1. Introduction to Navigation Systems
Joseph Hennawy
Computer Engineer

2. 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.

3. Inertial Navigation History
Inertial Guidance System of SAGEM used in the Air-Surface Medium-
range missile

4. Dead Reckoning

5. Early Compasses

6. Surveyor’s Compass--1820

7. Jean Bernard Léon Foucault
Originator of the Foucault pendulum
1819-68

8. Foucault's gyroscope (1851)

9. Mechanical Dead Reckoning Computer: Early 20th century

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

11. Charles Stark Draper
Gyroscopic Apparatus - Spinning Gyroscope
Born 2 October 1901
Died 25 July 1987

12. First Successful All-Inertial Navigator (1954)

13. Professor Arnold Nordiseck Holding Early Electrostatically Suspended
Gyroscope (1959)

14. Honeywell Advertisement for Electrostatically Suspended Gyroscope,
1962

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

16. Laser Gyro

17. 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 Error
1 milli-G Accel Error
• Housing identical to HG1700 IMU
<35 cubic inches

18. 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
[ ]IVm
dt
d
F

=

19. 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

20. Accelerometers

23. FORCER VERTICAL
PIVOT
PICKOFF
AMPLIFIER
Simple Pendulum Accelerometer

24. Torque Balance Pendulous Accelerometer Schematic

25. EMERGING ACCELEROMETER TECHNOLOGY APPLICATIONS

26. MEMS/MOEMS
Mech.
Silicon
Quartz

29. WSN-7 Accelerometer

30. 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

31. 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-g
Misalignment 412 arcsec
Scale Factor 500 ppm
Second Order Scale Factor Non-Linearity 60 micro-g/g2
Additional Terms
Notes

32. Accelerometer Name $20K
Part of System Name $20K
Where Found IMU Performance vs. Cost
Velocity Random Walk 0.03 (meters/sec)/√(rt-hr)
Bias 100 micro-g
Misalignment 10.3 arcsec
Scale Factor 10 ppm
Second Order Scale Factor Non-Linearity 3 micro-g/g2
Additional Terms
Notes

33. Velocity Random Walk 0.0003 (meters/sec)/√(rt-hr)
Bias 100 micro-g
Misalignment 3 arcsec
Scale Factor 100 ppm
Second Order Scale Factor Non-Linearity 0.5 micro-g/g2
Additional Terms
Notes
Accelerometer Name $100K
Part of System Name $100K
Where Found IMU Performance vs. Cost

34. Gyroscopes

35. INERTIAL ROTATION SENSOR TECHNOLOGY
E;CoursesGyros

37. INERTIAL SENSOR APPLICATION
1 5 25 125 625 3125
1e-005
0.0001
0.001
0.01
0.1
1
10
WEIGHT
SENSORPERFORMANCE(deg/hr)
TACTICAL
MISSILES
GBI / ASAT
RV
MEDIUM ACCURACY
AIRCRAFT
COMMERCIAL
AIRCRAFT
HIGH ACCURACY
AIRCRAFT
ICBMSDI
POINTING
SURFACE
SHIP
SUB

38. Inertial Sensor Technology Comparison
Inertial Acronym Definitions
ESG Electrostatic Gyro
FOG – Fiber Optic Gyro
HRG – Hemispherical Resonator Gyro
MS – Multisensor
MEMS – Micromachined
Electromechanical Sensor
QRS – Quartz Rate Sensor
RLG – Ring Laser Gyro
Inertial Acronym Definitions
ESG Electrostatic Gyro
FOG – Fiber Optic Gyro
HRG – Hemispherical Resonator Gyro
MS – Multisensor
MEMS – Micromachined
Electromechanical Sensor
QRS – Quartz Rate Sensor
RLG – Ring Laser Gyro
ESG
RLG
FOG
MS
QRS
HRG
MEMS
G yroD rift
(deg/hr)
Submarines
Strategic MX
Surface Ships
Aircraft
Cruise Missles
UAVs
Precision Guided Munitions (PGM)
SCUD-B
NO-DONG
Unguided
GGP
FOG
EGI
SLAM-ER
SLAM
F-18
TLAM
JDAM AGM-L30 EKGM
All sensor perf ranges are estimates based on available information
All sensor perf ranges are estimates based on available information

