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©Pisacane, 2013

CASSINI-HUYGENS Interplanetary Mission to Saturn

• Saturn surrounded by Rings and 62 Moons • Cassini launched in October 1997 arrived at Saturn June 2004 • The mission has been extended through September 2017

©Pisacane, 2013

Planned 6 Oct

Planned 20 June

Planned 16 August

@ 1,170 km

Planned 1 Dec

Planned 21 April

Planned 30 Dec 2000

Planned 1 July

CASSINI-HUYGENS Trajectory

©Pisacane, 2013

NEAR Configurations

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RISK MANAGEMENT NASA’s Approach to Risk Management

NASA identifies two activities critical to risk management

Risk-Informed Decision Making (RIDM) – Selection of alternatives based on assessment of requirements including risk

Continuous Risk Management (CRM) – Systematic identification, assessment, and management of all risks

From: NASA Risk-Informed Decision Making Handbook, NASA/SP-2010-576 Version 1.0 Apr 2010

©Pisacane, 2013

SYSTEM DEVELOPMENT NASA Project Life Cycle Reviews

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SYSTEM TESTING Sample NASA Payload Test Requirements

From: NASA-STD-7002A ─ Payload Test Requirements

©Pisacane, 2013

SPACECRAFT FAILURES NOAA Spacecraft Radiation Induced Failures May 1998

Data from NOAA GOES (Geostationary Operational Environmental Satellite) constellation

Equator-S failure attributed to latch-up in central processor as result of a week or more of elevated relativistic electron (top figure)

POLAR processor loss of 6 hours of data attributed to

single-event upset (SEU) in processor from increased proton flux (bottom figure)

Galaxy 4 processor failure likely caused, by the energetic electron environment most likely due to deep dielectric, (or bulk) charging (top figure)

Space Environmental Conditions During April and May 1998: An Indicator

for the Upcoming Solar Maximum

D.N. Baker, J.H. Allen, S. G. Kanekal, and G.D. Reeves

©Pisacane, 2013

The standard life test for flight hardware parts is the dynamic (power on) burn-in test for 1000 hours (41.7 d) at an ambient temperature of 125oC (257oF)

The Acceleration Factor (Af) is the test time multiplier derived from the Arrhenius equation for operation at another temperature

Activation energy (Ea) is an empirical value of the minimum energy required to initiate a specific type of failure mode that can occur within a technology type – Failure modes include: oxide defects, bulk silicon defects, mask defects, electro-

migration, and contamination

Typical values of Ea for electronic devices are 0.5-1.0 eV, typically > 0.7

Table shows acceleration factors and equivalent durations

Ea, eV

Acceleration Factors For use temperatures

Equivalent Duration, y

25oC 77oF

35oC 95oF

45oC 113oF

25oC 77oF

35oC 95oF

45oC 113oF

0.5 133 71 39 15 8 2

0.6 353 165 81 40 19 9

0.7 938 387 169 107 44 19

0.8 2,492 907 352 284 103 40

0.9 6,624 2,125 732 756 242 84

1.0 17,607 4,979 1,524 2,008 568 174

FAILURE ANALYSES Burn-in Tests at Elevated Temperatures

testuse

af

T

1

T

1

k

EexpA

Parameters Ea = Activation Energy of the failure mode, eV k = Boltzmann's Constant, 8.617 x 10-5 eV K-1

Tuse = Use Temperature, K Ttest = Test Temperature, K

testusef

testusef

T Tif 1AT Tif 1A

©Pisacane, 2013

Failure Modes, and Effects Analysis (FMEA)

System: Part Name Reference Drawing Mission

Date Sheet X of X

Compiled by: XXXX Approved by: XXXX

Item

Function or

Require-ment

Potential Failure Modes

Potential Causes

of Failure Mode

Potential Effects of Failure Mode

Detection and

Mitigating Factors

O c c u r r e n c e

D e t e c t i o n

S e v e r i t y

RPN Actions

Recomm- endations

Respon- sibility Local

Effects

Inter-mediate Effects

End Effects

Battery

Provide adequate

relay voltage

Fails to provide

adequate power

Voltage drops to

zero

Battery plates

shorted

Instrument not

functional

Mission Aborted

Test battery prior to launch

4 4

0.5+

0.3 X5 = 4

64 XXX XXX

…

FAILURE IDENTIFICATION Sample FMEA Worksheet Failure Modes and Effects Analysis (FMEA)

