zotero/storage/7GYP58XR/.zotero-ft-cache

PUBLIC RELEASE VERSION
NAVSTAR GPS USER EQUIPMENT
INTRODUCTION
SEPTEMBER 1996
PUBLIC RELEASE VERSION

CONTENTS
Page
CHAPTER 1: SYSTEM OVERVIEW..............................................................................................1-1 1.1 General Description .........................................................................................................1-1 1.2 System Overview.............................................................................................................1-2 1.2.1 Space Segment..................................................................................................1-2 1.2.2 Control Segment ...............................................................................................1-3 1.2.3 User Segment....................................................................................................1-5 1.3 GPS Services ...................................................................................................................1-5 1.3.1 Precise Positioning Service................................................................................1-5 1.3.2 Standard Positioning Service.............................................................................1-6 1.4 GPS Theory of Operation................................................................................................1-6 1.4.1 GPS Satellite Signals.........................................................................................1-7 1.4.1.1 C/A-Code..........................................................................................1-7 1.4.1.2 P(Y)-Code ........................................................................................1-7 1.4.1.3 Navigation Message..........................................................................1-7 1.4.1.4 Satellite Signal Modulation ..............................................................1-8 1.4.2 GPS Receiver Operation ...................................................................................1-9 1.4.2.1 Satellite Selection............................................................................1-10 1.4.2.2 Satellite Signal Acquisition..............................................................1-11 1.4.2.3 Down Conversion...........................................................................1-12 1.4.2.4 Code Tracking ................................................................................1-13 1.4.2.5 Carrier Tracking and Data Detection ..............................................1-13 1.4.2.6 Data Demodulation.........................................................................1-14 1.4.2.7 P(Y)-Code Signal Acquisition.........................................................1-14 1.4.2.8 PVT Calculations............................................................................1-14 1.4.2.9 Degraded Operation and Aiding .....................................................1-17 1.5 Program Management ...................................................................................................1-17 1.5.1 System Development and Management ..........................................................1-17 1.5.2 System Requirements, Planning, and Operations ............................................1-17 1.6 GPS Program History....................................................................................................1-18 1.6.1 Pre-Concept Validation (1960s-1972) ............................................................1-18 1.6.2 Phase I - Concept Validation (1973-1979)......................................................1-18 1.6.3 Phase II - Full Scale Development (1979-1985)..............................................1-19 1.6.4 Phase III - Production and Deployment (1986 to Present)..............................1-20 1.6.4.1 Space Segment (1986 to Present) ...................................................1-20 1.6.4.2 Control Segment (1986 to Present) ................................................1-21 1.6.4.3 User Segment (1986 to Present) .....................................................1-22
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CONTENTS (Continued)
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CHAPTER 2: TYPES OF GPS RECEIVERS AND THEIR INTENDED APPLICATIONS .....................................................................................................2-1
2.1 GPS Receiver Architectures ............................................................................................2-1 2.1.1 Continuous Receivers........................................................................................2-1 2.1.2 Sequential Receivers..........................................................................................2-1 2.1.2.1 One-Channel Sequential Receivers....................................................2-1 2.1.2.2 Two-Channel Sequential Receivers...................................................2-2 2.1.3 Multiplex (MUX) Receivers..............................................................................2-2
2.2 "All-In-View" Receivers ..................................................................................................2-2 2.3 Autonomous Integrity Monitoring Techniques................................................................2-3 2.4 Time Transfer Receivers..................................................................................................2-3 2.5 Differential GPS (DPGS) Receivers ................................................................................2-3 2.6 Surveying Receivers.........................................................................................................2-5 2.7 Analog/Digital Receivers .................................................................................................2-7 2.8 GPS As A Pseudorange/Delta Range Sensor ..................................................................2-8
CHAPTER 3: MINIMUM PERFORMANCE CAPABILITIES OF A GPS RECEIVER .....................................................................................................3-1
3.1 Basic Considerations........................................................................................................3-1 3.1.1 GPS System Accuracy Characteristics ..............................................................3-1 3.1.2 GPS PPS System Range-Error Budget.............................................................3-2 3.1.2.1 GPS UE Range-Error Budget...........................................................3-3 3.1.3 Geometric Dilution of Precision........................................................................3-4
3.2 Receiver Position Accuracy.............................................................................................3-6 3.3 Receiver Velocity Accuracy.............................................................................................3-7 3.4 Receiver Time Accuracy..................................................................................................3-7 3.5 Time-To-First-Fix............................................................................................................3-8
3.5.1 Warm Start, Cold Start, and Hot Start..............................................................3-9 3.5.2 Receiver Warm-Up ...........................................................................................3-9 3.5.3 Almanac Collection .........................................................................................3-10 3.5.4 Initial Uncertainties..........................................................................................3-10 3.5.5 Ephemerides Collection...................................................................................3-10 3.5.6 Enhanced Acquisition Techniques...................................................................3-10 3.5.7 Direct P(Y)-Code Acquisition.........................................................................3-11 3.5.8 TTFF Requirements ........................................................................................3-11 3.5.9 Satellite Reacquisition .....................................................................................3-11
CHAPTER 4: GPS RECEIVER INTERFACE AND ANCILLARY EQUIPMENT ..........................................................................................................4-1
4.1 Introduction .....................................................................................................................4-1 4.2 General Purpose Interfaces ..............................................................................................4-1
4.2.1 MIL-STD-1553 Multiplex Data Bus.................................................................4-1 4.2.2 ARINC 429 Digital Information Transfer System.............................................4-2
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CONTENTS (Continued)
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4.2.3 Uses of the MIL-STD-1553 and ARINC 429 Interfaces ..................................4-2 4.2.3.1 Control-and-Display Unit..................................................................4-2 4.2.3.2 Data Loader System..........................................................................4-5 4.2.3.3 Flight Instrument Interface Unit........................................................4-6 4.2.3.4 Inertial Navigation Systems...............................................................4-8
4.3 Precise Time and Time Interval Interface.........................................................................4-9 4.3.1 Introduction.......................................................................................................4-9 4.3.2 Precise Time Inputs...........................................................................................4-9 4.3.3 Precise Time Outputs ........................................................................................4-9
4.4 Roll/Pitch/Heading/Water-Speed Analog Input Interface..............................................4-10 4.5 Instrumentation Port Interface .......................................................................................4-10 4.6 RS-232 Interface............................................................................................................4-10 4.7 Barometric Altimeter Interface ......................................................................................4-10 4.8 GPS Interface Options...................................................................................................4-11
4.8.1 Introduction.....................................................................................................4-11 4.8.2 Implementing a New Interface in an Existing GPS Receiver...........................4-11 4.8.3 Redesign of HV Interfaces to Accommodate an
Existing GPS Receiver .................................................................................4-11 4.8.4 Separate Development of an Interface Box.....................................................4-11
CHAPTER 5: ANTENNA SUBSYSTEMS......................................................................................5-1 5.1 Introduction .....................................................................................................................5-1 5.2 FRPA...............................................................................................................................5-1 5.2.1 General Characteristics......................................................................................5-1 5.2.2 FRPA Types......................................................................................................5-2 5.3 CRPA Equipment............................................................................................................5-4
CHAPTER 6: SERVICE COVERAGE, SERVICE AVAILABILITY, AND SERVICE RELIABILITY; SATELLITE SELECTION CRITERIA AND FIGURE OF MERIT DESCRIPTION..........................................................................................6-1
6.1 Service Coverage, Service Availability, And Service Reliability ......................................6-1 6.1.1 Parameter Definitions........................................................................................6-1 6.1.2 Service Coverage Characteristics ......................................................................6-3 6.1.2.1 Service Coverage Standards .............................................................6-3 6.1.2.2 The GPS 24-Satellite Constellation...................................................6-3 6.1.2.3 Expected Service Coverage Characteristics ......................................6-4 6.1.3 Service Availability Characteristics....................................................................6-5 6.1.3.1 Service Availability Standards...........................................................6-5 6.1.3.2 Satellite Outage Effects on Service Availability ................................6-5 6.1.3.3 Expected Service Availability Characteristics....................................6-6 6.1.4 Service Reliability Characteristics......................................................................6-8 6.1.4.1 Service Reliability Standards.............................................................6-8 6.1.4.2 GPS Service Failure Characteristics..................................................6-9
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CONTENTS (Continued)
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6.1.4.3 Failure Frequency Estimate...............................................................6-9 6.1.4.4 Failure Duration Estimate .................................................................6-9 6.1.4.5 Failure Magnitude and Behavior.....................................................6-10 6.1.4.6 User Global Distribution and Failure Visibility................................6-10 6.1.4.7 Satellite Use in the Position Solution ..............................................6-10 6.1.4.8 Failure Effect on Position Solution..................................................6-11 6.1.4.9 Expected Service Reliability Characteristics....................................6-11 6.1.5 Additional Commentary ..................................................................................6-11 6.1.5.1 24 Operational Satellites and Service Availability ...........................6-11 6.1.5.2 PDOP Less Than Six ......................................................................6-13 6.1.5.3 Four-Satellite Solution and Five-Degree Mask Angle.....................6-13 6.1.5.4 Integrity Checking...........................................................................6-14 6.1.5.5 Summary of the Commentary .........................................................6-15 6.2 Satellite Selection Criteria..............................................................................................6-15 6.2.1 Introduction.....................................................................................................6-15 6.2.2 Satellite Health ................................................................................................6-15 6.2.3 Geometric Dilution of Precision......................................................................6-16 6.2.4 User Range Accuracy......................................................................................6-16 6.2.5 Satellite Elevation Angle .................................................................................6-16 6.2.6 External Aids...................................................................................................6-16 6.3 Figure Of Merit (FOM) .................................................................................................6-17
CHAPTER 7: AIDING OPTIONS FOR A GPS RECEIVER ..........................................................7-1 7.1 Types of Aiding ...............................................................................................................7-1 7.2 Aiding During Initial Acquisition .....................................................................................7-2 7.2.1 Position and Velocity Aiding.............................................................................7-2 7.2.2 Time Aiding.......................................................................................................7-2 7.2.3 Almanac Data....................................................................................................7-2 7.2.4 Effect On TTFF.................................................................................................7-2 7.3 Aiding to Translate Navigation Solution..........................................................................7-3 7.4 Aiding to Replace a Satellite Measurement......................................................................7-3 7.4.1 Clock Aiding .....................................................................................................7-4 7.4.2 Altitude Aiding..................................................................................................7-4 7.5 Aiding to Maintain Satellite Track...................................................................................7-4
CHAPTER 8: POSSIBLE INTEGRATIONS OF GPS.....................................................................8-1 8.1 Introduction .....................................................................................................................8-1 8.2 Mission Requirements......................................................................................................8-2 8.3 Integration Architectures .................................................................................................8-3 8.3.1 GPS Stand-Alone/Baro/Clock Aided................................................................8-3 8.3.2 GPS/INS Integrations .......................................................................................8-4 8.3.3 GPS and Mission Computer/Databus Emulator................................................8-5 8.3.4 GPS in a 1553 Databus Configuration ..............................................................8-5
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CONTENTS (Continued)
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8.3.5 Embedded GPS.................................................................................................8-6 8.4 GPS and Transit/Omega/Loran-C....................................................................................8-6
CHAPTER 9: GPS AND KALMAN FILTERING...........................................................................9-1 9.1 Introduction .....................................................................................................................9-1 9.2 Kalman Filter Principle.....................................................................................................9-1 9.2.1 Kalman Filter Model..........................................................................................9-2 9.2.1.1 The System Dynamics Process..........................................................9-2 9.2.1.2 The Measurement Process ................................................................9-2 9.2.2 Kalman Filter Algorithm....................................................................................9-3 9.2.2.1 Propagation.......................................................................................9-3 9.2.2.2 Update ..............................................................................................9-4 9.2.2.3 Initial Conditions...............................................................................9-6 9.3 Kalman Filtering for Unaided GPS ..................................................................................9-6 9.3.1 The GPS Navigation Process ............................................................................9-6 9.3.2 The GPS Navigation Equation..........................................................................9-7 9.3.3 The GPS Kalman Filter Model..........................................................................9-8 9.3.4 GPS Augmented Kalman Filter.......................................................................9-10 9.3.5 GPS Kalman Filter Tuning ..............................................................................9-10 9.4 Kalman Filtering for Aided/Integrated GPS...................................................................9-11 9.4.1 The Integrated Navigation Solution ................................................................9-11 9.4.2 Kalman Filtering and GPS/INS .......................................................................9-11 9.4.2.1 System Architecture........................................................................9-11 9.4.2.2 The INS Navigation Process...........................................................9-14 9.4.2.3 The INS Kalman Filter States .........................................................9-16 9.4.3 Kalman Filtering and GPS/Precise Clock ........................................................9-16 9.4.4 Kalman Filtering and GPS/Barometric Altimeter ............................................9-16 9.4.5 Kalman Filtering and GPS/AHRS ...................................................................9-17
CHAPTER 10: DIFFERENTIAL GPS............................................................................................10-1 10.1 Introduction ...................................................................................................................10-1 10.2 DGPS Concept ..............................................................................................................10-2 10.3 DGPS Implementation Types ........................................................................................10-3 10.3.1 Ranging-Code Differential...............................................................................10-3 10.3.2 Carrier-Phase Differential................................................................................10-4 10.3.3 DGPS Data Link Implementations..................................................................10-5 10.3.4 Local Area and Wide Area Systems................................................................10-6 10.4 Solution Error Sources ..................................................................................................10-6 10.5 System Block Diagram ..................................................................................................10-9 10.6 DGPS Integrity............................................................................................................10-10
CHAPTER 11: SPECIAL APPLICATIONS FOR NAVSTAR GPS .............................................11-1 11.1 Introduction ...................................................................................................................11-1
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CONTENTS (Continued)
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11.2 DGPS Applications........................................................................................................11-1 11.2.1 Potential Uses of DGPS ..................................................................................11-1 11.2.1.1 Instrument Approach ......................................................................11-1 11.2.1.2 All Weather Helicopter Operations.................................................11-1 11.2.1.3 Narrow Channel Maritime Operations............................................11-2 11.2.1.4 Reference Station for Testing/Calibration of Navigation Equipment................................................................11-2 11.2.1.5 Surveying for Mapping and Positioning..........................................11-2 11.2.1.6 Blind Take-Off................................................................................11-2 11.2.2 DGPS Data Link .............................................................................................11-2
11.3 GPS Used as an Attitude Reference System..................................................................11-3 11.3.1 Introduction.....................................................................................................11-3 11.3.2 Concept of Operation......................................................................................11-3 11.3.3 3-D Attitude Reference System.......................................................................11-4 11.3.4 Use of Multiple Receivers and a Reference Oscillator.....................................11-5 11.3.5 Error Sources and Degradation of Performance .............................................11-5 11.3.5.1 Absolute Position Uncertainty.........................................................11-5 11.3.5.2 PDOP..............................................................................................11-6 11.3.5.3 Antenna Location............................................................................11-6 11.3.5.4 Antenna Position Difference Uncertainty in the Body Frame ......................................................................11-6 11.3.5.5 Measurement Accuracy and Error Budget......................................11-6
11.4 Precise Time and GPS ...................................................................................................11-7 11.4.1 Introduction.....................................................................................................11-7 11.4.2 Applications of Precise Time...........................................................................11-7 11.4.3 Interrelationship Between Different Definitions of Time.................................11-7 11.4.3.1 Time Based on the Rotation of the Earth On Its Axis..................................................................................11-7 11.4.3.2 Atomic Time/UTC Time.................................................................11-8 11.4.3.3 GPS Time .......................................................................................11-9 11.4.4 Precise Time Dissemination from GPS............................................................11-9 11.4.4.1 Precise Time Dissemination Under Dynamic Conditions.................................................................................11-12 11.4.4.2 Reduced Time Accuracy Due to SA.............................................11-13 11.4.5 Time Transfer Using GPS .............................................................................11-14 11.4.5.1 Coordinated Simultaneous-View Time Transfer...........................11-14 11.4.5.2 Coordinated Simultaneous-View Time Transfer with USNO...............................................................................11-14
11.5 Satellite Orbit Determination Using GPS.....................................................................11-15
CHAPTER 12: GPS INTEGRITY AND CIVIL AVIATION ........................................................12-1 12.1 Introduction ...................................................................................................................12-1 12.2 Military Use of National Airspace..................................................................................12-2
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CONTENTS (Continued)
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12.3 Civil Aviation Authorities, Agencies, and Organizations ...............................................12-2 12.3.1 Regulatory Authorities ....................................................................................12-2 12.3.2 Advisory Groups.............................................................................................12-3 12.3.3 Industry Groups ..............................................................................................12-3 12.3.4 Civil Aviation Coordination with the U.S. and U.S. DoD...............................12-3
12.4 Primary Civil Aviation Concerns With GPS ..................................................................12-4 12.4.1 Integrity Requirements....................................................................................12-4 12.4.2 Required Navigation Performance ..................................................................12-5 12.4.3 Integrity Assurance..........................................................................................12-6
CHAPTER 13: DIGITAL MAPS....................................................................................................13-1 13.1 Introduction ...................................................................................................................13-1 13.2 What Is A Digital Map?.................................................................................................13-1 13.2.1 Digitized Paper Maps......................................................................................13-1 13.2.2 Digital Database Maps ....................................................................................13-2 13.2.3 HYBRID Maps...............................................................................................13-2 13.3 Navigation Maps and Tactical Maps..............................................................................13-2 13.3.1 Use of Digital Maps for Navigation.................................................................13-2 13.3.2 Use of Digital Maps for Tactical Displays.......................................................13-3 13.3.3 Improvement of Common Reference Grids.....................................................13-3 13.3.3.1 Improved Gridlock..........................................................................13-4 13.3.3.2 Geodetic Gridlock...........................................................................13-4 13.3.3.3 Sensor Calibration...........................................................................13-4 13.3.3.4 OTHT Operations...........................................................................13-4 13.4 Other Issues Concerning Digital Maps and GPS ...........................................................13-5 13.4.1 Electrical Interface Between the Digital Map Display and the GPS Receiver.....................................................................................13-5 13.4.2 Digital Maps Accuracy....................................................................................13-5 13.4.3 Map Datums....................................................................................................13-5
ANNEX A: GLONASS: RUSSIAN'S EQUIVALENT NAVIGATION SYSTEM........................A-1 A.1 Historical Perspective .....................................................................................................A-1 A.2 Purpose of Global Satellite Navigation Systems .............................................................A-1 A.3 System Accuracy ............................................................................................................A-2 A.4 Monitor and Control Subsystem.....................................................................................A-2 A.5 Space Segment ...............................................................................................................A-3 A.6 Maneuvering in Orbit......................................................................................................A-5 A.7 Spacecraft Description....................................................................................................A-6 A.8 Satellite Launch Program................................................................................................A-7 A.9 Transmission Frequencies.............................................................................................A-10 A.10 Transmission Powers and Protection Ratio ..................................................................A-11 A.11 Information Transmission, Bandwidth and Code Rates................................................A-11 A.12 Ranging Codes..............................................................................................................A-12
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CONTENTS (Continued)
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A.13 Navigation Data............................................................................................................A-12 A.14 Navigation Reference Frame.........................................................................................A-14 A.15 User Equipment............................................................................................................A-15 A.16 References ....................................................................................................................A-15
ANNEX B: WORLD GEODETIC SYSTEM 1984: A MODERN AND GLOBAL REFERENCE FRAME...........................................................................................B-1
B.1 Introduction ....................................................................................................................B-1 B.2 The Reference Frame......................................................................................................B-1 B.3 The Defining Parameters and Associated Constants.......................................................B-3 B.4 The Gravity Formula.......................................................................................................B-4 B.5 The Earth Gravitational Model .......................................................................................B-5 B.6 The Geoid.......................................................................................................................B-5 B.7 Relationship with Local Geodetic Datums......................................................................B-5 B.8 Accuracy.........................................................................................................................B-7 B.9 Summary.........................................................................................................................B-9 B.10 References ......................................................................................................................B-9
ANNEX C: BBS INFORMATION ..................................................................................................C-1 C.1 Introduction ....................................................................................................................C-1 C.2 BBS Listing ....................................................................................................................C-1
ANNEX D: IMPACT OF MULTIPATH.........................................................................................D-1
ANNEX E: DOCUMENTATION.................................................................................................... E-1 E.1 Introduction ....................................................................................................................E-1 E.2 ICDs ...............................................................................................................................E-1 E.3 Other Documentation ..................................................................................................... E-1 E.3.1 JPO Documents................................................................................................E-1 E.3.2 ION Documents...............................................................................................E-1 E.3.3 RTCM Document ............................................................................................E-1 E.3.4 RTCA Document.............................................................................................E-2 E.3.5 DoT Documents...............................................................................................E-2 E.3.6 Miscellaneous................................................................................................... E-2
ANNEX F: ABBREVIATIONS AND ACRONYMS...................................................................... F-1
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ILLUSTRATIONS

Figure

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1-1 Navstar GPS Major Segments...................................................................................................1-1 1-2 GPS Satellite Constellation........................................................................................................1-3 1-3 GPS Control Segment Locations...............................................................................................1-4 1-4 Monitor Station and Ground Antenna .......................................................................................1-5 1-5 The Navigation Message ...........................................................................................................1-8 1-6 Satellite Signal Modulation........................................................................................................1-9 1-7 GPS Signal Frequency Spectrum.............................................................................................1-10 1-8 Spread Spectrum Generation and Reconstruction ...................................................................1-11 1-9 Generic GPS Receiver Tracking System .................................................................................1-12 1-10 GPS Receiver Theory of Operation.........................................................................................1-16 2-1 Analog GPS Receiver Architecture ...........................................................................................2-7 2-2 Digital GPS Receiver Architecture ............................................................................................2-8 3-1 Dilution of Precision ..................................................................................................................3-4 3-2 Time-To-First-Fix (TTFF).......................................................................................................3-12 4-1 Example of a Dedicated CDU ...................................................................................................4-3 4-2 Example of a Multifunction CDU..............................................................................................4-4 4-3 Example of a Data Loader System ............................................................................................4-6 4-4 Flight Instruments and TACAN.................................................................................................4-8 4-5 Flight Instruments and GPS.......................................................................................................4-9 5-1 FRPA Spiral Helix .....................................................................................................................5-3 5-2 FRPA Bifilar Helix.....................................................................................................................5-3 5-3 FRPA Crossed Monopoles........................................................................................................5-3 5-4 FRPA Ground Plane..................................................................................................................5-4 6-1 Satellite Global Visibility Profile ................................................................................................6-4 7-1 Aiding Options for a GPS Receiver...........................................................................................7-1 8-1 GPS Stand-Alone Configuration ...............................................................................................8-3 8-2 GPS INS-Aided Configuration..................................................................................................8-4 8-3 Configuration with Mission Computer/Databus Emulator.........................................................8-5 8-4 GPS in 1553 Databus Configuration..........................................................................................8-6 9-1 Simplified Diagram of Kalman Filter..........................................................................................9-3 9-2 Geometry for GPS Measurement ................................................
and Defining Time ................................................................................................................11-8 11-3 Determination of GPS-UTC (USNO) Time Difference.........................................................11-10
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11-4 Uncoordinated Time Transfer Using GPS.............................................................................11-11 11-5 Coordinated Time Transfer Using GPS.................................................................................11-15 A-1 GLONASS Orbit Planes and Slots ...........................................................................................A-4 A-2 GLONASS L1 C/A and P(Y) Code Spectrum.......................................................................A-13 B-1 World Geodetic System 1984 Reference Frame.......................................................................B-2 B-2 WGS 84 Geoid (n=m=18 Truncation)......................................................................................B-6 D-1 Multipath Induced North Position Error...................................................................................D-3

