Download - Introduction to Navigation Systems
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Introduction to Navigation Systems
Joseph HennawyComputer Engineer
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Table of Contents History of Navigation Systems.
Accelerometer Sensors Technologies (Body Speed & Acceleration).
Gyroscope Sensors Technologies (Body Attitude).
Navigation Coordinate Systems.
GEODESY & DATUMS.
INS Systems Error Analysis.
GPS/INS Systems.
Current Navigation Systems
The Future of Navigation Technologies.
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Inertial Navigation History
Inertial Guidance System of SAGEM used in the Air-Surface Medium-range missile
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Dead Reckoning
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Early Compasses
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Surveyor’s Compass--1820
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Jean Bernard Léon FoucaultOriginator of the Foucault pendulum
1819-68
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Foucault's gyroscope (1851)
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Mechanical Dead Reckoning Computer: Early 20th century
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SG-66 Guidance System for the V-2 (1944)
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Charles Stark Draper Gyroscopic Apparatus - Spinning Gyroscope
Born 2 October 1901Died 25 July 1987
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First Successful All-Inertial Navigator (1954)
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Professor Arnold Nordiseck Holding Early Electrostatically Suspended Gyroscope (1959)
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Honeywell Advertisement for Electrostatically Suspended Gyroscope, 1962
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Warren Macek of Sperry Circa 1963 Demonstrating the Ring Laser Gyro Concept
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Laser Gyro
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Tactical Grade Closed-Loop FOG• Tactical FOG IMU funded by USAF
• HG1800 FOG IMU is pin-for-pin compatible with HG1700 RLG IMU
• Goals:1 deg/hour Gyro Error1 milli-G Accel Error
• Housing identical to HG1700 IMU<35 cubic inches
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INERTIAL NAVIGATION HISTORICAL EVENTS
• Newton’s second law: circa 1688
• Leon Foucalt: demonstration of earth rotation using a gyroscope 1852 Greek: “gyro”--rotation; “skopein”--to see
• G. Trouve: Mechanical gyroscope with electric motor 1865
• Anschutz: First gyrocompass 1904
• Schuler: Pendulum/gyroscope unaffected by ship/course/speed 1908
• Boykow(Austria): Mathematics of inertial navigation 1938
• Peenemunde Group(Germany): First operating inertial guidance on V2 1942• Autonetics: Under the ice Nautilus crossing of North Pole 1958• Autonetics: Transcontinental purely inertial flight 1958• AC-Delco, Litton, Honeywell, Sperry, Singer-Kearfott, Sagerm(French): 1960’s
Military bombers, ships, fighter, ballistic missiles• MIT/Delco: Apollo guidance system 1969• Honeywell: Electrically suspended gyro navigator 1967• Sperry: First ring laser gyro 1963
IVmdtdF
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INERTIAL NAVIGATION HISTORICAL EVENTS(2)
•Various: First inertial navigation systems in commercial aircraft late 60’s
• RLG: based strap down systems on commercial aircraft early 80’s
• RLG: based strapdown systems in military mid 80’s
• First Fiber Optic Gyro Based inertial systems early 90’s
• First Embedded GPS-INS systems early 90’s
• Low cost tactical microelectromechanical sensors(MEMS) NOW
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Accelerometers
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FORCER VERTICALPIVOT
PICKOFF
AMPLIFIER
Simple Pendulum Accelerometer
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Torque Balance Pendulous Accelerometer Schematic
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EMERGING ACCELEROMETER TECHNOLOGY APPLICATIONS
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MEMS/MOEMS
Mech.
