satellite bus platform subsystems

Technical Introduction to Geostationary Satellite Communication Systems Original Prepared by Telesat Canada

Slide Number 1Rev -, July 2001


Section 2.4


Slide Number 2Rev -, July 2001 Vol 2: Communication Satellites

Sec 4: Satellite Bus/Platform Subsystems

2.4.1: Introduction

2.4.1.1 What Is a Bus

2.4.1.2 Major Subsystems

2.4.1.3 Typical Spin Stabilized Spacecraft

2.4.1.4 Typical 3-axis Spacecraft

2.4.1.5 Historical Trends [Anik Spacecraft]

2.4.1.6 RF and Array Power Trend

2.4.1.7 TO Mass and Lifetime Trend

2.4.1.8 Technology Trends

Outline of This Part



What is a Bus?The bus is the platform that supports the payload and maintains

the satellite’s position in orbit.

The bus also provides the interface with the launch vehicle.

Bus Subsystems

Structure Electrical Power

Attitude Determination and Control

Telemetry &

Command

Propulsion Thermal Control

Subsystem

Mechanisms

2.4.1.1 Bus Subsystem

Vol 2: Communication Satellites, Sec 4: Satellite Bus/Platform Subsystems

Part 1: Introduction

2.4.1.1: What is a Bus?



Major Subsystemsprovides “real estate” for mounting all bus and payload units and the interface with the launch vehicle

provides electrical power to the payload and bus units

provides the control for achieving and maintaining orbit and pointing

provides the propulsive power for achieving and maintaining orbit



2.4.1.2: Major Subsystems

Structure

Electrical Power

Attitude Determination and Control

Propulsion



Major Subsystem

controls the spacecraft and monitors its health

maintains a benign operating environment

provides the means for deploying appendages which must be stored for launch, and the means to adjust appendages



2.4.1.2: Major Subsystems

Telemetry & Command

Thermal Control Subsystem

Mechanisms



Typical Spin Stabilized Spacecraft

2.4.1.3 Typical Spin Stabilized Spacecraft



2.4.1.3: Typical Spin Stabilized Spacecraft

Image Courtesy Image Courtesy of Boeing of Boeing

Satellite SystemsSatellite Systems



Typical 3-Axis Spacecraft

2.4.1.4 Typical 3-axis Spacecraft



2.4.1.4: Typical 3-Axis Spacecraft

Image Courtesy of Image Courtesy of Boeing Satellite Boeing Satellite SystemsSystems



Historical Trends [Aniks]

Anik A Anik B Anik C Anik D Anik E Nimiq Anik F1Prime Contractor Hughes RCA Hughes Spar Spar LM HughesSatellite Type HS-333 RCA-2000 HS-376 HS-376 GE-S5000 A2100AX HS-702Number 3 1 3 2 2 1 1Launch Vehicle Delta Delta STS/PAM-D Delta/STS Ariane Proton ArianeLaunch Date(s) 1972-75 1978 1982-85 1982-84 1991 1999 2000Transfer Orbit Mass (kg) 560 920 1140 1217 2930 3590 4600Array Power (W) 235 620 800 800 3900 8800 15000Life (years) 7 7 10 10 12 15 15Stabilization Spin 3-axis Spin Spin 3-axis 3-axis 3-axisTotal no. Channels 12 18 16 24 40 32 84Channels (C/Ku) 12/- 12/6 -/16 24/- 24/16 -/32 36/48HPA Power (W) (C/Ku) 5/- 10/20 -/15 11/- 12/50 -/120 40/115Total RF Power (W) 60 240 240 264 1088 3840 6960



2.4.1.5: Historical Trends (Anik Spacecraft)



0

4000

8000

12000

16000

1972 1982 1991 1999 2000

Arra

y Po

wer

0

1000

2000

3000

4000

5000

6000

7000

1972 1982 1991 1999 2000

RF P

ower

RF & Array Power Trend

2.4.1.6 RF & Array Power Trend



2.4.1.6: RF & Array Power Trend

(Wat

ts)

(Wat

ts)



0

1000

2000

3000

4000

5000

1972 1982 1991 1999 2000

Mas

s [k

g]T.O. Mass & Lifetime Trend

0

4

8

12

16

1972 1982 1991 1999 2000

Life

[yrs

]

2.4.1.7 T.O. Mass & Lifetime Trend



2.4.1.7: T.O. Mass & Lifetime Trend



Technological Trends

2.4.1.8 Technological Trends



2.4.1.8: Technology Trends

Image Courtesy of Telesat CanadaImage Courtesy of Telesat Canada




Vol 2: Communication Satellites

Electrical Power SubsystemPart 2




2.4.2.1 Introduction

2.4.2.2 Solar Arrays

2.4.2.3 Batteries

2.4.2.4 Power Electronics

2.4.2.5 Typical Failure Modes

2.4.2.6 Solar Array Analysis and Prediction Methods

2.4.2 Electrical Power Subsystem (EPS)





IntroductionA spacecraft power subsystem is designed to provide sufficient power to operate the spacecraft equipment over the life of the spacecraft.

For the majority of the mission, the source of electrical power is the solar arrays.

In eclipse, which occurs daily during 44-day periods twice a year for GEO spacecraft, electrical power is provided by batteries.

Regulation (control) of the variable source power is achieved with power interface electronics.



Part 2: Electrical Power Subsystems (EPS)



Typical EPS Configuration

SADM = Solar Array Drive Mechanism

Solar

Array (N)

Solar Array (S)

Power Control

Electronics

Battery

SADM

SADM

Fuse Box

Pyro Box

To spacecraft loads

To spacecraft Pyros

Battery discharge path

Battery charge

path




2.4.2.1 Typical EPS Configuration



CylindricalSolarArray




2.4.2.2a Cylindrical Solar Array

Image Courtesy Image Courtesy of Boeing of Boeing

Satellite SystemsSatellite Systems



Planar SolarArray




2.4.2.2b Planar Solar Array

Picture Courtesy of Telesat Canada



Solar Wing Assembly

Solar PanelSolar

cells Hinge

Yoke




2.4.2.2c Solar Wing Assembly



Solar Cell ConnectionsSolar cell assemblies are electrically configured on the panels in strings and circuits.

Strings consist of a number of solar cells in series to provide the required voltage.

Circuits are composed of a number of strings connected in parallel to provide the required current.

A solar cell assembly consists of a solar cell (silicon or gallium arsenide on germanium) and a cover glass bonded to its front surface.

Cell StringCircuit




2.4.2.2d Solar Cell Connections



Solar Panel Connections

Solar Panel

String

Cell

Circuit

+vBus

-vBus




2.4.2.2e Solar Panel Connections



ElectricalLoad

Collision with N-Type atom dislodgeselectron and results in electron migrationto negative terminal

-

+

-

+

e-

n

e-

Photon

N-TypeLayer

P-Type Layer e- Electron

Flow

Collision with P-Type atom dislodges electroncreating a vacancy spot called a hole whichwill migrate to positive thermal to accept electron

Solar Cell Operations




2.4.2.2f Solar Cell Operations



Current source

Rs

RloadRl

Rs = Source resistanceRl = Leakage resistanceRload = External load

Solar Cell Electrical Model




2.4.2.2g Solar Cell Electrical Model



Voltage

Current

Short circuit current (Isc)

Open Circuit voltage (Voc)

Voltage at Max Power Point (Vmp)

Current at Max Power Point (Imp)

Max Power Point

Constant current part of curve

Constant voltage part of curve

Solar Cell I-V Curve




2.4.2.2h Solar Cell I-V Curve



I = Isc * (1 - C1 * {exp[V / (C2 * Voc)] - 1})

Where:V is the bus voltage at which the current is to be calculated and C1 & C2 are constants as calculated below:C1 = [1-(Imp/Isc)] * {exp[-Vmp/(C2*Voc)]}C2 = [(Vmp/Voc) - 1] / [ln(1-Imp/Isc)]

Cell Voltage

Cell Current

Isc

VocVmp

Imp

Icell

Vcell

The Basic Cell Equation




2.4.2.2i Solar Cell I-V Curve



Voltage

Current

Short circuit current (Isc)

Open Circuit voltage (Voc)

BOL

EOL

Operating point

Array I-V Curve




2.4.2.2j Array I-V Curve



0.00

0.10

0.20

0.30

0.40

0.50

0 40 80 120 160 200

Panel Voltage, volts

Pan

el C

urre

nt, a

mps

Max Pow er PointPmax = 53.57 WImaxp = 0.3747 AVmaxp = 143.0 V

I-V Curve

Power Curve (I*V)

Max Power Point

I shunt

I load

Vref

Vbus

Iload

Iupper

array

Ishunt

Error amp

Ilower

array

Iload

Ireturn

Shunt eleme

nt

Spacecraft load

Upper section of solar array

Lower section of solar array

IshuntIload + Ishunt

Main bus

Shunt bus

Return bus

Shunt Regulator Operation




2.4.2.2k Shunt Regulator Operation



Array Output vs. Time




2.4.2.2l Array Output vs. Time



Batteries store electrical power for use during an eclipse or those short periods of time when there may not sufficient array power to support the full spacecraft load.

