chart 1 adcs design & hardwareme176: lecture 5 aaron rogers [email protected]

ADCS Design & Hardware ME176: Lecture 5 Chart 1

ME176: (Space!) Machine DesignADCS Design & HardwareADCS Design & Hardware

February 20th, 2003

Aaron [email protected]


Introductions and Overview


GNC-078

XECI

YECI

ZECI

line ofnodes

orbitnormalEarth's

equatorialplane

i

orbit

= right ascension of

the ascending node

i = inclination

Review of Last Section: Orbits


ascendingnode

descendingnode satellite's

position atepoch

rpra

ro

= argument of perigee

= true anomaly at epoch

rp = perigee radius

ra = apogee radius

a = semi-major axis = (ra + rp)/2

e = eccentricity = (ra - rp)/(ra + rp)

GNC-079

Earth

line of nodes

Review of Last Section: Orbits Cont.


R'

At T0=0 min, R' = RD + Rwhere RD=1.5km (nom. baseline),R > 300km (20% baseline)

V1 @ T0: Initiate Transfer OrbitV2 @ TF: Circularize Into Target Orbit

T ~ 24 min T ~ 71 min

Transfer Orbit:Apogee Alt=500km – Alt

At TF~95 min, R' = RD

Review of Last Section: Hohmann Transfer


Commercial Satellites at GEO


• 1 - Introduction• 2 - Propulsion & ∆V• 3 - Launch Vehicles• 4 - Orbits & Orbit Determination• 5 – Attitude Determination and

Control Sys. Design & Hardware– Coordinate Frames– Attitude Determination– Environmental Disturbances– Vehicle Stabilization Methods– Attitude Control– Control System Design– Assembly, Integration & Test– Simplifying ACS

• 6 - Power & Mechanisms

• 7 - Radio & Comms• 8 - Thermal / Mechanical

Design. FEA• 9 - Reliability• 10 - Digital & Software• 11 - Project Management

Cost / Schedule• 12 - Getting Designs Done• 13 - Design Presentations

Sporadic Events: •Mixers •Guest Speakers •Working on Designs •Teleconferencing

(Re) Orientation


Homework Questions for 2/27For your selected Mission:• Pick two attitude control

approaches that might work

• List the sensors and actuators necessary to implement each of them. How accurate / sensitive would each have to be? Any other special requirements ($, mass, volume, power, bandwidth etc.)

• Locate them on a “generic” spacecraft (e.g. a cube or faceted sphere

• Make a block diagram of the feedback control system you envision

• Pick your favorite of the two, and tell me why it’s your favorite (compare $, mass, complexity, performance…)

Comments:

• Eg. Gravity Gradient, Thompson Spin & TS with momentum storage are options for an earth pointer.

• Also search Web to locate actual components that might be candidates - not selections, but possibilities.

• ACS thrusters want to be in pairs and far from the CG. Torque coils don’t care. Sensors have to have a clear view out etc.

• What is the plant, what are the sensors, what is the actuator suite? What are the “set points?”

• Make a “trade table” listing specs / attributes of each to justify your selection.


For your Favorite of the two:

• List the ACS modes and the triggers to proceed from one to another. Diagram with a flow chart.

• Suggest a simple algorithm for your mission mode. Model it on Excel and show that it has a prayer of working

• List requirements your selected ACS imposes on the spacecraft

• List a candidate component suite and estimate the cost and labor to design, build & test.

Comments:

• Modes might include sleep, initial rate killer, sun or earth finder, rough point, tight point, and safe/hold.

• For instance, measure an error angle or rate and actuate something to reduce that error.

• For example: mass distribution, symmetry, power, siting, computation, magnetic / electromagnetic cleanliness

• Assume an engineering year costs $200,000 including the tools and toys necessary to play with ACS in the lab.

