Fields 54](#spacecraft-magnetic-fields)

[3.4.4 Atomic Oxygen 54](#atomic-oxygen)

[3.4.5 Spacecraft Charging from All Sources
54](#spacecraft-charging-from-all-sources)

[3.4.6 Launch Environment 54](#launch-environment)

[3.5 DESIGN AND CONSTRUCTION 54](#design-and-construction)

[3.5.1 Instrument Electrical Power Interface Design Requirements
54](#instrument-electrical-power-interface-design-requirements)

[3.5.2 Instrument Mechanical Interface Design Requirements
55](#instrument-mechanical-interface-design-requirements)

[3.5.2.1 Component Stiffness 55](#component-stiffness)

[3.5.2.2 Instrument Mechanisms 55](#instrument-mechanisms)

[3.5.2.3 Uncompensated Momentum 55](#uncompensated-momentum)

[3.5.2.4 Instrument Disturbance Allocations
55](#instrument-disturbance-allocations)

[3.5.2.4.1 Periodic Disturbance Torque Limits
55](#periodic-disturbance-torque-limits)

[3.5.2.4.2 Constant Disturbance Torque Limits
56](#constant-disturbance-torque-limits)

[3.5.2.4.3 Disturbance Torque Limits for Linear Forces
57](#disturbance-torque-limits-for-linear-forces)

[3.5.2.4.4 Spacecraft Non-Operational Deployment Disturbance
57](#spacecraft-non-operational-deployment-disturbance)

[3.5.2.4.5 Torque Profile Documentation
58](#torque-profile-documentation)

[3.5.2.4.6 Thrust Direction Definition 58](#thrust-direction-definition)

[3.5.3 Magnetics 58](#magnetics)

[3.5.4 Access 58](#access)

[3.5.4.1 Access Identification 58](#access-identification)

[3.5.5 General Access 58](#general-access)

[3.5.6 Mounting/Handling 58](#mountinghandling)

[3.5.6.1 Mounting Orientation 58](#mounting-orientation)

[3.5.6.2 Instrument-to-Spacecraft Integration and Test Mounting
59](#instrument-to-spacecraft-integration-and-test-mounting)

[3.5.6.3 Non-Flight Equipment 59](#non-flight-equipment)

[3.5.7 Venting 59](#venting)

[3.5.8 Contamination 59](#contamination)

[3.6 DOCUMENTATION 60](#documentation)

[3.6.1 Interface Control Drawings and Documents
60](#interface-control-drawings-and-documents)

[3.7 Interface Hardware Providers 60](#interface-hardware-providers)

Figures

[Figure 3.3-1. Spacecraft to Instrument Command, Telemetry and Data
Interfaces 16](#__RefHeading___Toc151257673)

[Figure 3.3.1.8-1. Spacecraft/Instrument Allocation of Knowledge (TBR)
27](#__RefHeading___Toc151257674)

[Figure 3.3.1.8-2. Spacecraft / Instrument Allocation of Geolocation
Knowledge Uncertainty 28](#__RefHeading___Toc151257675)

[Figure 3.3.1.8-3. Spacecraft / Instrument Allocation of Geolocation
Control Error 29](#__RefHeading___Toc151257676)

[Figure 3.3.3.2.1-1. Spacecraft to Instrument/Instrument to I-SSR
Electrical Interfaces \[TBR\] 47](#__RefHeading___Toc151257677)

[Figure 3.3.3.2.1-2. Spacecraft to I-SSR Electrical Interfaces \[TBR\]
47](#__RefHeading___Toc151257678)

[Figure 3.5.2.4.1-1. Allowed Transmitted Torque (TBR)
56](#__RefHeading___Toc151257679)

[Figure 3.5.2.4.2-1. Constant Torque vs. Duration of Application (TBR)
57](#__RefHeading___Toc151257680)

Tables

[Table 3.3.2.6.1-1. Thermal Control Hardware Responsibility
36](#__RefHeading___Toc151257681)

[Table 3.5.2.4.1-1. Allowable Transmitted Torque Transition Points (TBR)
56](#__RefHeading___Toc151257682)

# SCOPE

## IDENTIFICATION

This LDCM Interface Requirements Document (IRD) sets forth the general,
common interface requirements imposed on both the Operational Land
Imager (OLI) and Thermal InfraRed Sensor (TIRS) instruments and the
spacecraft for LDCM. This IRD also include requirements that apply to
the interfaces between the Instrument SSR (I-SSR) and the LDCM
instruments, and between the I-SSR and the LDCM spacecraft.

## Applicability

Throughout this document reference will be made to "the OLI", "the
TIRS", "the (LDCM) spacecraft", and "the instrument". References to
individual instruments, i.e. "the OLI" or "the TIRS" apply only to that
instrument; reference to "the spacecraft" or "the LDCM spacecraft"
applies only to the LDCM spacecraft; and reference to "the instrument"
or "the instruments" apply equally to both the OLI and the TIRS. The
I-SSR is considered an "instrument component" that is procured along
with the OLI Instrument, but is a separate component from the OLI
Instrument proper.

## TERMINOLOGY

Throughout this document the following definitions are used:

The term *spacecraft* is interchangeable with spacecraft bus or bus, and
refers to all parts of the space segment that are not the instruments.

The term *Observatory* refers to the fully or partially integrated
system, including the spacecraft and one or more instruments. This term
is used mainly in reference to testing to differentiate between the
spacecraft bus without instruments, and the integrated satellite with
instruments.

The term *instrument provider* refers to the organization / company
delivering the instruments for integration on the spacecraft. The term
is used interchangeably with the instrument contractor or sensor
subcontractor with respect to LDCM. For the purposes of this ICD, the
instrument provider is treated as a single entity, although in reality,
the integrating team will probably include representatives from many
organizations.

The expression 'separately-mounted instrument components' refers to each
part of an instrument which is separately mounted onto the spacecraft by
the spacecraft contractor. Where an instrument is divided into multiple
pieces, but is mounted onto the spacecraft via a single baseplate, that
is not considered 'separately-mounted' instrument components in this
document.

The term \"(TBD)\", which means \"to be determined\", applied to a
missing requirement means that the instrument contractor should
determine the missing requirement in coordination with the spacecraft
contractor.

The term \"(TBS)\", which means \"to be specified\", means that the
spacecraft contractor will supply the missing information in the course
of the contract. These serve as a placeholder for future requirements.
The instrument contractor is not liable for compliance with these
\"placeholder\" requirements, as insufficient information is provided on
which to base a design.

The term \"(TBR)\", which means \"to be refined/reviewed\", means that
the requirement is subject to review for appropriateness by both
contractors, and subject to revision. The instrument contractor is
liable for compliance with the requirement as if the \"TBR\" notation
did not exist. The \"TBR\" merely provides an indication that the value
is more likely to change in a future modification than requirements not
accompanied by a \"TBR\".

## DOCUMENT OVERVIEW

This IRD is controlled and maintained by the LDCM project office. The
IRD is imposed on both the instrument contractor(s) and the spacecraft
contractor, and will serve as a starting point for a detailed set of
Interface Control Documents (ICD) and Mechanical Interface Drawings
(MID) that will serve as the complete specification of the interfaces
involved.

The following Figure 1.4-1 shows the relationship between requirements
documents for the LDCM mission.

Figure 1.4-1. LDCM Requirements
Tree![](media/image3.wmf){width="8.792361111111111in"
height="5.614583333333333in"}

## CONFLICTS

ICDs to more fully define the interface requirements will be generated
by the various contractors during the LDCM implementation. This IRD will
be rendered obsolete when all the ICDs are baselined and put under
formal configuration management.

## REQUIREMENT WEIGHTING FACTORS

The requirements stated in this specification are not of equal
importance or weight. The following four paragraphs define the weighting
factors incorporated in this specification.

a\. ***Shall*** designates the most important weighting level; that is,
mandatory. Any deviations from these contractually imposed mandatory
requirements require the approval of the LDCM contracting officer.

b\. ***Will*** designates mandatory \"requirements\" that are mutually
agreed by the parties. Typically, *will* statements refer to tasks,
including document deliveries. *Will* statements are not verified in the
same manner as interfaces (e.g., analysis, inspection, test,
demonstration), but by mutually agreed schedules or action items. Unless
required by other contract provisions, noncompliance with the *will*
paragraphs does not require approval of the contracting officer, but
does require agreement by all affected parties, and does require
documented technical substantiation.

c\. ***Should*** designates suggestions that are requested, but are not
mandatory. Unless required by other contract provisions, noncompliance
with the *should* paragraphs does not require approval of the
contracting officer, but does require documented technical
substantiation.

d\. ***May*** designates the lowest weighting level. These *may*
statements designate intent and are often stated as examples of
acceptable designs, items, and practices. Unless required by other
contract provisions, noncompliance with the *may* paragraphs does not
require approval of the contracting officer and does not require
documented technical substantiation.

In the remainder of this document, those paragraphs which are intended
for information or clarification, but do not represent requirements,
will be shown in italics. Following such introductory text, it is
possible that the same basic information may exist as a requirement, but
it will be properly numbered as a requirement

# APPLICABLE DOCUMENTS

## GOVERNMENT REQUIREMENT DOCUMENTS

TBD

## NON-GOVERNMENT REQUIREMENT DOCUMENTS

TBD

## REFERENCE DOCUMENTS

TBD

# Observatory Interface REQUIREMENTS

## General

The uncertainty, repeatability and variation values specified in this
document are three (3) sigma unless specified otherwise. Ranges of
values are considered not-to-exceed ranges. All other specification
limits are not-to-exceed values unless specified otherwise.