Battery types used for commercial spacecraft include:

• Nickel Cadmium (older spacecraft)

• Nickel Hydrogen (modern spacecraft)

• Lithium Ion (next generation spacecraft)

Spacecraft Batteries

2.4.2.3 Batteries





A voltaic cell is the basic device for converting chemical energy into electrical energy.

It consists of two different metal plates immersed in a solution.

The metal plates are called positive and negative electrodes and the solution is called the electrolyte.

- +

Sulfuric acid

electrolyte

Zinc electrode

Copper electrode

External loadExternal current

flow

Basic Battery Chemistry

2.4.2.3 Batteries



2.4.2.3a Basic Battery Chemistry



Charge

Positive electrode: Ni(OH)2 + OH- NiOOH + H2O + e-

Negative electrode: H2O + e- ½H2 + OH-

Overcharge

Positive electrode: 2OH- ½O2 + H2O + 2e-

Negative electrode: 2H2O + 2e- 2OH- + H2

Recombination: ½O2 + H2 H2O

Chemical Equation For Nickel Hydrogen Cell

2.4.2.3 Batteries





Discharge

Positive electrode: NiOOH + H2O + e- Ni(OH)2 + OH-

Negative electrode: ½H2 + OH- H2O + e-

Chemical Equation For Nickel Hydrogen Cell

2.4.2.3 Batteries





A battery assembly consists of a number of battery cells connected in series, where the potentials of the individual cells add to give the total battery potential.

The chassis mechanically fixes the cells as well as provides a thermal conductive path for the heat generated by the cells.

Satellite batteries also include electrical heaters and cell bypass circuitry.

The electrical heaters maintain the battery cells at the desired temperature during the endothermic (heat absorbing) charge phase.

Battery

2.4.2.3 Batteries





Individual cell bypass circuitry provides an alternate conductive path should the associated cell fail open. This is simple circuitry consisting, typically, of diodes.

Most battery cells in use today are Nickel Hydrogen (NiH2) technology. This technology provides a significant improvement in cycle life and energy density compared to Nickel Cadmium.

A NiH2 battery cell is comprised of a stack assembly of “pineapple slice” electrodes, separators, gas screens and insulator rings mechanically fixed on a central core with end plates, a sealed cylinder pressure vessel with electrical axial electrical terminals, and an electrolyte solution.

Battery

2.4.2.3 Batteries





2.4.2.3 Batteries



2.4.2.3b Superbird 83-Ah Nickel-Hydrogen Cell



Battery Temperature Profile

Time (hours)

Temp

0 12 24

Charge Cool down Discharge

Heaters ‘on’Heaters ‘on’

2.4.2.3 Batteries



2.4.2.3c Battery Temperature Profile



The power control electronics regulates the power bus voltage during sunlight conditions by controlling the amount of solar array power that is passed to the spacecraft loads.

A control signal, which is generated by a comparison of the actual and desired power bus voltages, controls how much of the excess solar array power is redirected through the shunt switches.

The power control electronics may contain a dedicated shunt switch module for each solar array circuit. Contained in each module is a shunt switch and an isolation diode to protect the power bus from being short circuited when this switch is operational.

Power Control Electronics






Battery discharging, during eclipse or periods of limited solar array power availability, is achieved via the battery discharge electronics. This circuitry is designed to operate over a large range of battery input voltages and provides a regulated output voltage.

Activation of the circuitry is controlled by a voltage comparison control signal.

The power control electronics shunt switch modules are designed such that capability exists to disable a given module and permit continued use of the associated solar array circuit.

The charge, discharge, and control electronics consists of a number of independent modules that permit the loss of a single module without any operational impacts.

Power Control Electronics






Fuse BoxThe fuse box typically consists of dedicated fuses for each spacecraft load.

Redundant parallel-configured fuses are sometime used for each spacecraft load. The idea here is that if one fuse blows because of an internal defect rather than because of load current draw, the other fuse will remain intact and power to the load will not be interrupted. A real load failure, however, or current surge, will cause both fuses to blow, thus protecting the load.

Pyro BoxA pyro box typically consists of redundant transistor switches that permit firing of the pyrotechnic devices: hold-down straps and bolts that are fired to permit deployment of arrays and antennas.

Dedicated separation switches may be allocated to the primary and redundant pyros of some separation mechanisms.






Slip Ring Assembly

Solar Array harness Drive

MotorGear

Assembly

Slip Ring Assy

Spacecraft harness




2.4.2.4 Slip Ring Assembly



Broken solar cells: A broken cell contributes to an overall loss of power. Accelerated life testing should quantify expected failure rates; the acceptance test program, and visual inspection, should identify failures that occur on the ground. Power subsystem design should tolerate a number of broken cells in orbit.

Battery cell short or open: Again, this results in a loss of power. Life testing should demonstrate robust design, the acceptance test program should identify cell short failures on the ground, and power subsystem design should tolerate a limited number of in orbit battery cell failures.

Typical Failure Modes






Power control electronics module failure: This should have little or no impact for a single in orbit failure because of the built-in redundancy. The acceptance test program should identify a component failure on the ground prior to launch.

Blown fuse: Parallel redundant fuses insure that a single defective fuse will not prevent powering of associated equipment. Blowing both fuses is a strong indication of a problem with the associated equipment and, thus, performs the intended function of protecting the spacecraft power bus.

Typical Failure Modes






Array Performance




2.4.2.6a Array Performance



Electrical Configuration

Solar Panel

String

Cell

Circuit

+v Bus

-v BusCircuits 2..N

Panel voltage ~ number of cells in series (Ns x Vcell)

Panel current ~ number of strings in parallel (Np x Icell)

Circuit 1

Circuits diode ‘ored’

Isolating diodes




2.4.2.6b Electrical Configuration



Typical IV Curve

0.00

0.10

0.20

0.30

0.40

0.50

0 40 80 120 160 200

Panel Voltage, volts

Pan

el C

urre

nt, a

mps

Max Pow er PointPmax = 53.57 WImaxp = 0.3747 AVmaxp = 143.0 V

I-V Curve

Power Curve (I*V)

Max Power Point

Current available at operating voltage• Short circuit

current, Isc

• Open circuit voltage, Voc

• Current at max power point, Imp

• Voltage at max power point, Vmp




2.4.2.6c Typical IV Curve



Factors Affecting the Current

Sun intensity ƒ(orbit, t)

Panel assembly factors

Cell characteristicsVoc, Isc, Vmp, Imp ƒ(rad, temp)

Temperature effects ƒ(t)

Solar Panel Configuration Iarray = ƒ(Ns, Np)

Radiation ƒ(t)

Current vs timeCell Equation

Iarray = ƒ(N1.. Nn)

Solar Array Program

(SAP)




2.4.2.6d Factors Affecting the Current



The Basic Cell Equation

Cell Voltage

Cell Current

Isc

VocVmp

Imp

Icell

Vcell




2.4.2.6e Solar Cell I-V Curve

I = Isc * (1 - C1 * {exp[V / (C2 * Voc)] - 1})

Where:V is the bus voltage at which the current is to be calculated and C1 & C2 are constants as calculated below:C1 [1-(Imp/Isc)] * {exp[-Vmp/(C2*Voc)]}C2 = [(Vmp/Voc) - 1] / [ln(1-Imp/Isc)]



Modeled Vs Measured




2.4.2.6f Modeled Vs. Measured



Array Output Over Mission LifeArray current is typically calculated once a day.

Sun intensity is cyclic over the year and can be modeled.

Cell temperature is usually given at each of the 4 seasons at BOL & EOL and curve-fitted over mission life.

Cell degradation due to radiation effects is given with respect to 1 MeV electron fluence. Degradation over time is calculated as the product of the flux per day and the time on-orbit, which gives the fluence at that point in time.

Solar flares can be included as an SF-dose linearly applied over the mission, over part of the mission, or as discrete events specified by the designer or user.

SF Alphas are taken as 5% of the SF protons.






Sun Intensity




2.4.2.6g Sun Intensity

Sun

Inte

nsity

Fac

tor

April 1st



IV Curves for the 4 Seasons

Array Working Point




2.4.2.6h IV Curves for the 4 Seasons



Isc FactorsThe short circuit current (Isc) factors affecting the performance of a solar array are listed below:

Solar intensity - variation due to varying sun angle and distance from from the sun. Normalized to 135.5 mW/cm^2.

Coverglass transmission loss - caused by glassing of the solar cell and UV degradation when on-orbit.

Assembly loss - measurement error and scattering due to glassing and interconnects.

Isc temperature coefficient - change in cell current due to temperature variations.






Voc FactorsThe open circuit voltage (Voc) factors affecting the performance of the solar array are listed below:

Assembly loss - series resistance of cell interconnections and weld resistance that depresses the knee of the IV curve

Voc temperature coefficient - change in cell voltage due to temperature variations. Magnitude affected by radiation.






The ProcessThe process of performing a solar array prediction involves gathering specific data about the solar array to be analyzed. Such data includes:

• Cell front and back shielding values from the physical characteristics of the panel, materials, cell, cover etc.