Homework Questions Cont. for 2/27


Reading Suggestions for 2/27–Power:

• SMAD Chapter 11.4

• TLOM Chapter 13, 14

–Mechanisms (SMAD):• Chapter 11.6

–Extra ADCS Review (SMAD):• Chapt. 6.2: Orbit Perturbations

• Chapt. 6.2: Orbit Maintenance

• Chapt. 11.1: ADCS

–Extra ADCS Review (TLOM):• Chapters 6, 11


TT&C RFTT&C RF

TT&CBaseband

AttitudeAttitudeControlControl

AJTs LAE

Pyro Pyro PowerPowerPyro FirePyro Fire

RF, CMD, TLMRanging

Baseband CMD/TLM

MIL-STD 1553 Data BusTLM WordsAccepted

Commands

PropulsionPropulsion

REAs

TT&C Antennas

Pyro Control

Wheel Wheel ControlControl

70 VDC

Electrical PowerElectrical Power

Power BusPower Bus

ArcjetsSolar

Arrays

Deployment Mechanisms

Batteries

Power Regulation Unit (PRU)

Fuse Box

– Tanks– Valves– Lines

Mechanisms

Structure

ThermalControl- Heat Pipes- Heaters- OSRs

Pyro Relays

OBCs UDUP/L

RIUs

Wheels

SSA

IMU

RF Couplers

Receiver/Transmitter

ToToPayloadPayload

ESA

Bus RIUs

Spacecraft Bus Block Diagram


GN&C Coordinate Frames• Body frame

– Fixed in & rotates with the spacecraft– Reference for sensor & actuator alignments – Reference for control torque calculations

• Earth-Centered Inertial (ECI) frame– Constant orientation in inertial space– Used to define spacecraft & sun ephemeris for attitude determination

• Orbital frame– Earth-oriented coordinate frame defines nominal attitude– Orientation in space depends on spacecraft’s orbital location

• Target frame– The frame control system aligns the body frame with– Defined with commanded offsets relative to orbital frame


Body Frame• Coordinate system origin

– Geometric center of separation plane

• Xb-axis (yaw)– Perpendicular to the separation plane– Points away from the center of the

spacecraft

• Yb-axis (roll)– Perpendicular to the E & W panels– Points toward the east panel

• Zb-axis (pitch)– Perpendicular to the north & south

panels– Points toward the north panel

XB(Yaw)

YB (Roll)

(Pitch) ZB

EastPanel

NorthPanel

SeparationPlane

GNC-056


Earth-Centered Inertial (ECI) Frame• XECI

– Parallel to the intersection of Earth’s equatorial plane and the ecliptic plane

– Positive axis points toward the sun at the vernal equinox

• ZECI – Parallel to Earth’s polar axis– Positive axis points north

• YECI – Completes the right-handed triad


ECI Frame (Continued)

XECI

ZECI

YECI

sun

XECI

ZECI

YECI

XECI

ZECI

YECI

XECI

ZECI

YECIfirst dayof autumn

first dayof spring

first dayof winter

first dayof summer

(Seasons are for the Northern Hemisphere) GNC-050


ECI Frame & Orbital Frame

ZECI (north)

Orbit Normal

+Zo (pitch)

Direction of Orbit

XECI

+Yo (roll) +Xo(yaw)

YECI

Equator

GNC-020

• Yaw is toward zenith (straight up from Earth)

• Pitch is perpendicular to the orbit plane

• Roll is perpendicular to yaw and pitch and points in the direction that the satellite is moving


Target Frame–During normal operations, the

spacecraft body axes are controlled to the target coordinate frame

–The orientation of this frame relative to the orbital frame is defined by the enabled pointing offsets

• Constant• Earth-target• Fourier


Attitude Determination: The Problem

Where am I looking in space?!


Attitude Determination: Magnetometers

• Magnetometer measures applied magnetic field, outputs two or three magnitudes: B= [X, Y, Z].

• With known orbit model (IGRF2000) and ephemeris, can calculate attitude by comparing measured vs. expected field direction.