Typical FMEA worksheet is illustrated below for a spacecraft battery

©Pisacane, 2013

RELIABILITY, AVAILABILITY, MAINTAINABILITY, and SAFETY Derating Introduction

Derating increases the margin of safety between operating stress level and actual failure level for the part, providing added protection from unanticipated anomalies

Derating is employed in electrical and electronic devices, wherein the device is operated at lower than its rated maximum power dissipation, taking into account – Case/body temperature – Ambient temperature – Type of cooling mechanism

When derating, the application engineer applies a recommended derating factor

bases on the part specifications and operating environment

For microcircuits, major derating factors are – Supply voltage – Power dissipation – Signal input voltages – Output voltages – Output currents

©Pisacane, 2013

Series redundancy – Reliability Rs of the series chain is given

by

– If all components have the same reliability then Ri = R and

Parallel redundancy – The reliability of a parallel configuration

if only one device is needed is

– If all component s have the same reliability then Ri = R and

RELIABILITY, AVAILABILITY, MAINTAINABILITY, and SAFETY Calculating Reliabilities

n

sRR

n21

n

1iis

R1R1R11R11R

n321

n

1iis

RRRRRR

n

s R11R

©Pisacane, 2013

CELESTIAL MOTION Principal Motion of the Celestial Ephemeris Pole

(more accurate number is 25,780 yrs)

(average of 50.26 sec of arc per year or 0.1376 sec arc per day)

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COORDINATED UNIVERSAL TIME (UTC) Variation in the Length of Day 2/2

From: http://www.ucolick.org/~sla/leapsecs/dutc.html

25

©Pisacane, 2013

REFERENCE SYSTEM Geometrical Transformation Between GCRS and ITRS

Figure shows transformation between terrestrial (ITRS) to celestial (GCRS) taking into account (1) Pole Movement, (2) Earth Rotation , (3) Precession and Nutation – GCRS= Geocentric Celestial Reference System – ITRS = International Terrestrial Reference System – CIP = Celestial Intermediate Pole, instantaneous Earth spin axis – CTP = Conventional Terrestrial Pole, reference pole in ITRS (now average of pole positions from 1900 to 1905)

Modifiedfrom:ESA,http://navipedia.org/index.php/Transformation_bet

ween_Celestial_and_Terrestrial_Frames

©Pisacane, 2013

GRAVITATIONAL POTENTIAL Geometrical Representation of Spherical Harmonics

m = 0 no longitudinal

variation

n ≠ m and m ≠ 0 Tessarae (Tiles)

n = m no latitudinal

variation

n = 2, m = 2 n = 3, m = 3 n = 5, m = 0 n = 4, m =3

Pn,m(Cos q) Cos m(l – l n,m) has − (n–m) sign changes or zeros 0 q p (latitude of 180 degrees

− 2m zeros in interval 0 l < 2p (longitude of 180 degrees)

©Pisacane, 2013

TRAJECTORY PERTURBATIONS Mars Global Surveyor Aerodynamic Braking

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ROCKET PROPULSION Specific Impulse vs Thrust

From: http://dawn.jpl.nasa.gov/mission/images/CR-1845.gif

NH3 = Ammonia

N2H4 = Hydrazine

Grayed area are

realized

characteristics

©Pisacane, 2013

ROCKET PROPULSION de Laval Nozzle

The function of the nozzle is to convert the chemical-thermal energy produced in the combustion chamber into kinetic energy

Thrust is the product of mass time velocity so a very high gas velocity is desirable

The nozzle converts slow moving, high pressure, and high temperature gas in the combustion chamber into high velocity gas of lower pressure and temperature at the nozzle’s exit