• Reducing the radiation environment (electron, proton and solar flare) to the equivalent 1 MeV electron fluence

• Curve-fitting the cell degradation factors affected by radiation

• Developing computer code for that particular cell

• Determining the rest of input data such as panel configuration, seasonal temperatures, losses etc.






Concept of 1 MeV FluenceThe concept of damage-equivalent, normally-incident (DENI) mono-energetic 1 MeV fluence was developed by the solar cell industry to determine the degradation effects of a radiation environment with various energies and incident angles.

In this normalizing calculation, the actual damage due to electrons of various energies is related to the damage produced by 1 MeV electrons by the damage coefficients for electrons.

Likewise, proton damage is related to 10 MeV protons, which in turn is related to the damage produced by 1 MeV electrons.






Concept of 1 MeV FluenceOne 10 MeV proton does approximately the same damage as 3000 electrons of 1 MeV energy.

By combining the electron, proton and SF fluences, a single value of equivalent 1 MeV fluence can be used to determine cell degradation in a complex radiation environment.

Trapped protons in geostationary orbit are not modeled because their energies are low enough so that they are absorbed by the coverglass.

Typical solar cell IV curves before and after exposure to a heavy dose (1X1015 e/cm2) of 1 MeV electrons are shown on the next slide.






Temperature & Radiation Effects*Typical solar cell IV characteristics before and after irradiation.

Temperature effects are also shown.

With higher temperatures, current increases while voltage decreases




2.4.2.6i Temperature and Radiation Effects



1 MeV Fluence vs Cover Thickness*Shielding effectiveness changes with incident radiation, particle type and particle energy.

Figure 2.4.2.6j is a useful graph for estimating 1 MeV fluence for a given shielding.

PICTURE

Fluencevs

CoverThickness




2.4.2.6j 1 MeV Fluence vs Cover Thickness





Telemetry, Tracking & Command Subsystem

Part 3




2.4.3: Telemetry, Tracking & Command (TT&C) Subsystem


Outline of This Part2.4.3.1 TT&C Key Requirements

2.4.3.2 TT&C Equipment

2.4.3.3 TT&C Key Items

2.4.3.4 Command System Block Diagram

2.4.3.5 Command Format

2.4.3.6 Telemetry System Block Diagram

2.4.3.7 Telemetry Format

2.4.3.8 Creation of an 8-Bit Telemetry Word

2.4.3.9 Data Encoding

2.4.3.10 Failures, Degradations & Margins



• Receive, decrypt, authenticate, and process commands.• Collect, format, encrypt, and transmit satellite telemetry.• Support satellite control functions.

• Attitude determination and control• Battery charge management, solar array pointing• Autonomous configuration management

• Support range determination from ground station(s). • Provide antenna coverage for transfer & drift orbit operations

and during on-orbit attitude anomalies.• Be designed without any single point-of-failures.

2.4.3.1: TT&C Key Requirements


Part 3: Telemetry, Tracking & Command (TT&C)

A TT&C System must:



Typical TT&C Subsystem

TT&C RF Equipment Flight Software TT&C Baseband

Equipment

- CMD Receivers- CMD Horn Antenna(s)- TLM Horn Antenna(s)- CMD & TLM Omni Antenna- MISC RF H/W and Cabling

- CMD & TLM Database

- AD&C Software (Flight S/W)

- Encoder/Decoder Units- Remote Terminal Units

- Payload - Bus

- Computers- Harnesses

2.4.3.2: TT&C Equipment



Figure 2.4.3.2 TT&C Equipment



TT&C• CMD Uplink 500 bps

• TLM Downlink 4 kbps

• Encryption, Decryption

• Spacecraft Ranging

TT&C Omni Antenna

Arabsat 3A

TT&C On-station Antenna

2.4.3.3: TT&C Key Items



Figure 2.4.3.3 TT&C AntennasDrawings Used by Permission



Commanded functions include unit configuration, gain settings, redundancy settings, jet firings etc. The red and blue lines indicate main redundancy paths, while the black lines indicate redundancy switching options.

Command Receiver

Command Receiver

Command Decoder

Command Decoder

Remote Terminals

Remote Terminals

H

Ranging signal to tlm tx

Ranging signal to tlm tx

2.4.3.4: Command System Block Diagram



Figure 2.4.3.4 Command System Block Diagram

Cross Strapping



SYNCH/ADDRESS EXEC OP-CODE DATA WORD PARITY

SPACECRAFT COMMAND WORD

Commands are validated on-board prior to execution. Validation criteria are:

• Synchronization pattern

• Spacecraft address

• Command length

• Command segment order & content

• Parity

2.4.3.5: Command Format





Telemetered signals include unit status, temperatures, voltages, currents, register contents, etc.

Sensors

Signal Conditioning

Telemetry Tx

Telemetry Encoder

Telemetry Encoder

Remote Terminal

Unit

Sensors

Signal Conditioning

Telemetry Tx

H

Ranging signal from cmd rx

Ranging signal from cmd rx

2.4.3.6: Telemetry System Block Diagram



Figure 2.4.3.6 Telemetry System Block Diagram



Frame Synch SCID

Telemetry Words

TLM Mode Format ID Fixed

Variable Telemetry Words




Variable Telemetry Words Frame Count Checksum

SPACECRAFT TELEMETRY MINOR FRAME

2.4.3.7: Telemetry Format



• Telemetry transmission is composed of major and minor frames.• A major frame is a complete set of telemetry data.• The major frame is made up of a number of minor frames.• Each minor frame carries a number of Telemetry Words.

Framing



8-bit Digital Coding

2.4.3.8: Creation of an 8-Bit Telemetry Word





00 001 110 016 Your message00 110 110 066 Secret key11 000 111 307 Coded message

11 000 111 307 Coded message00 110 110 066 Secret key00 001 110 016 Your message

When:2 bits are the same, cipher text = 12 bits are different, cipher text = 0

2.4.3.9: Data Coding



Simple Example: The Exclusive “OR” Function



Typical TT&C designs offer low risk configurations :• No deployable antennas for transfer orbit operations• No RF switches in the command path(s)• Redundancy and cross-strapping of CMD/TLM/RNG signals• Multiple modes of operation, i.e. High & Low Power Transmitter

outputs• Positive RF link margins for CMD/TLM/RNG

On-orbit problems are generally due to H/W failures or degradation.

Operational recovery is achieved by a combination of cross-strapping signal paths and redundant equipment selection.

In a loss of earth-lock, Flight Software (FSW) typically reconfigures TLM transmission to high power, wide angle coverage to facilitate S/C recovery attempts.

2.4.3.10: Failures, Degradations & Margins







Attitude Control SubsystemsPart 4




2.4.4: Attitude Control Subsystems


Outline of This Part• Introduction• ACS Principles and Design• Sensors• Actuators• Spacecraft Processors• Operating Modes• Reliability and Risk• ACS Testing



The Attitude of a Spacecraft is its orientation in space.

Position and Velocity describe the translational motion of the center of mass of the spacecraft. Translational motion is motion from one location to another.

Attitude and attitude motion describe the rotational motion of the body of the spacecraft about the center of mass.

How is Attitude determined, and how is it controlled?

2.4.4.1: Introduction


Part 4: Attitude Control Subsystems

Initial Definitions



Attitude Determination is the process of computing the orientation of the spacecraft relative to a point of reference such as the Earth. This typically involves the use of several types of sensors and a means to process the resulting data.

Attitude Control is the process of orienting the spacecraft in a predetermined direction. This consists of stabilization, maintenance of an existing orientation, maneuver control, and controlling the reorientation of the spacecraft from one attitude to another.

Both of these functions are performed by the Spacecraft’s Attitude Control Subsystem (ACS).




2.4.4.1.1: Functional Definitions



Nominally box : + 0.05o longitude

+ 0.05o inclination

Worst case: + 0.1o for each

At 35,786 Km:

1o = 624 Km

0.05o =31Km

+ 0.05o = box 62 Km square

Station Keeping Box Maintain satellite in an orbit position so it always in the FOV of a non-tracking Earth Station

Stationkeeping box(0.05º or 0.1º)

Earth Station Beam Width




Figure 2.4.4.1.1 Station Keeping Box

2.4.4.1.1: Functional Definitions



Sensors(Side 1)

• Earth• Sun• Gyro

Sensors(Side 2)

• Earth• Sun• Gyro

Processor 1

Processor 2

Actuators(Side 1)

• Thrusters• Wheels

Actuators(Side 2)

• Thrusters• Wheels




2.4.4.1.2: Typical ACS Configuration

Figure 2.4.4.1.2 Typical ACS Configuration


Slide Number 76Rev -, July 2001 Vol 2: Communication Satellites, Sec 4: Satellite Bus/Platform Subsystems

2.4.4.1.3: Definition of Axes




YAW

Z

ROLL

X

PITCHY

EARTHSENSOR

THREE-AXIS (BODY STABILIZED

EARTH

Figure 2.4.4.1.3a 3-Axis



PITCH

(East-West)





Figure 2.4.4.1.3b Pitch



ROLL

(North-South)





Figure 2.4.4.1.3c Roll



YAW

(Beam Rotation)





Figure 2.4.4.1.3d Yaw



Center of Mass (c.o.m.) is the point where the satellite mass is considered to be “concentrated”. It is known as “that point at which the entire mass of an object may be considered to be located for purposes of understanding the object's motion.”1

Center of Gravity is the point where the force of gravity is considered to be acting. It may be different than the center of mass when mass distribution is not equidistant from the source of gravitational attraction.

