Section 12.0

Guidance, Navigation & Control

Samuel J. PlacanicaGN&C Lead Engineer

ST5 PDR June 19-20, 2001

5Space Technology

“Tomorrow’s Technology Today”GSFC

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Agenda

• Requirements

• Documentation

• Components

• Operational Modes

• Mission Planning

• Attitude Determination

• Dynamic Simulator

• Risk Mitigation

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General System Requirements (1 of 2)

Spin stabilized spacecraft control. (MRD10307000)

– Utilize attitude design techniques to maintain spin stabilization control

– Spin-to-transverse inertia ratio greater than 1.2

Provide ground-based attitude determination and maneuver planning capabilities. (MRD10308030)

– Utilize sun sensor and magnetometer data in conjunction with standard sun position and geomagnetic field (IGRF-2000) reference models

Provide the capability to perform spin axis attitude maneuvers and orbit adjustment maneuvers. (MRD10308000)

– Magnitude of attitude maneuvers calculated on the ground and based upon attitude determination results and mission planning

– Initiated by either real-time or stored program commands

– Uses the Cold Gas Micro Thruster System

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General System Requirements (2 of 2)

Implement a pre-programmed autonomous Sun Acquisition Mode (SAM). (MRD10305014)

– GN&C will provide to Flight Software an algorithm document which presents the SAM control equations

Passively control spacecraft nutation. (MRD10307020)

– Nutation will be dissipated using a ring-like shaped damper which will be fully filled with silicone oil

A Dynamic Simulator (DS) shall be designed to simulate a realistic orbital environment on the ground and emulate the functionality and performance needed for spacecraft level GN&C and science validation requirements. (MRD11100000)

– Based upon SMEX/Triana DS heritage

– Tenth order magnetic field model

– Includes science event modeling

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Specific Requirements (1 of 2)

Spin Rate Operation Range

– Perform autonomous SAM following launch vehicle release; spin rate greater than 20 rpm

– 20 rpm ± 10% following magnetometer boom and antenna deployments

SAM Performance

– Maintain the sun to within 5 degrees of the normal to the solar panels

– Spin rate knowledge: ± 10% (3 sigma)

– Spin axis pointing knowledge: ± 5 degree (3 sigma)

– Limit thruster actuation for any single SAM activity

Nominal Spacecraft Operation Performance

– Maintain spin axis to be within 5 deg (3 sigma) of the ecliptic pole

– Spin rate knowledge: ± 3% (3 sigma)

– Spin axis pointing knowledge: ± 1 deg (3 sigma)

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Specific Requirements (2 of 2)

Provide the capability for a full 540 deg attitude maneuver

– Includes margin for worst case precession maneuver scenario

Nutation Control

– Time constant less than 60 minutes

– Steady-state nutation angle less than 0.5 degrees

– Maximum nutation angle due to launch vehicle release is to be less than 10 degrees

Spacecraft relative separations between 100 and 1000 km at apogee

All spacecraft to have the same orbital period

No constellation station keeping is required

Comply with NASA Safety Standard 1740.14 regarding orbital debris

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Documentation

• GN&C Subsystem Specification ST5-495-020 Preliminary

• GN&C Algorithms Document ST5-495-056 Preliminary

• Magnetometer System Spec ST5-495-011 Baseline

• Magnetometer System SOW ST5-495-012 Baseline

• Magnetometer Mechanical ICD Draft

• Magnetometer Electrical ICD Draft

• Sun Sensor Specification ST5-495-034 Preliminary

• Sun Sensor Statement of Work ST5-495-035 Preliminary

• Sun Sensor ICD Draft

• Attitude Determination and Draft

Maneuver Planning Document

• Dynamic Simulator Users Guide Draft

• Nutation Damper Spec and ICD Draft

Document Number Status

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• 5/8 inch Damper Assembly GD-2035196 Draft

• 5/8 inch Tube GD-2035197 Draft

• Fill Adapter for 5/8 inch Tube GD-2035198 Draft

• Tube Clamp GC-2049776 Draft

• Test Stand GD-2049777 Draft

• Configuration Damper Fill GD-2049778 Draft

DrawingsNutation Damper Drawings Number Status

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Magnetometer

• Science-grade device provided by UCLA

• Component includes sensor head, electronics and interface cable

• Dynamic range

– 0 to 64,000 nT over two range bands

• Resolution

– 1 to 2 nT in 64,000 nT field

– 0.1 to 0.2 nT in 1000 nT field

• Sample rate

– 16 three-axis measurements per second

• Data will also be used for ground-based attitude determination activities

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Magnetometer

• Total mass: ~ 0.611 kg

– Electronics unit: ~ 0.220 kg

– Electronics chassis: ~ 0.250 kg

– Sensor head: ~ 0.075 kg

– Interface cable: ~ 0.066 kg

• Total power: 0.550 W

– Electronics unit: 0.500 W

– Sensor head: 0.050 W

• Volume/Length (cm)