De Laval nozzles consist of a convergent and divergent section

The section with minimum area is the nozzle throat

The nozzle is usually made long enough and the exit area large enough to reduce the high pressure in the combustion chamber to the ambient pressure at the nozzle exit to create maximum thrust

Typical DeLaval nozzle

T = temperature

p = pressures

v = speed

M = Mach number

From: http://en.wikipedia.org/wiki/Rocket_engine

©Pisacane, 2013

LAUNCH FLIGHT MECHANICS Available Launch Inclinations in the United States

37

114

©Pisacane, 2013

COLD GAS PROPULSION SYSTEMS Typical Cold Gas System Implementation

L L L L

GN2

L

P T

F

P

T

P

T

P

T

P

T

Latch Valve

Temperature sensor

Pressure Sensor

Pyrovalve

normally open

Pyrovalve

normally closed

Burst Valve

Latch Valve

Gas Regulator

Filter

Service valve

Access Port

L

P

T

F

NO

NC

L

Check Valve, arrow

direction of flow

Typical cold gas thruster

Propellants

Air, Carbon Dioxide,

Helium, Hydrogen,

Methane, Nitrogen, Freon

©Pisacane, 2013

LIQUID PROPULSION SYSTEMS Messenger Spacecraft Dual Mode Propulsion

S Wiley, K Dommer, L Mosher, Design and development of the

Messenger propulsion system, AIAA, PRA-053-03-14 July 2003

Illustrates the Messenger spacecraft propulsion system with 17 thrusters

Bipropellants ─ Hydrazine (N2H4) and Dinitrogen Tetroxide (N2O4)

Monopropellant ─ Hydrazine (N2H4)

©Pisacane, 2013

TRANSFER TRAJECTORIES Apollo 13 Circumlunar Free-Return Trajectory

CSM ─ Command Service Module, DPS – Descent Propulsion System EI – Entry Interface GET ─Ground Elapse Time LM – Lunar Module MCC ─ Mid-Course Correction PC – Pericynthion (closest point to moon) S-IV4B – Saturn IVB SM – Service Module TLI –Trans Lunar Injection

JL Goodman , Apollo 13 Guidance, Navigation, and Control Challenges AIAA SPACE 2009 Conference & Exposition, Sept 2009, Pasadena,, AIAA 2009-6455

©Pisacane, 2013

OVERVIEW Attitude Control Schematic

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ATTITUDE KINEMATICS Quaternion Mathematics 1/2

Addition and subtraction – Elements are added or subtracted

Multiplication – Not communicative, Q1Q2 ≠ Q2Q1 – Multiple each component

where

Equivalent quaternions – Reversing signs on all 4 elements yields an equivalent quaternion

─Q = Q

time

s 1 i j k

1 1 i j k

i i ─

1 k ─j

j J ─

k ─1 i

k k j ─i ─1

4,23,22,21,24,13,12,11,121

qkqjqiqqkqjqiqQQ

4,24,13,23,12,22,11,21,1 qqkqqjqqiqq

4,23,22,21,24,13,12,11,121

qkqjqiqqkqjqiqQQ

©Pisacane, 2013

ATTITUDE SENSORS ADCOL Two-Axis Digital Sun Sensor System

http://adcole.com/two-axis-dss.html

Two-Axis Digital Sun Sensor System – No of measurement axes:

• 2 each sensor) – Number of sensors

• 5 typical per electronics • 1 to 8 sensors can also be used • Electronics selects sensor that has

sun in field of view

Heritage – Many systems flown with 1 to 8 sensor

heads per processing electronics

Parameters – Field of view: ±64° x ±64°

• Note: 4π steradians (full sphere) coverage can be achieved with 5 sensors.

– Accuracy: ±0.25° (transition accuracy). – Least Significant Bit Size: 0.5° Sign bit

Most significant bit

Least significant bit Interpolating bits

©Pisacane, 2013

INTRODUCTION Function and Components of Spacecraft Power System

Power system functions – Supply electrical power to spacecraft loads – Distribute and regulate electrical power – Satisfy average and peak power demands – Condition and convert voltages – Provide energy storage for eclipse and peak demands – Provide power for specific functions, e.g., firing ordinance for mechanism

deployment – Ensure power to critical loads during critical phases and spacecraft anomalies – Ensure power for mission duration

Primary Power

Source

Energy

Conversion

Power

Regulation

Power

Distribution

Power

Regulation

Energy

Storage

Power

Regulation

Critical

Loads

Non-Critical

Loads ?