• Low cost and low power, though does require some EMI isolation.


ZSD

-

XSD

YSD

GNC-075

Sun lineof sight

Sun lineprojected on theXSD–ZSD plane

is the true detector elevation angle

is the measured angle

Sun lineprojected on theYSD–ZSD plane

Attitude Determination: Sun Sensors


SSA Mounting & Field of View

+X (Yaw Axis)

+Y (Roll Axis)

Ea

st F

ace

Earth FaceW

est

Fa

ce

SSAboresight

49.55°

49.55°35°

GNC-027

+Pitch

Sunline

AzimuthAngle

ElevationAngle

Projection of sunlineon yaw-roll plane

SSA

bore

sigh

t


SSA Problems: Earth Albedo• Problem: Earth Albedo at Low Altitude

– The SSA sun detection threshold is 20% of the nominal solar intensity

– At low altitude, Earth albedo can be 40% as bright as the sun– Albedo can trigger a false sun-presence indication and cause

erroneous sun azimuth and elevation readings

• Conditions that can cause the problem:– Transfer orbit perigee altitude below 5000 km– Perigee on Earth’s sunlight side

• Solution (GEO spacecraft)– Suspend use of sun sensor data when the spacecraft altitude is

below 5000 km.


SSA Problems: Moon Interference

• Problem: Moon Interference (Partial Solar Eclipse)

– The SSA’s detection threshold is 20% of the nominal solar intensity

• The SSA will detect the sun during a partial solar eclipse

– During a partial eclipse, the centroid of the visible solar crescent is offset from the sun’s true centroid

• This produces a 0.1° to 0.2 ° error in the measured sun angle

• Note– The sun's visible surface has an angular

diameter of 0.53 deg. as seen from Earth

GNC-077

MOON

SUN


Attitude Determination: Earth Sensors

Scan Mirror

Positive Pitch

PositiveRoll

North

North Scan

South ScanSenso

r Null A

xis

3.82±0.1°

East

3.82±0.1°

GNC-003

• Earth sensor assembly (ESA) provides Roll & Pitch attitude data

• Used to update inertial attitude reference– Data used indirectly in a highly

filtered form during normal operations

– Data used directly with little filtering during Earth acquisition


North

East

±15°SensorScan

Earth

No Roll or Pitch Offset

North

East

CenterReference

Pulse

Earth

Positive Pitch Offset

GNC-001

ESA Scans With a Pitch Offset

Pitch is determined from the offset between the center reference pulse and the center of the Earth.


North

East

±15°SensorScan

Earth

No Roll or Pitch Offset

North

East

Earth

Positive Roll Offset

GNC-002

ESA Scans With a Roll Offset

Roll is determined from the difference between the lengths of the north and south

scans across Earth.


ESA Problems: Multiple Targets• ESA response to this

condition:–Detects two targets in

the south scan–Automatically inhibits

the south scan–Uses north scan for

pitch angle calculations–Uses north scan and

standard chord for roll angle calculations

–Outputs sun presence bit = 1 (multiple targets detected)

Earth

Sun(3° effectivediameter) GNC-031

ESAscans


ESA Problems: Non-Distinct Targets

• ESA response to this condition:

– Detects only one target in the south scan

– Continues to use the south scan (scan is not inhibited)

– Outputs erroneous pitch and roll angles

– Outputs sun presence bit = 0 (only one target detected)

Earth

Sun(3° effectivediameter) GNC-032

ESAscans


Attitude Determination: Star Trackers

• Utilizes a light sensitive medium (CMOS, CCD)

• Pattern recognition of detected images against internal star catalog

• Acquisition, track modes• Extremely high precision

(typically high cost)• Sensitive to stray light (baffles)

Roll, Pitch, and YawAttitude(x, y, z)

Processor

Image

StarCatalog

Pattern Recognition

Software

Pinhole Lens

Active Pixel CMOS Imager

30° Field of

View


Attitude Determination: Propagation• Data output from the IMU CPU:

– The sampled angular outputs developed by each gyro

– The sampled acceleration outputs developed by each accelerometer

• Data is typically only available during significant orbit adjust maneuvers!