## Spacecraft to Ground Interface Requirements

### S-band to Ground Interface

#### LANDSAT Ground Network (LGN)

##### Transmit (downlink):

Narrative of the data content (e.g. safety and housekeeping, diagnostic,
etc.): TBD

Baseband data rate (nominal and maximum, if variable): TBD

Transmit frequency, bandwidth: TBD

Modulation and encoding: TBD

Ground G/T: TBD

BER required at ground: TBD

Polarization: TBD

##### Receive (uplink)

Narrative of the data content: TBD

Baseband data rate (nominal and maximum, if variable): TBD

Transmit frequency: TBD

Modulation, encoding, and encryption: TBD

Ground EIRP: TBD

BER required at spacecraft: TBD

Polarization: TBD

#### Space Network (SN)

#####  Transmit (downlink):

Narrative of the data content (e.g. safety and housekeeping, diagnostic,
etc.): TBD

Baseband data rate (nominal and maximum, if variable): TBD

Transmit frequency, bandwidth: TBD

Modulation and encoding: TBD

Ground G/T: TBD

BER required at ground: TBD

Polarization: TBD

##### Receive (uplink)

Narrative of the data content: TBD

Baseband data rate (nominal and maximum, if variable): TBD

Transmit frequency: TBD

Modulation, encoding, and encryption: TBD

Ground EIRP: TBD

BER required at spacecraft: TBD

Polarization: TBD

#### NASA Ground Network (GN)

##### Transmit (downlink):

Narrative of the data content (e.g. safety and housekeeping, diagnostic,
etc.): TBD

Baseband data rate (nominal and maximum, if variable): TBD

Transmit frequency, bandwidth: TBD

Modulation and encoding: TBD

Ground G/T: TBD

BER required at ground: TBD

Polarization: TBD

##### Receive (uplink)

Narrative of the data content: TBD

Baseband data rate (nominal and maximum, if variable): TBD

Transmit frequency: TBD

Modulation, encoding, and encryption: TBD

Ground EIRP: TBD

BER required at spacecraft: TBD

Polarization: TBD

### X-band to Ground Interface

The LDCM Observatory shall support up to three independent, simultaneous
wideband X-band downlinks that may be a combination of one or more of
the following:

a\. Realtime downlink of mission data to LGN.

b\. Playback of recorded mission data to LGN.

c\. Realtime downlink of mission data to International Cooperators.

#### Landsat Ground Network (LGN)

Transmit (downlink):

Narrative of the data content: TBD

Baseband data rate (nominal and maximum, if variable): TBD

Transmit frequency, bandwidth: 8025 -- 8400 MHz

Modulation and encoding: TBD

Ground G/T: TBD

BER required at ground: TBD

Polarization: TBD

#### International Cooperators (ICs)

Transmit (downlink):

Narrative of the data content: TBD

Baseband data rate (nominal and maximum, if variable): TBD

Transmit frequency, bandwidth: 8025 -- 8400 MHz

Modulation and encoding: TBD

Ground G/T: TBD

BER required at ground: TBD

Polarization: TBD

## Spacecraft to Instrument Interface Requirements

The system interfaces relevant to the instruments are depicted in Figure
3.3‑1. They are defined here in terms of the mechanical and electrical
requirements.

![](media/image4.wmf){width="6.634645669291339in"
height="5.011023622047244in"}

[]{#__RefHeading___Toc151257673 .anchor}Figure 3.3-1. Spacecraft to
Instrument Command, Telemetry and Data Interfaces

### Mechanical Interface Requirements

All mechanical requirements specified shall be met at the mechanical
interface; that is, at the surface(s) of the spacecraft where the
instrument is in contact with the spacecraft, unless otherwise
specifically indicated.

*Where kinematic mounts are used, the mechanical interface is the
surface of the mount that* contacts the spacecraft optical bench.

#### Dimensions

All interface documents shall use metric units. An exception is allowed
for interface documents in native imperial units. The metric units shall
be displayed above the converted units with the imperial units in
parentheses.

Dimensioning shall be in the as-designed units, and identified when
other than SI.

Analysis results produced by existing software tools may be in heritage
units.

The design of the instrument shall meet the dimensional envelope
constraints defined in the ICD under a combination of static, dynamic,
and thermal conditions encountered during factory assembly, system test,
transportation and handling, launch, deployment, and on-orbit
operations.

#### Reference Frames

##### Observatory Reference Frame

Reference Ellipsoid: The WGS 84 reference ellipsoid as defined by NIMA
TR8350.2.

Nadir: Defined at any point in the orbit as the direction toward the
center of mass of the Earth.

Inertial Velocity: The velocity relative to the ECI (J2000.0) reference
frame.

The LDCM Orbital Reference Frame shall be a right-handed, orthogonal,
XYZ coordinate system defined such that the origin is at the spacecraft
center of mass (CM), the +Z axis points toward Nadir, the +X axis is
coplanar with both the +Z axis and the spacecraft inertial velocity
vector (and is in the general direction of the spacecraft inertial
velocity vector), and the +Y axis completes the right-handed, orthogonal
coordinate system as shown in Figure 3.3.1.2.1-1.

![](media/image1.wmf)

Figure 3.3.1.2.1-1 Orbital Reference Frame

##### Spacecraft Body Frame

The LDCM Spacecraft Body Frame shall be a right-hand, orthogonal,
body-fixed XYZ coordinate system with the +Z direction pointing to
nadir.

The associated mechanical datums and origin shall be described in the
MID.

#### Instrument Fields of View

##### OLI Fields of View

###### Unobstructed Field of View

The LDCM spacecraft shall provide an unobstructed minimum nadir field of
view for OLI that provides a 15.0 degree cross track width by at least
1.7 degree along track length for any OLI optical aperture location that
falls within the X-Y dimension constraints of the instrument.

The boresight of the unobstructed field of view shall be parallel to the
nadir (+Z) axis.

###### Glint-Free Field of View

The LDCM spacecraft shall provide the OLI an unobstructed FOV within a
conical 25° (TBR) half angle of optical nadir for any OLI optical
aperture location that falls within the X-Y dimension constraints of the
instrument aperture.

The cone's circular intersection on the +Z face shall envelope the
oval-shaped nadir port aperture.

###### Unobstructed Solar Calibration Field of View

The LDCM spacecraft shall provide the OLI with a solar calibration field
of view unobstructed by spacecraft structures for any time of the year
across a range of solar elevation angles from -15° to -25° **(TBR)**
relative to the local horizontal for the post eclipse exit portion of
the orbit. 

To cover the seasonal variation in the solar azimuthal position, the
clear solar calibration field of view for the indicated solar elevation
range shall be 38°± 8° (TBR) in azimuth towards the sun side from the
velocity vector for the post eclipse exit portion of the orbit. 

###### Glint-Free Solar Calibration Field of View

The LDCM spacecraft shall provide the OLI an unobstructed FOV within a
region bounded by --10 to --30° **(TBR)** in elevation and 38°± 60°
(TBR) in azimuth (towards the sun side) from the edges of the solar
calibration port.

The bounding surface of this region's intersection with the instrument
shall envelope the rectangular shaped solar calibration aperture. The
spacecraft contractor shall accommodate this FOV except for the
obstructions shown in Figure 3-3.

##### TIRS Fields of View

TBD

#### Mass Properties

In the paragraphs below, the term \"actual\" refers to values which
apply to a given serial number. As indicated below, these are generally
provided in a data package associated with delivery of each individual
serial numbered instrument.

The terms \"typical\" and \"predicted\" are used to document a nominal
value in the ICD. These may be the result of calculations or
measurements on a non-flight unit which is expected to be representative
of the value for a flight unit.

##### Mass Allocation

###### OLI Mass Allocation

The total mass of the OLI and all associated hardware and harnessing
shall not exceed 350 kg (TBR).

The mass of the OLI shall be measured to an accuracy of ± 0.1 kg at
delivery to the spacecraft.

###### TIRS Mass Allocation

The total mass of the TIRS and all associated hardware and harnessing
shall not exceed 200 kg (TBR).

The mass of the TIRS shall be measured to an accuracy of ± 0.1 kg at
delivery to the spacecraft.

###### I-SSR Mass Allocation

The total mass of the I-SSR and all associated hardware and harnessing
shall not exceed 90 kg (TBR).

The mass of the I-SSR shall be measured to an accuracy of ± 0.1 kg at
delivery to the spacecraft.

##### Center of Mass

###### Center of Mass Location

The typical launch and on-orbit center of mass of each separately
mounted instrument component shall be documented in the MID, referenced
to the instrument coordinate axes.

The actual stowed (launch) and deployed (on-orbit) centers of mass of
each separately-mounted instrument component shall be measured by the
instrument contractor, reported to ± 6 mm (not-to-exceed), referenced to
the instrument coordinate axes and documented in the data package for
that serial number instrument.

The instrument coordinate axes shall be defined to be in the same
orientation as the spacecraft axes, but not necessarily the same origin.

##### Moments and Products of Inertia

The moments and products of inertia shall be measured or calculated for
each separately-mounted instrument component, using coordinates based on
the spacecraft axes but passing through the instrument component center
of mass.

These moments and products of inertia shall be provided to the
spacecraft contractor for documentation in the ICD.

###### Moments and Products of Inertia Accuracy

Moments and products of inertia values shall be accurate to within ± 5 %
for calculated values, and the lesser of 5% or 300 kg-cm² for measured
values.