SiliconQuartz
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WSN-7 Accelerometer
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Physical•Weight 1.54 pounds (700 grams)•Size 3.5 inches (8.9 cm) diameter by 3.35 inches (8.5 cm) high•Power 10 watts steady-state (nominal)•Cooling Conduction to mounting plate•Mounting 4 mounting bolts – M4
Activation Time 0.8 sec (5 sec to full accuracy)
Performance – Gyro•Bias Repeatability 1°/hr to 10°/hr 1σ•Random Walk 0.04 to 0.1°/√hr power spectral density (PSD) level•Scale Factor Stability 100 ppm 1σ•Bias Variation 0.35°/hr 1σ with 100-second correlation time•Nonorthogonality 20 arcsec 1σ•Bandwidth > 500 Hz
Performance – Accelerometer•Bias Repeatability 200 µg to 1 milli-g, 1σ•Scale Factor Stability 300 ppm 1σ•Vibration Sensitivity 17 µg/g2 1σ•Bias Variation 50 µg 1σ with 60-second correlation time•Nonorthogonality 20 arcsec 1σ•White Noise 50 µg /√Hz PSD level•Bandwidth > 500 Hz
Operating Range•Angular Rate ±1000°/sec•Angular Acceleration ±100,000°/sec/sec•Acceleration ±40g•Velocity Quantization 0.00169 fps•Angular Attitude Unlimited
Reliability (predicted) 23,345 hours MTBF (30°C missile launch environment)
Input/Output RS-485 Serial Data Bus (SDLC)
Data Latency < 1msec
Environmental•Temperature -54°C to +85°C operating•Vibration 11.9g rms – performance
17.9g rms – endurance•Shock 90G, ms terminal sawtooth
Summary of Ln-200 IMU Characteristics
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Accelerometer Name $2K(1)Part of System Name $2Ksystem(1)Where Found IMU Performance vs. Cost
Velocity Random Walk 0.60 (meters/sec)/√(rt-hr)Bias 1000 micro-gMisalignment 412 arcsecScale Factor 500 ppmSecond Order Scale Factor Non-Linearity 60 micro-g/g2
Additional Terms
Notes
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Accelerometer Name $20KPart of System Name $20KWhere Found IMU Performance vs. Cost
Velocity Random Walk 0.03 (meters/sec)/√(rt-hr)Bias 100 micro-gMisalignment 10.3 arcsecScale Factor 10 ppmSecond Order Scale Factor Non-Linearity 3 micro-g/g2
Additional Terms
Notes
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Velocity Random Walk 0.0003 (meters/sec)/√(rt-hr)Bias 100 micro-gMisalignment 3 arcsecScale Factor 100 ppmSecond Order Scale Factor Non-Linearity 0.5 micro-g/g2
Additional Terms
Notes
Accelerometer Name $100KPart of System Name $100KWhere Found IMU Performance vs. Cost
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Gyroscopes
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INERTIAL ROTATION SENSOR TECHNOLOGY
E;\Courses\Gyros
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INERTIAL SENSOR APPLICATION
1 5 25 125 625 31251e-005
0.0001
0.001
0.01
0.1
1
10
WEIGHT
SEN
SOR
PER
FOR
MA
NC
E (d
eg/h
r)
TACTICALMISSILES
GBI / ASAT
RV
MEDIUM ACCURACYAIRCRAFT
COMMERCIALAIRCRAFT
HIGH ACCURACYAIRCRAFT
ICBM SDI POINTING
SURFACESHIP
SUB
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Inertial Sensor Technology Comparison
Inertial Acronym Definitions
ESG Electrostatic GyroFOG – Fiber Optic GyroHRG – Hemispherical Resonator GyroMS – MultisensorMEMS – Micromachined Electromechanical SensorQRS – Quartz Rate SensorRLG – Ring Laser Gyro
ESGRLG
FOGMS
QRSHRG
MEMS
GyroDrift (deg/hr)
Submarines
Strategic MX
Surface ShipsAircraft
Cruise Missles
UAVs
Precision Guided Munitions (PGM)SCUD-B
NO-DONG
Unguided
GGPFOG
EGISLAM-ER
SLAMF-18
TLAM JDAM AGM-L30 EKGM
All sensor perf ranges are estimates based on available information
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Honeywell Gyro Technology Heritage1920 1960 1970 1980 2000 202020101990
Iron Gyros Optical Gyros MEMS
Optical Gyros Ring Laser Gyro Fiber Optic Gyro Digital Output Moderate Cost
Iron Gyros Spinning Wheel Analog Output High Cost
MEMS Gyros Silicon Sensor Analog or Digital
Output Low Cost
World’s first application gyros invented by Elmer Sperry
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IMU Product Evolution OverviewRLG FOG MEMS
• EGI • GGP • Future• MAPS • PSN Growth• Digital Laser Gyro
• HG1700 • HG1800 • HG1900
- in production - developmental - in development
TacticalGrade IMUs
NavigationGrade
Systemsand
Components
EGI Embedded GPS Inertial Integrated System - aircraft, et alMAPS Modular Azimuth & Positioning System - surface vehiclesGGP GPS Guidance Package - host of DoD platformsPSN Precision Strike Navigator - precision guided munitions
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Rate Gyro Principles and Designs
Type Principle
Rotor 1 and 221
Constancy of Angular Momentum
Sagnac Effect 11
Preservation ofPlane Vibration
1
Degrees ofFreedom
Design
Vibration
Optical
Hemispherical Resonant
Ring Laser.Fibre Optic.