– Integrates linear and angular rates in order to propagate state vectors

– Typically operates much faster than sensor measurements are taken

– Important when attitude update is not available (e.g. no sun).


IMU Functional Block Diagram

SensorProcessor

and I/OElectronics

1

PowerSupply

1

Input Power

GyroA

InterfaceElectronics

Accelerometer1

GyroB


GyroC


Accelerometer2

GyroD


SensorProcessor

and I/OElectronics

2

PowerSupply

2

Input Power

Relay Commands

Sensor outputs, ratemode commands, and

status (1553B data bus)

Sensor outputs, ratemode commands, and

status (1553B data bus)

Sensors

Sensors

GNC-029


Gyro & Accelerometer Alignments• Gyro A

– Positive sense axis is in the Y-Z plane, offset 125.3 from the +Z axis

• Gyros B, and C– Positive sense axes are offset

125.3 from the +Z axis

– Projections of the sense axes on the X-Y plane are 60 from the +Y axis (the positive roll axis)

• Gyro D– Senses rotation about the +Z axis

(the positive pitch axis)

• Accelerometers– Positive sense axes point in the +X

direction

GNC-033

+X(yaw)

+Y(roll)

+Z(pitch)

GyroD

GyroA

GyroC

GyroB


Attitude Determination Options

Option AccuracyOperational Flexibility

RecurringCost

ADCS Complexity

Design Impact

Development Risk

GPS + INS < 1 deg Low Low Low Low Medium

Sun Sensor + Magnetometer

< 5 deg Medium Low Medium MediumMedium

(Software)

Sun Sensor + Horizon Sensor

< 0.5 deg Medium Medium Medium LowMedium

(Software)

Star Tracker < 0.1 deg High High Low High Low

AeroAstro MiniStar Tracker

< 0.5 deg High Low Low LowMedium

(Software)


Environmental Disturbances• Atmospheric Drag (LEO)

– Function of Ballistic Coefficient, Altitude

• Solar radiation– Function of Surface Area, |CG – CSP|– Produces torque about all three axes– Varies with season and time of day

• Payload transmissions (recoil effect)– Primarily a pitch torque

Total EnvironmentalDisturbance Torque


Environmental Disturbances Cont.• Geomagnetic field (compass

needle effect)– Due to residual uncompensated

dipole, varies with R-3

– Primarily a yaw and roll disturbance

– Pitch torque produced only when solar storms temporarily distort the geomagnetic field

• Gravity Gradient (LEO)– Due to asymmetric mass

distribution

– Torques about pitch and roll axes

– Function off-nadir angle (theta), R-3

• Thermal radiation (recoil effect)– A function of the heat radiated

from various spacecraft surfaces

GravityGradientTorque

GeomagneticTorque


Environmental Disturbances Cont.• Solar wind (flow of charged particles)

– Very small effect– Earth’s magnetosphere deflects the solar wind before it

strikes the spacecraft• Force can increase temporarily (for a few hours) during strong solar

storms that distort the magnetosphere• Worst-case is still a small effect

• Micrometeoroid impact– Occasional small events (several times a year)– Most impacting particles are so small that the effects are

barely noticeable– Angular impulse almost always <0.5 in.lb.sec (<0.06 Nms)


Anarchy Happens: Unstabilized

Pros• Cheap, Simple, Reliable

• Can still determine attitude

Cons• Complicates

Radio Antennas

• Many missions impossible (eg imaging)

• How to ensure thermal balance?

• How to guard against spin-up?


Passive Stabilization: Gravity Gradient

• Cheap, Simple, Reliable

• Can still determine attitude

• Typical pointing performance:±5°

Pros

Cons• Complicates

Radio Antennas

• Many missions impossible (eg imaging)

• How to ensure thermal balance?