###### Moments and Products of Inertia Variation Documentation

If the instrument contains movable masses, expendable masses, or
deployables, the typical respective moments and products of inertia
variations during the deployment, specified at least at the beginning,
middle, and end of the deployment, shall be provided to the spacecraft
contractor for documentation in the ICD.

If an instrument contains movable masses or deployable items, the
respective moments and products of inertia variations shall be measured
or calculated and documented in the delivered data package.

#### Stowed and Critical Clearances

Both the spacecraft contractor and the instrument contractor must work
together to insure that the stowed, deploying, and final deployed
positions of the instrument will clear all obstacles, including
obstacles on the spacecraft, other instruments, and the launch vehicle.

The instruments shall fit within the static envelope of launch vehicle's
fairing.

Instrument components in all configurations shall be contained within
the instrument envelopes as allocated within the MID, but different
envelopes may be defined for launch configuration, on-orbit (deployed)
configuration, and during deployment.

The defined not-to-exceed envelope includes, but is not limited to, the
basic instrument structure, instrument-unique mounts (including
kinematic mounts, where used), instrument-unique mounting hardware,
mounting feet, thermal blankets (including billowing during launch and
ascent), connectors (instrument-side only), and intra-instrument harness
(with connectors), as well as any other items that constitute any part
of a deliverable instrument.

The intent of this requirement is that everything provided as part of a
deliverable instrument, including the items identified above, be
allocated against the instrument size and weight allocations, and that
everything provided by the spacecraft contractor be held against the
spacecraft allocation.

The instruments shall be laid out on the spacecraft so as to maintain
adequate clearance between the instrument and surrounding structures, in
order to provide access to instrument mounting hardware, access to
instrument connectors, and space for instrument interfacing harness
service loops.

#### Documentation in MID

The instrument subsystem envelope (including thermal blankets,
instrument connectors, mounting feet, instrument-supplied harnesses, and
unique mounts) will be documented in the MID.

Documentation is to be in the form of engineering drawings with a set of
\"not to exceed\" dimensions for the launch, deployed on-orbit, and
installation envelopes. The instrument installation envelope is defined
as the volume of the instrument including the attached lifting sling.

All instrument fields of view, including, where applicable, optical,
glint, thermal cryo-radiator, electronics radiator, and solar diffuser
fields of view, will be documented in the MID.

Information on field-of-view-related issues such as clearances,
multi-path tolerances, indirect effects such as glint and allowable
secondary radiation, and percent interference allowed will be documented
in the MID.

#### Mounting Provisions

##### Instrument Optical Bench

The spacecraft shall provide an optical bench as the mounting surface
for the TIRS and OLI instruments.

##### Instrument Structural Support

The spacecraft shall provide structural support for the instruments such
that the loads transmitted through the optical bench and across the
interface into the instruments do not exceed the interface limit loads
determined by the spacecraft contractor.

##### I-SSR Accommodation

The I-SSR shall not be mounted on the instrument optical bench.

The I-SSR shall be mounted on or within the main spacecraft structure.

The spacecraft shall provide the thermal path for all I-SSR heat
rejection.

##### Mounting Method

The mounting method shall accommodate manufacturing tolerance,
structural distortion, thermal distortions, and alignment requirements.

The spacecraft to instrument mounting interface will be established by
the spacecraft contractor from the instrument provided drill template
and shown in the MID. This process aligns the instrument mechanical
datums to the spacecraft body frame within the spacecraft alignment
control allocation precluding the need to shim or correct the instrument
orientation at instrument installation.

If required, the spacecraft contractor will provide contingency shims to
correct instrument alignment control error at installation. Shim
capability for out of plane rotations (spacecraft X & Y axis) will be
limited by pin / bolt clearance. There is no adjustment capability for
in plane rotation (about spacecraft Z axis) where close fit pins are
used.

Potential mounting interface gaps (prior to fastener full preload) will
be managed through instrument, spacecraft, and drill template flatness
and / or co-planarity requirements to preclude the need to shim for gaps
at installation.

If required, the spacecraft contractor will provide contingency shims to
close gaps at installation. Shim capability will be limited by pin /
bolt clearances and access required to inspect or perform gap
measurements.

The instrument mounting method shall require access only from outside
the spacecraft.

Specific requirements for pinning the instrument, if required, will be
documented in the MID.

##### Mounting Interface

###### Mounting Interface Documentation

The spacecraft mounting interface requirements for each
separately-mountable instrument component shall be defined in the MID.

##### Drill Templates

###### Drill Template Usage

The drill template is used during the manufacturing of the spacecraft
bus to establish instrument interface features. The drill template is
supported from a gantry (mechanical ground support equipment). The
template/instrument mounting datum features are aligned to the
spacecraft body frame with laser tracker and/or theodolite measurements.
With the template properly aligned, spacecraft secondary structure is
matched to the template through a combination of drill, bolt, and
bonding processes as required.

The drill template may also be used for interface and alignment
verification during the spacecraft manufacturing and AI&T phases.

If templates are used, then instrument equipment, spacecraft, and test
fixture interfaces shall be established using the same templates (i.e.
face to face master).

###### Drill Template Fabrication Requirements

Drill template design features (including, but not limited to surface
flatness, surface finish, surface treatment, etc.) will be documented in
the MID.

The material shall be aluminum (preferred), steel or other suitable
material.

Drill templates for each component that require alignment shall include
an optical alignment target or cube bonded to the template.

Optical alignment target / cube nominal location and orientation with
respect to mounting datum(s) will be as shown on MID.

Optical alignment target / cube nominal location and orientation with
respect to mounting datum(s) shall be similar to instrument.

##### Mounting Hardware

###### Mounting Hardware Provider

The instrument contractor will provide all unique instrument mounting
hardware (e.g., bolts, washers, etc. which have limited-off-the-shelf
availability or require a special fabrication lot).

The spacecraft contractor will provide all standard instrument mounting
hardware (e.g., bolts, washers, shims, etc. which are procurable
off-the-shelf from multiple hardware vendors).

###### Mounting Hardware Documentation

Instrument mounting hardware will be defined and documented in the MID,
including indication of the source of the hardware (spacecraft or
instrument).

###### Mounting Hardware Delivery

Delivery of mounting hardware with the instrument will be accomplished
with each separately-mounted or attached element to be separately
kitted.

Each unique hardware type or size will be separately packaged and
clearly identified with its relationship to the instrument top assembly
drawing.

###### Mounting Surface Requirements

Finish requirements for the mounting surfaces will be specified by the
spacecraft contractor and documented in the MID.

The spacecraft mounting surface flatness shall be as documented in the
MID.

The instrument mounting surface flatness shall be as documented in the
MID.

The spacecraft mounting surface planarity shall be as documented in the
MID.

The instrument mounting surface planarity shall be as documented in the
MID.

##### Mounting Location and Documentation

The instrument contractor will support the spacecraft contractor in
determining the location of the instrument on the spacecraft.

The instrument contractor will provide necessary mounting information to
the spacecraft contractor for documentation in the MID.

The instrument mounting location will be documented in the MID.

##### Special Mounts

Instrument components shall be mounted using kinematic mounts unless the
instrument contractor determines that kinematic mounts are not required.

The instrument will be delivered to the spacecraft contractor with the
flight kinematic mount installed.

#### Orbit Control, Pointing and Alignment

Pointing: The process of controlling the location or direction of a Line
of Sight with respect to an intended target location or direction.

Instrument Interface: The mechanical interface and associated datum(s)
at the location where the instrument attaches to the spacecraft.

Jitter: Rotations due to elastic as well as rigid body vibrations caused
by disturbance sources such as reaction wheels, solar array drive
assemblies and instrument mechanisms.

Spacecraft Attitude Determination Frame: The right-handed orthogonal
reference frame which approximates the Spacecraft Body Frame and of
which the Attitude Determination System on board the spacecraft
determines inertial attitude. For LDCM, the datums associated with the
Spacecraft Control Frame are (TBS). The Spacecraft Attitude
Determination Frame is defined by the relative orientation of these
datums with respect to the Spacecraft Body Frame, as established by
measurements during installation alignment or on-orbit calibration.

Spacecraft Target Frame: The right-handed, orthogonal reference frame
that, at any time, is the target attitude of the Spacecraft Attitude
Determination Frame.

3 Sigma: A set of values is considered to meet a 3 sigma requirement if
no fewer than 99.73% of the values are within the specified limits of
the requirement.

Root Sum Square (RSS): The square root of the sum of the squares of
components of error or uncertainty.

Alignment: The relative orientation of 2 reference frames.

Installation Alignment: The process of setting, measuring and/or
adjusting the relative orientation of an instrument or hardware
reference frame with respect to another reference frame in order to
satisfy required alignment criteria.

Alignment Shift: A change in the alignment between 2 reference frames
that occurs over one finite, defined, period of time, typically prior to
and not changing significantly during normal on-orbit operations. For
the alignment of the instruments relative to the spacecraft, alignment
shifts arise from structural distortion effects such as launch loads,
moisture out-gassing and the relief of gravity loads.

Alignment Drift: Changes in the alignment between 2 reference frames
that occur during normal on-orbit operations, excluding effects due to
jitter. For the alignment of instruments relative to the spacecraft, one
source of drift is structural distortions arising from orbital and
seasonal changes in thermal conditions.

Alignment Knowledge: The estimate of the relative orientation between 2
reference frames.

Alignment Knowledge Uncertainty: The uncertainty in the knowledge of the
relative orientation between 2 reference frames.