Rigid Rotors.Dry Tuned.
Nuclear Resonant.
Example
Etak
HitachiAndrews
MuradaDelcoDraperBosch
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CURRENT GYRO TECHNOLOGY APPLICATIONS
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Sagnac Effect
Active Approach Passive Approach
RING LASER FOG INTERFEROMETER
OPTICAL GYRO TECHNOLOGIES
c
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Suitability of RLG for Strapdown
•Wide Dynamic Measuring Range
•Direct Digital Output
•Excellent scale factoring Linearity and Repeatability
•Excellent Bias Repeatability
•Rapid Reaction
•No G Sensitivity
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GG 1320 Digital Ring Laser Gyro• Characteristics — <
5.5 cubic inches — < 1 lb — < 2.5 watts — DC power in (+ 15 and +5 Vdc) — Compensated serial digital data output — No external support electronics — All high voltages self-contained — Built on proven RLG technology
(> 60,000 RLGs delivered) — Proven mechanical dither
• Demonstrated better than 1.0 nmi / hr performance
— Low random walk — Excellent scale
factor stability — Superb bias stability — No turn-on bias
transients — Low magnetic sensitivity• Laser Block in full-scale production
(900 gyros in 1992, 1300 in 1993, 1400 in 1994)
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Honeywell Ring Laser Gyros (RLGs)
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Ring Laser Gyro Operation
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The Fiber Optic Gyro
• Consists of:1. Semiconductor laser
diode as light source.2. Beam splitter.3. Coil of optical fiber.4. Photodetector
The Fiber Optic Gyro (FOG) measures rotation by
analyzing the phase shift of light
caused by the signac effect
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Tactical Grade Closed-Loop FOG• Tactical FOG IMU funded by USAF
• HG1800 FOG IMU is pin-for-pin compatible with HG1700 RLG IMU
• Goals:1 deg/hour Gyro Error1 milli-G Accel Error
• Housing identical to HG1700 IMU<35 cubic inches
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Types/Characteristic Applications Ex. Manufacturer Accuracy(deg/hr)
Maturity Cable Length
(meters)
Commercial Grade Automotive,Camera
Andrews 100 Present 100
Tactical Grade Attitude/Hdgreferences;Short-term inertial (min)
Litton 200, Honeywell
1 Present 200
Avionic Grade Aircraft &Cruise missile inertial
Eg GGP (GPS Guidance Package) Honeywell & Litton
.01 - .1 Within next year or two
1000
Strategic Grade Long-term ship inertial
Honeywell .00001 Maybe within 5 – 10 years in fleet
5000 - 10000
Quick-Look FOG Status
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SAGNAC Effect (Phase Shift Measured in Nano Radians)
Computer Maintains Spatial Reference Uses Large Coil LD Product (5 Km Fiber) Rugged, High Shock Resistance No Precision Machining
Typical High-performance IFOG
GYROELECTRONICS
PUMPLASER WDM
Erbium dopedfiber
LIGHT SOURCE
IOCCOUPLER
X XXX
X
DET
FIBER COIL
ESG Spinner Assembly
ROTOR
TECHNOLOGY DIFFERENCESTECHNOLOGY DIFFERENCES
Spinning Mass (3600 RPS) Rotor Maintains Spatial Reference Small Size of Rotating Element 1 cm
Rotor) Not Rugged, Susceptible to Rotor
Crashes Expensive Technology, Precision
Machining
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IMU Product Evolution Summary• RLG IMUs and RLG systems are a growth industry with proven
track records in the field
• FOG Inertial Systems striving to be lower price than comparable RLG-based systems
• MEMS gyros offer the lowest price, smallest size, and lowest power for a tactical IMU