• Major deployable

• How simple & cheap is it?

• Very weakGG Torque = 32∆I = 3 x (2π/6000s) 2 x 1 kg-m2

= 3x10-6 N-m = 2 millionths of a foot pound


Passive Stabilization: Permanent Magnet

Pros• Cheap,

Simple, Reliable

• Pointing sideways often handy

• Passive yet strong

Cons• No yaw

control

• Flip 2x per orbit at poles

• Damping?

• Pointing typically ±5°


Passive Stabilization: Aerodynamic

Pros• Only simple

way to detect & point straight ahead

• Pointing sideways often handy

• Passive yet strong

• Pointing typically ±3°

Cons• Yaw damped

but not controlled

• Narrow altitude range => short lifetime

• May require deployables

• Damping may be necessary

Aero Torque = 1/2rAV2(cp-cg)

= 1/2 x 10-10kg/m3 x 1m2 x 70002 m2/s2 x 1 m (values @ 300 km)

= 2.5 x 10-3 N-m = 1.9 thousandths of a foot- pound


Spin-Stabilization

Pros• Disturbances

cancel / avg out

• Easy attitude determination

• Thermal rotisserie

• Typical pointing performance ±2°

Cons• Sensor

deconvolution

• Only one locale nadir pointing

• CG Control

A spinning body subjected to a torque impulse will precess its spin axis and otherwise will be unaffected

A non-spinning body subjected to a torque impulse will begin to tumble and continue doing so.


Spin Stabilization and Mom. of Inertia

Neutrally Stable

Unstable

AbsolutelyStable

IX < IY, IZ

IY < IX < IZ

IX > IY, IZ

Principal Axis

Intermediate Axis

For stable, minimum ACS design, SS prefers rotation about the principle axis aligned along the thrust vector, PX, where in general: IX > 2*(IY, IZ).

Rotation is possible, with nutation damping, about minimum axis, where:PX, where in general: IX < IY, IZ.

Minimum Axis


Thompson (non) Spinner

Pros• Disturbances &

thermal loads cancel / avg out

• Inherently stabile

• Antennas broadside to earth (+ 3 dB)

• No moving parts

• Scan pattern for sensors

• Whole earth nadir pointing spinner

Cons• Solar panel

usage (1/π)

• Non-spinner requires single mo. wheel

• CG Control

Non-spinner can stare and track subsatellite and lateral to subsatellite points

Momentum Wheel (non-spinner)


Pros• Huge electric power

gen.

• Stabile thermal / illumination environment

• High performance at low cost

• Pointing accuracy 0.2°

• Pointing knowledge 0.05°

Cons• Roll angle

hard to determine

• Attitude solution in umbra requires filter

• CG critical - difficult with deployables

Magnetometer

Fine Sun Sensors (2)

Z-coil

X-coil

Y-coil

Coarse SunSensor (12)

Flight Computer

Coil Driver

Horizon CrossingIndicator (HCI)

Non-Spinner: add just one wheel.

Q: On which axis?

Sun (non) Spinner


Three-Axis Stabilized

Pros• Arbitrary

pointing & staring

• Simple propulsion for station keeping

• Mass distribution not critical

Cons• Difficult

thermal control & power generation

• High power required

• Cost, mass & complexity

• Spin-up• Wheel

control• Lost wheel

• Torque noise

4 wheels divide three axes


Vehicle Stabilization Methods

Option AccuracyOperational Flexibility

Design Cost

ComplexityDevelopment

Risk

Passive Stabilization

Low Low Low Low Low• Magnetic• Atmospheric

Spin-Stabilized Medium Low Medium Low Low

3-Axis Control High High Medium Medium Medium


Spin vs. 3-Axis StabilizationRank Parameter Spin Stabilization 3-Axis Stabilization (Thrusters)

1Disturbance

Rejection

Directly proportional to stack MOI and spin-rate. High rpm might constrain AD, OD, and ConOps. Will require propellant for spin/de-spin.