2.4.4.2: ACS Principles and Design



2.4.4.2.1: Centers of Mass and Gravity



Moment of Inertia is a measure of resistance to change in rotational speed. It is a way of specifying the mass “distribution” about a certain axis.

Product of Inertia is a measure of the influence of an object’s geometry on its rotation. It is a way of defining the symmetry of an object about a plane defined by two axes.




2.4.4.2.2: Moments and Products of Inertia



Angular Momentum is a property of a rotating body,

H = [I] wIt is representative of a body’s moment of inertia [I], and rotation rate, w, and is usually measured in Newton-meter-seconds.

Angular momentum is a vector value, i.e., it has both magnitude and direction.

“Torque” is an external influence caused by forces acting about the center of mass, and is usually measured in N-m. It is caused by rotational devices (motors, shafts) and will affect the body’s angular momentum.




2.4.4.2.3: Angular Momentum and Torque

EQ. 2.4.4.2.3 Angular Momentum






2.4.4.2.4: Torque and Moment Arm

Fo rce F

C e n te ro f

M a s s

r

P e r p e n d ic u la rd is ta n c e

o rM o m e n t Ar m

Figure 2.4.4.2.4 Torque and Moment Arm



A “system” is a grouping of 2 or more “bodies.”

In the instance of a spacecraft, we have the spacecraft body itself and each of the momentum and reaction wheels.

Each “body” can rotate, and has its own angular momentum.

Individual “bodies” can affect the rotational state of each other (torque one another) and exchange angular momentum.




2.4.4.2.5: System Angular Momentum



System Angular Momentum is the “vector sum” of all “body” contributions.

It is affected only by torques “external” to the system. These torques can be interactions between the spacecraft and its environment, such as solar pressure on the solar panels.

It is not affected by torques internal to the system. Internal effects, such as antenna deployments, stay inside the system.




2.4.4.2.5: System Angular Momentum

Gyroscopic Stiffness Effect demonstrates how a rotating body tends to stay rotating in the same state (unless acted upon).

A spin axis will remain pointing in the same direction, this is a consequence of Newton’s laws.

2.4.4.2.6: Gyroscopic Stiffness Effect



Gyroscopic Stiffness is beneficial. It provides stability in orientation, reduces effects caused by external disturbances and it is the result of having angular momentum.

Spacecraft solutions that emerge as a result are:

• Spinning the satellite (“spin-stabilized”)

• Spinning wheels within satellite (“3-axis stabilized”)

• 3-Axis stabilization, but without a momentum bias




2.4.4.2.6: Gyroscopic Stiffness Effect



• Solar radiation pressure

• Gravity gradient and other gravitational sources

• Earth’s magnetic field

• Micro-meteoroid impacts

• Thrusters




2.4.4.2.7: Disturbance Sources and their Effects



Attitude effects resulting from disturbance torque include:

Precession is the rate of change in the direction of the angular momentum vector. This is caused by a torque acting over time.

If precession is slow and not corrected for, the whole satellite will drift in its orientation.

Nutation appears as coning type of motion when the spacecraft is disturbed from its equilibrium state; whenever precessional torques are applied the mode will be activated. The coning motion centers around the original direction of angular momentum.

It is the result of having too much rotational kinetic energy, which can be damped out actively (on-board controller), or passively (naturally)




2.4.4.2.7: Disturbance Sources and their Effects



In order to counteract disturbances, the ACS system includes:

• Devices that interact with the external environment, therefore affecting the system angular momentum. These include thrusters and magnetic torquers.

• Devices that act inside the spacecraft to redistribute the angular momentum within the spacecraft. Momentum/reaction wheels do this. Wheel “saturation”, however, requires momentum unloading.




2.4.4.2.8: Ways to Counteract Disturbances






2.4.4.2.9: Spin-Stabilized Satellites

H

SINGLE-SPIN(SPINNERS)

w

Figure 2.4.4.2.9a Single Spin (Spinners)






2.4.4.2.9: Spin-Stabilized Satellites

H

NONSPINNING

SPINNINGwFigure 2.4.4.2.9b Spin Stabilized Satellites






2.4.4.2.10: Three-Axis Stabilized Satellites

H

MOMENTUM-BIASED

w

Figure 2.4.4.2.10a 3-Axis Stabilized Satellite







Xs Zs

Ys

CE N T RA L R IG IDBO D Y

SO LA R P A NE L

M O M EN TU M W H E ELS

M O M EN TU M -B IAS ED (W IT H RE A C TIO N W H E E LS IN 3 AXE S)

Figure 2.4.4.2.10b Momentum-Biased







ZERO-MOMENTUM(W ITH REACTION W HEELS IN 3 AXES)

Figure 2.4.4.2.10c Zero-Momentum







Figure 2.4.4.2.10d Zero-Momentum (Gravity Gradient Stabilized)



Disturbances from Environment

Spacecraft Dynamics

ACS Subsystem

Sensors SCP Actuators




2.4.4.2.11: ACS Analytical Design

Figure 2.4.4.2.11 ACS Analytical Design



Pointing error budgets analyze the temporal behavior of the various error sources, usually divided into four categories:

• Constant

• Long term (longer than one day)

• Diurnal (one day)

• Short term (less than 10 minutes)

An addition of the above categories represents a conservative assessment.




2.4.4.2.12: Pointing Budgets



Typical error sources are:

• Thermal Distortion

• Disturbance Torques

• Misalignment Errors

• Structural Hysteresis

• Orbital Effects

• Sensor Noise




2.4.4.2.12: Pointing Budgets



Source of Error Normal Mode Stationkeeping Mode Roll Pitch Yaw Roll Pitch Yaw

Constant (or Fixed bias) ErrorsEarth sensor alignment 0.022 0.022 0.022 0.022Sun sensor alignment 0.015 0.015 0.015 0.015Antenna characterization errors 0.032 0.032 0.021 0.032 0.032 0.021Root Sum Square (RSS) of Constant Errors 0.041 0.041 0.021 0.041 0.041 0.021In-orbit calibration residual of Constant Errors 0.016 0.016 0.021 0.016 0.016 0.021Long-term ErrorsSensor long-term degradation 0.011 0.009 0.021 0.011 0.009 0.021Sensor seasonal variations 0.007 0.007 0.007 0.007Structure seasonal thermal distortions 0.017 0.017 0.008 0.017 0.017 0.008Orbital variations (East/West and North/South) 0.011 0.014 0.015 0.011 0.014 0.015Root Sum Square (RSS) of Long-term Errors 0.024 0.025 0.027 0.024 0.025 0.027Diurnal ErrorsGyro drift 0.105 0.105Sensor diurnal errors 0.033 0.039 0.045 0.033 0.039 0.045Ephemeris Error 0.002 0.004 0.018 0.002 0.004 0.018Structure diurnal thermal distortions 0.012 0.012 0.012 0.012Root Sum Square (RSS) of Diurnal Errors 0.035 0.041 0.115 0.035 0.041 0.115Short-Term ErrorsSensor noise 0.004 0.002 0.015 0.004 0.002 0.015Actuator transients 0.007 0.007 0.007 0.018 0.018 0.025Solar array tracking torque 0.002 0.004 0.002 0.002 0.004 0.002Maximum uncorrected disturbance torque 0.022 0.022 0.013 0.029 0.029 0.017Root Sum Square (RSS) of Short-term Errors 0.023 0.025 0.021 0.034 0.035 0.034Total Error (arithmetic sum of RSS terms) 0.123 0.132 0.184 0.134 0.142 0.197Total Error (arithmetic sum after in-orbit calibration) 0.098 0.107 0.184 0.109 0.117 0.197



Part 4: Attitude Control Subsystems2.

4.4.

2.12

: Poi

ntin

g B

udge

ts



In basic ACS systems, onboard control logic responds to two-axis attitude sensing while it provides three-axis control.

In more advanced systems, the three axes are being sensed and controlled directly and the total angular momentum is maintained near a defined nominal value.

Typical geosynchronous systems use sensors such as:• Earth sensors, Horizon sensors

• Sun sensors, Star trackers

• Gyros

• Accelerometers, magnetometers

• RF Beacons

2.4.4.3: Sensors



Introduction



Types of Earth Reference sensors are:

Infrared detectors (CO2 spectral band)– Detects Earth’s horizon

– Reference for roll, pitch, and spin rate

Beacon Tracking– RF pointing system that uses a beacon (RF signal generated from

the ground) to point the satellite antenna at the Earth

– Usually for geosynchronous orbit

– Requires cooperative Earth station

Magnetometer– Senses Earth’s magnetic field, requires ephemeris knowledge

2.4.4.3: Sensors



2.4.4.3.1: Earth Reference Sensors



• Spinners use the rotation of the satellite as a timebase to detect the difference between the IR value of deep space and that of Earth.