• MEMS gyro performance will improve to 1 deg/hr in the next few years; ManTech programs will enable affordable MEMS IMUs in quantities
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Coordinate Systems
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Coordinate FramesAXIS 1 AXIS 2 AXIS 3
Inertial(I) (vernal equinox (in equatorial plane) (polar)
Aries)
ECEF(E) (through (in equatorial
Greenwich) plane)
Local Level (north)(in meridian (East) (down)
North(N) plane)
3GHA
A B P
322
-
LoL
mG mG P
N E D
3-
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AXIS 1 AXIS 2 AXIS 3
Wander(WA) ( counterclockwise ( counterclockwise
from north) from east)
( chosen such that )
Body (point to bow in (point to starboard (deck to keel)
deck plane) in deck plane)
Train gunsight(T) (out through gun barrel) don’t care don’t care
Coordinate Frames cont’d
owB tbdS kD
LD ieWAIE sinˆ
DW ˆˆ VU
321 HPR
G
32 AzElv
NOTE: Names, ordering of axes, ordering of rotations are not universally accepted. They are conventions and definition
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Coordinate Systems Use
Navigation quantities, eg, Position, Velocity, Acceleration,Jerk…. are three dimensional vectors and must, when quantified, be expressed with respect to a reference frame (aka) coordinate system.Likewise navigation measurements, eg distances and angles are made with respect to origins and axes of a coordinate system.
Va = = (for example)51014
V1a
V2a
V3a
Meters/secExample:
Three scalar elements of velocity vector wrt a coordinate frame.
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GEODESY, DATUMS
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Conceptual Reasons for Studying Geodesy
• Three main reasons for studying Geodesy/Astronomy related to inertial navigation:
1.Understanding the meaning of inertial
coordinate frame.2.Knowing gravitational attraction.3.Knowing the shape of the earth to determine
Latitude, Longitude , and Height from ECEF position.
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The Ellipsoid of RotationZ
P
P’
Equatorial Plane
aa
F O F’
b
X
aa
22 ba
12
2
2
2
bZ
aX
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Shape of the Earth
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WGS-84 & WGS-72 Defining Parameters
For WGS-84 Ellipsoid
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WGS-84 Derived Geometric Constants
CONSTANT NOTATION VALUE
Flattening(ellipticity) f 1/298.257223563Semiminor Axis b 6356752.3142mFirst Eccentricity e 0.0818191908426First Eccentrity Squared e2 0.00669437999013Polar Radius of Curvature c 6399593.6258mAxis Ratio b/a 0.996647189335mMean Radius of Semiaxis R1 6371008.7714mEqual Area Sphere Radius R2 6371007.1809mEqual Volume SphereRadius
R3 6371000.7900
First Eccentricity Squared= (a2-b2)/a2
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Different datums may use different ellipsoids. Datums may also differ by the location of the center and orientation of the ellipsoid.
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Simply put, a datum is the mathematical model of the Earth we use to calculate the coordinates onany map, chart, or survey system. All coordinates reference some particular set of numbers for the size andshape of the Earth.
The problem for warfighters is that many countries use their own datum when they make their maps andsurveys--what we call local datums. Other nations' maps often use coordinates computed assuming theEarth is a completely different size and shape from what the Department of Defense uses, but we have tobe ready to fight around the world.