TABLES

Table

Page

1-1 NTS and Block I Satellite Launch Dates and Status................................................................1-19 1-2 Block II Satellite Launch Dates and Status..............................................................................1-21 3-1 GPS PPS System Range Error-Budget .....................................................................................3-3 3-2 Time Error Budget ....................................................................................................................3-8 3-3 Precise Time Output Accuracy (95%) for a Typical PPS
P-code Receiver......................................................................................................................3-8 3-4 Precise Time Output Accuracy (95%) for a Typical SPS
C/A-code Receiver .................................................................................................................3-9 6-1 Service Coverage Standards ......................................................................................................6-3 6-2 Service Availability Standards....................................................................................................6-6 6-3 Service Availability as a Function of Specified Satellite
Outage Conditions..................................................................................................................6-7 6-4 Example of 3-Day Global Service Availability with Component Failure
on Worst Day .........................................................................................................................6-7 6-5 Example of 30-Day Global Service Availability Without Component Failure ...........................6-8 6-6 Service Reliability Standards......................................................................................................6-8 6-7 Probability of Operational Satellites.........................................................................................6-12 6-8 Service Coverage of a Typical 24-Satellite Constellation ........................................................6-12 6-9 Availability of the Integrity Decision........................................................................................6-14 6-10 FOM/TFOM Numerical Values and Estimated Errors............................................................6-18 10-1 PPS DGPS Error Budget ........................................................................................................10-8 11-1 Uncoordinated Time Transfer Using GPS PPS Receivers.....................................................11-11 11-2 Coordinated Time Transfer Using GPS PPS Receivers.........................................................11-12 12-1 Typical Range of Integrity Parameters.....................................................................................12-5 A-1 GLONASS Satellite Launches..................................................................................................A-8 A-2 GLONASS Transmitted Power..............................................................................................A-11 A-3 Almanacs ................................................................................................................................A-14 B-1 WGS 84 Ellipsoid Four Defining Parameters ...........................................................................B-3 B-2 Relevant Miscellaneous Constants and Conversion Factors .....................................................B-4 B-3 Transformation Parameters Local Geodetic Systems to WGS 84 ............................................B-8 B-4 Methods of Determining and Accuracy of WGS 84 Coordinates.............................................B-9

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CHAPTER 1: SYSTEM OVERVIEW
1.1 GENERAL DESCRIPTION The Navstar Global Positioning System (GPS) is a space-based radio-positioning and timetransfer system. GPS provides accurate position, velocity, and time (PVT) information to an unlimited number of suitably equipped ground, sea, air and space users. Passive PVT fixes are available world-wide in all-weathers in a world-wide common grid system. Normally GPS contains features which limit the full accuracy of the service only to authorized users and protection from spoofing (hostile imitation). GPS comprises three major system segments, Space, Control, and User (see Figur-e11).
Figure 1-1. Navstar GPS Major Segments The Space Segment consists of a nominal constellation of 24 Navstar satellites. Each satellite broadcasts RF ranging codes and a navigation data message. The Control Segment consists of a network of monitoring and control facilities which are used to manage the satellite constellation and update the satellite navigation data messages. The User Segment consists of a variety of radio navigation receivers specifically designed to receive, decode, and process the GPS satellite
1-1

ranging codes and navigation data messages. The Space, Control, and User Segments are described in more detail in paragraph 1.2.
The ranging codes broadcast by the satellites enable a GPS receiver to measure the transit time of the signals and thereby determine the range between each satellite and the receiver. The navigation data message enables a receiver to calculate the position of each satellite at the time the signals were transmitted. The receiver then uses this information to determine its own position, performing calculations similar to those performed by other distance-measuring navigation equipment. Conceptually, each range measurement defines a sphere centered on a satellite. The common intersection point of the spheres on or near the earth's surface defines the receiver position.
For GPS positioning, a minimum of four satellites are normally required to be simultaneously "in view" of the receiver, thus providing four range measurements. This enables the receiver to calculate the three unknown parameters representing its (3-D) position, as well as a fourth parameter representing the user clock error. Treating the user clock error as an unknown enables most receivers to be built with an inexpensive crystal oscillator rather than an expensive precision oscillator or atomic clock. Precise time estimates are required for precise positioning, since a time error of 3 nanoseconds is roughly equivalent to a range error of 1 metre. Less than four satellites can be used by a receiver if time or altitude is precisely known or if these parameters are available from an external source. A more detailed explanation of the GPS theory of operation is provided in paragraph 1.4.
1.2 SYSTEM OVERVIEW
1.2.1 Space Segment
The GPS Space Segment consists of 24 Navstar satellites in semi-synchronous (approximately 12hour) orbits. The satellites are arranged in six orbital planes with four satellites in each plane. The orbital planes have an inclination angle of 55 degrees relative to the earth's equator. The satellites have an average orbit altitude of 20200 kilometres (10900 nautical miles) above the surface of the earth. Figure 1-2 illustrates the GPS satellite constellation.
The satellites complete one orbit in approximately 11 hours and 58 minutes. Since the earth is rotating under the satellites, the satellites trace a track over the earths surface which repeats every 23 hours and 56 minutes. A user at a fixed location on the ground will observe the same satellite each day passing through the same track in the sky, but the satellite will rise and set four minutes earlier each day, due to the 4 minute difference between the rotational period of the earth and two orbital periods of a satellite. The satellites are positioned in the orbital planes so that four or more satellites, with a good geometric relationship for positioning, will normally be observable at every location on earth. The effect of geometric relationships on GPS positioning accuracy is explained in further detail in Chapter 3.
1-2

Figure 1-2. GPS Satellite Constellation
The satellites transmit ranging signals on two D-band frequencies: Link 1 (Ll ) at 1575.42 MHz and Link 2 (L2) at 1227.6 MHz. The satellite signals are transmitted using spread-spectrum techniques, employing two different ranging codes as spreading fictions, a 1.023 MHz coarse/acquisition code (C/A-code) on L1 and a 10.23 MHz precision code (P-code) on both L1 and L2. Either the C/A-code or the P-code can be used to determine the range between the satellite and the user, however, the P-code is normally encrypted and available only to authorized users. When encrypted, the P-code is known as the Y-code. A 50 Hz navigation message is superimposed on both the P(Y) -code and the C/A-code. The navigation message includes satellite clock-bias data, satellite ephemeris (precise orbital) data for the transmitting satellite, ionospheric signal-propagation correction data, and satellite almanac (coarse orbital) data for the entire constellation. Refer to paragraph 1.4 for additional details regarding the ranging codes and navigation message.
1.2.2 Control Segment
The Control Segment primarily consists of a Master Control Station (MCS), at Falcon Air Force Base (AFB) in Colorado Springs, USA, plus monitor stations (MS) and ground antemas (GA) at various locations around the world. The monitor stations are located at Falcon AFB, Hawaii,
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Kwajalein, Diego Garcia, and Ascension. All monitor stations except Hawaii and Falcon AFB are also equipped with ground antennas (see Figure 1-3). The Control Segment includes a Prelaunch Compatibility Station (PCS) located at Cape Canaveral, USA, and a back-up MCS capability.
Figure 1-3. GPS Control Segment Locations
The MCS is the central processing facility for the Control Segment and is responsible for monitoring and managing the satellite constellation. The MCS functions include control of satellite station-keeping maneuvers, reconfiguration of redundant satellite equipment, regularly updating the navigation messages transmitted by the satellites, and various other satellite health monitoring and maintenance activities. The monitor stations passively track all GPS satellites in view, collecting ranging data from each satellite. This information is transmitted to the MCS where the satellite ephemeris and clock parameters are estimated and predicted. The MCS uses the ground antennas to periodically upload the ephemeris and clock data to each satellite for retransmission in the navigation message. Communications between the MCS the MS and GA are typically accomplished via the U.S. Defense Satellite Communication System (DSCS). The navigation message update function is graphically depicted in Figure 1-4.
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Figure 1-4. Monitor Station and Ground Antenna
The PCS primarily operates under control of the MCS to support prelaunch compatibility testing of GPS satellites via a cable interface. The PCS also includes an RF transmit/receive capability that can serve as a Control Segment ground antenna, if necessary. The U.S. Air Force Satellite Control Network (AFSCN) consists of a multipurpose worldwide network of ground- and spacebased satellite control facilities. Various AFSCN resources are available to support GPS but are not dedicated exclusively to GPS.
1.2.3 User Segment
The User Segment consists of receivers specifically designed to receive, decode, and process the GPS satellite signals. Receivers can be stand-alone, integrated with or embedded into other systems. GPS receivers can vary significantly in design and function, depending on their application for navigation, accurate positioning, time transfer, surveying and attitude reference. Chapter 2 provides a general description of GPS receiver types and intended applications.
1.3 GPS SERVICES
Two levels of service are provided by the GPS, the Precise Positioning Service (PPS) and the Standard Positioning Service (SPS).
1.3.1 Precise Positioning Service
The PPS is an accurate positioning velocity and timing service which is available only to authorized users. The PPS is primarily intended for military purposes. Authorization to use the PPS is determined by the U.S. Department of Defense (DoD), based on internal U.S. defense requirements or international defense commitments. Authorized users of the PPS include U.S.
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military users, NATO military users, and other selected military and civilian users such as the Australian Defense Forces and the U.S. Defense Mapping Agency. The PPS is specified to provide 16 metres Spherical Error Probable (SEP) (3-D, 50%) positioning accuracy and 100 nanosecond (one sigma) Universal Coordinated Time (UTC) time transfer accuracy to authorized users. This is approximately equal to 37 metres (3-D, 95%) and 197 nanoseconds (95%) under typical system operating conditions. PPS receivers can achieve 0.2 metres per second 3-D velocity accuracy, but this is somewhat dependent on receiver design.
Access to the PPS is controlled by two features using cryptographic techniques, Selective Availability (SA) and Anti-Spoofing (A-S). SA is used to reduce GPS position, velocity, and time accuracy to the unauthorized users. SA operates by introducing pseudorandom errors into the satellite signals. The A-S feature is activated on all satellites to negate potential spoofing of the ranging signals. The technique encrypts the P-code into the Y-code. Users should note the C/A code is not protected against spoofing.
Encryption keys and techniques are provided to PPS users which allow them to remove the effects of SA and A-S and thereby attain the maximum accuracy of GPS. PPS receivers that have not been loaded with a valid cryptographic key will have the performance of an SPS receiver.
PPS receivers can use either the P(Y)-code or C/A-code or both. Maximum GPS accuracy is obtained using the P(Y)-code on both L1 and L2. P(Y)-code capable receivers commonly use the C/A-code to initially acquire GPS satellites.
1.3.2 Standard Positioning Service
The SPS is a less accurate positioning and timing service which is available to all GPS users. In peacetime, the level of SA is controlled to provide 100 metre (95%) horizontal accuracy which is approximately equal to 156 metres 3D (95%). SPS receivers can achieve approximately 337 nanosecond (95%) UTC time transfer accuracy. System accuracy degradations can be increased if it is necessary to do so, for example, to deny accuracy to a potential enemy in time of crisis or war. Only the President of the United States, acting through the U.S. National Command Authority, has the authority to change the level of SA to other than peacetime levels.
The SPS is primarily intended for civilian purposes, although it has potential peacetime military use. Refer to "Technical Characteristics of the Navstar GPS" for additional details regarding SPS performance characteristics.

1.4 GPS THEORY OF OPERATION

The ranging codes broadcast by the satellites enable a GPS receiver to measure the transit time

of the signals and thereby determine the range between a satellite and the user. The navigation

message provides data to calculate the position of each satellite at the time of signal transmission.

From this information, the user position coordinates and the user clock offset are calculated using

simultaneous equations.

Four satellites are normally required to be

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simultaneously "in view" of the receiver for 3-D positioning purposes. The following paragraphs give a description of the GPS satellite signals and GPS receiver operation.
1.4.1 GPS Satellite Signals
1.4.1.1 C/A-Code
The C/A-code consists of a 1023 bit pseudorandom noise (PRN) code with a clock rate of 1.023 MHz which repeats every 1 millisecond. The short length of the C/A-code sequence is designed to enable a receiver to rapidly acquire the satellite signals which helps the receiver transition to the longer P-code. A different PRN is assigned to each GPS satellite and selected from a set of codes called Gold codes. The Gold codes are designed to minimize the probability that a receiver will mistake one code for another (minimize the cross-correlation). The C/A-code is transmitted only on L1. The C/A-code is not encrypted and is therefore available to all users of GPS.
1.4.1.2 P(Y)-Code
The P-code is a 10.23 MHz PRN code sequence that is 267 days in length. Each of the GPS satellites is assigned a unique seven-day segment of this code that restarts every Saturday/Sunday midnight GPS time (GPS time is a continuous time scale maintained within 1 microsecond of UTC, plus or minus a whole number of leap seconds). The P-code is normally encrypted into the Y-code to protect the user from spoofing. Since the satellites have the capability to transmit either the P- or Y-code, it is often referred to as the P(Y)-code. The P(Y)-code is transmitted by each satellite on both L1 and L2. On L1, the P(Y)-code is 90 degrees out of carrier phase with the C/A-code.
1.4.1.3 Navigation Message
A 50 Hz navigation message is superimposed on both the P(Y) code and the C/A-code. The navigation message includes data unique to the transmitting satellite and data common to all satellites. The data contains the time of transmission of the message, a Hand Over Word (HOW) for the transition from C/A-code to P(Y)-code tracking, clock correction, ephemeris, and health data for the transmitting satellite, almanac and health data for all satellites, coefficients for the ionospheric delay model, and coefficients to calculate UTC.
The navigation message consists of 25 frames of data, each frame consisting of 1,500 bits. Each frame is divided into 5 subframes of 300 bits each (see Figure 1-5). At the 50 Hz transmission rate, it takes 6 seconds to receive a subframe, 30 seconds to receive one data frame, and 12.5 minutes to receive all 25 frames. Subframes 1, 2, and 3 have the same data format for all 25 frames. This allows the receiver to obtain critical satellite-specific data within 30 seconds. Subframe 1 contains the clock correction for the transmitting satellite, as well as parameters describing the accuracy and health of the broadcast signal. Subframes 2 and 3 contain ephemeris (precise orbital) parameters used to compute the location of the satellite for the positioning equations.
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BIT NO. SUBFRAME
1
SUBFRAME 2
SUBFRAME 3
SUBFRAME 4
SUBFRAME 5

0

30

60

TELEMETRY WORD

HANDOVER WORD

300

330

360

TELEMETRY WORD

HANDOVER WORD

600

630

660

TELEMETRY WORD

HANDOVER WORD

900

930

960

TELEMETRY WORD

HANDOVER WORD

1200

1230

1260

TELEMETRY WORD

HANDOVER WORD

(MULTIPLEX) (MULTIPLEX)

*12.5 MINUTES BEFORE THE ENTIRE MESSAGE REPEATS

CLOCK CORRECTION
EPHEMERIS
EPHEMERIS
MESSAGE (CHANGES THROUGH 25 FRAMES) ALMANAC/HEALTH STATUS
(CHANGES THROUGH 25 FRAMES)

300 6 SEC
600 12 SEC
900 18 SEC
1200 24 SEC
1500 30 SEC

Figure 1-5. The Navigation Message
Subframes 4 and 5 have data which cycle through the 25 data frames. They contain data which is common to all satellites and less critical for a receiver to acquire quickly. Subframes 4 and 5 contain almanac (coarse orbital) data and low-precision clock corrections, simplified health and configuration status for every satellite, user text messages, and the coefficients for the ionospheric model and UTC calculation. A comprehensive description of the navigation message is provided in "Technical Characteristics of the Navstar GPS", together with the standard algorithms needed to use the data correctly.
1.4.1.4 Satellite Signal Modulation
The L1 carrier is BPSK modulated by both the C/A- and P(Y)-codes plus the navigation message superimposed on both codes. The L2 carrier is BPSK modulated by the P(Y)-code superimposed with the navigation message. The BPSK technique reverses the carrier phase when the modulating code changes from logic 0 to 1 or 1 to 0. On L1, the C/A-code is 90 degrees out of phase with the P(Y)-code. Figure 1-6 shows this modulation scheme in schematic form.

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Figure 1-6. Satellite Signal Modulation
The BPSK modulation spreads the RF signals by the code bandwidth. The result is a symmetrical spreading of the signal around the L1 and L2 carriers. The C/A-code spreads the L1 signal power over a 2.046 MHz bandwidth centered at 1575.42 MHz. The P(Y)-code spreads the L1 and L2 signal powers over a 20.46 MHz bandwidth centered about 1575.42 MHz on L1 and 1227.6 MHz on L2. Figure 1-7 shows the L1 and L2 signal spectrum as it appears at the 0 dB gain receiver antenna at the Earth's surface. The C/A-code component of L1 signal has a power of -160 dBW (decibels with respect to one watt), the L1 P(Y)-code signal has a power of -163 dBW, and the L2 P(Y)-code signal has a power of -166 dBW.
1.4.2 GPS Receiver Operation
In order for the GPS receiver to calculate a PVT solution, it must:
Search for a PRN C/A code lock C/A code track, carrier track Obtain bit synchronization with the navigation message Obtain frame synchronization, ie obtain HOW and Z count Decode GUV or CVw Remove SA Transition to P(Y)-code, -code lock, -carrier lock Data lock on P(Y) code Search, acquire and track 2nd to 4th SVs, up to all in view
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Figure 1-7. GPS Signal Frequency Spectrum
Take range and range rate measuerments Solve for range equations P(Y) code measurements L2 to remove ionospheric delays and refine navigation solution.
Details of the operations are expanded below.
1.4.2.1 Satellite Selection
A typical satellite tracking sequence begins with the receiver determining which satellites are visible for it to track. If the receiver can immediately determine satellite visibility, the receiver will target a satellite to track and begin the acquisition process. Satellite visibility is determined based on the GPS satellite almanac and the initial receiver estimate (or user input) of time and position. If the receiver does not have the almanac and position information stored, the receiver enters a "search the sky" operation that systematically searches the PRN codes until lock is obtained on one of the satellites in view. Once one satellite is successfully tracked, the receiver can demodulate the navigation message data stream and acquire the current almanac as well as the health status of all the other satellites in the constellation.
Depending on its architecture, a receiver selects either a "best" subset of the visible satellites to track or uses all healthy satellites in view to determine an "all-in-view" PVT solution. The all-inview solution is usually more accurate than a four satellite solution although it requires a
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more complex receiver and receiver processing. The all-in-view solution is also more robust, since the temporary loss of a satellite signal (for example due to a physical obstruction near the receiver) does not disrupt the flow of PVT data while the receiver attempts to reacquire the lost signal. Many receivers will track more than four satellites, but less than all-in-view, as a compromise between complexity, accuracy, and robustness. Receivers that select a "best" subset do so based on geometry, estimated accuracy, or integrity. More detailed discussion of specific satellite selection criteria is provided in Chapter 6. 1.4.2.2 Satellite Signal Acquisition The satellite signal power at or near the earth's surface is less than the receivers thermal (natural) noise level, due to the spread spectrum modulation of the signal, orbital height and transmitting power of the satellite. To extract the satellite signal the receiver uses code correlation techniques. An internal replica of the incoming signal is generated and aligned with the received satellite signal. The receiver shifts the replica code to match the incoming code from the satellite. When the codes match, the satellite signal is compressed back into the original carrier frequency band. This process is illustrated in Figure 1-8.
Figure 1-8. Spread Spectrum Generation and Reconstruction 1-11

The delay in the receiver's code is a measure of the transit time of the signals between the satellite and the receiver's antenna and hence, the range between the satellite position and receiver position. This measurement is called a pseudorange measurement, rather than a range measurement, because the receiver's clock bias has not been removed.

Receivers typically use phase-locked-loop techniques to synchronize the receiver's internally generated code and carrier with the received satellite signal. A code tracking loop is used to track the C/A- and P-code signals while a carrier tracking loop is used to track the carrier frequency. The two tracking loops work together in an interactive process, aiding each other, in order to acquire and track the satellite signals. A generic GPS receiver tracking system is illustrated in Figure 1-9.

ANTENNA PREAMPLIFIER

CARRIER
TRACKING CHANNEL

PSEUDODELTA RANGE MEASUREMENT
50 Hz NAVIGATION DATA

RF CONVERTER

IF SIGNAL

DOPPLER EST.

ON-TIME EST.