Requires appropriate sizing of thrusters and propellant.

2Sensitivity to Stack Moments of Inertia

Large dependency on payload mass properties (MP). Will likely require trim mass, frequent measurements of stack MP, and update of spin rate.

ACS can accommodate spacecraft plant through modification of software.

3Thrust

MisalignmentSee (1). Commanded thrust will precess and nutate spacecraft attitude.

See (1).

4 Slewing Requires “turning” stiff rotation vector. Easily accomplished; see (1).

5 PointingInertial pointing only; accuracy highly dependent upon MOI and spin rate (2).

Capable of high accuracy, tracking, offset commanding.

6 C.G. ManagementCG migration measurement must be extracted from off-axis rotation rates and may require analysis/update of spin rate (2).

CG migration can be determined from system performance and is easily accommodated; see (1).

7Propellant

Management

Inherently settles propellant for primary orbit adjust, but may inhibit fuel flow near end of life.

May require a short settling burn before start of primary ignition.

8 Slosh High spin rates can reduce slosh effects.Slosh effects handled through software and filter design.


Attitude Control H/W: Torque Rods/Coils

• A torquer consists of a coil (or two redundant coils) around a soft iron core.

• Coil magnetizes the iron core– Long, slender core magnetizes

easier and more uniformly than a short, “fat” bar.

– Longer bar uses less power and coil mass.

• Subject to hysteresis saturation effects.

• Typically with wheels and/or gravity gradient booms.


Attitude Control H/W: Reaction Wheels

– The positive momentum/torque axes are all 45 from the pitch axis

– The projections of the momentum/torque axes onto the yaw/roll plane are all 45 from the yaw and roll axes

GNC-036

Yaw

Roll

Pitch

RWA 1

RWA 2

RWA 3

RWA 4


Reaction Wheel vs. Momentum Wheel

Reaction Wheel• Bi-directional• Operates over a

large speed range (positive & negative)

• Generates torque by controlling motor current

Momentum Wheel• Unidirectional• Operates in a narrow

range about a high nominal speed

• Torque depends on difference between commanded speed & current speed

Mechanically, there is no difference between a reaction wheel and a momentum wheel.


Typical Pitch Momentum Profile.

0 1 2 3 4 5 6Elapsed Time (Days)

Pit

ch M

om

entu

m

GNC-041

– Daily oscillation is due mainly to solar radiation pressure

– Long-term slope is due mainly to payload transmissions


Yaw/Roll Momentum Interchange• If there are no disturbances

– The angular momentum vector has a constant magnitude and a constant direction in space

• The spacecraft rotates once per orbit– Necessary for the payload to continuously face Earth

• As viewed from the spacecraft coordinate system– Angular momentum is exchanged between the yaw and

roll axes– The wheel speeds vary once per orbit, with the yaw and

roll momentum 90 out of phase


Momentum Interchange (Continued)

GNC-042

Earth

yaw

roll

yaw

roll

yawroll

yaw

rollangular

momentum

angularmomentum

angularmomentum

angularmomentum

• If there are no disturbances, the angular momentum vector has a constant magnitude and a constant direction in space.

• If there are no disturbances, the angular momentum vector has a constant magnitude and a constant direction in space.


Typical Roll & Yaw Momentum Profiles

.