39. Honeywell Gyro Technology Heritage
1920 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

40. IMU Product Evolution Overview
RLG FOG MEMS
• EGI • GGP • Future
• MAPS • PSN Growth
• Digital
Laser
Gyro
• HG1700 • HG1800 • HG1900
- in production - developmental - in development
Tactical
Grade IMUs
Navigation
Grade
Systems
and
Components
EGI Embedded GPS Inertial Integrated System - aircraft, et al
MAPS Modular Azimuth & Positioning System - surface vehicles
GGP GPS Guidance Package - host of DoD platforms
PSN Precision Strike Navigator - precision guided munitions

41. Rate Gyro Principles and Designs
Type Principle
Rotor 1 and 2
2
1
Constancy of
Angular Momentum
Sagnac Effect 1
1
Preservation of
Plane Vibration
1
Degrees of
Freedom
Design
Vibration
Optical
Hemispherical Resonant
Ring Laser.
Fibre Optic.
Rigid Rotors.
Dry Tuned.
Nuclear Resonant.
Example
Etak
Hitachi
Andrews
Murada
Delco
Draper
Bosch

42. CURRENT GYRO TECHNOLOGY APPLICATIONS

43. Sagnac Effect
Active Approach Passive Approach
RING LASER FOG
INTERFEROMETER
OPTICAL GYRO TECHNOLOGIES
∆ƒ = (4Α/λΡ)Ω
∆Φ = (8πΝΑ/λ )Ωc

44. 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

45. 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)

46. Honeywell Ring Laser Gyros
(RLGs)

47. Ring Laser Gyro Operation

48. 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

49. 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 Error
1 milli-G Accel Error
• Housing identical to HG1700 IMU
<35 cubic inches

50. Types/Characteristic Applications Ex. Manufacturer Accuracy
(deg/hr)
Maturity Cable
Length
(meters)
Commercial Grade Automotive,
Camera
Andrews 100 Present 100
Tactical Grade Attitude/Hdg
references;
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

51.  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
GYRO
ELECTRONICS
PUMP
LASER
WDM
Erbium doped
fiber
LIGHT SOURCE
IOC
COUPLER
X XX
X
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
Ω=∆Φ
c
NA
λ
π8

52. 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

53. Coordinate Systems

54. Coordinate Frames
AXIS 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
↓
Aˆ Bˆ Pˆ
[ ]3
22
-
LoL





−
↓
π
mGˆ mG ′ Pˆ
Nˆ Eˆ Dˆ
[ ]3α-
↓

55. AXIS 1 AXIS 2 AXIS 3
Wander(WA) (α counterclockwise (α counterclockwise
from north) from east)
(α 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
owBˆ
tbdSˆ kDˆ
LD ie
WA
IE
sinˆ ω−=•Ω

DW ˆˆ =VˆUˆ
[ ] [ ] [ ]321 HPR
↓
Gˆ
[ ] [ ]32 AzElv
↓
NOTE: Names, ordering of axes, ordering of rotations are not universally accepted.
They are conventions and definition

56. 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)
5
10
14
V1
a
V2
a
V3
a
Meters/secExample:
Three scalar elements of velocity vector wrt a coordinate frame.

57. GEODESY, DATUMS

58. 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.

59. The Ellipsoid of Rotation
Z
P
P’
Equatorial
Plane
a
a
F O F’
b
X
a
a
22
ba +
12
2
2
2
=+
b
Z
a
X

60. Shape of the Earth

61. WGS-84 & WGS-72 Defining Parameters
For WGS-84 Ellipsoid

62. WGS-84 Derived Geometric
Constants
CONSTANT NOTATION VALUE
Flattening(ellipticity) f 1/298.257223563
Semiminor Axis b 6356752.3142m
First Eccentricity e 0.0818191908426
First Eccentrity Squared e
2
0.00669437999013
Polar Radius of Curvature c 6399593.6258m
Axis Ratio b/a 0.996647189335m
Mean Radius of Semiaxis R1 6371008.7714m
Equal Area Sphere Radius R2 6371007.1809m
Equal Volume Sphere
Radius
R3 6371000.7900
First Eccentricity Squared= (a2
-b2
)/a2