FREQUENCY SYNTHESIZER

LO IF

CODE
TRACKING CHANNEL

PSEUDO-RANGE MEASUREMENTS

REFERENCE OSCILLATOR
Figure 1-9. Generic GPS Receiver Tracking System
1.4.2.3 Down Conversion
The received RF signal is converted, usually through two intermediate frequencies (IF), down to a frequency near the code baseband, that can be sampled by an analogue to digital (A/D) converter. Inphase and quadrature digital samples are taken to preserve the phase information in the received signal. The samples are usually two bits to reduce conversion losses. The sampling rate must be higher than the code chipping rate for a non return to zero code, that is, greater than 10.23 MHz for the P(Y)-code. To ensure the phase of the received signal is maintained, all local oscillators are derived from, and phased locked through, a series of synthesizers derived from the receiver's master oscillator. Following the A/D conversion there
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is a final phase rotation circuit that enables the doppler in the satellite signal to be precisely tracked.
1.4.2.4 Code Tracking
The code tracking loop is used to make pseudorange measurements between the GPS satellites and the GPS receiver. The receiver's code tracking loop generates a replica of the C/A-code of the targeted satellite. The estimated doppler is removed by the phase rotation circuit prior to the correlator.
In order to align the received signal with the internally generated replica, the internally generated code is systematically slewed past the received signal. Typically the output of the correlator is integrated over 1 to 10 ms. If correlation is not detected the phase of the internally generated code is advanced by one chip. If correlation is not detected after the whole code has been searched the doppler is adjusted and the process repeated until correlation is achieved. Code synchronization is initially maintained by also correlating the received signal with half chip early and late codes. A simple feedback system keeps the prompt ("on time") code correctly positioned. To extract the carrier which is still modulated by the navigation message, the prompt code is subtracted from the incoming signal. The delay that the receiver must add to the replica code to achieve synchronization (correlation), multiplied by the speed of light, is the pseudorange measurement. Once the carrier is reconstructed, the center frequency of the replica code is adjusted using Doppler measurements from the carrier tracking loop to achieve a precise frequency lock to the incoming signal, thereby allowing more precise pseudorange measurements. The bandwidth of the code tracking loop is typically 0.1 Hz, which implies that independent measurements are available at approximately 10 s intervals.
1.4.2.5 Carrier Tracking and Data Detection
The receiver tracks the satellite carrier by adjusting the frequency synthesizers to produce a stationary phase at the output of the code tracking loop. The inphase and quadrature components are used to calculate the carrier's phase and doppler. A data bit is detected by a sudden change in the phase of the detected signal. The bandwidth of the carrier tracking loop is typically 6 Hz for a military airborne receiver, resulting in independent measurements being available every 150 ms.
Doppler is measured to provide an estimate of the relative velocity between the receiver and the satellite. These measurements are typically termed pseudorange rate measurements or they can be integrated over regular time intervals to produce deltarange measurements.
The receiver uses the doppler measurements from four (or more) satellites to determine the receiver velocity (in three dimensions) plus the receiver's master oscillator frequency bias. The deltarange measurements of the carrier tracking loop are also used to aid the code tracking loop to ensure code tracking is maintained during dynamic maneuvers where the simple code tracking system would be unable to maintain lock.
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1.4.2.6 Data Demodulation
Once the carrier tracking loop is locked, the 50 Hz navigation data message can be read. Each subframe of the navigation message begins with a preamble contained in the Telemetry Word, enabling the receiver to detect the beginning of each subframe. Each subframe is identified by bits contained in the Handover Word (HOW), enabling the receiver to properly decode the subframe data.
1.4.2.7 P(Y)-Code Signal Acquisition
The one millisecond C/A-code length permits a relatively narrow search window for code correlation even if the receiver must "search the sky" to find the first satellite. However the week long P(Y)-code sequence at 10.23 MHz does not allow the same technique to be used. Precise time must be known by the receiver in order to start the code generator within a few hundred chips of the correlation point of the incoming signal. The HOW contained in the GPS navigation message provides satellite time and hence the P(Y)-code phase information. A P(Y)-code receiver may attempt to acquire the P(Y)-code directly, without first acquiring the C/A-code, if it has accurate knowledge of position, time and satellite ephemeris from a recent navigation solution. External aiding and/or an enhanced acquisition technique are usually required to perform direct P(Y)-code acquisition.
1.4.2.8 PVT Calculations
When the receiver has collected pseudorange measurements, deltarange measurements, and navigation data from four (or more) satellites, it calculates the navigation solution, PVT. Each navigation data message contains precise orbital (ephemeris) parameters for the transmitting satellite, enabling a receiver to calculate the position of each satellite at the time the signals were transmitted. The ephemeris data is normally valid and can be used for precise navigation for a period of four hours following issue of a new data set by the satellite. New ephemeris data is transmitted by the satellites every two hours.
As illustrated in Figure 1-10, the receiver solves a minimum of four simultaneous pseudorange equations, with the receiver (3-D) position and clock offset as the four unknown variables. Each equation is an expression of the principle that the true range (the difference between the pseudorange and the receiver clock offset) is equal to the distance between the known satellite position and the unknown receiver position. This principle is expressed below mathematically using the same notation as Figure 1-10.
R - CB = c∆t - CB = (X - U X )2 + (Y - UY )2 + (Z - UZ )2
These are simplified versions of the equations actually used by GPS receivers. A receiver also obtains corrections derived from the navigation messages which it applies to the pseudoranges. These include corrections for the satellite clock offset, relativistic effects, ionospheric signal propagation delays. Dual frequency receivers can measure the delay between the L1 and L2 P(Y)-codes, if available, to calculate an ionospheric correction. Single frequency (either C/A-
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or P(Y)-code) receivers use parameters transmitted in the navigation message to be used in an ionospheric model. The receiver (3-D) velocity and frequency offset are calculated using similar equations, using deltaranges instead of pseudoranges.
The PVT calculations described here result in a series of individual point solutions. For receivers that are required to provide a navigation solution under dynamic conditions a smoothed or filtered solution that is less sensitive to measurement noise is employed. One of the most common types of filters used in GPS receivers is the Kalman filter. Kalman filtering is described in detail in Chapter 9.
The rate at which GPS receivers calculate the PVT solution is governed by their application. For flight control applications a 10 Hz rate is required whereas in handheld equipment a fix may only be required once every 4 to 5 seconds or at even longer intervals. A 1 Hz rate is typical for many equipments. In this scenario pseudorange measurements are typically only made every 4 to 5 seconds; pseudorange rate measurements are made more frequently and can be used to propagate the filter solution between updates. If a Kalman filter is used the measurements may be incorporated independently into the filter removing the requirement for symmetrical measurements from all channels. The filter also allows the solution to be extrapolated if measurements are interrupted, or data is available from other navigation sensors.
A minimum of four satellites are normally required to be simultaneously "in view" of the receiver, thus providing four pseudorange and four deltarange measurements. Treating the user clock errors as unknowns enable most receivers to be built with an inexpensive crystal oscillator rather than an expensive precision oscillator or atomic clock. Less than four satellites can be used by a receiver if time or altitude are precisely known or if these parameters are available from an external source.
GPS receivers perform initial position and velocity calculations using an earth-centered earthfixed (ECEF) coordinate system. Results may be converted to an earth model (geoid) defined by the World Geodetic System 1984 (WGS 84). WGS 84 provides a worldwide common grid system that may be translated into local coordinate systems or map datums. (Local map datums are a best fit to the local shape of the earth and not valid worldwide.) For more details regarding WGS 84, refer to Annex B. For more details regarding how a receiver uses WGS 84, refer to "Technical Characteristics of the Navstar GPS".
For navigation purposes, it is usually necessary for a GPS receiver to output positions in terms of magnetic North rather than true North as defined by WGS 84. For details regarding how the receiver calculates the magnetic variation from true North, refer to "Technical Characteristics of the Navstar GPS".
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Figure 1-10. GPS Receiver Theory of Operation 1-16

1.4.2.9 Degraded Operation and Aiding
During periods of high levels of jamming, the receiver may not be able to maintain both code and carrier tracking. The receiver normally has the capability to maintain code tracking even when carrier tracking is no longer possible. If only code tracking is available, the receiver will slew the locally generated carrier and code signals based on predicted rather than measured Doppler shifts. These predictions are performed by the receiver processor, which may have additional PVT information available from an external aiding source. See Chapter 7 for additional discussion of GPS receiver aiding.
1.5 PROGRAM MANAGEMENT
1.5.1 System Development and Management
The United States Air Force (USAF), Air Force Materiel Command, Space and Missile Center (SMC), Navstar GPS Joint Program Office (JPO) has total system responsibility for the GPS. The SMC and GPS JPO are located at the Los Angeles Air Force Base (AFB) in Los Angeles, California. The GPS JPO is manned by personnel from the USAF, US Navy, US Army, US Marine Corps, US Department of Transportation, US Defense Mapping Agency. NATO Nations and Australia may have representatives stationed at the JPO. The GPS JPO was responsible for development of the Control and Space Segments and is responsible for acquisition of replenishment satellites and common user equipment (UE) for all military services. The GPS JPO also provides technical support, security guidance, technical specification development, interface control documents, and implementation guidelines. NATO and other allied Nations have established Memoranda of Understanding with the United States which provides access to PPS, interchange of technical information, and the ability to purchase or locally manufacture PPS GPS UE.
The GPS JPO is supported by the Launch Vehicle System Program Office (SPO) and the Network SPO, also located at the SMC. The Launch Vehicle SPO provides the expendable boosters used to launch the Navstar satellites. The Network SPO is responsible for continuing development of the multi-use AFSCN. GPS JPO program management operations are also supported by the User Equipment Support Program Manager located at the Warner-Robbins Air Logistics Center in Warner Robbins, Georgia and by Detachment 25 (from the Sacramento Air Logistics Center) located at Colorado Springs, Colorado.
1.5.2 System Requirements, Planning, and Operations
The USAF Space Command (AFSPC) is responsible for GPS requirements, planning, and operations. Headquarters of the AFSPC and the requirements and planning functions are located at Peterson AFB in Colorado Springs, Colorado. Various agencies within the USAF Space Command (AFSPC) operate and maintain the Control Segment, prepare and launch the Navstar satellites, manage the operational constellation, and interface with the GPS user community. Elements of the AFSPC Fiftieth Space Wing (50SPW) are responsible for launch, early orbit support, and continued day-to-day operations of the GPS satellites.
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The First Space Operations Squadron (1SOPS) of the 50SPW, located at Falcon AFB in Colorado Springs, Colorado, provides launch and early-orbit support for the GPS satellites. The early orbit support includes control of the Navstar satellites to deploy solar arrays, perform stabilization maneuvers, and complete other procedures to make the satellites ready for service. The 1SOPS can also provide backup capability for critical day-to-day commanding procedures if necessary. When a satellite is ready for service, command is transferred to the Second Space Operations Squadron (2SOPS) of the 50SPW for payload turn-on and continued operations. The 2SOPS has responsibility for day-to-day operations and overall constellation management. The 2SOPS is also located at Falcon AFB.
The Forty-Fifth Space Wing (45SPW) of the AFSPC is responsible for management of Navstar pre-launch operations, including receiving of the satellites, storage on the ground if necessary, mating to the launch vehicle, and integration and compatibility testing. The 45SPW is located at Cape Canaveral Air Force Station, Florida, which is the launch site for the GPS satellites.
1.6 GPS PROGRAM HISTORY
1.6.1 Pre-Concept Validation (1960s-1972)
Since the early 1960s various U.S. agencies have had navigation satellite programs. The John Hopkins' Applied Research Laboratory sponsored the TRANSIT program and the U.S. Navy (USN) sponsored the TIMATION (TIMe navigATION) program. TIMATION was a program to advance the state of the art for two-dimensional (latitude and longitude) navigation. TRANSIT became operational in 1964 and is currently providing navigation service to low dynamic vehicles such as ships. It is scheduled to be phased out in 1996. The USAF conducted concept studies to assess a three-dimensional (latitude, longitude, and altitude) navigationsstyem called 621B.
1.6.2 Phase I - Concept Validation (1973-1979)
A memorandum issued by the US Deputy Secretary of Defense on 17 April 1973 designated the USAF as the executive service to consolidate the TIMATION and 621B concepts into a comprehensive all-weather navigation system named Navstar GPS. The Navstar GPS JPO was established on 1 July 1973.
Two experimental Navigation Technology Satellites (NTS) were built and launched to support concept validation of the GPS. The first true GPS signals from space came from NTS-2. NTS-2 was launched on an Atlas booster from Vandenberg AFB in June 1977 but malfunctioned after only 8 months. The first Navstar GPS Block I (research and development) satellite was launched in February 1978. A total of 11 Block I satellites were launched between 1978 and 1985. All of the Block I satellites were launched from Vandenberg AFB using the Atlas booster. Block I satellites did not incorporate SA or A-S features. As of June 1995 only one Block I satellite remained operational. Table 1-1 contains the launch dates and status (as of June 1995) of the NTS and Block I satellites.
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Table 1-1. NTS and Block I Satellite Launch Dates and Status

Navstar Number
NTS-1 NTS-2 I-1 I-2 I-3 I-4 I-5 I-6 I-7 I-8 I-9 I-10 I-11

Space Vehicle No.
(SVN)
1 2 3 4 5 6 7 8 9 10 11

PRN Code Number Launch Date

-

14 Jul 74

-

23 Jun 77

-

22 Feb 78

-

12 May 78

-

06 Oct 78

-

11 Dec 78

-

09 Feb 80

-

26 Apr 80

-

18 Dec 81

-

14 Jul 83

-

13 Jun 84

12

08 Sep 84

-

09 Oct 85

Status (June 95)
Deactivated Deactivated Jan 78 Deactivated 25 Jan 80 Deactivated 30 Aug 80 Deactivated 19 Apr 92 Deactivated 06 Sep 86 Deactivated 28 Nov 83 Deactivated 05 Mar 91 Launch Failure Deactivated 4 May 93 Deactivated 28 Feb 94 Operational Deactivated 14 Apr 94

The first Control Segment consisted of a control station, ground antenna, and monitor station located at Vandenberg AFB in California, supported by additional monitor stations located at Elmendorf AFB in Alaska, Anderson AFB in Guam, and the Naval Communications Station in Hawaii. This Phase I Control Segment was designated the Initial Control System (ICS).
The first user equipment (UE) testing began at Yuma Proving Ground (YPG) in March 1977 using ground transmitters to simulate the GPS satellites. As the Block I satellites were launched, a combination of satellites and ground transmitters were used for testing until December 1978, when four satellites were available to provide limited 3-D navigation capability. Shipborne UE was tested off the coast of California starting in October 1978 when three GPS satellites were available for two-dimensional (-2D) navigation.
1.6.3 Phase II - Full Scale Development (1979-1985)
In September 1980, a contract was awarded to upgrade and operate the ICS, as well as develop an Operational Control System (OCS). The ICS upgrades ensured continued support to the UE test team while the OCS was being developed. OCS equipment was delivered to Vandenberg AFB in May 1985. In October 1985, after installation and initial testing, the OCS conducted dual operations with the ICS. The OCS equipment was moved from Vandenberg to its permanent site at Falcon AFB by the end of 1985. In December 1980, the contractor was
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selected to provide 28 Block II (operational) Navstar GPS satellites. Development of the satellites continued throughout Phase II.
Phase II for the User Segment was divided into two parts. In Phase IIA, starting in July 1979, four contractors were selected to conduct performance analyses and preliminary design of UE. In Phase IIB, starting in 1982, two of the four contractors were selected to continue UE development. Phase IIB included design refinement, fabrication of prototypes, qualification testing, and extensive field testing of the UE. The UE was tested at YPG and at sea. Testing at sea was conducted by Naval Ocean Systems Center located in San Diego, California.
1.6.4 Phase III - Production and Deployment (1986 to Present)
1.6.4.1 Space Segment (1986 to Present)
The Block II satellites were originally designed to be launched aboard the Space Transportation System (Space Shuttle). Following an accident with the Space Shuttle Challenger in 1986, the Block II satellite-to-launch-vehicle interface was modified to enable launch aboard the Delta II booster. The first Block II satellite was launched on 14 February 1989. The combined constellation of Block I and Block II satellites achieved worldwide two-dimensional positioning capability in June 1991. Worldwide 3-D capability was achieved in 1993. The Initial Operational Capability (IOC) was declared on 8 December 1993. A full 24-satellite constellation of Block II satellites was achieved in April 1994. The military Full Operational Capability is planned for 1995. The remaining Block II satellites will be launched on demand. Table 1-2 is a summary of the Block II launch dates and status.
In June of 1989 a contract was awarded for 20 GPS replenishment satellites, designated Block IIR. The Block IIR satellites will have the capability to autonomously generate their own navigation messages. The Block IIR production schedule may allow a first launch as early as August 1996. In 1994, efforts were begun by the GPS JPO to procure additional Navstar satellites to sustain the GPS satellite constellation past the year 2000. These satellites are designated Block IIF (Follow-On). The contract to provide the Block IIF satellites is planned for November 1995. The planned production schedule supports a first launch in the year 2001.
In 1994 the GPS JPO also began studies for an Augmented GPS (AGPS). The AGPS concept is to enhance the availability, accuracy and integrity of the GPS system using up to six geostationary AGPS satellites. The satellites would broadcast integrity information and range corrections for all GPS satellites via GPS-like ranging signals.
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Table 1-2. Block II Satellite Launch Dates and Status

Navstar Number II-1 II-2 II-3 II-4 II-5 II-6 II-7 II-8 II-9 II-10 II-11 II-12 II-13 II-14 II-15 II-16 II-17 II-18 II-19 II-20 II-21 II-22 II-23 II-24 II-25 II-26 II-27 II-28

Space Vehicle No. (SVN) 14 13 16 19 17 18 20 21 15 23 24 25 28 26 27 32 29 22 31 37 39 35 34 36 TBD TBD TBD TBD

PRN Code Number 14 2 16 19 17 18 20 21 15 23 24 25 28 26 27 01 29 22 31 07 09 05 04 06 TBD TBD TBD TBD

Launch Date Status (June 95)

14 Feb 89

Operational

10 Jun 89

Operational

17 Aug 89

Operational

21 Oct 89

Operational

11 Dec 89

Operational

24 Jan 90

Operational

25 Mar 90

Operational

02 Aug 90

Operational

01 Oct 90

Operational

26 Nov 90

Operational

03 Jul 91

Operational

23 Feb 92

Operational

09 Apr 92

Operational

07 Jul 92

Operational

09 Sep 92

Operational

22 Nov 92

Operational

18 Dec 92

Operational

02 Feb 93

Operational

29 Mar 93

Operational

12 May 93

Operational

26 Jun 93

Operational

30 Aug 93

Operational

26 Oct 93

Operational

09 Mar 94

Operational

TBD

To Be Launched

TBD

To Be Launched

TBD

To Be Launched

TBD

To Be Launched

1.6.4.2 Control Segment (1986 to Present)
The GPS OCS achieved Full Operational Capability (FOC) in December 1986. In March 1986, the ICS at Vandenberg AFB was deactivated. In December 1989, verification of the OCS operational capability was completed by the USAF Operational Test and Evaluation Center. Turnover of the OCS to AFSPC was accomplished in June 1990. Since then, Control Segment development activities have been limited to upgrades of the operational software and additions
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to the equipment and facilities. The OCS has been augmented with a transportable GA capability and Back-Up MCS capability.
1.6.4.3 User Segment (1986 to Present)
1.6.4.3.1 GPS JPO Activities
In April 1985, the contractor was selected for the Phase III production GPS UE. Low rate initial production of the UE was begun and the first set was delivered to the JPO in June 1988. In January 1992, full rate production of the UE was approved. The Phase III production UE includes the 5-channel Receiver 3A (R-2332/AR) for airborne use, the 5-channel Receiver 3S (R2331/AR) for shipboard use, the 2-channel Receiver OH (R-2399/AR) and UH (R-2400/AR) for helicopter use, and the RPU-1 (R-2401/U) for manpack and ground vehicle use.
In 1989, a contract was awarded for 2-channel SPS C/A-code receivers to be used primarily for demonstration and training. These receivers are known as the Small Lightweight GPS Receiver (SLGR, AN/PSN-10). They are suitable for vehicle mounting or handheld use. In 1990, a large second purchase was made. Although originally intended for nontactical use, these receivers were used extensively in support of Operation Desert Shield and Operation Desert Storm.
In November 1990, a contract was awarded to develop a 5-channel 3/8 ATR (Air Transport Rack) size Miniature Airborne GPS Receiver (MAGR) for use in aircraft where space is severely limited. The contract to deliver operational models was awarded in April 1993 with the first delivery occurring in July 1994. Two versions of the MAGR have been produced. One version uses an RF interface directly from the antenna (R-2512/U) the other (R-2514/U) uses an IF (intermediate frequency) interface from an antenna electronics unit.
In February 1993, a contract was awarded to produce a hand-held PPS GPS receiver. Designated the Precision Lightweight GPS Receiver (PLGR, AN/PSN-11), it weighs less than 4 pounds, is self-contained as a handheld unit, and can be adapted for vehicle mounting. Delivery of the PLGR began in September 1993.
In the 1990s, the GPS JPO has continued to sponsor activities to improve the functions and performance of military GPS receivers. Activities are continuing that will improve anti-jamming performance of GPS antennas, antenna electronics units, and receiver signal processing. In 1994, procurement efforts were begun for a new Controlled Reception Pattern Antenna (CRPA). The new CRPA will be compatible with the form, fit, and function of the existing CRPA system procured by the JPO. Efforts are also underway that will allow Receiver Autonomous Integrity Monitoring (RAIM) to be implemented where enhanced GPS integrity or compatibility with civil aviation is desired. Other efforts are underway to add differential GPS (DGPS) to future military PPS receivers, to support new applications, such as precise positioning and aircraft precision approach. Additional programs that are underway or under consideration include a space-based GPS PPS receiver, a miniaturized PLGR, and a Survey GPS Receiver (SGR). Since 1993, the GPS JPO has been developing standards for a next generation PPS receiver module that can be embedded in other military systems. The JPO will not procure embedded GPS receivers (EGRs), but will provide technical support so that other military programs can procure the EGR as part of another system.
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The JPO has released an EGR Guidelines document which contains EGR interface, design, and performance requirements, as well as general guidance material regarding the EGR and host system. The document also includes specific guidance for integrating GPS with inertial or Doppler navigation systems.
The JPO EGR effort is evolving into a standard for a GPS Receiver Applications Module (GRAM). The GRAM will consist of a family of standard EGR modules suitable for a variety of embedded applications. The GRAM standard will define several EGR physical configurations conforming to standard modular architectures, such as the Standard Electronic Module (SEM) and Versa Module Europa (VME). The standard will include specifications for advanced functions, such as local- and wide-area DGPS corrections and receiver-based integrity enhancements (RAIM). The standard will also accommodate the next-generation GPS receiver security module known as the Selective Availability/Anti-Spoofing Module (SAASM).
1.6.4.3.2 International Military UE and Commercial UE
Phase III of the GPS program has seen a tremendous expansion in the development and production of international military UE and commercial UE. Military UE is being produced by participating NATO nations including Canada, France, Germany, Italy, and the United Kingdom.
In addition, a wide variety of commercial SPS UE has been developed by manufacturers around the world for many different applications. Some of these receivers have been acquired by Military and Government authorities for nontactical applications such as surveying, test support, and training.
1.6.4.3.3 User Equipment Testing
Development Test and Evaluation (DT&E) and OT&E have included test and evaluation of:
a. Integrated GPS/host vehicle navigation system performance b. Phase II and (early) Phase III deficiency correctiosn c. Reliability and maintainability of the GPS UE d. Operational effectiveness of the GPS UE against jamming and spoofing e. The SA and A-S features
In addition to U.S.-sponsored test efforts, Australia, Canada, Denmark, Germany, the Netherlands, Norway, and the United Kingdom conducted an extensive Phase III International Test Program in cooperation with the JPO. These countries were joined by France, Greece, Portugal, Spain, and Turkey for a subsequent International Test Program that focused exclusively on the PLGR.
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CHAPTER 2: TYPES OF GPS RECEIVERS AND THEIR INTENDED APPLICATIONS
2.1 GPS RECEIVER ARCHITECTURES
Modern military GPS receivers use predominantly a continuous satellite tracking architecture. However, some receivers use alternative architectures, either sequential or multiplex tracking to reduce hardware complexity.
2.1.1 Continuous Receivers
A continuous tracking receiver has five or more hardware channels to track four satellites simultaneously plus other channels to acquire new satellites. Due to their greater complexity, these receivers were traditionally the most expensive but offer the best performance and versatility. The multi-channel receiver uses the fifth channel to read the NAVigation (NAV) message of the next satellite to be used when the receiver changes the satel lite selections. It also uses the fifth channel in conjunction with each of the other four channels to perform dual frequency measurements as well as differential channel delay measurements. Individual, dedicated tracking channels enable the receivers to maintain accuracy under high dynamics, provide the best anti-jamming (A-J) performance, and have the lowest TTFF. This type of receiver is best suited for high-dynamic vehicles such as fighter aircraft, vehicles requiring low TTFF such as submarines, plus any user requiring good A-J performance.
2.1.2 Sequential Receivers
A sequential GPS receiver tracks the necessary satellites by typically using one or two hardware channels. The set will track one satellite at a time, time tag the measurements and combine them when all four satellite pseudoranges have been measured. These receivers are among the least expensive available, but they cannot operate under high dynamics and have the slowest time-to-first-fix (TTFF) performance.
2.1.2.1 One-Channel Sequential Receivers A 1-channel sequential receiver makes four pseudorange measurements on both the L1 and L2 frequencies in order to determine a position and compensate for ionospheric delay. The NAV message from each of the satellites must also be read to obtain ephemeris data. To determine an initial position, the receiver must perform the following operations, 1) C/A- code search for a SV, 2) C/A-code/carrier center, 3) data bit synchronization, 4) frame synchronization and Z-count, 5) HOW, 6) P-code carrier center, 7) data demodula tion and 8) ionospheric measurements. Once these operations are complete for one SV, the receiver must perform them again for three other SVs. The four pseudorange measurements must be propagated to the same reference time before a navigation solution is generated. Any movement of the Host Vehicle (HV) during the time the receiver collects the four pseudoranges will reduce the accuracy of the position, velocity, and time
2-1