Boresight Alignment Knowledge Uncertainty: For a specific instrument,
the uncertainty in the knowledge of the relative orientation between its
Instrument Boresight and the Spacecraft Attitude Determination Frame.
This uncertainty will be considered to be composed of the following
factors:

-   Static (not changing significantly during normal on orbit
    operations)

Measurement uncertainty during installation alignment.

Alignment Shifts.

-   Dynamic (changing significantly during normal on orbit operations)

Alignment Drifts.

Jitter.

Alignment Control: The process of controlling the maximum difference
between the actual orientation of a reference frame and its nominal
orientation.

Alignment Control Error: The maximum difference between the actual
orientation of a reference frame and its nominal orientation.

Boresight Alignment Control Error: For a specific instrument, the
maximum difference between the actual and the nominal orientation of its
Instrument Boresight relative to the Spacecraft Attitude Determination
Frame. This error will be considered to be composed of the following
factors:

-   Static (not changing significantly during normal on orbit
    operations)

Measurement uncertainty during installation alignment.

Limitations associated with fabrication and installation.

Alignment Shifts.

-   Dynamic (changing significantly during normal on orbit operations)

Alignment Drifts.

Jitter.

Figures 3.3.1.8.-1, 3.3.1.8.-2, and 3.3.1.8.-3 illustrate a framework
for representing the train of components of Knowledge Uncertainty and
Control Error between an Instrument Boresight and inertial references.

![](media/image2.wmf){width="4.761023622047244in"
height="3.136220472440945in"}

[]{#__RefHeading___Toc151257674 .anchor}Figure 3.3.1.8-1.
Spacecraft/Instrument Allocation of Knowledge (TBR)

![](media/image3.wmf){width="7.111805555555556in"
height="5.710416666666666in"}

[]{#__RefHeading___Toc151257675 .anchor}Figure 3.3.1.8-2. Spacecraft /
Instrument Allocation of Geolocation Knowledge Uncertainty

![](media/image4.jpeg){width="7.293055555555555in"
height="5.6819444444444445in"}

[]{#__RefHeading___Toc151257676 .anchor}Figure 3.3.1.8-3. Spacecraft /
Instrument Allocation of Geolocation Control Error

##### Orbit Control

###### Spacecraft State Knowledge

The spacecraft shall supply to the ground an estimate of Spacecraft
position and velocity for ground processing.

The spacecraft shall use the International Atomic Time (TAI) time as the
observatory time reference.

The spacecraft shall provide the orbital position and velocity in the
WGS84 Earth Centered Earth Fixed coordinate frame.

The spacecraft shall provide an estimate of position and velocity
time-tagged to an accuracy of 200.0 microseconds \[TBC\], 3-sigma, or
less, with a frequency of at least once per second.

The spacecraft shall provide orbital position knowledge accurate to 30 m
\[TBR\] radial, 30 m \[TBR\] in-track, and 30 m \[TBR\] cross-track --
all values are 3-sigma.

The spacecraft shall provide orbital velocity knowledge accurate to 0.3
m/sec \[TBR\] radial, 0.3 m/sec \[TBR\] in-track, and 0.3 m/sec \[TBR\]
cross track -- all values are 3-sigma.

##### Pointing

###### Spacecraft Attitude Knowledge

The spacecraft shall supply to the ground an estimate of Spacecraft
attitude for ground processing.

The spacecraft shall provide the attitude estimate as a quaternion
relative to the Earth Centered Inertial of epoch J2000.0 coordinate
frame.

The spacecraft shall provide inertial attitude knowledge of the
instrument mounting surface with an accuracy of 109 micro-radians,
3-sigma, per axis, for each instrument during imaging.

The spacecraft shall provide an estimate of the inertial attitude
knowledge time-tagged to an accuracy of 200.0 microseconds \[TBC\],
3-sigma, or less, with a frequency of at least once per second.

The spacecraft shall provide relative inertial attitude knowledge
accuracy over a 30 second period of 11 micro-radians or less, 3-sigma,
per axis, during imaging.

The spacecraft shall provide relative inertial attitude knowledge
accuracy over a 2.5 second period of 5 micro-radians or less, 3-sigma,
per axis, during imaging.

The spacecraft shall supply to the ground the attitude and/or attitude
rate measurements provided by the attitude determination sensors, at the
full attitude sensor data rate, for ground processing.

###### Spacecraft Attitude Control

During normal operation, the Spacecraft Attitude Control Error is the
difference between the orientation of the Spacecraft Target Frame and
the orientation of the Spacecraft Attitude Determination Frame. During
normal operation, the Spacecraft Attitude Control Rate Error is the
difference between the angular rate of the Spacecraft Target Frame and
the angular rate of the Spacecraft Attitude Determination Frame.

The LDCM Observatory pointing reference shall be with respect to the
Local Vertical / Local Horizontal reference frame.

The spacecraft shall limit pointing control error to less than 525
micro-radians \[TBR\], 3-sigma, per axis, during imaging periods.

The spacecraft shall limit pointing rate error to less than 145
micro-radians \[TBR\] / sec, 3-sigma, per axis, during imaging periods.

###### Jitter Sources

The spacecraft shall limit the high frequency disturbances from all
sources applied to the instrument interface to no greater than TBD,
3-sigma, per axis, during imaging.

##### Alignment

###### Installation Alignment Responsibilities

The instrument contractor shall provide an Instrument Alignment
Reference(s) composed of an alignment target or cube with associated
alignment data relative to the Instrument Boresight and the mounting
surface datum(s) associated with the instrument.

The spacecraft contractor will provide a Spacecraft Alignment
Reference(s) composed of an alignment target or cube with associated
alignment data relative to the Spacecraft Body Frame.

The alignment data for the Instrument Boresight relative to the
Instrument Alignment Reference(s) will be provided to the spacecraft
contractor 6 months prior to the delivery of the instrument.

The process for verifying gaps will be documented in the MID.

The spacecraft contractor will measure the alignment between the
Instrument Alignment Reference and the Spacecraft Alignment Reference.

###### Installation Alignment References

The Instrument Alignment References shall be viewable from two
orthogonal directions when integrated to the satellite.

The type, location, and orientation of the Instrument Alignment
References will be mutually agreed upon between the instrument
contractor and the spacecraft contractor, and documented in the MID.

Optical targets or cubes shall have a per-face surface area of at least
360 mm2.

At least two viewing surfaces of the optical alignment target or cube
shall be orthogonal to within ± 15 micro-radians, 3 sigma.

Each optical target or cube shall be covered with a flight quality (and
flight capable) cover.

The optical cube cover shall be removable during Integration and Test.

The optical cube cover shall be installed prior to integration onto the
launch vehicle.

The cover shall provide captive hardware to prevent loose pieces if the
cube comes loose.

###### Alignment Knowledge

All components of Boresight Alignment Knowledge Uncertainty between the
Spacecraft Attitude Determination Frame and the Instrument Interface
(mounting surface datums) will be allocated to the spacecraft.

All components of Boresight Alignment Knowledge Uncertainty between the
Instrument Interface (mounting surface datums) and the Instrument
Boresight will be allocated to the instrument.

Components of Boresight Alignment Knowledge Uncertainty resulting from
alignment shifts allowed by clearances across the mounting interface
(pin/fastener to hole tolerance) at the Instrument Interface (mounting
surface datums) will be allocated to the instrument.

###### Alignment Control 

All components of Boresight Alignment Control Error between the
Spacecraft Attitude Determination Frame and the Instrument Interface
(mounting surface datums) will be allocated to the spacecraft

For each instrument, the Boresight Alignment Control Error will be
controlled by an allocation from the Instrument Boresight to the
Instrument Interface mechanical datums, as well as an allocation from
the Instrument Interface mechanical datums to the Spacecraft Attitude
Determination Frame.

For each instrument, the Boresight Alignment Control Error allocations
will be documented in the instrument ICD.

All components of Boresight Alignment Control Error between the
Instrument Boresight and the Instrument Interface (mounting surface
datums) will be allocated to the instrument.

Components of Boresight Alignment Control Error resulting from alignment
shifts allowed by clearances across the mounting interface (pin/fastener
to hole tolerance) at the Instrument Interface (mounting surface datums)
will be allocated to the instrument.

###### Alignment Stability

The spacecraft shall maintain alignment stability of 9 microradians or
less, from the OLI mounting interface to the observatory attitude
reference, over any period of 30 seconds.

The spacecraft shall maintain alignment stability of 30 microradians or
less, from the OLI mounting interface to the observatory attitude
reference, over any period of 16 days.

The spacecraft shall maintain alignment stability of 28 (TBR)
microradians or less, from the TIRS mounting interface to the OLI
mounting interface, over any period of 16 days.

### Thermal Interface Requirements

#### General

The operating and survival temperatures, as well as the thermal
isolation requirements specified in this section shall be met at the
mechanical interface between the spacecraft and the instrument.

The instrument thermal design shall provide for utilizing spacecraft
survival bus power for maintaining the instrument at or above the
minimum survival temperature when the instrument is off (including
launch configuration) and at or above the minimum turn-on temperature
before the instrument is activated.

#### Thermal Recovery

The instrument shall take no more than 60 \[TBR\] minutes to return from
SAFE mode to a temperature condition capable of normal operation,
assuming no anomalies.

This time limitation does not include passive radiative detector
coolers.

The instrument shall take no more than 120 \[TBR\] minutes to return
from SURVIVAL mode to a temperature condition capable of normal
operation, after spacecraft application of operational power.

This time limitation does not include passive radiative detector
coolers.