• This change in IR level is used by the ACS Electronics to point the antennas towards the Earth.

• Three-axis satellites use torsional or vibrating mirrors to set-up an artificial timebase.

Deep space

Deep space

IR Radiation from Earth

North sensors scan “North” of equatorSouth sensors scan “South” of equator

NES

SES

2.4.4.3: Sensors




Sensors

Figure 2.4.4.3.1a Earth Reference Sensors



2.4.4.3: Sensors




NORMAL

PITCHERROR

YAW ERROR

ROLL ERROR

N

S

N

S

N

S

N

S

ERROR DETECTION

Figure 2.4.4.3.1b Error Detection



Oscillating Earth Sensor Static Earth Sensor

Earth Sensor output is proportional to area on Earth. Electronics use this to point the satellite body.

2.4.4.3: Sensors




Figure 2.4.4.3.1c Earth Sensor Positioning



2.4.4.3: Sensors



2.4.4.3.2: Beacon Sensors

PLANAR ARRAYANTENNA

ANTENNASYSTEM S SSM A

2 - 85 Hz TO NES

2 - 85 Hz TO NES

TCR#1

DAzDEL

TCR#2

DAzDEL

DAzDEL

DAzDEL

SCP#1

SCP#2

SH IELDED T W ISTEDPAIRS W ITH RETU RNS

TTC&R RFSUBSYSTEM

ATTITUDE CONTROLSUBSYSTEM

ANTENNASUBSYSTEM

Figure 2.4.4.3.2 Beacon Sensors



Sun Sensor

– For sun acquisition and tracking

– Good second reference for earth satellite

– For single-axis updating (yaw), or recalibration of gyro

Star Trackers (many varieties)

– Gimballed tracker (mechanically complex)

– Strapped-down mapper (substantial data processing required)

– Electronic tracker (image dissector)

– Provides a high-accuracy reference

2.4.4.3: Sensors



2.4.4.3.3: Stellar Reference Sensors



2.4.4.3: Sensors



2.4.4.3.3: Stellar Reference SensorsSUNLIGHT

AIR VACUUM

PATTERN MASK

PHOTOCEL

SLIT MASK

COARSE ANALOGSUN SENSOR

(CASS)

Figure 2.4.4.3.3a Stellar Reference Sensors



2.4.4.3: Sensors




Figure 2.4.4.3.3b Cass Output



2.4.4.3: Sensors




BIT 4F

BIT 3F

BIT 2F

BIT 1F

QUADRATUREPATTERNS (FINE BITS)(GRAY CODED BITS 1

THROUGH 3)

BIT 6CBIT 5C

BIT 4C

BIT 3C

BIT 2C

BIT 1C

COARSE BITS(GRAY CODED BITS

4 THROUGH 9)

ATA

PHOTOCELLS

DIGITAL SUN SENSOR (DSS) Figure 2.4.4.3.3c Digital Sun Sensor



2.4.4.3: Sensors



1/8°

2°

32°GRAY CODED

BIT 1

BIT 2

BIT 3

BIT 4

BIT 5

BIT 6

BIT 7

BIT 8

BIT 9

*AVERAGE VALUE

±1/8

±1/4

±1/2

±1

±2

±4

±8

±16

±32

DSS OUTPUTFI

NE

BIT

SC

OA

RSE

BIT

S2.4.4.3.3: Stellar Reference Sensors

Figure 2.4.4.3.3d DSS Output



Gyroscope

• Provides gyrocompass reference

• Provides a reference for active nutation damping

• 3-axis determination (position and rate)

• Requires independent updating to compensate for drift

Accelerometer

• Guidance reference to boost phase

• Phase reference for active nutation damping

2.4.4.3: Sensors



2.4.4.3.4: Inertial Reference Sensors (Gyros)



2.4.4.3: Sensors



AZ

(YAW)

AZ

(SKEW )

AZ

(PITCH)

AZ

(ROLL)

GYRO ASSEBBLIESARE FREQUENTLY

ARRANGED ASSINGLE UNITS

CONTAINING 3 OR 4INDEPENDENT

GYROS, ARRANGEDSO AS TO PROVIDE

POSITION AND RATEINFORMATION

ABOUT ANY AXIS.

a

b

a = 45°b = 35.3°90° - b = 54.7°AS = 1 (X+Y+Z)

3

A A A


Figure 2.4.4.3.4 Gyros



Accelerometers

• Accelerometers typically respond to linear acceleration.

• A miniature servo system responds to input acceleration along its sensitive axis.

• The movement is detected by a position error detector.

• An amplifier sends a feedback current via a restoring coil in a magnetic field. This applies a restoring force on the seismic system, returning it to its original position, nulling the position error detector.

• An analog voltage proportional to the input acceleration is measured and decoded into counts for the processor.

2.4.4.3: Sensors






Typical modern geosynchronous systems use actuators such as: • Two gimbaled momentum wheels / four reaction wheels in a

pyramid, all with momentum unloading (there are many variations)

• Propulsion torquers

• Magnetic actuators

The actuator (and sensor) controls, are performed through: • Central processor units (sensor processing, actuator drivers)

• Remote control units (sensor and actuator interfaces)

• Flight software (controls, service and fault protection algorithms)

2.4.4.4: Actuators



Introduction



Momentum and Reaction Wheels (MW, RW) are:• Used to stabilize the satellite throughout its orbital mission life

• Minimize roll and pitch errors

• Damp roll / yaw nutation

• Can provide gyroscopic stiffness

Three wheel assemblies are comprised of:• Two “momentum” wheels (primary mode)

• One “reaction” wheel (used in secondary mode)

• Each with its own wheel drive electronics (WDE)

2.4.4.4: Actuators



2.4.4.4.1: Momentum and Reaction Wheels



Four-wheel assemblies compromise of:

– Four momentum wheels in a pyramid configuration, all operating concurrently

Only three actually needed to meet the momentum storage requirements of the spacecraft, fourth wheel provides redundancy.

In all cases, momentum exchange permits cancellation of cyclic torques, primarily solar pressure torques, without employing attitude jets.

2.4.4.4: Actuators



2.4.4.4.1: Momentum and Reaction Wheels



2.4.4.4: Actuators



2.4.4.4.1: Momentum and Reaction WheelsFigure 2.4.4.4.1 Momentum and Reaction Wheel

Imag

e Co

urte

sy o

f Tel

esat

Can

ada

Imag

e Co

urte

sy o

f Tel

esat

Can

ada



As we have seen, apogee burns require a specific motor, the Liquid Apogee Engine (LAE).

Other, smaller thrusters are used for angular momentum unloading, stationkeeping, and large angle rotations.

Thrusters usually come in two (or more) redundant sets, with each set capable of providing torque about all 3 axes.

Each set is referred to as a “string” (A or B).

Each thruster has its redundant equivalent.

Each has its own latch valve (to isolate it).

2.4.4.4: Actuators



2.4.4.4.2: Thrusters



Three categories of thrusters are common:

• Solid propellant

• Liquid propellant

• Electric propulsion

2.4.4.4: Actuators



2.4.4.4.2: Thrusters

Figure 2.4.4.4.2 A Thruster at Work

Photo Courtesy of General Dynamics



Magnetic TorquerMagnetic coils—or electromagnets—are used to generate magnetic dipole moments for attitude and angular momentum control. These are also used to counteract residual spacecraft biases and attitude drift due to environmental disturbance torques.

A magnetic torquer consists of a simple coil which produces a magnetic dipole when current flows around the loop.

The magnetic dipole’s strength is proportional to the ampere-turns and area enclosed by the coil, and the direction is normal to the plane of the coil.

The torque acting on the spacecraft is the cross-product of the magnetic dipole of the coil and the Earth’s magnetic field. The direction of the control torque can be reversed by changing the direction of the current in the coils.

2.4.4.4: Actuators



2.4.4.4.3: Magnetic Actuators



2.4.4.4: Actuators



2.4.4.4.4: Solar Wing Positioners and Drivers

Zm

Xm

YmSOLAR RADIATION

PRESSURE MOMENT

EARTH

a

X

YZ

CM

CPr S

Figure 2.4.4.4.4a Solar Radiation Pressure Moment



2.4.4.4: Actuators



SPECU LAR YR EFLEC T ED

PH O T O N S

IN C O M IN GPH O T O NS

S

Fn n

SU R FAC E AR EA = ASO LAR PR ESSU R E = P

FT

SOLAR RADIATION FORCEON SURFACE

2.4.4.4.4: Solar Wing Positioners and Drivers

Figure 2.4.4.4.4b Solar Radiation Force on Surface



As in any feedback implementation, the sensed data has to be processed to establish deviations from the desired attitude and generate the necessary commands to the actuators.

Such computation, involving logic switching, control law implementation, and dynamic compensation, can be accomplished by unique electronic circuitry (analog and digital), a general purpose spacecraft computer, or by dedicated microprocessors.