calculations in the receiver. One-channel sequential receivers are limited to lowdynamic or stationary applications.
2.1.2.2 Two-Channel Sequential Receivers
Two-channel sequential receivers have been developed for use on medium-dynamic vehicles such as helicopters. During initial power-up each channel operates like a 1channel sequential receiver. After four SVs have been acquired, one channel is dedicated to navigation (pseudo range measurements, carrier tracking, etc.) while the other channel reads the NAV message from each satellite. Both channels are also used to perform dual frequency measurements to compensate for ionospheric delay and to measure differential channel delay. Two-channel sequential receivers decrease the time it takes to start navigating by better than one minute when compared to 1channel sequential receivers.
2.1.3 Multiplex (MUX) Receivers
A MUX receiver switches at a fast rate (typically 50 Hz) between the satellites being tracked, continuously collecting sampled data to maintain two to eight signal processing algorithms in software. In addition, the 50 Hz NAV message data is read continuously from all the satellites. In single channel MUX receivers the hardware channel is time shared and only one code generator and one carrier synthesizer is required to track the satellites. However, a multiplex receiver's measured carrier to noise ratio (C/N) for any satellite signal will be 10 log (n) (where n is the number of satellites being tracked) decibels (dB) below that of a continuous tracking receiver. Consequently, for military receivers, the MUX technique has the disadvantage of lower resistance to jamming and interference when compared to continuous tracking receivers. The MUX technique is more commonly found in commercial receivers where the reduced hardware cost can result in a less expensive product and where interference may be less of a concern.
2.2 "ALL-IN-VIEW" RECEIVERS
Traditionally, GPS receivers choose the four satellites of those available that give the best geometry to perform a position fix. However, in situations where one or more of the satellites are temporarily obscured from the antenna's view, the receiver will have to acquire additional satellite signals to generate a continuous PVT solution. The PVT solution degrades until the new satellites are acquired. One solution is to have a receiver which uses all available satellites in view to generate a solution. The inherent advantage of this receiver is that if it is tracking six or seven SVs and a satellite becomes obscured, the receiver will continue to provide a PVT solution with little, if any, degradation. In general, over-determined solutions improve accuracy of the receivers. If the receiver does not dedicate one hardware channel per satellite, then the receiver must use some sort of continual re-acquisition strategy (see MUX receivers paragraph 2.1.3).
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2.3 AUTONOMOUS INTEGRITY MONITORING TECHNIQUES
GPS receivers may track additional satellites for integrity monitoring purposes. This function is independent of receiver architecture. Integrity monitoring receivers derive multiple position solutions by excluding one satellite at a time. Inconsistencies in the results are used to identify and exclude a faulty satellite. In general, at least five satellites must be tracked to detect an integrity failure, and at least six satellites must be tracked to exclude an erroneous satellite. Other measurements, such as altitude or time, may be substituted for satellites in the integrity algorithms, much in the same manner as these measurements are substituted in the PVT solution. In doing so, the integrity of the aiding sources is checked as well. The integrity monitoring algorithms are commonly referred to as Fault Detection and Exclusion (FDE) algorithms or as Receiver Autonomous Integrity Monitoring (RAIM or AIM) algorithms. These algorithms are typically executed on each new set of measurements, thus protecting the integrity of each PVT data set output by the receiver. For additional discussion of integrity, refer to Chapter 12.
2.4 TIME TRANSFER RECEIVERS
One of the more common uses of GPS is for precise time dissemination applications. Several manufacturers offer this type of equipment commercially. These precise time GPS receivers need only one GPS satellite for precise time dissemination if the receiver is stationary on a precisely known location and the only "unknown" is its own clock offset from GPS time and therefore from UTC. To obtain the necessary precise position, the receiver either receives it as an operator input or uses four satellites to determine its own position. These receivers typically include an internal oscillator or an optional external frequency source (rubidium or cesium). Whenever the receiver is tracking a satellite, it generates 1, 5, or 10 MHz reference frequencies that are synchronized to UTC time. If no satellites are visible, the reference frequencies are derived from the internal or external frequency source. The receivers can provide either stand-alone (uncoordinated) or coordinated time-transfer operations. In SPS receivers, use of SA will reduce the time and position accuracy available. The manufacturers of time transfer receivers claim time accuracies in the 20 to 50 nanoseconds range, but this accuracy requires algorithms that average pseudorange measurements over time (10 - 60 minutes). A stand-alone PPS time receiver normally provides time accuracy in the 100 nanoseconds range. The advantage of having an external frequency source interface designed into the receiver is that the long term error in the frequency source can be adjusted when the receiver has satellites in view. A stationary PPS GPS receiver with a precise time and time interval (PTTI) interface should be able to provide UTC to an accuracy of 50 to 60 nanoseconds.
2.5 DIFFERENTIAL GPS (DGPS) RECEIVERS
DGPS receivers are used in applications where enhanced accuracy of the PVT solution is required or desired. DGPS is based on the principle that receivers in the same vicinity will see similar errors on a particular satellite ranging signal. In general, the DGPS technique
2-3

uses measurements from a reference receiver established at a known location, along with differencing algorithms, to remove common satellite and signal propaga tion errors from the PVT solutions of other (mobile) receivers operating in the vicinity of the reference station. The residual errors that remain uncorrected are due to multipath and noise in the receivers. DGPS techniques can be applied to the real-time PVT solution or to recorded measurement data. Real-time DGPS requires a data link pass the reference measurements to the mobile receiver(s). DGPS techniques can be applied to nondifferential receivers if the raw measurement data and navigation message are accessible. There are two primary variations of the differential techniques, one-based on ranging-code measurements and the other based on carrier-phase measurements.
Ranging-code DGPS (RCD) techniques can be applied to receivers with any of the tracking architectures described in the previous paragraphs. For RCD, measurements from the reference receiver are used at the receiver site to calculate corrections, which are then broadcast to the mobile receivers. The mobile receivers incorporate the corrections into their PVT solution, thereby removing the common errors and improving accuracy.
The reference receiver can develop corrections for the position solution or individual satellite ranging signals. If the corrections are provided for the position solution, the correction is simply the difference between the measured PVT solution and the "true" solution consisting of the surveyed location, zero velocity, and precise or smoothed time. However in this case, the reference and user receivers must either use the same satellites to calculate the same solution, or PVT corrections for each possible combination of satellites must be broadcast. It is usually more efficient and flexible to broadcast corrections based on individual satellite ranging errors, thereby allowing the user receiver to select the corrections that are applicable to the particular set of satellites that it is tracking. Real-time RCD is capable of producing accuracies on the order of 1 metre. Carrier-phase DGPS (CPD) systems essentially calculate the difference between the reference location and the user location using the difference between the carrier phases measured at the reference receiver and the user receiver. In realtime systems, carrier-phase data from the reference receiver is broadcast to the mobile receivers. The mobile receivers use double-differencing techniques to remove the satellite and receiver clock biases, then use the phase differences to determine the position of the mobile receiver with respect to the reference receiver location. Determining the initial phase offset (cycles plus fractional phase) between the reference station and the mobile receiver has traditionally been a process that required several minutes. Therefore, it is important to maintain phase-lock on the carrier signals to maintain a continuous flow of position data and avoid reinitialization. Consequently, CPD systems have traditionally used continuous tracking receivers. Receivers which gather measurements from more than four satellites are common since they add robustness in the event of loss-of-lock on one satellite and since additional satellites can reduce initialization time. The CPD techniques were originally developed for surveying applications where real-time data was not essential. However, near-real-time and real-ti me techniques are under development with the goal of supporting applications such as precisionapproach for
2-4

aircraft, as well as the original survey applications. Near-real-time and real-time range implementations can achieve centimeter accuracies (fractions of a carrier wavelength) and post-processing surveying techniques can achieve millimeter range accuracies. Surveying receivers are described in more detail in paragraph 2.6.
The accuracy of differential corrections developed at a single site will degrade with distance from the site due to increasing difference between the reference and mobile receiver ephemeris, ionospheric, and tropo spheric errors. Such systems are usually called local area differential GPS systems (LADGPS). The accuracy of the corrections can be extended over a larger area by using a network of reference receivers to develop the corrections, and by modifying the correction algorithms in the user receiver. RCD systems which compensate for distance degradations are usually called wide area differential GPS (WADGPS) systems. CPD systems which compensate for distance degradations are usually called very long baseline interferometry (VLBI) systems.
CPD techniques (interferometry) can also be used to determine platform attitude. In this case, the processing can be contained within one receiver using multiple antennas. The distinction is lost between which antenna is the "reference" and which is "mobile," since all are located at fixed positions on the platform and none are located at surveyed positions with respect to the earth. Since the antennas are separated by fixed distances, and since their relationship to the center-of-mass of the platform is known, it is possible to convert the carrier phase differences into angular differences between the antenna locations and the line of sight to a satellite. By using measurements from multiple satellites, or the position of the platform from a GPS position fix, these angular differ ences can then be transformed to represent the attitude of the platform with respect to the local vertical axis.
There are several standard (and numerous proprietary) broadcast protocols, receiver interfaces, data formats, data sets, and sets of algorithms that have been developed for DGPS applications. Consequently DGPS receivers are typically designed with a particular application in mind and may not be suitable for a different application. Similarly, proprietary systems may not be compatible for the same application. Therefore, DGPS requirements should be investigated thoroughly and candidate DGPS receivers or systems should be evaluated for suitability and compatibility.
2.6 SURVEYING RECEIVERS
Formal surveys are typically conducted with one surveying receiver located in a previously surveyed location and a second receiver at the new location to be surveyed. The receiver at the previously surveyed location acts as a DGPS reference receiver and the receiver at the new location acts as a DGPS "mobile" receiver. The "mobile" receiver is usually fixed at the new location for a period of time to collect redundant measurements and further improve the accuracy of the survey by post-processing to remove or reduce residual errors such as receiver measurement noise. The period of time can range from seconds to days depending on the survey accuracy required. Consequently, surveying
2-5

receivers must include considerable data recording capability. They may also include the capability to store additional information about the characteristics of the surveyed site. Many surveying receivers have the capability to do a "self-survey", that is, develop a smoothed or averaged position from non-differential GPS measurements. Nondifferential (absolute) surveys require considerably more time than DGPS surveys to develop the same accuracy. However, the technique can be useful to establish a reference point for subsequent DGPS (relative) surveys at locations where a formal reference point is inconvenient or unavailable. This capability can be especially valuable for tactical military survey applications where the relative location of the surveyed sites is more important than the absolute location or where centimeter accuracy is unnecessary.
Most surveying receivers can also function in some capacity as navigation receivers, thereby providing guidance for the surveyor to previously surveyed sites. Additional software functions may also be provided to support datum transformations, post-processing, and other related survey functions.
The signal processing techniques of GPS surveying receivers can be divided into four categories:
a. Non-differential GPS b. Ranging-Code Differential c. Carrier-Phase Differential (Interferometry) d. Codeless Carrier-Phase Differential
As described above, many surveying receivers have a non-differential GPS mode for navigation and self-surveys. The signal processing techniques and accuracies obtained are similar to other non-differential receivers as described in Chapter 1. Surveying receivers may use RCD to determine an initial survey position that aids the initialization process. The more accurate the initial position, the more quickly initialization can be completed for real-time applications. Even if the final results are post-processed to obtain maximum accuracy, real-time outputs can provide preliminary results that confirm the success of the survey in the field or enable the surveyor to detect and correct problems that may occur. The primary surveying mode of most surveying receivers is CPD. The carrier-phase measurements and algorithms enable centimeter and sub-centimeter accuracies in part due to significantly lower measurement noise when compared to pseudorange measurements. As in non-differential GPS, iono spheric errors can contribute significant errors. However, in surveying applications, dual frequency (P-code) measurements are almost essential to achieve surveying accuracies. Since the Pcode is normally only available as the Y-code, most surveying receivers use a "codeless" technique to perform ionospheric delay meas urements. One technique uses spectral compressors to compress the GPS signals into audio or subaudio bands. A processor is used to extract the frequency and phase of each satellite in view. Another technique is to split the received satellite signal and multiply it by itself to obtain a second harmonic of the carrier phase shift which does not contain the code modulation. Codeless tech niques can also be used to make CPD measurements but the C/A-code naviga tion message must also be read to obtain the satellite ephemeris if real-time outputs are desired.
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2.7 ANALOG/DIGITAL RECEIVERS
The majority of early GPS receiver designs made extensive use of analog signal processing techniques, however, most modern receivers incorporate digital signal processing to replace analog receiver functions wherever possible. The following examples are provided to give a description of the differences between these two design techniques. Figure 2-1 shows a multi channel GPS receiver in which code correlation is performed using analog mixing techniques at the intermediate frequency (IF). Each satellite signal to be tracked requires a separate hardware processing channel which consists of an analog correlator, code translator, IF stage, and base band converter. The bandwidth of the IF stage is designed to accommodate the GPS data rate and maximum carrier doppler-shifted frequency.

PRESELECTOR GAIN

ANALOG

DIGITAL

CORRELATOR

FINAL IF

DOWN CONVERSION

A/D

TRANSLATOR

CHANNEL 1

SIGNAL AND DATA
PROCESSING

CHANNEL 2

OTHER CHANNELS
Figure 2-1. Analog GPS Receiver Architecture
Figure 2-2 illustrates a GPS receiver using a largely digital architecture. Analog signal processing is limited to preselection and gain applied to the GPS signals during down conversion with fixed translation frequencies. The down-converted signals are digitized through sampling and are then ready for further digital processing. The digital signal processor (DSP) functions shown in Figure 2-2 include correlation, code and carrier acquisition, and data recovery.

2-7

In a digital receiver, analog to digital (A/D) conversion takes place at the receiver IF. Code correlation and further signal processing occurs digitally. Since the input signals remain code division multiplexed throughout the front end, this portion of the receiver can accommodate either a sequential or multiplexed tracking configuration for any number of satellites. Thus a digital receiver can easily be structured as an "all in view" receiver, whereas an analog equivalent would require a dedicated hardware correlation channel for each satellite in view. The digital architecture illustrated in Figure 2-2 also provides for a great reduction in complexity of the analog portion of a receiver. This in turn results in lower production costs for test, calibration, and maintenance.

ANALOG

DIGITAL

PRESELECTOR GAIN

DOWN

CONVERSION

A/D

DIGITAL SIGNAL PROCESSOR CHANNEL 1 CHANNEL 2
CHANNEL n

CONTROL
AND DATA PROCESSOR

Figure 2-2. Digital GPS Receiver Architecture
2.8 GPS AS A PSEUDORANGE/DELTA RANGE SENSOR
A GPS receiver need not necessarily be used as a PVT sensor in an integrated navigation system. An integrator may instead wish to use GPS to supply satellite pseudorange and deltarange measurements to an inte grated positioning solution. Additional measurements may be provided to the integrated solution by systems or equipment such as an Inertial Navigation System (INS) (position, velocity, acceleration, attitude), Doppler Navigation System (position, velocity), Inertial Aircraft Heading Reference System (velocity, acceleration, attitude), or a Central Air Data Computer (CADC) (baro-altitude and airspeed). The positioning processor can combine all the measurement data into one Kalman filter, to generate a system positioning solution.
Such a solution requires a sophisticated integration scheme. In order to use the pseudorange and deltarange data, additional information is needed such as satellite ephemeris and GPS receiver clock biases. Accurate system and subsystem clocks are needed to correct for differences between the time the calculations are performed and the
2-8

time that the measurements were taken. Alternatively, measurements can be deweighted in the integrated solution and latency errors added to the system error budget. If implemented correctly, a GPS sensor can still contribute to a navigation solution when less than four satellites are being tracked. Such a system is capable of incorporating a single satellite measurement into the system Kalman filter, thus bounding the navigation system solution in one dimension. The disadvantage of using GPS as a sensor in an integrated positioning solution is the high level of complexity involved in integrating such a system. In general, stand-alone GPS receivers do not allow corrected pseudorange and delta range data out of the receiver since it is classified data. Therefore, some receivers provide the capability to process the integrated solution within the GPS receiver. Many GPS sensors are now small enough to be embedded as a card or module into another system, such as an INS or Flight Management System (FMS). In such cases, the corrected pseudorange and delta range data may be permitted off the card or module since the classified data would be contained within a single unit. However, GPS technology now allows most of the GPS receiver functions to be performed by a single semiconductor chip or small chip set. Consequently, future security/processor devices such as the Selective Availability Anti-Spoofing Module (SAASM) may return to the "stand-alone" architecture, providing the capability to process the integrated solution aboard the device, while not allowing the corrected pseudorange and delta range data out of the device. For additional discussion of GPS integration architectures and related issues, refer to Chapter 8. For security design guidelines refer to Navstar GPS PPS Host Application Equipment (HAE) Implementation Guidelines.
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CHAPTER 3: MINIMUM PERFORMANCE CAPABILITIES OF A GPS RECEIVER
3.1 BASIC CONSIDERATIONS
There are a set of basic performance parameters that are useful for making comparisons between different GPS receivers. This set of parameters, together with others, can be used to determine what type of receiver one should choose for a particular application. The parameters of interest are:
a. Position accuracy b. Velocity accuracy c. Time accuracy d. TTFF
3.1.1 GPS System Accuracy Characteristics
There are a number of different ways in which to express GPS accuracy. All are expressed in statistical terms, with a probability assigned to the value given, and the number of dimensions expressed or implied. The two primary positioning accuracy requirements imposed on the GPS system by the U.S. DoD are 16 metres SEP for PPS, and 100 metres 95% horizontal for SPS. SEP represents a 50% probability. Note that the PPS requirement is a three-dimensional requirement specified at the 50% probability level and the SPS requirement is a two-dimensional requirement specified at the 95% probability level. Despite this inconsistency, these are the parameters and points on the accuracy distributions that the Control Segment has used to determine system management policies and methods.
GPS system positioning accuracy distributions are not spherical and are not Gaussian in the tails of the distributions. Consequently, conversions from the system accuracy requirements to other expressions of GPS accuracy, based on an assumption of a spherical distribution that is Gaussian in each dimension can be inaccurate, especially at the 95% probability level which is commonly used by NATO.
"Technical Characteristics of the Navstar GPS" gives conversions of the PPS positioning requirement for typical GPS system operating conditions as 37 metres 95% spherical accuracy and 21 metres 95% horizontal accuracy. Technical Characteristics of the Navstar GPS also provides 95% accuracy tables to facilitate comparisons of PPS and SPS spherical, horizontal, and vertical accuracies.
GPS exhibits statistical accuracy distributions because two important parameters determine the accuracy of the position solution. They are User Equivalent Range Error (UERE) and Geometric Dilution of Precision (GDOP). Both of these parameters are variable with time. UERE is a measure of the error in the range measurement to each satellite as seen by the receiver. UERE varies because of random variations in the satellite signal, signal propagation characteristics, and user measurement processes. Over the long term (days to months) UERE closely resembles a Gaussian distribution and is equivalent for each satellite. UERE tends
3-1

to be different for each satellite at any instant in time and tends to be at a minimum following a new navigation message upload.
GDOP is an instantaneous measure of the error contributed by the geometric relationship of the satellites as seen by the receiver. GDOP is a dimensionless multiplicative factor. For a given value of UERE, small GDOP values mean more precise position and/or time. GDOP varies because the satellites are in constant motion and their geometric relationships are constantly changing. Consequently GDOP can vary with time and user location. The "average" GDOP tends to induce a circular error distribution in the horizontal plane with the vertical contribution of error approximately 1.5 times the horizontal contribution. In real-time, GDOP can be asymmetrical in the three dimensions and vary significantly from the average or typical case, however, GDOP can be easily measured by the receiver, and is often used to select optimum combinations of satellites for the position solution or to develop real-time accuracy estimates.
GDOP distributions are not Gaussian, particularly in the tails of the distribution. The global distribution of GDOP can vary significantly at the 95% probability level due to temporary "vacancies" in the GPS constellation, while remaining relatively constant at the 50% probability level where the GPS PPS system accuracy requirement (16 metres SEP) is defined. Therefore, PPS 95% accuracy specifications derived from this requirement may not be rigorously maintained through all the possible states of the GPS constellation. However, although small variations in accuracy performance are likely with each change in the constellation state, worst-case situations are worst-case for all users and by all measures of system performance, and will therefore be avoided or quickly corrected by the Control Segment. (Temporary "vacancies" in the satellite constellation can be expected over the life of the system due to preventive maintenance, satellite endof-life failures and delayed replacements, or random satellite failures that are correctable by the Control Segment.)
UERE and GDOP are explained in more detail in paragraphs 3.1.2 and 3.1.3. It should be noted that these errors are constantly present as normal variations in accuracy, even with a complete GPS constellation and correctly operating satellites, Control Segment, and receiver.
3.1.2 GPS PPS System Range-Error Budget
The GPS PPS system range-error budget is presented in Table 3-1. The budget is expressed for the 95% probability level of the system UERE. This is a UERE averaged for all satellites over a 24-hour period. Therefore, the long-term (greater than 24 hours) onesigma UERE for an individual satellite can exceed this value and the system can still meet the accuracy requirements specified in the previous paragraph. The instantaneous UERE of all satellites will typically exceed this value at sometime during a 24 hour period. From the user point of view, the important values in this error budget are those allocated to the User Segment. These are excellent guidelines for the purchase or development of receivers because they are independent of the performance of the Space and Control segments.
3-2