#### Instrument Thermal Fields of View

Note: Thermal analysis will verify the heat transfer adequacy at the
final mounting locations.  Details of adjacent hardware within the
instruments' thermal FOV shall be provided by the spacecraft to a
mutually agreed upon level of fidelity.

##### OLI Thermal Fields of View

###### OLI Thermal FOV

The spacecraft shall provide to the cold face of the OLI a minimally
obstructed hemispherical FOV for heat dissipation.

##### TIRS Thermal Fields of View

The spacecraft shall provide to the cold face of the TIRS a minimally
obstructed hemispherical FOV for heat dissipation.

#### Heat Transfer

The instrument modules shall be thermally isolated to the maximum extent
possible to minimize heat transfer to the spacecraft and other adjacent
instruments.

To maintain flexibility in the placement of instrument modules on the
spacecraft, it is necessary to limit the total heat transfer (conduction
and radiation) to the spacecraft and other instruments. The thermal
isolation is accomplished by minimizing thermal conduction and radiation
using low thermal conductivity mechanical interface mounts and low
effective emissivity MLI blankets at all non-radiator instrument
surfaces. Radiation to the earth or to cold space is intended to be the
primary means of dissipating waste heat from the instrument.

The total orbit-average heat transfer rate (conducted and radiated)
between the instrument and the spacecraft shall not exceed 10 watts into
or out of the instrument mounting surface.

This includes heat conducted through the harness and ground straps.

Each instrument module total orbit average heat transfer rate (conducted
and radiated) between the instrument and the spacecraft, divided by the
footprint area, shall not exceed 15 watts per square meter into or out
of the instrument.

This includes heat conducted through the harness and ground straps.

*The term 'footprint area\" as used in this requirement means the area
under the instrument module, projected onto the spacecraft. It does not
refer to the area of the footprint of the kinematic mounts*.

#### Temperature 

Temperature limits for instrument components during ground test and
orbital operations shall be provided by the instrument contractors to
the spacecraft contractor for documentation in the ICD.

Turn-on temperature requirements, relative to the passive analog
temperature sensors, shall be provided by the instrument contractors to
the spacecraft contractor for documentation in the ICD.

##### Spacecraft Mounting Interface Temperature Requirements

In normal operational mode, the spacecraft shall maintain the
temperature of the mechanical instrument-to-spacecraft mounting surface
within the range specified in the ICD.

The instrument shall meet all specified performance requirements when
the temperature of the mechanical instrument-to-spacecraft mounting
surface is in the normal operating range.

The interface on the LDCM spacecraft shall remain in the range specified
in the ICD during survival mode.

##### Temperature Monitoring

The location of the spacecraft-provided interface temperature monitoring
sensors will be shown in the MID.

The location of all instrument contractor-supplied temperature sensors
and calibration data will be provided to the spacecraft contractor for
documentation in the MID.

###### Mechanical Mounting Interface Temperature Monitoring

The spacecraft shall monitor and report the temperatures of the
spacecraft at the instrument mechanical mounting interfaces in the
spacecraft telemetry.

###### Instrument Temperature Monitoring

All critical instrument temperatures shall be reported by the instrument
in the health and status telemetry data.

#### Thermal Control Design

##### Thermal Control Hardware

The instrument contractor will provide to the spacecraft contractor
information on instrument-provided thermal control hardware for
documentation in the ICD

The responsibility for providing the thermal control hardware is defined
in Table 3.2.4.7.6.1-1.

[]{#__RefHeading___Toc151257681 .anchor}Table 3.3.2.6.1-1. Thermal
Control Hardware Responsibility

  ---------------------------------------------------------------------------
  **Hardware**                                           **Responsibility**
  ------------------------------------------------------ --------------------
  Survival heaters                                       Instrument
                                                         Contractor

  Instrument thermal control hardware, including         Instrument
  blankets, louvers, and heat pipes                      Contractor

  Instrument-to-spacecraft interface operational and     Spacecraft
  survival heaters, thermistors, thermostats             Contractor

  Thermal close-out blankets to interface between the    Spacecraft
  instrument thermal blankets and the spacecraft thermal Contractor
  blankets                                               

  Instrument-to spacecraft interface MLI blankets        Spacecraft
                                                         Contractor
  ---------------------------------------------------------------------------

##### Survival Heaters

Instruments shall utilize survival heaters to maintain temperatures at
or above minimum survival limits, when the operational power bus has
been disconnected from the instrument.

Survival heaters shall not be required to provide heat when the
operational bus is active to the instrument.

In agreement with the instrument contractor, the spacecraft contractor
may elect to leave the survival heater bus powered during the initial
phase of warm-up upon exit from SURVIVAL mode, but the instrument must
be capable of achieving thermal control, once the operational power bus
is activated, without use of the survival heaters.

##### Operational Heaters

When the instrument operational power bus is active, operational heaters
located inside the instrument, and controlled by the instrument, shall
be used for normal thermal control.

##### Multilayer Insulation

The spacecraft contractor is responsible for approving the Multilayer
Insulation (MLI) selection in the Parts, Materials and Processes Control
Board (PMPCB) review process.

MLI used in thermal control design shall have the following provisions:
venting, interfacing with spacecraft thermal control surfaces, and
electrical grounding to prevent Electro-Static Discharge (ESD).

##### Surface Cleanliness

Thermal control surfaces shall be cleanable to visibly clean or better.

### Electrical Interface Requirements

#### Power

##### Power Bus Requirements

###### Electrical Interface Definitions 

Electrical Interface Location: All requirements apply at the electrical
interface, which is at the instrument end of the
instrument-to-spacecraft bus harness connector mating surfaces.

Operational Power: Operational power is used for instrument operational
modes such as Science Data Collection, Calibration, and Standby.

Peak Operational Power: Peak power is the maximum power required by an
instrument. Peak power does not include transients with a duration less
than 20 milliseconds.

Orbital Average Operational Power: The orbital average power is the
average power utilized by an instrument over any one-orbit period
commencing with the crossing of the night-to-day terminator.

Survival-Mode Power: Survival-mode power is power required by the
instrument in Survival Mode, in order to operate survival heaters.

Safe-Mode Power: Safe-mode power is the power required by the instrument
in instrument safe mode.

Launch-Phase Power: Launch-phase power is the power required by the
instrument in launch phase.

###### Power Allocation

The Spacecraft will provide suitable power harnesses to support the
allocated power levels specified at the instrument interface.

This section specifies the characteristics, connections, and control of
the Spacecraft power provided to the instrument and requirements at this
interface.

####### OLI Power Allocations

######## Operational Average Power Allocation

The orbital average OLI power consumed in operational modes shall be
less than 250W (TBR).

######## Operational Peak Power Allocation

The peak OLI power consumed in operational modes shall be less than 400W
(TBR).

######## Survival Power Allocation

The OLI power consumed in survival mode shall be less than 150W (TBR).

The OLI survival heaters shall be sized using an assumed minimum
spacecraft bus voltage of 25 V.

######## Launch-phase Power Allocation

The OLI power consumed in launch phase shall not exceed the survival
mode power allocation.

######## Power Interface Allocation

The number of power interfaces used by the OLI shall be consistent with
the number allocated in the ICD.

Power Bus connections and power profiles for all OLI modes will be
provided by the OLI contractor and documented in the ICD.

If the OLI power requirements vary as a function of orbit position, that
duty cycle will be supplied by the OLI contractor and documented in the
ICD.

####### TIRS Power Allocations

######## Operational Average Power Allocation

The orbital average TIRS power consumed in operational modes shall be
less than 250W (TBR).

######## Operational Peak Power Allocation

The peak TIRS power consumed in operational modes shall be less than
400W (TBR).

######## Survival Power Allocation

The TIRS power consumed in survival mode shall be less than 150W (TBR).

The TIRS survival heaters shall be sized using an assumed minimum
spacecraft bus voltage of 25 V.

######## Launch-phase Power Allocation

The TIRS power consumed in launch phase shall not exceed the survival
mode power allocation.

######## Power Interface Allocation

The number of power interfaces used by the TIRS shall be consistent with
the number allocated in the ICD.

Power Bus connections and power profiles for all TIRS modes will be
provided by the TIRS provider and documented in the ICD.

If the TIRS power requirements vary as a function of orbit position,
that duty cycle will be supplied by the TIRS provider and documented in
the ICD.

####### I-SSR Power Allocations

######## Operational Average Power Allocation

The orbital average I-SSR power consumed in operational modes shall be
less than 200W (TBR).

######## Operational Peak Power Allocation

The peak I-SSR power consumed in operational modes shall be less than
300W (TBR).

######## Survival Power Allocation

The I-SSR power consumed in survival mode shall be less than 100W (TBR).

The I-SSR survival heaters shall be sized using an assumed minimum
spacecraft bus voltage of 25 V.

######## Launch-phase Power Allocation

The I-SSR power consumed in launch phase shall not exceed the survival
mode power allocation.

######## Power Interface Allocation

The number of power interfaces used by the I-SSR shall be consistent
with the number allocated in the ICD.

Power Bus connections and power profiles for all I-SSR modes will be
provided by the I-SSR contractor and documented in the ICD.

###### Power Control

####### Power Connections

Three supply circuit types shall be provided by the spacecraft, with
specific allocations by instrument listed in the ICD:

a\) An operational power connection made through an appropriately-sized
overcurrent protection device and switch, both on the spacecraft side of
the interface

b\) A power connection (survival heater bus) made through an
appropriately-sized overcurrent protection device and a switch, both on
the spacecraft side of the interface

c\) A bus or set of isolated power services (safety bus) to supply power
for safety-related functions such as explosive bolts, thermal release
mechanisms, etc.