US forces now use datum called World Geodetic System 1984, or WGS 84. The National Imagery andMapping Agency (NIMA) produces all for its new maps with this system. Unfortunately, we reprint many ofour maps from products made by allied countries that use local datums. Our old maps were made on severaldifferent local datums, or sometimes WGS 72 (maps using this datum were often printed "World GeodeticSystem" with no year identification). So the old maps we're reproducing, and the foreign ones we reprint,might use those other datums.
WHAT’S A DATUM?
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Gravity Disturbance EffectsOn INS
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TLV = True Local VerticalPerpendicular to GeoidActual Gravity VectorAstronomic Vertical
REV = Reference-Ellipsoid VerticalPerpendicular to Reference EllipsoidTheoretical Gravity VectorGeodetic Vertical
Geodetic Latitude
Surface of the Earth
Dynamic Sea Level
Surface of Reference Ellipsoid
Surface of Geoid
Gravity Anomaly
Deflection of the Vertical
Astronomic Latitude
TLVREV
N
SST
N = Surface of Geoid - Surface of Ellipsoid
SST = Sea Surface TopographyDynamic Sea Level - Surface of
Geoid
Figure 1. Simplified Depiction of Gravity QuantitiesE:\Courses\Geophysical Navigation
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APPROACHES TO GRAVITY COMPENSATION
STORED MAP APPROACHPATROL AND PRELAUNCH PHASE USE DEFLECTION/GEOD MAPS
TARGET OFFSETS USED FOR INFLIGHT EFFECTS COMPUTED FROM A COMBINATION OF GLOBAL/LONG WAVELENTH GRAVITY MODELS AND HIGH FREQUENCY DATA MAPS
REAL-TIME COMPENSATIONGRAVITY GRADIOMETER/GRAVIMETER MAY BE USED TO LIMIT GRAVITY-INDUCED NAVIGATION ERRORS
LAUNCH POINT MEASUREMENTS MAY BE USED TO REDUCE INFLIGHT EFFECTS
6/10/99
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Gravity Compensation Techniques
GRAVITY COMPENSATON EMBODIES• MAP UTILIZATION/INTERPOLATION AND/OR• REAL-TIME MEASUREMENTS AND• SYSTEM INTEGRATION
FUNDAMENTAL ELEMENTS
OPTIMAL ESTIMATES
OF NAV QUANTITIES
NAVAIDS
INS
GRAVIMETER/GRADIOMETER
STOREDGRAVITY MAP
SYSTEMINTEGRATIONESTIMATOR
+
+
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INS Error Analysis
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Causes of Inertial Navigation Errors
• Initial Conditions– An inertial needs three dimensional position, velocity,
and attitude (theoretically wrt the inertial coordinate system, but practically wrt a local coordinate system).
– For self initialization, these initial condition errors (particularly initial attitude errors) can be caused by sensor errors.
– Initial position and velocity often obtained from GPS• Sensor Errors
– Gyro and Accelerometer Errors• Bias, Scale factor, Cross axis sensitivities, input axis
misalignments, environmental sensitivities
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Causes of Inertial Navigation Errors (cont’d)
• Inertial Sensor Assembly Misalignments– Each sensors orientation may be misaligned– In general, only one accelerometer input axis can arbitrarily be
taken to be correct• Environmental Effects
– Gravity Disturbance Errors• Vertical Deflection for horizontal loops• Gravity anomaly for vertical loop
• Aiding Sensor Effects– Errors in altimeter either due to instrument or environment; similarly
for EM Log or Doppler aiding • Other
– Generally small digital data processing (coning and sculling) and timing errors
– Latency, synchro conversion, vibration
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GPS/INS Systems
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Inertial Navigation
System
AidingSources
OptimalProcessor
Corrected Navigation
Output(Includes Models of
INS errors, aiding errors, and motion models)
Non-Complementary Navigation Integration Methodology
*
* Branches represent potentiallyindividual accels. or gyro. outputs
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Inertial Navigation
System
AidingSources
InertialError
Estimates
CorrectedInertial Outputs
KalmanFilter+
-
Inertial + Aiding errors errors
True navigation+ aiding errors
Standard Complementary Filter Methodology in Feedback Configuration
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Loosely Coupled GPS/INS Integration ArchitectureRF / IF / A/D
MU
LTI-C
HIP
CO
RE
LATO
R
CARRIERDISCRIMINATOR
90°
I & D
IE
IP
QE
QL
IL
QPL1 L2
I
Q
(1000 Hz)
IMU
KALMAN FILTER
MEASUREMENTPROCESSING
KALMAN FILTER
NAVIGATIONEQUATIONS
(CHIP/SEC)
(50 Hz)
(CYC/SEC)(50 Hz)
(1 Hz) (1 Hz).