Table 3-1. GPS PPS System Range Error-Budget

Segment

Error Source

Space

Frequency Standard Stability D-Band Delay Variation

Space Vehicle Acceleration Uncertainty Other

Control

Ephemeris Prediction and Model Implementation
Other

User

Ionospheric Delay Compensation

Tropospheric Delay Compensation Receiver Noise and Resolution

Multipath Other

Total (RSS) System UERE (metres, 95%)

UERE Contribution (metres, 95%)

P-Code

C/A-Code

6.5

6.5

1.0

1.0

2.0

2.0

1.0

1.0

8.2

8.2

1.8

1.8

4.5

9.8-19.6

3.9

3.9

2.9

2.9

2.4

2.4

1.0

1.0

13.0

15.7-23.1

3.1.2.1 GPS UE Range-Error Budget
The portion of the UERE allocated to the Space and Control segments is called the user range error (URE) and is defined at the phase center of the satellite antenna. The portion of the UERE allocated to the user equipment is called the UE error (UEE). Specifically, the UERE is the root-sum-square of the URE and UEE. The UEE includes residual errors after compensation for atmospheric delay, inherent receiver errors of noise and resolution, and multipath. Modern C/A-code receivers have demonstrated significant improvements in ionospheric delay compensation over the budgeted values. The values given for ionospheric delay compensation error are based on dual-frequency delay measurements for P-code and the singlefrequency ionospheric delay model for C/A-code (as specified in "Technical Characteristics of the Navstar GPS"). The budgeted values for C/A-code can be improved by use of a modified single-frequency model or code less dual-frequency measurements on the L1 and L2 carriers. Modern P-code and C/A-code receivers have both demonstrated significant improvements over the budgeted values for receiver noise, resolution, and multipath, using digital phase locking techniques and variable or narrow code correlation techniques.

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3.1.3 Geometric Dilution of Precision As described in paragraph 3.1.1, GDOP is a dimensionless multiplicative factor that is an instantaneous measure of the error in the positioning solution, contributed by the geometric relationships of the GPS satellites, as seen by the receiver. As an example, if two lines of position are necessary to establish a user position, the least amount of error is present when the lines cross at right angles. The greatest error is present as the lines approach parallel. (See Figure 3-1.) Similarly, for GPS, the greatest amount of error is present when the lines-of-sight between the user and 2 or more satellites approach parallel, or when all four satellites approach the same plane.
Figure 3-1. Dilution of Precision "Technical Characteristics of the Navstar GPS" contains the mathematical definition and derivation of GDOP. In short, if the one-sigma pseudorange measurement errors for all satellites are assumed to be unity, GDOP is defined to be the square root of the sum of the variances of the position and time error estimates. GDOP = (sx2 + sy2 + sz2 + c2st 2)1/2 (Where "c" is the speed of light and "t" is the user clock bias.) GDOP is therefore considered to relate the standard deviation of the satellite range errors (UERE) to the standard deviation of the position solution errors. GDOP is normally considered to be unitless; the units (metres) being carried by the range error and position solution errors. Expressed as a mathematical formula: sUERE x GDOP = sPOSITION SOLUTION ERROR
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Other dilution of precision factors can be defined which are a subset of GDOP and have a more specific physical meaning with respect to the x, y, and z axes in a local coordinate system. They include position dilution of precision (PDOP), horizontal dilution of precision (HDOP), vertical dilution of precision (VDOP) and time dilution of precision (TDOP). Mathematically they are defined as follows:
PDOP = (sx2 + sy2 + sz2)1/2
HDOP = (sx2 + sy2)1/2
VDOP = (sz2)1/2
TDOP = (st2)1/2
HDOP can be further resolved into its X and Y components. If the X axis is oriented in an East-West direction, an "East" DOP (EDOP) and "North" DOP (NDOP) can be defined as follows:
EDOP = (sx2)1/2
NDOP = (sy2)1/2
Similarly, if the Y axis is oriented along the track of a moving vehicle, a "cross-track" DOP (XDOP) and an "along-track" DOP (ADOP) can be defined:
XDOP = (sx2)1/2
ADOP = (sy2)1/2
The various elements of GDOP can also be calculated for an over-determined position solution, that is, where the available satellite or aiding measurements exceed the required minimum of four, and an "all-in-view" solution is calculated. The mathematical formulations are similar, and generally result in a lower value of GDOP (hence better solution accuracy) for each additional measurement that is added to the calculation.
GDOP can also be "weighted" with a vector of UERE values in the matrix calculations for real-time or short-term error estimates where the satellite (or aiding) UERE values are not equal. As mentioned previously, this is generally the case for instantaneous values of UERE, and especially true for SPS where large differences in instantaneous UERE can be caused by Selective Availability. This is also true for aiding situations where the equivalent "UERE" of the aid is usually different than the typical satellite UERE. This "weighted" variation of DOP is an estimate of User Navigation Error (UNE) and is sometimes termed "KDOP". KGDOP has the same definition as GDOP except that the statistical satellite range errors are not required to be equal. Similarly there are analogous subset definitions of KPDOP, KHDOP, etc.
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Which DOP value may be most relevant to a particular application is dependent on the mission and associated accuracy requirements of that mission. (K)HDOP may be most important for land and open ocean navigation where horizontal position location and rendezvous are primary mission requirements. (K)XDOP and (K)ADOP may be most important for air navigation where aircraft spacing is a primary safety consideration. (K)PDOP may be most important for aircraft weapons delivery, and (K)TDOP is obviously most important for time transfer applications. Note that the DOP values discussed here are instantaneous estimates of the geometric contribution to error for a particular location and time. System accuracy requirements often require estimates of long-term error distributions.
For long-term error estimates, the relationship between range error and position solution error should be determined by computer simulation. The standard deviation of the long-term position error distribution can be determined by using the standard deviation of GDOP and the standard deviation of UERE, but the relationship does not hold true for other probability levels, because the tails of the GDOP and position solution distributions are not Gaussian. The most effective method for determining long-term error distributions, for a particular constellation state or set of states, is by conducting a computer simulation.
Computer simulations can be performed to determine global, regional, or single location distributions, but they are often complex and time consuming. If a Monte Carlo simulation is performed assuming a one-metre standard deviation for UERE, the resultant normalized position error (NPE) distribution can be scaled by any UERE of interest, and examined at any probability level of interest. The simulation can iterate user locations around the globe and satellite orbital locations over time (24 hours) while simulating GPS receiver calculations to determine the NPE distribution. While NPE is analogous to GDOP in that it is a measure of the geometric characteristics of error, GDOP is an instantaneous measure and NPE is a statis tical measure. The 95% PPS and SPS accuracy values given in paragraph 5.1 of "Technical Characteristics of the Navstar GPS", were determined by an NPE simulation using an optimized 21 satellite constellation as a surrogate for the average or typical state of the GPS constellation.
It should be emphasized that it may be perfectly valid to translate user accuracy requirements between different dimensions and probability levels assuming a spherical error distribution and Gaussian error characteristics, if that is appropriate for the particular mission or application. The fact that GPS accuracy performance is nonspherical and non-Gaussian does not impose a similar condition on user requirements.
3.2 RECEIVER POSITION ACCURACY
As described in paragraph 3.1.2.1, the UEE is independent of the satellite and Control Segment errors, URE and receiver position accuracy are not. Therefore, receiver position accuracy must be specified for conditions of DOP and URE, in order to isolate the receiver contribution to position accuracy (e.g., UEE, filtering algorithms, and coordinate trans formations). Dynamic positioning accuracy requirements must take into account the effect of vehicle motion on the filter accuracy as well. Laboratory testing must control DOP and URE. Field testing must record DOP and URE. In general, testing is best performed when the system
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positioning accuracies can be achieved. Assuming the UERE error budget is maintained, this generally means DOP conditions of PDOP < 6, HDOP < 4, VDOP < 4.5, and TDOP < 2. URE and DOP are best measured during tests by a calibrated reference receiver. Computer programs which use the broadcast almanac to predict periods of favorable DOP can assist field test scheduling. The GPS System Effectiveness Model (SEM) is one such program developed for the GPS JPO and has been distributed to all NATO nations. Other similar programs are commercially available.
3.3 RECEIVER VELOCITY ACCURACY
GPS receivers typically calculate velocity by measuring the frequency shift (Doppler shift) of the GPS D-band carrier(s). Velocity accuracy can be scenario dependent, but 0.2 m/sec per axis (95%) is achievable for PPS receivers. SPS velocity accuracy is the same as PPS when SA is off. When SA is on, SPS velocity accuracy is degraded. The amount of degra dation of the velocity is classified. However, although not guaranteed, SPS velocity accuracies around 0.4 m/sec 95% have been observed by civilian users for the typical level of SA associated with normal peacetime operations and 100 metres 95% horizontal position ing accuracy.
Velocity accuracy can be effectively tested in a laboratory environment, but field testing can be difficult since a tracking system with 0.05 m/sec or better accuracy is required. The reader is urged to carefully consider the methods of testing if velocity accuracy is an important mission requirement.
3.4 RECEIVER TIME ACCURACY
A dedicated PTTI port should normally be used for precise time output from a GPS receiver. Significant time delays and uncertainties from microseconds to milliseconds can be introduced if time output is accomplished via a digital data interface. For a PPS P-code GPS receiver, tracking 4 satellites, an absolute time accuracy of better than 200 nanoseconds (95%) relative to UTC is possible in a stationary or low-dynamic situation at an unsurveyed location. Equivalent SPS C/Acode accuracy is 340 nanoseconds (95%). Higher dynamics will increase time error. Errors in the PTTI output result from errors in the GPS receiver as well as the Control and Space segments. The system time transfer error budget is shown in Table 3-2.
Processing errors in the GPS receiver and unaccounted time delays to propagate the timing pulses to the PTTI port can add another 60-100 nanoseconds (95%), depending on receiver design. Therefore, a total (RSS) time error of 209-224 nanoseconds (95%) can be expected.
Typical 95% time accuracies expected for precise time dissemination for different categories of GPS receivers are shown in Tables 3-3 and 3-4, assuming an RSS of 88 ns for the Control and Space Segment errors, and 78 ns for the PTTI error.
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3.5 TIME-TO-FIRST-FIX
Time-To-First-Fix (TTFF) is a measure of the elapsed time required for a receiver to acquire the satellite signals and navigation data, and calculate the first position solution. TTFF begins when initialization of the receiver is complete (including selftest, loading of PPS keys, and any required operator input) and the receiver is commanded to begin the positioning function. Some source material (U.S. DoD in particular) may refer to TTFF1 and TTFF2. TTFF1 is based on C/A-code acquisition with hand over to P-code tracking. TTFF2 is based on direct P-code acquisition. REAC (reaction time) is the term typically used to include both the initialization process and TTFF. Since initialization may necessitate operator action, REAC specifications or require ments may require assumptions of operator response times. TTFF is a function of the initial receiver state as well as receiver design. The following paragraphs describe the satellite acquisition and initial positioning processes in more detail.
Table 3-2. Time Error Budget

Error Component US Naval Observatory Measurement Component Control Segment Measurement Component GPS Time Predictability Navigation Message Quantization Satellite Orbit Satellite Clock Satellite Group Delay Downlink and User Equipment
Total (RSS) Time Transfer Error Budget

Error (ns, 95%) 137 59 92 6 22 63 12 65 199

Table 3-3. Precise Time Output Accuracy (95%) for a Typical PPS P-code Receiver

Receiver Mode
Stand-Alone, Stationary, or Low Dynamic

Receiver Output Instrumentation Port PTTI Port

S/A On 2 ms 127 ns

S/A Off 2 ms 127 ns

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Table 3-4. Precise Time Output Accuracy (95%) for a Typical SPS C/A-code Receiver

Receiver Mode and Output Stand-Alone (4 SVs), Stationary or Low Dynamic, PTTI Port Stand-Alone (1 SV), Stationary, Known Position, PTTI Port Coordinated Time Transfer, PTTI Port Instrumentation Port

S/A On 274 ns 255 ns 59 ns 2 ms

S/A Off 157 ns 147 ns 20 ns 2 ms

3.5.1 Warm Start, Cold Start, and Hot Start
Three different variations of TTFF are commonly defined and any one or all three can be specified or required for a particular receiver. A warm or normal start is based on the assumption that the receiver has an estimate of current time and position as well as a recent copy of the satellite almanac data. Typically, time should be known within 20 seconds of GPS time, position should be known within 100 kilometers, velocity within 25 metres per second, and the satellite almanac should have been collected within the past few weeks. TTFF1 for warm starts is typically in the 2 to 5.5 minute range.
A cold start occurs whenever there is a problem with these key data elements. This is typical of a receiver as delivered from a manufacturer, supply depot, or repair depot. Date and time will not be maintained if the receiver "keep alive" battery has been removed or drained. If the receiver clock and memory remains active, the last known position might be at a factory or depot thousands of kilometers from the present position, and the almanac may be several months old. Under such conditions, the receiver may have to systematically "search the sky" until it can find a satellite and retrieve time and a current almanac. A cold start can add at least 12.5 minutes to TTFF1 over that based on a warm start.
A hot start occurs when a receiver is provided with a standby feature to maintain oscillator temperature, time, position, and individual satellite ephemerides (as well as the almanac). When the receiver is commanded out of the standby mode, the time required to achieve the next full position fix is usually Termed Time to Subsequent Fix (TTSF) rather than TTFF. Typically, TTSF is on the order of 10 seconds for standby periods of a few hours.
3.5.2 Receiver Warm-Up
When a GPS receiver is initially turned on, time must be allowed for the receiver crystal oscillator to warm up and stabilize at its normal operating temperature. In a GPS receiver it typically takes up to 6 minutes to complete this process. If the receiver is provided with a mode that keeps the oscillator warm, this contribution to TTFF can be avoided.

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3.5.3 Almanac Collection
The first time a receiver is operated, it must perform an iterative search for the first satellite signal unless it can be loaded with a recent satellite constellation almanac, the approximate time and the approximate receiver location. The almanac gives the approximate orbit for each satel lite and is valid for long time periods (up to 180 days). The almanac is used to predict satel lite visibility and estimate the pseudorange to a satellite, thereby narrowing the search window for a ranging code. Once the first satellite signal is acquired, a current almanac can be obtained from the NAV msg. It takes up to 12 1/2 minutes to collect a complete almanac after initial acquisition. An almanac can be obtained from any GPS satellite. Most modern receivers can update the almanac periodically and store the most recent almanac and receiver position in protected memory. A clock can also be kept operating when the receiver is off or in standby mode, so as to minimize initial acquisition time for the next start-up.
3.5.4 Initial Uncertainties
The initial uncertainties associated with a GPS receivers initial position, velocity, acceleration, jerk and time inputs must be specified when satellite acquisition times are being tested. Acquisition and reacquisition times will vary depending on the accuracy of the receiver initial ization. Some military TTFF requirements that include jamming and other sensitive para meters in the start-up scenario may be classified.
3.5.5 Ephemerides Collection
Ephemeris data forms part of the 50 Hz NAV msg transmitted from the GPS satellites. Unlike almanac data which can be obtained for the whole constellation from a single satellite, ephemeris must be collected from each satellite being tracked on acquisition and at least once every hour. Ephemeris information is normally valid for 4 hours from the time of transmission, and a receiver can normally store up to 8 sets of ephemeris data in its memory. Acquisition and reacquisition times for a receiver will vary, depending on whether valid ephemeris data is already available to the receiver. When testing acquisition time it is necessary to specify whether a valid set of ephemerides is resident or not within the receiver. Depending on the NAV msg collection scheme employed in a particular receiver, it can take between 30 seconds and 3 minutes to collect the ephemeris information.
3.5.6 Enhanced Acquisition Techniques
A number of enhanced acquisition techniques have been developed for modern receivers. TTFF performance can be significantly improved by the use of multi-tap correlators and multi-channel search algorithms. Multi-tap correlators are essentially multiple correlators in the same package which greatly enlarge each search window for code correlation. Similarly, using all available receiver channels in the search for the first satellite can reduce TTFF by maximizing the effective search window of the receiver.
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3.5.7 Direct P(Y)-Code Acquisition Direct P(Y)-code acquisition can be effectively achieved using enhanced acquisition techniques to enlarge the search window and/or by us ing atomic clock aiding to reduce the initial time uncer tainty. Similarly, aiding from an inertial reference system can be used to reduce the initial velocity and position uncertainty. Downloading initialization data from another receiver can be used for direct P(Y)code acquisition as well. 3.5.8 TTFF Requirements Figure 3-2 is a decision chart for determining TTFF requirements for the various initial conditions described above, as well as the TTFF1 and TTFF2 acquisition strategies and different receiver designs 3.5.9 Satellite Reacquisition Satellite reacquisition assumes a temporary loss of a satellite signal due to masking or similar loss of satellite visibility. A satellite reacquisition time of 10 seconds or less is typically achievable. As described in paragraph 3.5.3, the accuracy of the receiver position estimate is a primary factor in determining satellite reacquisition time. Vehicle dynamics and elapsed time from loss of the signal are therefore important in determining the accuracy of the receiver position estimate, as is the presence of GPS aids such as an INS. Laboratory testing is recommended since these factors are difficult to control and predict in the field.
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START

YES ∆P <10km ∆V ≈ 0 ∆A ≈ 0 ∆J ≈ 2m/s3 ∆T <10µs
DIRECT ‘P’ CODE ACQUISITION
ADD 30 SECS
(TTFF2)

COLD START

YES

ADD 6 MINS FOR

CLOCK WARM-UP

NO

ALMANAC CLEARED
NO

YES

ADD 15 MINS FOR

TO COLLECT

ALMANAC

P.V.T.