©Pisacane, 2013

SECONDARY BATTERIES Candidate Technologies

http://www.clyde-space.com/products/spacecraft_batteries/useful_info_about_batteries/secondary_batteries

©Pisacane, 2013

SOLAR ARRAYS Solar Array Construction

Cells connected in series to achieve desired voltage

Cells connected in parallel to achieve desired power

Arrays organized to minimize current loops that result in dipole moment

©Pisacane, 2013

OVERVIEW NEAR Spacecraft Spacecraft Communication System

From: RS Bokulic, MKE Flaherty, JR

Jensen, and TR McKnight, The NEAR

Spacecraft RF Telecommunications System,

Johns Hopkins APL Technical Digest, Vol

19, No 2 (1998)

Transponder unifies a number of communication functions - receiver,

command detector, telemetry modulator, exciters, beacon tone

generator, and control functions

Diplexer is a device that can split and combine audio and video

signals

©Pisacane, 2013

ANTENNAS Typical Parabolic Antenna Pattern

©Pisacane, 2013

LINK ANALYSIS Example Link Analysis

dB3.38683

400

1000

1011038.1

63.097.01058.15.076.768.020

T

G

kR

LLEIRP

N

E623

116

s

RA

b

a

LossesOther

a

s

0

b

Transmitter power 20 W +13.0 dBW

Spacecraft cable loss 1dB ─1 dB

Antenna boresight

gain 76.76 +18.9 dB

EIRP 30.9 dBW

Antenna beamwidth 3 dB ─3.0 dB

Space loss at 10o

elevation @ 3000 km 1.58 x 1016 ─162.0 dB

Pointing error, 0.1 BW 0.12 dB ─0.12 dB

Atmospheric loss 0.1 dB ─0.2 dB

Receiver G/T 1000/400 K-1 4.0 dbK-1

Boltzmann constant,

k

1.38x10-23

JK-1

+228.6 dB J-

1K

Bit rate 106 bps ─60 dB s

Receiver Eb/No 38.2dB

76.7610x3

10x11π70.0

c

πDfG

2

8

92

boresight

Spacecraft antenna diameter = 1 m Frequency = 1 GHz Pointing error= 1/10 beamwidth Receiver gain = 30 dB Receiver system temperature = 400K Bit rate = 106 bps

16

2

8

962

l

s1058.1

103

1011034

c

rf4L

p

p

dB12.0

1.012dB

12L

2

2

dB3

dB3

i

2

i2

dB3

l q

qq

q

q

©Pisacane, 2013

THERMAL ANALYSES Analysis Process

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MULTILAYER INSULATION Gold and Black MLI

Gold Thermal Blanket − Outer layer is of a second surface mirror material with

high reflectivity and high emittance − Consists of multiple layers of silver coated Kapton film

that gives it a gold color − Except outer layers, all are perforated to allow entrapped

air to escape during launch and separated by a Dacron netting − Edges are finished with a tape prior to sewing − Individual blankets held together and to spacecraft by

dacron Velcro

Black Thermal Blanket − Black thermal blanket is used on the shade side of the

spacecraft − Identical to the gold blanket except for the outer layer

generally Kapton filled with carbon powder − Outer layer has a higher absorptance and lower

emittance than the gold Kapton − This layer is also electrically conductive because of

carbon fill − Grounding outer layer to the spacecraft frame dissipates

any charge build

Gold is multilayer insulation of

Cassini spacecraft; from

NASA

New Horizons spacecraft

http://www.boulder.swri.edu/pkb/ssr/ssr-

fountain.pdf

©Pisacane, 2013

DESIGN PROCESS Overall Development Flow Chart

Spin Balance and Environmental

Testing

Preliminary Launch Loads

Preliminary Natural Frequency Constraints

Thermal Analysis

Temperature Distribution

Structural Analysis Finite Element Model Dynamic Analysis Stress Analysis Thermal Distortion Assess Margins