0

0 1 2 3 4 5 6Elapsed Time (days)

Yaw

& R

oll

Mo

men

tum

RollYaw

GNC-043


Attitude Control H/W: Thrusters

Want 6-DOF Control


Attitude Control Options

OptionTorque

AuthorityPointingAccuracy

Recurring Cost

Disturbance Rejection

DevelopmentRisk

Gravity Gradient Low Low Low Low Low

Momentum Bias, 1 Reaction Wheel

Assembly (RWA)

High, Single-Axis

Medium Medium Medium Medium

RWA (3 or 4)High, All-

AxesHigh High Low Medium

Magnetic TorquersLow, All-

AxesLow Low Low Low

ThrustersHigh, All-

AxesMedium Medium High Medium

Thrusters + RWAsHigh, All-

AxesHigh High High High


GN&C Block DiagramGN&C Block Diagram

SensorProcessing

Logic

RedundancyMgt. Logic

ControlMode

SpecificLogic

AutomaticSwitching Logic

Attitude Determination and Ephemeris Propagation Logic

P I DController

ODDS LOGIC

Timed Pulse

MomentumMgt. Logic

AttitudeCommandProcessing

I M U(1)

RIU

Bus

Scheduler

0.2 lbREA(12)

5 lbREA(6)

LAE1

AJT(4)

Thruster

RWA

1553 Data Bus1553 Data Bus

ACS Algorithms (resides in OBC)ACS Algorithms (resides in OBC)

SSA(2)

ESA(2)

RWATach(4)

RWAs(4)

SensorsSensors ActuatorsActuators

RIU

Bus

Scheduler

1553 Data Bus1553 Data Bus


Spacecraft Feedback Controller

Plant(satellite)

Setpoint Error

ControlAlgorithm

Sensor

“point at sun”= 0 => V=0

V = Volts

Error Angle =Sun Sense - 0

T = -k+/- ThrusterTorque time

tTorque/I

Sun SensorV

Disturbances

Actuator

Actual


Angle Stabilization with position only feedback

-1.50

-1.00

-0.50

0.00

0.50

1.00

1.50

Time (seconds)

Angle ErrorError Angle Rate

0th Order Angle Controller


Angle Stabilization with position + rate feedback

-0.40

-0.20

0.00

0.20

0.40

0.60

0.80

1.00

1.20

Time (seconds)

Angle Error

Error Angle Rate

1st Order Angle Controller


Assembly, Integration & Test


ACS Test Setups

• Level 1 (sometimes all in 1 computer)

Controller Plant

Sensor Signals

Actuator Commands

• Actuator model• Actuator commands• Dynamic model• Orbit / Universe model• Sensor Model

=> Sensor Signals

ControlAlgorithm -->

• Level 2

SpacecraftController

Plant

Simulated

Actuator Commands

Real Control softwareReal sensorsReal output to actuators -->

• Actuator model• Dynamic model

• Orbit / Universe model• Sensor simulation=> simulated environment

Environment


More ACS Test Setups

• Level 3 (Big $)

• Level 4(on-orbit tweaking)

Spacecraft in vacuum on zero-mass, zero-friction

simulation mount


Simplifying ACS• Sensors

– Use simple ones: sun, magnetometers

– Use payload instruments as sensors• but beware accuracy limits and loop time

– Design using low-cost vendors, flight spares etc.

• Development and Test– Build testability into design

– Use Matlab or equivalent

– Simulate dynamics and sensors with external PC

– Use safe modes and assume final tweeks on orbit

• Managing the Payload– Can it search a little bit?

– Scanning vs. staring

– Larger apertures = shorter

– integration duration

– Duty cycling to avoid interference

– Self registration and non-real-time attitude reconstruction

• System Design– Choose simple modes - spinners, gg

– Avoid deployables

– Relax pointing / determination accuracy

– Use switching antennas and other techniques to eliminate some pointing requirements

– Basic autonomy / safing on board -– Handle anomalies on the ground

• Actuators– Air core torque coils where possible

– single wheel vs. 4-wheel momentum storage

– avoid propulsion (toxic, leaks, fluid handling, safety, lifetime limits)

• Alternative Approaches– Attitude Determination vs. Control

– Wide FOV instruments / multiple instruments

– Unstabilized and passive stabilization

chart 1 adcs design & hardwareme176: lecture 5 aaron rogers [email protected]

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