UNCERTAINTIES

NO

VERY

SMALL

YES

UNCERTAINTIES SMALL ENOUGH TO ALLOW DIRECT P-CODE ACQUISITION
OR ATOMIC AIDING

PVT UNCERTAINTIES LARGE ADD 30 SECS
ALMANAC
∆P <100km ∆V < 75m/s ∆A <10m/s2 ∆T <20s

NO
CA TO ‘P’ ACQUISITION ADD 60 SECS
(TTFF1)
CURRENT EPHEMERIS
IN SET
YES
FINISH

COLLECT EPHEMERIS ADDS
3 MINS FOR 1 CHAN SET 2 MINS FOR 2 CHAN 30 SECS FOR 5 CHAN 1 MIN FOR MULTIPLEX

Figure 3-2. Time-To-First-Fix (TTFF)

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CHAPTER 4: GPS RECEIVER INTERFACES AND ANCILLARY EQUIPMENT
4.1 INTRODUCTION
GPS receivers often require electrical interfaces with other components of the GPS receiver system or with other systems in a host vehicle (HV). Virtually all vehicle integrations will require interfaces with HV power and an external antenna. Many will require a crypto key interface and control-and-display interfaces between an equipment compartment and a crew compartment. Some will require interfaces between a data loader and the GPS receiver. Others may require interfaces between the GPS receiver and other navigation systems in order to develop an integrated position solution. In order to accommodate the varied requirements of different installations, a GPS receiver may be built with a variety of interfaces to aid integration. This chapter presents some thoughts on the ways of integrating GPS with other systems using the interfaces specified for many of the U.S. DoD receivers. These interfaces are also used by other NATO Nations and are provided by other manufacturers, and therefore give an indication of what type of interfaces could be available in a military GPS receiver. Examples of U.S. DoD ancillary equipment are also provided to clarify interface uses.
4.2 GENERAL PURPOSE INTERFACES
Two of the most used interfaces in a vehicle integration are the MIL-STD-1553 multiplex data bus and the ARINC 429 digital information transfer system. Both interfaces can be used to interconnect a GPS receiver with a wide variety of other equipment, for example, a control-and-display unit (CDU), data loader, flight instrument interface unit, or other navigation system such as an INS.
4.2.1 MIL-STD-1553 Multiplex Data Bus
Some GPS receivers are designed to communicate with other equipment via a MILSTD-1553 interface. The MIL-STD-1553 data bus is commonly used aboard military aircraft and can also be found aboard military ground vehicles, ships, and missiles. It is seldom used for civilian applications. The MIL-STD-1553 bus operates with one of the interconnected equipment units assigned as a bus controller. The bus controller controls the data flow on the bus in an asynchronous command/response mode, and also transmits and receives information. The other units are connected to the bus function as "slaved" remote terminals that receive and transmit information, but may also function as back-up bus controllers. The bus controller software program is specifically designed for each unique installation.
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4.2.2 ARINC 429 Digital Information Transfer System
The ARINC 429 data link is commonly used in commercial as well as military aircraft. It is a single-point to multi-point asynchronous half-duplex data link. That is, an equipment can transmit data to several other pieces of equipment. Each link is programmed to output specific data formats at specific data rates. The ARINC 429 specification defines standard data formats and rates for data transfer between a wide variety of commercial avionics equipment. However, the GPS data formats were designed for a commercial ARINC 743A GPS/ GLONASS receiver.
4.2.3 Uses of the MIL-STD-1553 and ARINC 429 Interfaces
The following paragraphs give several examples of ancillary equipment that might communicate with a GPS receiver over the MIL-STD-1553 or ARINC 429 data links.
4.2.3.1 Control and Display Unit
A CDU is often required when a GPS receiver in an equipment compartment must be controlled remotely from a crew compartment. The CDU allows the operator to enter initialization data and control parameters, display status and position data, and can provide access to related functions, such as, waypoint navigation functions. Examples of two types of CDUs procured by the U.S. DoD are discussed below to clarify typical CDU capabilities.
The U.S. DoD has procured dedicated CDUs as well as multifunction CDUs. A dedicated CDU (or "dumb" CDU) is essentially a remote control and display panel that possesses no processing capability, relying on the GPS receiver for all computation functions. A multifunction CDU (or "smart" CDU) is designed to control a GPS receiver, perform other navigation or control functions, and may interface with additional navigation equipment as well. The multifunction CDU includes onboard processing capability for functions, such as, calculating a composite positioning solution using GPS and other navigation sensors, or performing the waypoint navigation function.
4.2.3.1.1 Dedicated CDU
A view of the front panel of a dedicated CDU with a sample display is shown in Figure 4-1. The CDU has a four line, 13 character display controlled by two rotary switches, four line select keys, a display freeze key (Mark), a waypoint mode key, a page slew key, and an alphanumeric keypad. The MODE switch selects the receiver operating mode, the DATA switch selects which parameters are to be displayed, and the keyboard is used to make parameter entries.
In addition to the basic position, velocity, and time displays, the CDU also provides status information on various display pages. Some of this information is the external interface configuration, satellite tracking status, estimated position error, age of satellite almanac, PTTI 1 pulse per second time difference, and the Built-In-Test (BIT) fault log data. Control functions
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Figure 4-1. Example of a Dedicated CDU
include the selection of the lever arm source, flight instrument interface mode, and aiding sensor control.
4.2.3.1.2 Multiple Dedicated CDU Operation
Control and display for a GPS receiver may involve more than one dedicated CDU. The design of a GPS receiver may incorporate two dedicated CDU interfaces and may also provide a data link interface (e.g., MIL-STD-1553) that can also be utilized for control and display. However, only one interface should be able to control and manually initialize a GPS receiver at any given time. A master CDU can be designated by a software configuration connector strap as either a data bus or one of the dedicated CDUs. The master CDU is initially the "active" CDU when the receiver is powered up and may always regain control from another CDU if it has relinquished control to that unit. A designator indicating the current active CDU should be stored in non-volatile memory so that it will not change as a result of a accidental power outage. The active CDU has the sole responsibility for control and manual initialization and thus has sole responsibility for the following:
· Receiver Mode Commands · Rendezvous Mode Selection
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· Waypoint Activation · Destination Selection · Waypoint Definition · Mark Definition · Desired Track/Desired Vertical Angle Selection · Altitude Hold Activation · Stationary Mode Activation · Flight Instrument Scaling · Map Datum Selection · Acquisition Uncertainty Selection 4.2.3.1.3 Multifunction CDU A view of the front panel of a multifunction CDU, with a sample display, is shown in Figure 4-2. The CDU has an eight line, 22 character display controlled by standard and special function keys, full alphanumeric keypad, and eight line select keys. The CDU utilizes a menu driven approach for control, display, and data entry in lieu of the rotary switches of the dedicated CDU.
Figure 4-2. Example of a Multifunction CDU The CDU includes enhanced area navigation software and a dual-redundant MIL-STD1553 data bus. It is capable of operating as either a bus controller, backup bus controller, or remote terminal. The CDU can act as the MIL-STD-1553 bus controller and exchange data with the following equipment:
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· GPS Receiver · Attitude Heading Reference System (AHRS) · Central Air Data Computer (CADC) · Mission Data Loader (MDL) · Two Additional CDU Systems
4.2.3.1.4 Multiple Multifunction CDU Operation
The CDU MIL-STD-1553 bus logic can be designed to support an installation of two or more CDUs. In multiple CDU operations, one CDU is the bus controller and the other(s) are remote terminals and backup bus controllers. If the active bus controller fails, then another CDU becomes the bus controller and no degradation in system performance occurs. The CDU can be designed such that in multi-CDU installations, any CDU can become the "active" CDU and all can have independent control of data display.
4.2.3.2 Data Loader System
A GPS receiver (and/or multifunction CDU) may have the capability to load relevant data over a data link from a Data Loader System (DLS). The primary func tion of the interface is to provide the ability to input initialization data from an external nonvolatile memory device. This is almost essential for GPS avionics systems that must be compatible with civil aviation and use a large International Civil Aviation Organization (ICAO) standard waypoint and navaid data base. A data loader may also be useful for storing navigation, status, or mission data collected during a mission. The DLS may be used to store and load the following:
· Waypoints and Flight Plans · GPS Satellite Almanac Data · GPS Satellite Health/Status Data · Antenna Lever Arm Data · Instrumentation Port Parameters · SA/A-S Data · Sensor Configuration Data
An example data loader system is shown in Figure 4-3. The system consists of a memory device and a read/write/interface unit. The example memory device is a plug-in cartridge that contains solid state memory, memory addressing circuitry, serial input/output converters, and an alkaline cell to power the memory for data retention purposes. Other memory devices such as magnetic tape cassettes and computer diskettes are also common. The read/write/interface unit is installed in the HV and often resembles a small tape deck in size and appearance. It contains the appropriate circuitry to read from and write to the memory device, and contains interface circuitry to send and receive data from the data link (e.g., MIL-STD-1553 or ARINC 429).
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Figure 4-3. Example of a Data Loader System
4.2.3.3 Flight Instrument Interface Unit
Some GPS receiver designs will pass analog signals direct ly to the flight instruments, but many designs may have a digital-only output via an ARINC 429 interface. The reason for a digital-only design is the anticipation of all-digital flight instruments in the future. Aircraft with analog flight instru ments may require a separate digital-to-analog converter to convert the digital data to the synchro, analog and discrete signals needed to drive these instruments.
As an example, the Signal Data Converter (SDC) unit, developed for the U.S. DoD, performs this function. In concept, the SDC process is simple; the SDC takes the digital ARINC 429 data stream and converts those parameters to analog signals that can be handled by analog flight instruments. Not all of the parameters can be used (e.g. waypoint, latitude, and longitude) since the analog flight instruments have no way of processing or displaying such data. Data which can be used by analog flight instruments include:
· Distance to Waypoint · Waypoint Bearing · Desired Track (or radial) · Vertical/Horizontal Deviation From Selected Track (2-D or 3-D) · Data Validity Discretes · To/From Indication
The use of GPS for navigation in a mili tary aircraft is often seen as a substitute for the Tactical Air Navigation (TACAN) system. Therefore, it may be desirable to use TACAN procedures with GPS, and it may also be desirable for the GPS displays to emulate the TACAN displays. The SDC includes the capability to function as a TACAN digital-to-analog converter by means of a simple discrete switch. This provides a simplified method
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for GPS access to the analog flight instruments, using the existing TACAN wiring path (i.e., replace the existing TACAN D-to-A with the SDC).
Since GPS is still a relatively new system, some of the TACAN system characteristics need to be considered. Identified below are GPS flight instrument display and procedures comparisons to TACAN and other radio navigation aids.
4.2.3.3.1 Deviation Scale Factor
With TACAN, a 2-dot horizontal deviation displacement represents 10 degrees off the required radial. An Instrument Landing System (ILS) Localizer has a 2-dot displacement of approximately 3 degrees (runway dependent). In the case of the U.S. DoD equipment, the GPS 2-dot displacement represents either 4 nmi, 1 nmi, 0.3 nmi linear displacement, or 3 degrees depending on the scale factor selected (Enroute, Terminal, Nonprecision Approach, or Approach respectively).
These GPS horizontal scale factors were generally derived from airway track keeping requirements for the various phases of flight. The Enroute scale factor was derived from the typical ±4 nmi U.S. National Air Space (NAS) Airway width. The Terminal scale factor was selected based on U.S. Air Force Instrument Flight Center flight testing. The Non-Precision Approach scale factor corresponds with U.S. FAA nonprecision approach tolerance. The Approach scale factor simulates an ILS localizer display.
If 3-dimensional waypoints are used, then the U.S. DoD GPS receiver can present vertical deviation information. The vertical 2-dot deflections are 1000 ft, 500 ft and 200 ft linear displacement, and 0.7 degrees corresponding to the En Route, Terminal, NonPrecision Approach, and Approach scale factors respectively. The linear scale factors provide the opportunity to someday utilize GPS for vertical navigation in level flight. The Approach vertical scale factor simulates an ILS glideslope display.
4.2.3.3.2 TACAN and GPS Flight Procedural Differences
In the TO/FROM TACAN Navigation mode, the Omni Bearing Select (OBS) knob on the Horizontal Situation Indicator (HSI) allows the pilot to select the radial (to or from the current waypoint) along which he wishes to fly. As the knob is turned and the radial changes, the horizontal deviation bar swings to show the pilot whether he is left or right of that radial. In the case of TACAN, the OBS knob feeds back to the TACAN Digital-to-Analog Converter (DAC), where the left/right computation is carried out (see Figure 4-4). The deviation bar is driven by angular differences. The U.S. DoD SDC can mimic the TACAN DAC as shown in Figure 4-4.
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Figure 4-4. Flight Instruments and TACAN
In the case of U.S. DoD GPS receivers using GPS TO/FROM Navigation mode, the receiver is programmed with waypoint information which includes desired track. This can be analogous to selected TACAN station (waypoint) and OBS radial setting (desired track). The deviation bar deflection will be a function of linear distance (when not in approach mode) of the aircraft perpendicular to the desired track which was programmed in the receiver (see Figure 4-5). The SDC provides a desired track output synchro signal that can drive the HSI OBS to the appropriate radial setting. The pilot, however, can not turn the OBS knob to select a new GPS desired track (other similar products may choose to incorporate the OBS knob setting). The pilot wishing to change the desired track value must enter it into the CDU. The pilot alternatively can select the Direct-To navigation function to get a direct course to the waypoint.
Pilots generally steer magnetic headings. GPS is an inher ently "true" system. One must therefore be careful that the SDC always has a designated magnetic or true heading source and the GPS receiver has knowledge of local magnetic variation, or assigned magnetic variation (in the case of Navaids used as waypoints).
4.2.3.4 Inertial Navigation Systems
A GPS receiver integrated with an Inertial Navigation Systems (INS) forms a particularly effective navigation system. The GPS receiver can compensate for the long-term drift of an INS and an INS can compensate for the short-term noise and relatively low data rate of a GPS receiver. (Additional discussion of GPS integration architectures is provided in Chapter 8).
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Figure 4-5. Flight Instruments and GPS
4.3 PRECISE TIME AND TIME INTERVAL INTERFACE 4.3.1 Introduction
GPS is becoming recognized as the primary time dissemination system for military and commercial applications. An example of a system which may use time transfer from GPS is the calibration of atomic clocks.
4.3.2 Precise Time Inputs
A time input is used to reduce the uncertainty of the receivers initial time estimate and thus reduce TTFF, or it may be used instead of a satellite in the navigation solution. The precise time input to a GPS receiver is accomplished by using a 1 pulse per second rate representing UTC one-second-rollover and a Binary Code Decimal (BCD) time code describing the pulse per second time from an atomic clock. The pulse input indicates the moment of the time to UTC, and the BCD time code identifies what time it was at the UTC one-second-rollover.
The MIL-STD-1553 PTTI Input Message time transfer mechanism uses the same time rollover pulse input. However, instead of labe ling the time with a BCD time input, the HV supplies a PTTI input message via the MIL-STD-1553 MUX bus to label the time epoch.
4.3.3 Precise Time Outputs
The primary function of these outputs is to calibrate an atomic clock, or to support other systems that require precise time. The outputs are 1 pulse per second or 1 pulse per minute to indicate the one second or one minute rollover of UTC, and a BCD time code that indicates the time at the rollover epoch (Hours, Minutes, Seconds, Day of Year, Time Figure of Merit (TFOM)).
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Another means of precise time transfer from the GPS receiver is to use the 1 pulse per second output in conjunction with the PTTI output message available on the MIL-STD1553 multiplex bus.
4.4 ROLL/PITCH/HEADING/WATER-SPEED ANALOG INPUT INTERFACE
A shipborne receiver should be able to accept analog inputs of the ship's attitude and water speed in coarse and fine synchro for mat. The heading input signal can be used by the receiver to assist in satellite acquisition and tracking, and for relative course calculations. The roll/pitch input signal can be used by the receiver to compensate for antenna motion. The water speed input signal can be used by the receiver to aid in satellite acquisition and tracking, and for relative speed calcu lations.
4.5 INSTRUMENTATION PORT INTERFACE
GPS receivers typically have an interface for testing during development and manufacturing. If the configuration of this interface is documented and controlled, it may be useful for integration purposes. Several U.S. DoD GPS receivers have an instrumentation port interface. This interface can be used for some HV integration applications and for connection of test equipment used by maintenance and test activities. The interface is a full duplex RS-422 serial interface that can be con nected to a Smart Buffer Box for test instrumentation purposes, or to an Intermediate Level Test Set for maintenance purposes.
4.6 RS-232 INTERFACE
RS-232 is a common interface typically used to interface between computer equipment. The PLGR includes a RS-232 2-way serial port. This port provides the capability to control the PLGR remotely, and to transfer data between PLGRs or between a PLGR and a computer. This interface can also be used for reprogramming PLGR operational software.
4.7 BAROMETRIC ALTIMETER INTERFACE
A variety of barometric altimeter devices output digitally-encoded pressure altitude, referenced to the geoid or Mean Sea Level (MSL), with a pressure reference of 1013.2 hectoPascals (formerly millibars). This is the same encoded altitude as is used in Mode C altitude reporting via an air traffic control radar beacon (IFF transponder). Some U.S. DoD GPS receivers have a compatible baro-altimeter input. It is a parallel interface which consists of ten signal leads and one signal return. The seven most significant bits are a Gray Code representation of the barometric alti tude in feet, to the nearest 500 feet. The three least signifi cant bits are a binary code which indicates the 100 foot increment within the 500 foot interval.
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4.8 GPS INTERFACE OPTIONS
4.8.1 Introduction
Choice of interfaces for a GPS receiver are dependent on the system to which a GPS receiver shall be integrated, and are also dependent on the depth of the integration required. Alternative approaches to interfaces can be grouped as follows:
· Implement a new interface in an existing GPS receiver · Redesign of HV systems to accommodate an existing GPS receiver · Development of an interface box to adapt an existing GPS receiver to an
existing HV system.
4.8.2 Implementing a New Interface in an Existing GPS Receiver
Good design of a GPS receiver allows the partitioning of the receiver portion and the interface requirements. Often this can be accom plished by using a separate processor to manage interfaces, thus buffering the performance of the GPS receiver portion from the individual demands of a platform interface. This gives the ability to add new interfaces with minimum impact on the majority of receiver software design. Given the flexibility of the software design, an existing GPS receiver can have a new interface card inserted into a spare card slot, or if an existing interface is not used, then the new interface card can be substituted for it. This choice is constrained by the hardware limitations of wiring, output pin availability, etc.
4.8.3 Redesign of HV Interfaces to Accommodate an Existing GPS Receiver
Redesign of the HV interfaces to accommodate the GPS receiver with its current interface is a possibility; however, it may not be considered practical unless major components of the HV can be changed at the same time. With GPS becoming available as a sensor (rather than an LRU with interfaces), embedded GPS receiver alternatives (e.g., embedded in an INS) should also be considered when systems are being replaced.
4.8.4 Separate Development of an Interface Box
One approach that can have minimal impact on both an existing GPS receiver and HV systems is the design of a separate "box" that performs the interface functions. This "box" would accept existing interface inputs and outputs of a GPS receiver and convert them to the inputs and outputs normally used by the HV systems. This approach still requires the HV system's software to be changed to accept another navigation input, and the issues of space, weight, and power for the new "box" must be addressed. Of importance is the impact on the data senescence caused by the additional time delay necessary for the "box" to convert the data.
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CHAPTER 5: ANTENNA SUBSYSTEMS
5.1 INTRODUCTION
GPS users have different requirements for GPS system performance which demand a variety of antennas and antenna subsystems. There are three basic types of GPS antennas, a passive Fixed Radiation Pattern Antenna (FRPA), a FRPA with an integrated preamplifier, and a Controlled Radiation Pattern Antenna (CRPA). The requirement to drive a long cable run, with its associated signal loss between the antenna and the GPS receiver has resulted in a FRPA with an integrated amplifier. A CRPA is required to reduce the effects of RF interference which would otherwise jam the receiver's operation.
5.2 FRPA
5.2.1 General Characteristics
A FRPA has a fixed antenna radiation pattern which is only affected by the size and shape of the ground plane on which it is installed. As GPS antennas are typically narrow band the radiation pattern does not change over either the L1 or L2 bandwidth although due to the difference between the L1 and L2 wavelength there are significant differences in the radiation at the L1 and L2 frequencies.
Typical specifications for FRPAs include parameters for operating frequencies, impedance, Voltage Standing Wave Ratio (VSWR), radiation pattern, polarization, axial ratio and gain. These specifications impact receiver performance. The size, shape and weight of the FRPAs will vary with the application. A FRPA for an aircraft installation has a different form than a FRPA for a hand-held receiver. A number of FRPAs are discussed below. FRPAs are generally non -repairable units which require no adjustment over their lifetime. Passive FRPAs require no power. All FRPAs can be fitted with an external low noise amplifier should this be needed to overcome losses introduced by a long cable length. The amplifier will probably require a low power DC voltage.
An important parameter when selecting a GPS antenna is the gain. Gain is defined with respect to an isotropic radiator for circular polarization, expressed as dBic, and the sector of the sphere surrounding the antenna over which the gain can be maintained, expressed as the angle from the antenna boresight. (The boresight is the central axis of the antenna usually the direction of maximum gain).
To receive the signals from GPS satellites, which may be at any angle in the upper hemisphere, the gain must not drop below -5 dBic. In the case of an aircraft there is a significant problem of maintaining sufficient gain towards the satellites as the aircraft maneuvers through high angles of pitch and roll. Typically an aircraft's GPS antenna gain falls to -15 dBic below the azimuth plane, although a worst case gain of -20 dBic can be assumed.
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5.2.2 FRPA Types
There are many types of GPS FRPA antenna. The simplest is a resonant monopole approximately 5 cms in length. However, as the monopole has a toroidal radiation pattern and is vertically polarized, it is not optimum for use with the circularly polarized GPS transmissions. Gain is very low, -40 dBic on boresight and peaking to approximately 0 dBic at 70° from boresight depending on the conductivity of the ground plane.
Spiral Helix antennas are useful for several receiver applications where a small antenna is required that is generally unaffected by the presence or absence of a ground plane. The antenna can be configured to be low profile, but is not conformal and is therefore not suitable for fast aircraft. The antenna is less sensitive to the influence of the ground plane than some other FRPAs and is capable of being mounted on non-conducting surfaces, making it suitable for a variety of applications from vehicles to handheld receivers. Typically the gain is better than -4 dBic from boresight to 80°. The antenna's mechanical layout and typical dimensions are shown in Figure 51.
The FRPA Bifilar Helix is designed for hand-held applications and is capable of being integrated into a broad category of ground vehicles in addition to its main application on the Precise Lightweight GPS Receiver (PLGR). The antenna is insensitive to ground plane and installation location. Streamlined outer shell can be added to enable the device to be used in medium dynamic, for instance helicopter applications. It provides a gain of not less than - 3 dBic over 80° angle from boresight. The antenna 's mechanical layout and typical dimensions are shown in Figure 5-2.
To produce a conformal design for aircraft applications where minimal drag is required, a crossed slot or patch antenna can be used. The crossed slot is effectively four monopoles laid out at right angles with a suitable separation above the ground plane. Patches can take many formats. These antennas rely on the aircraft skin acting as a ground plane to achieve the required antenna performance. Gains of +2 dBic are typically achieved on boresight and, although the gain to circularly polarized radiation falls to -5 dBic at 90° (from boresight), the gain is sufficient to allow satellites to be tracked through medium dynamic aircraft maneuvers. Antennas can be made that are sensitive to L1 and L2 GPS frequencies. The mechanical layout and the dimensions of an example antenna are shown in Figure 5-3.
A special derivative of a FRPA crossed monopoles antenna is the FRPA Ground Plane. This special FRPA assembly (see Figure 5-4) is intended for shipborne mast applications where there is no ground plane. The assembly consists of a ground plane/mounting surface for the FRPA plus an environmentally sealed enclosure containing an integrated preamplifier. A derivation of the FRPA Ground Plane is employed for GPS Reference Stations in a differential system. In these applications a special choke ring is added to the antenna to reduce the gain in the direction of likely sources of multipath.
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Figure 5-1. FRPA Spiral Helix Figure 5-2. FRPA Bifilar Helix
Figure 5-3. FRPA Crossed Monopoles 5-3