Launch Vehicle Dynamic Model and

Forcing Functions

Coupled Launch Vehicle and

Spacecraft Dynamic Analysis

Spacecraft Dynamic

Model

Spacecraft Dynamic Response

Loads Acceleration

Functional Subsystem/Payloads

Requirements

Preliminary Spacecraft Structural

Design

Fabricate Spacecraft Structure

Launch Vehicle Constraints

Spacecraft Structural Configuration

Conceptual and

Preliminary

Design

Critical Design

Fabrication

Integration

launch

start

©Pisacane, 2013

STRUCTURAL CONFIGURATIONS Structural Categories

Structural components are categorized by the different types of requirements, environments, and methods of verification that drive their design – Primary structures are usually designed to survive steady-state accelerations and

transient loading during launch and for stiffness – Secondary and tertiary structures are usually designed for stiffness, positional

stability, and fatigue life

Primary structures: • body structure

• launch vehicle adapter

Secondary structures: • appendage booms

• support trusses

• platforms

• solar panels

• Antenna

• Extendibles

Tertiary structures: • brackets

• electronics boxes

©Pisacane, 2013

INTRODUCTION Space and Ground Based Systems

Reliability, complexity, development costs, and operational costs are affected by the partitioning of the computational load between the space and ground segment

From Wertz and Larson

©Pisacane, 2013

COMPUTER COMPONENTS Typical Spacecraft Computer Schematic

Figure is a simplified block diagram of a spacecraft computer system

One or more processing units have access through bus structures to – Read only memory, random access memory, and special purpose memory – Mass storage – Input/output ports to spacecraft subsystems and payloads – Spacecraft communication system – Numerical coprocessor to carry out floating point arithmetic faster

From Pisacane, Fundamentals of space systems, Oxford University Press,

2005

©Pisacane, 2013

FAULT TOLERANCE Summary Fault Tolerant Techniques

NMR = n-modular redundancy ECC = Error Correction Coding RESO = RE-computing with Shifted Operands; computation carried

out twice - once with usual input ─ once with shifted operands Self-purging = each module has a capability to remove itself from

the system if faulty Recovery blocks = Uses the concept of retrying the same

operation and expect the problem is resolved by the second or later tries

©Pisacane, 2013

SPACECRAFT PROCESSORS RAD6000 Processor

Characteristics − 35 Mbps at 33 MHz – Radiation Hardened 32-bit RISC – Super Scalar Single Chip CPU – 8K Byte Internal Cache – Simplex or Dual Lock-step (compares CPU

operations) – Low Power 3.3 Volt Operation – 72-bit (64 Data, 8 ECC) Memory Bus – Variable Power/Performance – Independent Fixed and Floating Point Units

Radiation Hardness Levels – Total Dose: 2x106 rads(Si) – Prompt Dose Upset: 1x109 rads(Si)/sec – Survivability: 1x1012 rads(Si)/sec – Single Event Upset: 1x10-10 Upsets/Bit-Day – Neutron Fluence: 1x1014 N/cm – Device Latchup: Immune

From Lockheed Martin Federal Systems RAD6000 Radiation Hardened 32-Bit

Processor

atc2.aut.uah.es/~mprieto/asignaturas/satelites/pdf/rad6000.pdf

COP = Common on-chip processor interface

FPGA = Field Programmable Gate Array

HMC = Hardware Management Console

RS232 = Serial binary single ended data connector

VME bus = VersaModular Eurocard bus

Dual Lock Step

A technique that achieves high

reliability by adding a second

identical processor that monitors and

verifies the operation of the system

processor

©Pisacane, 2013

INTEGRATION AND TEST PROCEDURES Integration and Test Procedure

From Spacecraft Computer Systems, JE Keesee ocw.mit.edu/courses/aeronautics-and.../l19scraftcompsys.pdf