introduction to navigation systems

Introduction to Navigation Systems

Joseph HennawyComputer Engineer

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.

Inertial Navigation History

Inertial Guidance System of SAGEM used in the Air-Surface Medium-range missile

Dead Reckoning

Early Compasses

Surveyor’s Compass--1820

Jean Bernard Léon FoucaultOriginator of the Foucault pendulum

1819-68

Foucault's gyroscope (1851)

Mechanical Dead Reckoning Computer: Early 20th century

SG-66 Guidance System for the V-2 (1944)

Charles Stark Draper Gyroscopic Apparatus - Spinning Gyroscope

Born 2 October 1901Died 25 July 1987

First Successful All-Inertial Navigator (1954)

Professor Arnold Nordiseck Holding Early Electrostatically Suspended Gyroscope (1959)

Honeywell Advertisement for Electrostatically Suspended Gyroscope, 1962

Warren Macek of Sperry Circa 1963 Demonstrating the Ring Laser Gyro Concept

Laser Gyro

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

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

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

Accelerometers

FORCER VERTICALPIVOT

PICKOFF

AMPLIFIER

Simple Pendulum Accelerometer

Torque Balance Pendulous Accelerometer Schematic

EMERGING ACCELEROMETER TECHNOLOGY APPLICATIONS

MEMS/MOEMS

Mech.

SiliconQuartz

WSN-7 Accelerometer

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

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

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

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

Gyroscopes

INERTIAL ROTATION SENSOR TECHNOLOGY

E;\Courses\Gyros

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

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

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

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

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

CURRENT GYRO TECHNOLOGY APPLICATIONS

Sagnac Effect

Active Approach Passive Approach

RING LASER FOG INTERFEROMETER

OPTICAL GYRO TECHNOLOGIES

c

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

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)

Honeywell Ring Laser Gyros (RLGs)

Ring Laser Gyro Operation

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

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

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

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

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

Coordinate Systems

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-

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

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.

GEODESY, DATUMS

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.

The Ellipsoid of RotationZ

P

P’

Equatorial Plane

aa

F O F’

b

X

aa

22 ba

12

2

2

2

bZ

aX

Shape of the Earth

WGS-84 & WGS-72 Defining Parameters

For WGS-84 Ellipsoid

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

Different datums may use different ellipsoids. Datums may also differ by the location of the center and orientation of the ellipsoid.

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?

Gravity Disturbance EffectsOn INS

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

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

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

+

+

INS Error Analysis

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

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

GPS/INS Systems

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

Inertial Navigation

System

AidingSources

InertialError

Estimates

CorrectedInertial Outputs

KalmanFilter+

-

Inertial + Aiding errors errors

True navigation+ aiding errors

Standard Complementary Filter Methodology in Feedback Configuration

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

Tightly Coupled GPS/INS Integration ArchitectureRF / IF / A/D

MU

LTI-C

HIP

CO

RE

LATO

R


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


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


Intimately Coupled GPS/INS Integration Architecture

RF / IF / A/D

MU

LTI-C

HIP

CO

RE

LATO

R


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


GPS RCVR/NAVPROCESSING

(256 HZ)

CODEGENERATOR

CODEDISCRIMINATOR

LOSPROJECTION

+

-

NAVIGATIONEQUATIONS

KALMAN FILTER

FILTER

FILTER

CARRIERNCO

CODENCO

, (1 Hz).

PV (1 Hz)

T (100 Hz)


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

Some Inertial Navigation Systems

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

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

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

• 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)

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


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

CD-132/WSN-7A(V)CD-133/WSN-7A(V)

Control Unit, Electronic


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


Equipment (Cont.)AN/WSN-7A(V) Red/Green RLGN


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

Evolution of Inertial Navigation

3-Axis Gyro Chip

3-Axis Accelerometer Chip

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

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• …….• ….• ..

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

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

introduction to navigation systems

Documents