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

64. Simply put, a datum is the mathematical model of the Earth we use to calculate the coordinates on
any map, chart, or survey system. All coordinates reference some particular set of numbers for the
size and
shape of the Earth.
The problem for warfighters is that many countries use their own datum when they make their maps
and
surveys--what we call local datums. Other nations' maps often use coordinates computed assuming
the
Earth is a completely different size and shape from what the Department of Defense uses, but we
have to
be ready to fight around the world.
US forces now use datum called World Geodetic System 1984, or WGS 84. The National Imagery
and
Mapping Agency (NIMA) produces all for its new maps with this system. Unfortunately, we reprint
many of
our maps from products made by allied countries that use local datums. Our old maps were made on
several
different local datums, or sometimes WGS 72 (maps using this datum were often printed "World
Geodetic
System" 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?

66. Gravity Disturbance Effects
On INS

67. TLV = True Local Vertical
Perpendicular to Geoid
Actual Gravity Vector
Astronomic Vertical
REV = Reference-Ellipsoid Vertical
Perpendicular to Reference Ellipsoid
Theoretical Gravity Vector
Geodetic 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
TLV
REV
N
SST
N = Surface of Geoid - Surface of
Ellipsoid
SST = Sea Surface Topography
Figure 1. Simplified Depiction of Gravity Quantities
E:CoursesGeophysical Navigation

68. APPROACHES TO GRAVITY COMPENSATION
STORED MAP APPROACH
PATROL 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 COMPENSATION
GRAVITY GRADIOMETER/GRAVIMETER MAY BE USED
TO LIMIT GRAVITY-INDUCED NAVIGATION ERRORS
LAUNCH POINT MEASUREMENTS MAY BE USED TO
REDUCE INFLIGHT EFFECTS
6/10/99

69. 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
STORED
GRAVITY MAP
SYSTEM
INTEGRATION
ESTIMATOR
+
+

70. INS Error Analysis

71. 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

72. 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

73. GPS/INS Systems

75. Inertial
Navigation
System
Aiding
Sources
Optimal
Processor
Corrected
Navigation
Output
(Includes Models of
INS errors, aiding errors,
and motion models)
Non-Complementary Navigation Integration Methodology
*
* Branches represent potentially
individual accels. or gyro. outputs

76. Inertial
Navigation
System
Aiding
Sources
Inertial
Error
Estimates
Corrected
Inertial
Outputs
Kalman
Filter+
-
Inertial + Aiding
errors errors
True navigation
+ aiding errors
Standard Complementary Filter Methodology in
Feedback Configuration

77. Loosely Coupled GPS/INS
Integration ArchitectureRF / IF / A/D
MULTI-CHIP
CORELATOR
CARRIER
DISCRIMINATOR
90°
I & D
IE
IP
QE
QL
IL
QP
}
L1 L2
I
Q
(1000 Hz)
IMU
KALMAN FILTER
MEASUREMENT
PROCESSING
KALMAN FILTER
Σ
NAVIGATION
EQUATIONS
(CHIP/SEC)
(50 Hz)
(CYC/SEC)
(50 Hz)
ρ (1 Hz) ρ (1 Hz)
.
PVT (1 Hz)∆θ,∆υ
PVAtt (1 Hz)
LOS VELOCITY
AIDING (50 Hz)
INERTIAL
SYSTEM
PROCESSING
1 of N
GPS
RCVR
CHANNELS
GPS RCVR
PROCESSING
+
-
GPS
NAV
PROCESSING
(256 HZ)
MEASUREMENT
PROCESSING
CODE
NCO
CARRIER
NCO
KFILTER
FILTER K
NAVIGATION
EQUATIONS
CODE
GENERATOR
CODE
DISCRIMINATOR
ΣΣ
LOS
PROJECTION
+
-
Σ
CARR. NCO
BIAS (1 Hz)
CODE NCO
BIAS (1 Hz)
E:CoursesGPS[10] GPS-INS