– Electronics chassis : 10 x 10 x 12

– Sensor head: 4 x 4 x 6

– Interface cable: 100 max

• Thermal operating environment

– Electronics unit: -20 to +40 deg C

– Sensor head: -20 to +40 deg C

– Interface cable: -100 to +40 deg C

• Thermal survival environment

– Electronics unit: -40 to +50 deg C

– Sensor head: -40 to +50 deg C

– Interface cable: -130 to +80 deg C

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Sun Sensor

• Accuracy: ± 0.25 deg

• Resolution: ± 0.125 deg

• Volume (estimated): 74 cm3

• Mass (estimated): 0.16 kg

• Power (estimated): 0.130 W

• Thermal environment

– Operating: -20 to +50 deg C

– Survival: -40 to + 60 deg C

• Radiation

– 100 Krad (Si) Total dose, SEU and latch-up immune

• Manufactured by Adcole Corp.

• 4 π steradian field of view (spinning)

• Will provide sun elevation and sun pulse data

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Nutation Damper (1 of 2)

• Passive device will damp nutation induced by both launch vehicle release and thruster firings

• GSFC in-house design and fabrication

• Fully-filled with silicone oil

• Will be mounted inside spacecraft along a wall

• Performance and environmental testing will be performed

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Nutation Damper (2 of 2)

• Size: 18 cm x 18 cm

• Tubing material: Aluminum Alloy 6061-T6

– 0.625 inch outer diameter

– 0.035 inch wall thickness

• Fill adapter material: Aluminum Alloy 6061-T6

• Fluid: Fully filled with Dow Corning 200 Silicone

– 5 centistoke viscosity at 25 deg C

• Mass (best current estimate): 0.2533 kg

• Power: None

• Thermal

– Operating: -20 to +50 deg C

– Survival: -40 to +60 deg C

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Operational Modes (1 of 2)

• Standby Mode

– Used during nominal spacecraft operations

– No thruster activity

• Sun Acquisition Mode (SAM)

– Can remain in mode for indefinite amount of time

– Uses on-board flight software to process sun sensor data

– SAM control logic will issue the appropriate commands to the cold gas thruster to orient spacecraft into a power positive attitude

– Activation methods:

• Autonomously entered following launch vehicle release

• On cold CPU reset, initialization will boot spacecraft CPU into SAM

• Ground command

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Operational Modes (2 of 2)

• Maneuver Mode

– Spacecraft spin axis is processed to a pre-determined orientation

– Required prior to a delta V in order to orient the spacecraft spin axis along the velocity vector

– Ground-based processing of sun sensor and magnetometer data results in thruster fire commands which are uploaded to the spacecraft and executed in open-loop fashion

• Delta V Mode

– Required to maneuver the three spacecraft into a mission orbit constellation in order to demonstrate inter-spacecraft communications/crosslink capabilities

– Ground command

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System Block Diagram

Sun Sensor

C&DH Card

Cold Gas

Micro

Thruster System

Thruster

Commands

Magnetometer

Sun Sensor

And

Mag Telemetry

Ground Station

Attitude/Orbit

Determination

Develop

Thruster Cmds

SAM FSW

Telemetry

Processing

Standby Mode

SAM

Maneuver &Delta V Modes

No Command

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Rhumb Line Precession (1 of 2)

InitialAttitude

FinalAttitude

Sun

InitialSpin Plane

RhumbLine

Y-Axis

X-Axis

Z-Axis

i

f

f

First used thirty years ago on Early Bird and still frequently employed to reorient spinning spacecraft.

A constant heading angle is maintained throughout the maneuver.

The heading angle is computed based on the known initial orientation and the desired final orientation.

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Rhumb Line Precession (2 of 2)

• Heading angle is used to compute a phase angle from the sun presence signal

• Phase angle is used with the estimated spin rate to compute a time delay between sun presence and the thruster pulse

Orientation in the Spin Plane at Sun Presence Orientation in the Spin Plane at Thruster Pulse

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Thruster and C&DH Interface

DigitalSun Sensor

Sun Elevation

Sun PresenceClock

Sun Presence Time

ThrusterDelay Time

ThrusterPulse Width

ThrusterEnable

FlightSoftware

C&DHSun

Presence

ThrusterDelay Time

ThrusterPulse Width

SunPresence

Spacecraft Spin Period

0

1Time

• Flight software sends the Delay Time, Pulse Width and Enable flag to C&DH

• C&DH produces the thruster commands in the form of a pulse train that repeats at the spin period in Sun Acquisition and Maneuver Modes and at 2 Hz in Delta V Mode