Typically, modern spacecraft include a flight processor that is used for many spacecraft functions, not only ACS.

2.4.4.5: Spacecraft Processors



Introduction






2.4.4.5.1: Typical ACS Hardware Architecture

GYROS

SUN SENSORS

TW O-AXISEARTH SENSORS

ATTITUDEREFERENCE SENSORS

PROCESSOR

CONTROLW HEELS

THRUSTERS

TORQUE ACTUATORSSCE

DCU-B

Figure 2.4.4.5.1 Typical ACS Hardware Architecture






2.4.4.5.2: Typical ACS Software Design

EphemerisEphemerisPredictionPrediction

EarthEarthSensorSensorControlControl

SunSunSensorSensorControlControl

GyroGyroControlControl

AttitudeAttitudePredictionPrediction

ThrusterThrusterControlControl

SolarSolarWing DriveWing Drive

ControlControl

MomentumMomentumWheelWheel

ControlControl

AttitudeAttitudeDeterminationDetermination

AttitudeAttitudeControlControl

ModeModeControlControl

GainGainSelectSelect

Figure 2.4.4.5.2 Typical ACS Software Design



ACS operation must consider spacecraft operation during all aspects of its mission, including:

• Launch/Ascent

• Transfer Orbits

• Sun/Earth Acquisitions

• In-Orbit Testing

• On-station Operations

• Station Keeping and Momentum Dumping

• Safing

2.4.4.6: Operating Modes



Introduction






2.4.4.6.1: Launch and AscentDuring launch and ascent, the spacecraft ACS will typically be inactive. The launch vehicle has its own ACS equipment.

Separation from the launch vehicle is passive: the launch vehicle will orient the satellite with the sun line incident on the stowed solar panels.

The launch vehicle may also impart a slow spin to the satellite to maintain it in this orientation. This spin is usually imparted either by rotating the launch vehicle stage prior to release, or by ejecting the satellite with the springs set to uneven tensions.

While the inactive attitude control is being performed, attitude determination is usually obtained through the gyros.






2.4.4.6.2: Transfer OrbitsAfter launch, the spacecraft will undergo a series of thruster firings to switch through one or more transfer orbits. These raise the orbit and then circularize it into a geosynchronous location.

The thrusters involved include the Liquid Apogee Engine and various Reaction Control thrusters to maintain the satellite in the desired orientation throughout the procedure.

In some cases, electric propulsion thrusters are also used to gradually circularize the orbit.

Also, typically in the course of transfer orbit operations, antenna and solar array deployments are performed.






2.4.4.6.3: Sun/Earth AcquisitionsThe final stage of transfer orbit operations, prior to arrival on-station, is to perform Sun and then Earth acquisition.

Sun acquisition is initiated by slewing the wings (using solar wing drive mechanisms) to search for the sun, using the solar panels’ output current as an indicator.

Once this is accomplished, the spacecraft is then slewed to put the sun in the field of view of the sun sensor (usually digital, for higher accuracy).

Spacecraft slews are then performed within the sun sensor field of view to more accurately position the spacecraft.






2.4.4.6.4: In-Orbit TestingOnce the spacecraft is on-station, in-orbit testing is performed to ensure the spacecraft is completely operational. As part of this testing, antenna pattern measurements, sometimes called antenna cuts, are performed.

This is done by using the momentum wheels to slew the spacecraft in pitch and then roll while a ground station measures the strength of a signal it is transmitting. This verifies the pattern edges and beam shape.

Once the antenna pattern measurements have been completed, a bias value can be obtained to alter the spacecraft orientation in pitch and roll to maximize antenna performance over the desired footprint.






2.4.4.6.5: On-Station OperationsIn its normal, on-station mode, the spacecraft usually uses data from the Earth sensor or, in some cases, from RF beacon sensors, to maintain highly accurate pointing of the antennas over the coverage area.

The Earth sensor provides pitch and roll information, while the gyros are used to provide yaw data. The gyros are recalibrated once per orbit using the sun sensor, when the sun is in the proper position. The solar wing drive mechanisms are used to allow the solar panels to follow the sun.

Momentum buildup due to disturbances is stored in the momentum wheels, which maintain the desired angular momentum direction, or in a zero-momentum system, as the case may be.






2.4.4.6.6: Station Keeping and Momentum DumpingSince the momentum wheels store momentum resulting from disturbance torques, a momentum dumping maneuver must be periodically performed in order keep the wheels within the linear region of their spinning characteristics.

This is generally done by firing thrusters, and allowing the wheels to spin down at the same time.

Often, momentum dumping is performed in conjunction with stationkeeping maneuvers in order to save thruster propellant.

However, momentum dumping may be needed more frequently if momentum buildup is faster than the station keeping interval.






2.4.4.6.6: Station Keeping and Momentum DumpingStation keeping is generally performed by firing the appropriate combination of thrusters to correct drifts in East/West or North/South position.

Propellant thrusters are often used for this purpose, but electrical propulsion thrusters are also common.

The interval between station keeping maneuvers may range from one day to two weeks.

Often, the maneuvers are performed automatically, especially for the more frequent cases.

During the maneuver, momentum wheels maintain the spacecraft’s pointing accuracy.






2.4.4.6.7: SafingThe spacecraft must also include a safing mode that activates in the case of failures of equipment, incorrect thruster firings, or other unexpected events.

The safing mode will vary depending on spacecraft design, but as a rule it must be entered automatically from any operating mode of the ACS, during transfer orbits and on-station. Its top priorities must be to maintain communication with Earth.

The safe mode will allow the spacecraft to stay in a benign state, where fault propagation is minimized, telemetry continues to be provided to the ground, and where ground command capability is available.

In the case that Earth pointing is lost, the safing mode will immediately attempt to reacquire Earth in the shortest possible time.



ACS reliability is typically 98% after 15 years.

For both actuators and sensors, unit and subsystem integrity depends on:

• Unit redundancy and subsystem cross-strapping• Design lifetime• Operational range of the units• Fields of View of the sensors• Interference to the sensors (Sun, Moon, Spacecraft reflections)• Controlled backup modes in the event of a failure• Workmanship (why we test)• Alignment and offset of the units• End to end polarity of the sensors to actuators

2.4.4.7: Reliability and Risk



2.4.4.7.1: Reliability



Typical Maneuvers (Control processor applies to all cases)

Transfer Orbit GNC Units RiskSpin Up / Down Thrusters / Sensors (1st) Med

Precession for Apogee burns Thrusters / Wheels / Sensors (1st) Med

Orbit Raising Single Engine High

Nutation Control Thrusters / Wheels / Pivot (1st) Med

Reorientation Thrusters / Wheels / Sensors Low

Solar Array/Ant. Deployments Mechanisms (1st) High

Sun / Earth AcquisitionsThrusters / Wheels / Sensors Med




2.4.4.7.2: Risk



Typical Maneuvers (Control processor applies to all cases)

Mission Operations GNC Units RiskRoll / Yaw / Pitch Control Wheels / Sensors Low

North / South Stationkeeping Thrusters / Wheels / Sensors Low

East / West Stationkeeping Thrusters / Wheels / Sensors Low

Momentum Unloading Thrusters / Wheels / Sensors Low

OtherIn-orbit Test All (1st) Med

Antenna Pattern Mapping Thrusters / Wheels / Sensors Med




2.4.4.7.2: Risk



The qualification of an ACS design entails analysis, simulation and testing on several different levels.

• Unit Testing

• Simulations

• Bench-level Validation

• Satellite-level Testing

• In-Orbit Testing

2.4.4.8: ACS Testing



Introduction






2.4.4.8.1: Unit TestingAll of the ACS units have their own unit-level test programs.

Newly-designed units undergo qualification testing.

Units with minor design modifications (from the qualified article) are subjected to protoflight testing.

Units with no changes from qualified designs are acceptance tested.

Each of these programs entails a suite of environmental and performance tests. Testing range and extent depend on whether the unit is undergoing qualification, protoflight, or acceptance testing.






2.4.4.8.2: SimulationsAt the spacecraft level, simulations are performed to support mission and failure mode analysis and to determine whether the ACS’s design is sound from a system point of view.

The simulations are performed in software running on a high performance platform.

Flight dynamics simulators are run in conjunction with mathematical models of the various sensors and actuators, with the control algorithms which reside in the spacecraft control processor.






2.4.4.8.3: Bench-Level ValidationOnce flight control software has been written and tested at the module level, validation testing needs to be performed at the bench level.

This is usually performed through closed-loop testing, in which a real spacecraft control processor is interfaced with simulators for sensors and actuators.

In this case, the real communications interfaces between the different spacecraft units are also tested.

This is done by ensuring that the sensor/actuator simulators are implemented in hardware platforms which replicate the real spacecraft’s electrical interfaces.






2.4.4.8.4: Satellite-Level TestingSatellite-level testing for the ACS is at the level of checkouts rather than design verification.

Functional checks are performed on the various units before, during, and after each phase of spacecraft testing.