Figure 5-4. FRPA Ground Plane
5.3 CRPA Equipment
CRPAs have been shown to be the only effective means of protecting GPS receivers against multiple wideband jammers. A CRPA has two components: an Antenna Control Unit (ACU) and an antenna array. Current aircraft CRPA's typically have seven antenna elements in the array with seven associated processing channels. CRPAs under development for missile may use only four or five elements.
The antenna array is composed of antenna elements which may be of any of the above FRPA types. However the vehicle environment significantly limits the choice. In the case of aircraft the array has to be conformal and is therefore usually made up of patch or crossed dipole antennas. The antenna elements are spaced at approximately half wavelength separation, at the shortest operational wavelength. It is essential for optimum operation of the CRPA that all the antenna elements in the array have omnidirectional performance with constant gain characteristics over as large a sector as possible.
The ACU controls the array's radiation pattern by adjusting the gains and phase from each antenna. First generation ACU employed analogue electronics with some digital control. Newer equipment digitizes the receiver signal in a similar manner to that used in a GPS receiver. The ACU contains a series of amplifiers and gain control systems for each channel, a set of weights that make up a beam former and a microprocessor and associated electronics that contains the control algorithm and drive the weights in the beamformer. Each weight is a phase shifter with gain control. The phase shift was initially performed by analogue components but it is now cost effective to employ digital multiplier circuits. The receiver signal is downconverted to near baseband and sampled into inphase and quadrature components. By adjusting the gain and sign of each component a 360 degree range is achieved.
As the GPS signal is below the thermal noise in the transmission bandwidth, any signal detected above the thermal noise level can be considered to be harmful to GPS operation. Initially the array's radiation pattern is set to omnidirectional, by
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adjusting the gain and phase in the ACU. Whenever a jamming signal is detected, the gain and phase of the beamformer is adjusted to form a null in the radiation pattern in the direction of the jammer with the result to cancel the effect of the jammer. A CRPA has one less degree of freedom than the number of elements (N), allowing N-1 independent jamming sources to be cancelled.
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CHAPTER 6: SERVICE COVERAGE, SERVICE AVAILABILITY, AND SERVICE RELIABILITY; SATELLITE SELECTION CRITERIA AND FIGURE OF MERIT DESCRIPTION
6.1 SERVICE COVERAGE, SERVICE AVAILABILITY, AND SERVICE RELIABILITY
This section describes the minimum performance an authorized user can expect to obtain from PPS receiver which is designed and operated in accordance with "Technical Characteristics of the Navstar GPS". Performance is specified in terms of minimum performance standards for each performance parameter. Each standard includes a definition of applicable conditions and constraints. The information provided in this section is derived and extracted from "The Global Positioning System (GPS) SPS Performance Specification", dated November 5, 1993, published by the U.S. DoD. Although the GPS SPS Performance Specification is directed toward SPS users of GPS, the specified performance of the system with respect to service coverage, service availability, and service reliability is the same for PPS users.
The data and associated statements provided in this chapter represent conservative performance expectations, based upon extensive observations of the system. The performance standards are limited to GPS Control Segment and Satellite contributions to the PPS signal-in-space characteristics and their effects on the position solution. The standards do not include enhancements or degradations to this service that might be provided by the UE or local environment. Examples of possible enhancements include altitude aiding, clock aiding, differential corrections, or integrity algorithms. Examples of possible local degradations include multipath, jamming, terrain masking, or receiver errors.
6.1.1 Parameter Definitions
The three parameters defined below are service coverage, service availability, and service reliability. These definitions and the relationships between them are different from traditional definitions of similar parameters. A dependent relationship is defined to exist between these performance parameters. Each successive layer of performance definitions are conditioned on the preceding layers. That is, coverage must be provided before the service may be considered available and it must be available before it can support service reliability requirements.
Service coverage is defined as the percentage of time over a specified interval that a sufficient number of satellites are above a specified mask angle and provide an acceptable position solution geometry at any point on or near the earth.
GPS coverage is viewed somewhat differently than coverage for existing terrestrial positioning systems. Traditionally coverage has been viewed as the surface area or volume in which a system may be operated. Since a terrestrial system's beacons are fixed, coverage does not change as a function of time. Since the GPS concept relies upon the dynamics of a satellite constellation, coverage must take into consideration a time dependency. GPS coverage is by definition intended to be global. GPS coverage is viewed alternatively as the percentage of time over a time interval that a user, anywhere in
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the world and at any time, can see a sufficient number of satellites to generate a position solution. Constraints are placed upon satellite visibility in terms of mask a ngle and geometry, to minimize the possibility of a GPS receiver generating a marginal position solution. Coverage characteristics over any given region vary slightly over time, due primarily to small shifts in satellite orbits.
Since GPS is a space-based system, coverage is defined as a function of each satellite's antenna beamwidth. The GPS satellite antenna's nominal beamwidth is approximately 28 degrees. If a user on the Earth's surface were to view a satellite which is just above the local horizon, the user could elevate from that location to an altitude of approximately 200 kilometers above the Earth's surface before effectively losing that satellite's signal. This condition defines the maximum altitude associated with the term "on or near the Earth."
Service availability is defined as the percentage of time over a specified time interval that a sufficient number of satellites are transmitting a usable ranging signal within view of any point on or near the earth, given that coverage is provided.
Just because a satellite is operational does not mean that it is currently transmitting a usable GPS ranging signal. Satellites will, on occasion, be removed temporarily from service for routine maintenance. As a result, the number of satellites actually transmitting usable ranging signals will vary over time. Service availability is the measure of how GPS coverage deviates from nominal conditions due to the temporary removal of satellites from service. This measurement represents the percentage of time that coverage is provided by those satellites which are transmit ting usable ranging signals to generate a position solution. Variations in service availability are a function of which satellites are removed from service, the length of the service outage, and where on the globe a user is located in relation to any resulting outage patterns.
Service reliability is defined as the percentage of time over a specified time interval that the instantaneous predictable horizontal error is maintained within the normal accuracy distribution at any point on or near the earth, given that coverage is provided and the service is available.
GPS can be used anywhere in the world. A failure in a system with such global coverage may affect a large percentage of the globe. A natural concern about using GPS is whether or not it provides a satisfactory level of service reliability. Service reliability as it is used in a GPS context is somewhat more restrictive than the classical definition, which includes times that the service is available as well as when it is performing within specified tolerances. GPS service reliability is viewed as a measure only of how well GPS maintains horizontal errors within the normal predictable PPS horizontal accuracy distribution. 100% service reliability is provided when the horizontal error remains within the normal accuracy distribution within the conditions specified for coverage and service availability. Periods where the service does not provide a sufficient number of satellites or adequate geometry to support position solution generation are assessed against the coverage service availability performance standard. 6.1.2 Service Coverage Characteristics
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This section defines the GPS coverage standards, GPS constellation design objectives, and the characteristics of GPS coverage which are expected with a 24 satellite operational constellation. The user is provided with general information concerning how coverage will vary over time on a global basis, and a worst-case projection of coverage on a regional basis. The data provided in the discussion is based upon a global assessment of grid points spaced equally, approximately 111 kilometers apart, every 30 seconds over a 24 hour period.
6.1.2.1 Service Coverage Standards
GPS Service will be provided in accordance with the coverage standards presented in Table 6-1.
Table 6-1. Service Coverage Standards

Coverage Standard ³99.9% global average
³96.9% at worst-case point

Conditions and Constraints
· Probability of 4 or more satellites in view over any 24 hour interval, averaged over the globe
· 4 satellites must provide PDOP of 6 or less · 5° mask angle with no obscure · Standard is predicated on 24 operational satellites, as the
constellation is defined in the almanac
· Probability of 4 or more satellites in view over any 24 hour interval, for the worst-case point on the globe
· 4 satellites must provide PDOP of 6 or less · 5° mask angle with no obscure · Standard is predicated on 24 operational satellites, as the
constellation is defined in the almanac

6.1.2.2 The GPS 24-Satellite Constellation
The 24 satellite constellation is designed to optimize global coverage over a wide range of operational conditions. Specific constellation design objectives are listed below:
· Provide continuous global coverage with specified geometry and mask angle constraints.
· Minimize coverage sensitivity to expected satellite orbital drift characteristics.
· Mitigate the effects on service availability of removing any one satellite from service.
Several factors affect GPS coverage. These factors must be taken into consideration in the constellation design. The factors are:

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· The difference between the planned orbit and the orbit actually achieved during the launch and orbit insertion process,
· Orbit variation dynamics, and · Frequency and efficiency of satellite station-keeping maneuvers. 6.1.2.3 Expected Service Coverage Characteristics Proper support of the first design objective (from above) requires that at least four satellites are continuously in view with an acceptable geometry and mask angle anywhere in the world. An impli cation of this requirement is that most of the time significantly more than four satellites will be visible. As shown in Figure 6-1, eight satellites will be visible on average for any location in the world, over 24 hours. Very seldom will a user see only four satellites when all 24 satellites are providing usable ranging signals. If the 24 satellites in the GPS constellation were all launched with no deviations into their planned orbits, and no drift were allowed, the constella tion would provide virtually 100% (0.99999714) four satellite coverage with a PDOP constraint of six.
Figure 6-1. Satellite Global Visibility Profile Unfortunately, variations in final orbits based upon launch uncertainties and routine drift do occur. The second design objective is supported by evaluating how changes in each satellite's orbital elements affect nominal coverage characteristics. Bounds are applied to orbital element deviation from the nominal orbit to ensure that constellation coverage does not degrade beyond allowed limits. Degraded coverage areas drift and change slightly in shape over time, but their average number and duration will remain approximately constant for a given constellation. Changes in the number of satellites or significant shifts in satellite orbits, however, can dramatically change the attributes of degraded coverage areas.
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Given a 24 satellite constellation, GPS will provide 100% four and five satellite coverage without a PDOP constraint (but with a mask angle of 5 degrees), and six satellite coverage greater than 99.9% of the time. However, four satellite coverage with a PDOP constraint of 6 can drop as low as 99.9%, with a worst-case dispersion of the 24 satellites with respect to their nominal orbits. Even in this event, most users will experience continuous coverage. A few isolated locations may experience foursatellite coverage as low as 96.9%, with a PDOP constraint of 6 and a mask angle of 5 degrees.
Satisfaction of the third design objective requires the ability to remove any individual satellite from the constellation, and still be able to provide as close to continuous global coverage as is practical. Satisfaction of this objective requires that at least five satellites be in view almost continuously. As shown in Figure 6-1, this is the case with the 24 satellite constellation design. Although an explicit requirement is not established to ensure that multiple combinations of satellites provide adequate solution geometry at any given time, most of the time at least two and usually more combinations of four satellites will support a Position Dilution of Precision (PDOP) constraint of 6 or less.
6.1.3 Service Availability Characteristics
This section defines the GPS availability standards and expected regional and global service availability characteristics. The user is provided with information concerning GPS service availability patterns on a global and regional basis. Service availability varies slightly over time, due to routine satellite maintenance requirements. Note that the regional service availability values provided below are based upon a global grid point spacing of approximately 111 x 111 kilometers, with 30 second intervals over 24 hours.
Service availability is described in two basic parts. The first part concerns the variation in service availability as a function of temporarily removing a number and specific combination of satellites from service. The second part of the assessment applies service availability variation characteristics to an operational scenario.
6.1.3.1 Service Availability Standards
GPS service will be provided in accordance with the availability standards specified in Table 6-2.
6.1.3.2 Satellite Outage Effects on Service Availability
Service availability varies predominantly as a function of the number and distribution of satellite service outages. With a 24 satellite constellation, the permutations and combinations of satellite service outages are rather large. Normally, no more than three satellites will be removed from service over any 24 hour interval. This ground rule bounds the problem to an analysis of the effects of removing each satellite and all combinations of two and three satellites from service for no more than 24 hours. The results of the analysis are summarized in Table 6-3.
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Table 6-2. Service Availability Standards

Service Availability Standard
³99.85% global average
³99.16% single point average
³95.87% global average on worst-case day ³83.92% at worst-case point on worst-case day

Conditions and Constraints
· Conditioned on coverage standard · Standard based on a typical 24 hour interval, averaged over the
globe · Typical 24 hour interval defined using averaging period of 30
days
· Conditioned on coverage standard · Standard based on a typical 24 hour interval, for the worst-case
point on the globe · Typical 24 hour interval defined using averaging period of 30
days
· Conditioned on coverage standard · Standard represents a worst-case 24 hour interval,averaged over
the globe
· Conditioned on coverage standard · Standard based on a worst-case 24 hour interval, for the worst-case
point on the globe

6.1.3.3 Expected Service Availability Characteristics
Table 6-3 defines what service availability characteristics will be like for a given satellite outage condition. Service availability projections over time may be generated by applying the information in Table 6-3 to expected satellite control operations scenarios. A satellite control operations scenario is based upon a conservative estimate of satellite maintenance activity frequency and duration. Satellite maintenance actions requiring service downtime include periodic cesium frequency standard maintenance, station keeping maneuvers to maintain orbits within tolerances, and responses to component failures. Given current routine maintenance requirements and component failure expectations, generally three, and no more than four satellites should be removed from service over any 30 day period. Once a satellite is removed from service, it is assumed that it will be down for no more than 24 hours.
The first service availability scenario to be defined represents a worst-case 30 day period. A summary of this scenario is provided in Table 6-4. The scenario is considered to be worst case from two perspectives: it includes a day with three satellites removed from service, and it includes a total of four satellite-down days. The three satellite-down scenario is based upon the simultaneous removal of two satellites for routine maintenance, accompanied with a component failure on a third satellite. Worst case global service availability on a day with three satellites removed from service is 95.87%; the associated worst case regional service availability is 83.92%. The resulting 30-day service availability values range from 99.85% to 99.99%, depending on which satellites make up the four which

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Table 6-3. Service Availability as a Function of Specified Satellite Outage Conditions

Global Average Worst Regional Service

Satellite Temporary Outage Condition Service Availability

Availability

No Satellites Out:

100.00%

100.00%

ONE SATELLITE OUT FOR MAINTENANCE OR REPAIR

Least Impacting Satellite Out:

99.98%

99.17%

Average Satellite Out:

99.93%

97.79%

Most Impacting Satellite Out:

99.83%

97.63%

TWO SATELLITES OUT FOR MAINTENANCE OR REPAIR

Least Impacting 2 Satellites Out:

99.93%

98.21%

Average 2 Satellites Out:

99.64%

95.71%

Most Impacting 2 Satellites Out:

98.85%

91.08%

THREE SATELLITES OUT FOR MAINTENANCE OR REPAIR

Least Impacting 3 Satellites Out:

99.89%

97.13%

Average 3 Satellites Out:

99.03%

93.38%

Most Impacting 3 Satellites Out:

95.87%

83.92%

Table 6-4. Example of 3-Day Global Service Availability with Component Failure on Worst Day

Ops Scenario Condition 1 Day - 3 Satellites Down 1 Day - 1 Satellite Down 28 Days - No Satellites Down
Average Daily Availability

Best Case 99.89% 99.98% 100.00% 99.99%

Average Case 99.03% 99.93% 100.00% 99.97%

Worst Case 95.87% 99.83% 100.00% 99.85%

experience downtime. The service availability service standard was established based upon this scenario, to ensure that the system can support standard compliance.
The second service availability scenario is shown in Table 6-5, and represents what may be considered to be a more common 30 day interval. In this scenario, three satellites were removed from service for up to 24 hours, each on separate days. Typical satellite maintenance operations are conducted on one satellite at a time, which means that the removal of two satellites for maintenance at the same time will be a rare occurrence.

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Global service availability on a day where the worst case satellite is removed from service is 99.85%; the associated worst case regional service availability is 97.63%. The resulting 30-day service availability values do not change much between the best and worst cases, with the worst case value being 99.98%.
Table 6-5. Example of 30-Day Global Service Availability without Component Failure

Ops Scenario Condition 3 Days - 1 Satellite Down 27 Days - No Satellites Down
Average Daily Availability

Best Case 99.98% 100.00% 99.99%

Average Case 99.93% 100.00% 99.99%

Worst Case 99.85% 100.00% 99.98%

6.1.4 Service Reliability Characteristics
This section defines conservative expectations for GPS service reliability performance. These expectations are based upon observed accuracy characteristics, the GPS service failure history to date, long-term failure rate projections, and current system failure response capabilities. The user is provided with information which indicates expected failure rates and their effects on a global and regional basis.
6.1.4.1 Service Reliability Standards
GPS service will be provided in accordance with the reliability standards presented in Table 6-6.

Table 6-6. Service Reliability Standards

Service Reliability

Conditions and Constraints

³99.97% global average

· Conditioned on coverage and service availability standards · Standard based on a measurement interval of one year; average of
daily values over the globe · Standard predicated on a maximum of 18 hours of major service failure
behavior over the sample interval

³99.79% single point average

· Conditioned on coverage and service availability standards · Standard based on a measurement interval of one year; average of
daily values from the worst-case point on the globe · Standard based on a maximum of 18 hours of major service failure
behavior over the sample interval

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6.1.4.2 GPS Service Failure Characteristics
A GPS service failure is defined as an excursion of unpredictable magnitude of the horizontal position solution due to a control segment or satellite fault which is unrelated to the normal predictable long-term PPS horizontal accuracy distribution. A GPS service failure is characterized by a large single-satellite range error which is unrelated to the normal long-term PPS range error distribution.
The characteristics of a service failure and the factors which affect service reliability are listed below. Each is discussed in more detail in the following sections.
· Ranging signal failure frequency · Failure duration · Failure magnitude and behavior · Distribution of user population around the globe · Probability that the failed satellite is used in the position solution · Effect that the failure has on the position solution, given the failed satellite's
contribution to solution geometry and the receiver's response to the failure condition.
6.1.4.3 Failure Frequency Estimate
The GPS satellite positioning service failure history over the past several years indicates a very low service failure rate (excluding Block I satellites). However, when a service failure does occur, it can result in extremely large position and/or velocity errors. This behavior will typically persist until action is taken to remedy the problem.
Based upon an historical assessment of Block II satellite and Control Segment failure characteristics, GPS should experience no more than three major service failures per year (excluding Block I satellites). This failure rate estimate is conservative expectations are on the order of one per year, based upon projected navigation payload component reliabilities and the assumption that action will be taken to switch redundancy configurations if early indications of an imminent failure are detected. An allocation of three per year allows for a possible increase in service failures as the Block II satellites reach the end of their operational life expectancy.
6.1.4.4 Failure Duration Estimate
The duration of a failure is a function of the following factors:
· Control Segment monitor station coverage · Control Segment monitor station, communications and Master Control Station
availability · Master Control Station failure detection efficiency and timeline · Timeline for correcting the problem or terminating the failed satellite's service.
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The combination of these factors results in a conservative system operator response timeline on the order of no more than six hours. In most cases the response to a failure will be much more prompt, but with any complex system such as the Control Segment, allowances must be made for varying system resource status and operational conditions.
6.1.4.5 Failure Magnitude and Behavior
GPS is designed to be fault tolerant - most potential failures are either caught before they manifest themselves, or their effects are compensated for by the system. The only failures to which the system seems susceptible are of two types:
· Insidious, long-term (day or more to become evident) performance deviations, or
· Catastrophic, almost instantaneous failures
Insidious failures do not propagate very quickly - failures of this type experienced to date have not affected the GPS ability to support accuracy performance standards. Insidious failures are typically due to a problem in the ephemeris state estimation process.
Catastrophic failures are due almost exclusively to satellite frequency generation hardware failures. These failures in general result in very rapid ranging error growth range errors can grow to several thousand metres in a very short period of time. Typically, a failure of this type will begin with a phase jump of indeterminate magnitude, followed by a large ramp or increased noise consistent with the behavior of a quartz oscillator.
6.1.4.6 User Global Distribution and Failure Visibility
For the purposes of reliability performance standard definition, the effect of a service failure is not weighted based upon user distribution - a uniform distribution of users over the globe is assumed.
Given a maximum failure duration of six hours, approximately 63% of the Earth's surface will have a failed satellite in view for some portion of the failure. The average amount of time that the failed satellite will be in view for those locations which can see it is approximately three hours.
6.1.4.7 Satellite Use in the Position Solution
Given a 24 satellite constellation, an average of eight satellites will be in view of any user on or near the Earth. The satellite visibility distribution for the nominal 24 satellite constellation is shown in Figure 6-1. With all satellites weighted equally, the probability of a failed satellite being in the position solution of any user located within the failure visibility region is 50%. Equal weighting is considered to be a reasonable assumption for use in global reliability computations. However, in the worst-case individual site computation it must be assumed that the receiver is tracking and using the failed satellite for the duration of the satellite visibility window.
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6.1.4.8 Failure Effect on Position Solution
Given the nature of catastrophic failures, it must be assumed that the inclusion of the satellite in the position solution will induce a service reliability failure independent of the satellite's geometric contribution. Some receivers will be capable of detecting and rejecting large instantaneous changes in a range residual which are indicative of a major service failure. The minimum receiver represented in the Signal Specification is not however, required to have this capability. For the purposes of service reliability standard definition, it must be assumed that if the receiver is capable of tracking the failed satellite and it supports the nominal position solution geometry, the receiver will use it in the position solution.
6.1.4.9 Expected Service Reliability Characteristics
When the system is performing nominally and the receiver design meets the minimum usage conditions established in Section 2.2 of the Signal Specification, predictable horizontal error will never reach the service reliability threshold. Service reliability on those days where GPS does not experience a major service failure will be 100%.
The estimated maximum of three major service failures per year, coupled with a maximum duration of six hours each, yields a maximum of 18 service failure hours per year. The worst-case site on the globe will be the place where all 18 service failure hours are observed and the failed satellites are used in the position solution. For this worst case condition, the daily average service reliability over a one year period will be no worse than 99.79%. The equivalent global daily average will be no worse than 99.97%.
6.1.5 Additional Commentary
(The following commentary is not derived from the GPS SPS Performance Stan dard.) It should be noted that several criteria used as conditions and constraints in the performance standards may not be applicable to many user applications. As examples, the coverage standard is based upon 24 operational satellites, a four-satellite position solution, a PDOP of 6 or less, and a 5 degree mask angle; the service availability standard is based on a "normal" operating scenario; and the service reliability standard is based on the assumption that the user does not perform integrity checking.
6.1.5.1 24 Operational Satellites and Service Availability
The assumption of 24 operational satellites may be optimistic rather than conservative. In the long term, the GPS constellation will be in a continuous cycle of satellite end-oflife failures and corresponding launch of replacements. It is expected that three to four satellites will reach end-of-life each year, based on experience with the Block I satellites and considering design improvements to the Block II satellites. This means that service coverage can change every few months, although end-of-life failures can be anticipated to some degree and some launches can be made prior to the actual failure. A number of studies have been conducted to determine the probability of a specific number of satellites in service at any given time, including some studies conducted for the U.S. DoD to help determine satellite replenishment strategies. One
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such study gives the long-term probabilities for the number of GPS satellites operational any given time.