PVT (1 Hz),
PVAtt (1 Hz)
LOS VELOCITYAIDING (50 Hz)
INERTIALSYSTEMPROCESSING
1 of NGPSRCVRCHANNELS
GPS RCVRPROCESSING
+
-
GPSNAV
PROCESSING
(256 HZ)MEASUREMENTPROCESSING
CODENCO
CARRIERNCO
KFILTER
FILTER K
NAVIGATIONEQUATIONS
CODEGENERATOR
CODEDISCRIMINATOR
LOSPROJECTION
+
-
CARR. NCOBIAS (1 Hz)
CODE NCOBIAS (1 Hz)
E:\Courses\GPS\[10] GPS-INS
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Tightly Coupled GPS/INS Integration ArchitectureRF / IF / A/D
MU
LTI-C
HIP
CO
RE
LATO
R
CARRIERDISCRIMINATOR
90°
I & D
IE
IP
QE
QL
IL
QPL1 L2
I
Q
(1000 Hz)
IMU
(CHIP/SEC)
(50 Hz)
(CYC/SEC)(50 Hz)
(1 Hz) (1 Hz).
PVT (1 Hz),
PVAtt (1 Hz)
LOS VELOCITYAIDING (50 Hz)
INERTIALSENSORPROCESSING
1 of NGPSRCVRCHANNELS
GPS RCVRPROCESSING
GPSNAV
PROCESSING
(256 HZ)
CODENCO
CARRIERNCO
KFILTER
FILTER K
MEASUREMENTPROCESSING
CODEGENERATOR
CODEDISCRIMINATOR
LOSPROJECTION
+
-
CARR. NCOBIAS (1 Hz)
CODE NCOBIAS (1 Hz)
NAVIGATIONEQUATIONS
KALMAN FILTER
PVAtt
PV
E:\Courses\GPS\[10] GPS-INS
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Intimately Coupled GPS/INS Integration Architecture
RF / IF / A/D
MU
LTI-C
HIP
CO
RE
LATO
R
CARRIERDISCRIMINATOR
90°
I & D
IE
IP
QE
QL
IL
QPL1 L2
I
Q
(1000 Hz)
IMU
(CHIP/SEC)
(50 Hz)
(CYC/SEC)(50 Hz)
PVT (1 Hz),
PVAtt (1 Hz)INERTIALSENSORPROCESSING
1 of NGPSRCVRCHANNELS
GPS RCVR/NAVPROCESSING
(256 HZ)
CODEGENERATOR
CODEDISCRIMINATOR
LOSPROJECTION
+
-
NAVIGATIONEQUATIONS
KALMAN FILTER
FILTER
FILTER
CARRIERNCO
CODENCO
, (1 Hz).