In addition to this, end-to-end polarity checks, unit alignments, and offset verifications are performed in order to verify that all equipment has been properly installed.

All redundant units are switched in and out of the operational configuration to verify the cross-strapping.

In particular, it is important to recheck alignments following major environmental tests, such as shock, vibration, and thermal vacuum, to ensure that the units are still within their fixed error tolerances.






2.4.4.8.5: In-Orbit TestingFinally, with regard to in-orbit testing, most tests are functional in nature.

Of course, since most ACS capabilities need to have been exercised in order to reach the on-station location, the in-orbit ACS test is relatively brief.

Cross-strapping capability is once again verified, and alignments of the various units are derived.

Antenna pattern measurements provide fixed error values at the antenna for which the ACS can then perform a bias correction.

After several weeks to a few months of observation, it is usually possible to derive values for key ACS characteristics, such as overall pointing performance and momentum buildup, in order to compare them to the values which were predicted by analysis.



Part OutlineIntroduction

Rocket Science

Design Approach

Technology

Risk Assessment




2.4.5: Propulsion


2.4.5.2 Functional Description

2.4.5.3 Rocket Equation

2.4.5.4 Typical Propellant Budget

2.4.5.5 Design Parameters and Processes

2.4.5.6 Typical Propulsion Subsystem

2.4.5.7 Electro-Magnetic Thrusters

2.4.5.8 Electro-Static Thrusters

2.4.5.9 Electro-Thermal Thrusters

2.4.5.10 Risk Areas, Impact, and Mitigation Plans




Part 5: Propulsion


IntroductionIn Conjunction with the Attitude Control Subsystem, the Propulsion Subsystem provides for control of:

• Orbit Insertion

• Orbit Maintenance

• Orbit Attitude Control

• Station Relocations

• End Of Life Deorbiting




Typical Transfer Orbit Functions Typical On Station Functions



Part 5: Propulsion

Figure 2.4.5.1a Typical Transfer Orbit Functions* Figure 2.4.5.1b Typical On Station Functions**



Functional DescriptionOrbit Insertion functions consist of:• Satellite attitude, stability and orbit adjustment during transfer

orbit.

• Corrections for reorientation, spin rate, orbit dispersion (an error associated with achieving the desired orbits), East and West drift, eccentricity and inclination during drift orbit and station acquisition.

Orbit Maintenance & Attitude Control consists of: • Orbit inclination, longitudinal position, drift, eccentricity, angular

momentum and Roll and Yaw attitude corrections on station.

2.4.5.2. Functional Description


Part 5: Propulsion



EQ. 2.4.5.3a, Thrust


Rocket ScienceThe basic rocket theory is based on Newton’s Third Law: “Every action has an equal and opposite Reaction”.

The two major Rocket Performance Parameters are the Thrust (F) and The Specific Impulse (Isp) .

• The Thrust is the amount of Force applied to the rocketF m.Ve [N]

m is Mass flow rate of the propellant [kg/sec]

Ve is Propellant Exhaust velocity [m/sec]

• The Specific Impulse is the ratio of the Thrust to the Weight flow rate of the propellant:

Isp F/m.g [sec]

g is 9.807 [m/sec2]

2.4.5.3. Rocket EquationPart 5: Propulsion

EQ. 2.4.5.3b, Specific Impulse

Figure 2.4.5.3a Thrust*



Rocket Science

2.4.5.3. Rocket Equation


Part 5: Propulsion

Figure 2.4.5.3b Rocket Science, Propellant System*



An Example of a Spacecraft Propellant BudgetAssumed Dispersion Magnitude = 3.00 SIGMA Calculated On Orbit Operational Lifetime= 10.34 yearsAssumed Initial orbit Parameters : Inclination = 25.70 degrees; Radius of Apogee = 27437.16 km; Radius of Perigee = 6544.00 km

Kinematics Burn Efficiency Effective Change in mass or Propellant RemainingManeuver Description Delta V Length Isp Propellant Used Remaining Mass

[m/sec] [min] [0] [sec] [Kg] [Kg] [Kg]Payload Lift off Mass 2215.80 3621.6

Spacecraft Separated Weight: (Payload Mass - LV adapter) 61.00 2215.80 3560.6ORBIT INSERTION FUNCTIONS

1st Sub-synchronus Attitude Control 2.30 2213.50 3558.3Liquid Apogee Motor Firing 1 16.18 1.95 0.99 311 18.99 2194.51 3539.3Liquid Apogee Motor Firing 2 175.9 21.07 0.9758 311 203.20 1991.31 3336.1Post Firing Attitude Control 1.60 1989.71 3334.52nd Sub-synchrous Attitude Control 1.00 1988.71 3333.5Liquid Apogee Motor Firing 3 175.9 19.81 0.9778 311 191.00 1797.71 3142.5Post Firing Attitude Control 1.20 1796.51 3141.3Transfer Orbit Attitude Control 1.40 1795.11 3139.9Liquid Apogee Motor Firing 4 960.7 88.49 0.9938 311 853.41 941.70 2286.5Post Firing Attitude Control 2.40 939.30 2284.1Transfer orbit Attitude Control 1.00 938.30 2283.1Liquid Apogee Motor Firing 5 744.06 51.38 0.9979 311 495.57 442.73 1787.5Post Firing Attitude Control 5.90 436.83 1781.6Transfer Orbit Attitude Control 1.00 435.83 1780.6Liquid Apogee Motor Firing 6 71.03 4.25 1 310 41.17 394.66 1739.5Post Firing Attitude Control 0.80 393.86 1738.7PRE ON-STATION Attitude Control 2.30 391.56 1736.4Liquid Apogee Motor Firing 7 3.57 2.52 0.9334 292 2.32 389.24 1734.0ON STATION MAINTENACE and ATTITUDE CONTROL FUNCTIONS

DISPERSION CORRECTION 60.30 328.94 1673.7In Orbit Test Attitude Control 0.80 328.14 1672.9In Orbit Tests Station Keeping (SK) 4.6 3.1 0.9439 295 2.82 325.33 1670.1North South SK FIRST SIX YEARS 258.06 165.84 0.9439 295 151.06 174.27 1519.1East West SK FIRST SIX YEARS 6 3.71 0.9334 292 3.41 170.86 1515.7

2.4.5.4. Typical Propellant Budget


Part 5: Propulsion



List aplicablespacacraft Functions

Orbit InsertionOrbit mainteneceAttitude Control

Determine TotalImpulse for Attitude

control.

Determine delta Vand Thrust level

constraints for orbitinsertion andMaintenace

Determine Thrustlevels for control

authority, duty cyclesand mission life

requirements

Determine propulsionsubsystem options

- combined or separate- Low or High thrust- Liquid, solid, electricor plasma

Deterime level ofredundancy and

overall configurationfor each option

Estimate Key parameters foreach option

- Effective Isp and Thrust fororbit and attitude control- Propellant mass andPressurant Volume

Estimate total massand power for each

option

Qualify hardware atcomponent andsubsytem level

Finalize design andprocure/manufacture

equipment

Inetgrate intospacecraft system

level AssemblyIntegration and test

Program

Are requirementsmet? Yes

NoNo

No

Propulsion Subsystem Design Process

2.4.5.5. Design Parameters & Process


Part 5: Propulsion

Figure 2.4.5.5a Propulsion Subsystem Design ProcessImage Courtesy of Telesat Canada



2.4.5.5. Design Parameters & Process


Part 5: Propulsion

Figure 2.4.5.5b Design Parameters & Process*



Chemical ThrustersChemical Spacecraft Thrusters could be either mono propellant or bi-propellant.

Bipropellant thrusters provide higher Isp than mono-propellant thrusters.


Part 5: Propulsion

Figure 2.4.5.6b Liquid Apogee Engine

Figure 2.4.5.6a Magellan Rocket Engine Module

Photo Courtesy of Telesat Canada

Courtesy of General Dynamics (Space

Systems)

2.4.5.6. Typical Propulsion Subsystem





Part 5: Propulsion

Figure 2.4.5.6.1a ACE Propulsion System

2.4.5.6.1: Typical Bipropellant System

Courtesy of General Dynamics (Space Systems)





Part 5: Propulsion

Figure 2.4.5.6.1b Typical Bipropellant System*

2.4.5.6.1: Typical Bipropellant System



2.4.5.6.2: Schematic• Fuel, oxidizer and

pressurant are loaded individually

• Fuel and oxidizer are hypergolic (i.e. they burn when mixed)

• Cannot launch with system pressurized because of tank design.



Part 5: Propulsion

Figure 2.4.5.6.2 Schematic of Galileo Propulsion System*



2.4.5.6.3: Monopropellant Vs Bipropellant Systems



Part 5: Propulsion

Figure 2.4.5.6.3a Typical Monopropellant System* Figure 2.4.5.6.3b Typical Bipropellant System**



2.4.5.6.4: Pressurant SystemsLiquid Apogee Motors operate in a regulated constant pressure mode to maintain high efficiency.

On-orbit chemical thrusters operate in a blow-down mode.



