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Mark ZimmermanTypewritten Texthttp://www.ATIcourses.com/schedule.htmhttp://www.aticourses.com/Fundamentals_Of_Space_Systems_Space_Subsytems.htm
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Mark ZimmermanTypewritten TextCourse Schedule:Course Outline:
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Mark ZimmermanTypewritten TextSPACE SYSTEMS AND SPACE SUBSYSTEMS - FUNDAMENTALS
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Mark ZimmermanTypewritten TextInstructor:
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Mark ZimmermanTypewritten TextDr. Vincent L. Pisacane
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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
Pisacane, 2013
RISK MANAGEMENT NASAs 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
Pisacane, 2013
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
ns 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)
Pisacane, 2013
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 (nm) 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
Pisacane, 2013
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 nozzles 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
Pisacane, 2013
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: 4 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
10x1170.0
c
DfG
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
dB12.01.012dB12L2
2
dB3
dB3
i
2
i2
dB3
l q
q
q
Pisacane, 2013
THERMAL ANALYSES Analysis Process
Pisacane, 2013
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