78. Tightly Coupled GPS/INS
Integration ArchitectureRF / IF / A/D
MULTI-CHIP
CORELATOR
CARRIER
DISCRIMINATOR
90°
I & D
IE
IP
QE
QL
IL
QP
}
L1 L2
I
Q
(1000 Hz)
IMU
(CHIP/SEC)
(50 Hz)
(CYC/SEC)
(50 Hz)
ρ (1 Hz) ρ (1 Hz)
.
PVT (1 Hz)
∆θ,∆υ
PVAtt (1 Hz)
LOS VELOCITY
AIDING (50 Hz)
INERTIAL
SENSOR
PROCESSING
1 of N
GPS
RCVR
CHANNELS
GPS RCVR
PROCESSING
GPS
NAV
PROCESSING
(256 HZ)
CODE
NCO
CARRIER
NCO
KFILTER
FILTER K
MEASUREMENT
PROCESSING
CODE
GENERATOR
CODE
DISCRIMINATOR
ΣΣ
LOS
PROJECTION
+
-
Σ
CARR. NCO
BIAS (1 Hz)
CODE NCO
BIAS (1 Hz)
NAVIGATION
EQUATIONS
KALMAN FILTER
PVAtt
PV
E:CoursesGPS[10] GPS-INS

79. Intimately Coupled GPS/INS Integration
Architecture
RF / IF / A/D
MULTI-CHIP
CORELATOR
CARRIER
DISCRIMINATOR
90°
I & D
IE
IP
QE
QL
IL
QP
}
L1 L2
I
Q
(1000 Hz)
IMU
(CHIP/SEC)
(50 Hz)
(CYC/SEC)
(50 Hz)
PVT (1 Hz)
∆θ,∆υ
PVAtt (1 Hz)INERTIAL
SENSOR
PROCESSING
1 of N
GPS
RCVR
CHANNELS
GPS RCVR/NAV
PROCESSING
(256 HZ)
CODE
GENERATOR
CODE
DISCRIMINATOR
LOS
PROJECTION
+
-
Σ
NAVIGATION
EQUATIONS
KALMAN FILTER
FILTER
FILTER
CARRIER
NCO
CODE
NCO
∆ρ, ∆ρ (1 Hz)
.
PV (1 Hz)
T (100 Hz)
E:CoursesGPS[10] GPS-INS

84. H-764G Embedded GPS/INS
H-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
* Will vary depending upon how the
expansion slots are populated

87. Some Inertial Navigation Systems

88. vendor units
model HG1900 HG1920 comments
volume 16 7.4 in³
Length/Diameter in
Width in
Depth in
mass 0.45 kg
power 3 w
temperature range
-55 to
+85
ºC
vibration
shock 10000 g
update rate 100 Hz
range 20 g
bias 1 .6-6.4 mg
scale factor 300 84-2700 ppm
nonlinearity 500 200 ppm
resolution µg
noise mg/√Hz
bandwidth Hz
random walk .19-.17 m/s/√hr
range 1440 º/sec
bias 30 09-76 º/hr
scale factor 150 91-524 ppm
nonlinearity ppm
resolution º/hr
noise deg/sec
bandwidth Hz
random walk 0.1 .02-.17 º/√hr
data source
gyro
http://content.honeywell.com/ds
Honeywell/Draper
imu
accelerometer
Honeywell/Draper

89. vendor units
model LN-200 comments
volume 32.2 in³
Length/Diameter 3.5 in
Width in
Depth 3.35 in
mass 0.7 kg
power 10 w
temperature
range
-54 to 85 ºC
vibration 18 g rms
shock 90 g
update rate Hz
range 40 g
bias 1 mg
scale factor 300 ppm
nonlinearity ppm
resolution µg
noise mg/√Hz
bandwidth Hz
random walk 0.012 m/s/√hr
range 1000 º/sec
bias 10 º/hr
scale factor 100 ppm
nonlinearity ppm
resolution º/hr
noise deg/sec
bandwidth 500 Hz
random walk 0.1 º/√hr
data source
gyro
imu
Northrup-Grumman
accelerometer
Northrup-Grumman