Each power interface shall consist of redundant two (isolated) power
feeds routed via two separate connectors. These connections will be
designated as power feeds A and B.

The nominal and maximum load currents for each power interface line, for
each instrument mode, shall be provided by the instrument contractor and
documented in the ICD.

The state of primary and redundant power supply during normal operations
shall be identified in the instrument telemetry.

If the power configuration is totally under the control of the
spacecraft, no additional power status telemetry is needed from the
instrument.

####### Power Application

In the absence of a fault (hardware, software, or operator) on either
power feed (i.e., an open fuse), the spacecraft shall apply power to
neither operational power feed (instrument OFF) or a single operational
power feed (instrument ON), but never to both operational power feeds
simultaneously.

####### Power Fault Tolerance

The instrument and spacecraft shall not propagate a single fault
occurring on either the A or B power interface circuit, on either side
of the interface, to the redundant interface or instrument.

####### Instrument Heater Power Separation

Instrument survival heaters and operational heaters shall be separate
and distinct, using separate electrical control, operating independently
of each other.

####### Instrument Internal Power

Usage and distribution of primary power shall be compatible with system
and subsystem Electromagnetic Compatibility (EMC) and magnetic field
performance requirements.

Secondary power distribution to power components shall be compatible
with system and subsystem EMC performance requirements.

####### Instrument External (Spacecraft) Power

Instruments shall be designed to operate from a 28 +6, -6 volt dc,
negative ground, unregulated power subsystem.

The spacecraft shall be able to remove bus power to all instruments if
the spacecraft power system indicates an emergency energy imbalance.
This (removal of power) will not apply to survival heaters unless
survival of the spacecraft is at stake.

All instrument components shall remain undamaged when subjected to
gradual ramp up monotonically (as low as 1 volt per minute) of voltage
in the range from 0-22 volts, for a duration not exceeding 10 seconds.

####### Unannounced Removal of Power

Excluding the thermal effects of removing instrument power, unannounced
removal of power shall not cause damage or degraded performance
(following re-application of power) to the instrument.

The exclusion refers specifically to the fact that unannounced removal
of ALL power (operational and survival) for an indefinite period could
have permanent detrimental thermal effects.

###### Electrical Power Interface Requirements

See the LDCM Environmental Verification Requirements document for
details on the operational bus electrical power interface environment.

####### Impedance

The power bus impedance will be documented in the ICD.

The Q and impedance of the input filters of the instrument components
will be documented in the ICD.

####### Survival Heater Bus

The spacecraft shall provide to the instruments power for survival
heaters, meeting the same voltage range requirements as the main 28 volt
bus.

Instrument design shall be such that having both primary and redundant
survival heater circuits enabled does not violate any thermal or power
requirement.

Note: Just because both survival circuits are enabled it does not mean
that both circuits are drawing current. Since either circuit has to be
able to meet the thermal requirements, if both circuits are drawing
current, the instantaneous power drawn would exceed the allowed maximum
by a factor of 2.

Survival power shall be used within the instrument only for resistive
heaters (and associated thermal control device) which maintain the
instrument at minimum turn-on temperature when the main power bus is
disconnected from the instrument.

Note: \"minimum turn-on\" refers to the temperature at which the
instrument can be sustained indefinitely without degradation in
performance, once operational power has been restored and a turn-on
sequence followed. The initial application of operational power by the
spacecraft (turn-on) can be made at this 'minimum turn-on' temperature.
Restoration of full operational status will likely involve a sequence of
events following that initial turn-on, and may involve intermediate
temperature constraints which must be observed.

The instrument contractor will identify (for documentation in the ICD)
if more than 2 circuits are required (1 primary and 1 redundant).

The instrument shall provide thermal control of the survival heaters
which is single-fault tolerant against both excessive and insufficient
application of heat, and document the extent of the redundancy in the
ICD.

The thermal control arrangement is intended to protect the spacecraft
from excessive current drain by the instrument if the control fails ON,
and to protect the instrument from becoming too cold if the control
fails OFF. Whether parallel redundancy, series redundancy, or both is
needed depends on the current drain, allowable instrument temperatures,
and other factors.

Survival heater circuit impedance will be documented in the ICD.

The survival heaters shall be functionally redundant.

Survival heater power buses shall be electrically isolated from each
other, from other instrument thermal control, from chassis, and shall
have independent power returns.

The spacecraft shall ensure that both the primary and redundant survival
heater circuits are normally enabled on-orbit when an instrument is off.

However, even the survival heaters may be turned off in the event of an
emergency where the survival of the spacecraft is in jeopardy.

######## Safety Bus Operation

The spacecraft will turn on the safety bus during those periods where it
is intended to be used for safety critical activities. The primary and
redundant safety bus power may or may not be on at the same time,
depending on spacecraft operations.

The instrument design shall not rely on presence of the Safety Bus at
any time other than for initial activation of safety-critical items.

##### Grounds, Returns, and References

###### Grounding Responsibility

The spacecraft shall employ a single-point-ground configuration.

The spacecraft contractor shall document the observatory grounding
scheme in the LDCM Observatory Electrical Requirements Document.

All interfaces from the instrument components to the spacecraft shall be
documented in the ICD.

###### External Ground Tie Point

Each instrument component will identify an external chassis ground tie
point to be used for external connections while the instrument is being
moved. This will be documented in the MID. This point may be the same
point used as the connection point to the spacecraft common ground path.

###### Thermal Blanket Grounding

All thermal blanket layers shall be grounded.

##### Electrical Harnesses and Connectors

###### Electrical Harnesses

####### Harness Wiring Requirements

Instrument power harnesses, within the instrument and within the
spacecraft harness, shall be appropriately sized to support the peak
allocated power levels and spacecraft fusing.

To the extent practical, power and signal circuits shall utilize
separate connectors.

Pyro power shall utilize connectors separate from those utilized for any
other purpose.

####### Harnesses Provider

All harnessing used on the spacecraft to connect the spacecraft to the
instrument will be provided by the instrument contractor.

Intra-instrument harnessing, exterior to the instrument housing and
connecting different parts of a single-unit instrument, or different
parts of a multiple-assembly instrument mounted on a single baseplate,
will be provided by the instrument contractor.

####### Harness Documentation

Harnesses, connectors, ground straps, and associated service loops will
be documented in the ICD, including all requirements for harness
construction, pin-to-pin wiring, cable type, etc.

####### Harness Tie Points

The provider and the locations of harness tie points will be an
agreement between the instrument contractor and the spacecraft
contractor, and documented in the ICD.

###### Electrical Connectors

####### General Considerations

Primary and redundant connectors shall be differentiated by clearly
marking all boxes and cables.

The instrument contractor will provide two sets of mated pairs of
interface connectors for each instrument.

The instrument contractor will provide static-discharging connector
covers, delivered in-place with the instrument.

The instrument interface mating connectors dimensions and locations will
be documented in the ICD.

For instrument connections to the spacecraft harness, there shall be
adequate clearance around the outside of mated connectors so that there
is no need for \"blind mate/demates\" or a need for removing adjacent
instruments in order to mate or demate a harness connection.

All spacecraft connectors for all instrument connections shall have
recessed contacts (e.g. female) to minimize inadvertent contact when the
instrument is not mated.

Provision of clearance around the outside of connectors is the
responsibility of the spacecraft contractor.

Captive covers shall be installed for all instrument connectors that are
not mated to harnesses or flight plugs.

All connectors that are not used for flight or for EMI/EMC testing shall
be covered with EMI tight covers.

####### Connector Location and Types

Connectors shall be located on the anti-cold-space side (sun side) of
the instrument, except in those cases where local conflicts exist on the
sun side, and that exception is documented in the ICD.

Connector locations, types, and orientations will be documented in the
ICD.

####### Keying

Connectors shall be different sizes, different types, different
orientation, color coded or uniquely keyed (in order of preference) to
the extent feasible in order to prevent improper connection.

####### Flight Plugs

Flight plugs requiring installation prior to launch shall be capable of
being installed at the Spacecraft level.

Flight plugs and their locations will be documented in the ICD.

####### Buffer Connectors and Connector Savers

Instrument buffer connectors and connector savers shall be utilized
prior to spacecraft-level system tests.

Note: The intention is to mitigate risk of bent pins or otherwise
damaged flight connectors.

####### Test Connectors:

Test point interface circuit and equipment failures shall not be capable
of propagating failures into the instrument. This includes credible
failures in GSE connected externally to the test point interface
connectors.

Captive flight quality and flight capable test connector covers shall be
installed whenever the test connector is not in use.

Test connectors and their locations will be documented in the ICD.

###### Breakout Boxes

Test-tees, interrupt boxes, and breakout boxes for
instrument-to-spacecraft interfaces will be provided by spacecraft
contractor.

#### Command and Data Handling

Unless specified otherwise, the requirements in this section shall apply
to all Command and Data Handling (C&DH) interfaces.

##### Electrical Interfaces

All electrical interfaces except test point interfaces shall be
functionally redundant.

The electrical interfaces (see Figures 3.3.3.2.1-1 and 3.3.3.2.1-2)
include the following. The requirements for each interface are
identified later in this Section:

> High Speed Science Data Bus (HSSDB)
>
> High Speed I-SSR Data Bus (HSIDB)
>
> 1553 Command and Data Bus
>
> 1 Hz Time of Day Pulse
>
> Pulse Command Interface
>
> Safety Bus
>
> Bilevel Telemetry Interface
>
> Passive Analog Telemetry Interface
>
> Operational Power Bus
>
> Survival Heater Power Bus
>
> Grounding Interface
>
> Test Point Interface.