Part 5: Propulsion

Figure 2.4.5.6.4 Pressurant Systems*



Electro-Magnetic Thrusters

2.4.5.7. Electro-Magnetic Thrusters


Part 5: Propulsion

Figure 2.4.5.7 Hall Current Thruster




Electro-Static Thrusters

2.4.5.8. Electro-Static Thrusters


Part 5: Propulsion

Figure 2.4.5.8a Electro-Static Thruster Schematic

Image Courtesy of the Associated

Plasma Laboratory (LAP), National

Space Research Institute (INPE),

Brazil



Ion Engine



Part 5: Propulsion

Electro-static thrusters are also known as Ion Engines.

Figure 2.4.5.8b Ion Engine Functional Diagram

Image Courtesy of the Associated Plasma Laboratory (LAP), National Space Research Institute (INPE), Brazil



Ion Engine



Part 5: Propulsion

Figure 2.4.5.8c Ion Engine Firing

Image Courtesy of the Associated Plasma Laboratory (LAP), National Space Research Institute (INPE), Brazil



Electro-Thermal Thrusters

2.4.5.9. Electro-Thermal Thrusters


Part 5: Propulsion

Figure 2.4.5.9bEHT(Electrically Heated

Thrusters) SystemFigure 2.4.5.9aMR-510 Arcjet System with Power Processor, 4 arcjets, and 4 cables

Two sub types of commercially available Electro-Thermal thrusters.

Courtesy of General Dynamics (Space

Systems)



Figure 2.4.5.9c MR-510 Arcjet Firing



Risk Areas, Impact & Mitigation Plans• Rupture of fuel systems

• Impact: Catastrophic failure• Mitigation Plans: Rigorous Qualification & Test Plan

• Liquid apogee motor under-performance• Pyro valve failures• Latch valve leakage• Thruster valve failures• Thruster failures

For the last five risk areas• Impact: Degraded mission• Mitigation Plans: Rigorous Qualification & Test Plan and stringent

workmanship processes

2.4.5.10. Risk Areas, Impact and Mitigation Plans


Part 5: Propulsion





Mechanical SubsystemsPart 6




2.4.6: Mechanical Subsystems

2.4.6.1 Structure

2.4.6.2 Mechanisms






2.4.6.1 Structure

2.4.6.1.1: IntroductionThe Mechanical Subsystem provides stable mechanical support to other subsystems and components.

Sustains all loads and environments during:• Fabrication & Transportation• Launch & Transfer Orbit• On orbit control maneuvers • Attitude control failures and

recoveries

The Mechanical Subsystem maintains dimensional stability for sensitive payload equipment.

Figure 2.4.6.1.1a Shuttle Launch


Part 6: Mechanical Subsystems

Figure 2.4.6.1.1b FLTSATCOM Structure*

Photo Courtesy of Telesat Canada



2.4.6.1.2: Design Parameters & Process


Figure 2.4.6.1.2 Design Process


2.4.6.1 Structure



2.4.6.1.3: Risk Areas, Impact & Mitigation PlansRisk Areas, Failures Occur due to:

• Inadequate design margins

• Poor workmanship

• Poor processes

• Overload situations

Impact of Failures & Mitigation Plans:• Mission Catastrophic Failures:

• As when the primary structure is affected, such as the main thrust tube or the launch vehicle interface region

Mitigating Plan: Robust Design margins, Rigorous Validation Plan



2.4.6.1 Structure



Impact of Failures & Mitigation Plans (cont):• Performance Degradation Type Failures

• Reduced RF performance when payload appendages are affected

Mitigating Plan: Adequate Redundancy, Rigorous Validation Plan

• Shorter life when bus appendages are affected

Mitigating Plan: Adequate Redundancy, Rigorous Validation Plan


2.4.6.1.3: Risk Areas, Impact & Mitigation Plans


2.4.6.1 Structure



2.4.6.2.1: IntroductionLaunch envelope constraints and environments dictate the need for deployment mechanisms.

Mechanism & actuator key requirements:• Minimum torque margin of 200 percent (3 to 1 ratio)

• Operational capability at least 1.5 times the number of cycles anticipated through ground test and on-orbit life

• Failure tolerant (w.r.t. primary initiator actuation)

• Capability to allow for end-to-end verification in a 1 g environment


2.4.6.2 MechanismsPart 6: Mechanical Subsystems



2.4.6.2.2: Design ApproachDesign considerations include:

• Minimize static and dynamic disturbances during launch, deployments and on-orbit operations

• Redundancy (primarily initiators and electronics)

• Clearance and accessibility issues due to tolerance stack-up (the incremental buildup of dimensional errors as parts are assembled), thermal environments and local obstructions

• Containment of any solid debris resulting from actuation and operation

• Operational life (brushes, seals, labyrinths, lubricants)

• Verification of design and workmanship process





2.4.6.2.3: TechnologyMechanism technology falls in two categories:

• Mechanized; motorized drives generally used for highly-cyclical applications such as:

• Antenna Pointing Mechanisms

• Solar Array Pointing and Tracking

• Attitude Control Reaction Wheels

• Gimbals for Propulsion units

• Passive; Springs and hinges with dampers, generally used for low cyclical applications such as:

• Antenna and Solar Array retention

• Antenna and Solar Array deployments

• Spacecraft launch vehicle separation systems





2.4.6.2.4: Risk, Failures & Mitigation PlansDeployable assembly “Hangs Up”:

• Impact: Can be a catastrophic failure

• Mitigation Plans: Rigorous qualification, robust torque margins and meticulous workmanship

Continuous use over single point or thermal design limitation:• Impact: Degraded mission

• Mitigation Plans: Rigorous qualification & test plans and stringent workmanship processes







ThermalPart 7




2.4.7: Thermal


2.4.7.2 Thermal Environment

2.4.7.3 Thermal Design Process

2.4.7.4 Thermal Design Approach

2.4.7.5 Thermal Key Items

2.4.7.6 Thermal Hardware

2.4.7.7 Thermal Equipment

2.4.7.8 Thermal Controller Functions

2.4.7.9 Failures, Degradation and Margins



Slide Number 178Rev -, July 2001 Vol 2: Communication Satellites, Sec 4: Satellite Bus/Platform Subsystems

2.4.7.1. IntroductionPart 7: Thermal

Figure 2.4.7.1 Thermal System


Drawing Drawing Courtesy TelesatCourtesy Telesat



2.4.7.2. Thermal Environment


Part 7: Thermal

Figure 2.4.7.2 Thermal Environment*



2.4.7.3. Thermal Design Process

Radia tion P rope rtyVa lues

Spacecra ft O rien ta tionand A ttitude

Spacecra ft G eom etry

E lectrica l P ow erD issipa tion

T herm ophyscia l P roperty Va lues

Requ irem ents- T em pera ture L im its- S urvivab ility

Rad ia tion C om puter P rogram

Therm al Ana lyzer C om puterP rogram

(S pacecra ft The rm al M ath M odel)

The rm al C ontro l H /W E lem ents(R adia to rs , Louvers, H eaters,

T herm al B lankets)

P red ic ted T herm a l P erform ance

Com parison

R adia tion E xchange Factorsand V iew Facto rs

Rad ia tion Absorbed onExterna l Surfaces

Com ponent-Leve l T ests

System -Leve l T ests


Part 7: Thermal

Figure 2.4.7.3 Thermal Design Process



Part 7: Thermal


2.4.7.4. Thermal Design Approach

Figure 2.4.7.4 Thermal Design Approach




Slide Number 182Rev -, July 2001Figure 2.4.7.5 Thermal Key Items

Slide Number 182


2.4.7.

5.

Therm

al Key

Items



Thermal Hardware


Part 7: Thermal

2.4.7.6. Thermal Hardware

Figure 2.4.7.6 Two Views of the LM A2100 S/C Photos Provided by Lockheed Martin



• Heat pipe network embedded in equipment panels• North, South, Earth

• Enhanced thermal conductance joint between North-South panels

• OSR radiator surfaces on North and South panels

• Standard multi-layer insulation blankets• Black Kapton outer layer

• Germanium shields for RF sensitive items

• Heater controllers regulate temperatures

Thermal Equipment


Part 7: Thermal

2.4.7.7. Thermal Equipment



Computer controlled heaters:• Set points are ground commandable

• Temperature sensor groupings are ground changeable

• Control type (min., max., duty cycle, difference) are ground changeable

• Total ground override capability exists

• Controller commands every heater on/off state, if enabled

Bimetallic thermostats switch heater circuits on and off in response to temperature changes.

Thermal Controller Functions


Part 7: Thermal

2.4.7.8. Thermal Controller Functions



• Maintain all component temperatures within operational limits throughout mission life as thermal surface properties degrade, potentially resulting in failure.

• Fully redundant heater circuits with single-point-failure tolerance

• Bimetallic thermostats are configured in dual or quad redundant configurations

• 25% margin on heater output capability, or a 10oC heater sizing margin

• 5oC uncertainty margin added to raw temperature predictions

Failures, Degradation & Margin


Part 7: Thermal

2.4.7.9. Failures, Degradation & Margin

satellite bus platform subsystems

Engineering