PV (1 Hz)
T (100 Hz)
E:\Courses\GPS\[10] GPS-INS
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E:\Courses\GPS\[10] GPS-INS
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H-764G Embedded GPS/INSH-764G Features
• Small size: 7.0”H x 7.0”W x 9.8”L
• Light weight: 18 lbs*
• Low power: < 40 watts*
• High MTBF: > 6,500 hours*
• GPS/INS and two expansion slots in one small package
• Single i960 Microprocessor
• Mature, High-Performance Inertial Sensors
• 15-year Inertial Calibration Interval
• Collins GPS receiver Module
• Flight-Proven Ada Software
• Turn-Key System Missionization Environment
* Will vary depending upon how the expansion slots are populated
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Some Inertial Navigation Systems
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vendor unitsmodel HG1900 HG1920 commentsvolume 16 7.4 in³
Length/Diameter inWidth inDepth inmass 0.45 kgpower 3 w
temperature range -55 to +85
ºC
vibrationshock 10000 g
update rate 100 Hzrange 20 gbias 1 .6-6.4 mg
scale factor 300 84-2700 ppmnonlinearity 500 200 ppmresolution µg
noise mg/√Hzbandwidth Hz
random walk .19-.17 m/s/√hrrange 1440 º/secbias 30 09-76 º/hr
scale factor 150 91-524 ppmnonlinearity ppmresolution º/hr
noise deg/secbandwidth Hz
random walk 0.1 .02-.17 º/√hrdata source
gyro
http://content.honeywell.com/ds
Honeywell/Draper
imu
accelerometer
Honeywell/Draper
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vendor unitsmodel LN-200 commentsvolume 32.2 in³
Length/Diameter 3.5 inWidth inDepth 3.35 inmass 0.7 kgpower 10 w
temperature range -54 to 85 ºC
vibration 18 g rmsshock 90 g
update rate Hzrange 40 gbias 1 mg
scale factor 300 ppmnonlinearity ppmresolution µg
noise mg/√Hzbandwidth Hz
random walk 0.012 m/s/√hrrange 1000 º/secbias 10 º/hr
scale factor 100 ppmnonlinearity ppmresolution º/hr
noise deg/secbandwidth 500 Hz
random walk 0.1 º/√hrdata source
gyro
imu
Northrup-Grumman
accelerometer
Northrup-Grumman
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vendor unitsmodel SiLMU01 commentsvolume 6.1 in³
Length/Diameter 2.36 inWidth inDepth 1.79 inmass 0.26 kgpower 5 w
temperature range
-40 to +72 operating ºC
vibrationshock 100 11 ms, .5 sine g
update rate Hzrange 50 ± gbias 2 1 σ mg
scale factor 2000 1 σ ppmnonlinearity 1500 ppmresolution µg
noise 5 mg rms in band mg/√Hzbandwidth 75 Hz
random walk 1 m/s/√hrrange 1000 ± º/secbias 100 º/hr
scale factor 400 accuracy ppmnonlinearity 100 ppmresolution º/hr
noise 0.5 rms inband deg/secbandwidth 75 Hz
random walk 1 º/√hrdata source http://www.baesystems-
BAE
imu
accelerometer
gyro
BAE
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• The AN/WSN-7 was designed as a form, fit, and function replacement for the AN/WSN-1, and -5 for installation on DDG 51, CG 47, CV, CVN, LHA 1 and LHD 1 Class platforms.
• The AN/WSN-7A was designed as a form, fit, and function replacement for the AN/WSN-3 on SSN688 Class platforms.
• Provides attitude (roll, pitch, and heading), position, and velocity data to ship system users.