In most cases, a satellite vacancy from the full constellation of 24 satellites will result in reduced service coverage. For convenience, the lack of a four satellite positioning solution or a condition where PDOP > 6 will be termed an "outage". In general, the number of outages, individual durations of outages, and areas affected by outages will increase with each additional vacancy from the constellation. As long as the U.S. can maintain 21 or more satellites on orbit, and worst-case situations can be avoided, the service coverage is likely to remain between 99% and 100%. Table 6-8 below gives some representative values of service coverage for a 24-Satellite constellation with "typical" deviations from the nominal orbit positions. During the worst-case threesatellite-failure condition, the worst location in the world may experience as low as 86% average positioning availability over a 24 hour period, while the best location may still experience 100% availability.

Table 6-7. Probability of Operational Satellites

Number of Satellites
24 23 22 21 20 19

State Probability
0.72 0.17 0.064 0.026 0.0116 0.0064

Cumulative Probability
0.72 0.89 0.954 0.980 0.9916 0.9980

Table 6-8. Service Coverage of a Typical 24-Satellite Constellation

Number of Satellites
23 22 21 20

Best Global Service Coverage
100.00% 100.00% 99.98% 99.97%

Average Global Service Coverage
99.99% 99.93% 99.69% 99.05%

Worst Global Service Coverage
99.97% 99.61% 97.69% 94.75%

As suggested above, there are several options the U.S. DoD may employ to minimize the impact of reduced service coverage. Suc h options include launches in anticipation of satellite end-of-life failures, planning normal maintenance to minimize service availability impact, deferring normal maintenance, and even minor rephasing of certain satellites in the constellation. In this respect, the standards quoted above for service availability under "normal" operating conditions have some flexibility to compensate for reduced service coverage and still maintain a high composite availability of a position solution.
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6.1.5.2 PDOP Less Than Six
PPS users are much less sensitive to large values of DOP than SPS users. Many PPS users will have sufficient position accuracy using GPS as a stand-alone system even if PDOP is greater than six. For example, for navigation missions, horizontal position accuracy is usually a more appropriate measure than PDOP. As a general rule of thumb, a PDOP of six is typically equivalent to an HDOP of four (although PDOP obviously contains a vertical component which can vary). This means that an approximate worst case PPS error for "normal" horizontal variations would be around 160 metres (assuming a three-sigma URE of 20 for all satellites and a maximum geometric effect of 2 X HDOP = 8). Many PPS users of GPS can navigate safely with a horizontal position accuracy of a kilometer or more, for example, ships in open ocean and aircraft enroute at altitude, and can therefore tolerate much higher values of HDOP (and PDOP). Therefore, "areas of reduced accuracy" is often a more appropriate term than "outage" for conditions of large PDOP, since the accuracy of the position solution may be reduced but still adequate for the mission requirements.
This suggests that the user should evaluate the performance standards with respect to the anticipated mission requirements. If the mission requirements are significantly different than the constraints used to develop the performance standards, an independent assessment of service coverage via computer simulations may be warranted. One method of determining the real-time effect of prevailing range errors and satellite geometry is calculation of a FOM described in paragraph 6.3 below. The user can then reduce the uncertainty associated with global averages and long-term statistics by comparing the current accuracy estimate to the mission accuracy requirements and thereby significantly improve the probability of success of the mission.
Most military GPS users will have to contend with the possibility of GPS "outages," due to hostile local conditions, for example, terrain masking or intentional jamming. One solution for some applications is an integrated navigation system. For example, if a GPS receiver is integrated with an inertial navigation unit, an intermittent GPS solution can be sufficient to maintain continuous high-accuracy positioning. For other applications, vertical aiding can be used as a pseudo-satellite to enhance availability, or differential GPS can be used to minimize range errors and correspondingly reduce sensitivity to DOP.
6.1.5.3 Four-Satellite Solution and Five-Degree Mask Angle
In effect, the performance standards are based on a "model" GPS receiver that calculates a four-satellite PVT solution and is constrained by a five-degree mask angle. In evaluating the impacts of these constraints, the user must consider the type of equipment that he is actually employing. Significant gains in service coverage can be achieved by the use of aiding, for example, from an altitude source or precision clock. Similarly, significant gains in service coverage can be expected if the satellite mask angle actually implemented by the receiver and GPS antenna is lower than five degrees. Correspondingly, a higher mask angle
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will reduce service coverage. In the event that the actual receiver differs significantly from the "model" receiver used to develop the standards, an independent assessment of service coverage may be advisable by means of computer simulations.
6.1.5.4 Integrity Checking
The service reliability concept defined here is closely related to the NATO concept of integrity. Consequently, user equipment that employs integrity checking algorithms may be able to detect the majority of "service failures" and continue to maintain a valid position solution by choosing a set of satellites which excludes the one experiencing the service failure. Various integrity monitoring algorithms have been developed by the civil aviation community which are well documented in open technical literature, and most receiver manufacturers are familiar with them. Most of these algorithms are based on the principle of a consistency check using additional range measurements and developing multiple solutions for comparison purposes (aiding measurements can be included). However, when such algorithms are employed, a minimum of five measurements are usually required, rather than the four required for a minimum position solution. Therefore, the overall system availability is likely to be determined by the availability of the integrity decision, rather than the availability of the navigation solution. Fortunately, the availability of an integrity decision based on PPS measurements is extremely high, since PPS is not subject to SA "noise" which can make SPS integrity decisions more difficult. Table 6-9 gives some results for the availability of an integrity decision from a recent study of a PPS integrity algorithm for military aviation which included pressure altimeter aiding. The results are based on a five-degree mask angle and a 556 metre position error threshold, suitable to protect the accuracy required for a nonprecision approach. The probability of detecting a service failure for this algorithm is 0.999, which when multiplied by the probability of occurrence of a service failure yields an overall level of integrity in excess of 0.99999.
Table 6-9. Availability of the Integrity Decision

Number of Satellites
24 23 22 21

Best Global Availability
N/A 99.998% 99.993% 99.94 %

Average Global Availability
100.000% 99.985% 99.866% 99.37 %

Worst Global Availability
N/A 99.965% 99.391% 97.55 %

Again, an assessment of the mission requirements is warranted to determine the integrity threshold, probability of residual "service failures", and duration of integrity "outages" that can be tolerated. For example, an application that involves safety of life may require that a position solution be declared invalid unless a positive confirmation of integrity is achieved. In contrast, a weapons delivery system might allow the position solution to be valid unless a negative assertion of integrity is determined, with the residual loss of integrity considered

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a minor overall detriment to weapon effectiveness when compared to the alternative loss of weapon availability.
6.1.5.5 Summary of the Commentary
If there are significant differences from the "model" receiver implied by the performance standards, different constraints applicable to the application, or different mission requirements, an independent assessment of the performance standards or similar parameters is probably warranted via computer simulation. In addition, real-time integrity checking and calculation of a figure-of-merit can significantly reduce the uncertainty associated with global averages and long-term statistics, and significantly improve the probability of success of a given military mission.
6.2 Satellite Selection Criteria
6.2.1 Introduction
The criteria used for satellite selection is a very important factor in GPS receiver design. Different receivers perform satellite selection using different algorithms. The important satellite criteria to be considered include:
a. Satellite health b. Geometric dilution of precision c. User range accuracy d. Elevation angle e. Availability of external aids.
6.2.2 Satellite Health
The NAV msg contains satellite health information for all the satellites in the GPS satellite constellation. Each satellite broadcasts health summaries for all (up to 32) GPS satellites, in page 25 of subframes 4 and 5. Each summary consists of 1 bit indicating the health of the NAV msg and 5 bits indicating the health of the satellite signals. (Refer to "Technical Characteristics of the Navstar GPS" or ICD-GPS-200PR for additional details). A satellite should never be used in a Nav-solution if its Navmessage is indicated to be unhealthy. If the NAV msg health is good, the five-bit signal status message should be compared against valid operating modes for the receiver to determine if the satellite can be used. For example, a P-code receiver could use a satellite broadcasting L1 only, if an ionospheric model can be used instead of dual frequency measurements to make the ionospheric corrections.
The NAV msg also contains a health message in subframe 1 which indicates the health of the broadcasting satellite. Since the data in subframes 4 and 5 are updated less frequently than subframe 1, subframe 1 may be used to indicate short-term health problems or may be updated before subframes 4 and 5. Therefore, after a satellite is acquired, the health data in subframe 1 should also be checked to deter mine if the satellite can be used.
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6.2.3 Geometric Dilution of Precision
As described previously in Chapters 2 and 3, GDOP is an important factor in determining the accuracy of the position (or time) solution. The combination of satellites which gives the lowest DOP value will provide the most accurate solution, assuming that all satellites have the same pseudorange error. Depending on the user mission, best PDOP, HDOP, or TDOP can be used as a satellite selection criterion.
6.2.4 User Range Accuracy
Each satellite broadcasts a user range accuracy (URA) value in subframe 1 of the NAV msg. URA is a prediction of the pseudorange accuracy obtainable from the satellite signal in space. URA is based on recent historical data and is therefore most accurate immediately following an upload. It does not include the UEE and therefore does not include ionospheric compensation error if the ionospheric model is used instead of dual frequency measurements. These additional errors should be added to URA for the best estimate of pseudorange accuracy, especially if the receiver is capable of performing dual frequency measurements on some satellites and must use an ionospheric model for others. (Refer to "Technical Characteristics of the Navstar GPS" or ICD-GPS-200PR for a more detailed explanation of URA.) URA can be used in conjunction with DOP to choose the best combination of satellites when the satellites have significantly different pseudorange errors. This is done by using URA as a weighting factor in the covariance matrix for user position and clock bias errors. Since URA is a prediction, it is not a guarantee of range accuracy, however, it can be used to help deselect satellites with known large pseudorange errors.
6.2.5 Satellite Elevation Angle
Selecting satellites by computing a minimum DOP will favor the use of satellites at low elevation angles. However, signals from satellites at a low elevation angle must travel a longer distance through the ionosphere and troposphere than signals from higher angles. They will therefore incur additional pseudorange error due to ionospheric and tropospheric delay. Many receivers will not use satellites below an arbitrary elevation angle. Five degrees is a typical lower limit. This also helps to reduce multipath problems.
6.2.6 External Aids
When an external aid is available to the GPS receiver, it can be incorporated into the satellite selection algorithm. It can be incorporated as a fixed mode of operation, an optional mode of operation when only three satellites are visible, or it can be treated as an additional "satellite" to be selected when the best combination of satellites includes the aid. Decision logic for the first two cases is relatively simple. If the aid is treated as an additional satellite, the expected error and geometry must be modelled and included in the satellite selection algorithm. For example, mean sea level (MSL) aiding can be considered to be equivalent to a satellite at the center of the earth with a UERE on the order of a typical satellite (6-7 metres). Other aiding schemes can be more complex, depending on the complexity of the integration, error model, and equivalent geometry. Barometric altimeter
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aiding should be treated with extra caution. Barometric altimeters are excellent devices for measuring pressure altitude, but pressure altitude can vary widely and non-linearly from geometric altitude. The resulting vertical errors should be modeled carefully since the errors can depend on meteorological conditions and vehicle dynamics. For additional discussion of GPS aids, refer to Chapter 7.
6.3 FIGURE OF MERIT (FOM)
A FOM is an indicator of receiver positioning or time accuracy which may be displayed to the operator or communicated to an integrated system. A FOM may be either a qualitative or quantitative measure, depending on the accuracy and integrity of the data used to calculate the FOM. In general, a FOM is not suitable for making integrity decisions where safety of life is concerned. However, a qualitative FOM may be perfectly suitable for integrity decisions regarding unmanned missions. (Refer to Chapter 12 for additional discussion of integrity.)
A FOM is typically calculated as the root-sum-square of the estimated errors contributing to the solution accuracy. Example criteria include:
a. GPS receiver state (e.g., carrier tracking, code tracking, acquisition) b. Carrier to noise ratio c. Satellite geometry (DOP value) d. Satellite range accuracy (URA value) e. Ionospheric measurement or modelling error f. Receiver aiding used g. Kalman filter error estimates.
The resultant FOM can be presented as a numerical value, for example from 1 to 9, where 1 indicates the best navigation performance. It can also be presented directly as an error estimate in metres, at a specified probability level, or even as a simple pass/fail indication. A time figure of merit (TFOM) can also be calculated to indicate the quality of the precise time information available from the GPS receiver via the PTTI interface (see paragraph 4.3.3). Table 6-10 gives the FOM and TFOM numerical assignments and equivalent estimated errors for the Rockwell-Collins family of receivers developed for the GPS JPO.
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Table 6-10. FOM/TFOM Numerical Values and Estimated Errors

FOM/TFOM

Estimated Position Error (EPE, metres)

Estimated Time Error (ETE, UTC)

0

Not Used

(Note 1)

1

EPE < 25

ETE £ 1 ns

2

25 < EPE £ 50

1 ns < ETE £ 10 ns

3

50 < EPE £ 75

10 ns < ETE £ 100 ns

4

75 < EPE £ 100

100 ns < ETE £ 1 µs

5

100 < EPE £ 200

1 µs < ETE £ 10 µs

6

200 < EPE £ 500

10 µs < ETE £ 100 µs

7

500 < EPE £ 1000

100 µs < ETE £ 1 ms

8

1000 < EPE £ 5000

1 ms < ETE £ 10 ms

9

EPE > 5000

10 ms < ETE, or Fault

10 to 14 15

Not Used Not Used

Not Used ETE Not Available

Note 1: External time source indicates proper/normal operation by TFOM = 0.

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CHAPTER 7: AIDING OPTIONS FOR A GPS RECEIVER
7.1 TYPES OF AIDING
Aiding a GPS receiver is done by incorporating inputs from external sources and is performed to enhance the following operations:
a. Acquisition of initial satellite signals, b. Translate the navigation solution to a position in the HV other than the GPS
antenna, c. Replace a satellite measurement in case of limited visibility or bad satellite
geometry, d. Maintain satellite tracking by increasing the tolerance of the GPS receiver to
interference, jamming or high HV dynamics.
Figure 7-1 illustrates some options. It should be noted that these are options and that not all GPS receivers presently have the capabilities described.

EXAMPLE DISCRETE INTERFACE

CDU

ARINC 429

DATA LOADER INS/AHRS

RS-422 ARINC 575

GPS RECEIVER

PTTI ALTIMETER

ARINC 572

MEMORY

EXAMPLE MUX BUS

MISSION COMPUTER HV CDU
HV DATA LOADER INS
AHRS DOPPLER
CADC OTHER SYSTEMS

Figure 7-1. Aiding Options for a GPS Receiver

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7.2 AIDING DURING INITIAL ACQUISITION
7.2.1 Position and Velocity Aiding
When a GPS receiver is first initialized for operation, approximate position and velocity of the receiver are required to minimize satellite acquisition time. The accuracy requirement of the U.S. DoD program for position is < 100 km of actual receiver location, and for velocity is < 100 m/s of actual receiver velocity, to ensure that satellite acquisition is within specification.
7.2.2 Time Aiding
Time aiding can be used during the initialization process, similar to position and velocity data. The time accuracy requirement is < 20 seconds relative to UTC. This is to ensure that satellite acquisition time is within specification.
Time aiding, if sufficiently accurate, can also be used to enable a direct P(Y) -code acquisition without first acquiring the C/A -code. This type of time aiding is relevant to HVs such as submarines where minimum exposure time of the GPS antenna on the ocean surface is of prime importance. An atomic time standard is one way to enable direct P(Y)-code acquisition.
7.2.3 Almanac Data
Normal satellite acquisition requires the availability of a current satellite almanac, stored in the receiver memory. If there are no significant changes in the satellite constellation, then the almanac is valid for several weeks.
If no stored or valid satellite almanac data are available, the GPS receiver starts to search the sky attempting to locate and lock onto any satellite in view. Depending on the receiver search strategy and on the actual satellite constellation, this process may take 15 -60 minutes. When one satellite is being tracked, the receiver can download and read the almanac information about all the other satellites in the constellation.
7.2.4 Effect On TTFF
Dependent on the type of integration (position, velocity and time) aiding data to the GPS receiver during the initialization process are provided as follows:
a. Manually by the operator via the GPS CDU or HV CDU, b. Automatically from INS/AHRS, PTTI or the HV mission computer (via 1553 -bus), c. Default by using the shut-down values stored in the receiver memory.
Initial acquisition performance can be expressed by the TTFF. In general terms, the TTFF is the time from when the receiver attempts to track the satellite signals until a navigation solution is determined. Knowing the position and velocity of the receiver, current time, and the positions of the satellites will all help to reduce the TTFF. Conversely, a lack o f
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reasonably accurate knowledge of any of these parameters will increase the TTFF. The amount of increase is dependent on the particular quantity and level of uncertainty.
7.3 AIDING TO TRANSLATE NAVIGATION SOLUTION
The navigation solution of an unaided GPS receiver is referenced to its antenna position. An aided GPS receiver can reference its navigation solution to another location. For example, the GPS navigation state can be resolved at the IMU instrument axes center in the case of an INS. To perform the calculations, the receiver needs to be aided with attitude information and a lever arm vector.
The attitude information in the form of roll, pitch and heading is provided in most cases by an INS or AHRS. A GPS receiver usually does all internal calculations in ECEF before carrying out any coordinate transformations. Using latitude and longitude in conjunction with attitude, the transformation between the GPS ECEF navigation frame and the HV body frame can be determined. Onboard ships, attitude aiding is also used to compensate for antenna motion and, together with water speed information, to do relative course and speed calculations.
A lever arm vector is provided to the GPS receiver as a vector between the GPS antenna and the HV reference point. If attitude aiding is removed from the GPS receiver, the navigation solution should revert back to the GPS antenna location. Often, more than one set of lever arm corrections may be stored in the GPS receiver. This is useful for installations having more than one INS aiding source or, in the case of big ships, where position and/or velocity information for different locations onboard may be of interest. However, only one attitude aiding source should be used by the GPS receiver at any one time. Hence the propagated navigation solution will only incorporate the one set of lever arm corrections applicable to the particular aiding source that is providing aiding data to the GPS receiver. Should the aiding source be changed, the lever arm corrections will change accordingly.
7.4 AIDING TO REPLACE A SATELLITE MEASUREMENT
During normal receiver operations, four satellite measurements are required inputs to solve the equations for position (Ux, Uy, Uz) and clock offset Dt. In case of limited satellite visibility or poor satellite geometry, one or more of the four satellite inputs may be replaced by inputs from an external aiding source.
When the GPS receiver is shipborne, or has barometric altimeter aiding or has a known height, then only three satellites are needed. Additional aiding by a precise clock can supplement the measurements in a two-satellite situation.
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7.4.1 Clock Aiding
A GPS receiver uses its own internal clock or may use a more accurate external clock as time reference. If only three (instead of four) satellites are available, then the GPS receiver can assume that its time reference is correct ( Dt = known) and treat the three available satellite range measurements as actual ranges instead of pseudoranges. In this case, the accuracy of the position derived from the pseudorange measurements will correspond to the equivalent time reference error.
If the GPS receiver clock or the external clock can be monitored during a previous period in which the receiver navigates with four satellites, then the clock phase bias and drift can be calculated. The resulting corrections for clock errors can be used to provide very accurate GPS time during a satellite outage and an accurate GPS position can be maintained for several minutes.
The method of using a clock instead of a satellite is not recommended as a permanent solution, but rather to help the GPS receiver operate during short periods when only a limited set of satellites is available. A GPS receiver should be capable of receiving (and providing) precise time via a dedicated PTTI interface or via the 1553 -bus.
7.4.2 Altitude Aiding
Similar to the clock aiding discussed in the previous paragraph, an airborne GPS receiver can use a barometric altimeter as aiding to replace a satellite measurement. Long-term altimeter errors are calibrated during periods of four satellite operation. Subsequently, when less than four satellites are being tracked, the calibrated baro-altimeter data are used as a known Uz-value in the 4 unknowns of (Ux, Uy, Uz) and Dt. Conceptually, the barometric altitude added to the earth radius provides a range measurement from a satellite with position at the center of the earth. An accurate GPS position can be maintained for as long as the estimated baro altitude errors are valid. Since the barometric altitude errors are generally slowly varying, both in time and distance, reasonable position accuracy can usually be maintained for 10-15 minutes, or within a radius of roughly 10 nmi of the position of the last 4 satellite solution. A gradual loss of position accuracy, especially in the vertical channel, can be expected. Depending on the algorithm used to compute altitude from pressure, the loss of accuracy may be hastened by altitude changes in a nonstandard atmosphere, particularly if no temperature compensation is used.
7.5 AIDING TO MAINTAIN SATELLITE TRACK
In normal receiver operation, the code and carrier tracking loops are both being tracked in phase lock. There is a symbiosis between the code and carrier tracking loops where each loop aids the other. In a high jamming environment, the receiver may lose its ability to track the carrier. Subsequent accelerations will cause the carrier frequency of the received GPS signal to vary due to a change in the Doppler shift. The Doppler shift of the frequency of the received carrier signal is proportional to the relative velocity of the receiver with respect to the satellite along the line-of-sight from the receiver to the satellite.
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Without some type of information to indicate this change in frequency, the center frequency of the receiver's replicated code signal will be different from the frequency of the actual received signal, which may then cause loss of code track as well. A receiver may be able to maintain code track in this case even while losing carrier track if it is aided with velocity. The primary function of aiding in this degraded mode of operation is to maintain code-loop tracking. The velocity data replaces the carrier tracking loop output as the source for code tracking loop aiding. Possible sources of velocity include INS, AHRS, and Doppler navigation systems. Requirements on the accuracy of the velocity will determine the allowable amounts of senescence, synchronization error, and aiding source absolute error under varying dynamic conditions. For example, higher dynamics will generally mean tighter restrictions on data senescence, which in turn can mean higher aiding rates. Typical accuracy requirements on the aiding velocity in order to maintain code track when the carrier is lost are on the order of 2-3 m/second.
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