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Sun Acquisition Mode Block Diagram

FilterSun Elevation

ThresholdComparison

With HysterisisThruster Enable Flag

If pos then = 270°If neg then = 90°

Sun Presence Time Spin RateEstimation

ThrusterDelay

CalculationThruster Delay Time

ComputeVariance

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Mission Planning

• Constellation - Formation defined as inter-spacecraft distances greater than 100 km and less than 1000 km at apogee

• Lifetime - 3 months, 6 month goal

• On-board maneuvers to final configuration

• Orbit Debris mitigation - NASA NSS-1740.14

• Orbit Determination - 10 km knowledge

• All spacecraft in same orbit plane, identical periods

• Delta V required to deploy from geosynchronous transfer orbit no greater than 1.6 m/s

• No provision for orbit maintenance

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Orbit Characteristics

• Assumptions

– Launch November 2003

– Nominal Orbit: from empirical survey of launch vehicle histories

• 200 x 38,000 km altitude (Period: 10.5 hours)

• 28 deg inclination

• perigee at descending node (typical LV injection for GTO)

• apogee sun side (noon)

• Eclipse History

– 30 minutes max eclipse, 143 days from launch

• Lifetime - nominal scenario

– +2 sigma solar flux = 1.1 years, meets 6 month goal

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Lifetime Estimate for Initial 200 x 38,000 km Orbit

ST-5 Lifetime HistoryNov 2003 Launch, 2s solar flux (09/00 Schatten),

M=25 kg, A=0.2 m2

0

5000

10000

15000

20000

25000

30000

35000

40000

Nov-03 Feb-04 Jun-04 Sep-04 Dec-04 Mar-05 Jul-05 Oct-05 Jan-06 May-06

Date

Ap

og

ee (

km)

Nominal 25 kg

30 kg

20 kg

25 kg 30 kg20 kg

Preliminary

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Lifetime Estimate - Perigee Variation

ST-5 Lifetime HistoryNov 2003 Launch, 2s solar flux (09/00 Schatten),

M=25 kg, A=0.2 m2

0

5000

10000

15000

20000

25000

30000

35000

40000

Nov-03 Mar-05 Aug-06 Dec-07 May-09 Sep-10 Feb-12 Jun-13

Date

Ap

og

ee (

km)

250 km

225 km

215 km

200 km

200 km

215 km 225 km

250 km

Preliminary

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Attitude Determination

• Shall accurately determine rate and attitude of the spacecraft for both the normal on-orbit mode and the orbit and attitude adjustment modes

• The ground-based Attitude Determination System (ADS) will reuse components from the MultiMission Spin-Axis Support System (MSASS)

• MSASS is a Matlab-based attitude determination system which has been used to support several GSFC missions, including FAST, Wind, and Polar

• ADS will utilize magnetometer and sun sensor telemetry data

• GN&C will provide hardware and analysis tools to support two redundant strings in the mission operation control center

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Dynamic Simulator (1 of 2)

• Simulator core simulates single spacecraft

• Simulator core code is hosted in a single PC-compatible

• Simulator core code is a ‘C’ program running under Linux

• Interface is through commercial PCI plug-in cards (via custom hardware as needed)

• ASCII text configuration files are used extensively to allow rapid re-configuration as needed

• Internal data is stored in a mnemonic-driven database

• Simulator core is derived from SMEX/Triana heritage code (especially Triana and FAST)

• One copy of simulator to be built for each of three test environments (Flight S/W, C&DH S/W, and FlatSat – no testing at spacecraft level intended)

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Dynamic Simulator (2 of 2)

• Simulator can be configured for “open loop” operation (i.e., sensor data is static regardless of actuators and dynamics, and interfaces can be tested in this fashion)

• Simulator can be configured for “mixed loop” operation (i.e., sensor data reflects the response of the dynamics to constant actuation regardless of actual commands, and interfaces can be tested in this fashion)

• Simulator can be configured for “closed loop” operation (i.e., sensor data reflects the response of the dynamics to the actuator commands received)

• Regardless of loop configuration, individual sensors and actuators can be disconnected from the simulator and actual hardware substituted (it will not then be “in the loop”)

• No simulation of GPS or inter-spacecraft signals intended

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Risk Mitigation

• Risk: Magnetometer electronics parts availability

- Mitigation: GSFC will assist UCLA in parts procurement

• Risk: Sun sensor shock environment

- Mitigation: Adcole will perform analysis and redesign to strengthen the sun sensor in the areas of recticle hold-down and recticle/solar cell separation

• Risk: Rapidly decaying orbit

- Mitigation:

• Provide larger delta V capacity

• Acquire higher perigee altitude due to launch vehicle insertion

• Accommodate through system-level design