WSN-7 Information
Courtesy Spawar Systems Center, Norfork (Carvil, Galloway)
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CN-1695/WSN-7(V)CN-1696/WSN-7(V)CN-1697/WSN-7(V)
Ring Laser Gyro Navigator
MX-11681/WSNInertial Measuring Unit
(Inside Cabinet)
IP-1747/WSN Display Unit, Control
EquipmentAN/WSN-7(V) 1/2/3 RLGN
Courtesy Spawar Systems Center, Norfork (Carvil, Galloway)
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Install Schedule
SHIPCLASS
FY02 FY03 FY04 FY05 FY06 FY07 TOCOMPLETE
CG 47 CG 48CG 49
DDG 51 DDG 51 DDG 61DDG 53 DDG 65DDG 56 DDG 73DDG 59 DDG 74
DDG 52
DD 963
LHA LHA 5 LHA 3LHA 1
LHD LHD 4 LHD 1LHD 2
LHD 3 LHD 6
AGF/LCC LCC 19LCC 20
CV/CVN CVN 65
DDG DDG 93DDG 94DDG 95
DDG 97 DDG 102DDG 103DDG 104
CVN CVN 67
LHD LHD 8
TOTALSHIPS
18 7 2 4
OP
NS
CN
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CD-132/WSN-7A(V)CD-133/WSN-7A(V)
Control Unit, Electronic
IP-1747/WSN Display Unit, Control
CY-8827/WSN-7A(A)Enclosure Assembly, Inertial
Measuring UnitMX-11681/WSN
Inertial Measuring Unit
MX-11682/WSN-7A(V)Support, Electronics Unit
MX-11682/WSN-7A(V)Support, Electronics Unit
IP-1746/WSN Display Unit, Secondary Control
IP-1747/WSN Display Unit, Control
Equipment (Cont.)AN/WSN-7A(V) Red/Green RLGN
Courtesy Spawar Systems Center, Norfork (Carvil, Galloway)
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Install Schedule (Cont.)
SHIPCLASS
FY02 FY03 FY04 FY05 FY06 FY07 TOCOMPLETE
SSN 688 SSN 690 SSN 763SSN 719 SSN 767SSN 721 SSN 768SSN 722 SSN 771SSN 754 SSN 772SSN 756
SSN 701SSN 757SSN 760
SSN 713SSN 715
SSN 709SSN 715SSN 752 SSN 756SSN 761SSN 764
SSN 698SNN 699SSN 720
SSN 769
SSN 21 SSN 21 SSN 22
SSN 21
SSN 774 SSN 778 SSN 779
SSN 780 SSN 784
SSN 782 SSN 783
SSN 784SSN 785
SSN 786 thru SSN 803
SSGN SSGN 726SSGN 728
SSGN 727SSGN 729
TOTALSHIPS
11 3 7 11 5 3 18
OP
NS
CN
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Evolution of Inertial Navigation
3-Axis Gyro Chip
3-Axis Accelerometer Chip
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Evolution of Inertial Navigation Technology
• Size ,cost,power of Inertial Systems greatly reduced by technology developments• MEMS Technology promises the next major step in Inertial System evolution
Litto
nSi
GyT
MS/
N #
0004
FPGA
Gimbaled Technology
Strapdown Technology
Ring Laser Technology
Fiber Optic Technology
MEMS Technology
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Low Cost Guidance and Navigation
• Low Cost Guidance Package enables cost effective precise positioning to be embedded in low value, high volume quantity systems
GPS
Low Cost Guidance and Navigation Package
MEMSInertial Sensors
DSP’sProcessorsElectronics
Applications
• Air/Ground Manned /Unmanned Platforms• Guided Rockets• Guided Munitions• Soldier Man Pack• Re-supply Vehicles• …….• ….• ..
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2000 200320022001
LN 205G
ATK SAASM GPS
•Leveraging LN 200 series development reduces MEMS time-to-market
LN 205
LN 200 IMU
LN 300
LN 300GLi
tton
SiA
cTM
S/N
#00
01
LittonSiAcTM
S/N #0001
Litto
nSi
AcT
M
S/N
#00
01
LittonSiGyTM
S/N #0001
Litto
nSi
GyT
M
S/N
#00
04
ANALOG DEVICES
AN
ALO
G D
EVI
CE
S
ANALOG DEVICES
ANALOG DEVICES
DigitalAsic
AnalogAsic
LN 200G IMU
LN300 /LN 200 MEMS INS/GPS Roadmap
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The Future• Over the next 3 to 5 years, the applicability of
MEMS for high-g tactical applications will be conclusively demonstrated.
• From 5 to 10 years, the insertion of high-volume production MEMS IMUs and INS/GPS into tactical systems will occur at an ever-increasing rate.
• The realization of 3 gyros on a chip and 3 accelerometers on a chip, represents the next order-of-magnitude size reduction.
• Commercial applications will exploit the development MEMS technology into quantities of billions.
3-Axis Gyro Chip
3-Axis Accelerometer Chip