90. vendor units
model SiLMU01 comments
volume 6.1 in³
Length/Diameter 2.36 in
Width in
Depth 1.79 in
mass 0.26 kg
power 5 w
temperature
range
-40 to +72 operating ºC
vibration
shock 100 11 ms, .5 sine g
update rate Hz
range 50 ± g
bias 2 1 σ mg
scale factor 2000 1 σ ppm
nonlinearity 1500 ppm
resolution µg
noise 5 mg rms in band mg/√Hz
bandwidth 75 Hz
random walk 1 m/s/√hr
range 1000 ± º/sec
bias 100 º/hr
scale factor 400 accuracy ppm
nonlinearity 100 ppm
resolution º/hr
noise 0.5 rms inband deg/sec
bandwidth 75 Hz
random walk 1 º/√hr
data source http://www.baesystems-
BAE
imu
accelerometer
gyro
BAE

92. • 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)

93. CN-1695/WSN-7(V)
CN-1696/WSN-7(V)
CN-1697/WSN-7(V)
Ring Laser Gyro Navigator
MX-11681/WSN
Inertial Measuring Unit
(Inside Cabinet)
IP-1747/WSN
Display Unit, Control
Equipment
AN/WSN-7(V) 1/2/3 RLGN
Courtesy Spawar Systems Center, Norfork (Carvil, Galloway)

94. Install Schedule
SHIP
CLASS
FY02 FY03 FY04 FY05 FY06 FY07 TO
COMPLETE
CG 47 CG 48
CG 49
DDG 51 DDG 51 DDG 61
DDG 53 DDG 65
DDG 56 DDG 73
DDG 59 DDG 74
DDG 52
DD 963
LHA LHA 5 LHA 3
LHA 1
LHD LHD 4 LHD 1
LHD 2
LHD 3 LHD 6
AGF/LCC LCC 19
LCC 20
CV/CVN CVN 65
DDG DDG 93
DDG 94
DDG 95
DDG 97 DDG 102
DDG 103
DDG 104
CVN CVN 67
LHD LHD 8
TOTAL
SHIPS
18 7 2 4
OPNSCN

95. 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 Unit
MX-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)

96. Install Schedule (Cont.)
SHIP
CLASS
FY02 FY03 FY04 FY05 FY06 FY07 TO
COMPLETE
SSN 688 SSN 690 SSN 763
SSN 719 SSN 767
SSN 721 SSN 768
SSN 722 SSN 771
SSN 754 SSN 772
SSN 756
SSN 701
SSN 757
SSN 760
SSN 713
SSN 715
SSN 709
SSN 715
SSN 752
SSN 756
SSN 761
SSN 764
SSN 698
SNN 699
SSN 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 784
SSN 785
SSN 786 thru
SSN 803
SSGN SSGN 726
SSGN 728
SSGN 727
SSGN 729
TOTAL
SHIPS
11 3 7 11 5 3 18
OPNSCN

97. Evolution of Inertial Navigation
3-Axis Gyro Chip
3-Axis Accelerometer Chip

99. 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
Litton
SiGyTM
S/N#0004
FPGA
Gimbaled
Technology
Strapdown
Technology
Ring Laser
Technology
Fiber Optic
Technology
MEMS
Technology

100. 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
MEMS
Inertial
Sensors
DSP’s
Processors
Electronics
Applications
• Air/Ground Manned
/Unmanned Platforms
• Guided Rockets
• Guided Munitions
• Soldier Man Pack
• Re-supply Vehicles
• …….
• ….
• ..

101. 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 300GLitton
SiAcTM
S/N#0001
Litton
SiAcTM
S/N #0001
Litton
SiAcTM
S/N#0001
Litton
SiGyTM
S/N #0001
Litton
SiGyTM
S/N#0004
ANALOG
DEVICES
ANALOG
DEVICES
ANALOG
DEVICES
ANALOG
DEVICES
Digital
Asic
Analog
Asic
LN 200G IMU
LN300 /LN 200 MEMS INS/GPS Roadmap

102. 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