![](media/image5.wmf){width="3.636220472440945in"
height="4.011023622047244in"}

[]{#__RefHeading___Toc151257677 .anchor}Figure 3.3.3.2.1-1. Spacecraft
to Instrument/Instrument to I-SSR Electrical Interfaces \[TBR\]

![](media/image6.wmf){width="3.260236220472441in"
height="3.260236220472441in"}

[]{#__RefHeading___Toc151257678 .anchor}Figure 3.3.3.2.1-2. Spacecraft
to I-SSR Electrical Interfaces \[TBR\]

###### Interface Fault Tolerance

A single fault on either the spacecraft side of the interface or the
instrument side of the interface shall not prevent the full and complete
digital data transfer needed to fulfill mission requirements.

##### Data Bus Requirements

###### Bus Functions

Spacecraft, Instruments and the I-SSR shall interface to the
MIL-STD-1553 data bus.

Instruments shall interface to the High Speed Science Data Bus (HSSDB).

The I-SSR shall interface to the High Speed Science Data Bus and the
High Speed I-SSR Data Bus (HSIDB).

The following functions, where appropriate, shall be provided between
the spacecraft and the instrument components (including the I-SSR).

a\) Spacecraft to instrument transfers consisting of:

\- real-time ground commands

\- stored commands

\- memory loads

\- state of health indications

\- ancillary data (e.g., time code, satellite ephemeris and attitude)

b\) Instrument to spacecraft transfers consisting of:

\- instrument health and status telemetry

\- instrument transition to safe mode indicator

\- instrument diagnostic data

\- memory dumps

\- survival mode temperatures (not part of the data bus)

c\) Instrument to I-SSR transfers consisting of:

> \- mission data, including ancillary data received from spacecraft

##### Interface Characteristics

###### High Speed Science Data Bus (HSSDB)

The HSSDB is defined as the set of interfaces from the OLI instrument
and the TIRS instrument to the I-SSR.

Instrument Mission Data shall be transferred to the I-SSR via the High
Speed Science Data Bus.

The details of the HSSDB will be documented in the LDCM Instruments to
I-SSR ICD.

The HSSDB may consist of a single or multiple physical interfaces, and
any necessary handshaking or flow control to manage the interface.

####### HSSDB Data Rate

TBD

####### HSSDB Physical Layer

TBD

####### HSSDB Data Format

The HSSDB shall include the instrument data and information on data type
and data timing (timestamp) as necessary to support file operations of
the I-SSR.

###### High Speed I-SSR Data Bus (HSIDB)

The HSIDB is defined as the interface from the I-SSR to the spacecraft
X-band system.

The HSIDB shall consist of three independent channels that each support
a realtime or playback data stream.

The HSIDB may consist of a single or multiple physical interfaces, and
any necessary handshaking or flow control to manage the interface.

The details of the HSIDB will be documented in the LDCM Spacecraft to
I-SSR ICD.

####### HSIDB Data Rate

The combined output rate of Mission Data and diagnostic data on the
Instrument to I-SSR interface shall not exceed the 230 Mbps (TBR) per
channel for each of the three independent output channels, including all
CCSDS packetization, framing, and coding overhead.

Note: When the data rate allocation in the LDCM Instruments to I-SSR ICD
is finalized it will supersede this number.

####### HSIDB Physical layer

TBD

####### HSIDB Data Format

The HSIDB data shall be compliant with CCSDS recommendation 131.0-B-1
Channel Access Data Units (CADUs).

The CADUs shall contain randomized codeblocks encoded with LDPC per
CCSDS experimental specification 131.1-O-1.

Convolutional coding shall not be applied.

###### MIL-STD-1553 Characteristics

####### Electrical Interface

The spacecraft contractor will define the bus implementation
characteristics in the LDCM Observatory 1553 Bus ICD.

Unspecified parameters will not be assumed or assigned by the instrument
contractor without concurrence from the spacecraft contractor, at which
time they are to be documented in the LDCM Observatory 1553 Bus ICD.

###### Pulse Commands

Pulse commands from the spacecraft shall not be used for functions which
could be accomplished by the instrument itself, or by command from the
spacecraft via the data bus.

The primary means of commanding the instrument is through the command &
data interface. If necessary, and as approved by the spacecraft
contractor and documented in the ICD, selected commands may utilize
pulse (relay driving) commands, having the characteristics given below.

Instrument discrete pulse commands shall have redundant interfaces per
function, one primary, and one redundant, each controlled by one wire
pair (command/return).

The spacecraft contractor shall define the pulse command characteristics
in the ICD.

###### Time of Day Pulse

The spacecraft shall provide a 1 Hz Time of Day pulse to indicate the
point in time at which to apply the time code which was previously
transmitted over the data bus.

The instrument shall utilize the rising edge of the Time of Day pulse,
together with the time code data, in order to establish the time
reference for instrument data.

The spacecraft contractor shall define the Time of Day pulse command
characteristics in the ICD.

The spacecraft shall issue the Time of Day pulse rising edge
(non-inverting side of differential interface) within 5 microseconds
(TBR) of each spacecraft 1-second time occurrence.

###### Synchronization

Where instruments require synchronization to each other that is more
precise than the onboard time correlation requirement, the spacecraft
shall provide a means whereby synchronization can be accomplished via
the data bus interface, as detailed in the ICD for the affected
instrument.

##### Instrument Commands and Data Load

The standard means of transmitting commands and data loads from the
spacecraft to the instrument shall be via the 1553 data bus.

The spacecraft shall, through the MIL-STD-1553 bus, deliver the
instrument command and data load to the specified instrument RT-receive
address/sub-address by conducting BC-to-RT Transfers or RT-to-RT
Transfers (from a spacecraft RT to an instrument RT).

The spacecraft shall provide discrete commands for instrument
operational and survival heater power switching, fault recovery, and
squib functions as necessary.

The spacecraft shall provide auxiliary data and spacecraft ephemeris and
attitude data, as appropriate and detailed in the ICD.

All commands to be transferred to the instrument via the point-to-point
(discrete) command interfaces and the command and telemetry bus, will be
documented in the ICD.

###### Commands

The instrument contractors will provide command information to the
spacecraft contractor for inclusion in the LDCM Observatory 1553 Bus
ICD.

Commands to instruments from the spacecraft will be documented in the
LDCM Observatory 1553 Bus ICD.

No single point failure shall prevent the instrument components from
receiving commands.

##### Instrument Health and Status Telemetry 

The instrument contractors will provide instrument health and status
telemetry information to the spacecraft contractor for inclusion in the
LDCM Observatory 1553 Bus ICD.

Instrument health and status telemetry will be documented in the LDCM
Observatory 1553 Bus ICD.

No single point failure shall prevent spacecraft access to critical
telemetry points.

###### Point-to-Point Telemetry

Point-to-point telemetry is intended to represent \"static\" information
(for example: relay status associated with pulse commands where
knowledge of the relay position is required even when the instrument is
off) such that sampling rates in the seconds, or tens of seconds, is
adequate.

All spacecraft-instrument point-to-point telemetry interfaces will be
documented in the ICD.

Instrument component point-to-point telemetry shall include redundant
passive analog temperature measurement devices at a number of locations
per instrument as documented in the ICD.

Redundancy may be accomplished by thermally overlapping regions for
non-critical measurements. These analog lines are separate and in
addition to any state of health (SOH) input being transmitted over the
command and data interface and are intended to provide insight during
periods when the instrument operational power is not present; therefore,
excitation of the passive analog temperature measurement devices is
provided by the spacecraft.

###### Command Verification

Receipt of individual commands via the Command & Data bus shall be
verifiable via instrument Health and Status telemetry.

Receipt of individual Pulse commands shall be directly verifiable via
Point-to-Point Telemetry or some other means.

###### Instrument Memory Dump

Instrument Memory Dump data shall be transferred to the spacecraft C&DH
via the 1553 data bus as specified in the ICD.

###### Telemetry Monitor

Any requirement for the spacecraft to monitor selected telemetry points
(including telemetry from the command and telemetry bus and
point-to-point interfaces) and initiate action to the instrument based
upon a pre-determined telemetry state will be negotiated with the
spacecraft contractor and documented in the ICD.

All instrument telemetry points to be monitored by the spacecraft, for
spacecraft action, will be documented in the ICD, along with the action
to be taken and the algorithm for such action, as specified by the
instrument contractor.

The Spacecraft Telemetry Monitor (TMON) function shall perform threshold
checking on specific data items as described in the ICD.

The Spacecraft Telemetry Monitor (TMON) function shall be capable of
activating a stored command sequence when errors are detected (i.e. a
defined threshold is exceeded).

The TMON data thresholds shall be defined in tables in the flight
software.

The TMON tables shall be modifiable by the ground via a Memory Load
operation.

The TMON function shall compare the collected data with the predefined
limits defined in the tables.

Any payload telemetry parameter that requires monitoring by the
spacecraft shall be placed in a TMON packet message.

The format description of this TMON packet shall be defined in the ICD.

> The instrument contractors shall provide to the spacecraft contractor
> the autonomous fault response algorithm for each TMON monitor point.

##### Command & Data Interface Test Packets

The instrument shall be capable of generating and transmitting, on
command, a continuous sequence of packets containing a fixed data
pattern.

Instrument test packets shall have a unique APID.

Test Packet APIDs, patterns and lengths will be documented in the ICD.

## Environmental Conditions

This section specifies the environment characteristics in the presence
of which the spacecraft and the instrument components (referred to as
space equipment in this section) must meet all other requirements.

### Radiation Environment

The space equipment shall be capable of meeting all performance
requirements for the ionization and displacement damage levels produced
by the natural radiation environments for their operational orbits for
the operational lifetime of the mission.

### Meteoroid and Debris Environments

Space equipment and materials that are directly exposed to free space,
such as cables, propulsion lines, pressurized tanks, and sensor optics,
shall be designed to meet reliability requirements to remain operable
within their performance specifications over the mission lifetime in the
meteoroid and space debris environments in the operational orbit.

### Spacecraft Magnetic Fields

The space equipment shall not exhibit any malfunction, degradation of
performance or deviation from the specified indications beyond the
tolerances indicated in their individual equipment specifications as a
result of being exposed to DC magnetic levels not exceeding the natural
Earth field at sea level. (TBR)

### Atomic Oxygen

The space equipment shall meet performance requirements during exposure
to atomic oxygen (AO) experienced over the mission lifetime in the
environments in the operational orbit.

### Spacecraft Charging from All Sources

The space equipment shall operate without performance degradation due to
the surface charging, bulk charging, and deep charging environment.

### Launch Environment 

The baseline LDCM launch vehicle is a Delta-II.

The LDCM spacecraft and instrument components shall be compatible with
the launch environments as specified in the LDCM Launch Vehicle ICD.

## DESIGN AND CONSTRUCTION

### Instrument Electrical Power Interface Design Requirements

The instruments shall have EMI input filters installed on the instrument
side of the power interface.

The filters shall provide both common-mode and differential-mode
filtering.

### Instrument Mechanical Interface Design Requirements

#### Component Stiffness

Except as permitted below, each separately-mounted instrument component,
configured for launch, shall have a fixed-base fundamental resonant mode
frequency of greater than 50 Hz (TBR).

Fixed-base is defined as follows: Each mounting point shall be
constrained in those translational degrees of freedom which are rigidly
attached to the spacecraft and shall be free in those translational
degrees of freedom for which kinematic mounts or flexures provide
flexibility.

The fundamental resonant mode frequency for an instrument in its
on-orbit configuration shall be 6 Hz (TBR) or greater.

#### Instrument Mechanisms

All instrument mechanisms which require restraint during launch shall be
caged during launch without requiring power to maintain the caged
condition.

Instrument mechanisms which require caging and/or uncaging during test
and launch site operations shall be capable of being caged or uncaged by
command and by manual actuation of accessible locking/unlocking devices.

Instrument mechanisms which require uncaging and/or caging on-orbit
shall be capable of being caged and uncaged by command.

#### Uncompensated Momentum

Each instrument having movable components shall not exceed an
uncompensated momentum contribution of +/- 0.5 N-m-sec (TBR) per axis.

The typical uncompensated momentum contribution of the instrument will
be documented in the instrument ICD.

#### Instrument Disturbance Allocations

Early estimates of gimbaled masses, inertias, and cg will be provided to
the spacecraft contractor to permit sizing the control components to
meet pointing and stability requirements.

##### Periodic Disturbance Torque Limits

The instrument contractor will provide, in electronic format,
time/magnitude plots of their disturbances, for decomposition by the
spacecraft contractor, of the sum of the instrument-induced periodic
disturbance torques, in order to produce the corresponding magnitude
spectrum.

The magnitude of the periodic disturbance torque, including the torque
resulting from linear forces reacting from the instrument to the
spacecraft) shall be in the acceptable range of Figure 3.5.2.4.1-1 for
all frequencies.

The instrument ICD will specify the predicted disturbance torque
contributions to the spacecraft, if any.

![](media/image2.wmf)

[]{#__RefHeading___Toc151257679 .anchor}Figure 3.5.2.4.1-1. Allowed
Transmitted Torque (TBR)

[]{#__RefHeading___Toc151257682 .anchor}Table 3.5.2.4.1-1. Allowable
Transmitted Torque Transition Points (TBR)

  --------------------------------------- -------------------------------
  **Frequency (rad/sec)**                 **Torque (N-m)**

  \< 0.001                                0.03

  0.01                                    0.003

  1.0                                     0.003

  \> 33.3                                 0.1
  --------------------------------------- -------------------------------

##### Constant Disturbance Torque Limits

Instrument-induced constant disturbances of the same polarity, separated
by more than 200 seconds, shall not exceed the torque limit defined in
Figure 3.5.2.4.2-1 if the duration of application is in excess of 10
seconds. For constant torques of 10 seconds duration or less, the
impulse limit is 0.04 N‑m‑s (TBR).

For constant torques of 400 seconds duration or more, the torque limit
is maintained at the 400 second limit shown in Figure 3.5.2.4.2-1.

![](media/image2.wmf)

[]{#__RefHeading___Toc151257680 .anchor}Figure 3.5.2.4.2-1. Constant
Torque vs. Duration of Application (TBR)

##### Disturbance Torque Limits for Linear Forces

The same requirements specified above for constant and periodic torques
shall apply when linear forces are converted to torques, assuming a
moment arm of 2 m for motion along the pitch or yaw axes and 3 m for
motion along the roll axis.

The torques about the spacecraft CM resulting from instrument linear
forces will be documented in the instrument ICD.

##### Spacecraft Non-Operational Deployment Disturbance

The on-orbit non-operational Spacecraft disturbances produced by the
uncaging or deployment of any mechanism or device shall be less than 158
N-m (TBR) at the point of attachment.

##### Torque Profile Documentation

The instrument contractor will provide the typical instrument torque
versus time profile to the spacecraft contractor for documentation in
the ICD.

##### Thrust Direction Definition

The instrument contractor will provide to the spacecraft contractor the
magnitude and direction of net thrust, timing of expulsion events, and
expulsion material resulting from the expulsion of expendables by the
instrument, for documentation in the ICD.

### Magnetics

The LDCM spacecraft and instruments shall minimize the use of magnetic
materials where possible.

If magnets are inherent to the instrument design, early estimates of
magnetic fields and residual magnetic dipole moments will be provided to
the spacecraft contractor for documentation in the ICD.

### Access

#### Access Identification

The instrument contractor will provide to the spacecraft contractor
access requirements for documentation in the MID.

### General Access

All items to be installed, removed, or replaced at the spacecraft level
shall be easily accessible without disassembly of the unit.

Access requirements and clearances will be mutually agreed-upon between
the spacecraft contractor and the instrument contractor.

### Mounting/Handling

#### Mounting Orientation

Instrument design shall be such that the instrument is capable of being
mounted to the spacecraft with the spacecraft interface in the
horizontal position.

#### Instrument-to-Spacecraft Integration and Test Mounting

Each separately-mountable instrument component shall be capable of being
installed and removed without removal of other instruments or
subsystems, and without demating or removal of harness (including coax)
other than for the unit being removed.

Connector access for removal and reassembly of an instrument shall be
accomplished without requiring removal of any equipment.

Instrument mounting hardware shall be installed from the instrument side
of the spacecraft - instrument mechanical interface.

Installation/removal of any separately-mountable component shall not
require rotation during installation/removal.

#### Non-Flight Equipment

The instrument contractor will provide to the spacecraft contractor
information on all items to be installed or removed prior to flight, for
identification in the ICD.

### Venting

Instrument contractors will provide to the spacecraft contractor the
location, size, path, venting forces and impulses, direction of net
thrust, and operation time of vents in the instruments for inclusion in
the ICD.

The spacecraft contractor shall place the instrument such that the
contamination products from the vents of one instrument cannot directly
impinge on another instrument\'s contamination-sensitive surface nor
directly enter another instrument\'s aperture.

### Contamination

The contamination interface requirements will be provided to the
spacecraft contractor for inclusion in the ICD.

This includes contamination sensitivity of the instrument, as well as
identification and characterization of all sources of contamination that
can be emitted from the instrument.

## DOCUMENTATION

### Interface Control Drawings and Documents

Each instrument developer is responsible for providing inputs to the
ICDs in accordance with contractual requirements.

After award of the spacecraft integration contract, each instrument
developer and the spacecraft contractor will jointly write the Interface
Control Documents (ICDs) which define the details of the
instrument-to-spacecraft interface and instrument accommodation
information.

The spacecraft contractor will be responsible for developing and
controlling the ICD except as noted otherwise.

Drawings defining/documenting the interface between the instrument and
the spacecraft which are not otherwise included in the ICD will be
contained in an MID controlled by the spacecraft contractor.

## Interface Hardware Providers

Interface hardware will be provided by the organization shown below.

  -------------------------------------------------------------------------
  **HARDWARE**                                               **PROVIDER**
  ---------------------------------------------------------- --------------
  Standard mounting hardware (e.g., std. fasteners, shims)   Spacecraft

  Special mounting hardware (instrument-unique)              Instrument

  Mounts (e.g., kinematic, flexure, etc.)                    Instrument

  Optically aligned component drill template                 Instrument

  Non-optically aligned component drill template             Instrument

  Instrument optical alignment target /cube and cover        Instrument

  Holding and lifting fixtures                               Instrument

  GSE - special, electrical and mechanical                   Instrument

  GSE - general, electrical and mechanical                   Spacecraft

  Shipping containers                                        Instrument

  Flight connectors both sides of the electrical interface   Instrument
  -------------------------------------------------------------------------